FR3069042A1 - THERMAL INSULATION FOR PRESSURIZED CRYOGENIC RESERVOIR - Google Patents

Abstract

The invention relates to a cryogenic reservoir (10) pressurized for liquid propellant, the reservoir (10) comprising a liquid outlet (14), a plurality of thermally insulating elements and a filter (22) disposed on the liquid outlet (14). ) and configured to retain the thermally insulating elements in the reservoir (10), the thermally insulating elements being freely moving relative to one another in the reservoir (10) and configured to form a thermal insulation layer (20) between a liquid propellant (16) intended to be contained in the tank (10) and a gas head (18).

Description

BACKGROUND OF THE INVENTION The present disclosure relates to a pressurized tank for liquid propellant, for example for a lower stage of a rocket engine or for terrestrial storage.

When using liquid propellants, used as oxidant and fuel in a rocket engine, the liquid propellants are stored in tanks. These tanks are pressurized to guarantee their structural strength and ensure a stable supply to the engines.

This pressurization is usually carried out by injecting a heated gas into the gas headspace of the tanks. The heating of the pressurization gases makes it possible to lower the mass of gas contained in the tanks at the end of propulsion which is equivalent to a dry mass penalizing the performance of the launcher.

This pressurization can be autogenous, that is to say that the propellant itself is used in gaseous form by combining a sample from the engine, for example at the pump outlet, and a vaporization and heating the propellant by appropriate means, for example using a heater placed in the hot gas piping on a gas generator cycle engine.

This pressurization can be non-autogenous and a neutral gas such as helium or nitrogen is generally used.

[0006] Thus, in flight, for a liquid oxygen tank, the tank comprises liquid oxygen at the bottom, and a gaseous sky comprising helium and / or gaseous oxygen. For example, liquid oxygen is at a temperature of 90.5 K (Kelvin), while gaseous helium and oxygen gas injected into the upper part of the tank are initially at temperatures of the order of 300 K. Even if the helium gas is cooled in contact with liquid oxygen until it reaches a temperature close to that of liquid oxygen, its density remains lower than that of pressurized gaseous oxygen. Thus, the helium gas tends to rise in the gaseous sky towards the top of the tank, so that the gaseous oxygen comes into contact with the liquid oxygen. Consequently, the gas / liquid interface is an interface between liquid oxygen and gaseous oxygen. The heat exchanges which occur at this interface have the effect of reducing the temperature of the gaseous oxygen, which increases its density. As a result, the mass of gaseous oxygen in the sky above the tank increases, which overall increases the mass of the tank. In addition, these exchanges cause heating of the liquid oxygen which, at the end of propulsion, can reach an unacceptable temperature for the engine, in particular for its supply pump.

[0007] Thus, in a rocket engine, for example of a lower stage of a rocket engine, there may remain in the tanks several hundred kilograms of propellant which are not used, also called thermal consumables.

Furthermore, in flight, the propellant tanks are subjected to significant vibrations which can modify the interface between the liquid phase and the gaseous phase in the tank and therefore modify the mass and heat exchange rates. between the liquid phase and the gas phase.

These modified exchanges can cause, in the autogenous pressurization configuration, an increase in the condensation of the gaseous species and therefore require increased flow rates of pressurization gas which may not be acceptable for heat exchangers. In a non-autogenous pressurization configuration, the dissolution of the pressurization gas in the liquid propellant can cause cavitation phenomena in turbopumps and inductors.

OBJECT AND SUMMARY OF THE INVENTION The present description aims to remedy these drawbacks at least in part.

To this end, the present disclosure relates to a pressurized cryogenic reservoir for liquid propellant, the reservoir comprising a liquid outlet, a plurality of thermally insulating elements and a filter disposed on the liquid outlet and configured to retain the thermally elements. insulating in the tank, the thermally insulating elements being free to move relative to each other in the tank and configured to form a layer of thermal insulation between a liquid propellant intended to be contained in the tank and a gaseous sky.

Thanks to this thermal insulation layer present between the liquid propellant and the gas headspace of the tank, the heat exchanges between the liquid phase (liquid propellant) and the gas headspace (pressurization gas) are reduced. The phenomena of thermal convection and condensation between the pressurization gas and the liquid propellant are greatly reduced.

Furthermore, the presence of this thermal insulation layer also reduces the formation of droplets and / or the surface of the liquid / gas interface in the tank, that is to say the interface between the liquid propellant and pressurizing gas, even when the tank is subjected to vibrations, for example at the sloshing frequency. The sloshing phenomena are commonly called “sloshing” phenomena, with reference to the English term for these phenomena. Indeed, the thermal insulation layer being at the liquid / gas interface in the tank, the phenomena linked to "sloshing" are greatly reduced.

One can thus advantageously remove or reduce the number of elements in the tank which are usually present in a tank of the prior art to reduce the phenomena related to "sloshing".

The pressurization gas therefore has less tendency to condense or dissolve in the liquid propellant. The mass of the pressurizing gas is thus reduced compared to a tank of the prior art.

The temperature of the liquid propellant also tends to increase less. Also, the mass of thermal consumables is reduced compared to a tank of the prior art. We can therefore use a larger volume of the propellant on board.

Thus, the volume occupied by the thermally insulating elements is compensated for by the reduction in volume of the thermal non-consumables. Furthermore, some of these thermally insulating elements can be completely submerged when the tank is completely filled and part of them can emerge when the tank is going to empty of propellant. Advantageously, it is not possible to modify the reservoir, in particular its volume.

It is understood that this thermal insulation layer being present between the liquid propellant and the gas canopy of the tank, the thermally insulating elements which form it can float, at least partially, on the surface of the liquid propellant.

Thus, the thermally insulating elements being free to move relative to one another, the thermally insulating elements are free to move at the liquid / gas interface of the tank and to rearrange each other depending on movement of the liquid / gas interface. The thermal insulation layer can therefore adapt to the surface of the liquid / gas interface. The thermally insulating elements can also be rearranged to adapt to the shape of the reservoir 10 which may not be of constant section and also to adapt to the presence, in the reservoir 10, of structural reinforcements, level probes, temperature sensors, etc.

In addition, the presence of the thermally insulating elements makes it possible to reduce the exchanges with the pressurization gas which tends to condense and enter the liquid phase in the form of fine droplets of liquid.

The tank is a tank for a lower stage of a rocket engine or for terrestrial storage.

The thermally insulating elements may have a thermal conductivity less than or equal to 0.100 W / mK (watt per meter-kelvin), preferably less than or equal to 0.075 W / mK, even more preferably less than or equal to 0.050 W / mK

It is understood that the thermal conductivity is the thermal conductivity of the thermally insulating element. Thus, it is conceivable to use a material whose thermal conductivity is greater than the values mentioned above but which would include porosity, so that the thermal conductivity of the thermally insulating element would be less than or equal to the above values.

The thermally insulating elements can be spheres.

The spherical shape allows the thermally insulating elements to move easily relative to each other. The spherical shape allows a relatively compact stack of thermally insulating elements. The spherical shape also allows the thermally insulating elements to better resist the pressure of the pressurization gas which can be between 3 bar and 6 bar, regardless of the liquid propellant used.

The spheres can be hollow.

One can thus use materials having a thermal conductivity greater than the above values. It is therefore possible to widen the choice of materials which can be used to manufacture the thermally insulating elements.

The thermally insulating elements can be closed cell foam or non-porous material.

The foam cells being closed or the material being non-porous, the liquid propellant cannot penetrate into the thermally insulating element.

The thermally insulating elements can form a plurality of layers of thermally insulating elements in the tank to form a stack of thermally insulating elements.

The thermal insulation layer is formed by a plurality of layers of thermally insulating elements. The thermal insulation between the liquid propellant and the pressurization gas is thus improved. Furthermore, the insulating layer being formed by a stack of thermally insulating element, these elements are in contact with one another, which makes it possible to better dissipate and reduce the oscillations of the liquid / gas interface in the tank. .

The stack of thermally insulating elements may include an immersed part of the stack of thermally insulating elements in a liquid propellant contained in the tank and an emerged part.

The emerged part of the stack of thermally insulating elements may have a height greater than or equal to 10 mm (millimeter) and less than or equal to 30 mm.

This condition is generally satisfied before using the liquid propellant contained in the tank.

The thermally insulating elements may have a density less than or equal to a quarter of the density of the liquid propellant intended to be contained in the tank, preferably less than or equal to a fifth, even more preferably less than or equal to one sixth.

It is understood that the density of the liquid propellant is the density at the temperature at which the propellant is liquid. Thus, it is conceivable to use a material whose density is greater than the values mentioned above but which would include porosity, so that the density of the thermally insulating element would be less than or equal to the above values.

The filter may be a grid comprising orifices, the orifices having a dimension less than a smaller dimension of the thermally insulating elements.

The thermally insulating elements cannot leave the tank through the orifices of the filter and, for example, enter a turbopump.

Brief Description of the Drawings Other characteristics and advantages of the invention will emerge from the following description of embodiments of the invention, given by way of nonlimiting examples, with reference to the appended figures, in which:

- Figures IA, IB and IC are schematic representations of a reservoir according to the description;

- Figure 2 is an enlargement of a thermal insulation layer according to the description.

Detailed description of the invention [0040] Figure IA represents a reservoir 10 for liquid propellant. This tank 10 can be a tank for a lower stage of a rocket engine, also called a booster, or for terrestrial storage.

The reservoir 10 is delimited by a wall 12 and has an opening serving as a liquid outlet 14 from the reservoir 10. The reservoir 10 may have other openings which are not shown in FIG. IA. The reservoir 10 may for example include a filling opening of the reservoir 10, an opening for injecting pressurization gas, etc.

The tank 10 of Figure IA is shown when it contains a liquid propellant 16, for example liquid oxygen. Liquid oxygen is at a temperature of around 90.5 K.

The tank 10 comprises, above the liquid propellant 16, a gaseous sky 18. Depending on whether the pressurization is autogenous or non-autogenous, the composition of the pressurization gas forming the gaseous sky 18 of the tank 10 may vary. The pressurization gas is at a temperature of approximately 300 K. When the pressurization is autogenous, the composition of the gas overhead 18 is gaseous oxygen. When the pressurization is not autogenous, the composition of the overhead gas 18 comprises gaseous oxygen mixed with an incondensable gas such as helium or nitrogen.

As shown in FIG. 2, between the liquid propellant 16 and the gas headspace 18, there is a liquid / gas interface 24.

In Figures IA, IB and IC, the liquid / gas interface 24 is shown schematically by a line in dashed lines. However, the liquid / gas interface 24 as shown in Figure 2 is present in the reservoir 10 of Figures IA, IB and IC.

The reservoir 10 includes a thermal insulation layer 20 disposed at the liquid / gas interface 24.

As shown in Figure 2, the thermal insulation layer 20 is formed of a plurality of thermally insulating elements 26 stacked on each other. The thermally insulating elements 26 are free to move relative to one another.

In the embodiment of Figures IA, IB and IC, the thermally insulating elements 26 may be solid spheres of closed cell foam. The solid spheres can also be made of non-porous material.

In the embodiment of Figure 2, the thermally insulating elements 26 are hollow spheres comprising a shell 26A, for example in closed cell foam or non-porous material, and an internal cavity 26B, disposed at the interior of hull 26A. When the shell 26A is made of non-porous material, it is possible to introduce a gas into the shell 26A, for example it is possible to inject nitrogen, which is an inert gas, at a pressure for example of 1 bar.

Figure IB shows the reservoir 10 when part of the liquid propellant 16 has been used to power the rocket engine.

Figure IC shows the reservoir 10 at the end of use. The reservoir 10 comprises a quantity of liquid propellant 16 present between the thermally insulating elements 26. As can be seen, the reservoir 10 comprises a filter 22 disposed on the liquid outlet 14 and configured to retain the thermally insulating elements 26 in the reservoir 10. In the embodiment of FIG. 1, the filter 22 comprises two grids 22A, 22B comprising orifices, the orifices having a dimension less than the diameter of the thermally insulating elements 26.

Also shown in Figure 2, the heat exchanges between the different elements present in the tank 10. In particular, arrows 28 represent the heat flow from the pressurization gas to the liquid propellant 16. There is also represented the phenomena of natural convection 30 which take place in the gas trapped in the internal cavity 26B of each hollow sphere, the convective exchanges 32 taking place between the spheres, the heat exchanges by direct contact or conduction 34 taking place from sphere to sphere, the heat exchanges 36 by direct contact between the wall 12 of the tank 10 and a sphere, as well as a heat exchange by convection 38 in a cavity close to the wall 12 of the tank 10.

In the embodiment of the solid spheres, only the natural convection phenomena 30 are absent in the tank 10.

As shown in Figure 2, the thermally insulating elements 26 form a thermal insulation layer 20. The thermally insulating elements 26 are stacked on each other and form a plurality of layers of thermally insulating elements 26 , three layers of thermally insulating elements 26 in the embodiment of FIG. 2.

The presence of the thermally insulating elements 26 at the liquid / gas interface 24 of the reservoir 10 makes it possible to reduce the phenomena of "sloshing", even if the layer of thermally insulating elements 26 is not continuous. Also, the advantages obtained thanks to the presence of the thermal insulation layer 20 at the liquid / gas interface 24 of the reservoir 10 can be observed even when the thermal insulation layer 20 is formed from a single layer of elements. thermally insulating 26 or when the layer of thermally insulating elements 26 is not continuous.

It will be noted that the presence of three layers of thermally insulating elements 26 makes it possible to significantly reduce the convective heat exchanges between the pressurization gas and the liquid propellant 16.

Preferably, only the first layer of thermally insulating elements 26 is partially immersed in the liquid propellant 16, when the tank 10 is at rest. Thus, we can determine the Archimedes' thrust which allows the apparent weight of the thermal insulation layer 20 to be balanced.

It will be noted that the rocket acceleration during its first flight phase has no effect on the apparent weight of the thermal insulation layer 20.

Thus, the following formula will be used:

Vjmmergées total X Pergol ⁼ Vgphères X Psphère (1) WHERE

Vjmmergées total is the total volume of the immersed spherical caps of radius R, it is equal to the sum of the volume of each immersed spherical cap (V _im mergée) / the volume of each immersed spherical cap being calculated for each partially immersed sphere;

V _S spheres is the total volume of the spheres of radius R;

Pergoi is the density of liquid propellant;

Psphère is the density of the thermally insulating elements 26, that is to say the ratio of the mass of a thermally insulating element to its volume.

V _immersed = ÎÜJBzS. (2) where

R is the radius of the spheres;

h is the height of the submerged spherical cap.

Using the immersed volume of a thermally insulating element 26, in this example a sphere, and taking into account the density of the liquid propellant 16 and the density of the thermally insulating element 26, we can determine that the thermally insulating elements 26 have a density less than or equal to a quarter of the density of liquid propellant 16 intended to be contained in the reservoir 10, preferably less than or equal to a fifth, even more preferably less than or equal to one sixth.

It has also been determined that the emerged height H of the thermal insulation layer 20 is preferably between 10 mm and 30 mm in this example.

For a pressurized cryogenic tank 10 comprising liquid methane, the thermally insulating elements 26 preferably have a density less than 422 kg / m ³ (kilogram per cubic meter) divided by 4, or 105.5 kg / m ³ , preferably less than 422 kg / m ³ divided by 5, or 84.4 kg / m ³ , even more preferably less than 422 kg / m ³ divided by 6, or 70.3 kg / m ³ , the density liquid methane being from about 422 kg / m ³ to about 112 K.

For a pressurized cryogenic tank 10 comprising liquid oxygen, the thermally insulating elements 26 preferably have a density less than 1139 kg / m ³ divided by 4, or 284.75 kg / m ³ , preferably less at 1139 kg / m ³ divided by 5, or 227.8 kg / m ³ , even more preferably less than 1139 kg / m ³ divided by 6, or approximately 189.83 kg / m ³ , the density of the liquid oxygen being from about 1139 kg / m ³ to about 90.5 K.

Depending on the density of the material chosen to form the thermally insulating elements 26, we will choose solid spheres or hollow spheres.

Thus a material such as H920A marketed by Air Liquide has a density of 51 kg / m ³ and a thermal conductivity of 0.033 W / mK between 20 and 400 K or such as R82-60 marketed by AIREX which has a density of 60 kg / m ³ and a thermal conductivity of 0.036 W / mK between 79 K and 433 K could be used in the form of solid spheres with liquid methane, liquid oxygen, liquid kerosene of type RP- 1, liquid ethanol or nitrogen peroxide. On the other hand, polychlorotrifluoroethylene (PCTFE) which has a density of 2160 kg / m3 and a thermal conductivity of 0.25 W / mK will be used in the form of hollow spheres with these same liquid propellants. These materials have good chemical compatibility with liquid propellant. These examples are not limitative.

For example, for a tank 10 comprising liquid methane whose internal diameter is 5.4 m, one can choose thermally insulating elements 26 in the form of solid spheres, in closed cell foam, such as H920A.

Taking into account the density of H920A and the density of liquid methane, for a stack of three layers of thermally insulating elements 26 uniformly distributed over the liquid / gas interface 24 of the tank 10, the values presented in Table 1.

Table 1

Radius of spheres (mm) Submerged height (mm) Total emergent height (mm) Total mass thermal insulation layer (kg) 10 9.0 43.6 55.0 5 4.5 21.8 27.5 2.5 2.2 10.9 13.7

Also, in this example, 27.5 kg of H920A spheres with a radius of 5 mm would be a good compromise.

For example, for a tank 10 comprising liquid oxygen whose internal diameter is 5.4 m, one can choose thermally insulating elements 26 in the form of solid spheres, in closed cell foam, such as the H920A.

Taking into account the density of H920A and the density of liquid oxygen, for a stack of three layers of thermally insulating elements 26 uniformly distributed over the liquid / gas interface 24 of the tank 10, we obtains the values presented in Table 2.

Table 2

Radius of spheres (mm) Submerged height (mm) Total emergent height (mm) Total mass thermal insulation layer (kg) 10 5.03 47.6 55.0 5 2.51 23.8 27.5 2.5 1.26 11.9 13.7

Also, in this example, a good compromise would be to choose H920A spheres with 2.5 mm radius, or even choose H920A spheres with 3.0 mm radius.

To overcome certain stacking defects of the thermally insulating elements 26, it is conceivable to add 10 to 20% by mass of thermally insulating elements 26 in the tank 10.

For hollow PCTFE spheres, it is possible, for example, to use hollow spheres with an external radius of 10 mm and an internal radius of 9.5 mm. Thus, the shell 26A has a thickness of 0.5 mm, the density of a sphere then being approximately 309 kg / m ³ . It will be noted that the density of the hollow spheres in PCTFE is greater than a quarter of the density of liquid oxygen.

It is understood that before being introduced into a tank 10 for rocket engine liquid propellant, the thermally insulating elements 26 are checked. For example, it is understood that the materials are compatible with the cryogenic conditions present in the tank 10.

It is also verified that the thermally insulating elements 26 are not likely to break in the tank 10, at the risk of forming debris which could leave the tank 10 through the liquid outlet 14 passing through the filter 22.

In particular, in the case of hollow spheres, it is verified that the hollow spheres can withstand pressures greater than the pressure in the tank 10, for example, one subjected to thermal and mechanical cycles followed by pressurization hollow spheres at a pressure greater than or equal to twice the pressure in the reservoir 10. All the spheres which pass these tests without deformation and without collapsing can be introduced into the reservoir 10.

We can also sort the thermally insulating elements 26 on the basis of their size to ensure that no thermally insulating element 26 will leave the tank 10 by the liquid outlet

14.

Although the present description has been described with reference to a specific embodiment, 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 discussed can be combined in additional embodiments. Therefore, the description and the drawings should be considered in an illustrative rather than restrictive sense.

It is understood that depending on the need and the mission, the terminals given for emerged height H of the thermal insulation layer 20 may be different.

Claims (10)

1. cryogenic pressurized tank for liquid propellant, the tank (10) comprising a liquid outlet (14), a plurality of thermally insulating elements (26) and a filter (22) disposed on the liquid outlet (14 ) and configured to retain the thermally insulating elements (26) in the reservoir (10), the thermally insulating elements (26) being free to move relative to one another in the reservoir (10) and configured to form a layer of thermal insulation (20) between a liquid propellant (16) intended to be contained in the tank (10) and a gaseous sky (18). 2. Tank (10) according to claim 1, in which the thermally insulating elements (10) have a thermal conductivity less than or equal to 0.100 W / mK, preferably less than or equal to 0.075 W / mK, even more preferably less than or equal to 0.050 W / mK 3. Tank (10) according to claim 1 or 2, wherein the thermally insulating elements (26) are spheres. 4. Tank (10) according to claim 3, wherein the spheres are hollow. 5. Tank (10) according to any one of claims 1 to 4, in which the thermally insulating elements (26) are made of closed cell foam. 6. Tank (10) according to any one of claims 1 to 5, wherein the thermally insulating elements (26) form a plurality of layers of thermally insulating elements (26) in the tank (20) to form a stack of 'thermally insulating elements (26). 7. Tank (10) according to claim 6, in which the stack of thermally insulating elements (26) comprises an immersed part of the stack of thermally insulating elements in a liquid propellant contained in the tank and an emerged part. 8. Tank (10) according to claim 7, in which the emerged part of the stack of thermally insulating elements has a height (H) greater than or equal to 10 mm and less than or equal to 30 mm. 9. Tank (10) according to any one of claims 1 to 8, in which the thermally insulating elements (26) have a density less than or equal to a quarter of the density of the liquid propellant (16) intended to be contained in the tank, preferably less than or equal to one-fifth, even more preferably less than or equal to one-sixth. 5 10. Tank (10) according to any one of claims 1 to 9, wherein the filter (22) is a grid comprising orifices, the orifices having a dimension less than a smaller dimension of the thermally insulating elements (26).