Airplane with drive and power system for low emission cruising

Abstract

The invention relates to a hybrid electric drive system (10) for a multi-engine aircraft (20). The drive system comprises at least first and second hybrid electric drive units (31, 32) each having an internal combustion engine (41, 42), an electric generator unit (71, 72) for transmitting drive power to a propeller (61, 62). The drive system (10) has a fuel cell (73) to supply the electric generator units (71, 72) with electric power. The fuel cell is in turn supplied with hydrogen via a fuel tank (74). In the fuel cell (73) the hydrogen is converted into electricity, which then supplies the power converters (81, 82) and the electric generator units (71, 72) via transmission means (80) with electric power in order to drive the propellers (61, 62). Advantage: Based on a turbo-prop aircraft with about 40 to 90 passengers, 40% of the energy can be represented during a 1-hour mission by hydrogen and fuel cell without emissions. This means less CO2 during cruising and also no climate-damaging exhaust gases and condensation trails at cruising altitude (FL 250), which represent the main proportion of aviation emissions.

Description

Technical Field

[0001] The present invention relates to a hybrid power drive system for a multi-engine aircraft, a multi-engine aircraft, and a method for operating a twin-engine aircraft. Background Technology

[0002] Aircraft should be understood, in particular, as motor-driven fixed-wing aircraft. However, the term "aircraft" also includes rotorcraft (rotor aircraft, helicopters) and motor-driven gliders. Aircraft and their propulsion systems can be distinguished according to applicable construction and approval regulations. EASA (European Aviation Safety Agency) Regulation CS-23 applies to light fixed-wing motor-driven aircraft. It covers aircraft in the "general, utility, or aerobatic" category with up to nine seats (no (or more) pilots and a maximum takeoff mass of 5,670 kg) and aircraft in the "commuter" category with up to 19 seats (no (or more) pilots) and a maximum takeoff mass of 8,618 kg. CS-25 is also a design specification issued by EASA for type approval of large aircraft, particularly large turbine-powered aircraft. In this case, a multi-engine aircraft certified according to the CS-25 construction specification is considered.

[0003] The main characteristics of regional transport aircraft are their straight, unswept wing design and cruising speed of 500 to 700 km / h. Currently, turboprop propulsion is the primary application area for civil regional transport aircraft. A prominent representative of this aircraft category is the Dornier 328-100 (Dornier 328 TP.2020. Available at: https: / / 328.eu / wp-content / uploads / 2020 / 09 / D328-100.pdf [accessed 09 / 28 / 2020]).

[0004] A turboprop engine (a man-made word combining "turbojet" and "propeller") is the common name for a turboprop turbine propeller (PTL), often simply called a propeller turbine. A turboprop engine is a continuous internal combustion heat engine (heat flow engine) primarily used for aviation propulsion. In layman's terms, aircraft powered by PTLs are commonly referred to as "turboprop aircraft."

[0005] This type of propulsion system is characterized by relatively low fuel consumption, which is why it is primarily used in transport and short-haul aircraft. Another area of ​​application in civil aviation is smaller business jets, such as the TBM-850. In the military, turboprop aircraft are mainly used for tactical transport. The maximum speed limit for turboprop aircraft is 80% of the speed of sound (Mach 0.8), which is equivalent to approximately 870 km / h at an altitude of 8,000 meters under normal conditions. Within this speed range, turboprop engines are more economical than pure turbine propulsion.

[0006] A turboprop thruster consists of a gas turbine, which typically drives a propeller via a reduction gearbox. The thrust of the thruster is primarily generated by the propeller; the working gas exiting the diffuser accounts for at most 10% of the total thrust. Its propulsion principle differs significantly from that of a turbojet engine, and is more similar to that of a turbofan engine. The propeller pushes a large amount of air to generate thrust, but compared to a turbojet engine, this is only a slight acceleration. On the other hand, in the case of a pure turbojet thruster, a significantly smaller amount of driving medium provides much greater acceleration.

[0007] Depending on flight speed, altitude, and load, the angle of attack of the propeller blades will vary to ensure that the turbine and propeller operate as consistently as possible within the optimal speed range.

[0008] The energy to drive the propeller is provided by a gas turbine. It draws in air, which is compressed in an axial or radial (usually multi-stage) turbocompressor. The compressed air then enters the combustion chamber, where fuel is burned. The now-heated, high-energy combustion gases flow through the main axial and multi-stage turbines, where they expand and cool. The energy transferred to the turbines drives the turbocompressor via the shaft on one hand, and the propeller via the gearbox on the other. Exhaust gases are discharged to the rear.

[0009] Turbines are typically optimized for the primary flight phase, usually cruise, as this is the phase with the highest energy consumption share throughout the mission. It's impossible to maintain equally high operating efficiency across all flight phases. While maximum efficiency can be designed and achieved for cruise flights that are clearly dominant for medium- and long-haul routes, the operating conditions for short- and ultra-short-haul routes are far fewer and more varied. Therefore, regional and short-haul aircraft engines operate at optimal efficiency far less frequently throughout the mission compared to short- and long-haul flights, resulting in lower fuel consumption per passenger. Different declines in propulsive and thermal efficiency also play a role here, for example, because propeller aircraft fly lower during cruise and experience less thrust loss due to high dependence compared to long-haul aircraft equipped with turbofan thrusters. Therefore, long-haul engines operate more consistently at high power and are more efficient than during climb, while regional aircraft have a wider power range, such as during takeoff, climb, and cruise.

[0010] Therefore, the thrust requirements of missions that can be designed and operated more efficiently by combining heat engines with electric motors are very different, especially for regional aircraft with propellers, but also applicable to other twin-engine aircraft with paddle wheels or rotors.

[0011] In addition to these challenges, there are the issues of reducing carbon dioxide, nitrogen oxides, and noise emissions. Decarbonization is a major challenge for the aviation industry. The aviation industry emits over 900 million tons of carbon dioxide (CO2) annually. Assuming the industry grows by 3% to 4% annually and efficiency improves by 2% annually, emissions will more than double by 2050. During the same period, the aviation industry (Air Transport Action Group – ATAG) has pledged to reduce CO2 emissions by 50% (compared to 2005 levels). Furthermore, the European Union (EU) has set a carbon neutrality target for itself through the Green Deal. Besides CO2, aircraft also impact the climate by emitting nitrogen oxides (NOx), soot and water vapor, contrails, and cirrus clouds. Therefore, their "comprehensive" contribution to global warming is far greater than just CO2 emissions. (A fact-based study on the technological, economic, and climate impacts of hydrogen-powered aviation to 2050, May 2020. Available at: https: / / www.fch.europa.eu / sites / default / files / FCH%20Docs / 20200507_Hydrogen%20Powered%20Aviation%20report_FINAL%20web%20%28ID%208706035%29.pdf [Accessed September 24, 2020]).

[0012] The current question is whether electric or hydrogen-powered aircraft will be used in the future to meet the previously proposed requirements. This is a possibility. Airbus, Rolls-Royce, General Electric, and Siemens believe they can address the issue of reducing CO2, NOx, and noise emissions by replacing turbofan engines with electric motors, following the path of electric or at least hybrid vehicles in the automotive industry (“FlightPath 2050: European Aviation Vision,” [online], available at: https: / / ec.europa.eu / transport / sites / transport / files / modes / air / doc / flightpath2050.pdf [accessed 3 April 2018]).

[0013] As disclosed in EP 3 421 760 A1, GE International is researching a corresponding hybrid turbofan drive for twin-engine commercial aircraft. In this design, electric motors are connected to the high-voltage shaft of one turbofan actuator and the low-voltage shaft of a second turbofan actuator. An energy storage device is provided to power the electric motors, allowing them to provide additional drive power to the connected turbofans under certain operating conditions. A similar solution was proposed by SNECMA in publication WO 2009 / 153471 A2.

[0014] However, the power-to-weight ratio of currently available battery technologies capable of providing significant electric drive power within the aforementioned concepts remains problematic. In short, current battery technologies cannot provide sufficiently high energy density; furthermore, the power-to-weight ratio is not high enough. For example, the energy density of combustible fuels such as kerosene is approximately 40 MJ / kg, or about 12,000 Wh / kg. The energy density of lithium-ion batteries is about 60 times lower than that of first-generation electric fans. Therefore, the specific energy of batteries is only about 2% of that of liquid fuels. As a reminder, a 600kg electric fan with a 67kg battery flew at low speeds for about an hour. In contrast, the empty weight of the Bae 146 is approximately 24,000kg. These figures seem to suggest that the battery weight of an electric aircraft is 60 times the actual fuel weight of the aircraft for the same flight. (Batteries vs. Fossil Fuels, https: / / batteryuniversity.com / learn / archive / batteries_against_fossil_fuel (accessed June 17, 2020)).

[0015] To mitigate its climate impact, the industry is exploring other concepts, such as radical new technologies that utilize sustainable aviation fuel (SAF) on a large scale as a temporary balance between bulk or a combination of both. Hydrogen-powered systems are one such technology.

[0016] In the 1980s, as part of the further development of the Tu-154, alternative fuels for jet engines were tested in Tupolev. This resulted in the Tu-155 prototype, powered by either liquid hydrogen or natural gas. In this three-engine machine, the right engine was not driven by kerosene, but by either hydrogen or natural gas. However, the lessons learned indicated that converting large transport aircraft to hydrogen required a redesign of large, heavy LH2 storage tanks. Furthermore, the increased weight of long-haul aircraft would significantly increase energy consumption, thus increasing costs. Summary of the Invention

[0017] Therefore, the object of this invention is to provide a hybrid power drive system that can further optimize the use of internal combustion engines, especially turbines, and reduce emissions. Furthermore, a multi-engine aircraft and its operating method will be specified, in which the hybrid power drive system optimizes typical flight operation phases.

[0018] According to the present invention, the purpose of the relevant drive system is achieved by a drive system.

[0019] Advantageous embodiments of the present invention for hybrid drive systems are derived from the description and drawings.

[0020] The propulsion system for multi-engine aircraft according to the present invention is based on a specific propulsion system architecture using two different energy sources or fuels, which enables the aircraft to operate with zero emissions (zero-emission cruise) during the basic phase of a typical mission. Due to the specific system architecture and the combination of two power sources, the impact of new technologies and power sources on the configuration of today's in-service aircraft is minimized, and deployment can be carried out much earlier than current zero-emission system proposals.

[0021] An important aspect of this invention is designing such a system based on the desired mission and the overall structure of the aircraft. Without considering the entire aircraft system, it is impossible to determine the dimensions and representation of the required efficiency.

[0022] A key component of this invention is an aircraft architecture capable of achieving "zero-emission" cruise flight, thus eliminating aviation-specific non-CO2 emissions in the upper atmosphere. This invention describes a solution for how the high thrust demands during startup can be distributed between two energy sources, allowing the weight and volume of an emission-free drive system to be integrated into existing technologies already used in current aircraft platforms. This invention also discloses an integrated aircraft system solution comprising a hybrid power drive system, a fuel cell system, a hydrogen tank system, and a control and regulation unit.

[0023] The overall system under discussion consists of a drive system with two different energy sources. The drive system comprises two electric-hybrid internal combustion engines, turbines, or other thermal combustion engines, each coupled with a propeller and a special regulation and control unit for the heat engine and motor. On one hand, fuel is used as the energy source, compatible with standard refueling, standard aviation fuel, or decarbonized, synthetic, or biofuels (SAF), as well as hydrogen for supplying the fuel cell system. Essentially, the advantage of this invention is the ability to achieve long-distance flights without emissions. It is important to consider both CO2 and non-CO2 emissions simultaneously. In particular, the climate-damaging greenhouse effect caused by combustion at high altitudes is taken into account.

[0024] To achieve the goal of zero-emission cruise using this system, while also enabling early application possibilities, the system's size design and the use of both energy sources during missions are crucial to ensuring compatibility with current aircraft concepts. At least two important factors are at play:

[0025] - The amount of hydrogen and the resulting tank volume have no effect on the size of the aerodynamic surface area (no additional drag, therefore no performance loss).

[0026] - The performance of fuel cell systems is primarily suited for cruise, so limited use during cruise and descent is an optimized use without the need for additional heavy-duty electrical energy sources such as batteries or supercapacitors (any increase in the power of a fuel cell system has an exponential effect on system weight and cooling requirements).

[0027] In summary, based on today's assessable prospects, the drive system is poised to incorporate new technologies into aviation within the next decade:

[0028] 1. Applicable to today's aircraft concepts (wing-fuselage, with significant but today assessable changes). For example, further utilizing the wing as a fuel tank for the internal combustion engine is essential.

[0029] 2. It is largely convertible, therefore applicable to aircraft-level approval rules. Verification of new technologies and energy sources, but meeting basic requirements.

[0030] 3. Even without hydrogen and fuel cell systems, aircraft with a given system design can be used, for example in areas without hydrogen infrastructure.

[0031] 4. Since two energy sources can be used, the impact on the system is limited.

[0032] 5. Use fuel cell systems as a continuous energy source.

[0033] 6. Improve operational safety in the event of a safety-related failure in a heat engine or energy source.

[0034] 7. Fuel in internal combustion engines is essentially not burned at cruising altitude, or only within a short time window, thus avoiding the climate-damaging greenhouse effect.

[0035] Regarding the primary objective, namely emission reduction, the following CO2 and non-CO2 emissions can be reduced through this system architecture and specific energy usage during the mission, using a specific example of the Do328 turboprop engine. For reference, the system is designed based on a one-hour mission with the following energy usage:

[0036] a) A sustainable, CO2-neutral fuel (SAF) thermal internal combustion engine primarily used for takeoff, climb, approach, and landing.

[0037] b) Hydrogen / fuel cell system for cruise and descent.

[0038] The energy requirements of this reference mission result in approximately 60% synthetic fuel (a) and approximately 40% hydrogen (b). In short, this is based on the fact that the overall efficiency of the heat engine is roughly the same as that of the fuel cell system, including cooling and the necessary electronics. This assumption can, of course, vary depending on integration factors and the respective levels of current technology, but it is in the range of approximately 5% to 10%. This will affect the detailed classification of carbon-neutral and zero-emission flights. From today's perspective, the technological outlook sees greater potential in improving the efficiency of fuel cell systems compared to the potential for improvement in turboprop drives.

[0039] Emissions reduction summary:

[0040] - Based on a turboprop aircraft carrying approximately 40 to 90 passengers, the use of hydrogen and fuel cells can generate about 40% of the energy in a one-hour mission with zero emissions. This means that no carbon dioxide is emitted during cruise flight, and there are no climate-damaging exhaust effects or condensation contamination at cruise altitude (FL 250), which account for a large proportion of aviation emissions.

[0041] Furthermore, when switching from conventional fuels to synthetic fuels, approximately 60% of the combustion share can produce CO2 neutrality. Therefore, this aircraft and system architecture enables 100% decarbonized and CO2-neutral flight, as well as approximately 40% emission-free flight at cruise altitude in terms of CO2, greenhouse gases, and contrail effects.

[0042] This architecture can be implemented using two different variations:

[0043] a) The heat engine is designed for takeoff, climb, and landing, while the fuel cell system is dedicated to cruise and descent. The required hydrogen tank volume is therefore minimized to optimize integration into the aircraft. The aircraft retains the fundamental ability to operate without the hydrogen and fuel cell systems.

[0044] b) By expanding the use of fuel cell systems to support the internal combustion engine during takeoff, climb, and landing, the power requirements of the heat engine can be reduced (miniaturization). This means that smaller internal combustion engines can be used. By reducing the mission block energy share from fuel combustion, the share of emissions-free flight can be further increased to approximately 70%. It is necessary to adapt the hydrogen storage tank volume to the increased energy demands.

[0045] In general, this invention relates to a drive system for small and large transport aircraft (CS-23 and CS-25) featuring a dual-engine drive (piston or turbine) that converts thermal energy into mechanical drive shaft power to drive propellers (propellers, impellers, rotors) to generate thrust. Converting to primarily propeller-driven regional aircraft, a size class of approximately 30 to 90 passengers appears to be the most economical. This corresponds to approximately 3,000 kW to 8,000 kW of propeller shaft power (the sum of all propeller shafts – total power).

[0046] The feasibility of this architecture takes into account the following factors, particularly its potential for integration into current aircraft architectures:

[0047] - Energy requirements at different stages of flight

[0048] - Power requirements at different flight phases

[0049] - Optimize the distribution between thermal combustion / fuel utilization and electricity / hydrogen utilization

[0050] - Hydrogen weight and canister volume

[0051] - Overall system and component weight

[0052] - Thermoelectric efficiency fuel cell system

[0053] In order to consider the feasibility of the invention in certain aircraft categories, it is based on a technological outlook regarding performance and maturity, which is achievable from today's perspective.

[0054] Of course, as technological prospects and the timing of introduction change, the applicability of this system architecture extends to higher power levels and aircraft sizes.

[0055] The following technical values ​​are for reference:

[0056] - Overall efficiency of the fuel cell system: 50%

[0057] - Total weight of fuel cell system: 1500 kg (including electric drive)

[0058] For liquid hydrogen storage tanks, a weight index (GI) of 20% or higher is desirable. The tank's GI is calculated by dividing the mass of stored hydrogen by the sum of the mass of stored hydrogen and the weight of the empty tank. A GI of 50% means that the empty tank weighs the same as the stored hydrogen.

[0059] Essentially, compared to known existing drive systems, this invention can achieve at least four significant advantages:

[0060] a) Fuel consumption is reduced by optimizing and adjusting the use of the internal combustion engine at different stages of flight. Hybrid power generation of the drive unit is crucial here, including a variable-switching electric motor-generator unit on a common transmission. This allows for a gradual transition of power from the electric motor to the engine during the transition from climb to cruise flight.

[0061] (b) The transmission plays a further role in the drive system according to the invention, its input shaft being capable of achieving an optimized speed range for the heat engine and motor in order to optimally transmit torque to the propeller shaft. This achieves weight-optimized design and loss-optimized operating conditions for the motor and heat engine.

[0062] c) Another important aspect of the invention is reducing the operating time of the heat engine during flight to lower maintenance costs and extend maintenance intervals. Typically, for the same rated power, a motor requires less maintenance work and costs because an electric generator unit generally does not have "hot" components.

[0063] d) Finally, the drive system can improve safety in the event of a "single engine failure". Especially during critical phases of flight, takeoff, initial climb, and landing (takeoff, initial / climb, and approach), the lost power can be immediately and symmetrically distributed to both sides via the motors.

[0064] Compared to the high-capacity, battery-based hybrid concepts currently under development, the result is a significant optimization of the weight design, as current battery concepts still only have relatively low specific energy density.

[0065] The required redundancy is generated by two drive units, including an electric generator unit. Both drive units have identical performance and zoning, providing the same thrust across the entire mission distribution. The primary dimensional error condition: complete drive failure is considered in the design of each drive unit and allows for maneuvering the aircraft at specified limits at each phase of flight until a safe landing.

[0066] Thrust adjustments during a flight mission are made to the same degree by both drives, with the option of additional active blade adjustments. Turbines are typically optimized for the main flight phase, primarily cruise, as this is the largest share of mission energy consumption. It is impossible to achieve equally high operating efficiency in all flight phases. While maximum efficiency can be designed and achieved for cruise flight, which is clearly dominant, for medium- and long-haul routes, operating conditions are far less frequent and more varied for short- and ultra-short-haul routes. Therefore, regional and short-haul aircraft have far fewer instances of engines operating at optimal efficiency throughout a mission compared to short- and long-haul flights, resulting in lower fuel consumption per passenger.

[0067] The varying degrees of decline in propulsive and thermal efficiency also play a role here, because, for example, propeller-driven aircraft fly at lower altitudes and experience less thrust loss during cruise compared to long-haul aircraft equipped with turbofan engines. Therefore, long-haul engines operate more consistently at higher power and with greater efficiency compared to climb and cruise, while regional aircraft have a wider power range, for example, during takeoff, climb, and cruise.

[0068] Therefore, the thrust requirements of missions that can be designed and operated more efficiently by combining heat engines with electric motors are very different, especially for regional aircraft with propellers, but also applicable to other twin-engine aircraft with paddle wheels or rotors.

[0069] While previous aircraft hybrid propulsion systems and architectures aimed to integrate additional or differently arranged propulsion elements (propellers, rotors, rotors) or additional alternative energy sources such as batteries or fuel cells, this invention achieves higher efficiency without requiring additional propulsion units or energy sources. Compared to previous hybrid concepts, such as the arrangement of multiple propulsion units distributed across the wing span to generate better lift at low speeds—a significant advantage in the short flight phases of takeoff and landing—or the use of energy systems with insufficient power density to meet the demands of today's large aircraft, the propulsion system of this invention can be implemented with significant advantages by leveraging current technologies and aircraft concepts.

[0070] Assuming uniform alternation of operation during cruise and descent, the uptime of a single hot engine can be reduced by approximately 30% (for a 60-minute mission reference), which translates one-to-one to extended maintenance intervals and reduced hot engine maintenance costs.

[0071] The drive architecture described in this invention can also be sized and integrated as a modification of existing aircraft. Attached Figure Description

[0072] Other features, advantages, and effects of the present invention arise from the following description of preferred exemplary embodiments of the invention, as shown in the figures. In the figures:

[0073] Figure 1a shows a top view of a twin-engine aircraft with a schematic diagram of a hybrid power drive system;

[0074] Figure 1b shows a side view of a twin-engine aircraft with a hybrid power drive system according to a schematic diagram of Figure 1a;

[0075] Figure 2 A system sketch of a hybrid drive system with a schematic diagram of the system architecture is shown.

[0076] Figure 3a shows the power requirements and flight altitude of an aircraft with a conventional gas turbine and electric motor transmission unit during the operational phase of a typical 200 NM mission.

[0077] Figure 3b shows a graph of the cumulative energy demand of an aircraft with a conventional gas turbine and electric motor transmission unit during the operational phase of a typical 200 NM mission.

[0078] Figure 4a shows the power requirements and flight altitude of an aircraft with a smaller gas turbine and electric motor transmission unit during the operational phase of a typical 200 NM mission.

[0079] Figure 4b shows a graph of the cumulative energy demand of an aircraft with a smaller gas turbine and electric motor transmission unit during the operational phase of a typical 200 NM mission.

[0080] Figure 5 A system sketch of a hybrid drive system in its main operating modes is shown, along with a schematic diagram of the system architecture; and

[0081] Figure 6 A system sketch of a hybrid drive system in the third operating mode is shown, along with a schematic diagram of the system architecture. Detailed Implementation

[0082] The typical installation configuration of the hybrid power drive system 10 for a twin-engine regional aircraft 20 can be seen in the examples of the Dornier 328-100 in Figures 1a and 1b. The aircraft 20 is designed as a conventional upper wing aircraft with a T-tail 21 at the rear. The partially cylindrical fuselage 22 houses a pressurized compartment, in which the cockpit 23 and passenger cabin 24 are housed. At the rear, the pressurized compartment is pressure-sealed closed by a pressure cover. Further aft, the fuselage 22 tapers and supports the T-tail 21. In the Dornier 328-100 known in the prior art, one of the baggage racks is located in the tapered transition area.

[0083] In a conventional high-wing configuration, wing 26 is tangentially attached to the fuselage tube above the fuselage. Hybrid electric drive units 31 and 32 are housed in thruster nacelles 33 and 34, one of which is attached to the left and right wings 26, respectively. Multi-bladed adjustable propellers 61 and 62 are driven by reduction gearboxes also integrated into thruster nacelles 33 and 34. To prevent unwanted icing on the propeller blades, they can be electrically heated, with the blades receiving power for heating from a transmission device 80.

[0084] The system architecture of the drive system 10 integrated in the twin-engine aircraft 20 shown in Figure 1a can be... Figure 2As seen in more detail below. The drive system 10 includes two hybrid electric drive units 31 and 32 that can operate independently of each other. Each hybrid electric drive unit 31, 32 has a gas turbine 41, 42 with a flange-mounted reduction gearbox 51, 52, and in each case, a propeller 61, 62 with variable blade adjustment is connected to it. The corresponding gas turbine 41, 42 with integrated reduction gearbox 51, 52 can be obtained, for example, from Pratt & Whitney Canada, designated PW 119C. Integrated fuel tanks 43, 44 are formed in the wing 26 for the left and right wings, which supply fuel to the two gas turbines 41, 42 via fuel lines and systems not shown.

[0085] Each drive unit 31, 32 is assigned an electric generator unit 71, 72; these are respectively connected to reduction gearboxes 51, 52 on the drive side. Depending on the operating phase, the electric generator units 71, 72 can operate as either electric motors or generators. In drive mode, the electric generator units 71, 72 transmit drive power to the corresponding propellers 61, 62 via the reduction gearboxes 51, 52. In generator mode, the electric generator units 71, 72 generate electricity, which is fed to the transmission device 80 for further distribution or storage. The functional components of the transmission device 80 are two power converters 81 and 82, one of which is assigned to each of the electric generator units 71, 72.

[0086] To supply electrical energy to the electric generator units 71 and 72, the drive system 10 includes a fuel cell 73, which in turn supplies hydrogen via a fuel tank 74. In the fuel cell 73, the hydrogen is converted into electrical energy, which is then transmitted via a transmission device 80 to power converters 81 and 82 and the electric generator units 71 and 72 to drive propellers 61 and 62. Currently, the most advanced and suitable fuel cell for aviation is the cryogenic proton exchange membrane fuel cell (PEM fuel cell). Adding optional energy storage devices such as batteries to the system helps ensure rapid load point tracking and power peaking to optimize the size of the fuel cell system.

[0087] Hydrogen can typically be stored in compressed gas or liquid form. While gaseous storage may be suitable for shorter flights and is commercially available, this invention focuses on liquid hydrogen (LH2) storage tanks because they require approximately half the volume and are therefore much lighter than gaseous hydrogen tanks. Since LH2 must be kept cryogenic and heat transfer must be minimized to prevent hydrogen evaporation, spherical or cylindrical tanks are required to maintain low losses. In the configurations shown in Figures 1a and 1b, a spherical fuel tank 74 is housed in a conical aft fuselage 27, which can be used as a cargo hold when the fuel tank 74 is removed.

[0088] The unit, consisting of controller 90, transmission device 80, and fuel cell 73, is located in the nose side region of fuselage 22, forward of the wing and outside the pressurized cabin. The unit and the fuel tank 74 for supplying fuel cell 73 form a substantially neutral torque balance relative to the center of gravity SP of aircraft 20 (see Figure 1b).

[0089] In a secondary function, waste heat removed from fuel cell 73 by the cooling unit is used to de-ice exposed surfaces of aircraft 20, such as the wing leading edge, the air intakes of gas turbines 41 and 42, and the leading edge of the T-tail.

[0090] A central controller 90 is provided to control the drive power generated by heat and electricity. It is connected to power converters 81 and 82 and gas turbines 41 and 42. On one hand, the controller 90 controls the output of the electric generator units 71 and 72, the electric drive power, and the electrical energy to be generated via the power converters 81 and 82 according to the operating phase. On the other hand, it controls the drive power generated by heat from the gas turbines 41 and 42. Typical parameters controlled and monitored by the controller 90 include fuel supply, power output, high-pressure shaft speed, and turbine temperature of the gas turbines 41 and 42.

[0091] In another embodiment, Figure 2 An architecture based on a DC voltage network 101 and AC-DC converters 81, 82 is shown. Depending on the operating mode and power requirements, the power output of the fuel cell 73 can be fed to the first and second electric generator units 71 or 72 via a transmission device 80 and DC-AC converters 81, 82, and the drive power can be transmitted to the propellers 61 or 62 via a reduction gearbox 51 or 52.

[0092] The graph in Figure 3a illustrates the differences between an aircraft 20 with the propulsion system 10 according to the invention and, for example, a Dornier 328 equipped with two conventional thrusters, using cruise flight as an example. The maximum power of the heat engine is designed for takeoff and landing, while the use of the fuel cell system is dedicated to cruise and descent. The required hydrogen tank volume is thus minimized to optimize the possibility of integration into the aircraft. The graph in Figure 3a shows the power requirements (in kilowatts) and flight altitude (in feet) over time for a typical 200 NM flight mission. The main phases of a typical flight mission are explained below:

[0093] take off:

[0094] Takeoff is the flight phase in which the aircraft transitions from ground movement (taxiing) to flight, typically starting from the runway. Usually, the thrusters operate at full power during takeoff.

[0095] Climb:

[0096] After takeoff, the aircraft climbs to a certain altitude (25,000 feet in this case) and then flies safely and economically to its destination.

[0097] Cruise flight:

[0098] Cruise is a part of air travel during which flight is the most fuel-efficient phase. It occurs between the ascent and descent phases and typically represents the largest part of the journey. Technically, cruise is conducted at a constant airspeed and altitude. Cruise ends as the aircraft approaches its destination, and the descent phase, in preparation for landing, begins. In most commercial airliners, the cruise phase consumes the majority of the fuel.

[0099] decline:

[0100] The descent during flight is the portion of the aircraft that loses altitude. Descent is an important part of the landing approach. Other parts of the descent may be to avoid traffic, avoid adverse flight conditions (turbulence or severe weather), avoid cloud cover (especially under visual flight rules), enter warmer air (if there is a risk of icing), or take advantage of wind direction at different altitudes. A normal descent occurs at a constant airspeed and a constant angle of descent. The pilot controls the angle of descent by changing engine power and angle of attack (nose lowering) to maintain airspeed within a specified range. During the initial and descent phases, the thrusters will operate at low power.

[0101] Approach and landing:

[0102] When an aircraft returns to the ground, the approach and landing are the final parts of the flight. During landing, airspeed and descent speed are reduced to maintain the designated skid path to the runway landing point (3 degrees for most airports). This speed reduction is achieved by decreasing thrust and / or using flaps, landing gear, or air brakes to generate greater drag. As the aircraft approaches the ground, the pilot performs a pull-up maneuver to initiate a soft landing. The landing and approach process is primarily performed using the Instrument Landing System (ILS).

[0103] Line A (dashed line): Power requirements of propulsion system 10 during missions using two hybrid electric drive units 31, 32. Power requirements are highest during takeoff and decrease during subsequent flight phases. Power requirements for takeoff, climb, and approach and landing during the flight phases are provided solely by gas turbines 41, 42 (“Primary Operating Mode”). During cruise and descent, electric generator units 71, 72 are driven by fuel cell 73 (“Third Operating Mode”). During flight, controller 90 controls the operation of propulsion units 31, 32 and the transition from the primary operating mode to the secondary operating mode, and vice versa. The dashed area under line A represents fuel consumption of gas turbines 41, 42, and the square area represents hydrogen consumption of electric generator units 71, 72 during the operating phase. The energy requirements of this reference mission result in approximately 60% fuel (SAF) and approximately 40% hydrogen.

[0104] Line B (solid line) describes the flight altitude (in feet) during the mission. The highest flight altitude is during cruise flight, at approximately 25,000 feet.

[0105] In the chart according to Figure 3b, the cumulative energy demand associated with the above-mentioned operating phases and operating modes is shown by solid and dashed lines C and D. Line C represents the energy demand during the primary operating mode, and line D represents the energy demand during the third operating mode.

[0106] The graphs in Figures 4a and 4b represent values ​​for a second variation of the invention. Here, by expanding the use of the fuel cell system to support the internal combustion engine during takeoff and landing, a reduction in the power demand of the heat engine (miniaturization) is achieved. In this case, there is a combined operation and power output of the gas turbines 41 and 42 and the electric generator units 71, 72 coordinated by controller 90 (“second operating mode”). One advantage is that a smaller internal combustion engine can be used. This results in a further increase in the emission-free ratio of flight to approximately 70% (the grid area in Figure 4a).

[0107] Line E (dashed line): According to Figure 4a, the power requirements of the drive system 10 in missions using the two hybrid electric drive units 31, 32 are, in principle, the same as those shown in Figure 3a. Similarly, the power requirements are highest during takeoff and decrease in subsequent flight phases.

[0108] Line F (solid line) describes the flight altitude (in feet) during the mission. As shown in Figure 3a, the flight altitude is highest during cruise flight and is approximately 25,000 feet.

[0109] The diagrams in Figures 3a and 3b clearly illustrate an important aspect of the invention: for many aircraft, particularly propeller-driven regional aircraft, the thrust required for takeoff is significantly higher than that for cruise flight, and therefore a large portion of the flight missions of a hot engine can only operate at about half its carrying capacity. This means that conventional turboprop engines operate outside their optimal operating point, closer to their maximum power output point. In contrast, hybrid-electric drive systems can be optimized for two different operating modes.

[0110] Regarding the operation phase of drive system 10, the following basic operating states are generated:

[0111] 1. Figure 5 A system schematic diagram of takeoff, climb, and approach (takeoff, climb, and approach / landing) is shown. Both gas turbines 41 and 42 are operational and drive propellers 61 and 62 (solid lines) via intermediate reduction gearboxes 51 and 52. Thrust regulation is centrally executed on gas turbines 41 and 42 via a hybrid propulsion controller 90 (primary operating mode). Electric generator units 71 and 72 do not generate drive power (dashed lines).

[0112] 2. Secondary operating modes for takeoff and climb, such as Figure 2 As shown. In this combined operating mode, the propellers (61, 62) obtain drive power from the first and second internal combustion engines (41, 42) and the first and second electric generator units (71, 72). The relevant power demand and cumulative energy consumption can be seen in the graphs of Figures 4a and 4b. The reduction in fuel consumption of the gas turbine becomes apparent here due to the additional drive power from the electric generator units 71, 72.

[0113] 3. System status for cruise and descent can be found in Figure 6 As seen in the figures, power / torque demand decreases to cruise levels (Figures 3a and 4a), where the power from the two gas turbines 41 and 42 is shut off, while drive power from the two electric generator units 71 and 72 is evenly distributed to the two propellers 61 and 62 via connected transmissions 51 and 52. During cruise and subsequent descent, thrust demand is adjusted by the hybrid propulsion controller 90. This is responsible for thermal and electrical regulation (third operating mode).

[0114] 4. Even in the event of a severe fault or internal combustion engine failure, the system architecture according to the invention can achieve symmetrical thrust due to power distribution on the power network. On the side of the faulty internal combustion engine 41, 42, the electric generator unit 71 or 72 can be switched on, thus at least partially compensating for the failed thrust.

[0115] List of reference numerals

[0116] 10 Drive System

[0117] 20 aircraft

[0118] 21 T-shaped tail fin

[0119] 22 fuselage

[0120] 23 Cockpit

[0121] 24 cabins

[0122] 26. Wings, left and right

[0123] 27 Rear fuselage

[0124] Drive units 31 and 32, left and right

[0125] 33, 34 Thruster compartments, left and right

[0126] 41 and 42 Gas turbines, left and right

[0127] 43, 44 Integrated fuel tanks, left and right

[0128] 51, 52 Reduction gearboxes, left and right

[0129] Propellers 61 and 62, left and right

[0130] 71 and 72 are electric generator units, on the left and right.

[0131] 73 Fuel Cells

[0132] 74 Fuel Tank

[0133] 80 Transmission device

[0134] 81, 82 Power Converters

[0135] 90 controller

[0136] Lines A, B, C, D, E, and F

[0137] SP Center of Gravity

Claims

1. A multi-engine aircraft having a hybrid power drive system (10), the multi-engine aircraft comprising: • At least the first and second hybrid electric drive units (31, 32) each have an internal combustion engine (41, 42) and an electric generator unit (71, 72) for transmitting drive power to the thrusters (61, 62). • A wing (26) housing the drive units (31, 32) and fuel tanks (43, 44), and a fuselage (22). The drive units (31, 32) consist of turboprop propellers each equipped with a gas turbine, which can be coupled to a reduction gearbox (51, 52) to drive the propellers (61, 62). The electric generator units (71, 72) can be coupled to the gearbox (51, 52) in a controlled manner via a controller (90) according to the operating mode. The thrusters (61, 62) are coupled to the internal combustion engines (41, 42) and / or the electric generator units (71, 72) to transmit drive power, and the first and second electric generator units (71, 72) are connected to the transmission device (80) for power distribution. • Fuel cell (73), said fuel cell being used to provide electrical energy to the first and / or second electric generator units (71, 72), • A controller (90), connected to the internal combustion engine (41, 42) and / or the transmission device (80) and / or the electric generator unit (71, 72) and / or the fuel cell (73), the controller (90) being used to control the drive power generated by heat and electricity. • Separate fuel tanks (43, 44) for supplying fuel to the internal combustion engine (41, 42) or cryogenic hydrogen to the fuel cell (73), Its features are, At least the main volume of the fuel tanks (43, 44) for supplying the internal combustion engines (41, 42) is integrated in the wing (26), and at least the main volume of the fuel tank (74) for supplying the fuel cell (73) is integrated in the rear region of the fuselage (22).

2. The multi-engine aircraft according to claim 1, characterized in that, The hybrid electric drive units (31, 32) • In the primary operating mode, the thrusters (61, 62) obtain drive power mainly or entirely from the internal combustion engines (41, 42). • In the secondary combined operation mode, the thrusters (61, 62) obtain drive power from the first and second internal combustion engines (41, 42) and the first and second electric generator units (71, 72), and • In the third operating mode, the thrusters (61, 62) obtain driving power from the first and second electric generator units (71, 72).

3. The multi-engine aircraft according to claim 1 or 2, characterized in that, In operation mode, the controller (90) symmetrically distributes the drive power to the thrusters (61, 62).

4. The multi-engine aircraft according to claim 1 or 2, characterized in that, The electric drive power of the first or second electric generator unit (71, 72) can be variably switched on during the transition between operating modes.

5. The multi-engine aircraft according to claim 1 or 2, characterized in that, Each of the hybrid electric drive units (31, 32) has a transmission (51, 52) for transmitting drive power, wherein the internal combustion engine (41, 42) and the electric generator unit (71, 72) can be connected to the propeller (61, 62) via the transmission (51, 52).

6. The multi-engine aircraft according to claim 1 or 2, characterized in that, The transmission of drive power from the internal combustion engines (41, 42) and the electric generator units (71, 72) changes successively to keep the drive power transmitted by the propellers (61, 62) and the common hybrid electric drive units (31, 32) the same.

7. The multi-engine aircraft according to claim 1 or 2, characterized in that, The propellers (61, 62) are designed as propellers with blade adjusters, and a controller (90) for controlling the drive power is connected to the blade adjusters.

8. The multi-engine aircraft according to claim 1 or 2, characterized in that, In another operating mode, the driving power of the first or second internal combustion engine (41, 42) is completely or primarily lost and the first or second electric generator unit (71, 72) is powered by the fuel cell (73) via the transmission device (80).

9. The multi-engine aircraft according to claim 1 or 2, characterized in that, The transmission device (80) is designed as an alternating current network.

10. The multi-engine aircraft according to claim 1 or 2, characterized in that, The transmission device (80) is designed as a DC network, wherein each electric generator unit (71, 72) is assigned an AC / DC voltage converter (81, 82), which is connected to a controller (90) for controlling the rotational speed of the thrusters (61, 62).

11. The multi-engine aircraft according to claim 1 or 2, characterized in that, The internal combustion engines (41, 42) operate using sustainable aviation fuel.

12. The multi-engine aircraft according to claim 1, characterized in that, The fuselage (22) has a space in the rear region for forming a cargo hold, in which a fuel tank (74) for supplying fuel cells (73) is provided.

13. The multi-engine aircraft according to claim 1 or 2, characterized in that, The unit, consisting of a controller (90) and / or a transmission device (80) and / or a fuel cell (73), is arranged in the nose side region of the fuselage (22).

14. The multi-engine aircraft according to claim 13, characterized in that, The unit and the fuel tank (74) for supplying the fuel cell (73) form a torque balance that is essentially neutral with respect to the center of gravity (SP) of the aircraft (20).

15. The multi-engine aircraft according to claim 1 or 2, characterized in that, The fuel cell (73) has a cooling unit in which waste heat is used to de-ice the exposed surfaces of the aircraft.

16. A method for operating a multi-engine aircraft (20) according to any one of claims 1 to 15, characterized in that, The drive system (10) operates in primary, secondary and third operating modes, wherein • The aircraft (20) taxis on the apron and taxiway in primary or third operating mode. • Take off and climb to cruise altitude in primary or secondary operating mode. • Cruise and descend to approach altitude in secondary or third operating modes. • Arrival and landing in primary or secondary operating modes, and • If the internal combustion engine (41, 42) fails, the flight will continue in the secondary or third operating mode.

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