Patent Description:
Gas turbine engines, such as turbofan engines, may be used for aircraft propulsion. Condensation trails (contrails) are typically formed when water molecules present in the exhaust gas from a turbofan engine becomes supersaturated after mixing with ambient air. These water molecules condense and quickly freeze into ice particles/crystals. Document <CIT> discloses a contrail suppression system.

Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.

The term "at least one of" in the context of, e.g., "at least one of A, B, and C" refers to only A, only B, only C, or any combination of A, B, and C.

Furthermore, the terms "upstream" and "downstream" refer to the relative direction with respect to fluid flow in a fluid pathway.

The term "turbomachine" or "turbomachinery" refers to a machine including one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output. The term "gas turbine engine" refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

The present disclosure is generally related to a contrail suppression system for a gas turbine engine. The formation of condensation trails (contrails) from gas turbine engine exhausts are suspected to have negative impact on climate conditions. Contrails are typically formed when water molecules present in exhaust gas becomes supersaturated after mixing with ambient air. These water molecules then condense and quickly freeze into ice particles/crystals. Contrails are known to form cirrus clouds and could sometimes persist for hours over several miles. Contrail formation is expected to be even more severe in hydrogen-fueled aircraft as compared to fossil-fueled aircraft since hydrogen combustion can release over <NUM> times more water.

This disclosure provides a contrail suppression system which incorporates a three-fluid shell and tube device such as a heat exchanger to suppress contrails from aircraft gas turbine engines, while minimizing exhaust gas pressure drop, thereby reducing any impact on overall engine thrust. The three-fluid heat exchanger works in suppressing contrails by first cooling the exhaust gas in a waste heat recovery (WHR) section and then dehumidifying the exhaust gas in a condenser section of the heat exchanger. The heat recovered in the WHR section is used to preheat compressed air from a low-pressure compressor to increase fuel efficiency while condensate water extracted from the condenser section of the heat exchanger may be injected either into the combustion chamber to help with nitrogen oxide (NOx) reduction, into a high-pressure turbine of the gas turbine engine for turbine blade cooling or may be stored in a storage tank or vessel aboard the aircraft or on the gas turbine engine.

In other embodiments, the heat exchanger for the contrail suppression system may be without the WHR section. It would only have the condenser section for dehumidifying the exhaust gas, which would eliminate the benefit of increased fuel efficiency and may result in lower effectiveness of contrail suppression.

Referring now to the drawings, <FIG> is a perspective view of an exemplary aircraft <NUM> that may incorporate at least one exemplary embodiment of the present disclosure. As shown in <FIG>, the aircraft <NUM> has a fuselage <NUM>, wings <NUM> attached to the fuselage <NUM>, and an empennage <NUM>. The aircraft <NUM> further includes a propulsion system <NUM> that produces a propulsive thrust to propel the aircraft <NUM> in flight, during taxiing operations, etc. Although the propulsion system <NUM> is shown attached to the wing(s) <NUM>, in other embodiments it may additionally or alternatively include one or more aspects coupled to other parts of the aircraft <NUM>, such as, for example, the empennage <NUM>, the fuselage <NUM>, or both. The propulsion system <NUM> includes at least one engine. In the exemplary embodiment shown, the aircraft <NUM> includes a pair of gas turbine engines <NUM>. Each gas turbine engine <NUM> is mounted to the aircraft <NUM> in an under-wing configuration. Each gas turbine engine <NUM> is capable of selectively generating a propulsive thrust for the aircraft <NUM>. The gas turbine engines <NUM> may be configured to burn various forms of fuel including, but not limited to unless otherwise provided, jet fuel/aviation turbine fuel, and hydrogen fuel.

<FIG> is a cross-sectional side view of a gas turbine engine <NUM> in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of <FIG>, the gas turbine engine <NUM> is a multi-spool, high-bypass turbofan jet engine, sometimes also referred to as a "turbofan engine. " As shown in <FIG>, the gas turbine engine <NUM> defines an axial direction A (extending parallel to a longitudinal centerline <NUM> provided for reference), a radial direction R, and a circumferential direction C extending about the longitudinal centerline <NUM>. In general, the gas turbine engine <NUM> includes a fan section <NUM> and a turbomachine <NUM> disposed downstream from the fan section <NUM>.

The exemplary turbomachine <NUM> depicted generally includes an engine casing <NUM> that defines an annular core inlet <NUM>. The engine casing <NUM> at least partially encases, in serial flow relationship, a compressor section including a booster or low-pressure (LP) compressor <NUM> and a high-pressure (HP) compressor <NUM>, a combustion section <NUM>, a turbine section including a high-pressure (HP) turbine <NUM> and a low-pressure (LP) turbine <NUM>, and a jet exhaust nozzle <NUM>.

A high-pressure (HP) shaft <NUM> drivingly connects the (HP) turbine <NUM> to the (HP) compressor <NUM>. A low-pressure (LP) shaft <NUM> that drivingly connects the (LP) turbine <NUM> to the (LP) compressor <NUM>. The compressor section, combustion section <NUM>, turbine section, and jet exhaust nozzle <NUM> together define a working gas flow path <NUM> through the gas turbine engine <NUM>.

For the embodiment depicted, the fan section <NUM> includes a fan <NUM> having a plurality of fan blades <NUM> coupled to a disk <NUM> in a spaced apart manner. As depicted, the fan blades <NUM> extend outwardly from disk <NUM> generally along the radial direction R. Each fan blade <NUM> is rotatable with the disk <NUM> about a pitch axis P by virtue of the fan blades <NUM> being operatively coupled to a suitable pitch change mechanism <NUM> configured to collectively vary the pitch of the fan blades <NUM>, e.g., in unison.

The gas turbine engine <NUM> further includes a power gear box <NUM>. The fan blades <NUM>, disk <NUM>, and pitch change mechanism <NUM> are together rotatable about the longitudinal centerline <NUM> by the (LP) shaft <NUM> across the power gear box <NUM>. The power gear box <NUM> includes a plurality of gears for adjusting a rotational speed of the fan <NUM> relative to a rotational speed of the (LP) shaft <NUM>, such that the fan <NUM> and the (LP) shaft <NUM> may rotate at more efficient relative speeds.

Referring still to the exemplary embodiment of <FIG>, the disk <NUM> is covered by rotatable front hub <NUM> of the fan section <NUM> (sometimes also referred to as a "spinner"). The front hub <NUM> is aerodynamically contoured to promote an airflow through the plurality of fan blades <NUM>. Additionally, the exemplary fan section <NUM> includes an annular fan casing or outer nacelle <NUM> that circumferentially surrounds the fan <NUM> and/or at least a portion of the turbomachine <NUM>. The nacelle <NUM> is supported relative to the turbomachine <NUM> by a plurality of circumferentially spaced struts or outlet guide vanes <NUM> in the embodiment depicted. Moreover, a downstream section <NUM> of the nacelle <NUM> extends over an outer portion of the turbomachine <NUM> to define a bypass airflow passage <NUM> therebetween.

It should be appreciated, however, that the exemplary gas turbine engine <NUM> depicted in <FIG> is provided by way of example only, and that in other exemplary embodiments, the gas turbine engine <NUM> may have other configurations. For example, although the gas turbine engine <NUM> depicted is configured as a ducted gas turbine engine (i.e., including the outer nacelle <NUM>), in other embodiments, the gas turbine engine <NUM> may be an unducted or non-ducted gas turbine engine (such that the fan <NUM> is an unducted fan, and the outlet guide vanes <NUM> are cantilevered from the engine casing <NUM>).

Additionally, or alternatively, although the gas turbine engine <NUM> depicted is configured as a geared gas turbine engine (i.e., including the power gear box <NUM>) and a variable pitch gas turbine engine (i.e., including a fan <NUM> configured as a variable pitch fan), in other embodiments, the gas turbine engine <NUM> may be configured as a direct drive gas turbine engine (such that the (LP) shaft <NUM> rotates at the same speed as the fan <NUM>), as a fixed pitch gas turbine engine (such that the fan <NUM> includes fan blades <NUM> that are not rotatable about a pitch axis P), or both. It should also be appreciated, that in still other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure may (as appropriate) be incorporated into, e.g., a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine.

During operation of the gas turbine engine <NUM>, a volume of air <NUM> enters the gas turbine engine <NUM> through an associated inlet <NUM> of the nacelle <NUM> and fan section <NUM>. As the volume of air <NUM> passes across the fan blades <NUM>, a first portion of air <NUM> is directed or routed into the bypass airflow passage <NUM> and a second portion of air <NUM> is directed or routed into the working gas flow path <NUM>, or more specifically into the (LP) compressor <NUM>. The ratio between the first portion of air <NUM> and the second portion of air <NUM> is commonly known as a bypass ratio. Pressure of the second portion of air <NUM> is then increased as it is routed through the (HP) compressor <NUM> and into the combustion section <NUM>, where it is mixed with fuel and burned to provide combustion gases <NUM>.

The combustion gases <NUM> are routed through the (HP) turbine <NUM> where a portion of thermal and/or kinetic energy from the combustion gases <NUM> is extracted via sequential stages of (HP) turbine stator vanes <NUM> that are coupled to a turbine casing and (HP) turbine rotor blades <NUM> that are coupled to the (HP) shaft <NUM>, thus causing the (HP) shaft <NUM> to rotate, thereby supporting operation of the (HP) compressor <NUM>. The combustion gases <NUM> are then routed through the (LP) turbine <NUM> where a second portion of thermal and kinetic energy is extracted from the combustion gases <NUM> via sequential stages of (LP) turbine stator vanes <NUM> that are coupled to a turbine casing and (LP) turbine rotor blades <NUM> that are coupled to the (LP) shaft <NUM>, thus causing the (LP) shaft <NUM> to rotate, thereby supporting operation of the (LP) compressor <NUM> and/or rotation of the fan <NUM>.

The combustion gases <NUM> are subsequently routed through the jet exhaust nozzle <NUM> of the turbomachine <NUM> to provide propulsive thrust. Simultaneously, the pressure of the first portion of air <NUM> is substantially increased as it is routed through the bypass airflow passage <NUM> before it is exhausted from a fan nozzle exhaust section <NUM> of the gas turbine engine <NUM>, also providing propulsive thrust. The (HP) turbine <NUM>, the (LP) turbine <NUM>, and the fan nozzle exhaust section <NUM> at least partially define a hot gas path <NUM> for routing the combustion gases <NUM> through the turbomachine <NUM>.

<FIG> is a schematic view of a portion of an exemplary gas turbine engine <NUM> shown in <FIG>. <FIG> depicts, in serial flow order, the fan section <NUM>, the (LP) compressor <NUM>, the (HP) compressor <NUM>, the combustion section <NUM>, the (HP) turbine <NUM> and the (LP) turbine <NUM>. <FIG> further depicts the nacelle <NUM> including the bypass airflow passage <NUM> and the jet exhaust nozzle <NUM>. In exemplary embodiments, as illustrated in <FIG>, the gas turbine engine <NUM> includes a contrail reduction or suppression system <NUM> herein denoted as system <NUM>. The system <NUM> includes at least one heat exchanger <NUM> in fluid communication with the jet exhaust nozzle <NUM>.

<FIG> provides and forward-looking view of an aft end of the jet exhaust nozzle <NUM> according to various embodiments of the present disclosure. As shown in <FIG>, the heat exchanger(s) <NUM> may be disposed inside the jet exhaust nozzle <NUM> at various locations. For example, the heat exchanger <NUM> may be positioned inside of the jet exhaust nozzle <NUM> or along an inner wall <NUM> of the jet exhaust nozzle <NUM>. In embodiments wherein there are multiple heat exchangers <NUM>, as show in <FIG>, the heat exchangers <NUM> may be circumferentially spaced along the inner wall <NUM>. In other embodiments, the heat exchanger <NUM> or one or more heat exchangers <NUM> may be positioned outside of the inner wall <NUM> of the jet exhaust nozzle <NUM>.

<FIG> provides a schematic view of an exemplary heat exchanger <NUM> according to various embodiments of the present disclosure. As shown in <FIG> and <FIG> collectively, the heat exchanger <NUM> includes a shell <NUM> having an exhaust gas inlet <NUM>, an exhaust gas outlet <NUM>, a condensate drain <NUM>. The heat exchanger <NUM> further includes a flow chamber <NUM> defined within the shell <NUM> and in fluid communication with the exhaust gas inlet <NUM>, the exhaust gas outlet <NUM> and the condensate drain <NUM>. The exhaust gas inlet <NUM> and the exhaust gas outlet <NUM> are in fluid communication with the jet exhaust nozzle <NUM>.

As further shown in <FIG> and <FIG> collectively, the heat exchanger <NUM> further includes a first tube bundle <NUM> disposed within the shell <NUM>, and more particularly, within the flow chamber <NUM> downstream from the exhaust gas inlet <NUM>. The first tube bundle <NUM> includes an inlet <NUM> and an outlet <NUM>. The inlet <NUM> of the first tube bundle <NUM> is fluidly connected to a first cooling medium source <NUM>.

The heat exchanger <NUM> further includes a second tube bundle <NUM> disposed within the shell <NUM>, and more particularly, within the flow chamber <NUM> downstream from the first tube bundle <NUM> and upstream from the exhaust gas outlet <NUM>. The second tube bundle <NUM> is fluidly isolated from the first tube bundle <NUM>. The second tube bundle <NUM> includes an inlet <NUM> and an outlet <NUM>. The inlet <NUM> of the second tube bundle <NUM> is fluidly connected to a second cooling medium source <NUM>. In particular embodiments, the second cooling medium source <NUM> includes the nacelle <NUM> and or the bypass airflow passage <NUM>.

In certain embodiments, the heat exchanger <NUM> includes one or more baffle plates <NUM>, or other means for flow smoothing and/or for generating turbulence of exhaust gases <NUM> flowing from the low-pressure turbine <NUM>. The baffle plate(s) <NUM> may have multiple perforations to enable uniform flow distribution as the exhaust gases <NUM> flow across them. The baffle plate(s) <NUM> is/are disposed within the shell <NUM>, and more specifically, within the flow chamber <NUM>. The baffle plate(s) <NUM> may increase heat transfer effectiveness of the heat exchanger <NUM> but may also result in slightly increased pressure drop of the exhaust gases in the shell <NUM>.

In exemplary embodiments, the first cooling medium source <NUM> includes any one or any combination of the low-pressure compressor <NUM>, the high-pressure compressor <NUM>, or any other cooling medium source, such as but not limited to, an auxiliary cooling or waste heat recovery system of the gas turbine engine <NUM>. In an exemplary embodiment, as shown in <FIG>, the first cooling medium source includes the low-pressure compressor <NUM> with the inlet <NUM> of the first tube bundle <NUM> fluidly coupled thereto.

In an exemplary embodiment, as shown in <FIG>, the outlet <NUM> of the first tube bundle <NUM> is fluidly coupled to the low-pressure compressor <NUM>. In other embodiments, the outlet <NUM> of the first tube bundle <NUM> may be fluidly coupled to any one or any combination of the low-pressure compressor <NUM>, the high-pressure compressor <NUM> or any other area of the gas turbine engine <NUM>.

In exemplary embodiments, as shown in <FIG> and <FIG> collectively, the second cooling medium source <NUM> for providing a second cooling medium <NUM> comprises the bypass airflow passage <NUM>. As shown in <FIG>, the outlet <NUM> of the second tube bundle <NUM> is fluidly coupled to the bypass airflow passage <NUM>. In addition or in the alternative, the outlet <NUM> of the second tube bundle <NUM> is in fluid communication with the jet exhaust nozzle <NUM>.

In exemplary embodiments, as shown in <FIG> and <FIG> collectively, the condensate drain <NUM> may be fluidly coupled to at least one of the combustion section <NUM>, the high-pressure turbine <NUM>, and/or the low-pressure turbine <NUM>. In addition, or in the alternative, the condensate drain <NUM> may be fluidly coupled to a storage tank <NUM>. The storage tank <NUM> may be stored aboard the aircraft <NUM> or mounted to the gas turbine engine <NUM>. The storage tank <NUM> may be removable or hard mounted.

<FIG> is a schematic cross-sectional view of an exemplary heat exchanger <NUM> according to exemplary embodiment of the present disclosure. In the exemplary embodiment shown in <FIG>, the heat exchanger <NUM> includes a shell <NUM> having an exhaust gas inlet <NUM>, an exhaust gas outlet <NUM>, a condensate drain <NUM> and a flow chamber <NUM> in fluid communication with the exhaust gas inlet <NUM>, the exhaust gas outlet <NUM> and the condensate drain <NUM>. The exhaust gas inlet <NUM> and exhaust gas outlet <NUM> are in fluid communication with the jet exhaust nozzle <NUM> of the gas turbine engine <NUM>. In certain embodiments, the heat exchanger <NUM> includes one or more baffle plates <NUM>, or other means for flow smoothing and/or for generating turbulence of exhaust gases <NUM>.

A tube bundle <NUM> is disposed within the flow chamber <NUM> downstream of the exhaust gas inlet <NUM> and upstream of the exhaust gas outlet <NUM>. The tube bundle <NUM> includes an inlet <NUM> and an outlet <NUM>. The inlet <NUM> and the outlet <NUM> of the tube bundle <NUM> are fluidly connected to the bypass airflow passage <NUM>. The condensate drain <NUM> is fluidly coupled to at least one of the combustion section <NUM>, the high-pressure turbine <NUM>, the low-pressure turbine <NUM>, and a storage tank <NUM>. The storage tank may be stored abord the gas turbine engine <NUM> or the aircraft <NUM>.

As shown in <FIG>, the heat exchanger(s) <NUM> may be disposed inside the jet exhaust nozzle <NUM> at various locations. For example, the heat exchanger <NUM> may be positioned inside of the jet exhaust nozzle <NUM> or along an inner wall <NUM> of the jet exhaust nozzle <NUM>. In embodiments wherein there are multiple heat exchangers <NUM>, as show in <FIG>, the heat exchangers <NUM> may be circumferentially spaced along the inner wall <NUM>. In other embodiments, the heat exchanger <NUM> or one or more heat exchangers <NUM> may be positioned outside of the inner wall <NUM> of the jet exhaust nozzle <NUM>. In exemplary embodiments the system <NUM> may include both embodiments of heat exchanger <NUM> and <NUM>.

In operation, the exhaust gases <NUM> flow from the low-pressure turbine <NUM> into the jet exhaust nozzle <NUM>. At least a portion of the exhaust gases <NUM> are directed into the exhaust gas inlet <NUM> and into the flow chamber <NUM> of the heat exchanger <NUM>. The first tube bundle <NUM> is charged with the first cooling medium <NUM> from the first cooling medium source <NUM>. The exhaust gases <NUM> flow across the first tube bundle <NUM> to reduce the temperature of the exhaust gases <NUM> upstream from the second tube bundle <NUM>. The second tube bundle <NUM> is charged with the second cooling medium <NUM> from the second cooling medium source <NUM>. In other embodiments, fluid pump(s) (not shown) may be fluidly connected between the first cooling medium source <NUM> and the first tube bundle inlet <NUM>, and/or between the second cooling medium source <NUM> and second tube bundle inlet <NUM>, to boost the flow pressure of the first and/or second cooling media <NUM>, <NUM>.

The cooled exhaust gases <NUM> flow across the second tube bundle <NUM>, thereby causing moisture from the exhaust gases to condense into water droplets. The water droplets flow out of/are removed from the shell <NUM> via the condensate drain <NUM>. Condensate <NUM> may be collected in the storage tank <NUM>, injected into the combustion section <NUM> to help reduce nitrogen oxide (NOx) emissions, injected into the high-pressure turbine <NUM> and/or the low-pressure turbine <NUM> for cooling purposes, or vented to atmosphere at an appropriate altitude or speed so as to not contribute to contrail formation. The storage tank <NUM> may be connected to a water treatment system (not shown) aboard the aircraft <NUM> or mounted to the gas turbine engine <NUM>.

Claim 1:
A contrail suppression system (<NUM>), comprising:
a shell (<NUM>) having an exhaust gas inlet (<NUM>), an exhaust gas outlet (<NUM>), a condensate drain (<NUM>) and a flow chamber (<NUM>) in fluid communication with the exhaust gas inlet (<NUM>), the exhaust gas outlet (<NUM>) and the condensate drain (<NUM>), wherein the exhaust gas inlet (<NUM>) is in fluid communication with a jet exhaust nozzle (<NUM>) of a gas turbine engine (<NUM>);
a first tube bundle (<NUM>) disposed within the flow chamber (<NUM>) downstream from the exhaust gas inlet (<NUM>), wherein the first tube bundle (<NUM>) includes an inlet (<NUM>) and an outlet (<NUM>), wherein the inlet of the first tube bundle (<NUM>) is fluidly connected to a first cooling medium source (<NUM>); the system being characterised in that it comprises
a second tube bundle (<NUM>) disposed within the flow chamber (<NUM>) downstream from the first tube bundle (<NUM>) and upstream from the exhaust gas outlet (<NUM>), wherein the second tube bundle (<NUM>) includes an inlet (<NUM>) and an outlet (<NUM>), wherein the inlet of the second tube bundle (<NUM>) is fluidly connected to a second cooling medium source (<NUM>).