AIRCRAFT HAVING A DRIVE-AND-ENERGY SYSTEM FOR LOW-EMISSION CRUISING FLIGHT

Abstract

The invention relates to a hybrid electric drive system (10) for multi-motor aircraft (20). The hybrid electric drive system comprises at least a first and a second hybrid electric drive unit (31, 32), each of which comprises: an internal combustion engine (41, 42), a motor-generator unit (71, 72) and a gear box (51, 52) for transmitting drive power to a propeller (61, 62). In order to supply the motor-generator units (71, 72) with electrical energy, the drive system (10) has a fuel cell (73), which in turn is supplied with hydrogen by means of 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 by means of the transmission device (80) and power converters (81) and (82), in order to drive the propellers (61, 62). Advantages: On the basis of a turboprop aircraft (20) with approximately 40 to 90 passengers, approximately 40% of the energy during a 1-hour mission can be provided emission-free by means of hydrogen and fuel cell. This means no CO2 emissions at all during the cruising flight and also no climate-damaging exhaust-gas and contrail effects at cruising altitude (FL250), which are a significant share of aviation emissions.

Claims

1. A hybrid propulsion system for multi-engine aircraft having: at least one first and one second hybrid-electric propulsion unit, each having an internal combustion engine and a motor-generator unit for transmitting propulsion power to a propulsor, wherein the propulsor can be coupled to the internal combustion engine and/or the motor-generator unit for the transmission of propulsion power, the first and second motor-generator units are connected to a transmission device for distributing electric power, a fuel cell for supplying the first and/or second motor-generator unit with electrical energy, a controller for controlling the thermally and electrically generated propulsion power is connected to the internal combustion engines and/or the transmission device and/or motor-generator units and/or the fuel cell, separate fuel tanks for supplying the internal combustion engines with fuel or the fuel cell with cryogenic hydrogen.

2. The propulsion system according to claim 1, characterized in that in the hybrid-electric propulsion unit: in a primary operating mode, the propulsors receive the propulsion power predominantly or entirely from the internal combustion engines, in a secondary, combined operating mode, the propulsors receive the propulsion power from the first and second internal combustion engines and from the first and second motor-generator units, and in a third operating mode the propulsors receive the propulsion power from the first and second motor-generator units.

3. The propulsion system according to claim 1, characterized in that, in the operating modes, the controller brings about symmetrical distribution of the propulsion power to the propulsors.

4. The propulsion system according to claim 1, characterized in that the electrical propulsion power of the first or second motor-generator unit can be variably switched on on transition between the operating modes.

5. The propulsion system according to claim 1, characterized in that the hybrid-electric propulsion units each have a gearbox for transmitting the propulsion power, wherein the internal combustion engine and the motor-generator unit can be coupled to the propulsor by means of the gearbox.

6. The propulsion system according to claim 1, characterized in that the change in the transmission of the propulsion power of the internal combustion engine and the propulsion power of the motor-generator unit takes place successively in such a way that the propulsion power output to the propulsor of the common propulsion unit remains approximately the same.

7. The propulsion system according to claim 1, characterized in that the propulsors are designed as propellers with blade adjuster and the controller for controlling the propulsion power is connected to the blade adjuster.

8. The propulsion system according to claim 1, characterized in that in a further operating mode the propulsion power of the first or second internal combustion engine has failed completely or predominantly and the first or second motor-generator unit is provided with electrical power by the fuel cell via the transmission device.

9. The propulsion system according to claim 1, characterized in that the transmission device takes the form of an AC network.

10. The propulsion system according to claim 1, characterized in that the transmission device takes the form of a DC network, each motor-generator unit being assigned an AC/DC converter which is connected to the controller to control the speed of the propulsor.

11. The propulsion system according to claim 1, characterized in that the internal combustion engines are operated with sustainable aviation fuel.

12. A multi-engine aircraft having a hybrid propulsion system according to claim 1, a wing accommodating the propulsion units and a fuel tank, and a fuselage, characterized in that the propulsion unit is formed of a turboprop engine with in each case one gas turbine which can be coupled to a speed-reducing gearbox to drive a propeller, wherein the motor-generator unit can be coupled to the gearbox in a controlled manner via the controller, depending on operating mode.

13. The multi-engine aircraft according to claim 12, characterized in that at least the predominant volume of the fuel tanks for supplying the internal combustion engines is integrated in the wing and at least the predominant volume of the fuel tanks for supplying the fuel cell is integrated in a rear area of the fuselage.

14. The multi-engine aircraft according to claim 12, characterized in that the fuselage has a space in a rear area for forming a cargo hold, in which the fuel tank for supplying the fuel cell is arranged.

15. The multi-engine aircraft according to claim 12, characterized in that units consisting of a controller and/or transmission device and/or fuel cell are arranged in a bow-side area of the fuselage.

16. The multi-engine aircraft according to claim 15, characterized in that the units and the fuel tank for supplying the fuel cell form a moment equilibrium which is essentially neutral with respect to the center of gravity (SP) of the aircraft.

17. The multi-engine aircraft according to claim 12, characterized in that the fuel cell has a cooling unit, the waste heat being used to de-ice exposed surfaces of the aircraft.

18. A method for operating a multi-engine aircraft according to claim 12, characterized in that the propulsion system is operated in a primary, secondary and in a third operating mode, wherein: taxiing of the aircraft, in particular on aprons and taxiways, takes place in the primary or third operating mode, take-off and climb to cruising altitude take place in the primary or secondary operating mode, cruising and descent to approach altitude take place in the secondary or third operating mode, approach and landing take place in the primary or secondary mode of operation, and if an internal combustion engine fails, the flight continues in the secondary or third operating mode.

Description

[0073] Further features, advantages and effects of the invention result from the following description of preferred exemplary embodiments of the invention, as shown in the drawings. In the figures:

[0074] FIG. 1a shows a plan view of a twin-engine aircraft with a schematic representation of a hybrid propulsion system,

[0075] FIG. 1b shows a side view of the twin-engine aircraft according to FIG. 1a with a schematic representation of a hybrid propulsion system,

[0076] FIG. 2 shows a system diagram of a hybrid propulsion system with a schematic representation of the system architecture,

[0077] FIG. 3a shows a diagram of the power requirement and flight altitude during the operating phases of a typical 200 NM mission by an aircraft with conventional gas turbines and electric motor-gearbox units,

[0078] FIG. 3b shows a diagram of the accumulated energy requirement during the operating phases of a typical 200 NM mission by an aircraft with conventional gas turbines and electric motor-gearbox units,

[0079] FIG. 4a shows a diagram of the power requirement and flight altitude during the operating phases of a typical 200 NM mission by an aircraft with smaller gas turbines and electric motor-gearbox units,

[0080] FIG. 4b shows a diagram of the accumulated energy requirement during the operating phases of a typical 200 NM mission by an aircraft with smaller gas turbines and electric motor-gearbox units,

[0081] FIG. 5 shows a system diagram of a hybrid propulsion system with a schematic representation of the system architecture in the primary operating mode, and

[0082] FIG. 6 shows a system diagram of a hybrid propulsion system with a schematic representation of the system architecture in the third operating mode.

[0083] A typical installation configuration of a hybrid propulsion system 10 for a twin-engine regional aircraft 20 is shown in FIG. 1a and FIG. 1b, taking the Dornier 328-100 as example.

[0084] The aircraft 20 takes the form of a conventional high-wing aircraft with a T-tail unit 21 at the rear. The fuselage 22, which is cylindrical in sections, is designed with a pressurized cabin in which the cockpit 23 and the passenger compartment 24 are accommodated. At the rear, the pressurized cabin is closed in pressure-tight manner by a pressure bulkhead. Further aft, the fuselage 22 tapers conically and carries the T-tail unit 21. In the Dornier 328-100, known from the prior art, one of the luggage compartments is provided in the conical transition area.

[0085] In a conventional high wing configuration, the wings 26 are attached to the fuselage tube, at an overhead tangent thereto. The hybrid-electric propulsion units 31 and 32 are accommodated in the engine nacelles 33 and 34, of which one is attached to each of the left and right wings 26. The multi-bladed and adjustable propellers 61, 62 are driven via reduction gearboxes likewise integrated in the engine nacelles 33 and 34. To avoid undesirable icing phenomena on the propeller blades, these can be heated electrically, the propeller blades receive the power for heating from a transmission device 80.

[0086] The system architecture of the propulsion system 10 integrated in a two-engine aircraft 20 in FIG. 1a can be seen in further detail in FIG. 2. The propulsion system 10 comprises two hybrid-electric propulsion units 31 and 32 that can be operated independently of one another. Each hybrid-electric propulsion unit 31, 32 has a gas turbine 41, 42 with flange-mounted reduction gearbox 51, 52, via which in each case a propeller 61, 62 with variable pitch adjustment is coupled. Corresponding gas turbines 41, 42 with integrated reduction gearbox 51, 52 are available, for example from Pratt & Whitney Canada under the designation PW 119C. In the wings 26, left and right wing integral tanks 43, 44 are formed, which supply the two gas turbines 41, 42 with fuel via fuel lines and systems not described in any greater detail.

[0087] Assigned to each propulsion unit 31, 32 is a motor-generator unit 71, 72, each of which are coupled to the reduction gearbox 51, 52 on the propulsion side. Depending on the operating phase, the motor-generator unit 71, 72 can be operated as an electric motor or as a generator. In propulsion mode, the motor-generator unit 71, 72 transmits propulsion power via the reduction gearbox 51, 52 to the respective associated propeller 61, 62. In generator mode, the motor-generator unit 71, 72 generates electrical power, which is fed to a transmission device 80 for further distribution or storage. Two power converters 81 and 82, one of which is in each case assigned to each motor-generator unit 71, 72, constitute a functional component of the transmission device 80.

[0088] In order to supply the motor-generator units 71, 72 with electrical energy, the propulsion system 10 has a fuel cell 73, which is in turn supplied with hydrogen via a fuel tank 74. In the fuel cell 73, hydrogen is converted into electricity, electric power then being supplied via the transmission device 80 and power converters 81 and 82 to the motor-generator units 71, 72 to drive the propellers 61, 62. Currently most advanced and best suited to aviation are low-temperature proton exchange membrane fuel cells (PEM fuel cells). The addition of an optional energy storage device such as a battery to this system helps ensure rapid load follow-up and power peak shaving to optimize fuel cell system dimensioning.

[0089] In general, 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 significantly lighter than gaseous hydrogen tanks. Since LH2 must remain cold and heat transfer must be minimized to avoid hydrogen vaporization, spherical or cylindrical tanks are required to keep losses low. In the configuration shown in FIG. 1a and FIG. 1b, the spherical fuel tank 74 is accommodated in the conical rear fuselage 27, which can be used as a cargo hold when the fuel tank 74 is removed.

[0090] The units consisting of controller 90, transmission device 80 and fuel cell 73 are arranged in a bow-side area of the fuselage 22 in front of the wing and outside the pressurized cabin. The units 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 SP of the aircraft 20 (see FIG. 1b).

[0091] In a secondary function, the waste heat removed from the fuel cell 73 by means of a cooling unit serves to de-ice 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.

[0092] A central controller 90, which is connected to power converters 81 and 82 and the gas turbines 41, 42, is provided for controlling the thermally and electrically generated propulsion power. On the one hand, depending on operating phase, the controller 90 controls the motor-generator unit 71, 72 via the power converters 81 and 82, the delivery of electrical propulsion power and the electrical energy to be generated, and on the other hand the thermally generated propulsion power of the gas turbines 41, 42. Typical controller 90 parameters to be controlled and monitored are the fuel supply, the speeds of the power and high-pressure shaft and the turbine temperature of the gas turbines 41, 42.

[0093] In a further embodiment, an architecture is shown in FIG. 2 which is based on a DC voltage 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 fed via the transmission device 80 and the AC/DC converters 81, 82 to the first and second motor-generator units 71 and 72, respectively, and the propulsion power can be transmitted via the reduction gearboxes 51 and 52 to the propellers 61 and 62, respectively.

[0094] Taking a Dornier 328 as example, the diagram in FIG. 3a shows the difference for cruising between an aircraft 20 with a propulsion system 10 according to the invention and the prior art, taking a Dornier 328 equipped with two conventional engines as an example. The maximum power of the thermal engines is designed to carry out take-off and landing, with use of the fuel cell system being intended exclusively for cruising and descent. The required tank volume for hydrogen is thus minimized in order to optimize the possibilities of integration into the aircraft. The diagram in FIG. 3a shows the power requirement in kW and flight altitude in ft over time for a typical 200 NM flight mission. The most important phases of a typical flight mission are explained below:

[0095] Take-off:

[0096] Take-off is the phase of flight in which the aircraft 20 makes the transition from moving along the ground (taxiing) to flying in the air, typically starting on a runway. As a rule, the engines are operated at full power during take-off.

[0097] Climb:

[0098] After take-off, the aircraft climbs to a certain altitude (in this case 25,000 ft) before flying safely and economically to its destination at that altitude.

[0099] Cruising:

[0100] Cruising is the portion of air travel where flying is at its most fuel efficient. It takes place between the ascent and descent phases and usually constitutes the majority of a journey. Technically, cruising is performed at constant airspeed and altitude. Cruising ends as the aircraft approaches its destination, with the descent phase starting in preparation for landing. In most commercial passenger aircraft, the cruising phase consumes most of the fuel.

[0101] Descent:

[0102] The descent during a flight is the portion where an aircraft loses altitude. The descent is an essential part of the landing approach. Other partial descents may serve to evade traffic, avoid bad flight conditions (turbulence or bad weather), avoid clouds (especially under contact flight rules), enter warmer air (if there is a risk of icing), or take advantage of the wind direction at a different altitude. Normal descents take place at constant airspeed and constant descent angle. The pilot controls the angle of descent by varying engine power and angle of attack (nose-down) to maintain airspeed within the specified range. At the start and during the descent phase, the engines will be operated at low power.

[0103] Approach & Landing:

[0104] Approach and landing are the final part of a flight when the aircraft returns to the ground. For landing, the airspeed and the rate of descent are reduced to the extent that a specified glide path (3 degree final approach at most airports) to the touchdown point on the runway is maintained. The reduction in speed is achieved by reducing thrust and/or creating greater drag using flaps, landing gear or air brakes. As the aircraft approaches the ground, the pilot performs a landing flare to initiate a soft landing. Landing and approach procedures are mostly carried out using an instrument landing system (ILS).

[0105] Line A (dashed): Power requirement of the propulsion system 10 over the mission with two hybrid-electric propulsion units 31, 32. The power requirement is highest during take-off and reduces over the subsequent flight phases. The power requirement for the flight phases take-off, climb and approach and landing is served by the gas turbines 41, 42 alone (“primary operating mode”). When cruising and descending, propulsion is provided by the motor-generator units 71, 72 supplied by the fuel cell 73 (‘third operating mode’). During flight, the controller 90 controls operation of the propulsion units 31, 32 and the transitions from primary to secondary mode of operation and vice versa. The dotted area under line A represents the consumption of fuel by the gas turbines 41, 42 and the checkered area the consumption of hydrogen by the motor-generator units 71, 72 during the operating phases. The energy requirement for this reference mission results in a split of approx. 60% for fuel (SAF) and 40% for hydrogen.

[0106] Line B (solid) describes the flight altitude in ft over the time of the mission. The flight altitude is highest during cruising and is about 25,000 ft.

[0107] In the diagram according to FIG. 3b, the associated accumulated energy requirement is shown by the solid and dashed lines C and D during the aforementioned operating phases and operating modes. Line C represents the energy requirement during the primary operating mode, and line D, during the third operating mode.

[0108] The diagrams of FIG. 4a and FIG. 4b represent the values for a second variant of the invention. Here, the internal combustion engines are supported during take-off and landing by extending the use of the fuel cell system in order to achieve a reduction in the power requirement of the thermal machines (downsizing). In this case, combined operation and power output of gas turbines 41 and 42 and motor-generator units 71, 72 takes place, 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 proportion of a flight which is emission-free to approximately 70% (checkered area in FIG. 4a).

[0109] Line E (dashed): The power requirement of the propulsion system 10 over the mission with two hybrid-electric propulsion units 31, 32 according to FIG. 4a is in principle the same as shown in FIG. 3a. Here, too, the power requirement is highest during take-off and reduces over the subsequent flight phases.

[0110] Line F (solid) describes the flight altitude in ft over the time of the mission. The flight altitude is highest during cruising and is approximately 25,000 ft, as in FIG. 3a.

[0111] The diagrams in FIG. 3a and FIG. 3b clarify an essential aspect of the invention, namely that for many aircraft, in particular for regional aircraft with propeller propulsion, the required thrust for take-off is significantly higher than for cruising and the thermal machines are therefore only operated at around half of their capacity over a large proportion of the flight mission. This means that conventional turboprop propulsion systems are operated outside of the optimum operating point, which is closer to the point of maximum power output. In contrast, a hybrid-electric propulsion system can be optimized for the two different operating modes.

[0112] With regard to the operating phases of the propulsion system 10, the following basic operating states result: [0113] 1. FIG. 5 shows a system diagram for take-off, climb and approach/landing. Both gas turbines 41 and 42 are in operation and drive the propellers 61, 62 via the interposed reduction gearboxes 51, 52 (solid lines). Thrust control takes place centrally via hybrid propulsion controller 90 to gas turbines 41 and 42 (‘primary operating mode’). The electric motor-generator units 71 and 72 do not produce any propulsion power (dashed lines). [0114] 2. The secondary mode of operation for take-off and climb is shown in FIG. 2. In this combined operating mode, the propulsors (61, 62) receive propulsion power from both the first and second internal combustion engines (41, 42) and from the first and second motor-generator units (71, 72). The associated power requirement and cumulative energy consumption can be seen in the diagrams in FIG. 4a and FIG. 4b. The reduced fuel consumption for the gas turbines due to the additional propulsion power from the electric motor-generator units 71, 72 becomes clear here. [0115] 3. The system state for cruising and descent can be seen in FIG. 6. The power/torque requirement drops to cruising level (FIGS. 3a and 4a) and the power of the two gas turbines 41, 42 is reduced while the propulsion power from the two motor-generator units 71, 72 can be transmitted to both propellers 61 and 62 in an equally distributed manner via the coupled gearboxes 51, 52. During cruising and subsequent descent, the thrust requirement is adjusted via the hybrid propulsion controller 90. This is responsible for thermal and electrical control (‘third operating mode’). [0116] 4. The system architecture according to the invention enables a symmetrical thrust due to the power distribution over the electrical network even in the case of a critical fault, failure of an internal combustion engine. On the side of the failed internal combustion engine 41, 42, the motor-generator unit 71 or 72 can be switched on and thus at least part of the failed thrust can be compensated.

LIST OF REFERENCE SIGNS

[0117] 10 Propulsion system [0118] 20 Aircraft [0119] 21 T-tail [0120] 22 Fuselage [0121] 23 Cockpit [0122] 24 Passenger cabin [0123] 26 Wings, left and right [0124] 27 Fuselage tail [0125] 31, 32 Propulsion units, left and right [0126] 33, 34 Engine nacelles, left and right [0127] 41, 42 Gas turbines, left and right [0128] 43, 44 Wing integral tanks, left and right [0129] 51, 52 Reduction gearboxes, left and right [0130] 61, 62 Propellers, left and right [0131] 71, 72 Motor-generator units, left and right [0132] 73 Fuel cell [0133] 74 Fuel tank [0134] 80 Transmission device [0135] 81, 82 Power converters [0136] 90 Controllers [0137] A, B, C, D, E, F Lines [0138] SP Center of gravity