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WO2025229474A1 - Aéronef à deux carburant avec réservoirs d'hydrogène intégrés - Google Patents

Aéronef à deux carburant avec réservoirs d'hydrogène intégrés

Info

Publication number
WO2025229474A1
WO2025229474A1 PCT/IB2025/054361 IB2025054361W WO2025229474A1 WO 2025229474 A1 WO2025229474 A1 WO 2025229474A1 IB 2025054361 W IB2025054361 W IB 2025054361W WO 2025229474 A1 WO2025229474 A1 WO 2025229474A1
Authority
WO
WIPO (PCT)
Prior art keywords
aircraft
fuel
tank
hydrogen
wing
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
PCT/IB2025/054361
Other languages
English (en)
Inventor
Alexander Frühbeis
Periklis LOLIS
Ingmar GEISS
Thomas AHN
Michael Shamiyeh
Johan Johnson
Florian Troeltsch
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Deutsche Aircraft GmbH
Original Assignee
Deutsche Aircraft GmbH
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from DE102024111926.8A external-priority patent/DE102024111926A1/de
Application filed by Deutsche Aircraft GmbH filed Critical Deutsche Aircraft GmbH
Publication of WO2025229474A1 publication Critical patent/WO2025229474A1/fr
Pending legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D37/00—Arrangements in connection with fuel supply for power plant
    • B64D37/30—Fuel systems for specific fuels
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/10—Aircraft characterised by the type or position of power plants of gas-turbine type 
    • B64D27/12—Aircraft characterised by the type or position of power plants of gas-turbine type  within, or attached to, wings
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D37/00—Arrangements in connection with fuel supply for power plant
    • B64D37/02—Tanks
    • B64D37/04—Arrangement thereof in or on aircraft
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/20—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
    • F02C3/22—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure

Definitions

  • the invention relates to a propulsion system for an aircraft and a transport aircraft with a dual fuel propulsion system.
  • the aircraft according to the invention features a dual fuel system to enable hydrogen fuelled operation with optimised payload efficiency.
  • the aircraft propulsion system architecture which is using a dual fuel engine technology in combination with the required fuel system and fuel tanks system is reducing significantly emissions by payload efficiency and opens the opportunity balance mission range with zero CO2 and low emission operation.
  • the term aircraft refers in particular to a multi engine motor-driven fixed-wing aircraft.
  • the term aircraft also includes, for example, rotorcraft (rotor aircraft, helicopters) and powered gliders. Aircraft and their flight propulsion systems can be differentiated regarding the applicable construction and certification regulations.
  • Regulation CS-23 issued by EASA is applicable to light fixed-wing powered aeroplanes. It applies to: 'normal, utility or aerobatic' category aeroplanes with a maximum of 9 seats (excluding pilot(s)) and a maximum take-off mass of 5,670 kg and to 'commuter' category aeroplanes with a maximum of 19 seats (excluding pilot(s)) and a maximum take-off mass of 8,618 kg.
  • the CS-25 is also a design specification issued by EASA for the type certification of large aircraft, in particular large turbine-powered aircraft. Multi-engine aircraft to be certified according to the CS-25 are considered here.
  • Regional transport aircraft are predominantly characterised by a design with straight, unswept wings and a cruising speed of 500 to 700 km/h.
  • turboprop engines are the main application for regional transport aircraft in civil aviation.
  • One of the first commercially successful turboprop driven aircraft was the Fokker F. 27 with a high-wing configuration and with two turboprop engines attached to the left and right side of the wing and hosted in an under-wing nacelle.
  • the advantage of this design was that it allowed for sufficient clearance and ground clearance for the large propellers to be used, and there would be less damage in the event of an emergency landing. All the fuel was stored in the wing, and there was a clear view from the cabin. Available at https://designer.home.xs4all.nl/aircraft/af27/af27.htm, downloaded on 04.02.2025.
  • D328 Turboprop (D 328 TP. 2020. available: https://www.tikaircraft.com/products/d328-turbo- prop#SAF [downloaded 30.01.2025]).
  • a turboprop engine is a turbine engine that drives an aircraft propeller.
  • a turboprop engine consists of an intake, reduction gearbox, compressor, combustor, turbine, and a propelling nozzle. The air enters the intake and is compressed by the compressor. Kerosene fuel is then added to the compressed air in the combustor, where the fuelair mixture then combusts. The hot combustion gases expand through a set of turbine stages, generating power at the point of exhaust. Some of the power generated by the turbine is used to drive the compressor and electric generator. The gases are then exhausted from the turbine. In contrast to a turbojet or turbofan, the engine's exhaust gases do not provide enough energy to create significant thrust, since almost all of the engine's power is used to drive the propeller via the gearbox.
  • nacelle design typically, in a multi engine aircraft the turboprop engine is attached to the aircraft wings and hosted in a nacelle assembly.
  • the main requirement for nacelle design is that it must be constructed and supported so that it can resist any vibration, inertia and air loads to which it may be subjected in operation.
  • nacelle design has to assure easy cowling removal and installation, access to engine equipment for maintenance and overhaul and assure engine ventilation and fire protection requirements.
  • the turbomachinery is usually optimised for the dominant flight phase, usually the cruise phase, as this also accounts for the highest share of energy consumption over the mission. Equivalent, high operating efficiency is not possible over all flight phases. While for medium and long haul the maximum efficiency can be designed and achieved for the clearly dominant cruise phase, the operating conditions for short and ultra-short haul are much less dominant and more diverse. As a result, engines for regional and short-haul aircraft are operated much less at optimal efficiency over the entire mission and have poorer specific fuel consumption per passenger relative to short- and long-haul aircraft.
  • the different drop in propulsive and thermal efficiency plays a role here, since a propeller aircraft, for example, flies lower and experiences lower, altitude-dependent thrust losses in cruise flight than, comparatively, a long-haul aircraft with a turbofan engine. This results in long-range engines running much more consistently at high power and good efficiency than in climb, while regional aircraft have very significantly wider power ranges, e.g. at take-off, climb and cruise.
  • the dual fuels propulsion architecture as it is known for example from aircraft engine manufacturer General Electric (WO 2012/045028 A1 and US 2013/0192246A1) is based on the combination of classical aircraft fuel, e.g., available Jet A-1 compatible fossil and synthetic fuels or full sustainable aviation fuels (SAF) in combination with hydrogen.
  • classical aircraft fuel e.g., available Jet A-1 compatible fossil and synthetic fuels or full sustainable aviation fuels (SAF) in combination with hydrogen.
  • SAF full sustainable aviation fuels
  • the hydrogen storage tank is foreseen in a tail section of the aircraft fuselage behind a passenger section.
  • the objective of the invention is to provide a propulsion system that allows the combustion of H2with gas turbines, piston engines, etc. and thus provides emission-free flight and a sufficient range.
  • a further objective of the invention is to provide an option that enables a simple retrofit or derivative of existing aircraft, i.e. a simple conversion from a kerosene-powered aircraft to a zero-emission aircraft. Furthermore, the invention enables increased range for H2 powered aircraft in a simple manner to provide the necessary flexibility for the operator or reserve for emergencies.
  • the idea of the invention is to be able to fly aeroplane provided with one turbine as propulsion engine at least with climate-neutral F ⁇ only by replacing the propulsion engine and other components in the engine nacelle.
  • the first tank i.e. the tank for H2
  • the tank for H2 is dimensioned in such a way that the tank has sufficient quantity for typical missions.
  • Typical missions are usually flight distances of at least 100, preferably at least 200, particularly preferably at least 300, especially preferably at least 400 nautical miles (NM).
  • NM nautical miles
  • Typical missions are in the range of 100 nautical miles to 600 nautical miles. It is very important that a propulsion system for an aircraft can continue to be powered even when the climate-neutral fuel has been used up, for example in emergencies, or when alternative airports are assigned.
  • the invention therefore means that typical missions, which also depend on the size of the aircraft, can be carried out entirely with climate-neutral fuel.
  • the tanks are dimensioned accordingly for this purpose.
  • the engine is operated with kerosene or SAF, whereby a change of fuel type from H2to kerosene or SAF is possible at any time.
  • the H2 can be transported in different states of aggregation.
  • the H2 is in gaseous form, preferably at high pressures of 1000 bar to 350 bar, particularly 700 bar to 350 bar, preferably 700 bar or 350 bar.
  • the hydrogen can also be present in cryogenic liquid form or as liquid hydrogen, for example at a pressure of 2 bar or as liquid hydrogen at a pressure of less than 20 bar.
  • cryogenic liquid hydrogen or liquid hydrogen the tank volume can be significantly reduced.
  • filling an empty tank with liquid hydrogen requires pre-cooling.
  • pressurised tanks can also be used, in which the hydrogen is stored in gaseous form. The disadvantage of pressurised tanks is that they are heavier. On the other hand, they require fewer additional devices, and the hydrogen can be stored for longer.
  • the problem of heavy tanks is solved by the propulsion system according to the invention with two liquid fuels, a so-called "dual fuel” propulsion system.
  • a propulsion system according to the invention has two separate tanks systems for each of the two fuels, a first and a second tank system, with the first tank system holding hydrogen H2fuel.
  • the invention provides a first tank to be dimensioned in such a way that the H2tank weighs a maximum of 7% of the maximum aircraft take-off weight. With an exemplary take-off weight of 15,000 kg, the maximum fuel-tank weight would be 1050 kg.
  • the H2 is pressurized up to less than 20 bar.
  • the H2 in liquid form is stored under pressure and the tank is highly thermally insulated by a relatively thick insulating layer.
  • the H2 tank is accommodated in the nacelle of the internal combustion engine, preferably behind the internal combustion engine.
  • the H2tank is designed to be very compact to achieve a very low weight. It is preferable for the tank to be arranged directly in the vicinity of the engine, preferably the gas turbine, namely behind the internal combustion engine, preferably the aircraft turbine in the nacelle away from the aircraft nose.
  • an arrangement of the tanks for H2 in the engine nacelle has the advantage that such a configuration can be implemented on conventional aircraft. So-called retrofits or derivatives are available.
  • 25.04.2025 ising the hydrogen directly after the tank and transporting the gaseous hydrogen allows the use of less complex and heavy compressors instead of liquid hydrogen pumps.
  • the concept according to the invention is further characterised by fact the that the typical ranges for the aircraft type are operated with hydrogen. Typical ranges are significantly shorter than the maximum ranges of the aircraft. Typical ranges of the aircraft are ranges that cover most of the routes flown in everyday operation of the aircraft. Maximum ranges are ranges that are flown very rarely. Infrequent but important routes can also be flown and a second type of aircraft, which would be significantly more expensive to maintain, does not have to be kept available for this purpose.
  • the operator can, as an example, operate the regional aircraft with around 40 - 50 passengers at short ranges of less than 400 or 300 nautical miles (NM), which can be executed almost exclusively with hydrogen fuel. Longer flights and range extensions are also possible.
  • the propulsion system is switched from H2 to the second fuel carried, preferably kerosene or SAF. This concept is also known as the "dual fuel concept".
  • the second fuel is conventionally arranged in fuel tanks within the wing.
  • the first tank is holding the H2 with a weight of up to 4 % of the maximum take-off weight of the aircraft. E.g. with a take-off weight of 15,000 kg for the aircraft, this results in a weight of less than 600 kg H2.
  • the first tank has a length of less than 5 metres.
  • a tank of this size allows it to be installed in the nacelle of a propulsion unit, which saves space. In particular, it is not necessary to accommodate the tank in the fuselage of the aircraft.
  • the first tank for H2 has a diameter of 200 %, in particular preferably at up to 100 % of the turbine diameter and a length of less than 500
  • a further advantage is that a tank as described above can easily be added, which enables retrofit design or derivative. Such an arrangement also has the advantage that only a few modifications to the wing are necessary for a retrofit. Due to the small volume of the H2tank according to the invention, integration is simplified.
  • the drive power of the at least one internal combustion engine is transmitted to a propulsor.
  • the propulsor can be a propeller, a paddle wheel or a rotor.
  • the changeover from the first fuel (H2 takes place) to the second fuel (kerosene) and vice versa successively for the internal combustion engine is conducted in such a way that the drive power delivered to the propulsor remains approximately the same.
  • the invention also provides an aircraft with such a propulsion unit.
  • the propulsion unit as an internal combustion engine can be at least one turboprop engine with a gas turbine, a piston engine or a radial piston engine.
  • the propulsion unit is a turboprop engine.
  • At least the first tank for the first fuel H2 is arranged in the nacelle for the propulsion unit, wherein particularly preferably the first tank is arranged in the nacelle behind the internal combustion engine.
  • the aerodynamic drag of the engine nacelle is not or only marginally increased, as the larger surface area of the nacelle can be compensated by a better aerodynamic shape factor. This means a greater degree of slenderness.
  • the tank can be integrated as a fixed installation or as a modular solution. The modular solution enables a quick change of delivered tanks. "Podded" kits simplifies logistics, especially at small airports where it is more difficult to set up an economic H2 infrastructure.
  • the tanks are located close to the aircraft's centre of gravity. This means that the aircraft's centre of gravity changes only minimally during operation when the tanks are emptied and does not cause a problem for the operation or stability of the aircraft. In particular, a large shift in the centre of gravity when the hydrogen is consumed is avoided. Furthermore, the tanks are located outside the propeller and rotor burst zone. This is very advantageous, as according to certification guidelines no fuels are allowed within the cones of the propeller or rotors.
  • Another objective of the invention is to provide an optimised solution of a commercial transport aircraft with a dual fuel propulsion architecture without scarifying the aircraft’s payload and performance characteristics too much.
  • a propulsion system is to be proposed, which offers an almost non-CC>2 emission operation in typical cruise flight altitudes.
  • a further task of the invention is to provide a propulsion system which enables a deri- vate, a retrofit or conversion of existing aircraft, i.e. a conversion from a kerosene- powered aircraft to an almost zero-emission aircraft with a dual fuel propulsion architecture.
  • An essential component of this invention is an aircraft architecture that enables an almost 'zero emission' cruise flight and thus also reduces the aviation-specific non-CC>2 emissions in high air layers.
  • the invention describes a solution for allocating the fuel types to the respective flight phases and thus for integrating the dimensioning of the dual propulsion system in weight and volume with existing technology in an existing aircraft platform.
  • the invention further discloses a solution for an overall aircraft system comprising the essential elements of a hybrid propulsion system, a hydrogen tank system and a control and regulation unit.
  • the invention provides the concept to design and operate an aircraft powered by a conventional turboprop engine with a hybrid dual fuel concept using a first kerosenebased jet fuel and a liquid hydrogen fuel.
  • the two different fuel types are used selectively for individual sections of a typical flight mission.
  • the aircraft provides separated
  • 25.04.2025 sets of fuel storage tanks One first set of fuel tanks for the kerosene fuel and one second set of fuel tanks for the liquid hydrogen fuel. Both sets of fuel storage tanks are seized and configured that the fuel required for the reserve fuel, which is typically 45 min flight time, is provided by the kerosene fuel tanks.
  • the invention according to claim 1 envisages that the first tank, i.e. the tank for H2, should be dimensioned so that the tank can hold a sufficient quantity of H2 for typical missions.
  • Typical missions usually have flight distances of 100 to 600 nautical miles (NM).
  • the typical missions are at max 500 nautical miles, and thus usually significantly less than the design mission of the aircraft type. It is important that an aircraft propulsion system can continue to be powered even after the hydrogen fuel has been consumed, for example during deviations from the flight plan, e.g. when circling is necessary before landing, or when alternative airports are assigned.
  • the invention ensures that the propulsion engine can be supplied with conventional fuel, such as kerosene or SAF, for such cases.
  • conventional fuel such as kerosene or SAF
  • the invention thus consists in the fact that the typical missions, which also depend on the size of the aircraft, can be carried out entirely with climate-neutral fuel.
  • the tanks are sized accordingly.
  • the engine is fuelled with kerosene, whereby a change of fuel type from H2 to kerosene is possible at short notice.
  • the hydrogen carried can be stored in different aggregate states.
  • the hydrogen is in gaseous state, preferably at high pressures of 350 bar to 1.000 bar, in particular 700 bar to 350 bar, preferably 700 bar or 350 bar.
  • the hydrogen can also be stored in cryogenic liquid form or as liquid hydrogen, for example at a pressure of 2 bar or as liquid hydrogen at a pressure up to 20 bar.
  • cryogenic liquid hydrogen or liquid hydrogen the tank volume can be significantly reduced.
  • refuelling an empty tank with liquid hydrogen requires pre-cooling of the tank and pipes.
  • the hydrogen must be converted into its gaseous state before feeding to the combustors of the engine, which requires additional equipment such as heat exchangers or heating elements.
  • pressurised tanks can also be used to store hydrogen in its gaseous state.
  • pressurised tanks are inexpensive and inexpensive.
  • the invention provides that the first tank is dimensioned in such a way that the H2 tank weights a maximum of 7% of the maximum aircraft take-off weight, depending on the size of the aircraft. With a take-off mass of 15,000 kg, this would be a maximum of 1050 kg of tank mass.
  • the H2 is stored in liquid form under a pressure of less than 20 bar. To prevent the liquefied hydrogen from evaporating when heated, the H2 tank is very heavily thermally insulated with a relatively thick layer of insulation.
  • the H2 tank is accommodated in the nacelle assembly of the combustion engine, preferably behind the combustion engine.
  • nacelle design is designed to resist any vibration, inertia and air loads to which it may be subjected in operation. Also, nacelle design has to assure easy cowling removal and installation, access to engine equipment for maintenance and overhaul and assure engine ventilation and fire protection requirements.
  • the nacelle external shape is a surface resulting from the following conditions:
  • Nacelle compartments are defined starting from fire protection requirements (i.e. engine firewalls position), air inlet system and engine inertial separator design. Nacelle panels position must assure easy access to engine accessories.
  • An arrangement of the hydrogen tanks in the proximity of the engines, namely in the same nacelle assembly has many advantages, as described below.
  • a fuel tank configuration in which the tanks for H2 are located in the engine nacelle has many advantages, as described below.
  • a fuel tank configuration in which the tanks for H2 are located in the engine nacelle has
  • the pipes feeding fuel to the engines are shorter and lighter when they are routed close to the engine in the engine nacelle, and the tanks, especially for H2, compared to the usual integration in the fuselage, wing or externally. If the tanks are located in the fuselage, more effort is required to ensure an equivalent level of safety as with the tanks located in the engine nacelle.
  • the short lines also allow the hydrogen to be transported in gaseous form. With tanks in the fuselage, it is necessary to transport the hydrogen to the engines in liquid form, since otherwise the larger diameter for transporting gaseous hydrogen would result in heavy pipes. Another problem is that, at the present time, there are no liquid hydrogen pumps suitable and certified for aviation. Vaporising the hydrogen directly after the tank and transporting the gaseous hydrogen makes it possible to use less complex and lighter compressors instead of liquid hydrogen pumps.
  • the aircraft comprises an Active Tank Control System (ATCS).
  • ATCS Active Tank Control System
  • the Active Tank Control System (ATCS) is designed to utilize excess heat from the gas turbine engine to regulate the pressure of a cryogenic hydrogen tank system. More specifically, the invention pertains to a pressure management system that enables direct pressure feeding of gaseous hydrogen to the engine, optimizing fuel efficiency and ensuring operational safety.
  • ATCS Active Tank Control System
  • 25.04.2025 • an hybrid fuel operation system, wherein the aircraft is initially started using kerosene or sustainable aviation fuel (SAF) before transitioning to hydrogen.
  • SAF sustainable aviation fuel
  • the ATCS optimizes energy utilization by harnessing waste heat from the gas turbine engine, thereby maintaining the hydrogen storage system within a desired pressure range.
  • the proximity of the gas turbine engine to the hydrogen tank system facilitates efficient heat transfer, reducing reliance on external energy sources.
  • Dual Fuel Redundancy The ability to switch back to kerosene at any flight phase ensures a high level of safety and operational reliability.
  • Operation of the engines is as follows: During initial startup, the gas turbine engine is ignited using kerosene or SAF. Once stable combustion is established and the nominal operating pressure in the hydrogen tank is, the system transitions to gaseous hydrogen supplied through direct pressure feed from the cryogenic tank.
  • the ATCS maintains optimal pressure levels by leveraging waste heat, ensuring continuous and stable fuel flow. If needed, the system can seamlessly revert to kerosene operation at any stage of flight, guaranteeing safety-of-flight compliance.
  • An essential aspect of the invention is the design of such a system on the basis of an intended mission and the overall architecture of the aircraft. Dimensioning and representation of the required efficiency is not possible without consideration of the overall aircraft system.
  • the reserve fuel is hardly used during a usual flight mission profile, it is allocated to the kerosene fuel tanks which are the usual fuel tanks in the wing of a turbo-prop aircraft like a D328-200. For a given 300nm mission this quantity already presents a significant volume of about 50% of the total required fuel quantity of a mission. In consequence the majority, if not the full remaining mission can be flown by using hydrogen fuel alone. In this way the required fuel storage for the mission can be limited to the actual flight time which is reducing the weight and volume of the cryogenic fuel tank significantly. In consequence bulky modifications of the aircraft architecture can be avoided so the important aerodynamic drag/lift characteristic of a given aircraft design stays largely untouched.
  • the dual-fuel propulsion system for aircraft is based on a specific propulsion system architecture using two different energy sources or fuels, which enables emission-free operation of an aircraft during essential phases of a typical mission (zero emission cruise, climb etc.). Due to the specific system architecture and the combination of two energy sources, the impact of the new technology and energy sources on today's aircraft configurations already in service is minimised and enables a significantly earlier deployment possibility compared to the current emission- free system proposals.
  • the proposed aircraft architecture is using a dual fuel engine technology in combination with the required dual fuel system and fuel tanks system to reduce significantly the weight and required space impact of hydrogen tanks and fuel systems for an aircraft.
  • This invention is enabling hydrogen operation with reduced impact on payload and I or passengers vs conventional aircraft with the same block fuel [BF] equivalent block energy [BE] for defined missions.
  • the dual-fuel combustion propulsion system architecture is supporting the following targets for future aircraft and its operation:
  • the dual-fuel combustion propulsion system is using jet-A1 , jet-A1 compatible sustainable aviation fuel or 100% sustainable aviation fuel (SAF) for low altitude mission phases and for required fuel margin and hydrogen for cruise.
  • SAF sustainable aviation fuel
  • the sizing of the fuel tank systems is primarily adapted to climb, cruise flight, descent and landing therefore the limited use of kerosene fuel predominantly for reserve is an optimisation of the use without having to oversize the LH2 fuel tanks and to take advantage of given fuel reservoirs hosted internally in the wings.
  • Direct air capture technologies extract CO2 directly from the atmosphere.
  • the CO2 can be permanently stored in deep geological
  • 25.04.2025 formations (thereby achieving negative emissions or carbon removal) or it can be used, for example in food processing or combined with hydrogen to produce synthetic fuels, e.g. SAF.
  • hydrogen fuel can be stored as a pressurised gas or in liquid form. While gaseous storage may be suitable for shorter flights and is commercially available, the invention focuses on liquid hydrogen storage tanks (LH2) as they require about half the volume and are consequently much lighter than tanks for gaseous hydrogen.
  • LH2 liquid hydrogen storage tanks
  • tanks with a gravimetric index of 20 per cent or higher should be aimed for.
  • the gravimetric index of a tank is calculated by dividing the mass of the stored hydrogen by the sum of the mass of the stored hydrogen and the empty tank weight.
  • a gravimetric index of 50 percent means that the empty tank is as heavy as the stored hydrogen.
  • the turboprop engine can be optimized for delivering propulsive power, where the LH2 is supplied to the combustor injectors at the appropriate temperature via a heat exchanger in the engine pylon.
  • the propulsion system enables the use of new technologies for aviation but on the basis of today's assessable outlook within the next decade:
  • the invention relates to engines for small and large transport aircraft (CS- 23 and CS-25) with twin-engine drives (piston engines or turbo engines), which convert thermal energy into mechanical drive shaft power and drive a propulsor (propeller, paddle wheel, rotor) to generate thrust.
  • twin-engine drives propulsion engines or turbo engines
  • the most economical implementation appears to be on primarily propeller-driven regional aircraft in the size class from approx. 30 passengers to 90 passengers. This corresponds to a propeller shaft power (sum of all propeller shafts - total power) of about 3,000 kW up to 8,000 kW.
  • Turbomachinery optimisation is usually done for the dominant phase of flight, primarily during the cruise phase, as this also represents the largest proportion and energy consumption over the mission. Equivalent, high operating efficiency is not possible across all flight phases. While maximum efficiency can be designed and achieved for the clearly dominant cruise flight for medium and long distances, the operating conditions for short and ultra-short distances are much less dominant and more varied. As a result, engines for regional and short-haul aircraft are operated much less at optimal efficiency over the entire mission and have worse specific fuel consumption per passenger relative to long-haul.
  • the different drop in propulsive and thermal efficiency plays a role here, since, for example, a propeller aircraft flies at lower altitudes and experiences lower, altitude-dependent thrust losses in cruise flight than, comparatively, a long-haul aircraft with a turbofan engine.
  • long-range engines run much more consistently at high power and are operated with good efficiency, both in climb and cruise, while regional aircraft have much wider power ranges, e.g. in take-off, climb and cruise.
  • the invention is applicable and usable in general for all segments of aircraft, airplanes and rotorcraft, for which hydrogen fuel is targeted objective for operation. This can be driven by fuel cost, existing infrastructure and to eliminate CO2 emissions. In all these cases the regulatory requirements concerning reserve fuel can significantly limiting the efficiency or missions can be operated.
  • the invention gives the possibility of choosing the best (lowest impact) fuel for each flight level and flight phase.
  • the aircraft operator can vary and choose the fuel according to the lowest emission impact of the aircraft depending on route and flight level. It will be assisted by a software supported system in the aircraft cockpit. The result will be an emission optimised aircraft operation.
  • the aircraft operator can vary and choose the fuel according to the lowest emission impact of the aircraft depending on route and flight level. It will be assisted by a software supported system in the aircraft cockpit. The result will be an emission optimised aircraft operation.
  • Another important aspect of the invention is the structural integration of the propulsion system within a given aerostructure of a transport aircraft.
  • the propulsion system together with the hosting nacelle assembly is attached to the wing of the aircraft the structural components of a wing is explained in more detail.
  • the wing structure on which the engines and hydrogen tanks are attached on the left and right, is divided into a load-bearing primary structure and an attached secondary structure with movable control surfaces, such as ailerons and flap systems.
  • ailerons and flap systems movable control surfaces
  • a wing box is used to take up bending and torsion loads.
  • the latter result, among other things, from the air loads and the weight and acceleration loads of the components attached to the wingbox, such as engines and hydrogen tanks.
  • Other components of the primary structure are ribs, spars and wingboxes.
  • the spars run along the spanwise direction of the wing and absorb the bending moment caused by lift, as well as the weight and acceleration load of the attached engines and tanks.
  • the spars are designed as I-beams or integrated into a wingbox.
  • the upper and lower areas, which are loaded with normal forces, are referred to as the spar cap.
  • the vertical walls, known as the spar web are subjected to shear and normal forces perpendicular and parallel to the chord direction.
  • Spars in sheet metal construction are designed as tension box girders.
  • a shear web with vertical posts subjected to compression is inserted between the spar caps, which are made of extruded sections or are milled integral components.
  • the ribs run either perpendicular to the spar or in the direction of flow from the wing nose to the trailing edge. They stiffen the wing skin and thus ensure the aerodynamic profiling of the wing. Spars are used to reinforce the torsion nose or, in the downstream area, as bearing spars to attach the aileron or flap hinges.
  • a closed box called a wing box
  • a spar box consisting of a front and rear spar or by a shell construction.
  • a wing box between 15% and 65% of the wing depth is sufficient to meet the torsional and bending strength requirements.
  • the wing box is usually extending in spanwise direction of the wing, having a polygonal, preferably rectangular cross section defined by an upstream and downstream wing spar and a upper and lower spar cap.
  • Another aspect of the invention is the effect of wing torsion due to landing shock.
  • wing torsion due to landing shock.
  • wing torsion One of the critical effects of landing is wing torsion, which can have structural and aerodynamic consequences.
  • the first is shock absorption and load transfer: The landing gear absorbs most of the impact forces, but some of the loads are transmitted to the aircraft structure, including the wings.
  • the second are ground reaction forces: If the aircraft experiences an uneven landing (e.g., one wheel touching down before the other), asymmetric forces induce torsional loads on the wing structure.
  • the third are aerodynamic forces: During landing, the wings may still be generating some lift, and the interaction between aerodynamic and structural forces can contribute to torsional stress.
  • Wing torsion can also lead to aileron reversal, where the movement of control surfaces (like ailerons) becomes less effective or even opposite to the pilot’s input due to aerodynamic distortion. Flutter risk is another behavioural consequence. If the torsion induces oscillatory motion, it can contribute to aeroelastic effects like flutter, where vibrations become self-sustaining and may lead to catastrophic failure. Repeated exposure to landing-induced torsional stress can accelerate material fatigue, requiring more frequent inspections and maintenance.
  • the effect of the invention by balancing the weight from the propulsion system attached to the wing box has a significant effect to mitigate torsion. I.e. by installing the internal combustion engine upstream the wing box and the hydrogen fuel tank downstream the wing box reduces the torsion effects on the wing box, minimizing the need- for additional and heavy reinforcement of the wing box or other structural elements of the wing.
  • flap systems arises from the conflicting demands for the highest possible cruising speed and the lowest possible take-off and landing speed, both of which are to be achieved with one and the same wing.
  • the cruise flight is performed with the flaps retracted (clean wing), while various flap systems can be extended in stages or continuously at the wing leading edge and/or at the wing trailing edge for take-off and climbing as well as for the approach and landing.
  • lift-increasing flaps is based on an increase in the camber of the airfoil and an increase in the wing area.
  • the main landing gear is located into lateral pods outside the fuselage cross section. This will keep space in the fuselage available for precious payload and the engine pylons often used for main landing gear storage is made available for ample hydrogen tank volume.
  • a tank mounting structure attached to the upstream side of the wing box, preferably to the upstream wing bar or center section of the ribs. Loads from the fuel tank mass can be more directly transferred to the primary structure of the wing. The reduces weight and complexity of the adaptation of the aerostructure.
  • a tank mounting structure similar to the upstream installed engine mounting structure is provided to hold the hydrogen fuel tank on each side of the propulsion system installation of the wing.
  • the tank mounting structure is fully integrated to the wingbox. Torsion and bending are predominantly transmitted to ribs and wing panels, querschub (shear force) predominantly to wing rear-spar.
  • querschub shear force
  • the installation downstream of the wing box and below the wing surface minimizes the drag effects of the installation. As an effect precious space in the fuselage remains untouched and remains fully available for passengers, freight and other payload or necessary systems.
  • the propulsion architecture described in this invention can also be sized and integrated as a retrofit or derivative variant for existing aircraft.
  • Fig. 1 a top-down view of a twin-engine regional transport aircraft featuring a dual-fuel propulsion system
  • Fig. 2 a schematic illustration of the propulsion system installation on one side of the wing of the aircraft according to Fig. 1 ,
  • FIG. 3 an enlarged cut-off illustration of the propulsion system according to Fig. 2,
  • Fig. 4a a side view of the propulsion system according to Fig. 2,
  • Fig. 4b a top view of the propulsion system according to Fig. 2,
  • Fig. 5a an enlarged top view cut-off illustration of the propulsion system according to Fig. 2, illustrating the flap actuator within the tank mounting structure,
  • Fig. 5b an enlarged side view cut-off illustration of the propulsion system according to Fig. 2, illustrating the flap actuator within the tank mounting structure,
  • Fig. 6 a chart representing power over mission time and altitude over mission time of a typical 300 nm mission profile
  • Fig. 7 a schema of the propulsion system illustrating the piping and components of both fuel systems
  • Fig. 9a to 9b an arrangement and size of an H2 tank in an aircraft with exclusive combustion of H2 (Fig. 2a) and with a “dual fuel” concept (Fig. 2b) and combustion of F ⁇ and kerosene,
  • Fig. 10 the diameter of a possible tank in an aircraft cabin
  • Fig. 11a a comparison of ranges with H2 flight alone and "dual fuel" flight
  • Fig. 11 b a comparison of ranges for different types of propulsion
  • Fig. 12 an arrangement of H2 tanks in a dual fuel aircraft
  • Fig. 13 an arrangement of H2 tanks in a dual fuel aircraft with kerosene tanks in the wing
  • Fig. 14 a side view of a propulsion engine for dual fuel operation
  • Fig. 15 a system diagram for a dual fuel aircraft
  • Fig. 16a to 16b an aircraft with a dual fuel engine and critical zones in the area of the turboprop engine.
  • FIG. 1 A typical installation configuration of a dual fuel propulsion system 10 for a two-engine regional transport aircraft 20 is shown in the embodiment according to Fig. 1 , using the legacy airplane D328-100 as an example for the airframe.
  • the airplane 20 is designed as a conventional shoulder-wing aircraft 20 with a tail-sided T-tail empennage 21.
  • the port and starboard side wings 26 are attached to the fuselage 22 via a wingbox 29.
  • the fuselage 22, is an assembly of cylindrical and conical sections and designed with a pressurized cabin, a cockpit 23 in the front nose of the fuselage 22 and the passenger compartment 24. At the rear, the pressurized cabin is closed off by a pressure-tight pressure bulkhead 25. Further aft, the fuselage 22 tapers conically and carries the tail cone 27.
  • the wings 26 are attached to the fuselage's 22 upper side, tangent to the fuselage's upper side.
  • the freight or baggage compartment 201 located with an outward facing door 202 to handle freight and baggage.
  • the pressure bulkhead 25 is located between the tail cone 27 which carries the tail empennage 21 and the freight I baggage compartment 201.
  • the two gas turbine engines 31 and 32 are housed in the engine nacelles 33 and 34 respectively, one each mounted on the left and right wing 26.
  • the gas turbine engines 31 and 32 are housed in the engine nacelles 33 and 34 respectively, one each mounted on the left and right wing 26.
  • the gas turbine engines 31 and 32 are housed in the engine nacelles 33 and 34 respectively, one each mounted on the left and right wing 26.
  • the gas turbine engines 31 and 32 are housed in the engine nacelles 33 and 34 respectively, one each mounted on the left and right wing 26.
  • 31 and 32 comprises a compressor driven by a turbine, a combustor that generates hot gases that drive the turbine and a set of fuel nozzles that supplies fuel to the combustor.
  • Corresponding gas turbines 31 , 32 with integrated reduction gearboxes 35, 36 are available, for example, under the designation PW 119C from Pratt & Whitney Canada turboprop engine.
  • the multi-bladed and adjustable propellers 37, 38 are driven by the engines 31 and
  • propeller blades receive the power for heating from a transmission unit.
  • a spinner 39 is provided on the hub of the propeller 37, 38 to reduce the drag of the hub.
  • Left- and right-wing kerosene tanks 40 are integrated in the wings 26 to supply fuel to the two gas turbines 31, 32 via fuel lines and systems not shown further.
  • the aggregates consisting of controller 60 and transmission device 61 are arranged in a bow-side area of the fuselage 22 in front of the wing 26 and outside the pressurised cabin.
  • the central controller 60 is provided to control the separate fuel systems.
  • the controller 60 controls the supply the engines with the first and second fuel.
  • Typical parameters of the controller 60 to be controlled and monitored are the fuel flow and pressure, the speeds of the power and high-pressure shafts and turbine temperature of the gas turbines 41 , 42.
  • a key conceptual feature of the transport aircraft 20 is its dual fuel propulsion system 10 as illustrated in the Fig. 2, 3, 4a, 4b, 5a and 5b.
  • the dual fuel propulsion system 10 enables the aircraft 20 to perform a flight mission - as illustrated in Fig. 6 - comprising sections of at least take off, climb, cruise, decent, approach and landing during which first liquid hydrogen fuel and a second kerosene-based jet fuel is selectively useable.
  • the gas turbine engines 31, 32 are capable of generating propulsive thrust using selectively the first liquid hydrogen fuel and the second kerosene-based jet fuel.
  • the aircraft 20 provides two independent sets of fuel systems, including a. a first fuel system 50, feeding the first fuel - liquid hydrogen fuel - from a first set of hydrogen fuel storage tanks 41 to the gas turbine engines 31, 32 and b. a second fuel system 51 feeding the second fuel - kerosene based jet fuel - from a second set of fuel storage tanks 40 to the gas turbine engines 31, 32.
  • LH2 or first fuel system for the distribution, vaporization, and feeding of LH2 from the first set of tanks 50 to the turbines 31 , 32: LH2 requires cryogenic
  • the gas turbines 31 , 32 are having hydrogen direct burning combustors (for H2 combustion): In H2 combustion airplanes LH2 is vaporized in an evaporator 55 and vaporous H2 burned in the combustor of the gas turbines 31 , 32, much like kerosene.
  • the operating pressure of the hydrogen tank 41 is kept above 3 bars to enable operation without hydrogen-pump on the A/C-side. Also, no pump is required before the evaporator 55. Excess-heat from gas turbines 31 , 32 is used to evaporate hydrogen.
  • Set of hydrogen tanks 50 hydrogen can be stored as pressurized gas or in liquid form. While gaseous storage can be suitable for shorter flights and is commercially available, the embodiment focuses on liquid hydrogen (LH2) storage tanks 50 as they require roughly half as much volume and consequently, they are significantly lighter than tanks for gaseous hydrogen. This is especially important for short-range segments, where the aircraft 20 will carry several hundred kg of hydrogen per flight. Compared to kerosene, LH2 tanks are still about four times as big. Since LH2 needs to remain cold and heat transfer must be minimized to avoid vaporization of hydrogen, spherical or cylindrical tanks 41 are provided to keep losses low. The first set of tanks 41 use lightweight material for double-insulated tank walls and insulation 52.
  • the tank mass is expressed by the gravimetric index and is defined as the weight of the LH2 fuel mass in relation to the full weight of the LH2 tank filled with maximum LH2 fuel.
  • the latest concepts for commuter aircraft 20 have a gravimetric index of up to 35 percent.
  • Illustrations Fig. 2, 3, 4a, 4b, 5a and 5b are focusing on the structural integration of the dual fuel propulsion system 10 in the aerostructure of the starboard side wing 26.
  • the integration on the port side of the wing 26 is similar if not almost identical.
  • the gas turbine engines 31 , 32 are supported by an engine mounting structure 35 attached to the upstream side of the wing box 29. Also, the hydrogen fuel storage tank 41 is supported by a tank
  • the nacelle assembly’s external shape is a surface resulting from the following conditions:
  • the wingbox 29 is extending in spanwise direction of the wing 26, having a rectangular cross section defined by an upstream and downstream wing spar 291 and a upper and lower spar cap 292.
  • the wing 26 furthermore comprises ribs 293 to stiffen the wing skin 294 extending perpendicular to the spar 291 and which are separated in spanwise direction, whereas the wing box 29 is supported by a centre section 2931 of the ribs 293.
  • the engine mounting structure 35 also known as the pylon, transfers loads from the engines 31 , 32 to the wing box 29 in several ways.
  • the primary loads include thrust, weight, and aerodynamic forces.
  • Thrust Loads These are longitudinal forces generated by the engine's thrust. The pylon transfers these forces to the wing box 29, ensuring the engine stays in place and contributes to the aircraft's forward motion.
  • Weight Loads The weight of the engine is supported by the pylon, which transfers this load vertically to the wing box 29.
  • Aerodynamic Loads These include forces due to airflow around the engine 31 , 32. The pylon must handle these loads and distribute them to the wing box 29 without causing undue stress.
  • tank mounting structure 36 transfers loads from hydrogen fuel storage tanks to the wing box 29.
  • the engine mounting structure 35 is attached to the upstream side of the wingbox, namely to the upstream wing spar 291 and to the centre section 2931 of the ribs. It also transfers loads to wing skin 294.
  • the tank mounting structure 36 is attached to the downstream side of the wingbox 29, preferably to the downstream wing spar 291 and to the centre section of the ribs 2931. It also transfers loads to wing skin 294.
  • the wing 26 comprises two adjoining ribs 293 which are extending upstream into the wingbox 29 and downstream towards the wing trailing edge 261 , whereas downstream rib sections 2932 defining the left and right side of the tank mounting structure 36 and the hydrogen tank 41 is attached to the downstream rib sections 2932.
  • Inertial forces from the hydrogen tank 41 in flight direction are introduced over a significant share of the tank-support box 36 length in order to avoid local tension peaks.
  • the hydrogen tank 41 is attached to the lower side of the tank mounting structure 36 and is strapped to the tank mounting structure 36 and primarily to the rib downstream section 2932 by a fail-safe design of 3 straps 361.
  • the wing 26 To increase the wing’s lift during take-off and landing the wing 26 features trailing flaps 262 on each side. It is a flap 262 which pivots back and downwards, increasing the camber and chord of the wing and providing a gap between flap 262 and wing 26, thus increasing CA of the wing 26. When extended, the flap 262 forms part of the upper surface of the wing 26, like a plain flap 262, it slides rearward before lowering.
  • the flaps 262 are divided into two segments, an inboard and an outboard flap segment 263, 264 to ensure proper aerodynamic performance.
  • the outer flap segment 264 extends between the tank mounting structure 36 and the wing tip and the inner flap segment 263 extends between the tank mounting structure 36 and the fuselage 22.
  • Each flap segment 263, 264 is attached to the wing box 29 through a series of flap tracks 265 and carriages.
  • the flap segments 263, 264 are mounted on flap tracks 265 that allow them to move aft and down when deployed. These tracks 265 are securely fastened to the wing box
  • Hinges and mechanical linkages connect the flap segments 263, 264 to the wing box 29. These components help the flaps 262 deploy smoothly and maintain alignment with the wing 26.
  • the flaps 262 are extended and retracted using electric or hydraulic actuators 266. These actuators 266 are linked to the flaps 262 and are also mounted to the wing box 29, ensuring controlled movement and driving the inner and outer flap segment 263, 264.
  • the flap actuator 266 is hosted inside the tank mounting structure 36.
  • the flap 262 is separated in the vicinity of the tank mounting structure 36. Both the inbound and outbound flap segments 263, 264 are connected by a flap spar 267 which is connected to the flap actuator 266 to provide a uniform movement of both flap segments 263, 264 when extracting and retracting. Together, these elements ensure that the flap segments are securely attached to the wing box 29 and can function effectively to increase lift during take-off and landing.
  • the evaporator 55 is integrated into tank-mounting structure 36. Access panels on the upper side of the wing skin 294 give access to the inside of the tank-mounting structure 36.
  • the ratio between length IT of the hydrogen tank (41) and the nacelle length IN, which is defined as the length between spinner (39) and the tail section 331 of the nacelle assembly is defined by the formula
  • the engine exhaust pipe 311 exits the nacelle 33, 34 between the gas turbine engine 31 , 32 and the hydrogen tank 41 and is tilted outboards giving a thrust component on the flight axis.
  • the diagram in Fig. 6 shows a typical 300 nm mission profile for the aircraft 20 configured for SAF as first fuel and H2 fuel hosted in multiple tank modules.
  • the fuel and weight calculation are as follows:
  • the diagram and the table demonstrate the required tank volume for hydrogen is thus minimised to maximise the possibilities for integration into the aircraft.
  • the diagram also shows the power requirement in kW and flight altitude in ft over the time of a typical 300 NM flight mission. The most important phases of the typical flight mission are explained below:
  • Take-off is the phase of flight in which the aircraft 20 undergoes a transition from moving on the ground (taxiing) to flying in the air, usually starting on a runway.
  • the engines are operated at full power during take-off.
  • the aircraft After take-off, the aircraft climbs to a certain altitude (in this case 25,000 ft) before flying safely and economically to its destination at this altitude.
  • a certain altitude in this case 25,000 ft
  • Cruise flight is the portion of air travel where flying is most fuel efficient. It takes place between ascent and descent phases and is usually the largest part of a journey. Technically, cruise flight is conducted at a constant airspeed and altitude. Cruise flight ends when the aircraft approaches the destination, where the descent phase begins in preparation for landing. For most commercial passenger aircraft, the cruise phase consumes most of the fuel.
  • the descent during a flight is the section in which an aircraft loses altitude.
  • the descent is an essential part of the landing approach.
  • Other partial descents may be to avoid traffic, to avoid poor flying conditions (turbulence or bad weather), clouds (especially under visual flight rules), to enter warmer air (if there is a risk of icing) or to take advantage of the wind direction of another altitude.
  • Normal descents take place at a constant airspeed and angle of descent.
  • the pilot controls the angle of descent by varying the engine power and angle of attack (lowering the nose) to keep the airspeed
  • the approach and landing are the final part of a flight where the aircraft returns to the ground.
  • airspeed and rate of descent are reduced enough to maintain a predetermined glide path (3-degree final approach to most airports) to the touchdown point on the runway.
  • Speed reduction is achieved by reducing thrust and/or creating greater drag using flaps, landing gear or airbrakes.
  • the pilot performs an intercept (flare) to initiate a soft landing.
  • Landing and approach procedures are mostly performed using an instrument landing system (ILS).
  • ILS instrument landing system
  • FIGS 8a to 8c show the possible arrangement of H2 tanks in an aircraft.
  • the H2 can be present in different states of aggregation.
  • the density of H2 is 0.090 kg/m 3 .
  • the density can be significantly increased for compressed or pressurised hydrogen.
  • Compressed hydrogen can be stored in hydrogen tanks with pressures of 350 bar to 700 bar for mobile hydrogen storage. At a pressure of 700 bar for compressed hydrogen, the density rises sharply to 42 kg/m 3 .
  • This type of hydrogen is referred as CGH hydrogen. Cooling to very low temperatures is not necessary, but the weight of the tanks increases considerably.
  • the hydrogen can also be liquefied at very low temperatures.
  • This type of hydrogen is known as LH.
  • the density is then 71 kg/m 3 and is higher than that of compressed hydrogen
  • the acceptable density makes it possible to use liquid F ⁇ with a temperature below -252.87 °C and a pressure of 20 bar to drive the propulsion engine, i.e. the turbine or piston engine, i.e. liquid H2 cooled to low temperatures is used.
  • the tank for the liquid H2 is very heavily insulated.
  • Figures 8a to 8c show possible locations of the tanks in the fuselage 10 of an aircraft 1 where tanks for liquid H2 could be arranged.
  • a first possibility shown in Fig. 8a is to locate the tank in the rear part 20.1 of the fuselage.
  • An accommodation in the rear part of the fuselage 20.1 is shown in Fig. 8c, but here further tanks are still arranged
  • a tank 20.1 .1 is in the rear part of the fuselage and a tank 20.1.2 in the front part of the aircraft.
  • liquid H2 as fuel for aircraft engines, e.g. turbines with propellers, which are arranged below the wings and labelled with reference number 30 and tanks in the fuselage of the aircraft, is, that long supply lines are required. Long lines increase the weight, require more interfaces, which can only be sealed with great effort in the case of hydrogen, and make it more difficult to pump the liquid hydrogen, which can vaporise especially during take-off procedures and when the lines are still warm, making it more difficult to pump the hydrogen.
  • H2 fuel is lighter than kerosene
  • the H 2 tanks are designed to be considerably heavier than in the case of kerosene, which is why the H2 tanks account for a large proportion of the weight.
  • FIG. 9a shows a tank 200 arranged in a fuselage 10 of an aircraft 1. If such a hydrogen tank is installed in an aircraft fuselage required for acceptable ranges greater than 250 NM, fuelled exclusively with H2, a very large tank 200 is required. The seating capacity in the fuselage decreases to only 18 cabin seats at sufficient ranges.
  • the weight of the H2 due to the installation of the tank in the configuration shown in Fig. 9a, which allows operation solely with a H2 tank 200 in Fig. 9a is 845 kg, the diameter of the tank is 1 metre, and the tank length is 8 metres.
  • the size and weight of the tank can be significantly reduced if, as shown in Fig. 9b, hydrogen is not used the entire range of the flight two propulsion systems.
  • a fuselage designed for two different types of fuel is shown in Fig. 9b. Flying with two different types of fuel is referred to as "dual fuel" flying.
  • the tank which is designed for a flight with H2 is significantly smaller than the tank in Fig. 9a, so the weight of the H2 tank that has to be carried is only 474 kg instead of 845 kg.
  • the diameter of the tank is again 1 meter, the tank length decreases from 8 m to 4.70 m and the seating capacity increases by 10 seats to a total of 28 seats for a dual fuel aircraft with H2 tanks in the fuselage of the aircraft.
  • the reference numbers in Fig. 9b are again 1 meter, the tank length decreases from 8 m to 4.70 m and the seating capacity increases by 10 seats to a total of 28 seats for a dual fuel aircraft with H2 tanks in the fuselage of the aircraft.
  • 25.04.2025 are the same as in Fig. 9a, except that the differently sized tanks labelled 200 for are the pure H2tank and 300 for the reduced H 2 tank in operation with two different types of fuel for distances greater than 250 NM (nautical miles).
  • Figure 10 shows a cross-section through the fuselage 10 of aircraft 1 , with the H 2 tank 200, 300 with a diameter D of 1 m, which is installed in the fuselage.
  • the space required for the tank in the fuselage is clearly recognisable.
  • Figure 11a shows the increase in range due to the so-called “dual fuel” concept, aircraft.
  • a DO 328 Figure 11a is used as an example shows the payload range over distance.
  • the pay load range chart is a representation of the to compare transport performance of aircraft and is used them.
  • the "pay load range” indicates the weight of freight and/or passengers that can be transported over which range.
  • the aircraft can only cover a relatively short distance. For a D328 aircraft this is approximately 600 NM. For longer distances, the cargo must be reduced so that more kerosene can be loaded into the aircraft.
  • all "pay load range curves" follow the same course. All curves for both a conventional kerosene- fuel led D328 aircraft (curve 400) and a dual-fuelled aircraft (curve 410) show two inflection points K1 and K2. The first inflection point K1 shows the distance from which the cargo or number of passengers must be reduced so that the range can be increased by loading more kerosene.
  • kink point K2 As the tank after a further is full refuelling, greater ranges can only be achieved if the cargo is further reduced (kink point K2). As no further fuel, e.g. kerosene, can be carried, is the curve after point K2 very steep. The area after point K2 is uneconomical and is only for special flights used or ferry flights or for special purposes in a non-civilian environment.
  • FIG. 11a The curve "pay load range” over distance is shown and labelled in Figure 11a as 400 for a conventional, kerosene-only aircraft.
  • the aircraft in question is a D328 ECO.
  • Fig. 11a also shows the payload range of an aircraft that is operated with two fuels according to the invention, namely H 2 and kerosene.
  • the payload range in "dual fuel" operation over the distance in NM is labelled with reference number 410. It can be seen that short distances of less than 250 nautical miles (NM) can be operated by burning only H 2 and thus without CO 2 emissions. Longer flights require the use of kerosene.
  • the range of emission-free flight is labelled 500 and in Fig. 11a.
  • Fig. 11b shows the "pay load range" over distance curves for different types of propulsion. As in Fig. 11a the
  • curve labelled with 400 shows the payload curve over distance for a conventionally powered aircraft is shown.
  • the curve 410 shows the "payload curve" for an aircraft in "dual fuel” operation.
  • K1 and K2 denote the same points as in Fig. 11a in Fig. 11b.
  • the curve 460 in Fig. 11 b denotes schematically the "payload curve" for an aircraft operated only on hydrogen without the option for a second fuel.
  • the curve 450 shows schematically the curve powered only by fuel H2 with a fuel cell instead of direct combustion in a gas turbine.
  • . 470 shows the payload curve for an aircraft powered by a battery. A retrofit with these alternative propulsion technologies was assumed for Fig. 11b.
  • Fig. 12 shows an aircraft 1 , e.g. a D328, with two engines 610, 620.
  • Aircraft 1 is again an aircraft that is operated with two types of fuel, namely H2 as the first fuel and kerosene or SAF as the second fuel.
  • the internal combustion engines that serve as propulsion engines 610, 620 can be turboprop engines, piston engines or rotary piston engines, such as those used in regional aircraft, for example the D328.
  • the tank sizes for H2 in a dual fuel system can be minimised, it is possible to make the H2 tanks significantly smaller.
  • Such small tanks can be in the area of the nacelle 700.1 , 700.2 of the two propulsion units, preferably be arranged in the rear end with a maximum of 300 kg H2 in liquid form, preferably 200 kg H2.
  • a further advantage is that structural changes to the aircraft, particularly the wing 810, are not necessary or only simple local reinforcements of the inner wing.
  • a further advantage is that very short fuel lines can be used due to the proximity of the tank and the propulsion engine, and the fuel is fed practically directly from the tank to the propulsion engine. The advantage of shortening the lines is that increased integration and certification costs avoided and weight is saved.
  • FIG 13 like Fig. 12, shows an overall view of an aircraft 1 with a propulsion system according to the invention.
  • the same components as in Fig. 12 are labelled with the same reference numbers.
  • the two tanks or tank systems namely the first tank(s) 800.1 , 800.2 for holding H2, are arranged in the area of the nacelle 700.1 , 700.2.
  • the engines, particularly the turbines, are again labelled 610, 620.
  • the two tanks or tank systems namely the first tank(s) 800.1 , 800.2 for holding H2
  • the engines, particularly the turbines are again labelled 610, 620.
  • the two kerosene tanks i.e. the second tanks 900.1 , 900.2, are arranged in the wing 810.
  • Figure 14 shows a sectional view of a nacelle 700.1 with first tank 800.1 for H2. Also shown is the propulsion engine 610, which is designed as a turbo-prop engine. Also shown is the wing 810, in which the kerosene tank 900.1 is arranged.
  • Figure 15 shows the system diagram of an aircraft according to the invention.
  • the fuselage 1000 of the aircraft 1 is shown, where no tank is located.
  • the nacelle 700.1 accommodates the propulsion engine 610 and the first tank 800.1 for liquid H2.
  • the wing 810 contains the kerosene tank 900.1.
  • Figures 16a and 16b show an aircraft with dual fuel propulsion according to the invention.
  • the same components bear the same reference numerals.
  • Figs. 16a and 16b show the areas 1000.1 and 1000.2 in which no tanks are installed for safety and complexity reasons. This concerns the propeller area (1000.1) and the wing area (1000.2).
  • the areas 1000.1 and 1000.2 are selected so that in the event of mechanical failure of the propeller or turbine or compressor blades, parts cannot penetrate the fuel tanks. For safety reasons, fuel tanks are therefore not located in possible impact areas.
  • an aircraft with the "dual fuel” propulsion system according to the invention has the further advantage that no complications during certification arise in other areas (passenger safety, extinguishing concepts, avoidance of influence on other systems, change in centre of gravity).

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Abstract

L'invention concerne un aéronef de transport doté d'un système de propulsion à deux carburants. L'aéronef 20 est capable d'effectuer une mission de vol comportant au moins des sections de décollage, d'ascension, de croisière, de descente, d'approche et d'atterrissage pendant lesquelles le premier ou le second carburant sont sélectivement utilisables. L'aéronef 20 comprend • au moins un moteur à combustion interne 31, 32 apte à générer une poussée propulsive, qui est enfermé dans un ensemble de nacelle 33, 34 et de préférence supporté par une structure de montage de moteur 35 fixée au caisson de voilure 29, • deux ensembles indépendants de systèmes de carburant 50, 51, comprenant un premier système de carburant 50 qui alimente en carburant sous forme d'hydrogène, depuis un premier ensemble de réservoirs de stockage de carburant 41, le moteur à combustion interne 31, 32 et un second système de carburant 51, qui alimente en carburant pour avion à réaction depuis un second ensemble de réservoirs de stockage de carburant 40, le moteur à combustion interne 31, 32, • le moteur à combustion interne 31, 32 utilise sélectivement un premier carburant pour avion à réaction à base de kérosène et un second carburant sous forme d'hydrogène et • le premier réservoir de stockage de carburant 40 est logé dans l'ensemble de nacelle 33, 34 et de préférence supporté par une structure de montage de réservoir 36 fixée au caisson de voilure 29. L'invention offre une gamme d'avantages significatifs : Vol à impact 002-minimal (~ -90% sur des missions typiques) et proche de la pleine charge utile en cas de rétroinstallation ou de dérivation (> 85% charge utile reste). Système d'hydrogène accessible et facile (bonne maintenance). CG du réservoir d'hydrogène à proximité de l'aéronef CG-> seul petit décalage du poids vide cg, seul très petit décalage en vol lors de la combustion de l'hydrogène.
PCT/IB2025/054361 2024-04-29 2025-04-26 Aéronef à deux carburant avec réservoirs d'hydrogène intégrés Pending WO2025229474A1 (fr)

Applications Claiming Priority (6)

Application Number Priority Date Filing Date Title
US202463639900P 2024-04-29 2024-04-29
DE102024111926.8 2024-04-29
DE102024111926.8A DE102024111926A1 (de) 2024-04-29 2024-04-29 Antriebssystem insbesondere für mehrmotorige Fluggeräte, bevorzugt Flugzeuge
US63/639,900 2024-04-29
DE102025111554.0 2025-03-25
DE102025111554 2025-03-25

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