Hybrid expander cycle with pre-compression cooling and turbo-generator

A gas turbine engine system includes a gas turbine engine and a fuel turbine system. The gas turbine engine includes an air inlet, compressor, combustor, turbine, and heat exchange system. The heat exchange system is configured to transfer thermal energy from an inlet air flow or exhaust air flow to a fuel to produce a gaseous fuel that is used to drive a fuel turbine and fuel pump and used for combustion in the gas turbine engine. The fuel turbine is in fluid communication with the heat exchange system and the combustor and configured to extract energy from expansion of the gaseous fuel. The fuel pump is configured to be driven by the fuel turbine and is in fluid communication with the heat exchanger system.

BACKGROUND

The present disclosure relates generally to a gas turbine engine of an aircraft and, more specifically to a gas turbine engine using non-traditional cooled liquid fuel with pre-compression inlet air cooling and a turbo-generator.

Aircraft engines are being simultaneously challenged to provide increases in thermal efficiency, electrical power generation, and thermal management while reducing environmental emissions. Shaft power extraction impacts sizing of turbomachinery components and can have an adverse impact on performance and operability. Thermal management (e.g, providing a heat sink for engine and external systems) is limited by engine internal temperatures and can result in excessive pressure losses as heat is rejected using heat exchangers or other devices. Use of existing hydrocarbon fuel as a heat sink is limited to fuel temperature typically under 450° F. due to coking and decomposition of fuels at high temperatures. Thermal efficiency improvement trends typically involve providing a higher overall pressure ratio (OPR) of the compression system with associated increases in compressor discharge pressure (P3) and accompanying temperature (T3). The OPR is increased by increasing a compressor discharge pressure (P3). As pressure increases across the compressor, temperature also increases. Current aircraft designs are generally limited by operational temperature limits of materials used for gas turbine structures. As aircraft speeds increase, inlet temperatures (T2) increase, which poses a challenge for providing improvements in thermal efficiency as T2greatly impacts T3and thereby the achievable OPR. For example, for a MIL-E-5007D inlet recovery, the temperature (T2) entering into the engine at subsonic cruise of Mach 0.8 at 45,000 feet is −20° F. (−29° C.), while the inlet temperature at the same altitude at Mach 3.0 is in excess of 630° F. (332° C.). This increased temperature greatly impacts the ability to operate at a high OPR due to limitations in material capability at high temperature. In addition, the relationship between T2, OPR, and T3may result in the formation of harmful emissions including nitrogen oxides (NOx), which can be produced in the compression cycle of the engine with production compounded in subsequent combustion and turbine operations. This will be of particular concern for potential future commercial supersonic transport or business jet aircraft flying at higher altitudes (i.e., greater than 60,000 feet) where NOx emissions can have an increased environmental impact. While emission reductions in NOx, as well as carbon monoxide and particulates is desirable, it often runs counter to desired cycle characteristics and can be difficult to achieve with current hydrocarbon fuels. New engine concepts that could fundamentally alter the impact of these trends are desired.

SUMMARY

In one aspect, a gas turbine engine system includes a gas turbine engine and a fuel turbine system. The gas turbine engine includes an air inlet, compressor, combustor, turbine, and heat exchange system. The heat exchange system is configured to transfer thermal energy from an inlet air flow or exhaust air flow to a fuel to produce a gaseous fuel that is used to drive a fuel turbine and fuel pump and used for combustion in the gas turbine engine. The fuel turbine is in fluid communication with the heat exchange system and the combustor and configured to extract energy from expansion of the gaseous fuel. The fuel pump is configured to be driven by the fuel turbine and is in fluid communication with the heat exchanger system.

In another aspect, a method of operating a gas turbine engine system includes cooling an inlet air flow via an inlet heat exchange system by transferring thermal energy to a liquid fuel to produce a cooled inlet air flow. The cooled inlet air flow is compressed to produce a compressed air for combustion. Energy is extracted from expansion of a gaseous fuel through a multi-stage fuel turbine, which produces a gaseous fuel having a pressure greater than a pressure of the compressed air flow. The fuel turbine is fluidly connected to the inlet heat exchange system. A mixture of the gaseous fuel from an outlet of the fuel turbine and compressed air flow is combusted. A pump driven by the fuel turbine pumps liquid fuel to the inlet heat exchange system.

In yet another aspect, a fuel turbine system includes a heat exchange system configured to transfer heat to a liquid fuel to produce a gaseous fuel, a fuel turbine in fluid communication with the heat exchange system and configured to extract energy from expansion of the gaseous fuel, and a fuel pump in fluid communication with the heat exchange system and configured to deliver the liquid fuel to the heat exchange system. The fuel pump is mechanically coupled to the fuel turbine and configured to be driven by the fuel turbine.

DETAILED DESCRIPTION

The present disclosure combines the use of a non-traditional fuel, such as methane or hydrogen, stored in a cooled liquid state, to drive a hybrid cycle of a gas turbine engine system—the hybrid cycle consisting of a conventional Brayton cycle with pre-compression inlet air cooling and an expander cycle, which utilizes waste heat added to the fuel to drive a turbo-generator. The disclosed embodiments are directed to an augmented military-style gas turbine engine with a low bypass ratio, however, could be adapted for use in a high speed commercial aircraft engine with or without thrust augmentation. In the cooled liquid state, the fuel provides a heat sink for high temperature inlet air. The liquid fuel can be further heated by exhaust gas waste heat of the gas turbine engine to form a high-pressure gaseous fuel, which is used to drive a multi-stage fuel turbine, liquid fuel pump, and motor/generator. Fuel expanded through the multi-stage fuel turbine is then used in the gas turbine engine for combustion.

The disclosed system can use plentiful and cleaner burning fuel as compared to current hydrocarbon fuels to allow engine cycle and OPR/thermal efficiency gains to be established independent of vehicle speed and inlet compression characteristic (i.e., temperature) while enabling high speed engine design using existing materials. Additionally, electrical generation using regenerative (i.e., waste heat) input can be provided with reduced impact on turbomachinery sizing, performance, and operability. Additional benefits may also be realized with new combustor concepts enabled by the disclosed system, including reduced combustion length, staged combustion, and reduced emissions. The disclosed system could also be used to reduce augmenter length for applications where increased specific thrust is required.

FIG. 1is a schematic diagram of one embodiment of gas turbine engine system10with pre-compression cooling and expander cycle. System10includes gas turbine engine12and turbo-generator14. Gas turbine engine12includes inlet heat exchanger16, fan section18, compressor section20(including low pressure compressor (LPC)21and high pressure compressor (HPC)22), combustor section24, turbine section26(including high pressure turbine (HPT)27and low pressure turbine (LPT)28), optional augmentor29, and optional exhaust heat exchanger30. Fan section18drives inlet air flow FI. Compressor section20draws air in along a core flow path where air is compressed and communicated to combustor section24. In combustor section24, air is mixed with fuel and ignited to generate a high-pressure combustion exhaust gas stream that expands through turbine section26where energy is extracted and utilized to drive fan section18and compressor section20.

Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines, including, for example, a turbine engine including a three-spool architecture. While the present disclosure focuses on utilization of a twin spool, axial flow gas turbine fan-jet military-style engine, it will be appreciated that it has utility in other types of engines, such as turbojets and engines used in nonmilitary, high speed applications (e.g., commercial supersonic transport).

Turbo-generator includes fuel turbine32, fuel pump34, and motor/generator36. Fuel turbine32is a multi-stage turbine with multiple stages of turbine blades driven by the expansion of high-pressure gaseous fuel. Fuel turbine32, fuel pump34, and motor/generator36are coupled to rotor shaft38such that fuel pump34and motor/generator36are mechanically driven by the rotation of fuel turbine32. Fuel pump34is configured to deliver fuel through system10. Motor/generator36can be configured to supply power for system10components and/or other engine systems and power needs.

System10additionally includes fuel tank40configured to contain a cryogenic fuel and a series of conduits configured to deliver the fuel in liquid and/or gaseous phase via fuel pump34through system10. A plurality of valved fuel lines A-H can be used to deliver and control the flow of fuel through system10via controller42. Controller42can be operatively coupled (e.g., electrically and/or communicatively) to components shown inFIG. 1as well as components not depicted (e.g., additional valves, sensors, etc.) to control operation of system10. Controller42can be configured to receive, transmit, and/or process sensor data and/or signals for the operation of system10. Controller42can include a processor, such as a microprocessor, programmable logic controller, or the like.

Additionally, system10can include a plurality of temperature and/or pressure sensors S1-S3configured to detect a temperature and/or pressure of the fuel at various locations in system10or air flow through gas turbine engine12. As shown inFIG. 1, sensors S1and S2detect a temperature and/or pressure of fuel in fuel lines E/G and J, while sensor S3detects a temperature and/or pressure (T2/P2) of air entering fan18. In some embodiments, sensor S4can be used to detect leakage of fuel from inlet heat exchanger16. In some embodiments, system10can also include auxiliary fuel tank46, configured to deliver an auxiliary supply of gaseous fuel to combustor24.

System10is configured for use with gas turbine engines operating at high speed (i.e., supersonic speeds typically greater than Mach 2) with inlet air temperatures generally exceeding 250° F. (121° C.). System10can be used to improve thermal efficiency of gas turbine engine12by reducing inlet temperatures T2and thereby allowing an increased level of overall pressure ratio (OPR) for a given T3limit. The OPR is increased by increasing the compressor outlet pressure P3. As pressure increases across compressor section20, temperature also increases. Current aircraft engine designs are generally limited by material capability and compressor outlet temperature T3limits. As aircraft speeds increase, inlet temperatures T2increase, which poses a challenge for providing improvements in thermal efficiency as T2greatly impacts T3and thereby the achievable OPR. This increased temperature greatly impacts the ability to operate at a high OPR due to limitations in material capability at high temperature. System10allows OPR/thermal efficiency gains to be established independent of vehicle speed and inlet air temperature T2.

In addition to improving thermal efficiency of gas turbine engine12, system10can be configured to reduce combustion related emissions as compared to engines that burn traditionally used fossil fuels, and to generate power for operating components of system10, including fuel pump34, as well as other engine systems from heat supplied by inlet air flow and/or exhaust gas from gas turbine engine12.

As illustrated inFIG. 1, a cryogenic liquid fuel is stored in fuel tank40at low temperature and pressure. Suitable fuels can include, but are not limited to, liquefied methane (i.e., liquefied natural gas (LNG)) and liquid hydrogen. Tank40can be configured in any manner and made of any material suitable for storing cryogenic fuels as known in the art. The temperature of the fuel is sufficiently low to provide inlet air cooling, but can vary significantly depending on system10configuration, inlet heat exchanger16configuration, and inlet air temperature T2. For example, inlet air temperature T2at Mach 3 can be greater than 630° F. (332° C.). Generally, it will be desired to reduce the inlet air temperature T2to 250° F. (121° C.) or less. In one non-limiting example, liquid hydrogen fuel stored at −425° F. (−254° C.) and 25 psi (172 kPa) can be used effectively for inlet air cooling at Mach 3 operating conditions. As will be understood, fuel temperature can be varied considerably from the example provided depending on system configurations and operational conditions (e.g., lower fuel temperatures may be used as speeds approach or exceed Mach 5).

Fuel pump34is configured to pump liquid fuel from tank40through fuel lines A and C to inlet heat exchanger16. Fuel pump34increases the pressure of liquid fuel entering inlet heat exchanger16. In some non-limiting embodiments, the pressure of fuel entering inlet heat exchanger16from fuel pump34can be in excess of 550 psi (3,792 kPa). It will be understood by one of ordinary skill in the art to design pump34and the fuel circuit of system10(i.e., fuel lines A-J) to provide effective circulation of the fuel through system10. Liquid fuel can be pumped to inlet heat exchanger16when inlet air cooling is needed. Generally, inlet heat exchanger16is needed only during high speed flight when inlet air temperatures exceed 250° F. (121° C.) and generally is not needed during takeoff and subsonic flight or when temperatures are below 250° F. (121° C.). Valve C can be used to control fuel flow into inlet heat exchanger16based on aircraft operation. Inlet heat exchanger16is positioned in a primary inlet of gas turbine engine12and configured to substantially cover the primary inlet to provide cooling to a substantial portion of inlet air while also allowing passage of inlet air. Inlet heat exchanger16can be a plate-fin, shell-and-tube, or other suitable air-to-liquid heat exchanger as known in the art. To substantially cover the primary inlet to gas turbine engine12, inlet heat exchanger16can have a web-like or grid-like configuration with a network of cooling channels extending radially, crosswise, and/or in concentric rings over the primary inlet to provide cooling to a substantial portion of inlet air entering gas turbine engine12.

A pressure drop can occur across inlet heat exchanger16. This drop in pressure can be avoided or mitigated when inlet heat exchanger16is not in use through use of auxiliary inlet44. Auxiliary inlet44can be used to partially bypass the primary inlet and inlet heat exchanger16. Auxiliary inlet44can include one or more openings arranged around a circumference of a nacelle of gas turbine engine12. Inlet air entering gas turbine engine12through auxiliary inlet44will be at a higher pressure than inlet air entering gas turbine engine12from inlet heat exchanger16, which causes air to be drawn preferentially through auxiliary inlet44. Auxiliary inlet44can be closed when inlet heat exchanger16is in use.

Inlet heat exchanger16is configured to place inlet air flow FIand liquid fuel in thermal communication such that thermal energy from the inlet air is transferred to the liquid fuel. The temperature of fuel exiting inlet heat exchanger16can vary depending on the temperature of the fuel and inlet air entering inlet heat exchanger16. Depending on the conditions, the fuel may remain in a liquid state or may vaporize when heated by inlet air. In some embodiments, gaseous fuel exiting inlet heat exchanger16can be directed through fuel line G directly to turbo-generator14to drive fuel turbine32. For example, inlet air temperature for an aircraft operating at Mach 4.5 to 5 may be sufficiently high to produce a gaseous fuel capable of driving fuel turbine32. In other instances, additional heat may be required to fully utilize the cooling and/or work extraction potential of the heated fuel. If additional heat is required, fuel exiting inlet heat exchanger16can be pumped through fuel line F to exhaust heat exchanger30where heat from exhaust gas exiting gas turbine engine12can be transferred to the fuel. Temperature and/or pressure sensor S1can be used determine whether fuel exiting inlet heat exchanger16has a sufficient temperature and/or pressure to drive fuel turbine32. Valves E and G can be opened or closed, accordingly, to direct fuel to exhaust heat exchanger30or fuel turbine32, respectively.

Exhaust heat exchanger30can be a plate-fin, shell-and-tube, or other suitable air-to-liquid heat exchanger as known in the art. In some embodiments, exhaust heat exchanger30can be disposed in an exhaust case wall of gas turbine engine12and heat can be transferred through a wall to fuel circulating in tubing coiled or otherwise distributed around the exhaust case. In other embodiments, exhaust heat exchanger30can be located in a flow path of the exhaust gas FE, as shown in phantom inFIG. 1. Exhaust heat exchanger30is configured to heat fuel from inlet heat exchanger16with waste heat from the exhaust gas of gas turbine engine12. During some operations exhaust gas can have a temperature greater than 1500° F. (816° C.). When augmentor29is in use, exhaust gas temperatures can exceed 3200° F. (1760° C.). In a non-limiting example, fuel exiting exhaust heat exchanger30and entering fuel turbine32can have a temperature of about 1300° F. (704° C.) and pressure of about 515 psi (3,551 kPa). The temperature and pressure of fuel exiting exhaust heat exchanger30can be detected by sensor S2positioned in fuel line J.

Fuel directed to turbo-generator14through fuel lines G or I expands through multi-stage fuel turbine32, driving rotation of fuel turbine32and thereby fuel pump34and motor/generator36, which can be located on common shaft38or otherwise mechanically coupled. Gaseous fuel exiting fuel turbine32can be supplied to combustor24through fuel line J and supplied to augmentor29through fuel line D. Combustor24is the primary burner for gas turbine engine12. Augmentor29adds additional thrust to the engine by burning fuel in a secondary combustor section. Fuel turbine32is configured to maximize turbine work extraction, while providing a fuel pressure sufficient to overcome an operating pressure P3. As such, the pressure of gaseous fuel exiting fuel turbine32must be greater than the pressure P3of compressed air entering combustor. Fuel turbine32can be sized to deliver the gaseous fuel at a pressure greater than P3. In some embodiments, fuel turbine32can include a discharge outlet50to enable discharge of fuel at a higher pressure than complete turbine discharge would provide as one element of a control mechanism to ensure the pressure of fuel delivered to combustor24exceeds P3. Combustor24and augmentor29include suitable fuel injectors for delivering gaseous fuel. Controller42can be used to regulate the amount of fuel delivered to combustor24and augmentor29to maintain optimum operation. The fuel pressure required for augmentor29is much lower than P3(e.g., may be only be 20% of P3) and, therefore, may be extracted from a final outlet of fuel turbine32(i.e., after a final turbine stage).

Fuel turbine32drives fuel pump34and motor/generator36, which are mechanically coupled to fuel turbine shaft38. Fuel pump34produces a continuous cycling of fuel through system10. Motor/generator36can be used to provide power to engine systems and components, including components of system10. In some embodiments, motor/generator36can be used to drive fuel pump34when fuel turbine32is not in operation. In addition, power extracted or input from motor generator36can be varied as one element of a control architecture used to ensure fuel discharge pressure from fuel turbine32is adequate to overcome P3.

Fuel can be routed through varying fuel lines depending on the mode of operation (e.g., startup and in flight) and operating conditions. As previously disclosed, fuel can bypass inlet heat exchanger16during climb and low speed operations when inlet air temperatures are sufficiently low enough to not require inlet air cooling. During high speed operations in which inlet temperatures are high enough to produce a gaseous fuel suitable for driving fuel turbine32, fuel exiting inlet heat exchanger16can bypass exhaust heat exchanger30via fuel line G to be routed directly to fuel turbine32. During engine startup, liquid fuel can be supplied directly to combustor24through fuel line I. During startup, no fuel is routed through inlet or exhaust heat exchangers16,30and, therefore, fuel turbine32is not in operation. In this instance, fuel pump32can be driven by motor/generator36or an alternative power source (not shown). In an alternative embodiment, an auxiliary fuel tank46can supply fuel to combustor through fuel line F during startup. Fuel stored in auxiliary fuel tank46can be the same as fuel stored in primary fuel tank40in a liquid or pressurized gaseous state. Fuel line H can be provided as a safety mechanism in the event of fuel leakage from inlet heat exchanger16. Sensor S4can be used to detect leakage of fuel from inlet heat exchanger16and signal controller42to close valved fuel line C and open bypass fuel line H to redirect fuel to exhaust heat exchanger30. Exhaust heat exchanger30transfers waste heat from exhaust gas to the fuel to vaporize fuel as necessary to drive turbo-generator14. From exhaust heat exchanger30, the gaseous fuel is delivered to fuel turbine32via fuel line J.

In alternative embodiments, an intermediate heat exchanger can be used to avoid bypassing inlet heat exchanger16in the event of fuel leakage.FIG. 2is a schematic diagram of gas turbine engine system50with intermediate inlet heat exchanger52. Intermediate inlet heat exchanger52can be located external to gas turbine engine12or outside of a core air flow to prevent fuel leakage into the core of gas turbine engine12. As illustrated inFIG. 2, fuel can be used to cool a separate inert (non-combustible) working fluid, such as helium or carbon dioxide that can circulate through inlet heat exchanger16. Intermediate inlet heat exchanger52is configured to receive liquid fuel from fuel pump32and place the liquid fuel and the working fluid in thermal communication. Inlet heat exchanger16is configured to receive the working fluid from intermediate inlet heat exchanger52and place the working fluid in thermal communication with the inlet air flow FI, such that thermal energy is transferred indirectly from the inlet air flow FIto the liquid fuel. As illustrated inFIG. 2, working fluid is circulated between intermediate inlet heat exchanger52and inlet heat exchanger16via pump34. Liquid fuel is pumped from tank40via fuel pump34to intermediate inlet heat exchanger52, where the fuel serves as a heat sink for working fluid from inlet heat exchanger16. From intermediate inlet heat exchanger52, fuel is circulated through exhaust heat exchanger30, and expanded through fuel turbine32before being delivered to combustor24or augmentor29for combustion in gas turbine engine12.

With the exception of routing liquid fuel through inlet heat exchanger16, system50can operate in a manner consistent with system10. As previously disclosed and illustrated inFIG. 2, fuel can be routed through various fuel lines depending on operational mode and operating conditions. It will be understood that the various fuel paths disclosed with respect to system10and illustrated inFIG. 1also apply to system50as illustrated inFIG. 2. With the addition of intermediate inlet heat exchanger52, liquid fuel in system50simply bypasses inlet heat exchanger16through fuel line E, which fluidly connects intermediate inlet heat exchanger52with exhaust heat exchanger30.

In another alternative embodiment, an additional intermediate exhaust heat exchanger can be used to indirectly transfer heat from exhaust gas to the fuel.FIG. 3is a schematic diagram of gas turbine engine system60including intermediate exhaust heat exchanger62. Intermediate exhaust heat exchanger62is in fluid communication with the fuel from intermediate inlet heat exchanger52and exhaust heat exchanger30. Intermediate exhaust heat exchanger62is configured to receive the liquid fuel from intermediate inlet heat exchanger52and place the fuel in thermal communication with a separate inert (non-combustible) working fluid, such as helium or carbon dioxide. Exhaust heat exchanger30is configured to receive the working fluid from the intermediate exhaust heat exchanger62and place the working fluid in thermal communication with the exhaust gas FE, such that thermal energy is transferred indirectly from the exhaust gas to the fuel. As illustrated inFIG. 3, working fluid is circulated between intermediate exhaust heat exchanger62and exhaust heat exchanger30via pump64. Fuel is pumped from tank40through intermediate inlet heat exchanger52and through intermediate exhaust heat exchanger62. From intermediate exhaust heat exchanger62, fuel is directed through fuel turbine32and to combustor24or augmentor29for combustion in gas turbine engine12.

With the exception of routing fuel through exhaust heat exchanger30, system60can operate in a manner consistent with system50. As previously disclosed and illustrated inFIG. 3, fuel can be routed through various fuel lines depending on operational mode and operating conditions. It will be understood that the various fuel paths disclosed with respect to systems10and50and illustrated inFIGS. 1 and 2, respectively, also apply to system60as illustrated inFIG. 3. With the addition of intermediate exhaust heat exchanger52, fuel in system60simply bypasses exhaust heat exchanger30through fuel line J, which fluidly connects intermediate exhaust heat exchanger62with fuel turbine32.

The working fluid used in intermediate inlet heat exchanger52is non-flammable fluid capable of removing thermal energy from inlet air and transferring thermal energy to the liquid fuel. The working fluid used in intermediate exhaust heat exchanger62can also be a non-flammable fluid and is capable of removing thermal energy from exhaust gas and transferring thermal energy to the fuel. As previously disclosed, working fluid for both intermediate inlet and intermediate exhaust heat exchangers52,62can include helium or carbon dioxide, but is not limited to the fluids disclosed.

The disclosed systems10,50,60can use plentiful and cleaner burning fuel to enable increases in aircraft maximum speed while controlling effective inlet air temperature T2, allowing continued use of existing fan, compressor, and hot section materials, and generating energy using regenerative (i.e., waste heat) input with reduced impact on turbomachinery sizing, performance, and operability.

Summation

Any relative terms or terms of degree used herein, such as “substantially”, “essentially”, “generally”, “approximately” and the like, should be interpreted in accordance with and subject to any applicable definitions or limits expressly stated herein. In all instances, any relative terms or terms of degree used herein should be interpreted to broadly encompass any relevant disclosed embodiments as well as such ranges or variations as would be understood by a person of ordinary skill in the art in view of the entirety of the present disclosure, such as to encompass ordinary manufacturing tolerance variations, incidental alignment variations, transient alignment or shape variations induced by thermal, rotational or vibrational operational conditions, and the like. Moreover, any relative terms or terms of degree used herein should be interpreted to encompass a range that expressly includes the designated quality, characteristic, parameter or value, without variation, as if no qualifying relative term or term of degree were utilized in the given disclosure or recitation.

Discussion of Possible Embodiments

A gas turbine engine system includes a gas turbine engine and a fuel turbine system. The gas turbine engine includes an air inlet, compressor, combustor, turbine, and heat exchange system. The heat exchange system is configured to transfer thermal energy from an inlet air flow or exhaust air flow to a fuel to produce a gaseous fuel that is used to drive a fuel turbine and fuel pump and used for combustion in the gas turbine engine. The fuel turbine is in fluid communication with the heat exchange system and the combustor and configured to extract energy from expansion of the gaseous fuel. The fuel pump is configured to be driven by the fuel turbine and is in fluid communication with the heat exchanger system.

A further embodiment of the gas turbine engine system, wherein the heat exchange system can include an inlet heat exchange system configured to transfer thermal energy from the inlet air flow to the fuel.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the inlet heat exchange system can include an inlet heat exchanger disposed in the air inlet of the gas turbine engine and in an inlet air flow path, wherein the inlet heat exchanger can be in fluid communication with the fuel pump.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the inlet heat exchange system can include an inlet heat exchanger disposed in the air inlet of the gas turbine engine and in an inlet air flow path, and an intermediate inlet heat exchanger disposed external to the inlet air flow path and in fluid communication with the fuel pump and the inlet heat exchanger. The intermediate inlet heat exchanger can be configured to place the fuel and a first working fluid in a thermal communication and wherein the inlet heat exchanger can be configured to receive the first working fluid from the intermediate inlet heat exchanger and place the first working fluid in thermal communication with the inlet air flow, such that thermal energy is transferred indirectly from the inlet air flow to the fuel.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the heat exchange system can further include an exhaust heat exchange system in fluid communication with the inlet heat exchange system and configured to transfer thermal energy from the exhaust gas flow to the fuel, wherein the fuel turbine can be in fluid communication with the exhaust heat exchange system.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the exhaust heat exchange system can include an exhaust heat exchanger in fluid communication with the inlet heat exchange system.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the exhaust heat exchange system can include an exhaust heat exchanger, and an intermediate exhaust heat exchanger in fluid communication with the inlet heat exchange system and the exhaust heat exchanger. The intermediate exhaust heat exchanger can be configured to place the fuel in thermal communication with a second working fluid, and wherein the exhaust heat exchanger can be configured to receive the second working fluid from the intermediate exhaust heat exchanger and place the second working fluid in thermal communication with the exhaust gas, such that thermal energy is transferred indirectly from the exhaust gas to the fuel.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the fuel turbine can include multiple stages and is configured to produce the gaseous fuel at the second pressure, wherein the second pressure is greater than a pressure of the compressed air flow.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the fuel turbine system can further include a combination motor/generator configured to be driven by the fuel turbine.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the fuel pump and combination motor/generator can be mechanically coupled to a rotor shaft of the fuel turbine.

A further embodiment of the gas turbine engine system of any of the preceding paragraphs, wherein the fuel pump can be in fluid communication with a cryogenic fuel.

A method of operating a gas turbine engine system includes cooling an inlet air flow via an inlet heat exchange system by transferring thermal energy to a liquid fuel to produce a cooled inlet air flow. The cooled inlet air flow is compressed to produce a compressed air for combustion. Energy is extracted from expansion of a gaseous fuel through a multi-stage fuel turbine, which produces a gaseous fuel having a pressure greater than a pressure of the compressed air flow. The fuel turbine is fluidly connected to the inlet heat exchange system. A mixture of the gaseous fuel from an outlet of the fuel turbine and compressed air flow is combusted. A pump driven by the fuel turbine pumps liquid fuel to the inlet heat exchange system.

The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations, additional components, and/or steps:

A further embodiment of the method, wherein the gaseous fuel can be produced by transferring thermal energy to the liquid fuel via the inlet heat exchange system.

A further embodiment of the method of any of the preceding paragraphs can further include heating fuel received from and heated by the inlet heat exchange system via an exhaust heat exchange system to produce the gaseous fuel, wherein the heating process comprises transferring thermal energy from an exhaust gas of the gas turbine engine to the fuel received from the inlet heat exchange system.

A further embodiment of the method of any of the preceding paragraphs can further include pumping the liquid fuel through an inlet heat exchanger of the inlet heat exchange system, wherein the inlet heat exchanger is positioned in a flow path of the inlet air flow.

A further embodiment of the method of any of the preceding paragraphs can further include pumping the liquid fuel through an intermediate inlet heat exchanger in fluid communication with the fuel pump and the inlet heat exchanger, pumping a first working fluid through the inlet heat exchanger and the intermediate inlet heat exchanger, transferring thermal energy from the inlet air flow to the first working fluid, transferring thermal energy from the first working fluid to the liquid fuel.

A further embodiment of the method of any of the preceding paragraphs can further include pumping the fuel received from the intermediate inlet heat exchanger through an intermediate exhaust heat exchanger in fluid communication with the intermediate inlet heat exchanger and the exhaust heat exchanger, pumping a second working fluid through the exhaust heat exchanger and the intermediate exhaust heat exchanger, transferring thermal energy from the exhaust gas to the second working fluid, and transferring thermal energy from the second working fluid to the fuel received from the intermediate inlet heat exchanger.

A further embodiment of the method of any of the preceding paragraphs, wherein the liquid fuel can be cryogenic.

A further embodiment of the method of any of the preceding paragraphs can further include extracting power from the fuel turbine via a combined motor/generator.

A fuel turbine system includes a heat exchange system configured to transfer heat to a liquid fuel to produce a gaseous fuel, a fuel turbine in fluid communication with the heat exchange system and configured to extract energy from expansion of the gaseous fuel, and a fuel pump in fluid communication with the heat exchange system and configured to deliver the liquid fuel to the heat exchange system. The fuel pump is mechanically coupled to the fuel turbine and configured to be driven by the fuel turbine.

The fuel turbine system of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations, and/or additional components:

A further embodiment of the fuel turbine system, wherein the heat exchange system can include an inlet heat exchange system of a gas turbine engine with the inlet heat exchange system configured to cool an inlet air flow, and wherein the fuel turbine can be fluidly connected to a combustor of the gas turbine engine and configured to deliver the gaseous fuel to the combustor.