Patent Description:
Conventional aircraft must inert a vapor space above jet fuel in fuel tanks to mitigate the risk of fires and explosions. These tanks are typically vented to the outside air to control their pressure. However, the transition to using alternative fuels, such as hydrogen or ammonia, means the risk mitigation strategy must change. For example, tanks containing hydrogen or ammonia may not be able to be vented directly to the atmosphere. Additionally, hydrogen or hydrogen-enriched fuels have wider flammability limits and lower minimum ignition energies than conventional jet fuel, requiring accommodation to prevent ignition and flame propagation. Similarly, although ammonia has a narrower flammability limit, it is a toxic fuel and thus cannot be released to ambient when an aircraft is on the ground. Accordingly, improved fuel systems and storage systems thereof may be useful for further aircraft configurations. <CIT> relates to a double-walled container. <CIT> relates to a fuel storage system. <CIT> relates to a fuel system comprising a cryogenic fuel tank located within a compartment provided with a nitrogen-rich stream of gas from an on board inert gas generating system. The compartment includes an overboard vent.

According to some embodiments, aircraft fuel systems are described. The aircraft fuel systems include a fuel vessel containing a non-mixture fuel, a protective vessel arranged about the fuel vessel such that the fuel vessel is contained within the protective vessel and a protective space is defined between an outer surface of a vessel wall of the fuel vessel and an inner surface of a vessel wall of the protective vessel, at least one mounting structure fixedly positioning the fuel vessel within the protective vessel, a fuel consumption device configured to consume the non-mixture fuel, a fuel output fluidly connecting an interior of the fuel vessel to the fuel consumption device, the fuel output fluidly isolated from the protective space, and a relief output fluidly connecting the protective space to a relief flow path, the relief output and relief flow path configured to vent gas from the protective space and remove any non-mixture fuel from the protective space. The aircraft fuel systems include that the protective space is filled with an inert gas and include an auxiliary system configured to receive the inert gas from the protective space.

In addition to the features described above, further embodiments of the aircraft fuel systems may include that the non-mixture fuel is cryogenic liquid hydrogen.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include that the non-mixture fuel is pressurized gaseous hydrogen.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include that the non-mixture fuel is pressurized ammonia or liquid ammonia at ambient pressure.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include an inerting agent supply device configured to generate inert gas and supply said inert gas into the protective space.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include that the at least one mounting structure comprises a heat exchanger configured to transfer heat from the fuel vessel to the inert gas.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include that the auxiliary system is a fire suppression system.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include that the fuel consumption system is one of a fuel cell or a non-mixture fuel burning engine.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include a second fuel vessel containing a second, different non-mixture fuel.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include that the relief flow path includes a relief valve, a flame arrestor, and an external vent.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft fuel systems may include that the relief flow path includes a combustor configured to combust vented non-mixture fuel from the protective space prior to venting of the combusted non-mixture fuel.

According to some embodiments, aircraft are described. The aircraft include a fuselage, wings, and an aircraft fuel system. The aircraft fuel system includes a fuel vessel containing a non-mixture fuel, a protective vessel arranged about the fuel vessel such that the fuel vessel is contained within the protective vessel and a protective space is defined between an outer surface of a vessel wall of the fuel vessel and an inner surface of a vessel wall of the protective vessel, at least one mounting structure fixedly positioning the fuel vessel within the protective vessel, a fuel consumption device configured to consume the non-mixture fuel, a fuel output fluidly connecting an interior of the fuel vessel to the fuel consumption device, the fuel output fluidly isolated from the protective space, and a relief output fluidly connecting the protective space to a relief flow path, the relief output and relief flow path configured to vent gas from the protective space and remove any non-mixture fuel from the protective space. The aircraft fuel system includes that the protective space is filled with an inert gas and includes an auxiliary system configured to receive the inert gas from the protective space. The fuel consumption system is installed to at least one of the fuselage and the wings.

In addition to one or more of the features described above, further embodiments of the aircraft may include that the non-mixture fuel is one of hydrogen or ammonia.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft may include that the fuel consumption system is one of a fuel cell system configured to generate power for flight of the aircraft and a non-mixture fuel-fuel burning engine configured to generate power for flight of the aircraft.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft may include a second fuel vessel containing a second, different non-mixture fuel.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft may include that the at least one mounting structure comprises a heat exchanger configured to transfer heat from the fuel vessel to the inert gas.

In addition to one or more of the features described above, or as an alternative, further embodiments of the aircraft may include that the relief flow path includes a relief valve, a flame arrestor, and an external vent.

Referring to <FIG>, a schematic illustration of an aircraft <NUM> that may incorporate embodiments of the present disclosure is shown. The aircraft <NUM> includes a fuselage <NUM>, wings <NUM>, and a tail <NUM>. In this illustrated embodiment, the aircraft <NUM> includes wing-mounted aircraft propulsion systems <NUM>. The wing-mounted aircraft propulsion systems <NUM> may be convention gas turbine engines, fuel-cell powered electrical propulsion systems, or other propulsion systems as known in the art. In other configurations, aircraft employing embodiments of the present disclosure may include fuselage-mounted and/or tail-mounted configurations. Further, any number of fuel-cell powered propulsion and/or power generation systems, such as an auxiliary or emergency power system, may be employed, from one to four or more, depending on the aircraft configuration and power and thrust needs thereof. The propulsion systems <NUM> may be used to generate thrust for flight and may also be used to generate onboard electrical power, particularly in a fuel cell configuration. The aircraft <NUM> may also include auxiliary power units <NUM> that may be fuel cell based, or otherwise configured to generate power. In other embodiments, the propulsion systems <NUM> may be configured to burn or consume fuel to drive a shaft and fan similar to conventional gas turbine engines and the like.

Fuel cell based power systems (e.g., for power generation and/or for propulsion) and/or combustion engines (e.g., for power generation and/or for propulsion) may employ various types of fuel, including hydrogen and/or ammonia. The fuel cell systems may employ a catalytic reaction to consume the fuel whereas the combustion systems may employ air mixing and ignition for consumption of the fuels. In either configuration, the aircraft must include onboard storage of the fuel to be consumed.

For example, turning now to <FIG>, a schematic diagram of an aircraft propulsion system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft propulsion system <NUM> is a non-combustion system, and includes a fan <NUM>, a drive shaft <NUM>, a motor <NUM>, and an aircraft power generation system <NUM>. The fan <NUM> is operably coupled to and configured to be rotated by the drive shaft <NUM> to generate thrust, similar to a fan and fan section of a conventional gas turbine engine. However, in the fuel cell configuration of <FIG>, there is no core flow path and no turbine section(s) driven by combusted and expanded gas. In contrast, the drive shaft <NUM> that drives rotation of the fan <NUM> is operably coupled to and driven by the motor <NUM>. The motor <NUM> may be an electric motor that converts electrical power to mechanical (rotational) energy. The motor <NUM> receives power from the aircraft power generation system <NUM> along an electrical connection <NUM>. The aircraft propulsion system <NUM> may be configured to operate within similar limits and envelops as a conventional gas turbine engine.

The fan <NUM>, the drive shaft <NUM>, and the motor <NUM> may be arranged along a propulsion system central longitudinal axis A. The fan <NUM>, the drive shaft <NUM>, the motor <NUM>, and the aircraft power generation system <NUM> can be mounted, installed, or otherwise housed within a propulsion system housing <NUM> (e.g., a nacelle for wing-mounted applications) which includes an exit nozzle <NUM> for directing an airflow therethrough for the purpose of driving flight of an aircraft (e.g., generating thrust). The propulsion system housing <NUM> may be configured to be mounted to a wing or fuselage of an aircraft.

The aircraft power generation system <NUM> may be a fuel cell or similar power source (e.g., a solid oxide fuel cell). The aircraft power generation system <NUM> can be configured to not only power the motor <NUM> but also may be used as a power source for other propulsion system components and/or other aircraft electrical systems and components. In one non-limiting example, the aircraft power generation system <NUM> may be configured to output about <NUM> to about <NUM> MW electrical power. In accordance with embodiments of the present disclosure, the aircraft power generation systems may be configured to generate at least <NUM> kW of electrical power (e.g., less power may be used if the system is not used for propulsion). It will be appreciated that when used as a propulsion configuration, the aircraft power generation systems described herein are configured to generate, at least, sufficient power to drive the fan <NUM> and provide sufficient thrust and propulsion for flight at cruise altitudes. The amount of electrical power may be selected for a given aircraft configuration (e.g., size, operating envelope requirements, etc.).

Whether used for propulsion or only onboard electrical power, the aircraft power generation system <NUM> may be configured to combine hydrogen (e.g., liquid, compressed, supercritical, etc.) or other organic fluids as a fuel source using a fuel cell for generation of electricity. In some embodiments, in operation, hydrogen may be heated by fuel cell waste heat (e.g., water output) via a heat exchanger and then expanded through a turbine connected to a generator to extract some electric power from the hydrogen before it is used in the fuel cell. The hydrogen can also be used as the cold sink to cool aircraft environmental control system working fluids and/or provide other onboard thermal management, prior to being supplied to the fuel cell. In some embodiments, the fuel cell of the aircraft power generation system <NUM> can be configured to provide base electric power (e.g., suited for cruise operation). In some non-limiting configurations, some fuel (hydrogen) may be directed to bypass the fuel cell and be used in a small gas turbine to generate additional power for take-off and climb peak power needs.

Turning now to <FIG>, a schematic diagram of an aircraft power generation system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft power generation system <NUM> includes a fuel cell <NUM> and a fuel source <NUM> (such as a hydrogen fuel source). The fuel cell <NUM> is configured to generate electricity, as will be appreciated by those of skill in the art (e.g., a solid oxide fuel cell). In this illustrative configuration, the fuel cell <NUM> includes an anode <NUM>, a cathode <NUM>, and an electrolyte membrane <NUM> arranged therebetween. The fuel cell <NUM> is supplied hydrogen (H<NUM>) from the fuel source <NUM>. The fuel source <NUM> may be a container or tank that houses liquid, compressed, supercritical fluid (e.g., the hydrogen in this example). The fuel cell <NUM> is supplied with oxygen (O<NUM>) from an oxygen source at an inlet <NUM>. In some embodiments, the O<NUM> may be supplied from ambient air, such as using an intake or scoop on a housing assembly, as will be appreciated by those of skill in the art. The O<NUM> and the H<NUM> are combined within the fuel cell <NUM> across the electrolyte membrane <NUM>, which frees electrons for electrical power output <NUM>. The combined O<NUM> and H<NUM> results in the formation of water (H<NUM>O), which may be passed through an outlet <NUM> and dumped overboard, supplied into an onboard water tank, or otherwise used onboard an aircraft, as will be appreciated by those of skill in the art. For example, in one non-limiting embodiment, the water may be injected into a supplementary take-off gas turbine compressor for mass flow augmentation.

The electrical power output <NUM> may be electrically connected to a motor that is configured to drive a drive shaft and a fan of a propulsion system to generate thrust (e.g., as shown in <FIG>). The electrical power output <NUM> may also or alternatively be electrically connected to other electrical systems of a propulsion system and/or aircraft system(s), as will be appreciated by those of skill in the art to provide electrical power thereto.

In conventional aircraft systems, a vapor space above jet fuel in the fuel tanks will be filled with an inert gas to mitigate the risk of flame propagation or explosions. These tanks are typically vented to the outside air to control their pressure. However, using cryogenic hydrogen (H<NUM>), or an alternative fuel such as ammonia (NH<NUM>) (e.g., pressurized ammonia or liquid ammonia at ambient pressure) or other non-mixture fuels, instead of jet fuel means the risk mitigation strategy must change. For example, tanks containing cryogenic fuels cannot be vented to the atmosphere. Further, H<NUM>-enriched fuels have wider flammability limits (FL) as compared to conventional fuels. In contrast NH<NUM> fuel, having narrower FL than H<NUM>, is also toxic and poses a risk of human exposure when released on the ground. Accordingly, prevention of such leaks from cryogenic fuel storage onboard aircraft is advantageous. The cryogenic fuels may be consumed in fuel cell systems, as described above, or may be combusted in a combustion chamber to drive a shaft, as done in conventional jet fuel gas turbine engines.

Non-mixture fuels, as employed by embodiments of the present disclosure, are fuels that are, at least, stored as a pure compound, such as just H<NUM> or just NH<NUM>, within a tank. The use or consumption of the fuel may be as the pure state (e.g., catalyzing or burning H<NUM>) or may be mixed downstream from the tank in order to be consumed. For example, in an ammonia (NH<NUM>) system, the ammonia may be stored in a tank onboard an aircraft, as described herein, and then converted or cracked into a mixture of ammonia, hydrogen, and nitrogen. In other embodiments, different fuels from different tanks may be sourced to then be mixed upstream of the fuel consumption device. As such, the term "non-mixture fuel" as used herein refers to the stored, in-tank state of the fuel, and not necessarily to the state of the fuel at the time of consumption within a consumption device.

Cryogenic fuels, such as H<NUM> and NH<NUM>, can be stored as liquids but used as fuels in a gaseous state (i.e., transitions from liquid to gas prior to consumption). Hydrogen, as a fuel, may also be stored as a high-pressure gas. Under ambient conditions, the lower flammability limit (LFL) H<NUM> is approximately <NUM>% (by volume in air). The LFL of ammonia is approximately <NUM>%. The upper flammability limit (UFL) for H<NUM> is <NUM>% and for NH<NUM> is <NUM>%. Ammonia also poses a toxicity risk when released. In some embodiments, ammonia may be stored as a liquid at room or ambient temperatures (non-cryogenic temperatures) but may require a pressure vessel to contain such a liquid. As described herein, the systems are configured to contain a fuel at high pressure (e.g., <NUM> psi (<NUM> kPa) or greater) and/or low temperature (e.g., <NUM> or less, <NUM> or less, <NUM> or less, etc.). For example, liquid hydrogen may be stored at <NUM> within a pressure vessel, pressurized hydrogen may be stored at higher temperature but at high pressures such as <NUM>,<NUM> psi or (<NUM> kPa) greater, and liquid ammonia may be stored at room temperatures as a liquid but contained at pressures of <NUM> psi (<NUM> kPa) or greater. The fuels stored in accordance with embodiments of the present disclosure may be referred to herein as "non-mixture fuels. " That is, the fuels are completely or substantially pure fuels of a single compound (e.g., hydrogen, ammonia, liquified methane, and the like) when stored in a fuel tank or storage tank. In some configurations, these fuels may be consumed directly through combustion or catalytic reactions, but do not comprise a mixture of different compounds in the stored state. As noted above, multiple non-mixture fuels may be stored in separate tanks, and then combined prior to being supplied into a fuel consumption device.

Enclosing a primary non-mixture fuel storage tank within another vessel enables a barrier layer or cavity to provide protection against leaks of the non-mixture fuels. For example, a void between a primary storage tank and a protective vessel can be filled with inert or low oxygen (O<NUM>) content gas or may be a vacuum. In configurations that employ a vacuum, leaks may be detected by monitoring pressure with sensors and identifying leaks through pressure rises from the vacuum level. As a result, any non-mixture fuel that leaks into the space between the two structures can be diluted to levels below the LFL and/or identified and addressed. In some embodiments, the inert gas can be O<NUM>-depleted air or carbon dioxide (CO<NUM>) that can also be used as a fire suppressant. Further, in some embodiments of the present disclosure, the external protective vessel can include a heat exchanger to enable thermal transfer between the non-mixture fuel and the inert gas. In some such embodiments, the heat exchanger(s) can be configured with microchannels to reduce weight. In some embodiments, a relief valve, flame arrestor, and optional catalytic combustor can allow safe venting of the non-mixture fuel outside the aircraft, when necessary.

In some embodiments, the non-mixture fuel lines of the system could be double-walled with the outer protective line/tube being vacuum insulated or incorporating a gas recovery line that can be purged with inert gases such as nitrogen (N<NUM>) or carbon dioxide (CO<NUM>) and connected to a fuel tank vent. Open vent lines can contain undesirable oxygen levels, and thus, these lines may be equipped with back pressure valves to prevent oxygen from entering the non-mixture fuel system and discharged only when certain pressure levels are reached.

A refrigerant could also be used as the inert agent and integrated as part of a vapor compression cycle on board the aircraft. Carbon dioxide (CO<NUM>) or other non-flammable refrigerant compounds could be employed to transfer heat from or to the fuel tank and then be used as a cold source elsewhere in the aircraft, such as in a condenser. Ground leaks of non-mixture fuels, such as ammonia (NH<NUM>), can pose toxicity as well as fire hazards.

Turning now to <FIG>, an aircraft fuel system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft fuel system <NUM> includes a fuel tank system <NUM> fluidly connected to a fuel consumption device <NUM> (e.g., a fuel cell, combustion engine, or the like). The fuel consumption device <NUM> may be configured to generate power and, in some embodiments, may be configured to consume one or more non-mixture fuels for generating power or thrust for flight of the aircraft. The fuel tank system <NUM> contains a non-mixture fuel <NUM>, such as hydrogen, ammonia, or liquified methane stored at high pressure and/or low temperature. The fuel tank system <NUM> may be configured to store the non-mixture fuels at temperatures ranging from <NUM> to room temperature and <NUM> bar pressure or greater, depending upon the specific non-mixture fuel. The fuel consumption device <NUM> may be a fuel cell, a hydrogen burning engine, an ammonia burning engine, an ammonia/hydrogen burning engine, or other type of consumption device.

The fuel tank system <NUM>, in this embodiment, includes a fuel vessel <NUM> having a respective vessel wall <NUM> and a protective vessel <NUM> having a respective vessel wall <NUM>. The fuel vessel <NUM> is arranged within the protective vessel <NUM> and affixed together by one or more mounting structures <NUM>, such as struts or other mounting mechanisms, as will be appreciated by those of skill in the art. A protective space <NUM> is defined between the vessel wall <NUM> of the fuel vessel <NUM> and the vessel wall <NUM> of the protective vessel <NUM>. Low thermal conductivity composites, such as thermoplastics or carbon matrix composites, as well as metals may be used as materials for the fuel vessel <NUM> and/or the protective vessel <NUM>. The protective space <NUM> is filled with an inert or low oxygen content gas. In some embodiments, an inerting agent supply device <NUM> may be fluidly coupled to the protective space <NUM> to provide inerting agent <NUM> therein. The inerting agent supply device <NUM> may be, for example and without limitation, an on-board inert gas generation system (OBIGGS), a CO<NUM> or other inert gas tank, a catalytic inerting system that catalyzes a hydrocarbon fuel to generate an inert gas, and the like. In some embodiments, other or additional flame retardant insulation materials may be used in the protective space <NUM> to provide additional thermal insulation for the internal fuel vessel <NUM>.

In this illustrative embodiment, the fuel tank system <NUM> has at least two flow outputs. A first flow output is a fuel output <NUM> that fluidly connects the fuel vessel <NUM> to the fuel consumption device <NUM>. Because the fuel vessel <NUM> is arranged inside the protective vessel <NUM>, the fuel output <NUM> is open to the interior of the fuel vessel <NUM>, passes through the protective space <NUM>, through the vessel wall <NUM> of the protective vessel <NUM>, and then fluidly connects to the fuel consumption device <NUM>. The second output is a relief output <NUM>. The relief output <NUM> is a vent or the like for venting gases from the protective space <NUM>. As such, unless there is a crack or leak from the fuel vessel <NUM> and in normal operation, the fuel vessel <NUM> is fluidly isolated from the relief output <NUM>. Similarly, during normal operation and without any failures or cracks, the protective space <NUM> between the fuel vessel <NUM> and the protective vessel <NUM> is fluidly isolated from the fuel output <NUM>.

The relief output <NUM> enables a venting of gases from the protective space <NUM>, whether non-mixture fuel is leaked from the fuel vessel into a vacuum space or an inerting agent space. The relief output <NUM> defines a start to a relief flow path <NUM>. A relief valve <NUM> is arranged proximate to the relief output <NUM> and may be a one-way valve to prevent gases from flowing back into the protective space <NUM>. The relief valve <NUM> may be a pressure-based valve that is configured to open and vent the protective space <NUM> if a pressure at the relief valve <NUM> reaches or exceeds a predetermined threshold value. Such valve may be based on a composition of the gases within the protective space <NUM>, such that a leak from the fuel vessel <NUM> causes an increase in pressure and if the percentage of non-mixture fuel within the protective space <NUM> reaches a specific value it will correspond to a pressure that causes the relief valve <NUM> to open and vent the protective space <NUM>.

Depending on the specific type of non-mixture fuel <NUM> within fuel vessel <NUM>, a flame arrestor <NUM> may be arranged downstream of the relief valve <NUM> along the relief flow path <NUM>. The relief flow path <NUM> may then split to direct the gases to an external vent <NUM> or to an optional combustor <NUM>. The combustor <NUM> may be included for systems that employ non-mixture fuels that require combustion prior to venting out the external vent <NUM> (e.g., ammonia (NH<NUM>)), particularly when an aircraft is on the ground. Alternatively, the combustor <NUM> may be replaced by a temporary holding tank to contain the gases until such gas can be removed safely (e.g., once in flight, or by appropriate ground-based removal techniques). If the combustor <NUM> is included, an oxygen or air source <NUM> may be provided for catalyzing the air or oxygen with the non-mixture fuel within the combustor <NUM>. The combustor <NUM> may be a catalytic combustor or catalytic reactor that is configured to catalyze the non-mixture fuel to form safe or safer gases prior to venting. A valve <NUM> may be arranged to control where the vented gas is directed, depending on the flight conditions when a venting is to occur. For example, if an aircraft is in flight, any leaked gas may be sent directly overboard through the external vent <NUM>. However, if the aircraft is on the ground, the valve <NUM> may be operated to direct the leaked gas to the combustor <NUM> (or holding tank) prior to venting through the external vent <NUM>.

In some embodiments, such as when the inerting agent <NUM> is supplied into the protective space <NUM> from the inerting agent supply device <NUM>. The inerting agent <NUM> is directed to an auxiliary system <NUM>. That is, the inerting agent <NUM> is used for multiple purposes onboard the aircraft. In one such configuration, a third output <NUM> may be arranged providing a fluid connection between the protective space <NUM> and the auxiliary system <NUM>. One or more valves <NUM> may be configured to control flow of the inerting agent <NUM> to the auxiliary system <NUM> and/or the external vent <NUM>. In some embodiments, the auxiliary system <NUM> may be a fire suppression system onboard the aircraft. In such configurations, the inerting agent <NUM> sourced from the inerting agent supply device <NUM> may be passed through the protective space <NUM> and then supplied to the fire suppression system to extinguish a fire onboard the aircraft. In another configuration, the inerting agent <NUM> can be used as a working fluid for thermal control in one or more heat exchangers onboard the aircraft. As the inerting agent <NUM> is sourced from the inerting agent supply device <NUM> it will flow along the fuel vessel <NUM> and thus cool down (or provide heat pick up from the fuel vessel <NUM> to maintain the cold temperatures) and then can be used as a cold sink for other systems onboard the aircraft.

In some embodiments, the third output <NUM> and the relief output <NUM> may be the same output such that an inert gas or the like is vented into the auxiliary system <NUM> rather than to the external vent <NUM>. In some such embodiments, the auxiliary system <NUM> may include a holding tank to contain the vented inert gas and/or leaked non-mixture fuel, depending on the nature of the auxiliary system <NUM>.

Turning now to <FIG>, an aircraft fuel system <NUM> in accordance with an embodiment of the present disclosure is shown. The aircraft fuel system <NUM> includes a fuel tank system <NUM> fluidly connected to a fuel consumption device <NUM> (e.g., a fuel cell, combustion engine, or the like). The fuel consumption device <NUM> may be configured to generate power and, in some embodiments, may be configured to consume a non-mixture fuel for generating power or thrust for flight of the aircraft. The fuel tank system <NUM> contains a non-mixture fuel <NUM>, such as hydrogen or ammonia stored at high pressure and/or low temperatures.

The aircraft fuel system <NUM> is substantially similar to that described with respect to <FIG>, and thus similar features and components may not be discussed in detail. The fuel tank system <NUM> has a double-vessel configuration, with a fuel vessel <NUM> having a respective vessel wall <NUM> arranged within a protective vessel <NUM> having a respective vessel wall <NUM>. A protective space <NUM> is defined between the vessel wall <NUM> of the fuel vessel <NUM> and the vessel wall <NUM> of the protective vessel <NUM>. In this embodiment, the protective space <NUM> is filled with an inerting agent <NUM> sourced from an inerting agent supply device <NUM>. A relief output <NUM> is provided in a similar manner as that shown and described above, and thus will not be described in further detail.

The primary difference between the embodiment of <FIG> and <FIG> is the nature of the mounting of the fuel vessel <NUM> within the protective vessel <NUM>. In the embodiment of <FIG>, the fuel vessel <NUM> is affixed to and mounted within the protective vessel <NUM> by a heat exchange structure <NUM>. The heat exchange structure <NUM> may provide for a mechanism for the inerting agent <NUM> to pick up heat from the fuel vessel <NUM> and remove such heat from the fuel vessel <NUM> to aid in keeping the temperatures of the non-mixture fuel <NUM> at appropriate levels. The heat exchange structure <NUM> my serve or function as a mounting structure while also encouraging heat transfer such that the inerting agent <NUM> removes heat from the system. For example, the heat exchange structure <NUM> may be a series of fins or plates that extend from an external surface of the vessel wall <NUM> of the fuel vessel <NUM> to an internal surface of the vessel wall <NUM> of the protective vessel <NUM>. In some such embodiments, a refrigerant could be used as the inerting agent and integrated as part of a vapor compression cycle on board the aircraft. For example, CO<NUM> or other non-flammable compounds could transfer heat with the fuel vessel <NUM> and then be used as a cold sink elsewhere on the aircraft, such as in a condenser, or other auxiliary system, such as described above.

Advantageously, embodiments of the present disclosure provide for means to safely inert new types of fuel tanks for non-mixture fuels onboard aircraft. In accordance with some embodiments, vacuum insulation allows system monitoring with pressure sensors to trigger a response such as inert purging that would provide effective mitigation due to the confined space of the protective space between the two vessels of the fuel systems. Further, advantageously, embodiments described herein can provide for an integrated means of exchanging heat between a stored non-mixture fuel and an inert gas. Advantageously, in such inert gas systems, total system weight may be required by using the inert gas for the fuel tanks as a source for fire suppression or other inert gas consumption system or auxiliary system.

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

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

While the present disclosure has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the present disclosure is not limited to such disclosed embodiments. Rather, the present disclosure can be modified to incorporate any number of variations, alterations, substitutions, combinations, sub-combinations, or equivalent arrangements not heretofore described, but which are commensurate with the scope of the present disclosure. Additionally, while various embodiments of the present disclosure have been described, it is to be understood that aspects of the present disclosure may include only some of the described embodiments.

Claim 1:
An aircraft fuel system (<NUM>) comprising:
a fuel vessel (<NUM>) containing a non-mixture fuel;
a protective vessel (<NUM>) arranged about the fuel vessel (<NUM>) such that the fuel vessel is contained within the protective vessel and a protective space (<NUM>) is defined between an outer surface of a vessel wall (<NUM>) of the fuel vessel and an inner surface of a vessel wall (<NUM>) of the protective vessel;
at least one mounting structure (<NUM>) fixedly positioning the fuel vessel (<NUM>) within the protective vessel (<NUM>);
a fuel consumption device (<NUM>) configured to consume the non-mixture fuel;
a fuel output (<NUM>) fluidly connecting an interior of the fuel vessel (<NUM>) to the fuel consumption device (<NUM>), the fuel output (<NUM>) fluidly isolated from the protective space (<NUM>);
a relief output (<NUM>) fluidly connecting the protective space (<NUM>) to a relief flow path (<NUM>), the relief output and relief flow path configured to vent gas from the protective space (<NUM>) and remove any non-mixture fuel from the protective space,
wherein the protective space (<NUM>) is filled with an inert gas and the aircraft fuel system comprises an auxiliary system (<NUM>) configured to receive the inert gas from the protective space (<NUM>).