System and method for cooling an aircraft turbine engine

Abstract

The invention relates to a system (S) for cooling a gas turbine engine (3), which system comprises: a first cooling circuit (CD1) configured to inject, into the gas turbine (33), a first air flow (D1) from the compressor (31), a second cooling circuit (CD2) configured to inject, into the gas turbine (33), a second air flow (D2) from the compressor (31), a loop (CF) for circulating a heat transfer fluid (F), a first heat exchanger (61) configured to cool the first air flow (D1) during circulation of the heat transfer fluid (F), a second heat exchanger (62) configured to cool the second air flow (D2) during circulation of the heat transfer fluid (F), and a third heat exchanger (63) configured to cool the heat transfer fluid (F) from the circulation of a cryogenic fuel flow (Q), the heat transfer fluid (F) circulating in the second heat exchanger (62), the first heat exchanger (61) and the third heat exchanger (63) successively.

Description

System and method for cooling an aircraft turbomachine

[0001] The present invention relates to the field of aircraft turbomachines and more particularly to the cooling of a gas turbine of an aircraft turbomachine.

[0002] Climate change is a major concern for many legislative and regulatory bodies around the world. Indeed, various carbon emission restrictions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies to both new aircraft types and those already in operation, requiring the implementation of technological solutions to ensure their compliance with current regulations. Civil aviation has been mobilizing for several years now to contribute to the fight against climate change.

[0003] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

[0004] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes and minimizing greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0005] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to ensure propulsion, and, as an essential complement to technological progress, aeronautical biofuels.

[0006] To this end, the invention is the result of technological research aimed at significantly improving the performance of aircraft and, in this sense, contributes to reducing the environmental impact of aircraft. For this purpose, the invention relates to the field of aircraft comprising a turbomachine powered by a fuel (more generally called combustible) stored in a cryogenic tank.

[0007] In a known manner, with reference to, an aircraft comprises a propulsion assembly 1 to enable its movement. The propulsion assembly 1 extends longitudinally along an axis X and is configured to participate in the propulsion of the aircraft by accelerating an air flow A which circulates from upstream to downstream. In this document, the terms "upstream" and "downstream" are understood with reference to the longitudinal axis X which extends from upstream to downstream. For this, the propulsion assembly 1 comprises a fan 2 configured to generate the air flow A to power a gas turbomachine 3 and enable the propulsion of the aircraft.

[0008] As is known, the gas turbomachine 3 comprises successively along the longitudinal axis X, at least one compressor 31, a combustion chamber 32 and at least one gas turbine 33. The compressor 31 is configured to receive the air flow A from the fan 2 and compress it so as to supply the combustion chamber 32. An exhaust air flow AE, from the combustion, in the combustion chamber 32, between the fuel and the compressed air flow, drives the gas turbine 33 in rotation which drives the compressor 31 with the fan 2.

[0009] In practice, at the outlet of the combustion chamber 32, the exhaust air flow AE has a very high temperature, generally of the order of 1700°C (2000K). Such a temperature requires cooling of the gas turbine 33 which receives the exhaust air flow AE.

[0010] To cool the gas turbine 33, it is known to provide a cooling circuit CD which takes a portion of the compressed air flow A, hereinafter “cooling air flow AR”, from the compressor 31 and injects it into the gas turbine 33. The cooling air flow AR, coming from the compressor 31, has a lower temperature than that of the exhaust air flow AE, which ensures efficient cooling.

[0011] With reference to the, to cool the gas turbine 33 more efficiently, it is known to provide a heat exchanger 6 in the cooling circuit CD in order to provide frigories to the cooling air flow AR. It has been proposed to circulate in the heat exchanger 6 a flow of fuel Q, in particular hydrogen, which is stored at a temperature of the order of -253 to -251°C (20 to 22 Kelvins) in a cryogenic tank R of the aircraft. The fuel Q then represents a significant cold source capable of providing frigories to the cooling air flow AR.

[0012] In practice, such an architecture may present risks of damage to the gas turbomachine 2 in the event of a leak in the heat exchanger 6. Fuel could thus be conducted by the cooling circuit near the gas turbine which is a hot zone of the gas turbomachine 2.

[0013] The invention thus aims to eliminate at least some of these drawbacks by proposing a simple and efficient gas turbine cooling system while guaranteeing optimal safety in the gas turbomachine. PRESENTATION OF THE INVENTION

[0014] The invention relates to a cooling system for an aircraft gas turbine engine, the gas turbine engine being configured to be powered by fuel stored in at least one cryogenic tank and comprising successively along a longitudinal axis at least one compressor, a combustion chamber and at least one gas turbine, a fuel flow circulating in a fuel circuit mounted between the cryogenic tank and the combustion chamber of the gas turbine engine, the cooling system comprising: a first cooling circuit configured to take a first air flow from the compressor and to inject it into the gas turbine, a second cooling circuit configured to take a second air flow from the compressor and to inject it into the gas turbine, at least one circulation loop for a heat transfer fluid, a first heat exchanger,mounted in the first cooling circuit and in the circulation loop, configured to cool the first air flow during the circulation of the heat transfer fluid,a second heat exchanger, mounted in the second cooling circuit and in the circulation loop, configured to cool the second air flow during the circulation of the heat transfer fluid, anda third heat exchanger, mounted in the circulation loop, and configured to cool the heat transfer fluid from the circulation of the fuel flow,the circulation loop being configured to circulate the heat transfer fluid successively in the second heat exchanger the first heat exchanger and the third heat exchanger.,

[0015] The cooling system according to the invention advantageously allows the gas turbine to be cooled efficiently and safely using air flows taken from the compressor. The heat transfer fluid loop eliminates any risk of mixing between air and fuel, thus ensuring a high level of safety.

[0016] In addition, two cooling circuits ensure that the gas turbine is cooled even if one of them fails, for example. Modulation is also possible to achieve a compromise between efficiency and the desired and optimal cooling.

[0017] The air for cooling the gas turbine is advantageously cooled efficiently, via the heat transfer fluid loop, using fuel from a cryogenic tank, which represents a significant source of cold. In this document, the term "fuel" refers to a fuel intended to be burned in the combustion chamber of a gas turbine engine.

[0018] A series circuit allows the first hot air flow from the compressor to be cooled, followed by the second air flow, while heating the heat transfer fluid, which can then advantageously heat the fuel flow. By heating the fuel flow, the heat transfer fluid is efficiently cooled before returning to the loop and cooling the air flow. Cascade cooling allows the temperature difference to be reduced quickly. In addition, since the fuel flow is heated by the heat transfer fluid, this eliminates the need for an independent conditioning system, thus limiting the aircraft's size and weight, thereby limiting greenhouse gas emissions and the impact on the environment.

[0019] Preferably, the compressor of the gas turbomachine comprising a low pressure stage and a high pressure stage, configured to compress the incoming air flow to supply the combustion chamber, the first air flow is a low pressure air flow from the low pressure stage and the second air flow is a high pressure air flow from the high pressure stage. Such an embodiment advantageously makes it possible to first cool the air flow from the high pressure compressor which is supercompressed and therefore hotter, then the air flow from the lower pressure compressor which has lower cooling requirements.

[0020] In one embodiment, each cooling circuit comprises a first circulation duct and a second circulation duct mounted in parallel respectively in the first cooling circuit and in the second cooling circuit. Each cooling circuit thus comprises redundancy of the circulation ducts, which advantageously makes it possible to ensure the circulation of the first air flow and the second air flow in the cooling circuits, making it possible to guarantee effective cooling of the turbine, even in the event of a fault in one of the circulation ducts, and therefore a high level of safety.

[0021] In one embodiment: the first cooling circuit comprises a first controllable distribution member configured to regulate the flow rate of the first air flow between the first circulation duct and the second circulation duct, the second cooling circuit comprises a second controllable distribution member configured to regulate the flow rate of the second air flow between the first circulation duct and the second circulation duct, and the cooling system comprises a computer configured to authorize circulation of the first air flow only in one of the first circulation duct or the second circulation duct of the first cooling circuit and circulation of the second air flow only in one of the first circulation duct or the second circulation duct of the second cooling circuit.

[0022] The two air flows from the compressor are thus not allowed to circulate simultaneously in the two circulation ducts of the same cooling circuit, which makes it possible to provide redundancy of the cooling circuits, thus making it possible to secure the cooling of the gas turbine even in the event of damage to one of the circulation ducts, for example. The gas turbine can also advantageously be cooled by an air flow from the compressor without penalizing the gas turbomachine by taking a limited fraction of the compressed air flow. The distribution member in fact advantageously makes it possible to regulate the flow rate in each cooling circuit.Thanks to the distribution components, it is also possible to allow, for example, operation in degraded mode in the event of damage to one of the elements of the cooling circuits while ensuring that the turbine is cooled to a minimum, so as to guarantee high levels of safety.

[0023] In one embodiment, the cooling system comprising at least a first member for measuring the flow rate of the first air flow in each circulation duct of the first cooling circuit and a second member for measuring the flow rate of the second air flow in each circulation duct of the second cooling circuit, the computer is configured to prohibit the circulation of the first air flow, respectively the second air flow, in one of the circulation ducts when the first flow rate measuring member, respectively the second flow rate measuring member, measures a flow rate lower than a predetermined flow rate threshold in said circulation duct. Safety in the cooling system is thus guaranteed by allowing the detection of a leak without the circulation of a cooled air flow being affected. The gas turbine can continue to be cooled thanks to the cooling circuit which does not have a fault.In fact, the computer is configured to, when the first distribution member prohibits the circulation of the first air flow in the first circulation duct following the measurement of an excessively high flow rate by the first measuring member, authorize the circulation of the first air flow in the second circulation duct. The same applies in the second cooling circuit.

[0024] Alternatively, the cooling system comprises at least a first member for measuring the temperature of the first air flow in each circulation duct of the first cooling circuit and a second member for measuring the temperature of the second air flow in each circulation duct of the second cooling circuit. The computer is configured to prohibit the circulation of the first air flow, respectively the second air flow, in one of the circulation ducts when the first temperature measuring member, respectively the second temperature measuring member, measures a temperature greater than a predetermined temperature threshold. Indeed, the more the pressure increases, the more the temperature increases.The computer advantageously controls the closure of the circulation pipe of the cooling circuit whose temperature is too high, indicating a fault in said circulation pipe, which allows the cooling circuit to continue to operate safely.

[0025] In one embodiment, the gas turbine engine comprising an oil circuit configured to circulate a flow of lubricating oil through the compressor so as to lubricate and cool it, the cooling system comprises a fourth heat exchanger, mounted in the heat transfer fluid circulation loop and in the oil circuit, configured to cool the flow of oil during the circulation of the heat transfer fluid. The single heat transfer fluid circulation loop can thus advantageously both cool the gas turbine, heat the fuel flow for injection and cool the lubricating oil.

[0026] In one embodiment, the gas turbomachine being configured to be powered by a first fuel flow circulating in a first fuel circuit and by a second fuel flow circulating in a second fuel circuit stored in at least one cryogenic tank, the cooling system comprising at least a first circulation loop of a first heat transfer fluid and a second circulation loop of a second heat transfer fluid, the first heat exchanger, the second heat exchanger and the third heat exchanger being mounted in the first circulation loop, the cooling system comprises: a fifth heat exchanger mounted in the second circulation loop and in an exhaust flow circuit, configured to heat the second heat transfer fluid during the circulation of the exhaust flow,anda sixth heat exchanger mounted in the second circulation loop and configured to heat the second heat transfer fluid during circulation of the second fuel flow.,

[0027] Using two separate fuel circuits allows only a fraction of the fuel required in the combustion chamber to be used to cool the air flows via the heat transfer fluid. It is also possible to adapt the flow rate of each fuel flow to promote either gas turbine cooling or fuel heating.

[0028] In one embodiment, the gas turbomachine comprising an oil circuit configured to circulate a flow of lubricating oil via the compressor so as to lubricate and cool it, the flow of oil is configured to participate in the heating of at least one of the fuel flows while being cooled by said fuel after its passage into the combustion chamber.

[0029] The invention also relates to an aircraft comprising at least: a cryogenic tank, a gas turbomachine comprising at least one compressor, a combustion chamber and a gas turbine, a fuel circuit mounted between the cryogenic tank and the combustion chamber of the gas turbomachine, a fuel flow circulating in the fuel circuit, and at least one system for cooling the gas turbomachine as described previously.

[0030] Finally, the invention relates to a method for cooling an aircraft gas turbine engine, the gas turbine engine being configured to be powered by fuel stored in a cryogenic tank and comprising successively along a longitudinal axis a compressor, a combustion chamber and a gas turbine, a fuel flow circulating in a fuel circuit mounted between the cryogenic tank and the combustion chamber of the gas turbine engine, the cooling method comprising the steps of:Taking a first air flow and a second air flow from the compressor of the gas turbine engine,Respectively cooling the second air flow in a second heat exchanger and / or the first air flow in a first heat exchanger during the circulation of the heat transfer fluid,the heat transfer fluid having been initially cooled in a third heat exchanger during the circulation of the fuel flow from a cryogenic tank, andInjecting, into the gas turbine at least the first air flow and / or the second flow, PRESENTATION OF FIGURES

[0031] The invention will be better understood upon reading the following description, given by way of example, and referring to the following figures, given by way of non-limiting examples, in which identical references are given to similar objects.

[0032] This is a schematic representation of a cooling system according to the prior art.

[0033] This is a schematic representation of a cooling system according to a first embodiment of the invention.

[0034] This is a schematic representation of a cooling system according to a second embodiment of the invention.

[0035] This is a schematic representation of a cooling system according to a third embodiment of the invention.

[0036] This is a schematic representation of a cooling system according to a fourth embodiment of the invention.

[0037] It should be noted that the figures set out the invention in detail to implement the invention, said figures can of course be used to better define the invention where appropriate. DETAILED DESCRIPTION OF THE INVENTION

[0038] With reference to the, there is shown a propulsion assembly 1 configured to allow the movement of an aircraft. The propulsion assembly 1 extends longitudinally along an axis X and is configured to participate in the propulsion of the aircraft by accelerating an air flow A circulating from upstream to downstream. The terms "upstream" and "downstream" refer to the axis X which extends from upstream to downstream of the propulsion assembly 1, as shown in the.

[0039] The propulsion assembly 1 comprises a fan 2 configured to rotate around the longitudinal axis X, so as to generate an incoming air flow A to supply a gas turbomachine 3 and generate thrust according to a bypass ratio.

[0040] For this, the gas turbomachine 3 comprises, successively from upstream to downstream, a compressor 31, a combustion chamber 32 and a gas turbine 33. The compressor 31 is configured to receive the incoming air flow A from the fan 2 and compress it, so as to supply the combustion chamber 32. In the combustion chamber 32, the combustion between the compressed air flow A and a fuel flow Q generates an exhaust air flow AE which drives the gas turbine 33 in rotation, thus driving the compressor 31 and the fan 2.

[0041] In particular, in this example, with reference to the, the compressor 31 comprises a low pressure stage 31B and a high pressure stage 31H. The low pressure stage 31B and the high pressure stage 31H make it possible to gradually compress the air flow A to reach a predetermined pressure ratio and allow the introduction of a sufficiently compressed air flow A into the combustion chamber 32. Such a gas turbomachine 3 is known to those skilled in the art and its operation will not be described in more detail in this document.

[0042] According to one aspect of the invention, the gas turbine engine 3 is configured to be powered by fuel Q stored in a cryogenic tank R. In particular, in the cryogenic tank R, the fuel Q is stored at a temperature of the order of -253 to -251°C (20 to 22 Kelvins). At this temperature, the flow of fuel Q is liquid. In this example, the fuel Q is hydrogen but the invention applies to other types of fuel, for example, liquid methane or liquefied natural gas.

[0043] As shown in the, a fuel circuit CQ connects the cryogenic tank R to the combustion chamber 32 of the gas turbomachine 3, so as to supply it with a flow of fuel Q circulating in the fuel circuit CQ.

[0044] Also shown is a cooling system S of the gas turbomachine 3 according to one embodiment of the invention. In particular, the cooling system S is configured to cool the gas turbine 33 of the gas turbomachine 3 from an air flow taken from the compressor 31.

[0045] For this, the cooling system S comprises a first cooling circuit CD1 and a second cooling circuit CD2. The first cooling circuit CD1 is configured to take a first air flow D1 from the compressor 31 and to inject it into the gas turbine 33. Similarly, the second cooling circuit CD2 is configured to take a second air flow D2, distinct from the first air flow D1, from the compressor 31 and to inject it into the gas turbine 33.

[0046] In this example in which the compressor 31 comprises a low pressure stage 31B and a high pressure stage 31H, the first air flow D1 is a low pressure air flow from the low pressure stage 31B and the second air flow D2 is a high pressure air flow from the high pressure stage 31H, as shown in the.

[0047] A cooling system S is shown comprising two cooling circuits CD1, CD2 mounted between the compressor 31 and the gas turbine 33, however, it goes without saying that the cooling system S could comprise a different number of cooling circuits, in particular a number greater than two.

[0048] According to one aspect of the invention, with reference to figures 2 and 3, the cooling system S comprises a circulation loop CF of a heat transfer fluid F. Preferably, the circulation loop CF is a closed loop in which the heat transfer fluid F, for example nitrogen, stored in a heat transfer fluid tank 4, circulates. The use of a heat transfer fluid F makes it possible to circulate an intermediate fluid between the fuel flow Q and the air flows D1, D2, as will be described in more detail below, advantageously making it possible to avoid any risk of contact between the fuel flow Q and the air flows D1, D2.

[0049] To allow the circulation of the heat transfer fluid F, the cooling system S comprises a pump 5, preferably with non-contact drive, for example a magnetic drive pump. Preferably, the pump 5 is configured to circulate the heat transfer fluid F at a predefined pressure and flow rate, making it possible, for example, to impose a pressure in the circulation loop CF greater than the pressure of the fuel flow Q in the fuel circuit CQ. For example, the pressure in the circulation loop CF is greater than or equal to 3.4x10 6 Pa (34 bar) and the temperature is greater than or equal to -147 °C (126 K), allowing the circulation of nitrogen in a supercritical state, which allows a stable flow of the heat transfer fluid F in the circulation loop CF.

[0050] To enable cooling of the first air flow D1 and the second air flow D2 before they are introduced into the gas turbine 33, as shown in FIGS. 2 and 3, the cooling system S comprises a first heat exchanger 61, a second heat exchanger 62 and a third heat exchanger 63 mounted in the circulation loop CF of the heat transfer fluid F.

[0051] In particular, the first heat exchanger 61 is mounted both in the first cooling circuit CD1 and in the circulation loop CF and is configured to cool the first air flow D1 during the circulation of the heat transfer fluid F. In other words, the first flow D1 is configured to be cooled by frigories transferred by the previously cooled heat transfer fluid F. Similarly, the second heat exchanger 62 is mounted both in the second cooling circuit CD2 and in the circulation loop CF and is configured to cool the second air flow D2 during the circulation of the heat transfer fluid F.

[0052] The third heat exchanger 63 is mounted both in the circulation loop CF and in the fuel circuit CQ and is configured to cool the heat transfer fluid F from the circulation of the fuel flow Q from the cryogenic tank R. Indeed, the fuel flow Q represents a significant cold source which makes it possible to cool the heat transfer fluid F allowing the latter to subsequently cool the first air flow D1 and the second air flow D2, as will be described in more detail later. In parallel, the fuel flow Q can advantageously be heated the air flows D1, D2 before being injected into the combustion chamber 32 of the gas turbomachine 3.

[0053] According to one aspect of the invention, the circulation loop CF is configured to circulate the heat transfer fluid F successively in the second heat exchanger 62, the first heat exchanger 61 and the third heat exchanger 63. The circulation loop CF thus makes it possible to successively cool the second air flow D2 and the first air flow D1 from the compressor 31. Thus, in the example in which the first air flow D1 comes from the low pressure stage 31B of the compressor 31 and the second air flow D2 comes from the high pressure stage 31H, the heat transfer fluid F is configured to firstly cool the second air flow D2 from the high pressure stage 31H which is hotter than the first air flow D1 from the low pressure stage 31B. Indeed, the higher the pressure, the higher the temperature of the air flow.In other words, the second, warmer air flow D2 at the outlet of the compressor 31 can be cooled more efficiently by the cold heat transfer fluid F. Incidentally, the first air flow D1, which is less hot than the second air flow D2, requires fewer frigories to be cooled and is therefore cooled by a less cold heat transfer fluid F since it has already transferred frigories to the second air flow D2.

[0054] In an embodiment shown in the, the first cooling circuit CD1 comprises a first circulation duct CD11 and a second circulation duct CD12 mounted in parallel in the first cooling circuit CD1 between the compressor 31 and the turbine 33. Similarly, the second cooling circuit CD2 comprises a first circulation duct CD21 and a second circulation duct CD22 mounted in parallel in the second cooling circuit CD2 between the compressor 31 and the turbine 33 (in the embodiment shown in the, each cooling circuit CD1, CD2 comprises a single circulation duct). The first air flow D1 is configured to circulate either in the first circulation duct CD11 or in the second circulation duct CD12 of the first cooling circuit CD1.The second air flow D2 is configured to circulate either in the first circulation duct CD21 or in the second circulation duct CD22 of the second cooling circuit CD2. In other words, each air flow D1 preferably does not circulate simultaneously in the two circulation ducts of the cooling circuit CD1, CD2.

[0055] For this, still with reference to the, the cooling system S preferably comprises a first controllable distribution member 71 mounted in the first cooling circuit CD1 and a second controllable distribution member 72 mounted in the second cooling circuit CD2. The first distribution member 71 is configured to regulate the flow rate of the first air flow D1 in each of the two circulation ducts CD11, CD12. Similarly, the second distribution member 72 is configured to regulate the flow rate of the second air flow D2 in each of the two circulation ducts CD21, CD22.Preferably, each distribution member 71, 72 is mounted respectively directly at the inlet of the first cooling circuit CD1 and of the second cooling circuit CD2, so as to limit the circulation of the first air flow D1, respectively of the second air flow D2, in one or other of the circulation ducts CD11, CD12, CD21, CD22 and therefore in each heat exchanger 61, 62 when necessary. In this example, each distribution member 71, 72 is in the form of a control valve.

[0056] The first distribution member 71 and the second distribution member 72 are preferably controllable by a computer 9. Preferably, the computer 9 is configured to allow circulation of the first air flow D1 only in one of the two circulation ducts CD11, CD12 of the first cooling circuit CD1 and circulation of the second air flow D2 only in one of the two circulation ducts CD21, CD22 of the second cooling circuit CD2. Thus, a limited flow of air is taken from the compressor 31 to be used in the cooling system S, which makes it possible not to penalize the performance of the gas turbomachine 3. The cooling system S thus advantageously has redundancy of the circulation ducts in each cooling circuit CD1, CD2.

[0057] Still with reference to, in order to control the state of each cooling circuit CD1, CD2, the cooling system S preferably comprises a first measuring member 81 of the flow rate of the first air flow D1 in each circulation duct CD11, CD12, CD21, CD22 of the first cooling circuit CD1 and a second measuring member 82 of the flow rate of the second air flow D2 in each circulation duct CD11, CD12, CD21, CD22 of the second cooling circuit CD2. Each measuring member 81, 82 is for example in the form of a flow meter. In this example, the computer 9 is configured to control the first distribution member 71 and the second distribution member 72, as a function of the measurement of the flow rate of each circulation duct by each measuring member 81, 82 of the flow rate.For example, when the flow rate in one of the circulation ducts CD11, CD12, CD21, CD22 is lower than an expected flow rate, signifying for example the presence of a leak, the computer 9 is configured to stop the circulation of the air flow D1, D2 in the circulation duct concerned, so as to limit any risk of damage to the gas turbomachine 3.

[0058] In one embodiment, the cooling system S comprises a measuring member 81, 82 for the flow rate upstream and downstream of each heat exchanger 61, 62, making it possible to detect a leak in the heat exchanger 61, 62. The terms “upstream” and “downstream” are understood here to mean the flow of the heat transfer fluid F in the circulation loop CF, in which the heat transfer fluid F circulates between an outlet and an inlet of the tank 4, successively via the second heat exchanger 62, the first heat exchanger 61 and the third heat exchanger 63.

[0059] Similarly, in an exemplary embodiment, the cooling system S comprises one or more measuring members (not shown) upstream and / or downstream of the third heat exchanger 63, so as to detect, for example, a leak in the fuel circuit CQ. In this embodiment, the cooling system S comprises a distribution member 73, mounted in the fuel circuit CQ, preferably at the outlet of the cryogenic tank R, and configured to stop the circulation of the flow of fuel Q in the fuel circuit CQ.

[0060] In an alternative embodiment, the cooling system S comprises one or more temperature measuring members (not shown). The computer 9 is then configured to detect a rise in temperature in each circulation duct CD11, CD12, CD21, CD22 of each cooling circuit CD1, CD2, so as to order the stopping of the circulation of the air flow D1, D2 in one of the circulation ducts CD11, CD12, CD21, CD22 which would have a fault. It goes without saying that the cooling system S could just as well comprise both one or more flow rate measuring members 81, 82 and one or more temperature measuring members on each cooling circuit CD1, CD2 allowing double monitoring to ensure the detection of a fault and thus guarantee the durability of the cooling system S and the gas turbomachine 3.

[0061] With reference to the, the propulsion assembly 1 also comprises an oil circuit CH configured to circulate a flow of lubricating oil H in the compressor 31 so as to lubricate and cool it. The oil circuit CH comprises an oil storage tank 9 and is configured to allow the circulation of the flow of oil H in a closed loop. Such an oil circuit CH is known to those skilled in the art and its operation will not be described in more detail in this document.

[0062] In an exemplary embodiment, still with reference to the, the cooling system S comprises a fourth heat exchanger 64, mounted in the circulation loop CF of the heat transfer fluid F and in the oil circuit CH. The fourth heat exchanger 64 is configured to cool the oil flow H during the circulation of the heat transfer fluid F. In this embodiment, the circulation loop CF is configured to circulate the heat transfer fluid F successively in the fourth heat exchanger 64, the second heat exchanger 62, the first heat exchanger 61 and the third heat exchanger 63. Thus, the oil flow H is efficiently cooled by the heat transfer fluid F directly at the outlet of the pump 5.

[0063] With reference to the, in an alternative embodiment, the gas turbomachine 3 is configured to be powered by two separate fuel flows Q1, Q2. In particular, the gas turbomachine 3 is configured to be powered by a first fuel flow Q1 circulating in a first fuel circuit CQ1 and by a second fuel flow Q2 circulating in a second fuel circuit CQ2. In one example, the first fuel flow Q1 and the second fuel flow Q2 are stored in the same cryogenic tank R. The first fuel circuit CQ1 and the second fuel circuit CQ2 are each mounted between the cryogenic tank R and the combustion chamber 32 of the gas turbomachine 3 to respectively power it with the first fuel flow Q1 and / or the second fuel flow Q2.Alternatively, as shown in the, the first fuel flow Q1 and the second fuel flow Q2 are stored in two separate cryogenic tanks R1, R2 mounted in the aircraft. Each fuel circuit CQ1, CQ2 is mounted respectively between each cryogenic tank R1, R2 and the combustion chamber 32 of the gas turbine engine 3.

[0064] Advantageously, the use of two fuel circuits CQ1, CQ2 makes it possible to use only a fraction of the fuel Q required by the turbomachine 3 to ensure the cooling of the gas turbine 33.

[0065] The cooling system S then comprises in this example, a first circulation loop CF1 of a first heat transfer fluid F1 and a second circulation loop CF2 of a second heat transfer fluid F2, the first heat exchanger 61, the second heat exchanger 62 and the third heat exchanger 63 described previously being mounted in the first circulation loop CF1. In other words, the first heat transfer fluid F1 is in this example cooled by the first fuel flow Q1 circulating in the first fuel circuit CQ1.

[0066] In this embodiment, the cooling system S comprises a first pump 51 for allowing the circulation of the first heat transfer fluid F1 (initially stored in a first cryogenic tank 41) in the first circulation loop CF1 and a second pump 52 for allowing the circulation of the second heat transfer fluid F2 (initially stored in a second cryogenic tank 42) in the second circulation loop CF2.

[0067] The cooling system S also comprises a fifth heat exchanger 65 and a sixth heat exchanger 66 mounted in the second circulation loop CF2. More specifically, the fifth heat exchanger 65 is mounted both in the second circulation loop CF2 and in a circuit of the exhaust flow AE of the gas turbine 33 and is configured to heat the second heat transfer fluid F2 during the circulation of the exhaust flow AE. The sixth heat exchanger 66 is configured to heat the second heat transfer fluid F2 during the circulation of the second fuel flow Q2.

[0068] Such an embodiment makes it possible to cool the air flows D1, D2 from the compressor 31 by means of the first fuel flow Q1 via the first circulation loop CF1 of the first heat transfer fluid F1, while efficiently heating the second fuel flow Q2 by means of the exhaust flow AE from the gas turbine 33 via the second circulation loop CF2 of the second heat transfer fluid F2.

[0069] The use of two fuel circuits CQ1, CQ2 with asymmetrical architecture makes it possible to increase robustness in order, if necessary, to adapt the circulation flow rates to promote cooling of the gas turbine 33 or heating of the fuel Q.

[0070] This document shows an alternative embodiment comprising an oil circuit H and another alternative embodiment comprising two fuel circuits CQ1, CQ2 and two circulation loops CF1, CF2 of two heat transfer fluids F1, F2, however, the cooling system S could alternatively comprise both two circulation loops CF1, CF2 of two heat transfer fluids F1, F2 for heating two fuel flows Q1, Q2, an oil circuit CH and a fourth heat exchanger 64, mounted both on the oil circuit CH and on one of the circulation loops CF1, CF2 of heat transfer fluids F1, F2 for cooling the oil flow H during the circulation of one of the heat transfer fluids F1, F2. The oil circuit CH could also alternatively comprise two heat exchangers to cool the oil flow H successively by means of the two circulation loops CF1, CF2 of heat transfer fluids F1, F2.

[0071] A method for cooling the gas turbomachine 3 will now be described with reference to, according to an embodiment of the invention. A flow of fuel Q circulates in the fuel circuit CQ between the cryogenic tank R and the combustion chamber 32 of the gas turbomachine 3. The pump 5 makes it possible to circulate the heat transfer fluid F in the circulation loop CF.

[0072] In a preliminary step E0, the fuel flow Q from the cryogenic tank R passes through the third heat exchanger 63 in which it cools the heat transfer fluid F circulating in the circulation loop CF.

[0073] The method then comprises a first step E1 of sampling the first air flow D1 and the second air flow D2 in the compressor 31 of the gas turbomachine 3, so as to circulate them respectively in the first cooling circuit CD1 and in the second cooling circuit CD2. In this example, a distribution member 71, 72 being mounted on each cooling circuit CD1, CD2, the first distribution member 71 authorizes the circulation of the first air flow D1 in a first circulation duct CD11 of the first cooling circuit CD1 and prohibits the circulation of the first air flow D1 in a second circulation duct CD12 of the first cooling circuit CD1.Likewise, the second distribution member 72 authorizes the circulation of the second air flow D2 in a first circulation duct CD21 of the second cooling circuit CD2 and prohibits the circulation of the second air flow D2 in the second circulation duct CD22 of the second cooling circuit CD2.

[0074] In a step E2, the heat transfer fluid F, which has been cooled by the fuel flow Q in the third heat exchanger 63, passes through the second heat exchanger 62 then the first heat exchanger 61. The first air flow D1 which circulates in the first cooling circuit CD1 is cooled in the first heat exchanger 61 by the circulation of the heat transfer fluid F.

[0075] At the outlet of the first heat exchanger 61, the first air flow D1 circulates in the first cooling circuit CD1 and is injected into the gas turbine 33 to cool it, in a step E3.

[0076] The heat transfer fluid F then passes again through the third heat exchanger 63, in a step E4, in which it exchanges calories with the fuel flow Q. In particular, in this step, the heat transfer fluid F is cooled by cryogenic fuel Q so that it can then return to the tank 4 before cooling the first air flow D1 in the first heat exchanger 61, while the fuel flow Q is reheated, so that it can be injected into the combustion chamber 32 of the gas turbomachine 3.

[0077] In one embodiment, in which the cooling system S comprises a flow rate measuring member 81 mounted on the first cooling circuit CD1, the method comprises a step E4 of measuring the flow rate of the first air flow D1 in each circulation duct CD11, CD12 of the first cooling circuit CD1 and comparing, by the computer 9, the measured flow rate with a predetermined flow rate threshold. When the measured flow rate is greater than the predetermined flow rate threshold, for example in the event of a leak in the first circulation duct CD11, the computer 9 controls, in a step E5, the first distribution member 71, so as to prohibit the circulation of the first air flow D1 in the first circulation duct CD11 of the first cooling circuit CD1 and to authorize the circulation of the first air flow D1 in the second circulation duct CD12 of the first cooling circuit CD1.The circulation of the first air flow D1 is thus advantageously ensured thanks to the redundancy of the circulation ducts, which makes it possible to ensure effective cooling of the gas turbine 33.

[0078] In a similar manner to steps E2 and E3, the heat transfer fluid F circulating in the circulation loop CF passes through the second heat exchanger 62 in which it transfers frigories to the second air flow D2. At the outlet of the second heat exchanger 62, the second air flow D2 circulates in the second cooling circuit CD2 and is injected into the gas turbine 33 to cool it.

Claims

Cooling system (S) for an aircraft gas turbine engine (3), the gas turbine engine (3) being configured to be powered by a fuel (Q) stored in at least one cryogenic tank (R) and comprising successively along a longitudinal axis (X) at least one compressor (31), a combustion chamber (32) and at least one gas turbine (33), a fuel flow (Q) circulating in a fuel circuit (CQ) mounted between the cryogenic tank (R) and the combustion chamber (32) of the gas turbine engine (3), the cooling system (S) comprising: a first cooling circuit (CD1) configured to take a first air flow (D1) from the compressor (31) and to inject it into the gas turbine (33), a second cooling circuit (CD2) configured to take a second air flow (D2) from the compressor (31) and to inject it into the gas turbine (33), at least a circulation loop (CF) of a heat transfer fluid (F),a first heat exchanger (61), mounted in the first cooling circuit (CD1) and in the circulation loop (CF), configured to cool the first air flow (D1) during the circulation of the heat transfer fluid (F), a second heat exchanger (62), mounted in the second cooling circuit (CD2) and in the circulation loop (CF), configured to cool the second air flow (D2) during the circulation of the heat transfer fluid (F), and a third heat exchanger (63), mounted in the circulation loop (CF), and configured to cool the heat transfer fluid (F) from the circulation of the fuel flow (Q), the circulation loop (CF) being configured to circulate the heat transfer fluid (F) successively in the second heat exchanger (62), the first heat exchanger (61) and the third heat exchanger (63)., Cooling system (S) according to claim 1, wherein the compressor (31) of the gas turbomachine (3) comprises a low pressure stage (31B) and a high pressure stage (31H), configured to compress the incoming air flow (A) to supply the combustion chamber (32), the first air flow (D1) is a low pressure air flow from the low pressure stage (31B) and the second air flow (D2) is a high pressure air flow from the high pressure stage (31H). Cooling system (S) according to one of claims 1 to 2, in which each cooling circuit (CD1, CD2) comprises a first circulation duct (CD11, CD21) and a second circulation duct (CD12, CD22) mounted in parallel respectively in the first cooling circuit (CD1) and in the second cooling circuit (CD2). Cooling system (S) according to claim 3, wherein:the first cooling circuit (CD1) comprises a first controllable distribution member (71) configured to regulate the flow rate of the first air flow (D1) between the first circulation duct (CD11) and the second circulation duct (CD12),the second cooling circuit (CD2) comprises a second controllable distribution member (72) configured to regulate the flow rate of the second air flow (D2) between the first circulation duct (CD21) and the second circulation duct (CD22),andthe cooling system (S) comprises a computer (9) configured to authorize circulation of the first air flow (D1) only in one of the first circulation duct (CD11) or the second circulation duct (CD12) of the first cooling circuit (CD1) and circulation of the second air flow (D2) only in one of the first circulation duct (CD21) or the second circulation duct (CD22) of the second cooling circuit (CD2)., Cooling system (S) according to one of claims 1 to 4, wherein, the gas turbomachine (3) comprising an oil circuit (CH) configured to circulate a flow of lubricating oil (H) via the compressor (31) so as to lubricate it, the cooling system (S) comprises a fourth heat exchanger (64), mounted in the circulation loop (CF) of the heat transfer fluid (F) and in the oil circuit (CH), configured to cool the flow of oil (H) during the circulation of the heat transfer fluid (F). Cooling system (S) according to one of claims 1 to 5, the gas turbomachine (3) being configured to be supplied by a first fuel flow (Q) circulating in a first fuel circuit (CQ1) and by a second fuel flow (Q2) circulating in a second fuel circuit (CQ2) stored in at least one cryogenic tank (R), the cooling system (S) comprising at least a first circulation loop (CF1) of a first heat transfer fluid (F1) and a second circulation loop (CF2) of a second heat transfer fluid (F2), the first heat exchanger (61), the second heat exchanger (62) and the third heat exchanger (63) being mounted in the first circulation loop (CF1), the cooling system (S) comprising: a fifth heat exchanger (65) mounted in the second circulation loop (CF2) and in an exhaust flow circuit (AE),configured to heat the second heat transfer fluid (F2) during the circulation of the exhaust flow (AE), and a sixth heat exchanger (66) mounted in the second circulation loop (CF2) and configured to heat the second heat transfer fluid (F2) during the circulation of the second fuel flow (Q2)., Aircraft comprising at least: a cryogenic tank (R), a gas turbomachine (3) comprising at least one compressor (31), a combustion chamber (32) and a gas turbine (33), a fuel circuit (CQ) mounted between the cryogenic tank (R) and the combustion chamber (32) of the gas turbomachine (3), a fuel flow (Q) circulating in the fuel circuit (CQ), and at least one cooling system (S) of the gas turbomachine (3) according to one of claims 1 to 6. A method for cooling an aircraft gas turbine engine (3), the gas turbine engine (3) being configured to be powered by a fuel (Q) stored in a cryogenic tank (R) and comprising successively along a longitudinal axis (X) a compressor (31), a combustion chamber (32) and a gas turbine (33), a fuel flow (Q) circulating in a fuel circuit (CQ) mounted between the cryogenic tank (R) and the combustion chamber (32) of the gas turbine engine (3), the cooling method comprising the steps of:Taking a first air flow (D1) and a second air flow (D2) from the compressor (31) of the gas turbine engine (3),Cooling respectively the second air flow (D2) in a second heat exchanger (62) and / or the first air flow (D1) in a first heat exchanger (61) during the circulation of the heat transfer fluid (F),the heat transfer fluid (F) having been initially cooled in a third heat exchanger (63) during the circulation of the fuel flow (Q) from a cryogenic tank (R), andInjecting, into the gas turbine (33) at least the first air flow (D1) and / or the second cooled air flow (D2).,

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