Gas turbine propulsion system with reduced start-up times

Abstract

Propulsion system (1) for an aircraft, comprising - at least one gas turbine (2), - at least one fluid system (4) for providing a heated fluid (6), - comprising an external auxiliary power unit (40) - wherein the at least one gas turbine (2) and the at least one fluid system (4) are coupled and arranged in such a fluid-conducting manner that at least one section (30, 34, 36) of the gas turbine (2) is supplied with the heated fluid (6), - wherein the fluid system (4) comprises a combustion unit (12) designed separately from the auxiliary engine (40), which is configured to produce the heated fluid (6) by burning fuel, - characterized in that a fluid supplied to the combustion unit (12) is bleed air from the auxiliary engine (40).

Description

[0001] The invention relates to a propulsion system, in particular for an aircraft, an aircraft comprising a propulsion system, and a method for reducing the start-up times of a propulsion system.

[0002] Propulsion systems are generally well-known; they are used, for example, to generate thrust for an aircraft. Such propulsion systems typically include a gas turbine, with thrust being generated by the recoil effect of the air and exhaust flow produced by the gas turbine (turbojet). The gas turbine typically also drives a shrouded propeller (turbofan) or an open propeller (turboprop), a process known as bypass airflow, which serves to improve the efficiency of the propulsion system. There are various ways to start such a propulsion system. These methods have in common that, once idle speed is reached after starting, it is maintained for a predetermined period, usually several minutes, to bring the individual components of the propulsion system up to operating temperature. This process consumes a significant amount of fuel.Furthermore, the drive system cannot be used to generate propulsion during this warm-up phase.

[0003] Another problem with known drive systems is that the compressor shaft typically bends after the drive system is switched off. This effect is also known as thermal bowing. The reason for this is that hot air usually rises, and therefore the upper part of the compressor shaft has a higher temperature than the lower part. A cooling phase can be incorporated to address this.

[0004] However, the long warm-up and cool-down phases cannot be combined, or can only be combined to a limited extent, with concepts such as electric taxiing or a descent with the gas turbines switched off. A taxi typically takes five to nine minutes to move an aircraft from a take-off position at the terminal to a position on the runway and / or back. If three minutes are required for both warming up and cooling the turbine, the gas turbine will be at least partially operating during these five to nine minutes, even with electric taxiing, thus reducing the fuel-saving effect of the electric taxi.

[0005] Documents US 2012 / 0153076A1 and US 2020 / 0040848A1 each disclose a propulsion system for an aircraft with a gas turbine and an auxiliary power unit. Document US 3677012A discloses a gas turbine.

[0006] It is therefore an object of the invention to provide a propulsion system, an aircraft, and a method that reduce or eliminate one or more of the aforementioned disadvantages. In particular, it is an object of the invention to provide a solution that reduces the start-up times of a propulsion system for an aircraft.

[0007] According to a first aspect, this problem is solved by a propulsion system, in particular for an aircraft, comprising at least one gas turbine, at least one fluid system for providing a heated fluid, wherein the at least one gas turbine and the at least one fluid system are coupled and arranged in such a fluid-conducting manner that at least one section of the gas turbine is supplied with the heated fluid, preferably by the heated fluid flowing into the at least one section of the gas turbine.

[0008] The propulsion system can be, for example, a jet engine or a turboprop. The gas turbine can, for example, be a component of a turbofan or a turboprop. The gas turbine can, for example, comprise a compressor, a combustion chamber, and / or a turbine. The heated fluid is directed into at least one of the aforementioned sections of the gas turbine. The compressor and / or the turbine can have multiple spools, for example, a high-pressure spool, a low-pressure spool, a booster spool, and / or additional spools.

[0009] The drive system includes the fluid system for supplying a heated fluid. The heated fluid is used to preheat the gas turbine, in particular its sections, preferably the compressor, combustion chamber, and / or the turbine itself. The heated fluid supplied to the gas turbine can, for example, have a temperature of more than 100°C, more than 200°C, or more than 500°C.

[0010] The fluid system can be located inside or outside a housing of the drive system in which the gas turbine is located. The sections of the gas turbine are preferably fluid-conducting and coupled to the fluid system, in particular by fluid lines. Preferably, a fluid line can lead from the fluid system to the gas turbine and there connect to a fluid distributor that distributes the fluid to the individual sections of the gas turbine. The fluid lines and / or the fluid distributor can also be part of the fluid system.

[0011] The invention is based on the understanding that existing drive systems and the methods used have or cause high fuel consumption. In particular, a significant amount of fuel, especially kerosene, is used due to the warming up of the drive systems at idle speed and the run-on time after operation of the drive system to prevent thermal bowing.

[0012] The invention significantly reduces fuel consumption by preheating the gas turbine before operation, specifically before the fuel is ignited in the combustion chamber, using the preheated fluid. This preheating of the gas turbine avoids or at least reduces the need to heat up the turbine during operation at idle speed.

[0013] A preferred embodiment of the drive system comprises at least one electric rotary unit for rotating one or two or more shafts of the drive system and a power supply unit for providing electrical power to the rotary unit. The electric rotary unit allows the shaft of the drive system, in particular a shaft of the gas turbine, to be rotated after the drive system has been used. By supplying the heated fluid and rotating the shaft of the drive system, the effect of thermal bowing is reduced or eliminated.

[0014] The drive system is further characterized by the fact that the fluid system includes a combustion unit, the combustion unit being configured to generate the heated fluid by burning fuel. The fuel can be, for example, kerosene. The drive system includes, for example, a fuel tank and preferably a supply line from the fuel tank to the combustion unit.

[0015] The combustion unit can be, for example, a conventional combustion chamber that mixes air with vaporized fuel and burns the mixture. Alternatively, the combustion unit can also be an afterburner that further heats already hot exhaust gases.

[0016] In a further preferred embodiment of the drive system, the fluid system includes a compressor, which supplies the combustion unit with a compressed fluid. This fluid is, in particular, air. The combustion unit thus has a compressed fluid and fuel available, enabling efficient combustion and the efficient generation of the heated fluid.

[0017] In a further preferred embodiment of the drive system, it comprises an auxiliary engine, wherein the auxiliary engine includes the supply unit, preferably configured as an electric generator. A fluid supplied to the combustion unit is drawn from the auxiliary engine. The bleed air is, in particular, a mixture extracted from the area downstream of the combustion chamber. The auxiliary engine is preferably a second gas turbine. The auxiliary engine is characterized, in particular, by the fact that, at comparable operating points, it has a significantly lower fluid flow rate, especially an air mass flow rate, than the drive system. The air mass flow rate of the auxiliary engine can, for example, be two, three, five, ten, or twenty times lower than that of the drive system.

[0018] Furthermore, the auxiliary power unit has a lower power output, which can be, for example, ten times lower than that of the propulsion system's gas turbine. The auxiliary power unit can be designed as an internal or external unit, meaning it can be integrated within the propulsion system's housing or located outside of it.

[0019] In a further preferred embodiment of the drive system, the fluid system comprises a fuel cell system, wherein the fuel cell system includes: a fuel cell, and preferably a reformer, wherein the fuel cell is or comprises the supply unit and is configured to generate electrical energy, in particular by means of fuel provided by the reformer, especially hydrogen, and wherein the fuel cell is coupled to the combustion unit in such a way that exhaust gases from the fuel cell are supplied to the combustion unit.

[0020] A compressor is preferably connected upstream of the fuel cell, which supplies the fuel cell with a compressed fluid. In particular, this can be the compressor described above.

[0021] Another preferred embodiment of the drive system is characterized by the fact that the fuel provided by the reformer is hydrogen and / or the reformer is operated with the fuel, in particular with a fuel other than hydrogen. The fuel is specifically the medium used to operate the gas turbine. The reformer can therefore be operated with kerosene, for example. This allows both the gas turbine and the fuel cell to be operated with the same energy storage medium, such as kerosene, and eliminates the need for additional equipment, such as a hydrogen tank.

[0022] A fuel cell is understood to be, in particular, a galvanic cell that converts the chemical reaction energy of a fuel and an oxidant, which is supplied continuously, into electrical energy. Specifically, the fuel cell described above is a hydrogen-oxygen fuel cell.

[0023] In a further preferred embodiment of the drive system, the rotary unit comprises a starter unit, in particular an electric, pneumatic, hydraulic and / or mechanical starter unit, or is designed as a starter unit, in particular an electric, pneumatic, hydraulic and / or mechanical starter unit, which is configured to rotate the at least one shaft or the two or more shafts of the drive system.

[0024] A further preferred embodiment of the drive system is characterized in that the rotary unit includes, or is designed as, an electric motor, which is preferably arranged directly on a shaft of the drive system, in particular on a shaft of the gas turbine. Since drive systems, especially for aircraft, can have two or more shafts, it is preferred that the rotary unit has several electric motors. This embodiment is advantageous because the electric motor(s) can be used to provide engine support during flight, in particular to switch off the gas turbine, for example, during a descent.

[0025] In another preferred embodiment of the drive system, it is provided that one or the shaft of the gas turbine is non-rotatably connected to a bladed wheel and preferably the drive system is designed as a ducted jet engine or as a turboprop.

[0026] According to another aspect, the aforementioned task is solved by an aircraft comprising a propulsion system according to one of the previously described design variants. The propulsion system can have two or more gas turbines, whereby the number of fluid systems can be less than the number of gas turbines, so that one fluid system can also supply more than one gas turbine.

[0027] According to a further aspect, the aforementioned problem is solved by a method for reducing the start-up times of a propulsion system, in particular for an aircraft, comprising the steps of: generating a heated fluid, in particular by a combustion unit, to which a compressed fluid and / or exhaust gases are supplied; supplying, in particular directing, the heated fluid to at least one section of a gas turbine of the propulsion system, and preferably rotating one or two or more shafts of the propulsion system, in particular the gas turbine.

[0028] The heated fluid is supplied, in particular, when the gas turbine is switched off. It is especially preferred that the generation and / or supply of the heated fluid takes place during a taxiing maneuver of the aircraft.

[0029] The process and its possible further developments exhibit characteristics and process steps that make them particularly suitable for use in a drive system and its further developments. For further advantages, implementation variants, and details of the other aspects and their possible further developments, please refer to the previously provided description of the corresponding characteristics and further developments of the drive system.

[0030] Preferred embodiments are explained by way of example with reference to the accompanying figures. These show: Fig. 1: a schematic view of an exemplary embodiment of a drive system; Fig. 2: a schematic view of another exemplary embodiment of a drive system; Fig. 3: a schematic view of another exemplary embodiment of a drive system; and Fig. 4: A schematic view of an exemplary procedure.

[0031] In the figures, identical or essentially functionally equivalent or similar elements are designated with the same reference symbols.

[0032] Fig. Figure 1 shows a propulsion system 1, which is designed in particular for an aircraft, for example an airplane. The propulsion system 1 comprises a gas turbine 2 and a fluid system 4. The fluid system 4 is designed to provide a heated fluid 6. The gas turbine 2 and the fluid system 4 are coupled and configured such that the heated fluid 6 enters at least one section 30, 34, 36 of the gas turbine 2. In the Fig. In the embodiment shown, the heated fluid 6 is directed into the compressor 30, the combustion chamber 34 and the turbine 36.

[0033] The fluid system 4 further comprises a combustion unit 12. The combustion unit 12 is configured to generate the heated fluid 6 by burning fuel. The fuel is supplied to the combustion unit 12 from a fuel tank 14 via a fuel supply line 16. Furthermore, the combustion unit 12 is fluidically coupled to a compressor 18. The compressor 18 supplies the combustion unit 12 with compressed fluid. The heated fluid 6 generated by the combustion unit 12 is conveyed to the gas turbine 2 via a fluid supply line 8. The fluid supply line 8 terminates in a fluid distributor 10, which directs the heated fluid to the various sections 30, 34, and 36 of the gas turbine 2. The drive system 1 further comprises a power supply unit 20, which is electrically coupled to a starter unit 22, which will be described in more detail below.

[0034] Fig. Figure 2 shows a drive system 1 with a differently designed fluid system 4. The fluid system 4 comprises, analogous to the fluid system 4 described above, a combustion unit 12, a fuel supply line 16, and a fuel tank 14. The fuel supply to the combustion unit 12 is essentially analogous to the embodiment described above.

[0035] A fluid is supplied to the combustion unit 12 via a bleed air line 46. For this purpose, the fluid system 4 includes an auxiliary engine 40. The auxiliary engine 40 can, for example, be a small gas turbine. The gas turbine 40 is driven by a fluid 42 and a combustion process, enabling the generation of electrical energy by a generator 47. A bleed air line 46 is coupled to the compressor of the auxiliary engine 40 in such a way that pre-compressed fluid can be drawn off and supplied to the combustion unit 12.

[0036] The generator 47 is coupled to a power distribution unit 50 by means of an electrical control point 48. The power distribution unit 50 can also be a supply unit. The power distribution unit 50 can supply electrical energy to a landing gear 52, an electric taxi drive 54, and / or air conditioning systems 56. In addition, the power distribution unit 50 can supply electrical energy to electric starter units 22. The electric starter units 22 are mechanically coupled to the gas turbine 2 and can drive a shaft of the drive system 1, in particular of the gas turbine 2.

[0037] Fig. Figure 3 shows a drive system 1 with a fluid system 4 comprising a fuel cell 60. Compressed fluid is supplied to the fuel cell 60 by means of a compressor 18. In addition, fuel supplied by a reformer 64 is provided to the fuel cell 60 via a fuel line 66. The reformer 64 is specifically arranged and configured to produce the fuel from a fuel that is supplied to the reformer 64, for example, from the fuel tank 14. The fuel cell 60 is coupled to the power distribution unit 50 described above. The exhaust gases from the fuel cell 60 are supplied to the combustion unit 12 via an exhaust line 62.

[0038] The drive system 1 also comprises a first electric motor 68 and a second electric motor 70, which are arranged and designed to drive the shaft of the drive system 1, in particular that of the gas turbine 2, in a rotary manner.

[0039] The characteristics of the in the Fig. 1, Fig. 2 to Fig. The 3 drive systems shown can be partially or completely combined with each other.

[0040] In Fig. Figure 4 shows a method for reducing the start-up times of a propulsion system for an aircraft. In step 100, a heated fluid 6 is generated. The generation of the heated fluid 6 is carried out in particular with a combustion unit 12.

[0041] In step 102, the heated fluid 6 is provided, in particular by directing it to sections 30, 34, 36 of a gas turbine 2 of the drive system 1. This can be done, for example, with a fluid supply line 8 that leads into a fluid distributor 10, so that the fluid distributor 10 directs the heated fluid 6 to the individual sections 30, 34, 36.

[0042] In step 104, a shaft of the drive system 1, in particular of the gas turbine 2, is rotated. This rotation prevents thermal bowing. Furthermore, an electric motor can also enable the drive system to start quickly. REFERENCE MARK LIST 1 Drive system 2 gas turbines 4 Fluid system 6 heated fluid 8 Fluid supply 10 fluid distributors 12 combustion units 14 Fuel tank 16 Fuel supply line 18 Compressor 20 Power supply unit 22 electric starter unit 30 compressors 34 Combustion chamber 36 Turbine 40 Auxiliary engine 42 Fluid 44 Fuel supply line 46 Bleed air line 47 Generator 48 electrical lines 50 power distribution unit 52 Chassis 54 Electric taxi drive 56 air conditioning systems 60 Fuel cell 62 Exhaust pipe 64 Reformers 66 Fuel supply line 68 first electric motor 70 second electric motor

Claims

[1] Propulsion system (1) for an aircraft, comprising - at least one gas turbine (2), - at least one fluid system (4) for providing a heated fluid (6), - comprising an external auxiliary power unit (40) - wherein the at least one gas turbine (2) and the at least one fluid system (4) are coupled and arranged in such a fluid-conducting manner that at least one section (30, 34, 36) of the gas turbine (2) is supplied with the heated fluid (6), - wherein the fluid system (4) comprises a combustion unit (12) designed separately from the auxiliary engine (40), which is configured to produce the heated fluid (6) by burning fuel, - characterized by , that a fluid supplied to the combustion unit (12) is bleed air from the auxiliary engine (40). [2] Drive system (1) according to claim 1, comprising - at least one electric rotary unit (22, 68, 70) for rotating one shaft or two or more shafts of the drive system (1) and a supply unit (47) for supplying electrical power to the rotary unit. [3] Drive system (1) according to claim 2, wherein the auxiliary drive (40) comprises the supply unit (47), and wherein the supply unit (47) is designed as an electric generator (47). [4] Drive system (1) according to claim 2 or 3, wherein the fluid system (4) comprises a fuel cell system, the fuel cell system comprising: - a fuel cell (60), and - a reformer (64), - wherein the fuel cell (60) is or comprises the supply unit and is configured to generate electrical energy using fuel supplied by the reformer (64), and wherein the fuel cell (60) is coupled to the combustion unit such that exhaust gases from the fuel cell (60) are supplied to the combustion unit (12). [5] Propulsion system (1) according to claim 4, wherein the fuel provided by the reformer (64) is hydrogen and / or the reformer (64) is operated with the fuel. [6] Drive system (1) according to any one of claims 2 to 5, wherein the rotary unit comprises an electrical, pneumatic, hydraulic and / or mechanical starter unit (22) or is configured as an electrical, pneumatic, hydraulic and / or mechanical starter unit (22) which is configured to rotate the at least one shaft of the gas turbine (2). [7] Drive system (1) according to one of claims 2 to 6, wherein the rotary unit comprises at least one electric motor (68, 70) or is designed as an electric motor (68, 70) which is arranged directly on the at least one shaft of the gas turbine (2). [8] Propulsion system (1) according to one of the preceding claims, wherein a shaft of the gas turbine (2) is non-rotatably connected to a blade wheel and the propulsion system (1) is designed as a ducted jet engine or as a turboprop. [9] Drive system (1) according to one of the preceding claims, wherein the fluid system (4) comprises a compressor (18), wherein the compressor (18) provides a compressed fluid to the combustion unit (12). [10] Aircraft comprising a propulsion system (1) according to any of the preceding claims. [11] Method for reducing the start-up time of a propulsion system (1) for an aircraft, comprising the steps: - Generating a heated fluid (6) by means of a combustion unit (12) to which a compressed fluid and / or exhaust gases are supplied; - Providing the heated fluid (6) to at least one section of a gas turbine (2) of the propulsion system (1); and - characterized by , that a fluid supplied to the combustion unit (12) is bleed air from an external auxiliary engine (40), and wherein the combustion unit (12) is designed separately from the auxiliary engine (40).

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