AIRCRAFT WITH A PROPULSION AND POWER SYSTEM FOR LOW-EMISSION CRUISE FLIGHT

Description

[0001] The invention relates to a multi-engine aircraft and a method for operating a multi-engine aircraft.

[0002] The term "aircraft" primarily refers to a motorized, fixed-wing aircraft. However, it also encompasses, for example, rotary-wing aircraft (rotor aircraft, helicopters) and motor gliders. Aircraft and their propulsion systems can be differentiated according to the applicable construction and certification regulations. EASA regulation CS-23 applies to light, fixed-wing, powered aircraft. It covers aircraft in the categories "normal," "utility," or "aerobatic" with a maximum of nine seats (excluding pilot(s)) and a maximum takeoff weight of 5,670 kg, as well as aircraft in the category "commuter" with a maximum of 19 seats (excluding pilot(s)) and a maximum takeoff weight of 8,618 kg. CS-25 is also an EASA construction regulation concerning type certification for large aircraft, particularly large, turbine-powered aircraft. This document focuses on multi-engine aircraft to be certified according to construction regulation CS-25.

[0003] Regional airliners are predominantly characterized by a design with straight, unswived wings and a cruising speed of 500 to 700 km / h. Today, turboprop engines are the primary power source for regional airliners in civil aviation. A prominent example of this aircraft category is the Dornier 328-100 (Dornier 328 TP. 2020. Available: https: / / 328.eu / wp-content / uploads / 2020 / 09 / D328-100.pdf [Accessed 28.09.2020]).

[0004] Turboprop (a portmanteau of turbojet and propeller) is a common term for a turboprop engine, often simply called a propeller turbine. A turboprop is a heat engine with continuous internal combustion (thermal turbomachine) and is primarily used for aircraft propulsion. Colloquially, an aircraft powered by a turboprop is also often referred to as a "turboprop."

[0005] This type of engine is characterized by relatively low specific fuel consumption, which is why it is primarily used in transport and short-haul aircraft. Another civilian application is in smaller business jets such as the TBM-850. In the military, turboprops are mainly used in tactical transport aircraft. Aircraft with turboprops are limited to speeds up to 80 percent of the speed of sound (Mach 0.8), which corresponds to approximately 870 km / h at an altitude of 8,000 m under normal conditions. In this speed range, turboprops are more economical than conventional turbine engines.

[0006] The turboprop engine consists of a gas turbine, which typically drives a propeller via a speed-reducing gearbox. The engine's thrust is primarily generated by the propeller; the working gas exiting the exhaust diffuser contributes only a maximum of 10% to the total thrust. This distinguishes the propulsion principle significantly from turbojet engines and makes it more similar to a turbofan. To generate thrust, the propeller moves a large volume of air, but this air is accelerated only weakly compared to turbojet engines. In contrast, pure turbojet engines accelerate significantly smaller volumes of the propellant to much greater speeds.

[0007] Depending on the airspeed, altitude and load, the angle of attack of the propeller blades is changed so that both the turbine and the propeller operate as consistently as possible within the optimal speed range.

[0008] The energy to drive the propeller is supplied by the gas turbine. It draws in air, which is compressed in an axial or radial, usually multi-stage, turbo compressor. This compressed air then enters the combustion chamber, where it is ignited by the fuel. The now hot, energy-rich combustion gas flows through the turbine, which is typically axial and multi-stage, expanding and cooling as it does so. The energy transferred to the turbine drives the turbo compressor via a shaft and, via a gearbox, the propeller. The exhaust gases are expelled to the rear.

[0009] Turboengine optimization is typically performed for the dominant flight phase, usually cruise, as this also accounts for the highest proportion of energy consumption over the mission. Equally high operating efficiency is not possible across all flight phases. While maximum efficiency can be designed and achieved for the clearly dominant cruise phase on medium- and long-haul flights, the operating conditions for short- and ultra-short-haul flights are significantly less dominant and more varied. Consequently, engines for regional and short-haul aircraft operate at optimal efficiency considerably less often over the entire mission and have a higher specific fuel consumption per passenger compared to short- and long-haul aircraft. The differing rates of decline in propulsive and thermal efficiency also play a role, as, for example,A propeller-driven aircraft flies lower and experiences lower altitude-dependent thrust losses during cruise flight than a comparable long-haul aircraft with a turbofan engine. This results in long-haul engines operating at a much more consistent high power output and good efficiency than during climb, while regional aircraft have significantly broader power ranges, for example, during takeoff, climb, and cruise flight.

[0010] Therefore, especially for regional aircraft with propellers, but also for other twin-engine aircraft with paddle wheels or rotors, very different thrust requirements arise over the course of the mission, which can be designed and operated more efficiently by hybridizing the thermal engines with an electric engine.

[0011] Added to this is the challenge of reducing CO2, NOx, and noise emissions. Decarbonization is a major challenge for aviation. The aviation sector emits more than 900 million tons of carbon dioxide (CO2) per year. Based on industry growth of 3 to 4 percent per year and an efficiency improvement of 2 percent per year, emissions would more than double by 2050. During the same period, the aviation industry (Air Transport Action Group - ATAG) has committed to a 50 percent reduction in CO2 emissions (compared to 2005). Furthermore, the European Union (EU) has set itself the goal of becoming carbon neutral with its Green Deal. Apart from CO2, aircraft influence the climate through emissions of nitrogen oxides (NOx), soot, and water vapor, as well as contrails and cirrus clouds. Therefore, their overall contribution to global warming is significantly higher than CO2 emissions alone.(Hydrogen-powered aviation A fact-based study of hydrogen technology, economics, and climate impact by 2050, May 2020. Available . https: / / www.fch.europa.eu / sites / default / files / FCH%20Docs / 20200507 Hydrogen%20Powered%20Aviation%20report FINAL%20web%20%2 8ID%208706035%29.pdf [Accessed: 24.09.2020]).

[0012] The question currently being asked is whether electric or hydrogen-powered aircraft will be used in the future to meet the aforementioned requirements. Possibly. Airbus, Rolls-Royce, GE, and Siemens believe they can solve the problem of reducing CO2, NOx, and noise emissions by replacing a turbofan engine with an electric motor, thus following the automotive industry on the path to electrically, or at least hybrid, powered vehicles ("Flightpath 2050 Europe's Vision for Aviation," [Online]. Available: https: / / ec.europa.eu / transport / sites / transport / files / modes / air / doc / flightpath2050.pdf [Accessed 14 March 2018]).

[0013] GE International is working on a corresponding hybridized turbofan propulsion system for twin-engine commercial aircraft, as revealed in disclosure EP 3 421 760 A1. In this system, an electric motor is coupled to the high-pressure shaft of one turbofan engine and to the low-pressure shaft of the other, second engine. An electrical energy storage system is provided to power the electric motors, enabling them to supply additional propulsion power to the coupled turbofan under certain operating conditions. Similar solutions are proposed by SNECMA in publication WO 2009 / 153471 A2 and UTC / PW in US 2029 / 0282121 A1. The aircraft manufacturer Boeing is also working on a similar hybrid propulsion concept, as evidenced by publications EP 3 556 659 A1 and EP 3 556 658 A1. Another example of an electro-hybrid propulsion system for an aircraft is described in DE 10 2014 224637 Al.WO 2018 / 175349 Al deals with the optimization of the flight profile taking into account the available energy resources on board a hybrid-electric powered aircraft.

[0014] However, the power-to-mass density of currently available battery technology remains problematic in order to provide significant electric drive power in concepts like those described above. Put simply, current battery technology does not offer a sufficiently high energy density; moreover, the power-to-weight ratio is not high enough. For example, Combustible fuels like kerosene have an energy density of approximately 40 MJ / kg, or about 12,000 Wh / kg. The energy density of the lithium-ion batteries used in the first E-Fan is about 60 times lower. The specific energy of the batteries is therefore only about 2% of that of the liquid fuel. As a reminder, the 167 kg batteries of the 600 kg E-Fan were sufficient for about one hour of low-speed flight. In comparison, the empty weight of a BAE 146 is about 24,000 kg. These figures seem to suggest that the battery weight for an electric aircraft would be 60 times greater than the fuel weight for a current aircraft for the same flight. (Batteries against Fossil Fuel, https: / / batteryuniversity.com / learn / archive / batteries_against_fossil_fuel (accessed: June 17, 2020). An example of a purely electric drive comes from Bell, as can be seen in publication US 2020 / 277069 Al.

[0015] To reduce the impact on the climate, the industry is exploring further concepts, such as a radically new technology that uses sustainable aviation fuels (SAF) in significant quantities as synthetic fuel (synfuel) temporarily as compensation in large quantities or a combination thereof. Hydrogen propulsion is one such technology.

[0016] In the 1980s, Tupolev tested alternative fuels for jet engines as part of the further development of the Tu-154. This resulted in the Tu-155 prototype, powered by liquid hydrogen or natural gas. In this three-engine aircraft, the right engine was powered not by kerosene, but by hydrogen or natural gas. In any case, the lessons learned from this experience show that converting large commercial aircraft to hydrogen requires a new design with large, heavy LH2 tanks. Furthermore, the increased weight of long-haul aircraft significantly increases energy consumption and thus costs.

[0017] Based on this, the object of the invention is to provide a multi-engine aircraft with a hybrid propulsion system, which nevertheless allows the use of internal combustion engines and, in particular, turbomachinery to be further optimized and emissions to be reduced. Furthermore, a method for operating the aircraft is to be provided, which, using a hybrid propulsion system, optimizes the typical phases of flight operation.

[0018] According to the invention, the problem is solved by a multi-engine aircraft having the features of claim 1 and by a method for operating a multi-engine aircraft according to claim 15.

[0019] Advantageous embodiments of the invention will become apparent from the dependent claims, the description and the accompanying drawing.

[0020] The propulsion system for multi-engine aircraft is based on a specific propulsion system architecture utilizing two different energy sources or fuels, enabling emission-free operation of an aircraft for significant phases of a typical mission (zero-emission cruise). Due to the specific system architecture and the combination of two energy sources, the impact of the new technology and energy sources on current, already operational aircraft configurations is minimized, allowing for significantly earlier deployment compared to current zero-emission system proposals. A key aspect of the invention is the design of such a system based on a target mission and the overall aircraft architecture. Sizing and demonstrating the required efficiency is not possible without considering the entire aircraft system.

[0021] A key component of this invention is an aircraft architecture that enables zero-emission cruising, thereby eliminating aviation-specific non-CO2 emissions at high altitudes. The invention describes in its claims a solution for distributing the high thrust requirement during takeoff between two energy sources, thus allowing the dimensioning of the emission-free propulsion system, in terms of weight and volume, to be integrated into an existing aircraft platform using existing technology. The invention further discloses a solution for a complete aircraft system comprising the essential elements of a hybrid propulsion system, a fuel cell system, a hydrogen tank system, and a control unit.

[0022] The overall system in question consists of a propulsion system with two different energy sources. The propulsion system comprises two electro-hybridized internal combustion engines, turbomachines, or other thermal combustion engines, each coupled with a propulsor and a special control unit for the thermal and electric machines. The energy sources include fuels compatible with standard refueling, standard aviation fuel, or decarbonized, synthetic, or biofuels (SAF), as well as hydrogen for supplying a fuel cell system. The fundamental advantage of the invention is the ability to fly emission-free for extended periods. This requires considering both CO2 and non-CO2 emissions, with particular attention paid to the climate-damaging greenhouse effects resulting from combustion at high altitudes.

[0023] To achieve the goal of emission-free cruising with this system, while simultaneously enabling early deployment, the system's sizing and the utilization of both energy sources throughout the mission are essential to ensure compatibility with current aircraft designs. At least two key factors are involved: The amount of hydrogen and the resulting tank volume should have no impact on the aerodynamic surface area (no additional drag and therefore no loss of performance). The performance of the fuel cell system is primarily adapted to cruise flight; therefore, limiting its use to cruise and descent optimizes its use without having to combine additional, heavy electrical power sources, such as batteries or supercapacitors (any increase in fuel cell system power has an exponential effect on system weight and cooling requirements).

[0024] In summary, the propulsion system enables the use of new technologies for aviation, but based on today's assessable outlook, this will only be possible within the next decade: 1. Applicability to current aircraft concepts (wing-fuselage, with significant but currently assessable modifications). Essential for this is, for example, the continued use of the wing as a fuel tank for the internal combustion engines. 2. Fundamentally transferable and therefore applicable certification rules at the aircraft level. Verification of new technologies and energy sources, but conformity to basic requirements. 3. Use of the aircraft even without hydrogen and fuel cell systems with the given system design, e.g., in regions without hydrogen infrastructure. 4. Limitation of the system impact due to the possibility of adapting the use of the two energy sources. 5. Use of the fuel cell system as a continuous energy source. 6. Increased operational reliability in the event of a safety-relevant failure of a thermal engine or energy source. 7. Combustion of fuel from the internal combustion engines essentially not at cruising altitude, or...only for short periods of time, thus also avoiding climate-damaging greenhouse effects.

[0025] Regarding the primary objective of reducing emissions, this system architecture and specific energy utilization can reduce the following CO2 and non-CO2 emissions over the course of a mission, using the Do328 turboprop as a concrete example. The system design is based on a one-hour mission and the following energy utilization: a) Thermal combustion engine using sustainable, CO2-neutral fuel (SAF) primarily for takeoff, climb, approach, and landing. b) Hydrogen / fuel cell system for cruise and descent.

[0026] The energy requirement for this reference mission results in a distribution of approximately 60% for synthetic fuel (a) and 40% for hydrogen (b). This calculation simplifies by assuming that the overall efficiency of the thermal engines is roughly equivalent to that of the fuel cell system, including cooling and the necessary electrical components. This assumption can, of course, vary depending on the integration factor and the specific state of the art, but this variation is in the range of approximately 5% to 10%. This influences the detailed breakdown of CO2-neutral and emission-free flight. From today's technological perspective, there is greater potential in improving the efficiency of fuel cell systems compared to the potential improvements of turboprop engines.

[0027] Summary of emission reduction: Based on a turboprop aircraft with approximately 40 to 90 passengers, about 40% of the energy required for a one-hour mission can be generated emission-free using hydrogen and a fuel cell. This means zero CO2 emissions during cruise flight and no climate-damaging exhaust effects or contrail formation at cruising altitude (FL 250), which represent a significant portion of aviation emissions. Furthermore, the combustion component of approximately 60% can be produced CO2-neutrally by switching from conventional fuel to synthetic fuels. This aircraft and system architecture thus enables 100% decarbonized and CO2-neutral flight and approximately 40% emission-free flight with regard to CO2, greenhouse gas, and contrail effects at cruising altitude.

[0028] This architecture can be implemented in two different variations: a) The combustion engines are designed to handle takeoff, climb, and landing approaches, while the fuel cell system is used exclusively for cruise and descent. This minimizes the required hydrogen tank volume to optimize its integration into the aircraft. The aircraft retains the fundamental capability to operate without hydrogen and the fuel cell system. b) The combustion engines are supported during takeoff, climb, and landing by expanding the use of the fuel cell system, thereby reducing the power requirement of the combustion engines (downsizing). This allows for the use of smaller combustion engines. As a result, the zero-emission flight share is further increased to approximately 70% by reducing the proportion of mission-related energy derived from fuel combustion.An adjustment of the hydrogen tank volume to the increased energy demand is necessary.

[0029] In general, the invention relates to propulsion systems for small and large transport aircraft (CS-23 and CS-25) with twin-engine propulsion systems (piston engines or turbomachines) that convert thermal energy into mechanical drive shaft power and drive a propulsor (propeller, paddle wheel, rotor) to generate thrust. Implementation on primarily propeller-driven regional aircraft in the size class of approximately 30 to 90 passengers appears to be the most economical application. This corresponds to a power output at the propeller shafts (sum of all propeller shafts - total power) of approximately 3,000 kW to 8,000 kW.

[0030] The following factors were taken into account to determine the feasibility of this architecture, particularly its integration into current aircraft architectures: Energy requirements in the different flight phases; Power requirements in the different flight phases; Optimized distribution between thermal combustion / fuel use and electrical energy / hydrogen use; Hydrogen weight and tank volume; Overall system and component weights; Thermal and electrical efficiency of the fuel cell system

[0031] To assess the feasibility of this invention for specific aircraft classes, a technology outlook regarding performance and maturity is used as a basis, which is achievable from today's perspective.

[0032] Naturally, the applicability of this system architecture expands to higher performance classes and aircraft sizes as the technology outlook and introduction time change.

[0033] The following technology values ​​are used as a reference: Overall efficiency of fuel cell system: 50% Total weight of fuel cell system: 1500 kg (including electric drive)

[0034] Regarding liquid hydrogen storage tanks, tanks with a gravimetric index of 20 percent or higher are desirable. 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 weighs the same as the stored hydrogen.

[0035] Essentially, the invention offers at least four significant advantages over drive systems known from the prior art: a) Reduction of fuel consumption through optimized, adapted use of the internal combustion engines for the different flight phases. Hybridization of the drive units is of crucial importance here, incorporating a variably switchable electric motor-generator unit on a common gearbox. Through a gradual transfer of drive power to the electric motor, it can take over the power of the combustion engine during the transition from climb to cruise flight. b) The gearbox also plays a key role in the drive system according to the invention. Its input shafts enable an optimized speed range for both the combustion engine and the electric motor, thus ensuring optimal torque delivery to the propulsion shaft. This allows for a weight-optimized design and loss-optimized operating conditions for both the electric and combustion engines.c) Another important aspect of the invention is the reduction of the operating time of the thermal machines over flight hours, thereby lowering maintenance costs and extending maintenance intervals. Generally, electric machines require less maintenance and lower costs for the same rated power, since the electric motor-generator unit typically has no "hot" components. d) Finally, the propulsion system allows for increased safety in the event of a single engine failure. Particularly during the critical flight phases of takeoff, initial climb, and approach, the missing power can be immediately redistributed symmetrically to both sides via the electric machines.

[0036] In comparison to high-capacity, battery-based hybrid concepts currently under development, this results in a significantly weight-optimized design, as current battery concepts still only have a low specific energy density.

[0037] The required redundancy is provided by the two propulsion units, including their electric motor-generator units. Both propulsion units have the same power output and distribute thrust equally across the entire mission profile. The primary design constraint – the complete failure of one propulsion unit – is factored into the design of each unit, allowing the aircraft to be maneuvered within predefined limitations at every stage of flight until a safe landing.

[0038] Thrust adjustments throughout the flight mission are made equally by both engines and optionally with additional active pitch control. Turbomachinery optimization is typically performed for the dominant flight phase, primarily cruise, as this phase also accounts for the largest share and energy consumption over the mission. Equally high operating efficiency is not achievable across all flight phases. While maximum efficiency can be designed and achieved for the clearly dominant cruise phase on medium- and long-haul flights, the operating conditions for short- and ultra-short-haul flights are significantly less dominant and more diverse. Consequently, engines for regional and short-haul aircraft operate at optimal efficiency considerably less often over the entire mission and have a higher specific fuel consumption per passenger compared to short- and long-haul aircraft.

[0039] Among other factors, the differing rates of propulsive and thermal efficiency decline play a role, as, for example, a propeller-driven aircraft flies at lower altitudes and experiences lower altitude-dependent thrust losses during cruise flight than a comparable long-haul aircraft with a turbofan engine. This results in long-haul engines operating much more consistently at high power and with good efficiency, both during climb and cruise, while regional aircraft have significantly broader power ranges, for example, during takeoff, climb, and cruise.

[0040] Therefore, especially for regional aircraft with propellers, but also for other twin-engine aircraft with paddle wheels or rotors, very different thrust requirements arise over the course of the mission, which can be designed and operated more efficiently by hybridizing the thermal engines with an electric engine.

[0041] While previous hybrid propulsion systems and architectures for aircraft aimed to integrate additional or different arrangements of propulsive elements (propellers, rotors, paddle wheels) or additional, alternative energy sources such as batteries or fuel cells, this invention achieves higher efficiency without additional propulsors or energy sources. In contrast to previous hybrid concepts, such as the arrangement of multiple propulsors distributed across the wing span to generate better lift at low speeds—a significant weight penalty for a benefit during the short flight phases of takeoff and landing—or the use of energy systems whose power density is currently insufficient for larger aircraft, this propulsion system can be implemented with current technology and aircraft concepts, offering significant advantages.

[0042] The operating hours of the individual thermal engines, assuming consistent, alternating operation during cruise and descent, can be reduced by approximately 30% (using a 60-minute mission as a reference), which translates directly into longer maintenance intervals and reduced maintenance costs for the thermal engines.

[0043] The drive architecture described in this invention can also be designed and integrated as a retrofit variant for existing aircraft.

[0044] Further features, advantages, and effects of the invention will become apparent from the following description of preferred embodiments of the invention, as illustrated in the drawing. The drawing shows: Fig. 1a a top view of a twin-engine aircraft with a schematic representation of a hybrid propulsion system, Fig. 1b a side view of the twin-engine aircraft according to Fig. 1awith a schematic representation of a hybrid propulsion system, Fig. 2; a system sketch of a hybrid propulsion system with a schematic representation of the system architecture, Fig. 3a; a diagram of power demand and altitude during the operating phases of a typical 200NM mission of an aircraft with conventional gas turbines and electric motor-gearbox units, Fig. 3a; a diagram of accumulated energy demand during the operating phases of a typical 200NM mission of an aircraft with conventional gas turbines and electric motor-gearbox units, Fig. 4a; a diagram of power demand and altitude during the operating phases of a typical 200NM mission of an aircraft with miniaturized gas turbines and electric motor-gearbox units, Fig. 4a; a diagram of accumulated energy demand during the operating phases of a typical 200NM mission of an aircraft with miniaturized gas turbines and electric motor-gearbox units, Fig.Fig. 5 a system sketch of a hybrid drive system with a schematic representation of the system architecture in the primary operating mode and Fig. 6 a system sketch of a hybrid drive system with a schematic representation of the system architecture in the third operating mode.

[0045] A typical installation configuration of a hybrid propulsion system 10 for a twin-engine regional aircraft 20 is shown using the Dornier 328-100 as an example in Fig. 1a and Fig. 1bThe aircraft 20 is a conventional high-wing monoplane with a T-tail 21 at the rear. The partially cylindrical fuselage 22 houses a pressurized cabin containing the cockpit 23 and the passenger compartment 24. At the rear, the pressurized cabin is sealed by a pressure dome. Further aft, the fuselage 22 tapers conically and supports the T-tail 21. In the Dornier 328-100, a baggage compartment is located in the conical transition area.

[0046] In a conventional shoulder-wing configuration, the wings 26 are attached to the fuselage tube, tangent to it above. The hybrid-electric propulsion units 31 and 32 are housed in the engine nacelles 33 and 34, one of which is mounted on each of the left and right wings 26. The multi-bladed, adjustable propellers 61 and 62 are driven via reduction gearboxes, which are also integrated into the engine nacelles 33 and 34. To prevent undesirable icing on the propeller blades, they are electrically heated; the current for heating is supplied to the propeller blades by a transmission unit 80.

[0047] The in Fig. 1a The integrated system architecture of the propulsion system 10 in a twin-engine aircraft is described in further detail in Fig. 2The propulsion system 10 comprises two independently operable hybrid-electric propulsion units 31 and 32. Each hybrid-electric propulsion unit 31, 32 has a gas turbine 41, 42 with a flanged reduction gearbox 51, 52, to which a propeller 61, 62 with variable pitch is coupled. Corresponding gas turbines 41, 42 with integrated reduction gearbox 51, 52 are available, for example, under the designation PW 119C from Pratt & Whitney Canada. Left and right wing integral fuel tanks 43, 44 are integrated into the wings 26, which supply the two gas turbines 41, 42 with fuel via fuel lines and systems (not shown).

[0048] Each drive unit 31, 32 is assigned a motor-generator unit 71, 72, which is coupled on the drive side to the reduction gear 51, 52. Depending on the operating phase, the motor-generator unit 71, 72 can be operated as an electric motor or as a generator. In drive mode, the motor-generator unit 71, 72 transmits drive power via the reduction gear 51, 52 to the respective assigned propeller 61, 62. In generator mode, the motor-generator unit 71, 72 generates electrical power, which is supplied to a transmission device 80 for further distribution or storage. Two power converters 81 and 82 are functional components of the transmission device 80, one of which is assigned to each motor-generator unit 71, 72.

[0049] To supply the motor-generator units 71, 72 with electrical energy, the propulsion system 10 includes a fuel cell 73, which in turn is supplied with hydrogen via a fuel tank 74. In the fuel cell 73, hydrogen is converted into electricity, which then supplies the motor-generator unit 71, 72 with electrical power via the transmission device 80 and power converters 81 and 82 to drive the propellers 61, 62. Low-temperature proton exchange membrane (PEM) fuel cells are currently the most advanced and best suited for aviation applications. Supplementing this system with an optional energy storage device such as a battery helps to ensure rapid load tracking and the coverage of power peaks, thus optimizing the dimensioning of the fuel cell system.

[0050] Generally, hydrogen can be stored as a pressurized gas or in liquid form. While gaseous storage may be suitable for shorter flights and is commercially available, the invention focuses on liquid hydrogen (LH2) storage tanks, as they require about half the volume and are consequently much lighter than tanks for gaseous hydrogen. Since LH2 must remain cold and heat transfer must be minimized to prevent evaporation of the hydrogen, spherical or cylindrical tanks are required to keep losses low. In the Fig. 1a and Fig. 1b In the configuration shown, the spherical fuel tank 74 is housed in the conical rear fuselage 27, which can be used as a cargo space when the fuel tank 74 is removed.

[0051] The assemblies, consisting of controller 90, transmission unit 80, and fuel cell 73, are arranged in a forward-facing area of ​​the fuselage 22, in front of the wing and outside the pressurized cabin. The assemblies and the fuel tank 74 for supplying the fuel cell 73 form a moment equilibrium that is essentially neutral with respect to the center of gravity (CG) of the aircraft 20 (see Fig. 1b ).

[0052] In a secondary function, the waste heat removed by means of a cooling unit from the fuel cell 73 serves to de-ic exposed surfaces of the aircraft 20, such as the wing leading edges, air inlets of the gas turbines 41, 42 and leading edges of the T-tail.

[0053] A central controller 90 is provided for controlling the thermally and electrically generated drive power. This controller is connected to power converters 81 and 82 and the gas turbines 41 and 42. Depending on the operating phase, the controller 90 controls the motor-generator unit 71 and 72 via the power converters 81 and 82, controlling the output of electrical drive power and the electrical energy to be generated. It also controls the thermally generated drive power of the gas turbines 41 and 42. Typical parameters controlled and monitored by the controller 90 include the fuel supply, the rotational speeds of the power and high-pressure shafts, and the turbine temperature of the gas turbines 41 and 42.

[0054] In another embodiment, in Fig. 2An architecture is shown that is based on a DC network 101 and AC / DC converters 81, 82. Depending on the operating mode and power requirement, the power output of the fuel cell 73 can be supplied to the first and second motor-generator units 71 and 72, respectively, via the transmission device 80 and the AC / DC converters 81, 82, and the drive power can be transmitted to the propellers 61 and 62, respectively, via the reduction gears 51 and 52.

[0055] The diagram of Fig. 3aThe diagram illustrates the difference between an aircraft 20 with a propulsion system 10 according to the invention and the prior art, using a Dornier 328 equipped with two conventional engines as an example. The maximum power of the combustion engines is designed to perform takeoff and landing approaches, while the fuel cell system is intended exclusively for cruise and descent. The required hydrogen tank volume is thus minimized to optimize integration possibilities into the aircraft. Fig. 3a This represents the power requirement in kW and flight altitude in ft over the duration of a typical 200 NM flight mission. The most important phases of a typical flight mission are explained below: Start:

[0056] Takeoff is the phase of flight in which the aircraft transitions from movement on the ground (taxiing) to flying in the air, usually taking off from a runway. Typically, the engines are running at full power during takeoff. Climb:

[0057] After takeoff, the aircraft climbs to a certain altitude (in this case 25,000 ft) before flying safely and economically to its destination at that altitude. Passenger flight:

[0058] Cruise flight is the portion of an air journey when flying is most fuel-efficient. It occurs between ascent and descent phases and typically constitutes the largest part of a trip. Technically, cruise flight is conducted at a constant airspeed and altitude. Cruise flight ends when the aircraft approaches its destination, at which point the descent phase begins in preparation for landing. For most commercial passenger aircraft, cruise flight consumes the majority of the fuel. Descent:

[0059] The descent phase of an aircraft flight is the phase in which it loses altitude. It is an essential part of the landing approach. Other partial descents can be used to avoid traffic, adverse flying conditions (turbulence or bad weather), clouds (especially under visual flight rules), to enter warmer air (in case of icing), or to take advantage of wind direction at a different altitude. Normal descents occur at a constant airspeed and a constant angle of descent. The pilot controls the angle of descent by varying engine power and the angle of attack (lowering the nose) to maintain the airspeed within the specified range. At the beginning and during the descent phase, the engines will operate at low power. Approach & Landing:

[0060] The approach and landing are the final part of a flight, during which the aircraft returns to the ground. For landing, the airspeed and rate of descent are reduced to maintain a predetermined glide path (3 degrees final approach to most airports) to the touchdown point on the runway. This reduction in speed is achieved by reducing thrust and / or generating more drag using flaps, landing gear, or airbrakes. As the aircraft approaches the ground, the pilot performs a flare to initiate a smooth landing. Approach and landing procedures are usually performed using an Instrument Landing System (ILS).

[0061] Line A (dashed line)Power requirement of the propulsion system 10 over the mission with two hybrid-electric propulsion units 31, 32. The power requirement is highest during takeoff and decreases over the subsequent flight phases. For the takeoff, climb, approach, and landing phases, the power requirement is met solely by the gas turbines 41, 42 (primary operating mode). During cruise and descent, propulsion is provided by the motor-generator units 71, 72, supplied by the fuel cell 73 (third operating mode). During flight, the controller 90 manages the operation of the propulsion units 31, 32 and the transitions between the primary and secondary operating modes. The dotted area below line A represents the fuel consumption by the gas turbines 41, 42, and the checkered area represents the hydrogen consumption by the motor-generator units 71, 72 during the operating phases.The energy requirement for this reference mission results in a distribution of approximately 60% for fuel (SAF) and 40% for hydrogen.

[0062] Line B The solid line indicates the flight altitude in feet over the course of the mission. The flight altitude is highest during cruise flight and is approximately 25,000 feet.

[0063] According to the diagram Fig. 3b The corresponding accumulated energy demand is shown by the solid and dashed lines C and D during the aforementioned operating phases and modes. Line C represents the energy demand during the primary operating mode, and line D during the third operating mode.

[0064] The diagrams of Fig. 4a and Fig. 4bThe values ​​represent a second variant of the invention. In this variant, the internal combustion engines are supported during takeoff and landing by expanding the use of the fuel cell system, thereby reducing the power requirement of the thermal engines (downsizing). This involves the combined operation and power output of gas turbines 41 and 42 as well as engine-generator units 71 and 72, coordinated by controller 90 (secondary operating mode). One advantage is that smaller internal combustion engines can be used. This results in a further increase in the emission-free flight share to approximately 70% (checkered area in the figure). Fig. 4a ).

[0065] Line E (dashed line): The power requirement of the propulsion system 10 over the mission with two hybrid-electric propulsion units 31, 32 after Fig. 4a is basically the same as in Fig. 3aAs illustrated, the power requirement is highest during takeoff and decreases over the subsequent flight phases.

[0066] Line F (Solid line) describes the flight altitude in feet over the duration of the mission. The flight altitude is highest during cruise flight and is the same as during Fig. 3a approximately 25,000 ft.

[0067] The diagrams of Fig. 3a and Fig. 3bThis highlights a key aspect of the invention: for many aircraft, especially regional propeller-driven aircraft, the required thrust for takeoff is significantly higher than for cruise flight. Consequently, for a large portion of the flight mission, the thermal engines operate at only about half their capacity. This means that conventional turboprop engines operate outside their optimal operating point, which is closer to the point of maximum power output. In contrast, a hybrid-electric propulsion system can be optimized for both different operating modes.

[0068] Regarding the operating phases of the drive system 10, the following basic operating states result: 1. The Fig. 5Figure 1 shows a system diagram for takeoff, climb, and approach / landing. Both gas turbines 41 and 42 are in operation and drive propellers 61 and 62 via the intermediate reduction gearboxes 51 and 52 (solid lines). Thrust control is centrally managed by the hybrid propulsion controller 90 for gas turbines 41 and 42. (primary operating mode). The electric motor-generator units 71 and 72 do not provide any drive power (dashed lines). 2. The secondary operating mode for takeoff and climb is in Fig. 2 In this combined operating mode, the propulsors (61, 62) receive drive power from both the first and second internal combustion engines (41, 42) as well as from the first and second motor-generator units (71, 72). The corresponding power requirement and cumulative energy consumption are shown in the diagrams of the Fig. 4a and Fig. 4bto be seen. The reduced fuel consumption for the gas turbines due to the additional drive power from the electric motor-generator units 71, 72 is clearly evident here. 3. The system state for cruise and descent is shown in Fig. 6 This can be seen. The power / torque requirement drops to cruise flight levels. ( Fig. 3a and 4a ), The power output of the two gas turbines 41, 42 is reduced, while the propulsion power from the two motor-generator units 71, 72 can be transmitted equally to both propellers 61 and 62 via the coupled gearboxes 51, 52. During cruise and subsequent descent, the thrust requirement is adjusted via the hybrid propulsion controller 90. This controller is responsible for thermal and electrical control. (third operating mode).4. The system architecture according to the invention enables symmetrical thrust even in the event of a critical fault, such as the failure of an internal combustion engine, due to the power distribution via the electrical network. On the side of the failed internal combustion engine 41, 42, the motor-generator unit 71 or 72 can be switched on, thus compensating for at least part of the lost thrust. Reference symbol list

[0069] 10 Propulsion system 20 Aircraft 21 T-tail 22 Fuselage 23 Cockpit 24 Passenger compartment 26 Wing, left and right 27 Rear fuselage 31, 32 Propulsion units, left and right 33, 34 Engine nacelle, left and right 41, 42 Gas turbine, left and right 43, 44 Wing integral fuel tank, left and right 51, 52 Reduction gearbox, left and right 61, 62 Propeller, left and right 71, 72 Engine-generator units, left and right 73 Fuel cell 74 Fuel tank 80 Transmission unit 81, 82 Power converter 90 Controller A, B, C, D, E, FLines SPFocus

Claims

1. Multi-engine aircraft having a hybrid propulsion system (10), comprising • at least a first and a second hybrid-electric propulsion unit (31, 32), each comprising an internal combustion engine (41, 42), a motor-generator unit (71, 72) for transmitting propulsive power to a propulsor (61, 62), • wherein the propulsor (61, 62) is connectable to the internal combustion engine (41, 42) and / or the motor-generator unit (71, 72) for the transmission of drive power, • the first and second motor-generator units (71, 72) are connected to a transmission device (80) for distributing electrical power, • a fuel cell (73) for supplying the first and / or second motor-generator unit (71, 72) with electrical energy, • a controller (90) which is connected to the internal combustion engines (41, 42) and / or the transmission device (80) and / or motor-generator units (71, 72) and / or the fuel cell (73) to control the thermally and electrically generated drive power, • separate fuel tanks (43, 44) for supplying the internal combustion engines (41, 42) with fuel and the fuel cell (73) with cryogenic hydrogen, • an airfoil (26) accommodating the drive units (31, 32) and a fuel tank (43, 44), a fuselage (22), • wherein the drive unit (31, 32) is formed by a turboprop engine having one respective gas turbine (41, 42) that can be coupled to a speed-reducing gearbox (51, 52) to drive a propeller (61, 62), and • the motor-generator unit (71, 72) can be coupled to the gearbox (51, 52) under the control of the controller (90) depending on the operating mode, characterized in that at least the predominant volume of the fuel tanks (43, 44) for supplying the internal combustion engines (41, 42) is integrated in the airfoil (26) and at least the predominant volume of the fuel tanks (74) for supplying the fuel cell (73) is integrated in a tail region of the fuselage (22).

2. Multi-engine aircraft according to claim 1, characterized in that the hybrid-electric propulsion unit (31, 32) • in a primary operating mode, the propulsors (61, 62) receive the propulsion power predominantly or entirely from the combustion engines (41, 42), • in a secondary combined operating mode, the propulsors (61, 62) receive the drive power from the first and second internal combustion engines (41, 42) and from the first and second motor-generator units (71, 72), and • in a third operating mode, the propulsors (61, 62) receive the drive power from the first and second motor-generator units (71, 72).

3. Multi-engine aircraft according to one of the preceding claims, characterized in that the controller (90) effects a symmetrical distribution of the drive power to the propulsors (61, 62) in the operating modes.

4. Multi-engine aircraft according to one of the preceding claims, characterized in that the electrical drive power of the first and second motor-generator unit (71, 72) can be variably connected during the transition between the operating modes.

5. Multi-engine aircraft according to one of the preceding claims, characterized the hybrid-electric propulsion units (31, 32) each have a gearbox (51, 52) for transmitting the drive power, wherein the internal combustion engine (41, 42) and the motor-generator unit (71, 72) can be coupled to the propulsor (61, 62) by means of the gearbox (51, 52).

6. Multi-engine aircraft according to one of the preceding claims, characterized in that the change in the transmission of the drive power of the internal combustion engine (41, 42) and of the drive power of the motor-generator unit (71, 72) is carried out successively in such a way that the power delivered to the propulsors (61, 62) of the common propulsion unit (31, 32) remains approximately the same.

7. Multi-engine aircraft according to one of the preceding claims, characterized in that the propulsors (61, 62) are designed as propellers with pitch-adjustable blades and the controller (90) for controlling the drive power is connected to the pitch-adjustment mechanism.

8. Multi-engine aircraft according to one of the preceding claims, characterized in that, in a further operating mode, the drive power of the first or second internal combustion engine (41, 42) has failed completely or predominantly and electrical power is applied to the first or second motor-generator unit (71, 72) via the transmission device (80) from the fuel cell (73).

9. Multi-engine aircraft according to one of the preceding claims, characterized in that the transmission device (80) is designed as an alternating current network.

10. Multi-engine aircraft according to one of the claims 1 to 9, characterized in that the transmission device (80) is designed as a DC network, wherein an AC / DC converter (81, 82) is associated with each motor-generator unit (71, 72) which is connected to the controller (90) to control the speed of the propulsor (61, 62).

11. Multi-engine aircraft according to one of the preceding claims, characterized in that the internal combustion engines (41, 42) are operated with sustainable aviation fuel (SAF).

12. Multi-engine aircraft according to one of the preceding claims, characterized in that the fuselage (22) is provided in a tail region with a space for the formation of a cargo compartment, in which the fuel tank (74) for supplying the fuel cell (73) is arranged.

13. Multi-engine aircraft according to one of the preceding claims, characterized in that units consisting of controller (90) and / or transmission device (80) and / or fuel cell (73) are arranged in a nose region of the fuselage (22).

14. Multi-engine aircraft according to claim 13, characterized in the units and the fuel tank (74) for supplying the fuel cell (73) form a substantially neutral moment equilibrium with respect to the center of gravity (SP) of the aircraft (20) and / or the fuel cell (73) has a cooling unit, wherein the waste heat serves to de-ice exposed surfaces of the aircraft.

15. Method for operating a multi-engine aircraft (20) according to claim 2 or according to one of the preceding claims dependent on claim 2, characterized in that the propulsion system (10) is operated in a primary, secondary and third operating mode, wherein • the taxiing of the aircraft (20), in particular on aprons and taxiways, takes place in the primary or third operating mode, • take-off and climbing to cruising altitude takes place in the primary or secondary operating mode, • cruising and descending to approach altitude takes place in the secondary or third operating mode, • approach and landing takes place in the primary or secondary operating mode, and • in the event of failure of one combustion engine (41, 42), the flight is continued in secondary or third operating mode.