EP2943674A1

EP2943674A1 - Propellant supply circuit and cooling method

Info

Publication number: EP2943674A1
Application number: EP14703132.2A
Authority: EP
Inventors: Didier Vuillamy; Gérard ROZ
Original assignee: SNECMA Services SA; SNECMA SAS
Current assignee: ArianeGroup SAS
Priority date: 2013-01-11
Filing date: 2014-01-10
Publication date: 2015-11-18
Also published as: WO2014108649A1; FR3000997A1; RU2647353C2; JP2016505764A; RU2015133524A; WO2014108651A3; WO2014108650A1; US20150354505A1; EP2943675B1; JP6352305B2; EP2943675A1; FR3000998A1; RU2015133553A; US10082106B2; WO2014108651A2; JP2016505765A; FR3000997B1; FR3000995B1; US20150337763A1; JP6352306B2

Abstract

The invention relates to the aerospace field and, in particular, that of rocket-engine-propelled vehicles. In particular, the invention relates to a supply circuit (6) for supplying a rocket engine (2) with at least a first liquid propellant, said circuit comprising: at least a first buffer tank (20) for the first liquid propellant, and a first heat exchanger (18) which is built into the buffer tank (20) and can be connected to a cooling circuit (17) of at least one heat source in order to cool the heat source by transferring heat to the first propellant.

Description

ERGOL SUPPLY CIRCUIT AND METHOD OF

COOLING BACKGROUND OF THE INVENTION

The present invention relates to the aerospace field, and in particular that of engines propelled by rocket engines.

In the following description, the terms "upstream" and "downstream" are defined with respect to the normal flow direction of the propellants in the supply circuits of a rocket engine.

One of the main concerns in this area is that of obtaining satisfactory cooling of the heat sources on board. Indeed, to ensure their optimal operation, it may be necessary to maintain the operating temperature of certain heat generating devices, such as fuel cells, but also batteries or electronic circuits, in a relatively narrow range. However, domain-specific constraints can make it very difficult to evacuate the heat generated by these devices. Especially in the space vacuum, the heat evacuation routes are extremely restricted.

Object and summary of the invention

The present description aims to remedy these drawbacks. In particular, it is intended to provide a power supply circuit of a rocket engine in at least a first liquid propellant, which at the same time allows the cooling of at least one heat source.

This object is achieved thanks to the fact that this supply circuit comprises at least one buffer tank for said first liquid propellant and a first heat exchanger, integrated in said buffer tank and connected to a cooling circuit of the at least one a source of heat. Thus, in operation, the heat generated by the heat source can be evacuated, through the cooling circuit and said first heat exchanger, to the liquid propellant in the power supply circuit of the rocket engine. Unlike the regenerative cooling of a rocket engine propellant chamber wall, in which the propulsive chamber is directly cooled by the propellant, this cooling is therefore performed with, as intermediate, a cooling circuit interposed between the heat source and the propellant flowing through the supply circuit, which allows a possible more precise regulation of the temperature of the heat source through a possible regulation of the flow of cooling fluid in the cooling circuit. The integration of the first heat exchanger in a buffer tank of the supply circuit makes it possible to increase the heat power absorbed, even when the supply circuit is closed and said first propellant does not circulate.

The present description also relates to the assembly of said supply circuit and of the heat source equipped with a cooling circuit connected to said first heat exchanger of the supply circuit. The heat source may in particular be a fuel cell. Such a fuel cell can for example be powered by the same propellants as the rocket engine to generate electricity for the embedded systems of a machine powered by the rocket engine. Alternatively, however, other types of on-board heat sources, such as batteries or electronic circuits, can be cooled in the same way.

The present description also relates to a machine comprising a rocket motor with said supply circuit and an onboard device generating heat with a cooling circuit connected to said first heat exchanger of the supply circuit. This device may be, for example, a space launcher stage, a satellite, or any other type of vehicle powered by a liquid propellant rocket engine.

According to a second aspect, said first liquid propellant may in particular be a cryogenic liquid, and in particular liquid hydrogen, thus providing an even more efficient cooling due to its low temperature.

In a third aspect, said supply circuit may comprise a pump upstream of said first heat exchanger, to ensure the circulation of the first propellant. This pump may be, for example, an electric pump or a turbopump. However, alternatively, the power circuit could be configured from in order to ensure the circulation of the first propellant by other means, such as for example by pressurizing an upstream reservoir.

The first ergol heated downstream of the first heat exchanger can be used not only to supply a propellant chamber of the rocket engine or possibly a gas generator or the heat source itself (for example when this heat source is a source of heat. fuel cell), but also, in the gaseous state, to maintain the internal pressure in at least one tank of the first propellant during its emptying through the supply circuit. For this, the supply circuit may comprise a branch opening on an upper part of the tank of the first propellant. The propellant in the gaseous state can thus be reinjected into this reservoir to maintain the internal pressure during its emptying.

According to a fourth aspect, said supply circuit may comprise, downstream of said first heat exchanger, a branch crossing a second heat exchanger. The second heat exchanger can thus ensure the passage to the gaseous state of a flow rate of the first propellant derived through said branch, even when the heat output of said heat generating device is, alone, insufficient for that. This gas flow can thus be used, for example, to maintain the internal pressure of a reservoir supplying the feed circuit with the first propellant during its emptying. The present description therefore also relates to the assembly comprising this supply circuit and a reservoir for said first liquid propellant, connected to this supply circuit upstream of said first heat exchanger, and to said branch downstream of said second heat exchanger.

According to a fifth aspect, said second heat exchanger can be integrated in a tank for a second liquid propellant so as to be able to heat the first liquid propellant by heat transfer of the second liquid propellant. In particular when the second liquid propellant has a boiling temperature substantially higher than the first liquid propellant (for example, when the first liquid propellant is liquid hydrogen and the second liquid propellant is liquid oxygen) this allows not only to ensure the gas phase of the first propellant in the second heat exchanger, but simultaneously cool the second propellant. This cooling of the second propellant makes it possible to avoid cavitation in a pump downstream of this second tank. The present description therefore also relates to an assembly of this supply circuit and a reservoir for a second liquid propellant, containing said second heat exchanger.

Finally, the present description also relates to a method of cooling a heat source, wherein a cooling circuit of said heat source transfers heat generated by the heat source to a first liquid propellant of a rocket engine through a first heat exchanger of a supply circuit of said rocket engine in at least said first liquid propellant. As specified above, this first heat exchanger is contained in a buffer tank of the first propellant supply circuit, and the heat source can be a fuel cell. In addition, after absorbing this heat in the first heat exchanger, a portion of the flow of the first liquid propellant can then be diverted through a second heat exchanger, in which it will absorb heat from a second propellant, to reach the gaseous state before being reinjected into a tank of the first propellant supplying the supply circuit.

Brief description of the drawings

The invention will be better understood and its advantages will appear better, on reading the detailed description which follows, of embodiments shown by way of non-limiting examples. The description refers to the accompanying drawings in which:

- Figure 1 is a schematic view of a machine according to a first embodiment of the invention;

FIG. 2 is a schematic view of a machine according to a second embodiment of the invention; and

- Figure 3 is a schematic view of a machine according to a third embodiment of the invention.

Detailed description of the invention

A machine 1, which may be, for example, a stage of a space launcher, is illustrated schematically in FIG. 1. For its propulsion, this machine 1 comprises a rocket engine 2 with liquid propellants, with a first reservoir 3 for a first propellant, a second reservoir 4 for a second propellant, a propellant chamber 5 for the combustion of a mixture of the two propellants and the acceleration of the combustion gases of this mixture, a first supply circuit 6 connected to the first reservoir 3 and to the propulsion chamber 5 in order to bring the first propellant from the first reservoir 3 to the propulsion chamber 5, and a second supply circuit 7 connected to the second reservoir 4 to the propulsion chamber 5 to bring the second propellant of the second reservoir 4 to the propellant chamber 5. These first and second propellants may be cryogenic propellants such as hydrogen and liquid oxygen. Each of the feed circuits 6, 7 includes a pump 8, 9 for circulating the respective propellant through each feed circuit 6, 7, and outlet valves 10, 11 to open and close the flow. propellants to the propulsion chamber 5. These pumps 8,9 may be, for example, electric pumps, or turbopumps.

In addition, the machine 1 also comprises, for the power supply of on-board equipment, an on-board fuel cell 16, adapted to generate electricity following a chemical reaction between the two propellants and connected to circuits of supply 12,13 to be fed with these two propellants. Each of these circuits 12, 13 comprises a micropump 14, 15 making it possible to regulate the propellant flow rate supplied to the fuel cell 16. Owing to the internal pressure of the tanks 3,4, the micropumps 14, 15 can, however, be optionally substituted. by variable flow valves, the internal pressure tanks 3,4 can then be sufficient to ensure the flow of the propellants to the fuel cell 16.

The fuel cell 16 is furthermore provided with a cooling circuit 17 containing a cooling fluid such as, for example, helium, and connected to a heat exchanger 18 integrated in a buffer tank 20 of the cooling circuit. supply 6 of the first propellant. In the machine 1 illustrated, the circulation of this cooling fluid in the cooling circuit 17 can be driven by and regulated through a variable flow forced circulation device 19, which, in the illustrated embodiment, takes the form of a fan. However, other alternatives can be envisaged both to impulse the circulation of the cooling fluid and to regulate it. So, the pulse of the cooling fluid could be provided by a thermosiphon, and the flow control by at least one variable flow valve.

In operation, after the valves 10 and 11 have been opened, the pumps 8, 9 drive the propellants feeding the propulsion chamber 5 through the supply circuits 6. The heat generated by this fuel cell 16, which is simultaneously fed with propellants through the supply circuits 12,13 in order to generate electricity, is evacuated, through the cooling circuit 17 and the heat exchanger 18, to the first propellant flowing through the circuit. 6. In particular, in the embodiment described, the very low temperature of this first propellant, when it is a cryogenic liquid, allows extremely efficient evacuation of this heat.

Thanks to the buffer tank 20, it is possible to evacuate towards the first propellant a larger amount of heat emitted by the fuel cell 16, even when the valves 10, 11 are closed and the pumps 8, 9 are shutdown. A volume V _t of 30 l of liquid hydrogen in the buffer tank 20 can thus absorb the amount of heat corresponding to a heat output P _c of 100 W for one hour with a temperature rise ΔΤ of only 17 K of the liquid hydrogen.

A machine 1 according to a second embodiment is illustrated in FIG. 2. This other machine 1 differs from that of the first embodiment in that the first feed circuit 6 comprises, downstream of the buffer tank 20, a connection 21 back to the top of the first tank 3 through a variable flow valve 22, and a second heat exchanger 23 which is integrated in the base of the second tank 4 near its connection to the second supply circuit 7. The second circuit 7 also comprises, downstream of the pump 9, a return connection 40 to the top of the second reservoir 4, passing through another heat exchanger 41 disposed around the propulsion chamber 5 so as to be heated by radiation or conduction by this one. This branch 40 also comprises, upstream of the heat exchanger 41, a valve 42, which may be a variable flow valve, thus allowing precise regulation of the flow rate through the branch 40. The other elements of this machine 1 are essentially equivalent to those of the first embodiment, and receive the same reference numerals.

In operation, after its heating by the heat exchanger 18, part of the flow of the first propellant discharged from the first tank 3 through the first supply circuit 6 is diverted through the branch 21 to the second heat exchanger 23, in which it absorbs an additional heating power of the second propellant, warmer, thus passing into the gaseous state, before being reinjected at the top of the first tank 3 and thus maintain its internal pressure during its emptying. If the first propellant is liquid hydrogen and the second propellant is liquid oxygen, the temperature difference between their respective boiling points, at atmospheric pressure, is almost 70 K, which allows a heat transfer more than enough to ensure the vaporization of the liquid hydrogen before their temperatures equalize, and this even for a large flow of liquid hydrogen relative to the volume of liquid oxygen contained in the second reservoir. At the same time, this absorption of heat by the first propellant in the second heat exchanger 23 cools the second propellant, which makes it possible to reduce the saturation pressure of the second propellant supplying the pump 9, so as to avoid cavitation phenomena. This also allows to more widely fluctuate the pressure and temperature of the second propellant in the second tank 4.

At the same time, to maintain the pressure in the second tank 4, a portion of the flow of the second propellant extracted from the second tank 4 through the second circuit 7 is diverted through the branch 40 and heated in the exchanger 41 by the thermal radiation the propellant chamber 5 or conduction, so that it goes into the gas phase before being reinjected into the second tank 4, in order to maintain the internal pressure. This flow bypass is regulated by the valve 42.

However, alternatively to the pumps 8 and 9 of these first two embodiments, the flow of the propellant to the propellant chamber can be provided by other means such as, for example, by pressurizing the tanks. Thus, in a third embodiment illustrated in FIG. 3, these pumps are replaced by a reservoir 24 of pressurized gas, for example helium, connected to the propellant tanks 3,4 through respective valves 26,27. Thus, in operation, the helium pressure of the pressurized gas reservoir 24 pushes the propellants, through their respective feed circuits 6, 7, to the propellant chamber 5. The propellant pressurization in the tanks 3, 4 allows also to overcome micropumps for propellant supply of the fuel cell 16, which supply is regulated in this embodiment by valves 28,29 variable rate in the circuits 12,13.

On the other hand, as in the second embodiment, the first supply circuit 6 comprises a buffer tank 20 and, downstream thereof, a return connection 21 to the top of the first tank 3 through a variable flow valve 22 and a second heat exchanger 23 which is integrated in the base of the second tank 4 near its connection to the second supply circuit 7, which reduces the pressure gas consumption of the tank 24 for the pressurizing the first propellant reservoir 3. Finally, in order to allow reinjection into the gaseous phase at the top of the first reservoir 3 of the propellant derived through the connection 21, this connection 21 comprises a forced circulation device 30, more specifically under as a fan or a pump. The other elements of this machine 1 are essentially equivalent to those of the second embodiment, and receive the same reference numerals.

Although the present invention has been described with reference to specific exemplary embodiments, it is obvious that various modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. In addition, individual features of the various embodiments mentioned can be combined in additional embodiments. Thus, for example, in a variant of the third embodiment, the machine could also comprise a reinjection connection of the second propellant in the gas phase in the second reservoir, as in the second embodiment, with a forced circulation device of this second propellant in the gas phase. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

1. Power supply circuit (6) of a rocket engine (2) in at least a first liquid propellant, characterized in that it comprises at least one buffer tank (20) for said first liquid propellant and a first heat exchanger (18) integrated in said buffer tank (20) and adapted to be connected to a cooling circuit (17) of at least one heat source.

2. Feeding circuit (6) according to any one of the preceding claims, wherein said first liquid propellant is a cryogenic liquid.

The feed circuit (6) of claim 2, wherein said first liquid propellant is liquid hydrogen.

A feed circuit (6) as claimed in any one of the preceding claims comprising a pump (8) upstream of said first heat exchanger (18).

5. Feeding circuit (6) according to any one of the preceding claims, further comprising, downstream of said first heat exchanger (18), a branch (21) passing through a second heat exchanger (23).

6. An assembly comprising a supply circuit (6) according to claim 5 and a reservoir (3) for said first liquid propellant, said reservoir (3) being connected to said supply circuit (6) upstream of said first heat exchanger (18), and also to said branch (21) downstream of said second heat exchanger (23).

7. An assembly comprising a supply circuit (6) according to claim 5 and a reservoir (4) for a second liquid propellant, wherein said second heat exchanger (23) is integrated.

8. An assembly comprising a supply circuit (6) according to any one of claims 1 to 5 and a heat source equipped with a cooling circuit (17) connected to said first heat exchanger (18) of the circuit of power supply (6).

The assembly of claim 8, wherein said heat source is a fuel cell (16).

Apparatus (1) comprising a rocket motor (2) with at least one feed circuit (6) according to any one of claims 1 to 5, and a heat source with a cooling circuit (17) connected to said first heat exchanger (18) of the supply circuit (6).

A method of cooling a heat source, wherein a cooling circuit (17) of said heat source transfers heat generated by the heat source to a first liquid propellant of a rocket engine (2). through a first heat exchanger (18) integrated in a buffer tank (20) of a supply circuit (6) of said rocket engine (2) in at least said first liquid propellant.