aircraft gas turbine

Abstract

An aircraft gas turbine (1) having a high-pressure turbine section (13) and a low-pressure turbine section (15) and having a turbine exhaust duct (19) formed in a ring around an engine axis (21), which is formed by a turbine outlet casing (17) and arranged downstream of the low-pressure turbine section, wherein the turbine exhaust casing forms an outer turbine exhaust duct wall (23) and an inner turbine exhaust duct wall (25) which delimit the annular turbine exhaust duct (19), and having an annular bypass duct (29) which is arranged downstream of the low-pressure turbine section (15) outside the outer turbine exhaust duct wall (23) and has an inlet (31) formed on the outer turbine exhaust duct wall (23), wherein a part of the exhaust gas flow can be guided into the bypass duct (29), and a heat exchanger (37) for Preconditioning of fuel by means of the part of the exhaust gas flow directed into the bypass channel (29),wherein the heat exchanger is arranged in or on the bypass channel (29).,

Description

The present invention relates to an aircraft gas turbine according to the preamble of claim 1.In aircraft gas turbines, fuel is burned in a combustion chamber in order to drive first a high-pressure turbine section and then a low-pressure turbine section by means of the exhaust gas flow which is produced. In order to achieve higher efficiency, preheating of the fuel is provided in some aircraft gas turbines.In order to significantly reduce a negative air conditioning effect of aviation, some concepts propose the combustion of hydrogen in aircraft gas turbines. In this case, hydrogen is carried as cryogenic liquid in tanks in the aircraft. In such concepts, preheating the fuel having temperatures of <30 K is of particular importance, since the hydrogen must be heated to about 150-250 K in order to be able to be burned in the combustion chamber. Since hydrogen has a high specific heat capacity, considerable heat outputs are required for the preheating.The systems described above are generally known to the applicant, but do not necessarily relate to a published state of the art.EP 4 269 765 A2 discloses an aircraft gas turbine in which hydrogen is preheated via a plurality of heat exchangers which are operated with oil, bleed air and exhaust gas.The known construction is very complicated and, moreover, in particular the heat exchangers operated with oil and bleed air are located in areas which are to be regarded as critical damage areas in which, in the event of, for example, disc failures, there is the risk that flying parts damage the hydrogen circuit, which represents a great safety risk. In this construction, provision is also made for the hydrogen to be conducted through the entire exhaust gas stream, which is likewise to be regarded as a safety risk. US 2023 / 0 076 757 A1 discloses an aircraft gas turbine having the features of the preamble of claim 1.It is therefore the object of the present invention to provide an aircraft gas turbine, in particular for operation with hydrogen, which has a fuel preconditioning and which is of simple construction. At the same time, the aircraft gas turbine should meet the highest possible safety requirements.The invention is defined by the features of claim 1.The aircraft gas turbine according to the invention has a high-pressure turbine section and a low-pressure turbine section and a turbine exhaust duct which is formed annularly about an engine axis. The turbine exhaust passage is formed by a turbine outlet casing and is disposed downstream of the low pressure turbine section. The turbine exhaust casing forms an outer turbine exhaust passage wall and an inner turbine exhaust passage wall defining the annular turbine exhaust passage. The invention is characterized in that an annular bypass duct is provided which is arranged downstream of the low-pressure turbine section outside the outer turbine exhaust duct wall and has an inlet which is formed on the outer turbine exhaust duct wall, wherein a part of the exhaust gas flow can be conducted into the bypass duct. Furthermore, a heat exchanger for preconditioning liquid or gaseous fuel by means of the part of the exhaust gas flow conducted into the bypass duct is arranged in or on the bypass duct. The fuel can be liquid or gaseous or in the under- or supercritical state.The inner turbine exhaust duct wall can be formed, for example, by an outlet cone.Due to the arrangement of the bypass duct and the heat exchanger outside the outer turbine exhaust duct wall and downstream of the low-pressure turbine section, there is sufficient space available to advantageously arrange the heat exchanger. Furthermore, the heat exchanger and thus the parts conducting the liquid or gaseous fuel are not located in a radial region of blades of the high-pressure or low-pressure turbine. This provides a high degree of safety. The line guide for the supply and discharge of the liquid or gaseous fuel to and from the heat exchanger can furthermore be effected in a simple manner, since these do not have to be guided by the hot exhaust gas flow which is conducted by the turbine exhaust gas duct. The bypass duct can in a targeted manner derive a part of the exhaust gas flow from the turbine exhaust gas duct and feed it to the heat exchanger. Thus, the fuel can be preconditioned in an advantageous manner.It is preferably provided that a flow throttling device for slowing down the flow of the part of the exhaust gas flow conducted in the bypass channel is formed in the bypass channel upstream of the heat exchanger. The flow throttling device can, for example, reduce the velocity of the part of the exhaust gas flow conducted into the bypass duct from a Mach number of M=0.5 to M<0.15. As a result, the downstream heat exchanger can be flowed through with low pressure losses.The flow throttling device can be designed, for example, as a diffuser.It is preferably provided that the heat exchanger is designed as a shell-and-tube heat exchanger with a plurality of individual tubes, wherein the liquid or gaseous fuel can be conducted through the individual tubes. Such an arrangement has proven to be particularly advantageous for heat transfer.Furthermore, the individual pipes through which the liquid or gaseous fuel is conducted can be designed in such a way that they allow pipe stretches that are thermally induced.For example, the individual tubes can have an outer diameter D, wherein 3 mm≤D≤12 mm applies. As a result, the tubes can form a large heat exchanger surface with a low weight, wherein these can at present be formed with a comparatively thin wall thickness and yet still withstand high pressures on the tube inner side. When using hydrogen as fuel, for example, pressures of >50 bar can occur, so that the configuration of the individual pipes according to the invention has proven to be particularly advantageous.The heat exchanger preferably has a cross-countercurrent arrangement. Such an arrangement has proven to be particularly advantageous for heat transfer.In a particularly preferred embodiment, it is provided that the individual tubes are arranged helically around the engine axis. In other words, the individual tubes are arranged in the annular bypass duct such that they are wound around the turbine exhaust gas housing. In this case, the individual tubes can each be bent with a constant radius to the engine axis. Such an arrangement is advantageous in particular for thermally induced pipe stretches, since the individual pipes can easily expand, whereby only the radius of the bend increases. Complex bearings can thus be avoided.Preferably, provision is made for supporting struts, so-called struts, to be arranged in the bypass duct. The support struts form an advantageous stability of the bypass duct.The support struts can be arranged, for example, in the flow throttling device and thus in a region upstream of the heat exchanger.According to the invention, it is provided that a supply device for compressed bleed air is arranged on the bypass channel. The bleed air can be taken from various take-off points in the engine. The channel can be flushed, for example, by means of the bleed air. This can be done, for example, when the aircraft gas turbine is started, in order to prevent thermal shocks or to reduce thermal loads. In the event of damage in the region of the heat exchanger, the bypass channel can also be purged by means of the bleed air in order to remove leaked fuel. In principle, a temperature regulation of the part of the exhaust gas flow conducted into the bypass channel can also take place via the bleed air.The supply device of compressed bleed air can have, for example, lines arranged in the support struts and bleed air outlets arranged on the support struts. Thus, the bleed air can be introduced into the bypass channel in an advantageous manner by means of the support struts. For example, the bleed air can be discharged into the bypass duct in the flow direction of the exhaust gas.The turbine exhaust duct can open into an exhaust nozzle and the bypass duct can open into a second exhaust nozzle surrounding the exhaust nozzle.Alternatively, it can be provided that the turbine exhaust gas duct opens into an exhaust gas nozzle and the bypass duct opens into a mixer which feeds the part of the exhaust gas flow conducted into the bypass duct to the exhaust gas flow in the turbine exhaust gas duct. By means of the mixer, the part of the exhaust gas flow conducted into the bypass duct can be mixed with the exhaust gas flow in the turbine exhaust gas duct after cooling in the heat exchanger.At the inlet, a splitter formed by the outer turbine exhaust duct wall can be arranged. This advantageously allows the part of the exhaust gas stream conducted into the bypass duct to be separated off.A flap, a so-called flap, can also be arranged at the inlet. The bypass duct can be closed via the flap, which can be arranged, for example, on the splitter, wherein the region of the fuel preconditioning can be shielded from exhaust gases that are too hot.The invention is explained in more detail below with reference to the following figures. The following are shown: FIG. 1 shows a schematic sectional illustration of an aircraft gas turbine according to the invention, and FIG. 2 shows a schematic sectional illustration of the turbine outlet housing with bypass duct and heat exchanger.FIG. 1 schematically shows a partial section of an aircraft gas turbine 1 according to the invention. The aircraft gas turbine 1 has an outer housing 3 (nacelle), in which a fan 5 is arranged. The outer casing 3 forms an air passage 7 for the bypass of the air.Behind the fan 5, the air is guided into the compressor region 9 of the propulsion unit and reaches the combustion chamber 11, The exhaust gas generated in the combustion chamber 11 by combustion of hydrogen is guided through a high-pressure turbine section 13 and a low-pressure turbine section 15 and reaches the outside through a turbine exhaust duct 19 (cf. FIG. 2 ) formed by a turbine outlet housing 17.The turbine outlet housing 17 is shown schematically in partial section in FIG. 2.The turbine exhaust duct 19 is formed annularly around the engine axis 21 and is bounded by an outer turbine exhaust duct wall 23 and an inner turbine exhaust duct wall 25. The inner turbine exhaust duct wall 25 is formed by an outlet cone 27. Downstream of the low-pressure turbine section 15 in the flow direction of the exhaust gas (shown by the arrows in FIG. 2 ), a bypass duct 29 is formed, which is located outside the outer turbine exhaust duct wall 23. The bypass duct 29 is likewise of annular configuration around the engine axis 21. An inlet 31 is arranged on the outer turbine exhaust duct wall 23, wherein the outer turbine exhaust duct wall 23 forms a splitter 33. As a result, a portion of the exhaust gas flow can be conducted into the bypass duct 29. For example, 5% of the exhaust gas flow can be conducted into the bypass channel 29.In the bypass duct 29, support struts 35 are arranged, which ensure the stability of the bypass duct 29.In the bypass duct, a heat exchanger 37 is furthermore arranged, which consists of a plurality of individual tubes 37 a. The individual tubes 37 acan extend in the circumferential direction of the bypass duct 29 and can be helical. The heat exchanger 37 can thus be operated in cross-countercurrent. For preconditioning, hydrogen is conducted into the individual tubes 37 aof the heat exchanger 37 through part of the exhaust gas stream conducted into the bypass duct 29 (cf. arrows).A flow throttling device in the form of a diffuser 39 is arranged adjacent to the inlet 31, by means of which the part of the exhaust gas flow conducted into the bypass duct 29 is decelerated in order to reduce the pressure loss during the flow through the heat exchanger 37.The support struts 35 are arranged in the region of the diffuser 39.A supply device 41 for compressed bleed air is arranged on the bypass duct 29. The compressed bleed air can be taken from the air duct 7, as is schematically indicated by a feed line 43. The compressed bleed air can be conducted through lines, not shown, into the support struts 35 and blown into the bypass duct 29 in the direction of the heat exchanger 37.The turbine exhaust duct 19 opens into an exhaust nozzle 19a, via which the exhaust gas is discharged. The bypass duct 29 opens into a second exhaust gas nozzle 29 a, which surrounds the exhaust gas nozzle 19 ain an annular manner. In principle, the bypass duct 29 can also alternatively open into a mixer, via which the part of the exhaust gas flow conducted into the bypass duct 29 is mixed with the exhaust gas flow of the turbine exhaust gas duct 19.At the inlet 31, a flap, not shown, can also be provided, by means of which the bypass duct 29 can be closed and the mass flow which is conducted through the bypass duct 29 can be adjusted.List of reference characters1 Aircraft gas turbine 3 Housing 5 Fan 7 Air duct 9 Compressor region 11 Combustion chamber 13 High-pressure turbine section 15 Low-pressure turbine section 17 Turbine outlet housing 19 Turbine exhaust gas duct 19 a Gas nozzle 21 Engine axis 23 Outer turbine exhaust gas duct wall 25 Inner turbine exhaust gas duct wall 27 Outlet cone 29 Bypass duct 29 a Gas nozzle 31 Inlet 33 Splitter 35 Support struts 37 Heat exchanger 37 a Einzel tubes 39 Diffuser 41 Feed device 43 Feed line

Claims

Aircraft gas turbine (1) having a high-pressure turbine section (13) and a low-pressure turbine section (15) and having a turbine exhaust duct (19) which is annularly formed about an engine axis (21) and is formed by a turbine outlet housing (17) and is arranged downstream of the low-pressure turbine section (15), wherein a turbine exhaust housing forms an outer turbine exhaust duct wall (23) and an inner turbine exhaust duct wall (25) which delimit the annular turbine exhaust duct (19), and also having an annular bypass duct (29) which is arranged downstream of the low-pressure turbine section (15) outside of the outer turbine exhaust duct wall (23) and has an inlet (31) which is formed on the outer turbine exhaust duct wall (23), wherein part of the exhaust gas stream can be directed into the bypass duct (29), and a heat exchanger (37) for preconditioning fuel by means of the part of the exhaust gas flow conducted into the bypass duct (29), wherein the heat exchanger (37) is arranged in or on the bypass duct (29), characterized in that a feed device for compressed bleed air is arranged on the bypass duct (29).Aircraft gas turbine according to Claim 1, characterized in that a flow throttling device for slowing down the flow of the part of the exhaust gas stream conducted into the bypass duct (29) is formed in the bypass duct (29) upstream of the heat exchanger (37).Aircraft gas turbine according to Claim 2, characterized in that the flow throttling device is designed as a diffuser (39).Aircraft gas turbine according to one of Claims 1 to 3, characterized in that the heat exchanger (37) is designed as a shell-and-tube heat exchanger having a multiplicity of individual tubes (37a), it being possible for the liquid or gaseous fuel to be conducted through the individual tubes (37a).Aircraft gas turbine according to Claim 4, characterized in that the individual tubes (37a) have an outer diameter D, where the following applies: 3 mm ≤ D ≤ 12 mm.Aircraft gas turbine according to Claim 4 or 5, characterized in that the heat exchanger (37) has a cross-countercurrent arrangement.Aircraft gas turbine according to one of Claims 4 to 6, characterized in that the individual tubes (37a) are arranged helically around the engine axis (21).Aircraft gas turbine according to claim 7, characterised in that the individual tubes (37a) are each bent with a constant radius to the engine axis (21).Aircraft gas turbine according to one of Claims 1 to 8, characterized in that support struts (35) are arranged in the bypass duct (29).Aircraft gas turbine according to one of the preceding claims, characterized in that the supporting struts (35) are arranged in the flow throttling deviceAircraft gas turbine according to one of Claims 1 to 10, characterized in that the feed device for compressed bleed air has lines arranged in the supporting struts (35) and bleed air outlets arranged on the supporting struts (35).Aircraft gas turbine according to one of Claims 1 to 11, characterized in that the turbine exhaust duct (19) opens into an exhaust gas nozzle (19a), and the bypass duct (29) opens into a second exhaust gas nozzle (29a) surrounding the exhaust gas nozzle (19a).Aircraft gas turbine according to one of Claims 1 to 11, characterized in that the turbine exhaust duct (19) opens into an exhaust gas nozzle (19a), and the bypass duct (29) opens into a mixer which feeds the part of the exhaust gas stream conducted into the bypass duct (29) to the exhaust gas stream in the turbine exhaust duct (19).Aircraft gas turbine according to one of Claims 1 to 13, characterized in that a splitter (33) formed by the outer turbine exhaust duct wall (23) is arranged at the inlet (31).Aircraft gas turbine according to one of Claims 1 to 14, characterized in that a flap is arranged at the inlet (31).

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