TURBINE ENGINE ASSEMBLIES

Turbine engine assemblies including a turbine engine assembly having a turbine core comprising a compressor section, a combustion section, a turbine section, and a nozzle section, which are axially aligned, wherein the combustion section comprises a generally annular case having inner and outer walls, a heat exchanger comprising multiple passages in proximity to at least one of the inner and outer walls, with the passages arranged about at least a portion of the case and in fluid communication with each other such that fluid may flow through the passages and a cryogenic fuel system having a cryogenic fuel tank with a supply line coupled to one of the passages, wherein cryogenic fuel may be supplied from the cryogenic fuel tank, through the supply line, to the passages of the heat exchanger, where the fuel in the passages may be heated by the combustion section. The heat exchanger may be a single or multistage vaporizer.

BACKGROUND OF THE INVENTION

The technology described herein relates generally to aircraft systems, and more specifically to aircraft systems using dual fuels in an aviation gas turbine engine and a method of operating same.

Certain cryogenic fuels such as liquefied natural gas (LNG) may be cheaper than conventional jet fuels. Current approaches to cooling in conventional gas turbine applications use compressed air or conventional liquid fuel. Use of compressor air for cooling may lower efficiency of the engine system.

Accordingly, it would be desirable to have aircraft systems using dual fuels in an aviation gas turbine engine. It would be desirable to have aircraft systems that can be propelled by aviation gas turbine engines that can be operated using conventional jet fuel and/or cheaper cryogenic fuels such as liquefied natural gas (LNG). It would be desirable to have more efficient cooling in aviation gas turbine components and systems. It would be desirable to have improved efficiency and lower Specific Fuel Consumption in the engine to lower the operating costs. It is desirable to have aviation gas turbine engines using dual fuels that may reduce environmental impact with lower greenhouse gases (CO2), oxides of nitrogen—NOx, carbon monoxide—CO, unburned hydrocarbons and smoke.

BRIEF DESCRIPTION OF EMBODIMENTS OF THE INVENTION

In one aspect, an embodiment of the invention relates to a turbine engine assembly including a turbine core comprising a compressor section, a combustion section, a turbine section, and a nozzle section, which are axially aligned, wherein the combustion section comprises a generally annular case having inner and outer walls, a heat exchanger comprising multiple passages in proximity to at least one of the inner and outer walls, with the passages arranged about at least a portion of the case and in fluid communication with each other such that fluid may flow through the passages, and a cryogenic fuel system having a cryogenic fuel tank with a supply line coupled to one of the passages, wherein cryogenic fuel may be supplied from the cryogenic fuel tank, through the supply line, to the passages of the heat exchanger, where the fuel in the passages may be heated by the combustion section.

In another aspect, an embodiment of the invention relates to a turbine engine assembly having a turbine core comprising a compressor section, a combustion section, a turbine section, and a nozzle section, which are axially aligned, a cryogenic fuel system having a cryogenic fuel tank with a supply line, and a multistage vaporizer comprising at least one passage fluidly coupled with the supply line such that cryogenic fuel supplied from the cryogenic fuel tank flows through the at least one passage of the multistage vaporizer, where the fuel in at least one passage may be heated.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

Referring to the drawings herein, identical reference numerals denote the same elements throughout the various views.

FIG. 1shows an aircraft system5according to an exemplary embodiment of the present invention. The exemplary aircraft system5has a fuselage6and wings7attached to the fuselage. The aircraft system5has a propulsion system100that produces the propulsive thrust required to propel the aircraft system in flight. Although the propulsion system100is shown attached to the wing7inFIG. 1, in other embodiments it may be coupled to other parts of the aircraft system5, such as, for example, the tail portion16.

The exemplary aircraft system5has a fuel storage system10for storing one or more types of fuels that are used in the propulsion system100. The exemplary aircraft system5shown inFIG. 1uses two types of fuels, as explained further below herein. Accordingly, the exemplary aircraft system5comprises a first fuel tank21capable of storing a first fuel11and a second fuel tank22capable of storing a second fuel12. In the exemplary aircraft system5shown inFIG. 1, at least a portion of the first fuel tank21is located in a wing7of the aircraft system5. In one exemplary embodiment, shown inFIG. 1, the second fuel tank22is located in the fuselage6of the aircraft system near the location where the wings are coupled to the fuselage. In alternative embodiments, the second fuel tank22may be located at other suitable locations in the fuselage6or the wing7. In other embodiments, the aircraft system5may comprise an optional third fuel tank123capable of storing the second fuel12. The optional third fuel tank123may be located in an aft portion of the fuselage of the aircraft system, such as for example shown schematically inFIG. 1.

As further described later herein, the propulsion system100shown inFIG. 1is a dual fuel propulsion system that is capable of generating propulsive thrust by using the first fuel11or the second fuel12or using both first fuel11and the second fuel12. The exemplary dual fuel propulsion system100comprises a gas turbine engine101capable of generating a propulsive thrust selectively using the first fuel11, or the second fuel21, or using both the first fuel and the second fuel at selected proportions. The first fuel may be a conventional liquid fuel such as a kerosene based jet fuel such as known in the art as Jet-A, JP-8, or JP-5 or other known types or grades. In the exemplary embodiments described herein, the second fuel12is a cryogenic fuel that is stored at very low temperatures. In one embodiment described herein, the cryogenic second fuel12is Liquefied Natural Gas (alternatively referred to herein as “LNG”). The cryogenic second fuel12is stored in the fuel tank at a low temperature. For example, the LNG is stored in the second fuel tank22at about −265° F. at an absolute pressure of about 15 psia. The fuel tanks may be made from known materials such as titanium, Inconel, aluminum or composite materials.

The exemplary aircraft system5shown inFIG. 1comprises a fuel delivery system50capable of delivering a fuel from the fuel storage system10to the propulsion system100. Known fuel delivery systems may be used for delivering the conventional liquid fuel, such as the first fuel11. In the exemplary embodiments described herein, and shown inFIGS. 1 and 2, the fuel delivery system50is configured to deliver a cryogenic liquid fuel, such as, for example, LNG, to the propulsion system100through conduits54that transport the cryogenic fuel. In order to substantially maintain a liquid state of the cryogenic fuel during delivery, at least a portion of the conduit54of the fuel delivery system50is insulated and configured for transporting a pressurized cryogenic liquid fuel. In some exemplary embodiments, at least a portion of the conduit54has a double wall construction. The conduits may be made from known materials such as titanium, Inconel, aluminum or composite materials.

The exemplary embodiment of the aircraft system5shown inFIG. 1further includes a fuel cell system400, comprising a fuel cell capable of producing electrical power using at least one of the first fuel11or the second fuel12. The fuel delivery system50is capable of delivering a fuel from the fuel storage system10to the fuel cell system400. In one exemplary embodiment, the fuel cell system400generates power using a portion of a cryogenic fuel12used by a dual fuel propulsion system100.

The propulsion system100comprises a gas turbine engine101that generates the propulsive thrust by burning a fuel in a combustor.FIG. 4is a schematic view of an exemplary gas turbine engine101including a fan103and a core engine108having a high pressure compressor105, and a combustor90. Engine101also includes a high pressure turbine155, a low pressure turbine157, and a booster104. The exemplary gas turbine engine101has a fan103that produces at least a portion of the propulsive thrust. Engine101has an intake side109and an exhaust side110. Fan103and turbine157are coupled together using a first rotor shaft114, and compressor105and turbine155are coupled together using a second rotor shaft115. In some applications, such as, for example, shown inFIG. 4, the fan103blade assemblies are at least partially positioned within an engine casing116. In other applications, the fan103may form a portion of an “open rotor” where there is no casing surrounding the fan blade assembly.

During operation, air flows axially through fan103, in a direction that is substantially parallel to a central line axis15extending through engine101, and compressed air is supplied to high pressure compressor105. The highly compressed air is delivered to combustor90. Hot gases (not shown inFIG. 4) from combustor90drives turbines155and157. Turbine157drives fan103by way of shaft114and similarly, turbine155drives compressor105by way of shaft115. In alternative embodiments, the engine101may have an additional compressor, sometimes known in the art as an intermediate pressure compressor, driven by another turbine stage (not shown inFIG. 4).

During operation of the aircraft system5(See exemplary flight profile shown inFIG. 7), the gas turbine engine101in the propulsion system100may use, for example, the first fuel11during a first selected portion of operation of propulsion system, such as for example, during take off. The propulsion system100may use the second fuel12, such as, for example, LNG, during a second selected portion of operation of propulsion system such as during cruise. Alternatively, during selected portions of the operation of the aircraft system5, the gas turbine engine101is capable of generating the propulsive thrust using both the first fuel11and the second fuel12simultaneously. The proportion of the first fuel and second fuel may be varied between 0% to 100% as appropriate during various stages of the operation of the propulsion system.

An aircraft and engine system, described herein, is capable of operation using two fuels, one of which may be a cryogenic fuel such as for example, LNG (liquefied natural gas), the other a conventional kerosene based jet fuel such as Jet-A, JP-8, JP-5 or similar grades available worldwide.

The Jet-A fuel system is similar to conventional aircraft fuel systems, with the exception of the fuel nozzles, which are capable of firing Jet-A and cryogenic/LNG to the combustor in proportions from 0-100%. In the embodiment shown inFIG. 1, the LNG system includes a fuel tank, which optionally contains the following features: (i) vent lines with appropriate check valves to maintain a specified pressure in the tank; (ii) drain lines for the liquid cryogenic fuel; (iii) gauging or other measurement capability to assess the temperature, pressure, and volume of cryogenic (LNG) fuel present in the tank; (iv) a boost pump located in the cryogenic (LNG) tank or optionally outside of the tank, which increases the pressure of the cryogenic (LNG) fuel to transport it to the engine; and (iv) an optional cryo-cooler to keep the tank at cryogenic temperatures indefinitely.

The fuel tank will preferably operate at or near atmospheric pressure, but can operate in the range of 0 to 100 psig. Alternative embodiments of the fuel system may include high tank pressures and temperatures. The cryogenic (LNG) fuel lines running from the tank and boost pump to the engine pylons may have the following features: (i) single or double wall construction; (ii) vacuum insulation or low thermal conductivity material insulation; and (iii) an optional cryo-cooler to re-circulate LNG flow to the tank without adding heat to the LNG tank. The cryogenic (LNG) fuel tank can be located in the aircraft where a conventional Jet-A auxiliary fuel tank is located on existing systems, for example, in the forward or aft cargo hold. Alternatively, a cryogenic (LNG) fuel tank can be located in the center wing tank location. An auxiliary fuel tank utilizing cryogenic (LNG) fuel may be designed so that it can be removed if cryogenic (LNG) fuel will not be used for an extended period of time.

A high pressure pump may be located in the pylon or on board the engine to raise the pressure of the cryogenic (LNG) fuel to levels sufficient to inject fuel into the gas turbine combustor. The pump may or may not raise the pressure of the LNG/cryogenic liquid above the critical pressure (Pc) of cryogenic (LNG) fuel. A heat exchanger, referred to herein as a “vaporizer,” which may be mounted on or near the engine, adds thermal energy to the liquefied natural gas fuel, raising the temperature and volumetrically expanding the cryogenic (LNG) fuel. Heat (thermal energy) from the vaporizer can come from many sources. These include, but are not limited to: (i) the gas turbine exhaust; (ii) compressor intercooling; (iii) high pressure and/or low pressure turbine clearance control air; (iv) LPT pipe cooling parasitic air; (v) cooled cooling air from the HP turbine; (vi) lubricating oil; or (vii) on board avionics or electronics. The heat exchanger can be of various designs, including shell and tube, double pipe, fin plate, etc., and can flow in a co-current, counter current, or cross current manner. Heat exchange can occur in direct or indirect contact with the heat sources listed above.

A control valve is located downstream of the vaporizer/heat exchange unit described above. The purpose of the control valve is to meter the flow to a specified level into the fuel manifold across the range of operational conditions associated with the gas turbine engine operation. A secondary purpose of the control valve is to act as a back pressure regulator, setting the pressure of the system above the critical pressure of cryogenic (LNG) fuel.

A fuel manifold is located downstream of the control valve, which serves to uniformly distribute gaseous fuel to the gas turbine fuel nozzles. In some embodiments, the manifold can optionally act as a heat exchanger, transferring thermal energy from the core cowl compartment or other thermal surroundings to the cryogenic/LNG/natural gas fuel. A purge manifold system can optionally be employed with the fuel manifold to purge the fuel manifold with compressor air (CDP) when the gaseous fuel system is not in operation. This will prevent hot gas ingestion into the gaseous fuel nozzles due to circumferential pressure variations. Optionally, check valves in or near the fuel nozzles can prevent hot gas ingestion.

An exemplary embodiment of the system described herein may operate as follows: Cryogenic (LNG) fuel is located in the tank at about 15 psia and about −265° F. It is pumped to approximately 30 psi by the boost pump located on the aircraft. Liquid cryogenic (LNG) fuel flows across the wing via insulated double walled piping to the aircraft pylon where it is stepped up to about 100 to 1,500 psia and can be above or below the critical pressure of natural gas/methane. The cryogenic (LNG) fuel is then routed to the vaporizer where it volumetrically expands to a gas. The vaporizer may be sized to keep the Mach number and corresponding pressure losses low. Gaseous natural gas is then metered though a control valve and into the fuel manifold and fuel nozzles where it is combusted in an otherwise standard aviation gas turbine engine system, providing thrust to the airplane. As cycle conditions change, the pressure in the boost pump (about 30 psi for example) and the pressure in the HP pump (about 1,000 psi for example) are maintained at an approximately constant level. Flow is controlled by the metering valve. The variation in flow in combination with the appropriately sized fuel nozzles result in acceptable and varying pressures in the manifold.

The exemplary aircraft system5has a fuel delivery system for delivering one or more types of fuels from the storage system10for use in the propulsion system100. For a conventional liquid fuel such as, for example, a kerosene based jet fuel, a conventional fuel delivery system may be used. The exemplary fuel delivery system described herein, and shown schematically inFIGS. 2 and 3, comprises a cryogenic fuel delivery system50for an aircraft system5. The exemplary fuel system50shown inFIG. 2comprises a cryogenic fuel tank122capable of storing a cryogenic liquid fuel112. In one embodiment, the cryogenic liquid fuel112is LNG. Other alternative cryogenic liquid fuels may also be used. In the exemplary fuel system50, the cryogenic liquid fuel112, such as, for example, LNG, is at a first pressure “P1”. The pressure P1is preferably close to atmospheric pressure, such as, for example, 15 psia.

The exemplary fuel system50has a boost pump52such that it is in flow communication with the cryogenic fuel tank122. During operation, when cryogenic fuel is needed in the dual fuel propulsion system100, the boost pump52removes a portion of the cryogenic liquid fuel112from the cryogenic fuel tank122and increases its pressure to a second pressure “P2” and flows it into a wing supply conduit54located in a wing7of the aircraft system5. The pressure P2is chosen such that the liquid cryogenic fuel maintains its liquid state (L) during the flow in the supply conduit54. The pressure P2may be in the range of about 30 psia to about 40 psia. Based on analysis using known methods, for LNG, 30 psia is found to be adequate. The boost pump52may be located at a suitable location in the fuselage6of the aircraft system5. Alternatively, the boost pump52may be located close to the cryogenic fuel tank122. In other embodiments, the boost pump52may be located inside the cryogenic fuel tank122. In order to substantially maintain a liquid state of the cryogenic fuel during delivery, at least a portion of the wing supply conduit54is insulated. In some exemplary embodiments, at least a portion of the conduit54has a double wall construction. The conduits54and the boost pump52may be made using known materials such as titanium, Inconel, aluminum or composite materials.

The exemplary fuel system50has a high-pressure pump58that is in flow communication with the wing supply conduit54and is capable of receiving the cryogenic liquid fuel112supplied by the boost pump52. The high-pressure pump58increases the pressure of the liquid cryogenic fuel (such as, for example, LNG) to a third pressure “P3” sufficient to inject the fuel into the propulsion system100. The pressure P3may be in the range of about 100 psia to about 1000 psia. The high-pressure pump58may be located at a suitable location in the aircraft system5or the propulsion system100. The high-pressure pump58is preferably located in a pylon55of aircraft system5that supports the propulsion system100.

As shown inFIG. 2, the exemplary fuel system50has a vaporizer60for changing the cryogenic liquid fuel112into a gaseous (G) fuel13. The vaporizer60receives the high pressure cryogenic liquid fuel and adds heat (thermal energy) to the cryogenic liquid fuel (such as, for example, LNG) raising its temperature and volumetrically expanding it. Heat (thermal energy) can be supplied from one or more sources in the propulsion system100. For example, heat for vaporizing the cryogenic liquid fuel in the vaporizer may be supplied from one or more of several sources, such as, for example, the gas turbine exhaust99, compressor105, high pressure turbine155, low pressure turbine157, fan bypass107, turbine cooling air, lubricating oil in the engine, aircraft system avionics/electronics, or any source of heat in the propulsion system100. Due to the exchange of heat that occurs in the vaporizer60, the vaporizer60may be alternatively referred to as a heat exchanger. The heat exchanger portion of the vaporizer60may include a shell and tube type heat exchanger, or a double pipe type heat exchanger, or fin-and-plate type heat exchanger. The hot fluid and cold fluid flow in the vaporizer may be co-current, or counter-current, or a cross current flow type. The heat exchange between the hot fluid and the cold fluid in the vaporizer may occur directly through a wall or indirectly, using an intermediate work fluid.

The cryogenic fuel delivery system50comprises a flow metering valve65(“FMV”, also referred to as a Control Valve) that is in flow communication with the vaporizer60and a manifold70. The flow metering valve65is located downstream of the vaporizer/heat exchange unit described above. The purpose of the FMV (control valve) is to meter the fuel flow to a specified level into the fuel manifold70across the range of operational conditions associated with the gas turbine engine operation. A secondary purpose of the control valve is to act as a back pressure regulator, setting the pressure of the system above the critical pressure of the cryogenic fuel such as LNG. The flow metering valve65receives the gaseous fuel13supplied from the vaporizer and reduces its pressure to a fourth pressure “P4”. The manifold70is capable of receiving the gaseous fuel13and distributing it to a fuel nozzle80in the gas turbine engine101. In a preferred embodiment, the vaporizer60changes the cryogenic liquid fuel112into the gaseous fuel13at a substantially constant pressure.FIG. 2aschematically shows the state and pressure of the fuel at various points in the delivery system50.

The cryogenic fuel delivery system50further comprises a plurality of fuel nozzles80located in the gas turbine engine101. The fuel nozzle80delivers the gaseous fuel13into the combustor90for combustion. The fuel manifold70, located downstream of the control valve65, serves to uniformly distribute gaseous fuel13to the gas turbine fuel nozzles80. In some embodiments, the manifold70can optionally act as a heat exchanger, transferring thermal energy from the propulsion system core cowl compartment or other thermal surroundings to the LNG/natural gas fuel. In one embodiment, the fuel nozzle80is configured to selectively receive a conventional liquid fuel (such as the conventional kerosene based liquid fuel) or the gaseous fuel13generated by the vaporizer from the cryogenic liquid fuel such as LNG. In another embodiment, the fuel nozzle80is configured to selectively receive a liquid fuel and the gaseous fuel13and configured to supply the gaseous fuel13and a liquid fuel to the combustor90to facilitate co-combustion of the two types of fuels. In another embodiment, the gas turbine engine101comprises a plurality of fuel nozzles80wherein some of the fuel nozzles80are configured to receive a liquid fuel and some of the fuel nozzles80are configured to receive the gaseous fuel13and arranged suitably for combustion in the combustor90.

In another embodiment of the present invention, fuel manifold70in the gas turbine engine101comprises an optional purge manifold system to purge the fuel manifold with compressor air, or other air, from the engine when the gaseous fuel system is not in operation. This will prevent hot gas ingestion into the gaseous fuel nozzles due to circumferential pressure variations in the combustor90. Optionally, check valves in or near the fuel nozzles can be used prevent hot gas ingestion in the fuel nozzles or manifold.

In an exemplary dual fuel gas turbine propulsion system described herein that uses LNG as the cryogenic liquid fuel is described as follows: LNG is located in the tank22,122at 15 psia and −265° F. It is pumped to approximately 30 psi by the boost pump52located on the aircraft. Liquid LNG flows across the wing7via insulated double walled piping54to the aircraft pylon55where it is stepped up to 100 to 1,500 psia and may be above or below the critical pressure of natural gas/methane. The Liquefied Natural Gas is then routed to the vaporizer60where it volumetrically expands to a gas. The vaporizer60is sized to keep the Mach number and corresponding pressure losses low. Gaseous natural gas is then metered though a control valve65and into the fuel manifold70and fuel nozzles80where it is combusted in an dual fuel aviation gas turbine system100,101, providing thrust to the aircraft system5. As cycle conditions change, the pressure in the boost pump (30 psi) and the pressure in the HP pump58(1,000 psi) are maintained at an approximately constant level. Flow is controlled by the metering valve65. The variation in flow in combination with the appropriately sized fuel nozzles result in acceptable and varying pressures in the manifold.

The dual fuel system consists of parallel fuel delivery systems for kerosene based fuel (Jet-A, JP-8, JP-5, etc) and a cryogenic fuel (LNG for example). The kerosene fuel delivery is substantially unchanged from the current design, with the exception of the combustor fuel nozzles, which are designed to co-fire kerosene and natural gas in any proportion. As shown inFIG. 2, the cryogenic fuel (LNG for example) fuel delivery system consists of the following features: (A) A dual fuel nozzle and combustion system, capable of utilizing cryogenic fuel (LNG for example), and Jet-A in any proportion from 0- to 100%; (B) A fuel manifold and delivery system that also acts as a heat exchanger, heating cryogenic fuel (LNG for example) to a gas or a supercritical fluid. The manifold system is designed to concurrently deliver fuel to the combustor fuel nozzles in a uniform manner, and absorb heat from the surrounding core cowl, exhaust system, or other heat source, eliminating or minimizing the need for a separate heat exchanger; (C) A fuel system that pumps up cryogenic fuel (LNG for example) in its liquid state above or below the critical pressure and adds heat from any of a number of sources; (D) A low pressure cryo-pump submerged in the cryogenic fuel (LNG for example) fuel tank (optionally located outside the fuel tank.); (E) A high pressure cryo-pump located in the aircraft pylon or optionally on board the engine or nacelle to pump to pressures above the critical pressure of cryogenic fuel (LNG for example). (F) A purge manifold system can optionally employed with the fuel manifold to purge the fuel manifold with compressor CDP air when the gaseous fuel system is not in operation. This will prevent hot gas ingestion into the gaseous fuel nozzles due to circumferential pressure variations. Optionally, check valves in or near the fuel nozzles can prevent hot gas ingestion. (G) cryogenic fuel (LNG for example) lines running from the tank and boost pump to the engine pylons have the following features: (1) Single or double wall construction. (2) Vacuum insulation or optionally low thermal conductivity insulation material such as aerogels. (3) An optional cryo-cooler to recirculate cryogenic fuel (LNG for example) flow to the tank without adding heat to the cryogenic fuel (LNG for example) tank. (H) A high pressure pump located in the pylon or on board the engine. This pump will raise the pressure of the cryogenic fuel (LNG for example) to levels sufficient to inject natural gas fuel into the gas turbine combustor. The pump may or may not raise the pressure of the cryogenic liquid (LNG for example) above the critical pressure (Pc) of cryogenic fuel (LNG for example).

III. A Fuel Storage System

The exemplary aircraft system5shown inFIG. 1comprises a cryogenic fuel storage system10, such as shown for example, inFIG. 3, for storing a cryogenic fuel. The exemplary cryogenic fuel storage system10comprises a cryogenic fuel tank22,122having a first wall23forming a storage volume24capable of storing a cryogenic liquid fuel12such as for example LNG. As shown schematically inFIG. 3, the exemplary cryogenic fuel storage system10has an inflow system32capable of flowing the cryogenic liquid fuel12into the storage volume24and an outflow system30adapted to deliver the cryogenic liquid fuel12from the cryogenic fuel storage system10. It further comprises a vent system40capable of removing at least a portion of a gaseous fuel19(that may be formed during storage) from the cryogenic liquid fuel12in the storage volume24.

The exemplary cryogenic fuel storage system10shown inFIG. 3further comprises a recycle system34that is adapted to return at least a portion29of unused gaseous fuel19into the cryogenic fuel tank22. In one embodiment, the recycle system34comprises a cryo-cooler42that cools the portion29of unused gaseous fuel19prior to returning it into the cryogenic fuel tank22,122. An exemplary operation of the cryo-cooler42operation is as follows: In an exemplary embodiment, boil off from the fuel tank can be re-cooled using a reverse Rankine refrigeration system, also known as a cryo-cooler. The cryo-cooler can be powered by electric power coming from any of the available systems on board the aircraft system5, or, by ground based power systems such as those, which may be available while parked at a boarding gate. The cryo-cooler system can also be used to re-liquefy natural gas in the fuel system during the dual fuel aircraft gas turbine engine101co-fire transitions.

The fuel storage system10may further comprise a safety release system45adapted to vent any high pressure gases that may be formed in the cryogenic fuel tank22. In one exemplary embodiment, shown schematically inFIG. 3, the safety release system45comprises a rupture disk46that forms a portion of the first wall23. The rupture disk46is a safety feature, designed using known methods, to blow out and release any high pressure gases in the event of an over pressure inside the fuel tank22.

The cryogenic fuel tank22may have a single wall construction or a multiple wall construction. For example, the cryogenic fuel tank22may further comprise (SeeFIG. 3for example) a second wall25that substantially encloses the first wall23. In one embodiment of the tank, there is a gap26between the first wall23and the second wall25in order to thermally insulate the tank to reduce heat flow across the tank walls. In one exemplary embodiment, there is a vacuum in the gap26between the first wall23and the second wall25. The vacuum may be created and maintained by a vacuum pump28. Alternatively, in order to provide thermal insulation for the tank, the gap26between the first wall23and the second wall25may be substantially filled with a known thermal insulation material27, such as, for example, Aerogel. Other suitable thermal insulation materials may be used. Baffles17may be included to control movement of liquid within the tank.

The cryogenic fuel storage system10shown inFIG. 3comprises the outflow system30having a delivery pump31. The delivery pump may be located at a convenient location near the tank22. In order to reduce heat transfer in to the cryogenic fuel, it may be preferable to locate the delivery pump31in the cryogenic fuel tank22as shown schematically inFIG. 3. The vent system40vents any gases that may be formed in the fuel tank22. These vented gases may be utilized in several useful ways in the aircraft system5. A few of these are shown schematically inFIG. 3. For example at least a portion of the gaseous fuel19may be supplied to the aircraft propulsion system100for cooling or combustion in the engine. In another embodiment, the vent system40supplies at least a portion of the gaseous fuel19to a burner and further venting the combustion products from the burner safely outside the aircraft system5. In another embodiment the vent system40supplies at least a portion of the gaseous fuel19to an auxiliary power unit180that supplies auxiliary power to the aircraft system5. In another embodiment the vent system40supplies at least a portion of the gaseous fuel19to a fuel cell182that produces power. In another embodiment the vent system40releases at least a portion of the gaseous fuel19outside the cryogenic fuel tank22.

According to an embodiment of the invention, foam stabilizers may be added to the cryogenic fuel delivery system50to minimize pressure pulses and flow instabilities in the fluid circuits allowing safe engine operation and enhanced system life. Foam stabilizers may also improve the vaporization stability of the LNG mixture by keeping the boiling process out of the film-boiling regime and by creating a pressure loss mechanism to isolate the upstream pump from the downstream fuel nozzles.

Typically, foam stabilizers may be positioned in transfer lines and in components in which it is beneficial to minimize pressure pulses and flow instabilities. The foam stabilizers of embodiments of the invention may be used in single or dual fuel engines.

In general, the foam stabilizers may include, but are not limited to solid materials having an open or closed cellular structure that have a large volume fraction of gas-filled pores. The pores may form an interconnected network that allows fluids to pass through it. The high surface area and turbulence created by the ligament structures of the foams may prevent or reduce the formation of a vapor film along the walls of a fluid passage.

Foam stabilizers may include, but are not limited to metal or composite materials, or a combination thereof. Metal foam stabilizers typically have high porosity that allows for a very lightweight material. For example, metals including, but not limited to aluminum, titanium, and tantalum may be used as foam stabilizers. According to an embodiment of the invention, foam stabilizers may be constructed by braising sheets of metal on either side of the foam, thereby creating a fluid passage for LNG.

The density and pore size of the foam stabilizer may be varied to achieve optimum system performance. For example, a foam stabilizer according to embodiments of the invention may have a density of about 0.1 to about 1.5 g/cm3 or about 0.4 to about 0.9 g/cm3. The pore size of the foam stabilizer may be about 0.5 to about 15 mm or about 1 to about 8 mm.

The exemplary operation of the fuel storage system, its components including the fuel tank, and exemplary sub systems and components is described as follows.

Natural gas exists in liquid form (LNG) at temperatures of approximately about −260° F. and atmospheric pressure. To maintain these temperatures and pressures on board a passenger, cargo, military, or general aviation aircraft, the features identified below, in selected combinations, allow for safe, efficient, and cost effective storage of LNG. Referring toFIG. 3, these include:

(A) A fuel tank21,22constructed of alloys such as, but not limited to, aluminum AL 5456 and higher strength aluminum AL 5086 or other suitable alloys.

(B) A fuel tank21,22constructed of light weight composite material.

(C) The above tanks21,22with a double wall vacuum feature for improved insulation and greatly reduced heat flow to the LNG fluid. The double walled tank also acts as a safety containment device in the rare case where the primary tank is ruptured.

(D) An alternative embodiment of either the above utilizing lightweight insulation27, such as, for example, Aerogel, to minimize heat flow from the surroundings to the LNG tank and its contents. Aerogel insulation can be used in addition to, or in place of a double walled tank design.

(E) An optional vacuum pump28designed for active evacuation of the space between the double walled tank. The pump can operate off of LNG boil off fuel, LNG, Jet-A, electric power or any other power source available to the aircraft.

(F) An LNG tank with a cryogenic pump31submerged inside the primary tank for reduced heat transfer to the LNG fluid. It is contemplated that the pump may be driven by an electric motor, which is co-located with the pump inside the tank; electric motor losses may be dissipated within the LNG, thereby helping to pressure the tank with additional boil off.

(G) An LNG tank with one or more drain lines36capable of removing LNG from the tank under normal or emergency conditions. The LNG drain line36is connected to a suitable cryogenic pump to increase the rate of removal beyond the drainage rate due to the LNG gravitational head.

(H) An LNG tank with one or more vent lines41for removal of gaseous natural gas, formed by the absorption of heat from the external environment. This vent line41system maintains the tank at a desired pressure by the use of a 1 way relief valve or back pressure valve39.

(I) An LNG tank with a parallel safety relief system45to the main vent line, should an overpressure situation occur. A burst disk is an alternative feature or a parallel feature46. The relief vent would direct gaseous fuel overboard.

A similar parallel safety relief system47may be installed for the vacuum-insulating space enveloping the cryogenic fuel tank in the event that the tank wall might rupture, thereby spilling fuel inventory into the vacuum space and flash vaporizing the spilled fuel such that a catastrophic overpressure pulse could result if the additional safety relief system were otherwise absent.

(J) An LNG fuel tank, with some or all of the design features above, whose geometry is designed to conform to the existing envelope associated with a standard Jet-A auxiliary fuel tank such as those designed and available on commercially available aircrafts.

(K) An LNG fuel tank, with some or all of the design features above, whose geometry is designed to conform to and fit within the lower cargo hold(s) of conventional passenger and cargo aircraft such as those found on commercially available aircrafts.

(L) Modifications to the center wing tank22of an existing or new aircraft to properly insulate the LNG, tank, and structural elements.

(M) An LNG fuel tank, with some or all of the design features above, whose geometry is designed to conform to and fit within chines, wing-mounted pods, or other aerodynamic structures external to the airframe of military aircraft or helicopters.

Venting and boil off systems are designed using known methods. Boil off of LNG is an evaporation process, which absorbs energy and cools the tank and its contents. Boil off LNG can be utilized and/or consumed by a variety of different processes, in some cases providing useful work to the aircraft system, in other cases, simply combusting the fuel for a more environmentally acceptable design. For example, vent gas from the LNG tank consists primarily of methane and is used for any or all combinations of the following:

(A) Routing to the Aircraft APU (Auxiliary Power Unit)180. As shown inFIG. 3, a gaseous vent line from the tank is routed in series or in parallel to an Auxiliary Power Unit for use in the combustor. The APU can be an existing APU, typically found aboard commercial and military aircraft, or a separate APU dedicated to converting natural gas boil off to useful electric and/or mechanical power. A boil off natural gas compressor is utilized to compress the natural gas to the appropriate pressure required for utilization in the APU. The APU, in turn, provides electric power to any system on the engine or A/C.

(B) Routing to one or more aircraft gas turbine engine(s)101. As shown inFIG. 3, a natural gas vent line from the LNG fuel tank is routed to one or more of the main gas turbine engines101and provides an additional fuel source to the engine during operation. A natural gas compressor is utilized to pump the vent gas to the appropriate pressure required for utilization in the aircraft gas turbine engine.

(C) Flared. As shown inFIG. 3, a natural gas vent line from the tank is routed to a small, dedicated vent combustor190with its own electric spark ignition system. In this manner methane gas is not released to the atmosphere. The products of combustion are vented, which results in a more environmentally acceptable system.

(D) Vented. As shown inFIG. 3, a natural gas vent line from the tank is routed to the exhaust duct of one or more of the aircraft gas turbines. Alternatively, the vent line can be routed to the APU exhaust duct or a separate dedicated line to any of the aircraft trailing edges. Natural gas may be suitably vented to atmosphere at one or more of these locations V.

(E) Ground operation. As shown inFIG. 3, during ground operation, any of the systems can be designed such that a vent line41is attached to ground support equipment, which collects and utilizes the natural gas boil off in any ground based system. Venting can also take place during refueling operations with ground support equipment that can simultaneously inject fuel into the aircraft LNG tank using an inflow system32and capture and reuse vent gases (simultaneous venting and fueling indicated as (S) inFIG. 3).

FIG. 4shows an exemplary dual fuel propulsion system100comprising a gas turbine engine101capable of generating a propulsive thrust using a cryogenic liquid fuel112. The gas turbine engine101comprises a compressor105driven by a high-pressure turbine155and a combustor90that burns a fuel and generates hot gases that drive the high-pressure turbine155. The combustor90is capable of burning a conventional liquid fuel such as kerosene based fuel. The combustor90is also capable of burning a cryogenic fuel, such as, for example, LNG, that has been suitably prepared for combustion, such as, for example, by a vaporizer60.FIG. 4shows schematically a vaporizer60capable of changing the cryogenic liquid fuel112into a gaseous fuel13. The dual fuel propulsion system100gas turbine engine101further comprises a fuel nozzle80that supplies the gaseous fuel13to the combustor90for ignition. In one exemplary embodiment, the cryogenic liquid fuel112used is Liquefied Natural Gas (LNG). In a turbo-fan type dual fuel propulsion system100(shown inFIG. 4for example) the gas turbine engine101comprises a fan103located axially forward from the high-pressure compressor105. A booster104(shown inFIG. 4) may be located axially between the fan103and the high-pressure compressor105wherein the fan and booster are driven by a low-pressure turbine157. In other embodiments, the dual fuel propulsion system100gas turbine engine101may include an intermediate pressure compressor driven by an intermediate pressure turbine (both not shown inFIG. 4). The booster104(or an intermediate pressure compressor) increases the pressure of the air that enters the compressor105and facilitates the generation of higher pressure ratios by the compressor105. In the exemplary embodiment shown inFIG. 4, the fan and the booster are driven by the low pressure turbine157, and the high pressure compressor is driven the high pressure turbine155.

The vaporizer60, shown schematically inFIG. 4, is mounted on or near the engine101. One of the functions of the vaporizer60is to add thermal energy to the cryogenic fuel, such as the liquefied natural gas (LNG) fuel, raising its temperature. In this context, the vaporizer functions as heat exchanger. Another, function of the vaporizer60is to volumetrically expand the cryogenic fuel, such as the liquefied natural gas (LNG) fuel to a gaseous form for later combustion. Heat (thermal energy) for use in the vaporizer60can come from or more of many sources in the propulsion system100and aircraft system5. These include, but are not limited to: (i) The gas turbine exhaust, (ii) Compressor intercooling, (iii) High pressure and/or low pressure turbine clearance control air, (iv) LPT pipe cooling parasitic air, (v) cooling air used in the High pressure and/or low pressure turbine, (vi) Lubricating oil, and (vii) On board avionics, electronics in the aircraft system5. The heat for the vaporizer may also be supplied from the compressor105, booster104, intermediate pressure compressor (not shown) and/or the fan bypass air stream107(SeeFIG. 4). An exemplary embodiment using a portion of the discharge air from the compressor105is shown inFIG. 5. A portion of the compressor discharge air2is bled out to the vaporizer60, as shown by item3inFIG. 5. The cryogenic liquid fuel21, such as for example, LNG, enters vaporizer60wherein the heat from the airflow stream3is transferred to the cryogenic liquid fuel21. In one exemplary embodiment, the heated cryogenic fuel is further expanded, as described previously herein, producing gaseous fuel13in the vaporizer60. The gaseous fuel13is then introduced into combustor90using a fuel nozzle80(SeeFIG. 5). The cooled airflow4that exits from the vaporizer can be used for cooling other engine components, such as the combustor90structures and/or the high-pressure turbine155structures. The heat exchanger portion in the vaporizer60can be of a known design, such as for example, shell and tube design, double pipe design, and/or fin plate design. The fuel112flow direction and the heating fluid96direction in the vaporizer60(seeFIG. 4) may be in a co-current direction, counter-current direction, or they may flow in a cross-current manner to promote efficient heat exchange between the cryogenic fuel and the heating fluid.

Heat exchange in the vaporizer60can occur in direct manner between the cryogenic fuel and the heating fluid, through a metallic wall.FIG. 5shows schematically a direct heat exchanger in the vaporizer60.FIG. 6ashows schematically an exemplary direct heat exchanger63that uses a portion97of the gas turbine engine101exhaust gas99to heat the cryogenic liquid fuel112. Alternatively, heat exchange in the vaporizer60can occur in an indirect manner between the cryogenic fuel and the heat sources listed above, through the use of an intermediate heating fluid.FIG. 6bshows an exemplary vaporizer60that uses an indirect heat exchanger64that uses an intermediary heating fluid68to heat the cryogenic liquid fuel112. In such an indirect heat exchanger shown inFIG. 6b, the intermediary heating fluid68is heated by a portion97of the exhaust gas99from the gas turbine engine101. Heat from the intermediary heating fluid68is then transferred to the cryogenic liquid fuel112.FIG. 6cshows another embodiment of an indirect exchanger used in a vaporizer60. In this alternative embodiment, the intermediary heating fluid68is heated by a portion of a fan bypass stream107of the gas turbine engine101, as well as a portion97of the engine exhaust gas99. The intermediary heating fluid68then heats the cryogenic liquid fuel112. A control valve38is used to control the relative heat exchanges between the flow streams.

(V) Method of Operating Dual Fuel Aircraft System

An exemplary method of operation of the aircraft system5using a dual fuel propulsion system100is described as follows with respect to an exemplary flight mission profile shown schematically inFIG. 7. The exemplary flight mission profile shown schematically inFIG. 7shows the Engine power setting during various portions of the flight mission identified by the letter labels A-B-C-D-E- . . . -X-Y etc. For example, A-B represents the start, B-C shows ground-idle, G-H shows take-off, T-L and O-P show cruise, etc. During operation of the aircraft system5(See exemplary flight profile120inFIG. 7), the gas turbine engine101in the propulsion system100may use, for example, the first fuel11during a first selected portion of operation of propulsion system, such as for example, during take off. The propulsion system100may use the second fuel12, such as, for example, LNG, during a second selected portion of operation of propulsion system such as during cruise. Alternatively, during selected portions of the operation of the aircraft system5, the gas turbine engine101is capable of generating the propulsive thrust using both the first fuel11and the second fuel12simultaneously. The proportion of the first fuel and second fuel may be varied between 0% to 100% as appropriate during various stages of the operation of the dual fuel propulsion system100.

An exemplary method of operating a dual fuel propulsion system100using a dual fuel gas turbine engine101comprises the following steps of: starting the aircraft engine101(see A-B inFIG. 7) by burning a first fuel11in a combustor90that generates hot gases that drive a gas turbine in the engine101. The first fuel11may be a known type of liquid fuel, such as a kerosene based Jet Fuel. The engine101, when started, may produce enough hot gases that may used to vaporize a second fuel, such as, for example, a cryogenic fuel. A second fuel12is then vaporized using heat in a vaporizer60to form a gaseous fuel13. The second fuel may be a cryogenic liquid fuel112, such as, for example, LNG. The operation of an exemplary vaporizer60has been described herein previously. The gaseous fuel13is then introduced into the combustor90of the engine101using a fuel nozzle80and the gaseous fuel13is burned in the combustor90that generates hot gases that drive the gas turbine in the engine. The amount of the second fuel introduced into the combustor may be controlled using a flow metering valve65. The exemplary method may further comprise the step of stopping the supply of the first fuel11after starting the aircraft engine, if desired.

In the exemplary method of operating the dual fuel aircraft gas turbine engine101, the step of vaporizing the second fuel12may be performed using heat from a hot gas extracted from a heat source in the engine101. As described previously, in one embodiment of the method, the hot gas may be compressed air from a compressor155in the engine (for example, as shown inFIG. 5). In another embodiment of the method, the hot gas is supplied from an exhaust nozzle98or exhaust stream99of the engine (for example, as shown inFIG. 6a).

The exemplary method of operating a dual fuel aircraft engine101, may, optionally, comprise the steps of using a selected proportion of the first fuel11and a second fuel12during selected portions of a flight profile120, such as shown, for example, inFIG. 7, to generate hot gases that drive a gas turbine engine101. The second fuel12may be a cryogenic liquid fuel112, such as, for example, Liquefied Natural Gas (LNG). In the method above, the step of varying the proportion of the first fuel12and the second fuel13during different portions of the flight profile120(seeFIG. 7) may be used to advantage to operate the aircraft system in an economic and efficient manner. This is possible, for example, in situations where the cost of the second fuel12is lower than the cost of the first fuel11. This may be the case, for example, while using LNG as the second fuel12and kerosene based liquid fuels such as Jet-A fuel, as first fuel11. In the exemplary method of operating a dual fuel aircraft engine101, the proportion (ratio) of amount of the second fuel12used to the amount of the first fuel used may be varied between about 0% and 100%, depending on the portion of the flight mission. For example, in one exemplary method, the proportion of a cheaper second fuel used (such as LNG) to the kerosene based fuel used is about 100% during a cruise part of the flight profile, in order to minimize the cost of fuel. In another exemplary operating method, the proportion of the second fuel is about 50% during a take-off part of the flight profile that requires a much higher thrust level.

The exemplary method of operating a dual fuel aircraft engine101described above may further comprise the step of controlling the amounts of the first fuel11and the second fuel12introduced into the combustor90using a control system130. An exemplary control system130is shown schematically inFIG. 4. The control system130sends a control signal131(S1) to a control valve135to control the amount of the first fuel11that is introduced to the combustor90. The control system130also sends another control signal132(S2) to a control valve65to control the amount of the second fuel12that is introduced to the combustor90. The proportion of the first fuel11and second fuel12used can be varied between 0% to 100% by a controller134that is programmed to vary the proportion as required during different flight segments of the flight profile120. The control system130may also receive a feed back signal133, based for example on the fan speed or the compressor speed or other suitable engine operating parameters. In one exemplary method, the control system may be a part of the engine control system, such as, for example, a Full Authority Digital Electronic Control (FADEC)357. In another exemplary method, a mechanical or hydromechanical engine control system may form part or all of the control system.

The control system130,357architecture and strategy is suitably designed to accomplish economic operation of the aircraft system5. Control system feedback to the boost pump52and high pressure pump(s)58can be accomplished via the Engine FADEC357or by distributed computing with a separate control system that may, optionally, communicate with the Engine FADEC and with the aircraft system5control system through various available data busses.

The control system, such as for example, shown inFIG. 4, item130, may vary pump52,58speed and output to maintain a specified pressure across the wing7for safety purposes (for example at about 30-40 psi) and a different pressure downstream of the high pressure pump58(for example at about 100 to 1500 psi) to maintain a system pressure above the critical point of LNG and avoid two phase flow, and, to reduce the volume and weight of the LNG fuel delivery system by operation at high pressures and fuel densities.

In an exemplary control system130,357, the control system software may include any or all of the following logic: (A) A control system strategy that maximizes the use of the cryogenic fuel such as, for example, LNG, on takeoff and/or other points in the envelope at high compressor discharge temperatures (T3) and/or turbine inlet temperatures (T41); (B) A control system strategy that maximizes the use of cryogenic fuel such as, for example, LNG, on a mission to minimize fuel costs; (C) A control system130,357that re-lights on the first fuel, such as, for example, Jet-A, only for altitude relights; (D) A control system130,357that performs ground starts on conventional Jet-A only as a default setting; (E) A control system130,357that defaults to Jet-A only during any non typical maneuver; (F) A control system130,357that allows for manual (pilot commanded) selection of conventional fuel (like Jet-A) or cryogenic fuel such as, for example, LNG, in any proportion; (G) A control system130,357that utilizes 100% conventional fuel (like Jet-A) for all fast accels and decels.

FIG. 8illustrates an exemplary embodiment of a liquid fuel vaporizer500located in the exhaust gas flow, illustrated as arrow502, of a jet engine. Specifically, the fuel is vaporized by means of a panel or panels504mounted via panel mounts506. The panel(s)504may be attached to the center body508and may be located in the aerodynamic wake of the turbine rear frame struts510.

FIGS. 9-12Billustrate that the panel504is constructed such that liquid fuel is supplied through a liquid supply line520having a liquid header521and vaporized fuel is returned through a gas return line522having a gas header523. The panel504may be located in the aerodynamic wake of the turbine rear frame strut510so it is a lower drag structure than one that might reside outside the wake of the turbine rear frame struts510. The exterior shape of the panel504is constructed such that the fore end of the panel504located near the turbine rear frame strut510is of a similar thickness to that of the width of the turbine rear frame strut510. The aft end of the panel504could be bluntly terminated or terminated with a more aerodynamic shape. The edges of the panel504close to the centerbody508or the exhaust nozzle may be bluntly terminated at the centerbody508or exhaust nozzle surface. If the panel height does not span the full distance between the centerbody508and the exhaust nozzle, then the panel504may also be terminated with more aerodynamic shapes. The length of each panel504could vary to whatever length necessary to fully vaporize the fuel and heat it to the desired temperature.

The exterior512of the panel504may be an impermeable shell512. The interior of the panel504may be a semi-hollow cavity with the capacity to contain fuel under pressure and direct it from the liquid inlet to the gaseous exhaust. It could be filled with and/or bonded to foamed metal514, baffles516, or some other structure or material(s) that react to the fuel pressure forces from one surface to the other maintaining the panel shape and integrity (as shown more clearly inFIG. 12Awhere fuel pressure is shown by arrows518and metal foam, baffles, or other internal structure reaction forces are shown by arrows519). The internal structures could also serve as thermal fins to enhance heat transfer between the fuel and the vaporizer wall surface. Foamed metal, baffles516, or other structures could be designed with dimensions and properties to stabilize the vaporization of the fuel, reducing fuel pressure pulsations inherent to vaporization systems. The baffles516or other structures connecting exterior faces of the panel504may direct fuel flow.

Attachment methods could utilize the turbine rear frame struts, centerbody, and exhaust nozzle surfaces.

The exterior of the panel504could be smooth and flat or could be modified to enhance heat transfer to the exhaust gas by adding fins, texture, devices to trip the boundary layer flow, or by twisting or curving the panel surface in such a way to redirect the exhaust gas flow.

The vaporizer could be limited to one panel504behind one turbine rear frame strut510or it could include multiple panels504behind multiple turbine rear frame struts510. These panels504could be independent, connected in parallel, in series, or with bypass valves to allow flow through some panels while others have no fuel flowing through them, thus regulating the vaporized gas temperature.

Placing the vaporizer panels504in the aerodynamic wake of the turbine rear frame struts510solves the problem of high drag aerodynamic losses when the vaporizer is installed in the exhaust system. The inclusion of internally bonded baffles516or foamed metal514allows the vaporizer to be thin and light weight while being able to handle the internal fuel. The foamed metal internal structure ensures stable vaporization, a major problem in vaporization systems.

The low aero loss vaporizer solution described by this invention offers a significant improvement in SFC over other vaporizer solutions that obstruct the exhaust flow due to the associated drag penalty. The technology of the foamed metal core bonded to the exterior thin metal envelope allows inherently stable vaporization in a light weight, self-supporting structure.

The present disclosure contemplates that, in some circumstances, it may be disadvantageous to supply certain liquid fuels (e.g., liquid natural gas, liquid hydrogen) to the combustor nozzles (also referred to as “fuel nozzles”) in a liquid form using current nozzle designs. Vaporizing such fuels prior to injection into the combustor may allow the fuels to ignite and burn more effectively.

The present disclosure contemplates that some vaporization systems may use a heavy intermediate fluid system to extract heat from other areas of the engine. Some example embodiments according to at least some aspects of the present disclosure may not require the use of an intermediate fluid system.

Some example embodiments according to at least some aspects of the present disclosure may relate to methods and apparatus for vaporizing a liquid fuel using the combustor and/or associated components of a jet engine as the heat source. In some example embodiments, heat from the combustion area (e.g., combustor90and/or combustor case (both shown inFIGS. 3 and 4)) may be absorbed by a vaporizer located in the general vicinity of the combustor. An example vaporizer may include a separate component attached (e.g., internally and/or externally) to the combustor case, which may be generally annular, and/or may integral to the inner and/or outer walls of the combustor case. An example vaporizer may include multiple passages (e.g., generally parallel passages), which may be connected by headers, one or more passages connected in series, and/or a combination of these and other configurations.

An example vaporizer may be a separate component mounted externally to the combustor case (e.g., to an exterior surface of the combustor case wall), a separate component mounted inside the combustor case (e.g., to an interior surface of the combustor case wall), and/or vaporizer passages may be manufactured integrally with the combustor case wall.

Generally, as liquid fuel is supplied to the inlet of the vaporizer, heat absorbed into the vaporizer from the combustion process may heat and/or boil the liquid fuel until it emerges from the vaporizer exit (e.g., as a gas). In some example embodiments, the gaseous fuel may be supplied to the combustor fuel nozzles.

The present disclosure contemplates that, typically, the combustor may be one of the hottest parts of the engine. A vaporizer disposed at or near the combustor may require less surface area to absorb a given amount of energy than at other locations in or on the engine. Accordingly, a vaporizer configured for mounting at or near the combustor may be able to be sized smaller and/or lighter than a vaporizer configured for mounting at a lower-temperature location.

FIG. 13is a cross sectional view of an example combustor case vaporizer mounted internally, according to at least some aspects of the present disclosure. More specifically, a combustor600having an outer case602formed by a combustor case wall604having an exterior surface606and an interior surface606. Air supply entering the combustor600is illustrated with arrow608and a fuel nozzle612having a fuel supply for the combustor is designated at610. The vaporizer620having vaporizer passages622is illustrated as being mounted internally on the combustor case wall604forming the outer case602. Liquid fuel may enter the vaporizer at630and exit the vaporizer at614. Generally, the vaporizer620may be referred to as being mounted internally if the vaporizer620is disposed within a volume at least partially defined by the combustor case wall604. In some example embodiments, the vaporizer620may be mounted substantially in direct contact with the interior surface606of the combustor case wall604. In some example embodiments, the vaporizer620may be mounted to, but may not lie directly against, the combustor case wall604. For example, some embodiments may include a gap allowing at least some airflow between the interior surface606of the combustor case wall604and the vaporizer620. In other words, the vaporizer620may be spaced apart from the interior surface606of the combustor case wall604. In some example embodiments, the vaporizer620may be mounted to and/or supported by the interior surface606of the combustor case wall604, but a spacer element may at least partially interpose the vaporizer620and the interior surface606of the combustor case wall604.

FIG. 14is a cross section view of an example combustor case vaporizer620mounted externally, according to at least some aspects of the present disclosure. Generally, the vaporizer620may be disposed in heat transfer contact with the combustor case wall604. In some example embodiments, at least a portion of the vaporizer620may be mounted substantially against the exterior surface605of the combustor case wall604to receive heat from the combustor case wall604by conduction.

FIG. 15is a cross section view of an example integral combustor case vaporizer620, according to at least some aspects of the present disclosure. In some example embodiments, at least some portions of the vaporizer620may be monolithically (e.g., as a single element) formed with the combustor case wall604. In some example embodiments, at least some portions of the vaporizer620may be embedded at least partially within the combustor case wall604. For example, tubing forming the vaporizer passages622may be at least partially embedded within the thickness of the combustor case wall604. In some example embodiments, at least a portion of the vaporizer620may form at least a portion of the interior surface606of the combustor case wall604and/or the exterior surface605of the combustor case wall604. Alternatively, the turbine section or nozzle section may include an annular case to which a portion of the vaporizer may be attached or imbedded such that fuel in the passages may be heated by the turbine section or nozzle section

AlthoughFIGS. 13-15illustrate a liquid fuel inlet at a generally forward location on the combustor case wall and the vaporized fuel exit at a generally aft location on the combustor case wall, it is within the scope of the disclosure to direct the liquid in a generally aft inlet and withdraw the vapor from a generally forward exit.

AlthoughFIGS. 13-15illustrate a vaporizer comprising passages having generally circular cross sections and substantially uniform diameters, it is within the scope of the disclosure to use vaporizer passages having cross sections of alternative shape(s) (e.g., generally square, generally rectangular, generally triangular, generally oval, etc.) and/or to use vaporizer passages of differing diameters (or effective diameters) in a serial and/or parallel flow arrangement.

The following is a non-exhaustive list of potential points of novelty: A fuel vaporizer disposed in heat transfer communication with a combustor of a gas turbine engine. A fuel vaporizer disposed in heat transfer communication with a combustor case wall of the gas turbine engine. A fuel vaporizer arranged to vaporize a liquid fuel flowing there through by transferring heat from a combustor to the fuel. A fuel vaporizer mounted within a combustor case on an interior surface of the combustor case wall. A fuel vaporizer mounted on an exterior surface of a combustor case wall. A fuel vaporizer integrally formed with a combustor case wall.

FIG. 16illustrates an embodiment similar to that disclosed inFIG. 4, except that compressor bleed flow may be utilized in the heat exchanger or vaporizer60for vaporization of liquid natural gas. The vaporizer60may be provided for heating a fluid from cold temperatures to any required system and/or combustion temperatures. Embodiments include those that utilize heat transfer associated with compressor discharge gases of the aircraft engine. As desired, embodiments provide that the fluid may undergo a phase change from liquid to gas in the heating as well as embodiments for which the fluid remains in a single phase. The phase may be selected from the group including liquid or gas. As such, embodiments and alternatives are provided that allow single or dual fuel combustion for airplane engines.

The illustrated example is not meant to be limiting of a vaporizer/heat exchanger that is capable of transferring heat to a fluid (with or without phase change, liquid to gas) that is routed through the vaporizer60and that uses compressor discharge gases of the aircraft engine. The vaporizer60may include embodiments that provide coiled, all axial, and/or a combination (coiled and axial) of tubes selected as desired to accomplish a desired heat transfer requirement to the fluid. Alternatives provide that the vaporizer/heat exchanger may be integral to the compressor case itself. Embodiments of the vaporizer60may be manufactured from materials to include metal, composite, or a combination thereof.

Utilization of the embodiments and alternatives herein provide for a variety of single and dual fuel engines for which the with the exhaust gasses are a heat source to bring the fuel temperature to system and/or combustion requirements while also achieving a minimal increase to specific fuel consumption.

With reference toFIGS. 17A-17C, embodiments provide a multistage liquid natural gas vaporizer. With respect toFIG. 17A, the multistage vaporizer system700includes a first heat exchanger702in parallel with a second heat exchanger704. The first heat exchanger702may utilize a hot air engine sink to heat the LNG. By way of example, the hot air may be selected from a group comprising compressor air, core exhaust air, and/or turbine bleed air. The second heat exchanger704, as illustrated, may utilize alternative fluid flow(s) to heat the LNG. It is contemplated that the second heat exchanger704may utilize a lower temperature air sink. A valve706may be utilized to direct the flow of the LNG to the first heat exchanger702and/or the second heat exchanger704. If the two heat exchangers are heating the LNG to different temperatures the valve706may be used to adjust the temperature of the LNG leaving the multistage vaporizer system700. A gas metering valve708may be utilized to modulate the flow of vaporized gas from the multistage vaporizer system700.

With respect toFIG. 17B, an alternative multistage vaporizer system710is illustrated, which includes a heat exchanger714, utilizing the core exhaust cross-flow as a heat source, in series with a heat exchanger712, utilizing hot air selected from a group comprising compressor air, core exhaust air, and/or turbine bleed air. A valve716may be utilized to modulate the flow into the multistage vaporizer system710and a gas metering valve718may be utilized to modulate the flow of vaporized gas from the multistage vaporizer system710.

During operation, the heat exchanger714changes the phase of the natural gas while the second heat exchanger712utilizes a lower temperature air sink to operate as a recuperater to lower or raise the temperature, as desired, of the vaporized liquid gas. For example, the second heat exchanger712may lower the temperature at low LNG flow rates or, alternatively, raise the temperature at high LNG flow rates.

Conversely, inFIG. 17C, an alternative multistage vaporizer system720is illustrated, which includes a heat exchanger722, utilizing hot air selected from a group comprising compressor air, core exhaust air, and/or turbine bleed air in series with a heat exchanger724utilizing the core exhaust cross-flow as a heat source. A valve726may be utilized to modulate the flow into the multistage vaporizer system720and a gas metering valve728may be utilized to modulate the flow of vaporized gas from the multistage vaporizer system720. With the hot air engine sink as the second heat exchanger in the series, the overall weight of the heat exchanger system may be reduced by allowing the use of materials that do not need to be capable of withstanding as high of a temperature as before these embodiments came into being.

FIG. 18illustrates an embodiment similar to that disclosed inFIG. 16, except that two vaporizers in series are utilized. As illustrated the first vaporizer60A utilizes fan bleed or compressor discharge pressure bleed to heat the natural gas and the second vaporizer60B utilizes a portion97of the gas turbine engine101exhaust gas99to heat the cryogenic liquid fuel112.

The above described embodiments allow for the utilization of different heat sinks for vaporization of liquid natural gas in an aircraft engine. The multistage vaporizer systems not only vaporize the liquid natural gas but also control its temperature. In the art, temperature control of the vaporized LNG is a challenge not overcome until the creation of the present embodiments. Before now, if a heat exchanger was sized to provide the vaporization required at high engine demand LNG flow rates then the fuel would have likely been over temped at lower engine demand LNG flow rates. Prior art designs further require that active control using a bypass system and valving is provided. Such prior art actively controlled designs are complex and add weight to the system. The prior art challenge of existing designs is overcome by the present embodiments in that no control system is required to ensure that the temperature fuel is within spec. Instead, embodiments for this passively controlled and therefore simpler system dispense with a need for extra valving. There are also some potential weight advantages for putting the lower temperature heat sink first in that it could enable the use of lower density materials such as aluminum.

Embodiments are provided for heating a fluid such as fuel in the form of liquid natural gas from cold temperatures to the required system and/or combustion temperatures with the exhaust gasses of airplane engines including turbo-fan, turbo-jet, turbo-prop, open-rotor, etc. As desired, embodiments provide that the fluid may undergo a phase change from liquid to gas in the heating as well as embodiments for which the fluid remains in a single phase. The phase may be selected from the group including liquid or gas. As such, embodiments and alternatives are provided that allow single or dual fuel combustion for airplane engines.

Exemplary embodiments include vaporizer/heat exchangers that are able to transfer heat to a fluid, with or without phase change from liquid to gas, in the exhaust gases of airplane engines. With reference toFIG. 19, the vaporizer800surrounds the exhaust center body802and includes tubes810. The tubes810may be selected from a group of tube descriptions as all coiled, all axial, and/or a combination of coiled and axial tubes, all selections made in order to accomplish a desired amount of heat transfer to the fluid as required with a minimal detrimental impact on specific fuel consumption. A hat band812may be attached to the shroud to carry the tube loads. Further, any number of stiffening rings may be utilized to provide rigidity to the tubes810.

With reference toFIG. 20, a vaporizer/heat exchanger is illustrated as including a panel design with and without internal axial, cross and switchback, or serpentine channels. To mitigate against vapor trapping along internal channels and associated heat transfer degradation or two-phase flow instability, the internal channel orientations are selectively tilted relative to gravity by local rotation of the channel flow direction about at least one Cartesian coordinate axis. The panels820may be attached to the exhaust center boy822and the exhaust nozzle shroud824. The panels820may be constructed internally with finned, metal/composite foam, and or engineered metal/composite foam (DMLS). Embodiments for the vaporizer/heat exchanger are manufactured from materials selected for desired properties at to temperature variations, weight, cost, etc. Alternatives are produced in metal, composite, or a combination of the two.

It will be understood that while the dual fuel system has been described and illustrated as including a first fuel system independent from the second fuel system that the dual fuel system may be structured in any suitable manner. For example, portions of the first and second fuel systems may be combined in any suitable manner, which may reduce the weight. By way of non-limiting example, such a system may include that the fuels may be mixed in one supply system. For example, the fuels may be mixed as a liquid, vaporized, and the resulting mixture may be supplied out of a single port fuel nozzle.

To the extent not already described, the different features and structures of the various embodiments may be used in combination with each other as desired. That one feature may not be illustrated in all of the embodiments is not meant to be construed that it may not be, but is done for brevity of description. Thus, the various features of the different embodiments may be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure.