Propulsion system for an aircraft, aircraft with a propulsion system and method for operating an aircraft

Description

[0001] The invention relates to a propulsion system for an aircraft, designed to operate with a liquid and a gaseous fuel. The invention further relates to an aircraft with a propulsion system and a method for operating the aircraft.

[0002] In so-called "dual-fuel" propulsion systems, the combustion chamber, and thus the engine, can be operated with either a liquid fuel, such as kerosene or similar, or a gaseous fuel, such as hydrogen or similar. Operation can be either alternative, exclusively with the liquid or the gaseous fuel, or simultaneously, with a mixture of the two fuels.

[0003] A gas turbine arrangement with a supply system for a "dual-fuel" engine is specified in CN 112 709 639 A.

[0004] A nozzle device for a "dual-fuel" engine is disclosed in US 2016 / 0 201 897 A1. Further generic gas turbine arrangements are disclosed in US 5069031 A, EP 4123145 A1 and EP 4095368 A1.

[0005] The "dual-fuel" principle is generally based on the combustion chamber being designed to operate with either fuel (alternatively or simultaneously) in any operating situation, i.e., during takeoff, cruise, landing, and the like, regardless of external conditions such as pressure, temperature, and altitude. This includes the engine being able to start with either gaseous or liquid fuel, again regardless of external conditions. This also applies to high-altitude ignition, which is typically performed in an emergency at altitude, for example, at a cruising altitude of 30,000 ft.

[0006] To ensure reliable high-altitude ignition of the combustion chamber even at cruising altitude with liquid fuel, the combustion chamber must have a certain volume. The volume required for high-altitude ignition is a key factor in the design, determining the combustion chamber volume and thus the residence time of the airflow within the combustion chamber. A long residence time, in turn, leads to increased NOₓ (nitrogen oxide) emissions. Furthermore, a large combustion chamber volume increases the space required in the engine, the weight, and the cost of the combustion chamber, and consequently, the overall engine.

[0007] The invention is based on the objective of providing a propulsion system of the type mentioned above, which is optimized with regard to emission characteristics and efficiency, as well as a corresponding aircraft and a method for operating the aircraft.

[0008] The problem is solved for the propulsion system with the features of claim 1. For the aircraft, the problem is solved with the features of claim 10, and for the method with the features of claim 11.

[0009] Advantageous variants are specified in the dependent claims.

[0010] According to the invention, the propulsion system is designed such that, in the event of high-altitude ignition, only gaseous fuel can be used or is used in the at least one engine, preferably in all engines present on the aircraft ("dual-fuel"). The use of liquid fuel is not possible.

[0011] The high-altitude ignition process describes the process of igniting and ramping up the respective engine to a specific operating condition, particularly idle, at high altitude, e.g., a cruising altitude of, for example, 30,000 ft, typically under ambient conditions with temperatures down to -40 °C and atmospheric pressures of 0.3 to 0.4 bar. The duration of the high-altitude ignition process for the respective engine (with, for example, up to 100 klbf full-load thrust) and ramping up to idle typically ranges from 60 to 180 seconds.

[0012] In particular, a control device associated with the propulsion system is configured to control and / or regulate the altitude ignition process exclusively with the gaseous fuel, not with the liquid fuel.

[0013] Gaseous fuels, such as hydrogen, methane, or a gas mixture containing at least one of these gases, generally exhibit a much higher reactivity (e.g., characterized by an ignition delay time or flame speed) and a significantly larger ignition range (e.g., characterized by the ignition limits) than liquid fuels, particularly kerosene. Therefore, the combustion of gaseous fuels can proceed much more efficiently even at low loads and in the subatmospheric range, such as at high altitudes. In contrast, the combustion chamber efficiency with kerosene is very low (< 30%) under high-altitude ignition conditions, resulting in large quantities of kerosene remaining unburned. Thus, the measure proposed according to the invention can advantageously optimize emissions and the efficiency of the propulsion system, particularly during the high-altitude ignition process.Furthermore, the combustion chamber volume can be designed for high-altitude ignition with the (comparatively) highly reactive gaseous fuel, rather than with kerosene. This allows for a smaller combustion chamber volume and thus a smaller combustion chamber, which has a positive effect on NOx emissions as well as the weight and space requirements of the combustion chamber, and therefore on the engine's efficiency.

[0014] The propulsion system stores or carries a minimum quantity of gaseous fuel, which is used or can be used exclusively in the event of high-altitude ignition. This minimum quantity preferably corresponds exactly to the amount required for at least one high-altitude ignition in the (dual-fuel) engine(s) on the aircraft. Alternatively, the minimum quantity may be designed for a different defined number of high-altitude ignitions, for example, for two or more ignitions. The required quantity per engine corresponds to a partial minimum quantity. The sum of all partial minimum quantities constitutes the total minimum quantity. The minimum quantity is preferably stored even if no other operation with the gaseous fuel is planned for the flight in question.

[0015] Maintaining the minimum quantity is possible, for example, through control technology, by keeping a minimum fill level in the (main) tank compartment for gaseous fuel within the propulsion system, which contains the gaseous fuel for the rest of the flight operation.

[0016] To store the minimum quantity of fuel, at least one additional fuel compartment is provided, which (in the case of altitude ignition) is preferably connected to the combustion chamber via a separate auxiliary fuel line. The auxiliary fuel compartment is specifically arranged separately from the other (main) fuel compartments of the propulsion system. A separate auxiliary valve and / or an auxiliary pressure reducing device are preferably arranged within the auxiliary fuel line. The auxiliary fuel line can be connected to a (main) fuel line of the propulsion system. The presence of the auxiliary fuel compartment and, if applicable, the separate auxiliary fuel line advantageously provides redundancy within the fuel system from a safety perspective, as well as simple controllability in emergency situations.The risk of unplanned withdrawal of the minimum volume for a purpose other than high-altitude ignition, resulting in excessive emptying of the (main) fuel compartment, is eliminated. Furthermore, the comparatively compact design allows the auxiliary fuel compartment to be advantageously positioned and / or filled independently of the (main) fuel compartment for gaseous fuel used in other flight operations.

[0017] A defined and space-saving storage of the minimum quantity is made possible if at least one additional fuel tank compartment is available for each engine (which is to be ignited during altitude ignition). Preferably, each additional fuel tank compartment contains the partial minimum quantity required for altitude ignition of the respective engine to which the additional fuel tank compartment is assigned.

[0018] An advantageous arrangement is possible if the additional fuel compartment is located either on the respective engine to which it is assigned (for storing the minimum partial quantity of fuel for that engine) or at another location within the aircraft. Preferably, the additional fuel compartment is thus arranged separately from the (main) fuel compartment for storing the remaining quantity of gaseous fuel. "On the engine" means, in particular, within the outer engine cowling.

[0019] Preferably, the additional tank compartment is formed within a pressure-resistant auxiliary tank. The maximum filling pressure can be, for example, between 300 bar and 700 bar. For a compact design, the pressure tank is preferably not thermally insulated.

[0020] In a compact and / or weight-optimized design, the additional tank container is made of metal, in particular stainless steel, and / or comprising metal, or made of plastic, in particular (e.g. with glass fiber) reinforced, e.g. polyamide, and / or comprising plastic.

[0021] A compact design is further enhanced if the volume of the auxiliary fuel tank is designed such that at least, preferably precisely, a minimum quantity of fuel is stored or can be stored for carrying out, preferably precisely, at least one high-altitude ignition (or, preferably precisely, two or more high-altitude ignitions) in the associated engine. The volume is thus determined by the characteristics of the associated engine, in particular its fuel consumption during the high-altitude ignition. Due to the small size of the auxiliary fuel tank resulting from the demand-oriented design, a significant overall weight saving is possible with a corresponding combustion chamber design featuring a reduced combustion chamber volume.

[0022] For an advantageous reduction in size and weight, the combustion chamber volume is preferably designed for high-altitude ignition with gaseous fuel, wherein the combustion chamber volume is at least 10% smaller, preferably at least 20% or at least 30% smaller, e.g. by up to 50% smaller, than in a design of a similar engine (with the same parameters, for example the thrust) for high-altitude ignition with liquid fuel.

[0023] Preferably, the drive system is designed for operation with a kerosene-based and / or kerosene-related fuel as a liquid fuel and / or for operation with hydrogen and / or methane or a gas mixture containing at least one of these gases as a gaseous fuel.

[0024] According to the invention, the method provides that in the case of high-altitude ignition, only gaseous fuel is used in the at least one engine during the high-altitude ignition process.

[0025] For the high-altitude ignition process, a minimum amount of gaseous fuel stored in the propulsion system is used.

[0026] In the case of altitude ignition, an additional valve is switched in an additional fuel line, through which a minimum partial quantity of gaseous fuel required for the altitude ignition process of the corresponding engine is directed from an additional tank compartment to the combustion chamber, whereby in particular the pressure of the fuel is throttled via an intermediate additional pressure reducing device to a pressure level required for altitude ignition, e.g. up to 15 bar overpressure.

[0027] Further implementation variants of the procedure are specified analogously in connection with the design variants of the drive system.

[0028] The invention will now be explained in more detail using exemplary embodiments and with reference to the drawings. The drawings show: Fig. 1 shows an aircraft with a propulsion system designed for operation with a liquid and a gaseous fuel according to the prior art in schematic representation, and Fig. 2 shows an aircraft according to the invention with a propulsion system comprising a separate additional tank compartment for storing gaseous fuel for high-altitude ignition in schematic representation.

[0029] Fig. 1 Figure 1 shows a schematic representation of an aircraft 1 with a propulsion system 11 designed for operation with a liquid fuel and a gaseous fuel, as is generally known from the prior art.

[0030] The liquid fuel is, in particular, a kerosene-based or kerosene-related (i.e., comparable to kerosene) fuel. The gaseous fuel is, in particular, hydrogen and / or methane or another highly reactive gas, or a gas mixture containing at least one of these gases.

[0031] The propulsion system 11 has a tank compartment 3 in a tank container 30 for gaseous fuel, which is connected via a fuel line 5.1 to a combustion chamber 6 of an engine 2 of the propulsion system 11 for supplying it with gaseous fuel during operation. For liquid fuel, the propulsion system 11 has a tank compartment 4 in a tank container 40, which is connected via a fuel line 5.2 to the combustion chamber 6 of the engine 2 for supplying it with liquid fuel during operation.

[0032] The aircraft 1 can be operated either alternatively with one of the two fuels, i.e. exclusively with the liquid or the gaseous fuel, or simultaneously, with a mixture of the two fuels.

[0033] For this purpose, at least one, or usually all, of the engines 2 in the aircraft's propulsion system 11 are designed as so-called "dual-fuel" engines 2. In this configuration, a gas turbine assembly with the combustion chamber 6 of the respective engine 2 is designed so that it can be operated with either fuel in any operating situation, i.e., during takeoff, cruise, landing, etc., regardless of external conditions such as pressure, temperature, and altitude. This also includes the ability for the engine 2 to start with either gaseous or liquid fuel, again regardless of external conditions, even at high altitudes (e.g., 30,000 ft), i.e., with high-altitude ignition.

[0034] To ensure reliable high-altitude ignition of combustion chamber 6 at, for example, 30,000 ft using liquid fuel, the combustion chamber of combustion chamber 6 must have a comparatively large volume. To guarantee safe flight operations, combustion chamber 6 is designed for emergency high-altitude ignition, and its combustion chamber volume is determined accordingly. Consequently, combustion chamber 6 has a relatively large size with a large combustion chamber volume, which negatively impacts NOx emissions, the space required in engine 2, the weight, and the cost of combustion chamber 6 and thus the entire engine 2. Furthermore, the efficiency of combustion chamber 6 when ignited with kerosene as liquid fuel under high-altitude ignition conditions is very low (less than 30%), meaning that large quantities of kerosene remain unburned during engine start-up.

[0035] To improve emission levels and combustion efficiency, the invention proposes that, in the case of high-altitude ignition, the combustion chamber 6 be ignited exclusively with the gaseous fuel. The gaseous fuel generally exhibits a much higher reactivity (e.g., characterized by an ignition delay time or a flame speed) and a significantly larger ignition range (e.g., characterized by the ignition limits) than the liquid fuel, so that the combustion of the gaseous fuel can proceed much more efficiently even at low load points and in the subatmospheric range, such as at high altitudes. Furthermore, the combustion chamber volume of the combustion chamber 6 can advantageously be designed for high-altitude ignition with hydrogen or another, generally (comparatively) highly reactive fuel, and not for high-altitude ignition with kerosene, which allows for a smaller combustion chamber volume and thus a smaller combustion chamber 6.

[0036] In this context, it is extremely important from a safety perspective to ensure that at all times a sufficient quantity of gaseous fuel is available for each engine 2 to be ignited during altitude ignition, regardless of how much gaseous fuel is still available for the remainder of the flight.

[0037] The sufficient quantity is hereinafter referred to as the minimum quantity. The minimum quantity represents the amount of gaseous fuel required to ignite all engines 2 to be fired during one altitude ignition event (or multiple altitude ignition events). A corresponding partial minimum quantity is required for the altitude ignition of a single engine 2.

[0038] The minimum partial quantity for an altitude ignition event per engine 2 is the quantity required to achieve ignition and idle speed of engine 2, and varies depending on the size, particularly the thrust, of engine 2 and the duration of the altitude ignition event. Typically, an altitude ignition event (ignition and idle speed of engine 2) takes, for example, 60 seconds to 180 seconds for the start-up of one engine 2 (with, for example, up to 100 klbf full-load thrust).

[0039] Fig. 2 Figure 1 shows an embodiment of the aircraft 1 according to the invention with a propulsion system 11 which is designed in such a way that the minimum amount of gaseous fuel is stored, which is used exclusively in the event of high-altitude ignition.

[0040] In Fig. 2According to a preferred embodiment, at least one additional fuel compartment 7 is provided, in which or in which the minimum quantity is stored. Preferably, at least one additional fuel compartment 7 is provided for each engine 2 to be ignited during altitude ignition. Each additional fuel compartment 7 stores the partial minimum quantity sufficient for at least one altitude ignition of the corresponding engine 2.

[0041] The combustion chamber volume of combustion chamber 6 is designed for high-altitude ignition with gaseous fuel, not for liquid fuel, and therefore has a combustion chamber volume that is, for example, at least 10% or 20% smaller, or up to 50% smaller, than that of a combustion chamber 6 designed for the same engine 2 for high-altitude ignition with liquid fuel.

[0042] The additional fuel compartment 7 can be located on the respective engine 2 to which the additional fuel compartment is assigned (in particular within the outer engine cowling), or at another location within the aircraft 1. The additional fuel compartment 7 is preferably arranged separately from the fuel compartment 3 and the fuel compartment 4.

[0043] The auxiliary fuel tank compartment 7 is connected to the combustion chamber 6 via a separate auxiliary fuel line 10. An auxiliary valve 8 and an auxiliary pressure reducing device 9 are arranged in the auxiliary fuel line 10.

[0044] In the event of high-altitude ignition, the auxiliary valve 8 is switched on, thus establishing the flow connection between the auxiliary fuel tank 7 and the combustion chamber 6 via the auxiliary fuel line 10. The gaseous fuel is fed to the combustion chamber 6 via the auxiliary fuel line 10, whereby the pressure of the gaseous fuel is reduced by the auxiliary pressure reducing device 9 to a pressure level required for ignition, for example up to 15 bar overpressure.

[0045] At the in Fig. 2In the example shown, the auxiliary tank compartment 7 is formed in a pressure-resistant auxiliary tank 70, which, separate from tanks 30 and 40, is arranged here by way of example on the engine 2. When hydrogen is used as the gaseous fuel, the auxiliary tank 70 is designed, for example, for a maximum storage pressure between 300 bar and 700 bar. The auxiliary tank 70 is made of metal, in particular stainless steel, and / or is entirely made of metal. Alternatively, the auxiliary tank 70 is made of, in particular, glass fiber reinforced plastic, e.g., polyamide, and / or is entirely made of plastic.

[0046] The volume of the auxiliary fuel tank compartment 7 is designed to store the minimum partial quantity of fuel required for at least one high-altitude ignition. Therefore, the volume of the auxiliary fuel tank compartment 7 is determined by the characteristics of the engine 2, in particular its fuel consumption during high-altitude ignition.

[0047] For example, the minimum required amount of hydrogen for an altitude ignition process to ignite and start engine 2 (full thrust approx. 20 klbf) to idle within 90 seconds is approximately 0.8 kg. This corresponds to a hydrogen volume of about 20 liters at a storage pressure of 700 bar. The additional tank compartment 7 is therefore designed for a volume of approximately 20 liters, and may, for example, have a diameter of approximately 20 cm and a length of approximately 70 cm.

[0048] For a larger engine, for example, the additional fuel tank compartment 7 can be designed to be larger, or there can be several additional fuel tank compartments 7 of the same size (in Fig. 2 (not shown).

[0049] The size of the additional tank space 7 or all required additional tank spaces 7 is / are so small that, when designing the combustion chamber 6 with the reduced combustion chamber volume, an overall weight saving is possible.

[0050] Maintaining the minimum quantity is also possible from a control engineering perspective by keeping a minimum fill level in tank space 3, whereby no additional tank space 7 is available (in Fig. 2 (not shown).

[0051] By modifying the design of the propulsion system 11 with regard to high-altitude ignition not for both fuels, but only with the gaseous fuel, it is possible to design the combustion chamber 6 with a significantly reduced volume. This results in a shorter residence time of the air-fuel mixture within the combustion chamber 6 and thus a reduction in NOₓ emissions. In addition, the engine weight and the associated costs can be reduced. Reference symbol list

[0052] 1 Aircraft 2 Engine 3 Fuel compartment 30 Fuel tank 4 Fuel compartment 40 Fuel tank 5.1 Fuel line 5.2 Fuel line 6 Combustion chamber 7 Auxiliary fuel compartment 70 Auxiliary fuel tank 8 Auxiliary valve 9 Auxiliary pressure reducing device 10 Auxiliary fuel line 11 Propulsion system

Claims

1. Propulsion system (11) for an aircraft (1), configured for operation with a liquid and a gaseous fuel, comprising - at least one engine (2) with a combustion chamber (6) for operation with the liquid and / or gaseous fuel and - at least one separate tank space (3, 4) for each fuel, which can be brought or is brought into a flow connection with the combustion chamber (6) via a fuel line (5.1, 5.2) to supply it with the corresponding fuel, - means to ensure that, in the event of a high-altitude ignition, exclusively gaseous fuel can be used or is used in a high-altitude ignition process in the at least one engine (2), - means wherein a minimum quantity of gaseous fuel can be stored or is stored that can be used or is used exclusively in the event of a high-altitude ignition, characterised in that - at least one additional tank space (7) is available for storing the minimum quantity.

2. Propulsion system (11) according to claim 1, characterised in that the additional tank space (7) can be brought or is brought into a flow connection with the combustion chamber (6) via a separate additional fuel line (10).

3. Propulsion system (11) according to claim 1 or 2, characterised in that at least one additional tank space (7) is available for each available engine (2).

4. Propulsion system (11) according to any of the preceding claims, characterised in that the additional tank space (7) is arranged on the respective engine (2) to which the additional tank space (7) is assigned, or at another location within the aircraft (1).

5. Propulsion system (11) according to any of the preceding claims, characterised in that the additional tank space (7) is configured in a pressure-tight additional tank container (70).

6. Propulsion system (11) according to claim 5, characterised in that the additional tank container (70) is configured from metal, in particular stainless steel, and / or comprises metal, or is configured from plastic, in particular reinforced plastic, e.g. polyamide, and / or comprises plastic.

7. Propulsion system (11) according to any of the preceding claims, characterised in that the volume of the, optionally respective, additional tank space (7) is designed in such a way that at least, preferably precisely, one partial minimum quantity can be stored or is stored for carrying out, preferably precisely, at least one high-altitude ignition process in the assigned engine (2).

8. Propulsion system (11) according to any of the preceding claims, characterised in that the combustion chamber volume within the combustion chamber (6) is designed for high-altitude ignition with the gaseous fuel, the combustion chamber volume being at least 10% smaller, preferably at least 20% or at least 30% smaller, e.g. by up to 50% smaller, than in the case of a design of a similar engine (2) for high-altitude ignition with the liquid fuel.

9. Propulsion system (11) according to any of the preceding claims, characterised in that the propulsion system (11) is designed for operation with a kerosene-based and / or kerosene-related fuel as a liquid fuel and / or for operation with hydrogen and / or methane or a gas mixture having at least one of these gases as a gaseous fuel.

10. Aircraft with a propulsion system (11) according to any of the preceding claims.

11. Method for operating an aircraft (1), which can be operated alternatively or simultaneously with a liquid and a gaseous fuel by means of a propulsion system, in which at least one combustion chamber (6) of at least one engine (2) is supplied with the fuel(s) used via a fuel line (5.1, 5.2) from a separate tank space (3, 4), characterised in that in the event of a high-altitude ignition, exclusively gaseous fuel is used in a high-altitude ignition process in the at least one engine (2), a minimum quantity of gaseous fuel stored in the propulsion system (11) being used for the high-altitude ignition process, characterised in that, in the event of the high-altitude ignition, an additional valve (8) is switched in an additional fuel line (10), via which a partial minimum quantity of gaseous fuel required for the high-altitude ignition process of the corresponding engine (2) is channelled from an additional tank space (7) to the combustion chamber (6).

12. Method according to claim 11, characterised in that the pressure of the fuel is throttled to a pressure level required for the high-altitude ignition via an intermediate additional pressure reduction device (9).