58results about "Fuel systems for specific fuels" patented technology

Hybrid propulsion systems that utilize liquid natural gas solid oxide fuel cells in a manner practical for use in aircraft that avoid the use of heavy batteries, provide transient response times suitable for use in aircraft, and / or simplify reactant pre-conditioning systems using a compressor and turbine pair operatively coupled to the solid oxide fuel cell. Such hybrid propulsion systems for an aircraft may include a liquid natural gas turboshaft engine, a turbo generator operatively coupled to the turboshaft engine, a liquid natural gas solid oxide fuel cell, and an electric fan. The electric fan may be configured to generate thrust for the aircraft and may be powered by the turbo generator and / or the solid oxide fuel cell. Fuel cell reactants may be pre-conditioned using turboshaft bleed air, and the liquid natural gas may be pre-heated by fuel exhaust from the solid oxide fuel cell.

A fuselage portion of or for an aircraft. The fuselage portion comprises an aircraft component which is installed in a shell of the fuselage portion by a plurality of connector elements. Therein, a respective first end of each connector element is held in an associated bore b running through the shell, and a respective second end of each connector element is joined to the aircraft component by a joint providing the respective connector element with a margin of swiveling relative to the aircraft component. Also an aircraft comprising such a fuselage portion, and a method for assembling such a fuselage portion.

The invention relates to a drive device (1) for an aircraft (2), having: a fuel cell device (3) which can be connected to a hydrogen tank (4, 4a, 4b, 4c, 4d); an electric drive motor (5) which is electrically connected to the fuel cell device and is designed to drive a propeller (6) about a drive axis (7); a first fire protection chamber (8) in which the fuel cell device and in particular a hydrogen tank are arranged; a first fire protection wall (8a) which extends completely or partially around the first fire protection chamber; a second fire protection chamber (9) in which at least the electric drive motor is arranged; and a second fire protection wall (9a) which extends completely or partially around the second fire protection chamber, wherein the first and second fire protection walls are at least partially different from one another, and wherein the first fire protection chamber is separated from the second fire protection chamber by at least one fire protection wall.

Hydrogen supply device comprising at least two tanks configured to store hydrogen in liquid and supercritical states, respectively. The invention relates to a hydrogen supply device (18) comprising: at least one first tank (20) configured to store hydrogen in a liquid state at a pressure below 13 bar; at least one second tank (22) configured to store hydrogen in a supercritical state at a pressure greater than or equal to 13 bar; at least one upstream portion (26) comprising at least one first hydrogen flow control system (32, 32', 32'') configured to regulate the amount of hydrogen flowing from the first tank (20) to the second tank (22); at least one downstream portion (28) connecting the second tank (22) and a hydrogen-powered system (16). The invention also relates to an aircraft comprising at least one such supply device. Figure 2

The application discloses a kind of hydrogen storage bottle fuselage integration structure manufacturing method of unmanned plane, the application is by constructing the three-dimensional model of target unmanned plane body and carries out simulation analysis, obtains bearing key area and non-bearing area;Selecting hydrogen storage bottle inner container, the hydrogen storage bottle inner container is divided into pressure-resistant winding area and fuselage combination area, and the inner container partition is wound with different thickness of prepreg, to obtain hydrogen storage bottle;For the installation parameters of three-dimensional model and gas bottle of fuselage, design mould;The bearing key area and non-bearing area are respectively wound with different thickness of prepreg, and the whole structure of hydrogen storage bottle and unmanned plane body integration is formed by mould pressing forming;The application can realize weight reduction while ensuring strength, which is conducive to optimizing unmanned plane load and endurance, integrating space to improve hydrogen storage capacity and startability, simplify process and adapt to large-scale production.The application is suitable for hydrogen energy unmanned plane field.

The present invention comprises: a hydrogen fuel tank (2); a plurality of fuel supply pipes (30) connecting the hydrogen fuel tank (2) and a hydrogen gas turbine engine (14); a plurality of pumps (32) that send liquid hydrogen to the hydrogen gas turbine engine (14); and a plurality of return pipes (38) that return a part of the liquid hydrogen flowing through the fuel supply pipes (30) to the hydrogen fuel tank (14). The plurality of fuel supply pipes (30) converge at a convergence portion (36) that is positioned downstream of the plurality of pumps (32), and each of the return pipes (38) has a first end (40) connected to the hydrogen fuel tank (2), and a second end (42) connected to the fuel supply pipe (30) between the pump (32) and the convergence portion (36).

Cooling architecture for a cryogenic and superconducting powertrain using liquid hydrogen on board and method for managing the cooling temperature of the powertrain components. The cooling architecture comprises at least one heat exchanger configured to transfer heat from a first fluid to a second fluid. The first fluid of each heat exchanger may be liquid hydrogen (H2), and the second fluid of each heat exchanger may be gaseous helium (He), liquid Neon or liquid Nitrogen (N2).

The invention relates to a fuel conditioning system (SC) configured to supply an aircraft propulsion system with fuel (Q) from a cryogenic tank (R), the conditioning system (SC) comprising a plurality of compression modules (2A-2C) configured to increase the temperature of the fuel (Q) in an elementary tank (3A-3C) in an isochoric manner, and a metering tank (5) having a fixed calibrated volume (VOLc); wherein each compression module (2A-2C) comprises an inlet valve (V1A-V1C) for supplying the elementary tank (3A-3C) with a volume of fuel (Q) corresponding to the calibrated volume (VOLc) of the metering tank (5), an outlet valve (V2A-V2C) and a degassing valve (V3A-V3C) connected to the elementary tank (3A-3C) and configured to circulate a gaseous fuel flow (G) through a return circuit (CR).

The invention relates to a system comprising a carrier aircraft (1) and a satellite aircraft (2), wherein the carrier aircraft and the satellite aircraft can be detachably connected to each other by means of a coupling device (3, 4, 5), wherein the carrier aircraft has a drive device (6) and is designed to carry the satellite aircraft during take-off from the Earth's surface, and wherein the satellite aircraft has a solar drive device (7a, 7b, 7c, 7d, 7e, 7f, 7g, 7h, 9a, 9b, 9c) and a hydrogen drive device (10, 11, 12, 15, 9a, 9b, 9c, 9d, 9e), each of which can be electrically connected to at least one electric drive motor (8).

The invention relates to a wing structure for an aircraft (1), comprising a wing having at least two half-wings (3), two wings (3), and at least two struts (4), wherein the wings (half-wings) and the struts (4) are connectable to an aircraft fuselage (2) of an aircraft (1), wherein the half-wings (3) are each supported at a respective support point (7) by at least one strut (4), characterized in that a tank (5) is arranged between each half-wing (3) and the strut (4) supporting it, the tank (5) being arranged to connect the strut (4) and the half-wing (3). The invention further relates to an aircraft (1) comprising an aircraft fuselage (2), in which the aircraft fuselage (2) is connected to a wing structure according to one of the preceding claims.

A cryogenic fueled aircraft propulsion system includes a core engine (22) where fuel is mixed with compressed air and ignited to generate a high energy exhaust gas flow, a fuel system (24) that is configured to supply the fuel to the core engine (22), a cryogenic fuel tank (26) for storing fuel cryogenically in a liquid phase, a gas collection system (28) where boil off gas (36) from the fuel system (24) is collected, a fuel cell system (34) where a flow of collected boil off gas (36) is utilized to generate electric power (52), and an electrical coupling (48) where electric power (52) generated by the fuel cell is communicated to location outside of the aircraft (20).

A fuel cell power system comprises at least one fuel cell (200), a ram air system and a cooling circuit in which coolant is intended to circulate for regulating temperature of the at least one fuel cell (200). The cooling circuit comprises a ram air heat exchanger (240) in the ram air system and the ram air system comprises a nozzle. The fuel cell power system further comprises a water tank (500) and the fuel cell power system is arranged to flow water from the water tank (500) to the ram air system so as to spray water in ram air via the nozzle. Thus, dimensioning of the ram air system which includes the ram air heat exchanger (240) is reduced.

The present invention relates to the protection of an aircraft fuselage. To provide a protective measure that is as lightweight as possible and easily integrated into the fuselage, a protective insert (10) for the tail section of an aircraft is provided. The protective insert comprises a protective panel (12), an absorption body (14) supporting the protective panel, and a force transmission device (16) holding the absorption body. The protective panel is designed for a pressure-loaded sliding contact with a runway in the event of a tail impact. The protective panel transmits a pressure force acting during the tail impact to the absorption body. The absorption body is designed to absorb the pressure force by damping it, whereby a portion of the energy due to the acting pressure force can be absorbed and converted into an elastic and / or plastic deformation of at least a part of the absorption body.The power transmission device is designed to hold the protective insert on the underside of a tail section of an aircraft and has a plurality of attachment points (18).

An aircraft includes a first tank and a second tank each storing low-temperature liquid fuel; a pump that receives supply of the low-temperature liquid fuel from the first tank and the second tank; and a controller. Each of the first tank and the second tank is switched between a fuel supply phase of supplying the low-temperature liquid fuel to the pump and a temperature restoring phase of cooling the low-temperature liquid fuel in each of the first tank and the second tank. The controller switches between a first mode where the first tank is in the fuel supply phase and the second tank is in the temperature restoring phase, and a second mode where the first tank is in the temperature restoring phase and the second tank is in the fuel supply phase.

The invention relates to a pressurized bulkhead for an aircraft, comprising a membrane (2) having a planar main surface and at least one hinge (3) for connecting the membrane (2) to the fuselage of the aircraft.

A method of assembling a tube assembly (100) includes the steps of providing a saddle clamp (104) having a first clamp half (109) extending for less than 50 percent of a circumference of a tube (102), and a second clamp half (108) extending for more than 50 percent of the circumference of the tube (102), placing a tube (102) on the first clamp half (109), resting on a static structure (106), and applying a lubricant (122) to a laterally inner surface (117) of the second clamp half (108) and then bringing the second clamp half (108) onto the tube (102), and then driving a fastener (110) to secure the first and second clamp halves (108, 109). A gas turbine engine (20) is also disclosed.

A power unit for power generation in aerospace and other systems. The power unit has a power generation device (101) coupled to a liquid hydrogen reservoir (102) via a first conduit (105) and to an oxygen supply via a second conduit (107). A heat exchanger (106) is provided between the two conduits, and a separator (109) is present in the second conduit for separation of gaseous nitrogen from liquid oxygen. The arrangement provides an efficient, lightweight and compact power generation unit.

Improved cryogenic fluid supply device for an aeronautical turbojet engine. Cryogenic fluid supply device (10) for an aeronautical turbojet engine (1), a tank (20) storing the fluid in liquid form, the device (10) comprising a first heat exchanger (30) capable of first heating the fluid drawn from the tank (20) via a sampling line (11) connecting the tank (20) to the first heat exchanger (30), and a main line (12) connected in a loop to the first heat exchanger (30), the device (10) comprising a bypass channel (14) comprising a first branch (141) branching from the main line (12) at a bypass point (140), to a mixer (50) capable of mixing a liquid phase and a gaseous phase of the fluid, and a second branch (142) connecting the mixer (50) to the sampling line (11), a bypass valve (60) disposed downstream of the bypass point. (140),and movable between a first position allowing a deviation of the fluid flow via the bypass channel (14), and a second position limiting this deviation. Fig. 3.

The application relates to a hydrogen cylinder fixing structure for a unmanned aerial vehicle and belongs to the technical field of unmanned aerial vehicles. The fixing structure comprises a support frame, a limiting mechanism and an elastic abutting mechanism. The support frame is used for being mounted on a hydrogen cylinder. The limiting mechanism and the elastic abutting mechanism are both used for being mounted on a unmanned aerial vehicle. The limiting mechanism and the elastic abutting mechanism are arranged at intervals. The support frame is inserted between the limiting mechanism and the elastic abutting mechanism. One end of the support frame in the length direction is in abutment with the limiting mechanism, and the other end is in abutment with the elastic abutting mechanism. The elastic abutting mechanism is used for abutting the support frame and the limiting mechanism. Through the cooperation of the support frame, the elastic abutting mechanism and the limiting mechanism, the support frame can be quickly mounted between the elastic abutting mechanism and the limiting mechanism. In addition, the support frame can also be quickly dismounted from between the elastic abutting mechanism and the limiting mechanism. Therefore, the convenience of dismounting and mounting the hydrogen cylinder can be improved, and the requirement that the hydrogen cylinder on the unmanned aerial vehicle is convenient to replace can be met.

A non-propulsive turbomachine (2) for an aircraft fuel conditioning system (Q) comprising: a first compressor (21) configured to supply a cabin air conditioning device (6) with a first pressurized airflow (A1p) via a circulation channel (31); a second compressor (22); a combustion chamber (23) configured to be supplied by a supply airflow (A) circulating in a supply channel (32) connecting to the second compressor (22); a single gas turbine (24) connected to the first compressor (21) and the second compressor (22) by a single turbomachine shaft (25); a bypass channel (33) connected to the circulation channel (31) and the supply channel (32); and a relief valve (4) mounted on the bypass channel (33) and configured to move between: a closed position and an open position, in which part of the first pressurized airflow (A1p) flows in the bypass channel (4).Figure from the summary: Figure 3.

A securing structure for a container closure (401) configured to prevent complete detachment of the container closure (401) from a container outlet (402) in the event of deformation of the outlet (402) and / or deformation of the closure (401), the securing structure comprising: an outer ring (200) configured to be disposed radially around the outlet (402); a retaining ring (300) configured to be disposed within the outer ring (200) and axially abut against an abutment surface (201) of the outer ring (200), where the retaining ring comprises: an outer body (301) and; an inner wedge element (302) configured to fit tightly into the outer body (301), and; a retaining plate (100) configured to axially cover the container outlet (402) in an axial direction of the container outlet (402), where the retaining plate (100) is configured to be attached to the outer ring (200) by means of an anchoring element (101).

An engine injector assembly for introducing a gaseous fuel, a liquid fuel and air into a combustion chamber, has a stem and main body oriented along an axis. The main body includes: a central air duct on the axis, having an outlet for directing a central air flow, an outer air duct running radially outside around the central air duct, having an outlet for directing an external air flow, a liquid fuel supply radially between the central and outer air ducts, having a liquid fuel duct and an outlet for introducing the liquid fuel, and a gas fuel supply radially between the liquid fuel supply and the outer air duct, having a gas fuel duct and an outlet for introducing the gaseous fuel. The outer air duct has, at its downstream end, a radially outwardly oriented end portion for directing radially outward the outer air flow.

A gas turbine engine 10 for an aircraft comprising a rich burn quick quench, lean burn (RQL) combustor 16 having a number of spray nozzles 403, 404, in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range of 2-6, with the engine configured to provide fuel comprising sustainable aviation fuel to the spray nozzles. The engine operation is characterised by specific non-volatile particulate emission ratio beign related to the bypass ratio. The fuel spray nozzles may comprise one or more duplex nozzles and one more single flow nozzles and preferably the combustor comprises 10-14 duplex fuel spray nozzles and 4-8 signle flow fuel spray nozzles. The duplex fuel spray nozzles are arranged in groups about the circumference of the combustor and optionally wherein the groups duplex fuel spray nozzles comprise at least two groups arraned dimametrically opposite each other. A corresponding method is also claimed. Fig 6

Multi-phase fluid fuel systems and related methods are disclosed. An example fuel system includes a liquid hydrogen storage tank to include hydrogen fuel in a liquid phase, and an at least one of gaseous or liquid hydrogen storage tank to include the hydrogen fuel in at least one of a gaseous phase or the liquid phase, a first portion of the hydrogen fuel in at least one of the gaseous phase or the liquid phase to exit the at least one of gaseous or liquid hydrogen storage tank and flow to a combustor of an engine, a second portion of the hydrogen fuel in at least one of the gaseous phase or the liquid phase to exit the at least one of gaseous or liquid hydrogen storage tank and flow to the liquid hydrogen storage tank.

The invention relates to an assembly (2) for reinforcing a fuselage (10) of an aircraft (1), said aircraft (1) comprising a fuselage (10) having an opening (12) and a removable panel (14) intended to be received in said opening (12), said assembly (2) comprising: - a first support (22) intended to be fixed to an outer face of the fuselage (10) as well as around and outside an opening (12) of the fuselage; - a sub-assembly (24) intended to be fixed to an outer face of a removable panel (14) of the fuselage, where said sub-assembly (24) is housed inside said first support (22); - first means (32) for fixing said first support to said fuselage; and - second means (34) for fixing said sub-assembly to said removable panel. This makes it possible to reinforce the fuselage of an aircraft when removing the removable panel, for example when maintenance of the hydrogen tanks is necessary.