AU2021427742A1

AU2021427742A1 - Fuel

Info

Publication number: AU2021427742A1
Application number: AU2021427742A
Authority: AU
Inventors: Dominic FREUDENMANN; Christoph KIRCHBERGER; Maxim KURILOV
Original assignee: Deutsches Zentrum fuer Luft und Raumfahrt eV
Current assignee: Deutsches Zentrum fuer Luft und Raumfahrt eV
Priority date: 2021-02-12
Filing date: 2021-11-25
Publication date: 2023-08-17
Also published as: JP2024509729A; US20240124372A1; WO2022171324A1; DE102021103380B3; EP4291545A1; KR20240007899A

Abstract

The invention relates to a liquid or gel-like fuel, in particular for rocket engines, comprising: - an inorganic salt as the oxidant, the inorganic salt having an oxygen content of at least 60 wt.%; and - a solvent as the combustible material, comprising a mono- or divalent alcohol, a nitroalkane and/or an ionic liquid with nitrate as the anion, and the solvent has an oxygen content of 30 to 55 wt.%, and the inorganic salt is dissolved in the solvent.

Description

Propellant

The present invention relates to a propellant in liquid or gel form, in particular for rocket engines.

Of propellants for rocket engines that operate by the recoil principle, a distinction is primarily made between solid and liquid propellant systems. Depending on the concrete area of application (launch vehicles, orbital engines, etc.), different propellant systems may be used advantageously, but all systems also have specific disadvantages.

Solid propellants are characterised in particular by a very high energy density, and are widely used in aerospace engineering and military engineering, for example in launch vehicles and military projectiles. It is possible to use both mixtures of solid fuels and oxidising agents (bipropellant) and monopropellants, in which an intramolecular redox reaction takes place (e.g. propellants based on HMX, RDX or a mixture of nitroglycerin and nitrocellulose). Rocket engines using solid propellants are simple to construct, since the propellant is introduced directly into the combustion chamber and so there is no need for a separate propellant tank or for supply systems for the propellant.

The disadvantage of solid propellants is primarily the lack of flexibility, both in scaling the engines and during operation, since it is not possible in practice to vary thrust by controlling burn. It is also a problem that the known solid propellants are highly explosive, which makes safe handling of the propellants before, during and after they are introduced into the combustion chamber very complex. When the solid propellant is introduced, a positively engaging connection with the combustion chamber wall has to be ensured, with gaps or tears possibly resulting in uncontrolled burn and, in extreme cases, total loss of the craft.

In the case of liquid propellants, the above-mentioned disadvantages do not occur, and in particular rocket engines having liquid propellants are substantially

20085958_1 (GHMatters) P122316.AU more flexible than those with solid propellants. Not only is scaling of the engines more easily possible, but so is thrust control by regulating the supply of propellant to the combustion chamber. On the other hand, this entails the disadvantage of a greater complexity of construction, since one or more propellant tanks and supply systems with pumps, valves and other auxiliary components are needed.

Monopropellants and bipropellants are also known in the case of liquid propellants. The latter include cryogenic or partially cryogenic propellants with liquid hydrogen, liquid methane or kerosene as the fuel and liquid oxygen as the oxidising agent. Storage and handling of these liquefied gases is complex and requires the most demanding safety measures, since any leak results in a risk of explosion.

As well as cryogenic systems, combinations of liquid fuels and oxidising agents are known, and these are used in particular in orbital engines for controlling the flight and positioning of satellites or space probes. In this context, the fuels used are typically hydrazine and derivatives thereof (mono- and dimethylhydrazine), in combination with nitric acid, dinitrogen tetroxide or hydrogen peroxide. A significant problem here is that hydrazines are highly toxic and carcinogenic, so for health and ecological reasons alternatives should be used where possible. Dinitrogen tetroxide is also of toxicological concern, but replacing it with hydrogen peroxide impairs the effectiveness of the propellants.

Less toxic alternatives to hydrazines are ammonium dinitramide (ADN) and hydroxylammonium nitrate (HAN). These monopropellants are highly explosive, so they are only handlable in the form of aqueous solutions, optionally in combination with methanol and/or ammonia as additional fuels. In this case too, the water content results in reduced ignitability, such that, unlike hydrazine engines, engines based on ADN or HAN are not capable of cold start, but have to be warmed up beforehand. Moreover, ADN and HAN are comparatively expensive.

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The object of the invention is to propose a propellant, in particular for rocket engines, by which the above-mentioned disadvantages of the prior art can be avoided as far as possible.

According to the invention, this object is achieved by a propellant in liquid or gel form that comprises the following: - an inorganic salt as the oxidising agent, wherein the inorganic salt has an oxygen content of at least 60 weight%; and - a solvent as the fuel, comprising a monovalent or divalent alcohol, a nitroalkane and/or an ionic liquid in which nitrate is the anion, wherein the solvent has an oxygen content of from 30 to 55 weight%, wherein the inorganic salt is dissolved in the solvent.

The propellant according to the invention is in liquid or gel form, wherein in the case of a propellant in gel form it is also pumpable (see below). In this way, the typical disadvantages of solid propellants are avoided. The propellant according to the invention is not cryogenic, which fundamentally simplifies storage, handling and the supply of the propellant to the engine. Finally, the components of the propellant according to the invention are of relatively little concern from a toxicological and ecological point of view, at least by comparison with hydrazine and its derivatives. By comparison with ADN or HAN, the propellant according to the invention offers a significant cost advantage.

Although the propellant according to the invention is a bipropellant, since the oxidising agent and fuel are separate chemical compounds, it is advantageously a homogeneous mixture that is supplied to the engine. Here, the invention exploits the fact that the inorganic salts used as the oxidising agent have relatively good solubility in various solvents that are suitable as fuel, with the result that the required mixing ratios of oxidising agent and fuel can be set. In this context, it is particularly advantageous that there may be used as solvents relatively low-energy fuels (monovalent or divalent alcohols) rather than relatively high-energy fuels (nitroalkanes/nitrates) in order to enable the specific

20085958_1 (GHMatters) P122316.AU properties of the propellant, such as oxygen balance, energy density, specific impulse, etc., to be adapted to different requirements and areas of use.

The propellant according to the invention and its components are typically not explosive, which simplifies handling and manufacture and makes the propellant safer. On the other hand, the propellant according to the invention has good ignition properties, wherein electrical, thermal, catalytic-thermal or catalytic ignition are fundamentally possible.

Th inorganic salt that is used as the oxidising agent is preferably selected from lithium nitrate, lithium dinitramide, lithium perchlorate or a mixture thereof. These salts have relatively good solubility in a plurality of suitable fuels, in particular alcohols. For example, the solubility of lithium nitrate in methanol is approximately 58 g/100 g, and the solubility of lithium perchlorate in methanol is approximately 182 g/100 g. In nitroalkanes and ionic liquids where nitrate is the anion, solubilities are less good, and in this case it is possible to increase the solubility of the salts by admixing alcohols or further solvents.

The proportion of inorganic salt in the propellant may be varied over a broad range, usually being in a range of from 15 to 65 weight%. On the one hand the proportion depends on the solvent selected and the solubility of the inorganic salt, as discussed above, and on the other it may be varied in dependence on the desired oxygen balance for combustion of the propellant.

In a preferred embodiment of the invention, the solvent comprises ethanol, methanol or n-butanol, in particular as a single constituent. Because of the good solubility in these monovalent alcohols, in this case a relatively high proportion of inorganic salt in the propellant may be selected, preferably from 50 to weight%.

In a further advantageous embodiment of the invention, the solvent comprises nitromethane or nitroethane, in particular as a single constituent. Because of the

20085958_1 (GHMatters) P122316.AU somewhat lower solubility, in this case the proportion of inorganic salt is preferably from 10 to 40 weight%, more preferably from 20 to 30 weight%.

As mentioned above, the propellant according to the invention may comprise a single alcohol or a single nitroalkane as the solvent or fuel. According to a further embodiment of the invention, the solvent additionally comprises a further alcohol, in particular n-butanol (where this is not the primary solvent) and/or ethylene glycol, and/or a carbonate ester, in particular dimethyl carbonate, diethyl carbonate, ethyl methyl carbonate and/or propylene carbonate. Carbonate esters are also low-energy fuels.

In a further preferred embodiment of the invention, the solvent comprises an ionic liquid, in particular ethylammonium nitrate. In this case, the proportion of inorganic salt in the propellant is preferably from 10 to 40 weight%, more preferably from 15 to 25 weight%.

In addition to the ionic liquid, the solvent preferably comprises an alcohol, in particular ethylene glycol and/or ethanol. A mixture of this kind enables the solubility of the inorganic salt (oxidising agent) to be increased by comparison with a single solvent. In the case of ethylammonium nitrate, the mixing ratio of ethylammonium nitrate to alcohol is preferably in the range of from 6:1 to 1:3.

Preferably, the propellant according to the invention contains no water, and in this differs in particular from the known propellants based on ADN or HAN. As a result of the absence of water, the propellants according to the invention have good ignitability.

According to a further preferred embodiment, the propellant according to the invention is a propellant in gel form. In order to achieve a gel-like consistency, the propellant comprises a thickening agent, which is preferably selected from polyacrylic acids, pyrogenic silicon dioxides, micro- to nano-scale metal powders, titanium dioxide nanoparticles and/or carbon nanotubes. In the case of metal

20085958_1 (GHMatters) P122316.AU powder, this is preferably selected from aluminium, magnesium, aluminium/magnesium alloys, boron, iron and zirconium.

Propellants in gel form have the advantage over liquid propellants that they tend to be safer, firstly because the vapour pressure of the liquid components is reduced, and secondly because the higher viscosity means that the discharge rate in the event of a leak is lower. A further advantage is that, in the case of a propellant in gel form, insoluble components that would settle in a liquid propellant may also be kept in suspension. This is true in particular of the metal powders which the propellant optionally contains, which in addition to their function as a thickening agent also provide an additional fuel and can be used to increase the energy density of the propellant.

The proportion of thickening agent in the propellant according to the invention is preferably up to 10 weight%, more preferably from 1 to 5 weight%. Here, the type and quantity of the thickening agent may favourably be selected such that the propellant according to the invention in gel form has substantially all the advantages of a liquid propellant, that is to say in particular good pumpability and flexibility (simple scalability, capacity for thrust control, and re-ignitability). As a result, the propellants in gel form according to the present invention differ substantially from known gel propellants of the prior art, of which the rheology has a shear thinning behaviour and which have a pronounced yield point such that they are not pumpable using the feed systems that are usual for liquid propellants.

Further, the propellant according to the invention may comprise one or more hydrides of light metals, which are preferably selected from AH 3, NaBH 4 and/or AILiH 4 .Metal hydrides are additional fuels by which the energy content and performance capability of the propellant may be modified.

According to a further embodiment of the invention, the propellant further comprises a further oxidising agent, which is preferably selected from the nitrates and perchlorates of ammonium, sodium and potassium. This is true in

20085958_1 (GHMatters) P122316.AU particular in the case of propellants in gel form, in which these oxidising agents are in suspended form because of their lower solubility.

The propellant according to the invention typically has a density in the range of from 900 to 1 700 kg/M 3, preferably in the range of from 1 100 to 1 400 kg/M 3

. For combustion, the propellant according to the invention favourably has an oxygen balance of from 0 to -50%, more preferably from -20 to -40%. At an oxygen balance of 0, combustion is completely stoichiometric, with the result that the energy content of the propellant is entirely used up. However, a negative oxygen balance, that is to say an excess of fuel with respect to the oxidising agent, is preferred in most cases in order to avoid an excessive tendency to spontaneous combustion (explosiveness) of the propellant.

Because of the above-mentioned possibilities for varying the qualitative and quantitative composition of the propellant in liquid or gel form within the context of the present invention, the specific impulse of the propellant may also lie within a broad range (e.g. within the range of from 150 to 300 s at a combustion pressure of 7 MPa and an expansion ratio of 70:1). Accordingly, the propellants according to the invention may be used in various types of rocket engines in aerospace engineering, for both main drives and auxiliary drives, in particular for launch vehicles, booster rockets, rocket stages or orbital engines. In addition, propellants having a specific impulse at the lower end of the above-mentioned range may also be used for the operation of gas generators in aerospace systems.

Apart from space travel, the propellant according to the invention may also be used for driving aircraft (e.g. for auxiliary power units on launch) or civil or military projectiles.

A further advantageous area of application for the propellant according to the invention is in mining, where the propellant may be used for example for cutting torches or borers. However, it is also conceivable, apart from mining, to use the

20085958_1 (GHMatters) P122316.AU propellant for driving for example machine tools for joining and separating metals.

These and further advantages of the invention are explained in more detail with reference to the examples below.

Examples of liquid propellants

In Table 1 below, in each case the percentage composition, specific impulse, density, adiabatic combustion temperature and oxygen balance are specified for four examples of liquid propellants according to the invention (Examples 1 to 4). As comparison examples here there serve conventional propellants based respectively on hydrazine (V1) and ammonium dinitramide (V2 and V3).

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Table 1 Example Composition in Specific Density Adiabatic Oxygen weight% impulse combustion balance temperature 1 30% lithium perchlorate 284s 1 280 2 472 K -31% 10% n-butanol kg/M 3 60% nitromethane 2 60% lithium perchlorate 246 s 1 100 1 682 K -40% 40% ethanol kg/m 3 3 18% lithium nitrate 230 s 1 280 1 198 K -50% 70% ethylammonium kg/m 3 nitrate 12% ethylene glycol 4 17% lithium nitrate 233 s 1 320 1 307 K -41% 11% ammonium nitrate kg/m 3 60% ethylammonium nitrate 12% ethylene glycol V1 100% hydrazine 233 s 1 010 880 K kg/M 3 V2 63% ADN 256s 1 240 1 863 K -16% 18% methanol kg/m 3 5% ammonia 14% water V3 65% ADN 262 s 1 360 2 184 K -0.4% 11% monomethyl kg/m 3 formamide 24% water

The values for specific impulse (at a combustion pressure of 5 MPa and an expansion ratio of 50:1) are comparable with those of conventional propellants

20085958_1 (GHMatters) P122316.AU or in some cases even higher. The densities of the propellants according to the invention are likewise in a similar range.

Like the specific impulse, the adiabatic combustion temperature was calculated using the NASA-CEA code (McBride & Gordon, 1996). In Examples 2 to 4, this value is also in a similar range to the comparison examples. This means that the materials used for construction of the engines can be very similar to those hitherto, which simplifies technical implementation of the new propellants. Example 1 is an exception. In this case, both the combustion temperature and the performance (specific impulse) are significantly higher, with the result that an adaptation to existing engineering of the engines might be required (construction materials that can withstand high temperature, in particular for catalyst devices), which does nonetheless appear worthwhile given the performance potential that is realised.

The oxygen balance of the propellants according to the invention is lower than in the case of the conventional "green propellants" based on ADN. This means on the one hand that combustion is less stoichiometric, resulting in incomplete conversion of the chemical energy into propulsion energy; on the other hand, it may be an indication that the newly developed propellants are less easy to detonate as a result of mechanical and thermal loads.

Figure 1 shows a graph of the performance potential (delta v) of the propellants listed in Table 1, expressed as a percentage deviation from the reference propellant, hydrazine (V1), in different spacecraft configurations. These are plotted against the burnout mass ratio (), which is shown on the x axis. As an example, a burnout mass ratio of 0.55 corresponds to a typical space probe fuelled by hydrazine (that is to say, the propellant accounts for 45% of the total spacecraft mass); a burnout mass ratio of 0.92 would correspond to an earth observation satellite fuelled by hydrazine.

If a denser propellant is used, more propellant can be carried with the same tank volume, that is to say that the burnout mass ratio falls and the delta v of the

20085958_1 (GHMatters) P122316.AU spacecraft, that is the amount of adjustment of the orbital velocity that is required for orbital manoeuvres, goes up. The lines on the graph show how much more delta v can be applied if propellants other than hydrazine are used in the same spacecraft. With conventional propellants based on ADN (V2 and V3), between 30 and 50% more delta v can be achieved, and hence the duration of operation of probes and satellites can be lengthened by up to 1.5 times. The liquid propellants according to the invention are in a similar range, or higher (Example 1), and so are competitive with conventional hydrazine substitutes.

Examples of propellants in gel form

In Table 2 below, in each case the percentage composition, specific impulse, density and C* combustion efficiency are specified for three examples of propellants in gel form according to the invention (Examples 5 to 7). As comparison examples there serve various conventional propellants (V4 to V7).

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Table 2 Example Composition in weight% Specific Density C* combustion impulse efficiency 5 15% lithium nitrate 269 s 1 420 45% 59% ethylammonium kg/m3 nitrate 10% ethylene glycol 15% aluminium powder 1% Carbopol 980 6 50% lithium perchlorate 263 s 1 280 55% 34% ethanol kg/m3 15% aluminium powder 1% Carbopol 980 7 23% lithium perchlorate 291 s 1 370 >70% 8% n-butanol kg/m3 50% nitromethane 15% aluminium powder 2% Carbopol 980 2% Aerosil 200 V4 69% ammonium 280 s 1 808 n.d. perchlorate kg/m3 19% aluminium powder 12% HTPB V5 Liquid oxygen / methane in 343 s 1 105 n.d. a mass ratio of 3.6:1 kg/m3 V6 Liquid oxygen / kerosene 337 s 1 049 n.d. in a mass ratio of 2.7:1 kg/M 3 V7 75% nitromethane 287 s 1 270 92% 15% aluminium powder kg/m 3 10% gelling agent and other additives

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The values for specific impulse are lower here than in the case of cryogenic or partially cryogenic bipropellants, and are more comparable with the energy characteristics of solid propellants. The densities of the propellants according to the invention are higher, however, though not as high as the density of solid propellants.

Figure 2 shows a graph of the performance potential (delta v) of the propellants listed in Table 2, expressed as a percentage deviation from the comparison example V4, in different spacecraft configurations. The burnout mass ratio values presented are typical of projectiles and high-altitude research rockets (0.15 to 0.35) and booster stages (0.3 to 0.45) or upper stages (0.4 to 0.65).

With the same size of stage, all the propellants shown in the graph would provide less delta v than the reference solid propellant (V4), which is used for example in the P80 booster stage of the Vega launch system. In the case of cryogenic or partially cryogenic propellants, there is between 1 and 20% less delta v. In the case of the propellants according to the invention, there is between 12 and 25% less delta v. Example 7 is an exception: the delta v performance of this propellant is in the range of conventional cryogenic and partially cryogenic bipropellants.

However, the fact that all the propellants tested would give less delta v than the reference propellant does not mean that their use is disadvantageous. This comparison does not take into account the differences between the individual drive systems. For example, it must be taken into account that the "tank" of a solid propellant engine is at the same time the combustion chamber, and for this reason has to withstand high pressures and at the same time high thermal loads, which, in particular in the case of very large stages with a small burnout mass ratio, results in relatively large structural masses and hence in higher minimum burnout mass ratios than in the case of liquid or gel propellants. An additional point is that, because of the cavity along the longitudinal axis in the centre of the fuel block, the maximum propellant filling level of solid propellant engines is smaller than that of liquid propellant stages.

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A further important aspect of the system can be seen in a comparison between mono- and bipropellants in liquid and gel form: in the former case, only one substance has to be stored and fed within a rocket stage, but in the latter case two. For this reason, monopropellant systems are significantly less complex than bipropellant systems. An additional point is that none of the monopropellants according to the invention is cryogenic. This similarly simplifies handling considerably. Unlike other storable propellants such as MON or hydrazine, the high-energy monopropellants according to the invention are not toxic, carcinogenic or hazardous to the environment.

If these aspects of the system are taken into account, drive systems that are operated using the propellants according to the invention have the potential to be superior to many conventional drive systems. In this context, attention should be drawn in particular to the propellant according to Example 7, which has the performance of bipropellants at the same time as the advantages of the system: thrust control and engine re-ignition can be realised easily. As described above, this is far more complex with solid propellant engines.

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Claims

1. A propellant in liquid or gel form, in particular for rocket engines, comprising: - an inorganic salt as the oxidising agent, wherein the inorganic salt has an oxygen content of at least 60 weight%; and - a solvent as the fuel, comprising a monovalent or divalent alcohol, a nitroalkane and/or an ionic liquid in which nitrate is the anion, wherein the solvent has an oxygen content of from 30 to 55 weight%, wherein the inorganic salt is dissolved in the solvent.

2. A propellant according to Claim 1, in which the inorganic salt is selected from lithium nitrate, lithium dinitramide, lithium perchlorate or a mixture thereof.

3. A propellant according to Claim 1 or 2, in which the proportion of inorganic salt in the propellant is from 15 to 65 weight%.

4. A propellant according to one of the preceding claims, in which the solvent comprises ethanol, methanol or n-butanol, in particular as a single constituent, and in which the proportion of inorganic salt in the propellant is preferably from to 65 weight%.

5. A propellant according to one of the preceding claims, in which the solvent comprises nitromethane or nitroethane, in particular as a single constituent, and in which the proportion of inorganic salt in the propellant is preferably from 10 to weight%, more preferably from 20 to 30 weight%.

6. A propellant according to Claim 4 or 5, in which the solvent additionally comprises a further alcohol, in particular n-butanol and/or ethylene glycol, and/or a carbonate ester, in particular dimethyl carbonate, diethyl carbonate, ethyl methyl carbonate and/or propylene carbonate.

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7. A propellant according to one of Claims 1 to 3, in which the solvent comprises an ionic liquid, in particular ethylammonium nitrate, and in which the proportion of inorganic salt in the propellant is preferably from 10 to weight%, more preferably from 15 to 25 weight%.

8. A propellant according to Claim 7, in which the solvent additionally comprises an alcohol, in particular ethylene glycol and/or ethanol, preferably in a weight ratio of ethylammonium nitrate to alcohol in the range of from 6:1 to 1:3.

9. A propellant according to one of the preceding claims, in which the propellant contains no water.

10. A propellant according to one of the preceding claims, in which the propellant further comprises a thickening agent, which is preferably selected from polyacrylic acids, pyrogenic silicon dioxides, micro- to nano-scale metal powders, titanium dioxide nanoparticles and/or carbon nanotubes.

11. A propellant according to Claim 10, in which the metal powder is selected from aluminium, magnesium, aluminium/magnesium alloys, boron, iron and zirconium.

12. A propellant according to Claim 10 or 11, in which the proportion of thickening agent is up to 10 weight%, preferably from 1 to 5 weight%.

13. A propellant according to one of the preceding claims, in which the propellant further comprises one or more hydrides of light metals, which are preferably selected from AH 3, NaBH 4 and/or AILiH 4 .

14. A propellant according to one of the preceding claims, further comprising a further oxidising agent, which is selected from the nitrates and perchlorates of ammonium, sodium and potassium.

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15. A propellant according to one of the preceding claims, in which the propellant has a density of from 900 to 1 700 kg/m 3, preferably from 1 100 to 1 400 kg/M 3; and/or in which, for combustion, the propellant has an oxygen balance of from 0 to -50%, preferably from -20 to -40%.

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