Aircraft propulsion

Abstract

The invention relates to an aircraft propulsion system (10) comprising a turbomachine (20) through which a gas flow (G) can flow and a suspension device (30), wherein the turbomachine (20) can be connected to an aircraft by means of the suspension device (30).

Description

[0001] The invention relates to an aircraft propulsion system comprising a turbomachine through which a gas flow can flow and a suspension device, wherein the turbomachine can be connected to an aircraft by means of the suspension device.

[0002] Future engine concepts, such as water-enhanced turbofan technology or fuel cell-powered systems, will require relatively high levels of heat to be dissipated into the environment. One goal of future developments is to improve thermal efficiency by recovering the waste heat contained in the exhaust gas. Energy recovery can be achieved, for example, through the use of evaporators in a bottoming cycle (Rankine or Cheng), as in the Water Enhanced Turbofan (WET) concept, or through exhaust gas recuperation, as in the Intercooled Recuperated Aero Engine (IRA). In these concepts, a heat exchanger is provided downstream of the turbine to recover the waste heat from the exhaust gas.

[0003] Since such heat exchangers typically have a large structural volume, this can result in the overall engine becoming larger, generating more drag due to the increased surface area. Gas flow routing downstream of the turbine to the heat exchanger must also be designed to minimize pressure losses in the exhaust flow. Overall, this can increase system complexity, the effort required to transport the heat to the heat exchanger, and thus the aircraft's drag.

[0004] Based on this, it is an object of the present invention to improve an aircraft propulsion system. In particular, an aircraft propulsion system is to be proposed that enables improved exhaust energy recovery. This is achieved according to the invention by the teaching of the independent claims. Advantageous embodiments of the invention are the subject of the dependent claims.

[0005] To achieve the object, an aircraft propulsion system is proposed, comprising a turbomachine through which a gas flow can flow and a suspension device by means of which the turbomachine can be connected to an aircraft, wherein a flow channel is formed on the suspension device downstream of the turbomachine in the flow direction.

[0006] By forming a flow channel on the suspension device, additional installation space can be created, which can be used, for example, to arrange subsystems of the aircraft propulsion system and / or the aircraft. The suspension device is, in particular, a pylon, or is referred to as a pylon, and is designed for the mechanical attachment of at least one engine or at least one turbomachine to an aircraft, and in particular to an aircraft wing. The suspension device is typically connected to the aircraft at structurally reinforced attachment points in order to be able to absorb and transmit any loads that occur. In addition to the mechanical attachment, the suspension device or the pylon can also have lines for flight operations, such as electrical or hydraulic systems and / or for fuel.

[0007] Such a suspension device or pylon is typically designed to be aerodynamically advantageous in order to reduce air resistance resulting from its flow.

[0008] The aircraft engine is designed specifically to generate propulsion for an aircraft. The turbomachine can, for example, be an engine powered by a combustion device and / or a fuel cell system. Typically, drawn-in ambient air is used to flow through the turbomachine or as a working fluid, and is compressed, burned, and expanded, for example, to perform work and generate propulsion for the aircraft.

[0009] The turbomachine is, in particular, an aircraft engine and, in some embodiments, may comprise a fan, a compressor, a combustion chamber, and a turbine, and may, for example, be configured as a turbofan engine. By means of the fan, ambient air can be drawn in as a working fluid or gas flow and compressed in the compressor to increase the pressure, particularly progressively in the direction of flow. In a combustion chamber arranged downstream of the compressor in the flow direction of the turbomachine, the compressed working fluid is usually combusted with a fuel to generate combustion gases at high pressure and high temperature. The combustion gases flow as a gas flow from the combustion chamber to the turbine, where they expand to perform work. In particular, the expansion of the combustion gases in the turbine section also drives at least one shaft or shaft device.

[0010] The invention is based, among other things, on the idea of ​​using the suspension device or the pylon to create an additional flow-through space downstream of the turbomachine and / or to be able to integrate this flow channel on the aircraft fuselage or on the wing of the aircraft. The flow channel can be a physically limited flow-through space extending in the direction of flow, into which the gas flow can enter after leaving the turbine. This flow channel can be used to structurally accommodate or integrate further, in particular flow-through, components and / or subsystems. These components and / or subsystems can be arranged and / or designed so that the gas flow can flow through and / or around it.By integrating such components and / or subsystems into a flow channel formed on the suspension system, the aircraft's drag during flight can be reduced, which can have a positive impact on energy and / or environmental performance. Arranging additional components in the flow channel can also make them more easily accessible for maintenance and / or repair.

[0011] In one embodiment, at least one heat exchanger device is arranged in the flow channel. The heat exchanger device can be configured to extract energy from the gas flow, whereby the temperature of the gas flow can be lowered in order to be able to recover water from the gas flow. In addition, the heat exchanger device can be configured to use this energy to heat and / or evaporate water, in particular the water extracted from the gas flow, in order to be able to make it available, in particular, for combustion in the combustion chamber of the turbomachine. In this way, the water to be supplied or supplied to the heat exchanger can be essentially kept in a circuit, whereby an additional water supply for a combustion process of a turbomachine designed as a WET engine can be eliminated or reduced.

[0012] In a turbomachine designed as a wet-temperature engine, the compressed working fluid or the compressed gas flow can be combined with fuel and water, particularly water vapor, which is extracted from the gas flow by means of the heat exchanger. This can be combusted in the combustion chamber to improve the combustion process. By adding water or water vapor to the gas flow or the combustion process, the efficiency of the turbomachine can be increased and / or undesirable nitrogen oxide (NOx) emissions can be reduced.

[0013] The heat exchange device can comprise, for example, a heat exchanger or evaporator, a condenser, and a water separator through which the gas flow can flow. The gas flow can be pre-cooled in the evaporator, and the water contained therein can condense in the condenser by further cooling the gas flow. Liquid water can be separated from the gas flow in the water separator. The heat exchange device can be designed so that the water supplied for combustion can be largely condensed and recovered.

[0014] In one embodiment, the heat exchanger device can be flowed through in a direction substantially parallel to the flow direction of the turbomachine or substantially parallel to a rotational axis of the turbomachine. By flowing through the heat exchanger device parallel to the flow direction of the turbomachine, pressure losses in the gas flow and / or the flow chamber can be reduced, thereby improving the efficiency of the aircraft propulsion system.

[0015] A guide system for guiding the gas flow and / or an air distribution device can be arranged in the flow channel, in particular upstream of the heat exchanger in the flow direction, to supply the gas flow to the heat exchanger. A particularly planar outlet or outlet nozzle can be provided downstream of the heat exchanger, wherein the outlet or outlet nozzle can, for example, have a narrowing in the spanwise direction of a wing or an aircraft wing. The spanwise direction extends between a fuselage-wing transition or a wing root and a wing tip.

[0016] In one embodiment, an inlet diameter of the flow channel substantially corresponds to the diameter of a turbine outlet of the turbomachine. This allows essentially the entire gas flow, as well as the energy / water contained therein, to enter the flow channel and be made available there for use or processing. In particular, the flow channel can be structurally attached directly to the turbine outlet, so that the flow channel and the turbine outlet can have substantially the same cross-section to enable flow guidance with as little loss as possible.

[0017] In one embodiment, a width of a flow-through cross-section of the heat exchanger device or of a guide system and / or distribution device corresponds at most to the diameter of the turbine outlet of the turbomachine. This can enable a loss-reduced or loss-free supply of the gas flow to the heat exchanger device. The width of the heat exchanger device extends in particular transversely to the flow direction or in the wingspan direction of the aircraft. Furthermore, it can be provided that a width of a flow-through cross-section of the heat exchanger device corresponds at least to the width or spanwise width of the suspension device or pylon between the flow channel and the aircraft or a wing of the aircraft, or is oriented towards it. This can enable dimensioning or aerodynamic design of the suspension device and / or the flow channel.

[0018] In one embodiment, the flow channel has a geometry extending in the flow direction of the turbomachine. A flow-through cross-section of the flow channel can taper and / or widen in the flow direction, in particular to achieve a desired flow pattern and / or to enable improved integration of components, such as the heat exchanger device. At its downstream end, the flow channel can have an outlet or an outlet nozzle. This outlet of the flow channel can, in particular, be planar and form at least one angle with the flow direction or the rotational axis of the turbomachine.

[0019] In one embodiment, the flow channel is formed by a housing or by at least one attachment of the suspension device arranged on the suspension device. The housing is in particular an outer cover of the suspension device, which is designed in a particularly aerodynamically favorable manner in order to reduce air resistance. The at least one attachment arranged on the suspension device can also be arranged outside a housing of the suspension device and, in particular, form a flow channel with a housing surface of the suspension device. The flow channel can be at least partially delimited by the housing and / or the at least one attachment, wherein the housing and / or the at least one attachment can assume gas flow-guiding functions.In particular, the housing of the suspension device and / or the at least one attachment part arranged thereon is designed in such a way that supersonic flows on its surface are avoided or even prevented in order to keep air or flow resistance as low as possible.

[0020] In one embodiment, the turbomachine is arranged in a nacelle, and the suspension device is configured to connect the nacelle to the aircraft. A nacelle generally has a tubular structure and / or a housing along a longitudinal axis and comprises an air inlet arranged at the front in the direction of flight, a central section arranged downstream of the air inlet, which surrounds the turbomachine, and a downstream section, which may have an outlet nozzle and is designed to discharge the gas flow, in particular in the turbine or turbine outlet region. The nacelle can form, provide, or surround the turbine outlet, so that the flow channel can connect to the downstream end of the nacelle. This enables a streamlined design of the aircraft propulsion system.

[0021] In one embodiment, the suspension device is configured to connect the nacelle or turbomachine to a wing or wing of the aircraft. In other embodiments, the suspension device can be configured to connect the nacelle or outer shell of the turbomachine to an aircraft fuselage. The suspension device thus represents a structural component of an aircraft.

[0022] According to a further aspect, an aircraft is proposed which has at least one aircraft propulsion system described herein. By means of an aircraft configured in this way, the effects and advantages of the previously described aircraft propulsion system can be utilized.

[0023] In one embodiment, the flow channel formed on the suspension device extends between the turbine outlet of the turbomachine and a trailing edge of the wing. This allows the suspension device to be extended in a longitudinal direction of the aircraft in such a way that an optimized design of the flow channel or an increase in installation space can be provided by means of the flow channel. Furthermore, fastening the suspension device or pylon designed in this way can enable improved connection to the wing and thus to the aircraft.

[0024] Further features, advantages, and possible applications of the invention will become apparent from the following description in conjunction with the figures. In general, features of the various exemplary aspects and / or embodiments described herein may be combined with one another, unless clearly precluded in the context of the disclosure.

[0025] In the following part of the description, reference is made to the figures shown to illustrate specific aspects and embodiments of the present invention. It is understood that other aspects may be utilized and structural or logical changes to the illustrated embodiments are possible without departing from the scope of the present invention. The following description of the figures is therefore not to be understood as limiting. It shows Fig. 1 is a first schematic diagram of an exemplary aircraft engine according to the present disclosure; Fig. 2 is another schematic diagram of the exemplary aircraft engine according to the present disclosure; Fig. 3 is another schematic representation of the exemplary aircraft propulsion system according to the present disclosure.

[0026] Fig. 1 shows an exemplary aircraft propulsion system 10 according to the invention with a turbomachine 20 and a suspension device 30 in a schematic representation.

[0027] The exemplary turbomachine 20 has a fan 22, a compressor 23, a combustion chamber 24, and a turbine 25, through which a gas flow G can flow in a flow direction S (illustrated by the arrow) or through which the gas flow G flows during operation of the turbomachine 20. The gas flow G can, for example, be a core flow of a turbofan engine. The turbomachine 20 is arranged in a nacelle 21, and the turbomachine 20 or the nacelle 21 can be attached to or connected to an aircraft (not shown) by means of the suspension device 30.

[0028] A flow channel 31 is formed on the suspension device 30 and is arranged downstream of the turbomachine 20 in the flow direction S of the gas flow G, in which flow channel 31, in the illustrated embodiment, a heat exchange device 18 is arranged. Downstream of the turbine 15 in the flow direction S, the heat exchange device has an evaporator 181 which is designed to cool the gas flow G and / or to evaporate water obtained from the gas flow G or to generate water vapor using energy from the gas flow G. This water vapor can be supplied to the gas flow G via a vapor supply 185, in particular with a fuel, for combustion in the combustion chamber 24, in particular by means of a suitable supply device. The heat exchange device 18 has a condenser 182 and a water separation device 183 downstream of the evaporator 181 in the flow direction, through which the gas flow G can flow or through which the gas flow G can flow.during operation of the turbomachine 20, the gas flow G flows through it. The gas flow G can leave the aircraft engine 10, the suspension device 30 or the flow channel 31 via an outlet 32 ​​and, in particular, be discharged to the environment.

[0029] The separated water can, for example, be fed via an optional water treatment system 186 into a water reservoir 187, where it can be available for further use. By means of a supply device 184, the water can be provided to the evaporator 181 to generate water vapor, which can be supplied to the gas flow G via the vapor supply 185.

[0030] Fig. 2 shows the exemplary aircraft engine 10 according to the invention from Fig. 1 with a turbomachine 20 and a suspension device 30 in a schematic longitudinal section.

[0031] In the illustrated embodiment, the turbomachine 20 is accommodated in a nacelle 21 and is arranged on a wing 11 of an aircraft by means of the suspension device 30 or is firmly connected to this. A flow channel 31 formed on the suspension device 30 extends in a longitudinal direction of the aircraft or of the turbomachine 20 and has a geometry extending in the flow direction S. The flow channel 31 adjoins a turbine outlet 25a of the turbine 25. An inlet diameter 33 or inlet cross section of the flow channel 31 can essentially correspond to a diameter or cross section of the turbine outlet 25a of the turbomachine 20. Starting from the turbine outlet 25a orthe inlet cross-section of the flow channel 31, the cross-section of the flow channel 31 can taper and / or widen in order, for example, to accommodate the heat exchange device 18 and to be able to arrange it therein in particular in a way that is favorable for flow.

[0032] Flow guide devices (not shown here) can be provided and / or configured in the flow channel 31 to guide the gas flow G to the heat exchange device 18. The heat exchange device 18 can be arranged and / or configured such that it can be flowed through in a direction parallel to the flow direction S of the turbomachine 20, whereby pressure losses in the gas flow G can be reduced.

[0033] In the illustrated embodiment, the flow channel 31 extends between the turbine outlet 25a of the turbomachine 20 and a trailing edge 12 of the wing 11, which can improve the stability of the attachment provided by the suspension device 30. The flow channel 31 can be formed or delimited by a housing of the suspension device 30, thereby achieving a weight-reduced integration.

[0034] Fig. 3 shows the exemplary aircraft propulsion system 10 according to the invention from Fig. 1 and Fig. 2 with a turbomachine 20 and a suspension device 30 in a schematic cross-sectional view.

[0035] The illustration shows that a maximum width B of a flow-through cross-section of the heat exchanger device 18 can essentially correspond to the diameter or cross-section of the turbine outlet 25a of the turbomachine 20. Furthermore, a minimum width B of the flow-through cross-section of the heat exchanger device 18 can be dependent on a particular minimum width b of the suspension device 30, for example, in a connection area to the wing 11 or the aircraft. This can enable a weight-optimized and / or aerodynamically favorable design of the aircraft propulsion system 10. LIST OF REFERENCE SYMBOLS 10 Aircraft propulsion 11 Wing 12 Trailing edge of the wing 20 Turbomachine 21 gondolas 22 fans 23 compressors 24 combustion chamber 25 turbines 25a Turbine outlet 30 Suspension device 31 flow channel 32 Outlet 33 inlet diameter 18 Heat exchanger device 181 evaporators 182 Capacitor 183 Water separation device 184 Feeding device 185 Steam supply 186 Water treatment system 187 water reservoirs b Width of the suspension device B Width of the heat exchanger G Gas flow S Flow direction

Claims

[1] Aircraft propulsion (10) comprising a turbomachine (20) through which a gas flow (G) can flow and a suspension device (30) by means of which the turbomachine (20) can be connected to an aircraft, wherein a flow channel (31) is formed on the suspension device (30) downstream of the turbomachine (20) in the direction of flow (S). [2] Aircraft propulsion (10) according to claim 1, wherein at least one heat exchange device (18) is arranged in the flow channel (31). [3] Aircraft propulsion (10) according to claim 2, wherein the heat exchange device (18) is permeable in a substantially parallel direction to the flow direction (S) of the turbomachine (20). [4] Aircraft propulsion (10) according to at least one of the preceding claims, wherein an inlet diameter (33) of the flow channel (31) substantially corresponds to a diameter of a turbine outlet (25a) of the turbomachine (20). [5] Aircraft propulsion (10) according to the preceding claim, wherein a width (B) of a flowable cross-section of the heat exchange device (18) corresponds at most to the diameter of the turbine outlet (25a) of the turbomachine (20). [6] Aircraft propulsion (10) according to at least one of the preceding claims, wherein a width (B) of a flowable cross-section of the heat exchange device (18) corresponds at least to the width of the suspension device (30). [7] Aircraft propulsion (10) according to at least one of the preceding claims, wherein the flow channel (31) has a geometry extending in the flow direction (S) of the turbomachine (20). [8] Aircraft propulsion (10) according to at least one of the preceding claims, wherein the flow channel (31) is formed by a housing of the suspension device (30). [9] Aircraft propulsion (10) according to at least one of the preceding claims, wherein the flow channel (31) is formed by at least one attachment part arranged on the suspension device (30). [10] Aircraft propulsion (10) according to at least one of the preceding claims, wherein the turbomachine (20) is arranged in a gondola (21) and the suspension device (30) is arranged to connect the gondola (21) to the aircraft. [11] Aircraft propulsion (10) according to claim 10, wherein the suspension device (30) is configured to connect the gondola (21) to a wing (11) of the aircraft. [12] Aircraft comprising at least one aircraft propulsion (10) according to at least one of the preceding claims. [13] Aircraft according to claim 12, wherein the flow channel (31) formed on the suspension device (30) extends between the turbine outlet (25a) of the turbomachine (20) and a trailing edge (12) of the wing (11).

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