Patent Description:
It is generally known in the art to power aircraft gas turbine engines with gases expelled from combustion chambers. In the gas turbine engine, a fuel is combusted in an oxygen rich environment. The fuel may be any appropriate fuel such as a liquid or gas. Exemplary fuels include hydrocarbons (for example methane or kerosene) or hydrogen. Generally, these combustion systems may emit undesirable compounds such as nitrous oxide compounds (NOx) and carbon containing compounds. It is generally desirable to decrease various emissions as much as possible so that selected compounds may not enter the atmosphere. In particular, it has become desirable to reduce NOx emissions to a substantially low amount. There is a need in the art, therefore, for improved systems and methods for reducing NOx emissions from aircraft gas turbine engines.

<CIT> discloses a prior art catalyst system. The catalyst system includes a catalyst bed that comprises a plurality of catalyst segments arranged such that an exhaust gas flow passes along a longitudinal axis of the catalyst system through the plurality of catalyst segments.

According to an aspect of the present disclosure, there is provided a gas turbine engine for an aircraft as set forth in claim <NUM>.

In any of the aspects or embodiments described above and herein, the gas turbine engine further includes a fixed structure surrounding at least a portion of the turbine section. The exhaust section further includes a diffuser nozzle mounted to the fixed structure downstream of the turbine section and configured to receive the exhaust gas stream from the turbine section. The monolithic catalyst structure is located within the diffuser nozzle.

In any of the aspects or embodiments described above and herein, the gas turbine engine may be a turboprop or a turboshaft gas turbine engine.

In any of the aspects or embodiments described above and herein, the monolithic catalyst structure may include a plurality of cells defining a respective plurality of channels extending therethrough.

In any of the aspects or embodiments described above and herein, the plurality of cells may include a catalytic washcoat.

In any of the aspects or embodiments described above and herein, the reducing agent may be an ammonia-based reducing agent.

In any of the aspects or embodiments described above and herein, the gas turbine engine further may include a nacelle defining an exterior housing of the gas turbine engine. The diffuser nozzle may be entirely located within the nacelle.

In any of the aspects or embodiments described above and herein, the diffuser nozzle may include a housing disposed about a nozzle axis and extending between a first nozzle end and a second nozzle end, the monolithic catalyst structure located in a first axial portion of the housing. A first diameter of the housing in the first axial portion may be greater than a second diameter of a nozzle inlet of the diffuser nozzle and a third diameter of a nozzle outlet of the diffuser nozzle.

According to another aspect of the present disclosure, there is provided a method for treating exhaust gases from a gas turbine engine for an aircraft as set forth in claim <NUM>.

In any of the aspects or embodiments described above and herein, the exhaust section may further include a diffuser nozzle configured to receive the exhaust gas stream from the turbine section and the monolithic catalyst structure may be located within the diffuser nozzle.

In any of the aspects or embodiments described above and herein, the step of injecting the reducing agent into the core flow path of the gas turbine engine may include injecting an ammonia-based reducing agent into the core flow path of the gas turbine engine.

In any of the aspects or embodiments described above and herein, the diffuser nozzle may be located entirely within a nacelle defining an exterior housing of the gas turbine engine.

In any of the aspects or embodiments described above and herein, the method may further include diffusing exhaust gas stream with the diffuser nozzle at a first axial location within the diffuser nozzle and subsequently concentrating the exhaust gas stream with the diffuser nozzle at a second axial location within the diffuser nozzle which is different than the first axial location.

The present disclosure, and all its aspects, embodiments and advantages associated therewith will become more readily apparent in view of the detailed description provided below, including the accompanying drawings.

<FIG> illustrates a gas turbine engine <NUM> of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication along a core flow path <NUM> through an air inlet <NUM>, a compressor section <NUM> for pressurizing the air from the air inlet <NUM>, a combustor <NUM> in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, a turbine section <NUM> for extracting energy from the combustion gases, and an exhaust section <NUM> through which the combustion exhaust gases exit the gas turbine engine <NUM>.

The gas turbine engine <NUM> of <FIG> generally includes a high-pressure spool <NUM>, a low-pressure spool <NUM>, and a power spool <NUM> mounted for rotation about an axial centerline <NUM> (e.g., a rotational axis) of the gas turbine engine <NUM>. The high-pressure spool <NUM> generally includes a high-pressure shaft <NUM> that interconnects a high-pressure compressor <NUM> and a high-pressure turbine <NUM>. The low-pressure spool <NUM> generally includes a low-pressure shaft <NUM> that interconnects a low-pressure compressor <NUM> and a low-pressure turbine <NUM>. The power spool <NUM> generally includes a drive output shaft <NUM> in rotational communication with a power turbine <NUM> having a forward end configured to drive a rotatable load <NUM>. The rotatable load <NUM> can, for instance, take the form of a propeller. In alternative embodiments, the gas turbine engine <NUM> may be configured such that the rotatable load <NUM> may include a rotor, such as a helicopter main rotor, driven by the drive output shaft <NUM>. The drive output shaft <NUM> may be connected to the rotatable load <NUM> through a gear assembly <NUM> to drive the rotatable load <NUM> at a lower speed than the power spool <NUM>. It should be understood that "low pressure" and "high pressure," or variations thereof, as used herein, are relative terms indicating that the high pressure is greater than the low pressure. The high-pressure shaft <NUM>, the low-pressure shaft <NUM>, and the drive output shaft <NUM> may be concentric about the axial centerline <NUM>. The gas turbine engine <NUM> of <FIG> further includes a nacelle <NUM> defining an exterior housing of the gas turbine engine <NUM>. The gas turbine engine <NUM> of <FIG> further includes an aircraft wing <NUM> mounted to and extending outward from the nacelle <NUM>.

The gas turbine engine <NUM> of <FIG> may be configured, for example, as a turboprop or a turboshaft gas turbine engine. It should be understood that the concepts described herein are not limited to use with turboprops as the teachings may be applied to other types of gas turbine engines such as turbofan gas turbine engines as well as those gas turbine engines including single-spool or two-spool architectures.

In some embodiments, the gas turbine engine <NUM> may include a diffuser nozzle <NUM> in the exhaust section <NUM> of the gas turbine engine <NUM>. The diffuser nozzle <NUM> is configured to direct combustion exhaust gases and to decelerate the combustion exhaust gases for post-combustion treatment to reduce or otherwise mitigate the emission of air pollutants from the gas turbine engine <NUM> including, but not limited to, nitrogen oxides (NOx). The gas turbine engine <NUM> may include a fixed structure <NUM> such as a casing or cowl surrounding at least a portion of the turbine section <NUM>. The diffuser nozzle <NUM> may be mounted to the fixed structure <NUM> axially downstream of the turbine section <NUM>. As shown in <FIG>, at least a portion of the diffuser nozzle <NUM> may be located within the nacelle <NUM> surrounding the gas turbine engine <NUM>. In some embodiments, the diffuser nozzle <NUM> may be entirely disposed within the nacelle <NUM>. Aspects of the present disclosure diffuser nozzle <NUM> maybe particularly relevant in for the treatment of combustion exhaust gases turboprop or turboshaft gas turbine engines, as the combustion exhaust gases may not be used to generate a substantial amount of thrust for an associated aircraft. Accordingly, treatment of the combustion exhaust gases to remove NOx may provide a valuable means of controlling emissions of air pollutants without restricting the operational capacity of the associated gas turbine engine. However, it should be understood that aspects of the present disclosure may also be relevant to other types of aircraft gas turbine engines such as turbofan and turbojet gas turbine engines.

Referring to <FIG>, the diffuser nozzle <NUM> includes a housing <NUM> disposed about a nozzle axis <NUM> and extending between a first nozzle end <NUM> and a second nozzle end <NUM>. The nozzle axis <NUM> may or may not be colinear with the axial centerline <NUM> of the gas turbine engine <NUM>. The housing <NUM> includes a nozzle inlet <NUM> located at the first nozzle end <NUM> and a nozzle outlet <NUM> located at the second nozzle end <NUM>. Combustion exhaust gases (schematically illustrated in <FIG> as exhaust gas stream <NUM>) are directed from the turbine section <NUM> to the nozzle inlet <NUM> and then through the diffuser nozzle <NUM> in a direction from the nozzle inlet <NUM> to the nozzle outlet <NUM>. The housing <NUM> radially surrounds and defines a nozzle duct <NUM> of the diffuser nozzle <NUM> extending from the nozzle inlet <NUM> to the nozzle outlet <NUM> and including the nozzle inlet <NUM> and the nozzle outlet <NUM>.

In an upstream-to-downstream direction as shown in <FIG>, the diffuser nozzle <NUM> may include the nozzle inlet <NUM>, a diffusing axial portion <NUM>, a treatment axial portion <NUM>, a concentrating axial portion <NUM>, and the nozzle outlet <NUM>. The treatment axial portion <NUM> includes a maximum cross-sectional area of the nozzle duct <NUM>. A diameter D1 of the housing <NUM> along the treatment axial portion <NUM> is greater than a diameter D2 of the housing <NUM> at the nozzle inlet <NUM> and a diameter D3 of the housing <NUM> at the nozzle outlet <NUM>. Within the diffusing axial portion <NUM>, the duct cross-sectional area of each duct section <NUM> gradually increases until reaching a maximum duct cross-sectional area within the treatment axial portion <NUM>. Within the concentrating axial portion <NUM>, the duct cross-sectional area of each duct section <NUM> gradually decreases from the maximum duct cross-sectional area of the treatment axial portion <NUM> until reaching the nozzle outlet <NUM>.

The present disclosure exhaust section <NUM> of the gas turbine engine <NUM> includes a monolithic catalyst structure <NUM> configured to treat air pollutants such as NOx from the exhaust gas stream <NUM> as the exhaust gas stream <NUM> passes through the monolithic catalyst structure <NUM>. In some embodiments, the monolithic catalyst structure <NUM> may be part of and located within the diffuser nozzle <NUM>, as shown in <FIG>. For example, the monolithic catalyst structure <NUM> may be located within the treatment axial portion <NUM> of the diffuser nozzle <NUM>. However, the present disclosure is not limited to the inclusion of the monolithic catalyst structure <NUM> in the diffuser nozzle <NUM> and the monolithic catalyst structure <NUM> may be included in the exhaust section <NUM> within the diffuser nozzle <NUM> of <FIG>. <FIG> illustrates a cross-sectional view of the treatment axial portion <NUM> of the diffuser nozzle <NUM> showing the monolithic catalyst structure <NUM>. The monolithic catalyst structure <NUM> may be disposed across all or substantially all of the duct cross-sectional area within the treatment axial portion <NUM> of the nozzle duct <NUM>.

The monolithic catalyst structure <NUM> may be made from a ceramic material forming a plurality of substrate cells <NUM>. The plurality of substrate cells <NUM> define a respective plurality of channels <NUM> extending through the monolithic catalyst structure <NUM> in a generally axial direction. The monolithic catalyst structure <NUM> includes a catalyst washcoat applied to the surfaces of the substrate cells <NUM>. The catalyst washcoat serves as a carrier for a catalyst such as, but not limited to, platinum, palladium, rhodium, and/or zeolite, which catalyst is used to stimulate and accelerate a NOx reduction chemical reaction of the monolithic catalyst structure <NUM>. As shown in <FIG>, the substrate cells of the plurality of substrate cells <NUM> may have a generally square cross-sectional shape. However, plurality of substrate cells <NUM> can have other cross-sectional shapes such as hexagons, circles, etc. Density of the plurality of substrate cells <NUM> may vary widely depending on the particular application of the diffuser nozzle <NUM> as well as other considerations such as the acceptable pressure loss through the diffuser nozzle <NUM> and the emissions reduction requirements for the diffuser nozzle <NUM>. Accordingly, the density of the plurality of substrate cells <NUM> may range from approximately <NUM> to <NUM> cells per square inch ( <NUM> to <NUM> cells per square cm). The plurality of substrate cells <NUM> may have an average wall thickness in a range of approximately <NUM> to <NUM> inches (i.e., <NUM>-<NUM>). The catalyst washcoat applied to the plurality of substrate cells <NUM> may have an average thickness in a range of approximately <NUM> to <NUM> inches (i.e., <NUM>-<NUM>).

Combustion exhaust gases of the exhaust gas stream <NUM> passing through the diffuser nozzle <NUM> are directed through the monolithic catalyst structure <NUM> where the exhaust gas stream <NUM> is treated through chemically interaction with the catalyst washcoat applied to the surfaces of the plurality of substrate cells <NUM>. Diffusion of the exhaust gas stream <NUM> within the diffusing axial portion <NUM> of the diffuser nozzle <NUM> from the nozzle inlet <NUM> to the maximum cross-sectional area provided by the treatment axial portion <NUM> provides for an increase in the static pressure of the exhaust gas stream <NUM> and a reduction in velocity of the exhaust gas stream <NUM>, within the treatment axial portion <NUM> of the diffuser nozzle <NUM>. By reducing the velocity of the exhaust gas stream <NUM> within the treatment axial portion <NUM>, the length of time for chemical interaction between the exhaust gas stream <NUM> and the monolithic catalyst structure <NUM> may be increased, thereby improving post-combustion treatment of the exhaust gas stream <NUM>. Moreover, the pressure losses of the exhaust gas stream <NUM> passing through the monolithic catalyst structure <NUM> are reduced. Concentration of the exhaust gas stream <NUM> within the concentrating axial portion <NUM> of the diffuser nozzle <NUM> from the treatment axial portion <NUM> to the nozzle outlet <NUM> provides for a decrease in the static pressure of the exhaust gas stream <NUM> and an increase in velocity of the exhaust gas stream <NUM> which exits the nozzle outlet <NUM> of the diffuser nozzle <NUM>, thereby providing some amount of usable thrust. Accordingly, the configuration of the diffuser nozzle <NUM> may provide a tradeoff whereby an axial length of the diffuser nozzle <NUM> may be decreased while a diameter of the diffuser nozzle <NUM> (e.g., the diameter D1 of the housing <NUM> along the treatment axial portion <NUM>) may be increased, while maintaining the post-combustion treatment capability of the diffuser nozzle <NUM> with respect to the exhaust gas stream <NUM>. The diffuser nozzle <NUM> may, therefore, provide a form factor which can more readily be incorporated into gas turbine engines such as the gas turbine engine <NUM> and, for example, be retained within a nacelle for the respective gas turbine engine.

Referring to <FIG>, <FIG>, the present disclosure gas turbine engine <NUM> further includes a reducing agent injection system <NUM> configured to inject a reducing agent (schematically illustrated in <FIG> as reducing agent <NUM>) into the core flow path <NUM> of the gas turbine engine <NUM>. The post-combustion introduction of a reducing agent into the core flow path <NUM> may further reduce exhaust emissions of NOx which may be found in the exhaust gas stream <NUM>. Reduction of NOx emissions may be accomplished through one or both of selective catalytic reduction (SCR) and/or selective non-catalytic reduction (SNCR) chemical reactions, as will be discussed in greater detail. The reducing agent may typically be an ammonia-based fluid including, for example, anhydrous ammonia (NH<NUM>) or aqueous ammonia (NH<NUM>OH), however, the present disclosure is not limited to any particular reducing agent.

In some embodiments, the reducing agent injection system <NUM> may be configured to implement an SCR process to treat NOx found within the exhaust gas stream <NUM> along the core flow path <NUM>. As shown in <FIG>, the reducing agent injection system <NUM> may be positioned within the gas turbine engine <NUM> to inject the reducing agent <NUM> into the core flow path <NUM> downstream of the turbine section <NUM> (e.g., downstream of a final turbine stage) for mixing with the exhaust gas stream <NUM>. For SCR, the NOx reduction reaction takes place as the mixed exhaust gas stream <NUM> and the reducing agent <NUM> pass through the monolithic catalyst structure <NUM> of the diffuser nozzle <NUM>. The chemical reactions for the SCR process may be generalized by the following equations [<NUM>], [<NUM>], [<NUM>] which convert the NOx constituents, nitric oxide (NO) and nitrogen dioxide (NO<NUM>), to nitrogen (N<NUM>) and water (H<NUM>O):.

<NUM>NO + <NUM>NH<NUM> + O<NUM> → <NUM>N<NUM> + <NUM>H<NUM>O     [<NUM>].

NO + NO<NUM> + <NUM>NH<NUM> → <NUM>N<NUM> + <NUM>H<NUM>O     [<NUM>].

<NUM>NO<NUM> + <NUM>NH<NUM> → <NUM>N<NUM> + <NUM>H<NUM>O     [<NUM>].

The SCR process uses the catalyst of the monolithic catalyst structure <NUM> to reduce the necessary activation energy for the above-noted SCR reduction reactions. Accordingly, the SCR process can eliminate as much as <NUM> percent of NOx within the exhaust gas stream <NUM>, with a sufficiently large and appropriately sized monolithic catalyst structure <NUM>.

The reducing agent injection system <NUM> is configured to implement a SCR process and a SNCR process to treat NOx found within the exhaust gas stream <NUM> along the core flow path <NUM>. As shown in <FIG>, the reducing agent injection system <NUM> is positioned within the gas turbine engine <NUM> to inject the reducing agent <NUM> into the core flow path <NUM> downstream of the combustor <NUM> but upstream of the turbine section <NUM> (e.g., upstream of the high-pressure turbine <NUM>), for mixing with the exhaust gas stream <NUM>. The SNCR process does not require a catalyst, but may only occur at elevated temperatures such as, for example, between <NUM>,<NUM> °F and <NUM>,<NUM> °F and, preferably, greater than approximately <NUM>,<NUM> ° F (respectively <NUM>, <NUM> and <NUM>).

Accordingly, the SNCR process may only occur in portions of the core flow path <NUM> of the gas turbine engine <NUM> with sufficiently high temperatures. The chemical reaction for the SNCR process may be represented by the following equation [<NUM>] which converts the NOx constituent, nitric oxide (NO), to nitrogen (N<NUM>) and water (H<NUM>O):.

Because of the very short time that the mixed exhaust gas stream <NUM> and reducing agent <NUM> may spend in the temperature range necessary for the SNCR process to occur, the SNCR process may result in a NOx reduction of less than <NUM> percent in aircraft gas turbine engine applications. Accordingly, the possible increased cost and complexity of positioning the reducing agent injection system <NUM> upstream of the high-pressure turbine <NUM> (in contrast to placement of the reducing agent injection system <NUM> downstream of the turbine section <NUM>) may be considered with the expected NOx reduction provided by the associated SNCR process, for the particular NOx emissions reduction application.

In some embodiments, the reducing agent injection system <NUM> may include an annular manifold <NUM>, as shown in <FIG>, which extends about the axial centerline <NUM> of the gas turbine engine <NUM>. The reducing agent injection system <NUM> may further include a plurality of nozzles <NUM> circumferentially spaced about the manifold <NUM> and configured to direct the reducing agent <NUM> into the exhaust gas stream <NUM> transiting the core flow path <NUM>. It should be understood, however, that the present disclosure reducing agent injection system <NUM> is not limited to the above-described configuration and other means for introducing the reducing agent to the exhaust gas stream <NUM> may be considered.

Claim 1:
A gas turbine engine (<NUM>) for an aircraft, the gas turbine engine (<NUM>) comprising:
a turbine section (<NUM>);
an exhaust section (<NUM>) configured to receive an exhaust gas stream (<NUM>) from the turbine section (<NUM>), the exhaust section (<NUM>) including a monolithic catalyst structure (<NUM>); and
a reducing agent injection system (<NUM>) configured to inject a reducing agent (<NUM>) into a core flow path (<NUM>) of the gas turbine engine (<NUM>), wherein the reducing agent injection system (<NUM>) is located upstream of the turbine section (<NUM>).