Patent Description:
Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section further drives the compressor section to rotate. In some configurations, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine.

However, such gas turbine engines generate waste products, such as CO<NUM>. It may be advantageous to have aircraft propulsion systems that do not generate waste byproducts or generate less waste, including CO<NUM>.

A prior art aircraft propulsion system having the features of the preamble of claim <NUM> is disclosed in <CIT>. <CIT> discloses another prior art aircraft propulsion system.

According to an aspect of the present invention, an aircraft propulsion system is provided in accordance with claim <NUM>.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the supercritical fluid is CO<NUM> that is passed through the turbine, the cooler heat exchanger, the compressor, and the recovery heat exchanger.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include a gear system coupled to the shaft between the turbine and the fan.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the closed loop-supercritical fluid system further comprises a recuperator heat exchanger arranged between the turbine and the cooler heat exchanger along the closed-loop flow path.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the recuperator heat exchanger is a supercritical fluid-to-supercritical fluid heat exchanger.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include a cryogenic fuel tank configured to supply fuel to the burner through a fuel flow path, wherein the cooler heat exchanger is a fuel-to-supercritical fluid heat exchanger.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include a turbo expander operably coupled to the shaft and arranged between the cryogenic fuel tank and the burner.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the cooler heat exchanger is an air-to-supercritical fluid heat exchanger.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the cooler heat exchanger is arranged within a bypass duct downstream of the fan.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include a bypass flow path defining a flow path of air that passes through an air inlet, through the fan coupled to a shaft, through a bypass duct, and out a bypass nozzle, a hot gas flow path defining a flow path of air that passes through the air inlet, through the fan, into the burner for combustion with fuel to generate combusted gas, and through the recovery heat exchanger, and out the exhaust nozzle, and a closed-loop flow path defining a closed-loop flow path of a supercritical fluid that passes through the turbine operably coupled to the shaft to drive rotation of the shaft, the cooler heat exchanger, the compressor coupled to the shaft, into the recovery heat exchanger, and back to the turbine. The turbine of the closed-loop flow path drives rotation of the shaft, the compressor, and the fan.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include a fuel flow path defining a flow path from a fuel tank to the burner of the hot gas flow path.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include a turbo expander coupled to the shaft and configured to expand a fuel prior to injection into the burner.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include a first gear system coupled to the shaft between the turbine and the turbo expander and a second gear system coupled to the shaft between the turbo expander and the fan.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that a fuel in the fuel flow path is cryogenic fuel.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the cryogenic fuel is hydrogen.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the cooler heat exchanger of the closed-loop flow path receives fuel to form a fuel-to-supercritical fluid heat exchanger.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft propulsion systems may include that the hot gas flow path further comprises a blower arranged between the fan and the burner to increase a speed of the air passing through the hot gas flow path.

The foregoing features and elements may be executed or utilized in various combinations without exclusivity, unless expressly indicated otherwise.

As illustratively shown, the gas turbine engine <NUM> is configured as a two-spool turbofan that has a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. The illustrative gas turbine engine <NUM> is merely for example and discussion purposes, and those of skill in the art will appreciate that alternative configurations of gas turbine engines may employ embodiments of the present disclosure. The fan section <NUM> includes a fan <NUM> that is configured to drive air along a bypass flow path B in a bypass duct defined within a nacelle <NUM>. The fan <NUM> is also configured to drive air along a core flow path C for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines.

In this two-spool configuration, the gas turbine engine <NUM> includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via one or more bearing systems <NUM>. It should be understood that various bearing systems <NUM> at various locations may be provided, and the location of bearing systems <NUM> may be varied as appropriate to a particular application and/or engine configuration.

The low speed spool <NUM> includes an inner shaft <NUM> that interconnects the fan <NUM> of the fan section <NUM>, a first (or low) pressure compressor <NUM>, and a first (or low) pressure turbine <NUM>. The inner shaft <NUM> is connected to the fan <NUM> through a speed change mechanism, which, in this illustrative gas turbine engine <NUM>, is as a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. A combustor <NUM> is arranged in the combustor section <NUM> between the high pressure compressor <NUM> and the high pressure turbine <NUM>. A mid-turbine frame <NUM> of the engine static structure <NUM> is arranged between the high pressure turbine <NUM> and the low pressure turbine <NUM>. The mid-turbine frame <NUM> may be configured to support one or more of the bearing systems <NUM> in the turbine section <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine central longitudinal axis A which is collinear with their longitudinal axes.

The core airflow through core airflow path C is compressed by the low pressure compressor <NUM> then the high pressure compressor <NUM>, mixed and burned with fuel in the combustor <NUM>, then expanded over the high pressure turbine <NUM> and low pressure turbine <NUM>. The mid-turbine frame <NUM> includes airfoils <NUM> (e.g., vanes) which are arranged in the core airflow path C. The turbines <NUM>, <NUM> rotationally drive the respective low speed spool <NUM> and high speed spool <NUM> in response to the expansion of the core airflow. It will be appreciated that each of the positions of the fan section <NUM>, the compressor section <NUM>, the combustor section <NUM>, the turbine section <NUM>, and geared architecture <NUM> or other fan drive gear system may be varied. For example, in some embodiments, the geared architecture <NUM> may be located aft of the combustor section <NUM> or even aft of the turbine section <NUM>, and the fan section <NUM> may be positioned forward or aft of the location of the geared architecture <NUM>.

The gas turbine engine <NUM> in one example is a high-bypass geared aircraft engine. In some such examples, the gas turbine engine <NUM> has a bypass ratio that is greater than about six, with an example embodiment being greater than about ten. In some embodiments, the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about five. In one non-limiting embodiment, the bypass ratio of the gas turbine engine <NUM> is greater than about ten, a diameter of the fan <NUM> is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about five. The low pressure turbine <NUM> pressure ratio is pressure measured prior to inlet of low pressure turbine <NUM> as related to the pressure at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle. In some embodiments, the geared architecture <NUM> may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>. It should be understood, however, that the above parameters are only for example and explanatory of one non-limiting embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including turbojets or direct drive turbofans or turboshafts.

The fan section <NUM> of the gas turbine engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>,<NUM> meters). "Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]^<NUM> (where °R = <NUM>/<NUM> x K).

As noted above, conventional gas turbine engines, such as shown in <FIG>, generate waste heat and chemicals. However, it is difficult to achieve zero carbon emission flight. Embodiments of the present disclosure are directed to propulsion systems that have reduced or no emissions and achieve a high thermal efficiency. Additionally, low altitude thrust lapse is a problem caused by a density drop that air experiences in a traditional jet engine. However, altitude-invariant engines do not currently exist. Embodiments, of the present disclosure are directed to closed loop CO<NUM> propulsion systems that can provide a system of effectively no lapse due to altitude.

Embodiments of the present disclosure are directed to supercritical fluid closed-loop driven turbofan arrangements to generate propulsion for an aircraft. In some embodiments, the supercritical fluid may be CO<NUM>. In operation, air will enter the system and a blower fan will be used as a booster for air that is heated by a burner (e.g., a convention combustion chamber). A duct behind the burner allows the air to even out or to become uniform (or more uniform), prior to entering an annular heat exchanger that will heat the fluid of a closed-loop supercritical fluid system. For example, the pressure and temperature profiles may become more uniform (e.g., inner diameter to outer diameter and circumferentially) upon entering the heat exchanger or otherwise substantially uniform across whatever shape defines an entrance to the heat exchanger. The duct or diffuser conduit diffuses the flow resulting in a reduced flow velocity and more even flow profile. The supercritical fluid system includes a turbine on a shaft that is operably connected to the blower and a bypass air fan. In some embodiments, one or more gear systems are arranged between the turbine and the blower and/or bypass air fan to ensure proper gearing and transition of power from the turbine to the blower/fan. For aircraft propulsion, the bypass air provides most of the engine thrust. In some embodiments, a turboexpander could be added to the system. In some such embodiments, the turboexpander may be used to expand a cryogenic fuel (e.g., hydrogen, methane, etc.) prior to combustion within the burner. The turboexpander may be driven on the same shaft driven by the closed-loop turbine, upstream relative to the burner, and may be arranged to extract additional work within the system. In some embodiments, heat not recovered by the supercritical fluid closed-loop system may be recovered by an additional heat exchanger downstream (prior to atmosphere), to then use that recovered heat within a cryogenic fuel system. In one non-limiting example in accordance with the present disclosure, a CO<NUM> closed-loop system will only be heated to a CO<NUM> system optimum efficiency (e.g., about <NUM>° F).

Turning now to <FIG>, a schematic diagram of an aircraft propulsion system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft propulsion system <NUM> may be configured to be mounted to an aircraft to generate propulsive force for flight. The aircraft propulsion system <NUM> includes an air inlet <NUM>, a fan <NUM> (and/or low pressure compressor), a burner <NUM>, a diffuser conduit <NUM>, and an exhaust nozzle <NUM>. The fan <NUM> is arranged along a shaft <NUM> that is driven by a turbine <NUM>. The turbine <NUM> is part of a closed loop-supercritical fluid system <NUM>. A gearing system <NUM> is arranged between the turbine <NUM> and the fan <NUM> along the shaft <NUM>. The gearing system <NUM> is configured to transition a high speed rotation from the turbine <NUM> to a lower speed rotation for the fan <NUM>.

The closed loop-supercritical fluid system <NUM> is a closed loop system containing a supercritical fluid within a fluid line <NUM>. In some non-limiting embodiments, the supercritical fluid may be CO<NUM>. Along the fluid line <NUM> the closed loop-supercritical fluid system <NUM> includes the turbine <NUM>, an optional recuperator heat exchanger <NUM>, a cooler heat exchanger <NUM>, a compressor <NUM>, and a recovery heat exchanger <NUM>. As the supercritical fluid passes through the fluid line <NUM> it will drive the turbine <NUM> which in turn drives rotation of the shaft <NUM>. The compressor <NUM> and the fan <NUM> are coupled to the shaft <NUM> and thus are driven by the turbine <NUM>.

Starting at the turbine <NUM>, the supercritical fluid will be directed to the optional recuperator heat exchanger <NUM>. The recuperator heat exchanger <NUM> is a supercritical fluid-to-supercritical fluid heat exchanger with both flow paths through the recuperator heat exchanger <NUM> containing the supercritical fluid. The supercritical fluid will then flow into the cooler heat exchanger <NUM>, or directly to the cooler heat exchanger <NUM> if no recuperator heat exchanger <NUM> is present. The cooler heat exchanger <NUM> is configured to cool the supercritical fluid. As such, the cooler heat exchanger <NUM> may be an air-to-supercritical fluid or fuel-to-supercritical fluid heat exchanger. In operation, the cooler heat exchanger <NUM> is employed to reject heat of the supercritical fluid (e.g., sCO<NUM>) into the air stream forward of the burner <NUM> and is then recovered into the air stream that feeds the recovery heat exchanger <NUM>. Optionally, in an air-to-supercritical fluid configuration, the cooler heat exchanger <NUM> may be arranged in a duct, such as a bypass duct or bypass stream <NUM> of the aircraft propulsion system <NUM>, such as downstream from the air inlet <NUM> and downstream of the fan <NUM>. In such a configuration the bypass duct would direct air from the air inlet <NUM> toward the exhaust nozzle <NUM> without the air interacting with other components and, in some embodiment, may be the primary thrust generator of the aircraft propulsion system <NUM>. In the fuel-to-supercritical fluid, the cooler heat exchanger <NUM> may receive a fuel to be provided into the burner <NUM> for combustion and provide heat exchange between the fuel and the supercritical fluid. In some embodiments, the fuel may be jet fuel, biofuels, or cryogenic fuels (e.g., hydrogen, methane, etc.).

The supercritical fluid will then pass into the compressor <NUM>, which is driven on the shaft <NUM> by the turbine <NUM>. The supercritical fluid will then flow through the optional recuperator heat exchanger <NUM>, if present, and into the recovery heat exchanger <NUM>. The recovery heat exchanger <NUM> may be an annular structure arranged about and/or proximate to the exhaust nozzle <NUM>. Hot gases from the burner(s) <NUM> will flow along the diffuser conduit <NUM> and into the recovery heat exchanger <NUM>. The supercritical fluid will experience heat pickup within the recovery heat exchanger <NUM> which takes advantage of the waste heat generated by the burner(s) <NUM>. The heated supercritical fluid will then enter the turbine <NUM> and drive rotation thereof.

There are two primary fluid paths through the aircraft propulsion system <NUM>. First, there is the closed loop-supercritical fluid system <NUM>. This closed loop-supercritical fluid system <NUM> provides the primary driving force for rotating the shaft <NUM>. The second is an air flow path that passes through the aircraft propulsion system <NUM>. The bulk of the air will enter at the air inlet <NUM>, be driven by the fan <NUM>, and then pass through a bypass channel to provide thrust. A portion of the air will be directed from the fan <NUM> into the burner <NUM> to be combusted with a fuel, such as jet fuel, hydrogen, or the like. The combusted fuel and air will be passed into and through the diffuser conduit <NUM> prior to entering the recovery heat exchanger <NUM> and then exit the aircraft propulsion system <NUM> through the exhaust nozzle <NUM>. This exhaust will generate some amount of thrust to supplement the primary thrust generator in a bypass, with the bypass being similar to that shown and described with respect to <FIG> and/or as otherwise described herein.

The aircraft propulsion system <NUM> is distinct from prior gas turbine engines, particularly in that the primary driving force to drive the fan <NUM> is not a convention combustor-turbine configuration. The combusted fuel is not the primary driving force, but rather the closed-loop cycle of the closed loop-supercritical fluid system <NUM> provides the motive force to drive the fan <NUM> and generate thrust for flight. Advantageously, because the closed loop-supercritical fluid system <NUM> is a closed-loop system, there is no requirement for a direct fuel injection or reliance upon the fuel itself. The closed-loop cycle enables elimination of efficiency dependency that is due to a density drop associated with altitude changes when using traditional jet engines. That is, the aircraft propulsion system <NUM> provides for an altitude-invariant engine. As a result, embodiments of the present disclosure provide for systems having effectively no efficiency lapse due to altitude.

Turning now to <FIG>, a schematic diagram of an aircraft propulsion system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft propulsion system <NUM> may be substantially similar to that shown in <FIG>. The aircraft propulsion system <NUM> includes an air inlet <NUM>, a fan <NUM> (and/or low pressure compressor), a burner <NUM>, a diffuser conduit <NUM>, and an exhaust nozzle <NUM>. The fan <NUM> is arranged along a shaft <NUM> that is driven by a turbine <NUM> of a closed loop-supercritical fluid system <NUM>, similar to that described above. A gearing system <NUM> is arranged between the turbine <NUM> and the fan <NUM> along the shaft <NUM>. The gearing system <NUM> is configured to transition a high speed rotation from the turbine <NUM> to a lower speed rotation for the fan <NUM>.

The closed loop-supercritical fluid system <NUM> is a closed loop system containing a supercritical fluid within a fluid line <NUM>. Along the fluid line <NUM> the closed loop-supercritical fluid system <NUM> includes the turbine <NUM>, an optional recuperator heat exchanger <NUM>, a cooler heat exchanger <NUM>, a compressor <NUM>, and a recovery heat exchanger <NUM>. Similar to the above described configurations, the cooler heat exchanger <NUM> may be arranged between the fan <NUM> and the burner <NUM>. As such, heat from the supercritical fluid may be rejected into the air stream forward of the burner <NUM> and be recovered into the air stream that feeds the recovery heat exchanger <NUM>. In some embodiments, the cooler heat exchanger <NUM> can be optionally arranged within a bypass stream of the system. As the supercritical fluid passes through the fluid line <NUM> it will drive the turbine <NUM> which in turn drives rotation of the shaft <NUM>. The compressor <NUM> and the fan <NUM> are coupled to the shaft <NUM> and thus are driven by the turbine <NUM>.

In this embodiment, a heat exchanger bypass <NUM> is arranged along the flow path of the air that passes from the fan <NUM> to the exhaust nozzle <NUM>. As such, a portion of the combusted gases exiting the burner <NUM> may be used to provide additional thrust, without a portion of the energy extracted within the recovery heat exchanger <NUM>. As such, this configuration can potentially generate additional thrust through air passing through the heat exchanger bypass <NUM>. In this configuration, the size of the recovery heat exchanger <NUM> may be smaller than the similar recovery heat exchanger <NUM> in the configuration of <FIG>. This is because less air is passing through the recovery heat exchanger <NUM> and thus weight savings may be achieved, while also gaining increased thrust potential.

Turning now to <FIG>, a schematic diagram of an aircraft propulsion system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft propulsion system <NUM> may be similar in function as those described above. In this illustrative embodiment, there are three main flow paths of fluids through the aircraft propulsion system <NUM>. The aircraft propulsion system <NUM> includes a bypass flow path <NUM>, a hot gas flow path <NUM>, and a closed-loop flow path <NUM>.

The bypass flow path <NUM> is an air flow path that generates the majority of the thrust for the aircraft propulsion system <NUM>. Air enters the bypass flow path <NUM> at an air inlet <NUM>. The air is driven by a rotating fan <NUM> into a bypass duct <NUM>. The air is then ejected or driven out of a bypass nozzle <NUM>.

The hot gas flow path <NUM> is also an air flow path that is used to generate heat for the aircraft propulsion system <NUM>. Air is sourced from the air inlet <NUM>, passed through the fan <NUM>, and is accelerated through an optional blower <NUM>. The accelerated air is then passed through a burner <NUM> where the air is mixed and combusted with a fuel. The hot combusted gas is then passed through a hot stream duct <NUM> where the flow evens out and reduces turbulence prior to entering a recovery heat exchanger <NUM> and ejected out a hot stream nozzle <NUM>.

The closed-loop flow path <NUM> is part of a closed loop-supercritical fluid system <NUM> and contains a supercritical fluid within a fluid line. The closed loop-supercritical fluid system <NUM> includes a turbine <NUM>, an optional recuperator heat exchanger <NUM>, a cooler heat exchanger <NUM>, a compressor <NUM>, and the recovery heat exchangers <NUM>. As the supercritical fluid passes through the closed-loop flow path <NUM> it will drive the turbine <NUM> which in turn drives rotation of a shaft <NUM>.

The compressor <NUM> of the closed-loop flow path <NUM>, the optional blower <NUM> of the hot gas flow path <NUM>, and the fan <NUM> of the bypass flow path <NUM> are each coupled to the shaft <NUM> and thus are driven by the turbine <NUM>. The shaft <NUM> may include one or more gear systems 438a, 438b. The gear systems 438a, 438b may be configured to step the rotational speed of the shaft to appropriately drive the optional blower <NUM> and/or the fan <NUM>. For example, in some non-limiting embodiments, the gear systems may be configured to provide an overall gear ratio across both gear systems 438a, 438b between <NUM>:<NUM> to <NUM>:<NUM>, although other gear ratios and gearing may be employed without departing from the scope of the present disclosure, as will be appreciated by those of skill in the art.

In this illustrative embodiment, the cooler heat exchanger <NUM> of the closed-loop flow path <NUM> is arranged within the bypass duct <NUM> of the bypass flow path <NUM>. As such, the cooler heat exchanger <NUM> in this embodiment is an air-to-supercritical fluid heat exchanger.

Turning now to <FIG>, a schematic diagram of an aircraft propulsion system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft propulsion system <NUM> may be similar in function as those described above. In this illustrative embodiment, there are four main flow paths of fluids through the aircraft propulsion system <NUM>. The aircraft propulsion system <NUM> includes a bypass flow path <NUM>, a hot gas flow path <NUM>, a closed-loop flow path <NUM>, and a fuel flow path <NUM>.

The hot gas flow path <NUM> is also an air flow path that is used to generate heat for the aircraft propulsion system <NUM>. Air is sourced from the air inlet <NUM>, passed through the fan <NUM>, and is accelerated through a blower <NUM>. The accelerated air is then passed through a burner <NUM> where the air is mixed and combusted with a fuel. The hot combusted gas is then passed through a hot stream duct <NUM> where the flow evens out and reduces turbulence prior to entering a recovery heat exchanger <NUM> and ejected out a hot stream nozzle <NUM>. In some embodiments, an optional pump <NUM> may be arranged upstream of recovery heat exchanger <NUM>. For example, the potential phase change to liquid within the closed-loop flow path <NUM> may enable the use of pumps. Such pumping may enable boosting the working fluid within the closed-loop flow path <NUM> to a higher pressure that may reduce the required compression work, thus improving overall efficiencies.

The fuel flow path <NUM> is configured to supply fuel from a fuel tank <NUM> to the burner <NUM> for combustion with the air in the hot gas flow path <NUM>. The fuel flow path <NUM> is part of a cryogenic fuel system <NUM>. The cryogenic fuel system <NUM> provides a supply of fuel that is stored in a cryogenic state (such as liquid hydrogen or the like). In this embodiment, the fuel is extracted from the fuel tank <NUM> (e.g., by a pump or the like) and passed through a power electronics heat exchanger <NUM> that begins the warming process of the fuel from its cryogenic state. In some embodiments, an additional heat exchanger may be arranged within a fan stream of the engine, as discussed above, which can result in reduced size heat exchangers throughout the system.

In some embodiments, the cooler heat exchanger <NUM> provides the cryogenic fuel as a cold sink side for the hot working fluid passing through the closed-loop flow path <NUM>. In some embodiments, an additional heat exchanger may be arranged downstream of the cooler heat exchanger <NUM>, in a similar fan stream position to, for example, cooler heat exchanger <NUM> shown in <FIG>. That is, there may be multiple cooler heat exchanges arranged along the closed-loop flow path <NUM> and arranged in different arrangements along the closed-loop flow path <NUM> based on thermal considerations, for example, and/or upon physical limitations of the specific engine configurations. The fuel is then passed through the cooler heat exchanger <NUM> of the closed-loop flow path <NUM> where the fuel picks up heat from the fluid of the closed-loop flow path <NUM>. The fuel is then passed through a turbo expander <NUM> and subsequently supplied into the burner <NUM> for combustion with the air of the hot gas flow path <NUM>. It will be appreciated that the cryogenic fuel could be passed through other engine and/or aircraft heat exchangers as a coolant prior to entering the combustor, and the present flow path is not to be limiting.

The compressor <NUM> of the closed-loop flow path <NUM>, the blower <NUM> of the hot gas flow path <NUM>, the fan <NUM> of the bypass flow path <NUM>, and the turbo expander <NUM> of the fuel flow path <NUM> are each coupled to the shaft <NUM> and thus are driven by the turbine <NUM>. The shaft <NUM> may include one or more gear systems 538a, 538b. The gear systems 538a, 538b may be configured to step the rotational speed of the shaft to appropriately drive the blower <NUM>, the turbo expander <NUM>, and/or the fan <NUM>.

It will be appreciated that the schematic illustrations herein are more diagrams rather than structural configurations. Further, although it may appear that two of each element are illustrated, such may not be the case, rather a single structure or a multiple of different components may be included. For example, in some embodiments, the ducts and fans illustrated as two separate structures may in fact be portions of an annular structure. In other instances, the elements may be representative of multiple components arranged about a central axis. For example, the burners may be a number of discrete and separate burners (e.g., combustion chambers) arranged about a central axis, which may be defined by the shaft of the propulsion systems described herein. Thus, the illustrations are not to be limiting but rather are provided for explanatory and illustrative purposes only, as described herein.

Turning now to <FIG>, a schematic illustration of an aircraft <NUM> that may incorporate embodiments of the present disclosure is shown. The aircraft <NUM> includes a fuselage <NUM>, wings <NUM>, and a tail <NUM>. In this illustrated embodiment, the aircraft <NUM> includes wing-mounted aircraft propulsion systems <NUM>. The wing-mounted aircraft propulsion systems <NUM> may be arranged as the aircraft propulsion systems shown and described above. It will be appreciated that other aircraft configurations may employ the propulsion systems of the present disclosure without departing from the scope of the present disclosure. For example, fuselage-mounted and/or tail-mounted configurations are possible. Further, any number of propulsion systems may be employed, from one to four or more, depending on the aircraft configuration and power and thrust needs thereof.

Advantageously, embodiments of the present disclosure provide for an alternative aircraft propulsion system that may generate less waste and/or pollutants while providing improved efficiency that is not dependent upon altitude. Such aircraft propulsion systems may be zero or near-zero CO<NUM> emission aircraft propulsion systems. A new engine architecture is provided herein where the primary motive force is provided from a closed-loop cycle, rather than the conventional air-breathing Brayton cycles that primarily rely upon combustion to drive rotation of a shaft. Accordingly, the amount of fuel may be reduced, the size of components may be optimized, and improved efficiencies may be achieved. Advantageously, an altitude invariant engine design is provided by embodiments of the present disclosure. Further, improved thermal efficiencies of flight-ready engines are provided by the configurations described herein.

As used herein, the term "about" is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, "about" may include a range of ± <NUM>%, or <NUM>%, or <NUM>% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein.

It should be appreciated that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," "radial," "axial," "circumferential," and the like are with reference to normal operational attitude and should not be considered otherwise limiting.

Claim 1:
An aircraft propulsion system (<NUM>; <NUM>; <NUM>; <NUM>) comprising:
a closed loop-supercritical fluid system (<NUM>; <NUM>; <NUM>; <NUM>) having a turbine (<NUM>; <NUM>; <NUM>; <NUM>), a cooler heat exchanger (<NUM>; <NUM>; <NUM>; <NUM>), a compressor (<NUM>; <NUM>; <NUM>; <NUM>), and a recovery heat exchanger (<NUM>; <NUM>; <NUM>; <NUM>) arranged along a closed-loop flow path (<NUM>; <NUM>; <NUM>; <NUM>) of a supercritical fluid;
a shaft (<NUM>; <NUM>; <NUM>; <NUM>) operably coupled to the turbine (<NUM>...<NUM>) and configured to be rotationally driven by the turbine (<NUM>... <NUM>), wherein the compressor (<NUM>...<NUM>) is arranged on the shaft (<NUM>...<NUM>) and configured to be rotationally driven by the shaft (<NUM>...<NUM>);
a fan (<NUM>; <NUM>; <NUM>; <NUM>) configured to generate thrust, the fan (<NUM>...<NUM>) operably coupled to the shaft (<NUM>...<NUM>) to be rotationally driven by the shaft (<NUM>...<NUM>); and
a burner (<NUM>; <NUM>; <NUM>; <NUM>) configured to combust a fuel and air from the fan (<NUM>... <NUM>) to generate a combusted gas and supply said combusted gas to the recovery heat exchanger (<NUM>...<NUM>) of the closed loop-supercritical fluid system (<NUM>...<NUM>) and out an exhaust nozzle (<NUM>; <NUM>; <NUM>; <NUM>),
characterized by:
a diffuser conduit (<NUM>; <NUM>) arranged between the burner (<NUM>; <NUM>) and the recovery heat exchanger (<NUM>; <NUM>) with the burner (<NUM>; <NUM>) arranged to direct combusted fuel and air into the diffuser conduit (<NUM>; <NUM>) and the recovery heat exchanger (<NUM>; <NUM>) arranged downstream from the diffuser conduit (<NUM>; <NUM>), the diffuser conduit (<NUM>; <NUM>) configured to slow down and even out a profile of a flow of the combusted gas prior to entry into the recovery heat exchanger (<NUM>; <NUM>).