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 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.

Typically, liquid fuel is employed for combustion onboard an aircraft, in the gas turbine engine. The liquid fuel has conventionally been a hydrocarbon-based fuel. Alternative fuels have been considered, but suffer from various challenges for implementation, particularly on aircraft. Hydrogen-based and/or methane-based fuels are viable effective alternatives which may not generate the same combustion byproducts as conventional hydrocarbon-based fuels. The use of liquid, compressed, or supercritical hydrogen and/or methane, as a gas turbine fuel source, may require very high efficiency propulsion, in order to keep the volume of the fuel low enough to feasibly carry on an aircraft. That is, because of the added weight associated with such liquid/compressed/supercritical fuels, such as related to vessels/containers and the amount (volume) of fuel required, improved efficiencies associated with operation of the gas turbine engine may be necessary.

<CIT> discloses a reaction propulsion system having a ram air intake, a combustion chamber having at its rearward end an impulse expansion outlet nozzle and a direct expansion turbine for driving a compressor.

According to a first aspect, a turbine engine system is provided. The turbine engine system includes a combustor arranged along a core flow path of the turbine engine; a drive shaft having at least a fan, a compressor section and a turbine section coupled thereto, with the fan configured to be rotationally driven through the drive shaft by the rotationally driven turbine section and the combustor is arranged between the compressor section and the turbine section along the core flow path; a cryogenic fuel tank configured to supply a fuel to the combustor; an expansion turbine mechanically coupled to the drive shaft and arranged in the core flow path downstream from the turbine section, the expansion turbine configured to receive fuel from the cryogenic fuel tank and expand said fuel, wherein expansion of said fuel by the expansion turbine drives rotation of the expansion turbine to provide power input to the drive shaft and supplement or augment rotation of the drive shaft; and a flow supply line fluidly connecting the cryogenic fuel tank to the combustor with the expansion turbine arranged between the cryogenic fuel tank and the combustor along the flow supply line.

Optionally, the drive shaft comprises a low spool and a high spool and the expansion turbine is mechanically coupled to the low spool.

Optionally, the fuel is one of liquid hydrogen and liquid methane.

Optionally, the expansion turbine is configured to impart work to the drive shaft during expansion of the fuel.

Optionally, a waste heat-heat exchanger is arranged downstream of the combustor along a core flow path, wherein the waste heat-heat exchanger is arranged along the flow supply line and configured to heat the fuel.

Optionally, the waste heat-heat exchanger is arranged upstream of the expansion turbine along the flow supply line.

Optionally, a power electronics cooling heat exchanger is arranged along the flow supply line between the cryogenic fuel tank and the expansion turbine.

Optionally, a supplemental cooling heat exchanger is arranged along the flow supply line and configured to cool at least one of engine oil, environmental control system fluids, pneumatic off-takes, and cooled cooling air fluids.

Optionally, at least one flow controller is arranged along the flow supply line and configured to control a flow of fuel through the flow supply line.

Optionally, a gear system is operably coupled to the drive shaft and configured to drive rotation of the fan.

According to another aspect, an aircraft engine system according to claim <NUM> is provided. The aircraft engine system includes a combustor arranged along a core flow path of the aircraft engine, a drive shaft having at least a compressor section and a turbine section coupled thereto, a fan operably coupled to the draft shaft, a cryogenic fuel tank configured to supply a fuel to the combustor, and an expansion turbine mechanically coupled to the drive shaft, the expansion turbine configured to receive fuel from the cryogenic fuel tank and expand said fuel, wherein expansion of said fuel by the expansion turbine drives rotation of the expansion turbine to provide power input to the drive shaft. A flow supply line fluidly connecting the cryogenic fuel tank to the combustor with the expansion turbine is arranged between the cryogenic fuel tank and the combustor along the flow supply line.

Optionally, the combustor, the drive shaft, the compressor section, and the turbine section are arranged as a turboshaft engine or a turboprop engine.

Optionally, the combustor, the drive shaft, the compressor section, and the turbine section are arranged as a turbofan engine.

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 in a fan case <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 engine <NUM> has a bypass ratio that is greater than about six (<NUM>), with an example embodiment being greater than about ten (<NUM>). 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 (<NUM>). In one non-limiting embodiment, the bypass ratio of the gas turbine engine <NUM> is greater than about ten (<NUM>:<NUM>), 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 (<NUM>:<NUM>). 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, turboshafts, or turboprops.

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>.

Gas turbine engines generate substantial amounts of heat that is exhausted from the turbine section <NUM> into a surrounding atmosphere. This expelled exhaust heat represents wasted energy, and can be a large source of inefficiency in gas turbine engines.

Turning now to <FIG>, a schematic diagram of a turbine engine system <NUM> in accordance with an embodiment of the present disclosure is shown. The turbine engine system <NUM> may be similar to that shown and described above, but is configured to employ a non-hydrocarbon fuel source, such as liquid/compressed/supercritical hydrogen and/or methane, or other types of cryogenic fuels, as will be appreciated by those of skill in the art. The turbine engine system <NUM> includes an inlet <NUM>, a fan <NUM>, a low pressure compressor <NUM>, a high pressure compressor <NUM>, a combustor <NUM>, a high pressure turbine <NUM>, a low pressure turbine <NUM>, a core nozzle <NUM>, and an outlet <NUM>. A core flow path is defined through, at least, the compressor <NUM>,<NUM>, the turbine <NUM>, <NUM>, and the combustor sections <NUM>. The compressor <NUM>, <NUM>, the turbine <NUM>, <NUM>, and the fan <NUM> are arranged along a shaft <NUM>.

As shown, the turbine engine system <NUM> includes a cryogenic fuel system <NUM>. The cryogenic fuel system <NUM> is configured to supply a fuel from a cryogenic fuel tank <NUM> to the combustor <NUM>. The fuel is supplied from the cryogenic fuel tank <NUM> to the combustor <NUM> through a fuel supply line <NUM>. The fuel supply line <NUM> may be controlled by a flow controller <NUM> (e.g., pump(s), valve(s), or the like). The flow controller <NUM> may be configured to control a flow through the fuel supply line <NUM> based on various criteria as will be appreciated by those of skill in the art. For example, various control criteria can include, without limitation, target flow rates, target turbine output, cooling demands at one or more heat exchangers, target flight envelopes, etc. As shown, between the cryogenic fuel tank <NUM> and the flow controller <NUM> may be an optional power electronics cooling heat exchanger <NUM>. The power electronics cooling heat exchanger <NUM> may receive the cryogenic fuel directly from the cryogenic fuel tank <NUM> as a first fluid and a power electronics working fluid for power electronics of the turbine engine system <NUM> (or other aircraft power electronics) as a second fluid. A relatively hot power electronics working fluid can pass through the power electronics cooling heat exchanger <NUM> and heat may be transferred into the cryogenic fuel. This may serve, in some configurations, to begin raising a temperature of the cryogenic fuel to a desired temperature for efficient combustion in the combustor <NUM>.

When the fuel is directed along the flow supply line <NUM>, the fuel will pass through a core flow path heat exchanger <NUM> (e.g., an exhaust waste heat recovery heat exchanger). The core flow path heat exchanger <NUM> is arranged in the core flow path downstream of the combustor <NUM>, and in some embodiments, downstream of the low pressure turbine <NUM>. In this illustrative embodiment, the core flow path heat exchanger <NUM> is arranged downstream of the low pressure turbine <NUM> and at or proximate the core nozzle <NUM> upstream of the outlet <NUM>. As the fuel passes through the core flow path heat exchanger <NUM>, the fuel will pick up heat from the exhaust of the turbine engine system <NUM>. As such, the temperature of the cryogenic fuel will be increased.

The heated fuel will then be passed into an expansion turbine <NUM>. As the fuel passes through the expansion turbine <NUM> the fuel will be expanded. The process of passing the fuel through the expansion turbine <NUM> will cause a phase change from liquid to gas and/or warm the liquid fuel and/or further expand gaseous fuel, which is aided by one or more heat exchangers along the fuel supply line <NUM>. The expanded fuel may then pass through an optional supplemental heating heat exchanger <NUM>. The supplemental heating heat exchanger <NUM> is configured to receive the heated (but potentially still relatively cold) fuel as a first fluid and as the second fluid may receive one or more aircraft system fluids, such as, without limitation, engine oil, environmental control system fluids, pneumatic off-takes, or cooled cooling air fluids. As such, the fuel will be heated as the other fluid may be cooled. The fuel will then be injected into the combustor <NUM> through one or more fuel injectors, as will be appreciated by those of skill in the art. Because the fuel is heated from the cryogenic state in the cryogenic fuel tank <NUM> through the various mechanisms along the flow supply line <NUM>, combustion efficiency may be improved.

In accordance with embodiments of the present disclosure, the expansion turbine <NUM> for the cryogenic fuel is arranged along and driven by the shaft <NUM>. The shaft <NUM> may be a two-spool shaft system, such as described with respect to <FIG>, having a low spool and a high spool. In some embodiments of the present disclosure, the expansion turbine <NUM> is configured to be driven by the low spool of the two-spool shaft system. In such configurations, the expansion of the cryogenic fuel within the expansion turbine <NUM> can be used to supplement or augment the cycle of the shaft <NUM>. That is, the expansion within the expansion turbine <NUM> can provide additional power input to the shaft <NUM> by mechanically tying the expansion turbine <NUM> to the shaft <NUM> (e.g., low spool shaft). The shaft <NUM> may be a shaft of a turbo shaft engine configuration or a shaft of a turbo fan engine configuration, as will be appreciated by those of skill in the art. In some embodiments, the expansion turbine <NUM> may be operably coupled to the shaft <NUM> through a gearbox or other geared system.

Turning now to <FIG>, a schematic illustration of a turboshaft engine or turboprop <NUM> in accordance with an embodiment of the present disclosure is shown. The turboshaft or turboprop engine <NUM> may be powered by combusting a fuel that is stored at cryogenic temperatures. The turboshaft or turboprop engine <NUM> includes a propeller <NUM>, a compressor section <NUM>, a combustor section <NUM>, a turbine section <NUM>, and an outlet <NUM>. The compressor section <NUM> and the turbine section <NUM>, at least, are arranged along a drive shaft <NUM>. The drive shaft <NUM> is operably connected to a gear system <NUM> that is configured to drive rotation of the propeller <NUM>. In some embodiments or configurations, the gear system <NUM> may be a gearbox or the like.

The turboshaft or turboprop engine <NUM> includes a cryogenic fuel system <NUM>. The cryogenic fuel system <NUM> is configured to supply a fuel from a cryogenic fuel tank <NUM> to the combustor section <NUM>. The fuel is supplied from the cryogenic fuel tank <NUM> to the combustor section <NUM> through a fuel supply line <NUM>. The fuel supply line <NUM> may be controlled by a flow controller <NUM> (e.g., pump(s), valve(s), or the like).

When the fuel is directed along the flow supply line <NUM>, the fuel will pass through a core flow path heat exchanger <NUM> (e.g., an exhaust waste heat recovery heat exchanger). The core flow path heat exchanger <NUM> is arranged in the core flow path downstream of the combustor <NUM>, and in some embodiments, downstream of the turbine section <NUM>. In this illustrative embodiment, the core flow path heat exchanger <NUM> is arranged within the outlet <NUM> of the turboshaft or turboprop engine <NUM>. As the fuel passes through the core flow path heat exchanger <NUM>, the fuel will pick up heat from the exhaust of the turboshaft or turboprop engine <NUM>. As such, the temperature of the cryogenic fuel will be increased.

The heated fuel will then be passed into an expansion turbine <NUM>. As the fuel passes through the expansion turbine <NUM> the fuel will be expanded. The process of passing the fuel through the expansion turbine <NUM> will cause a phase change from liquid to gas and/or warm the liquid fuel and/or further expand gaseous fuel, which is aided by one or more heat exchangers along the fuel supply line <NUM>. The fuel will then be injected into a combustor of the combustor section <NUM> through one or more fuel injectors, as will be appreciated by those of skill in the art. Because the fuel is heated from the cryogenic state in the cryogenic fuel tank <NUM> through the various mechanisms along the flow supply line <NUM>, combustion efficiency may be improved.

In accordance with embodiments of the present disclosure, the expansion turbine <NUM> for the cryogenic fuel is arranged along and driven by the shaft <NUM> of the turboshaft or turboprop engine <NUM>. As shown, an expansion shaft <NUM> may be operably coupled to the shaft <NUM> of the turboshaft or turboprop engine <NUM>. In such configurations, the expansion of the cryogenic fuel within the expansion turbine <NUM> can be used to supplement or augment the cycle of the turboshaft or turboprop engine <NUM>. That is, the expansion within the expansion turbine <NUM> can provide additional power input to the shaft <NUM> of the turboshaft or turboprop engine <NUM> by mechanically tying the expansion turbine <NUM> to the shaft <NUM> of the turboshaft or turboprop engine <NUM>.

In this embodiment, a secondary flow controller <NUM> (e.g., valves and/or pumps) may be arranged downstream from the expansion turbine <NUM>. The secondary flow controller <NUM> may be configured to control a fuel input into the combustor(s) of the combustor section <NUM>. Thrust generated by the turboshaft or turboprop engine <NUM> can be controlled, for example, through a combination of pitch of the propeller <NUM>, throttling the supply of liquid fuel to the expansion turbine <NUM> (through control of the flow controller <NUM>), and throttling gaseous fuel to the combustor (through the secondary flow controller <NUM>). The supply of fuel to the expansion turbine <NUM> can enable power to be input to the shaft <NUM> (e.g., directly or through a gearbox) through work extracted during the expansion process.

Turning now to <FIG>, a schematic illustration of a turbofan engine <NUM> in accordance with an embodiment of the present disclosure is shown. The turbofan engine <NUM> may be powered by combusting a fuel that is stored at cryogenic temperatures. The turbofan engine <NUM> includes a fan section, a compressor section, a combustor section, a turbine section, and an outlet, similar to that shown and described above. The fan section, the compressor section, and the turbine section, at least, are arranged along a drive shaft. The turbofan engine <NUM> includes a cryogenic fuel system <NUM>. The cryogenic fuel system <NUM> is configured to supply a fuel from a cryogenic fuel tank <NUM> to the combustor section. The fuel is supplied from the cryogenic fuel tank <NUM> to the combustor section through a fuel supply line <NUM>. The fuel supply line <NUM> may be controlled by a flow controller <NUM> (e.g., pump(s), valve(s), or the like).

When the fuel is directed along the flow supply line <NUM>, the fuel will pass through a core flow path heat exchanger <NUM> (e.g., an exhaust waste heat recovery heat exchanger). The core flow path heat exchanger <NUM> is arranged in the core flow path downstream of the combustor section, and in some embodiments, downstream of the turbine section. In this illustrative embodiment, the core flow path heat exchanger <NUM> is arranged within the outlet of the turbofan engine <NUM>. As the fuel passes through the core flow path heat exchanger <NUM>, the fuel will pick up heat from the exhaust of the turbofan engine <NUM>. As such, the temperature of the cryogenic fuel will be increased.

The heated fuel will then be passed into an expansion turbine <NUM>. As the fuel passes through the expansion turbine <NUM> the fuel will be expanded. The process of passing the fuel through the expansion turbine <NUM> will cause a phase change from liquid to gas and/or warm the liquid fuel and/or further expand gaseous fuel, which is aided by one or more heat exchangers along the fuel supply line <NUM>. For example, in this illustrative embodiment, the expanded fuel may pass through an optional supplemental heating heat exchanger <NUM>. The supplemental heating heat exchanger <NUM> is configured to receive the heated (but potentially still relatively cold) fuel as a first fluid and as the second fluid may receive one or more aircraft system fluids, such as, without limitation, engine oil, environmental control system fluids, pneumatic off-takes, or cooled cooling air fluids. As such, the fuel will be heated as the other fluid may be cooled. The fuel will then be injected into a combustor of the combustor section through one or more fuel injectors, as will be appreciated by those of skill in the art. Because the fuel is heated from the cryogenic state in the cryogenic fuel tank <NUM> through the various mechanisms along the flow supply line <NUM>, combustion efficiency may be improved.

In accordance with embodiments of the present disclosure, the expansion turbine <NUM> for the cryogenic fuel is arranged along and driven by the shaft of the turbofan engine <NUM>. In such configurations, the expansion of the cryogenic fuel within the expansion turbine <NUM> can be used to supplement or augment the cycle of the turbofan engine <NUM>. That is, the expansion within the expansion turbine <NUM> can provide additional power input to the shaft of the turbofan engine <NUM> by mechanically tying the expansion turbine <NUM> to the shaft of the turbofan engine <NUM> (e.g., low spool shaft).

In this embodiment, a secondary flow controller <NUM> (e.g., a valve and/or pump) may be arranged downstream from the expansion turbine <NUM>. The secondary flow controller <NUM> may be configured to control a fuel input into the combustor(s) of the combustor section. Thrust generated by the turbofan engine <NUM> can be controlled, for example, through a combination of throttling the supply of liquid fuel to the expansion turbine <NUM> (through control of the flow controller <NUM>) and throttling gaseous fuel to the combustor (through the secondary flow controller <NUM>). In some configurations, the expansion turbine <NUM> can be configured to add power to the low spool, which in turn can enable a reduction in the amount of fuel burned in the combustor. Such configurations can take advantage of energy stored in the cryogenic fuel that is released when the cryogenic fuel is expanded within the expansion turbine <NUM>. Thrust is generally controlled by fan speed, and thus, in such configurations, a control can be implemented to throttle the fuel burned to hold the fan speed.

It is noted that in the configurations shown in <FIG>, specific arrangements of components are shown and described. However, it will be appreciated by those of skill in the art that various other arrangements are possible, without departing from the scope of the present disclosure. It is noted that one or more optional heat exchangers can provide various cooling to fluids of other engine and/or aircraft systems. For example, as described, a heat exchanger can be provided to cryogenically cool power electronics by the fuel, adding some heat to the fuel before recovering heat from another source in the turbine engine (e.g., a source along the core flow path, referred to herein as core flow path heat exchangers). As described, a heat exchanger can be provided to cool the engine oil, aircraft ECS needs, pneumatic off-takes, and/or cooled cooling air, downstream of the expansion turbine prior to being injected into the combustor. Alternatively, an aircraft ECS cooler heat exchanger could be arrange between the cryogenic power electronics cooler and the core flow path heat exchangers. The engine oil cooler or cooled cooling air heat exchangers could alternatively be between the core flow path heat exchangers and expansion turbine.

In some embodiments, the flow controller of the systems may be configured to allow for a portion of the fuel to flow along one or more flow paths. In some such configurations, two or more separated flows of fuel may be recombined and mixed together prior to or at the point of entering the expansion turbine and/or the combustor. As such, a flow controller may be dynamically controlled to ensure a desired temperature of the fuel at the point of injection into the combustor of the turbine engine.

Advantageously, embodiments of the present disclosure are directed to improved turbine engine systems that employ non-hydrocarbon fuels at cryogenic temperatures. In accordance with some embodiments, the systems described herein may allow the cryogenic fuel to recover heat from various systems such as waste heat-heat exchangers, system component heat exchangers, and expansion turbines. Such expansion turbines, advantageously, may provide supplemental or augmentation to thrust generated by the engines. This may be achieved due to the liquid or cold fuel expanding within the expansion turbine, thus applying force to a shaft of the expansion turbine, which may be mechanically coupled to a shaft of the engine. This additional rotation can be used to generate additional thrust output from the engine.

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:
A turbine engine system (<NUM>; <NUM>; <NUM>), comprising:
a combustor (<NUM>; <NUM>) arranged along a core flow path of the turbine engine;
a drive shaft (<NUM>; <NUM>) having at least a fan (<NUM>; <NUM>), a compressor section (<NUM>, <NUM>; <NUM>) and a turbine section (<NUM>, <NUM>; <NUM>) coupled thereto, with the fan configured to be rotationally driven through the drive shaft by the rotationally driven turbine section and the combustor is arranged between the compressor section and the turbine section along the core flow path;
a cryogenic fuel tank (<NUM>; <NUM>; <NUM>) configured to supply a fuel to the combustor;
an expansion turbine (<NUM>; <NUM>; <NUM>) mechanically coupled to the drive shaft and arranged in the core flow path downstream from the turbine section, the expansion turbine (<NUM>; <NUM>; <NUM>) configured to receive fuel from the cryogenic fuel tank (<NUM>; <NUM>; <NUM>) and expand said fuel, wherein expansion of said fuel by the expansion turbine drives rotation of the expansion turbine to provide power input to the drive shaft and supplement or augment rotation of the drive shaft; and
a flow supply line (<NUM>; <NUM>; <NUM>) fluidly connecting the cryogenic fuel tank to the combustor with the expansion turbine arranged between the cryogenic fuel tank and the combustor along the flow supply line.