WO2022171324A1

WO2022171324A1 - Fuel

Info

Publication number: WO2022171324A1
Application number: PCT/EP2021/082929
Authority: WO
Inventors: Maxim KURILOV; Dominic FREUDENMANN; Christoph KIRCHBERGER
Original assignee: Deutsches Zentrum für Luft- und Raumfahrt e.V.
Priority date: 2021-02-12
Filing date: 2021-11-25
Publication date: 2022-08-18
Also published as: KR20240007899A; US20240124372A1; AU2021427742A1; EP4291545A1; DE102021103380B3; JP2024509729A

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

fuel

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

When it comes to propellants for rocket engines that work according to the recoil principle, a distinction is primarily made between solid and liquid propellant systems. Depending on the specific area of application (launchers, orbital engines, etc.), different fuel systems can be used to advantage, but all systems also have specific disadvantages.

Solid propellants are characterized in particular by a very high energy density and are widely used in aerospace and military technology, e.g. in carrier rockets and military missiles. Mixtures of solid fuels and oxidizers can be used (diergolic systems) as well as monergole fuels, in which an intramolecular redox reaction takes place (e.g. fuels based on HMX, RBX or a mixture of nitroglycerine and nitrocellulose). Solid propellant rocket engines are simple in design because the fuel is injected directly into the combustion chamber, eliminating the need for a separate fuel tank and fuel delivery systems.

The main disadvantage of solid propellants is the lack of flexibility, both in the scaling of the engines and in operation, since it is practically impossible to vary the thrust by controlling the burnup. Another problem is that the known solid propellants are highly explosive, which makes the safe handling of the propellants before, during and after introduction into the combustion chamber very complex. When introducing the solid propellant, a form-fitting connection with the combustion chamber wall must be ensured, whereby defects or crack formation can lead to uncontrolled combustion and, in extreme cases, to the total loss of the vehicle. In the case of liquid propellants, the disadvantages mentioned above do not occur; in particular, rocket engines with liquid propellants are much more flexible than solid propellants. It is easier to scale the engines, as well as to control the thrust by regulating the fuel supply to the combustion chamber. On the other hand, this means the disadvantage of a higher design effort, since one or more propellant tanks and supply systems with pumps, valves and other auxiliary components are required.

Monergole and diergole systems are also known for liquid fuels. The latter include cryogenic or partially cryogenic fuels with liquid hydrogen, liquid methane or kerosene as the fuel and liquid oxygen as the oxidizer. The storage and handling of these liquefied gases is expensive and requires the highest safety measures, since any leakage leads to a risk of explosion.

In addition to cryogenic systems, combinations of liquid fuels and oxidizers are known, which are used in particular in orbital engines for flight and attitude control of satellites or space probes. Hydrazine and its derivatives (mono- and dimethylhydrazine) are typically used as fuel in this context, in combination with nitric acid, dinitrogen tetroxide or hydrogen peroxide. The main problem here is that hydrazines are very toxic and carcinogenic, so for health and environmental reasons they should be replaced by alternatives whenever possible. Dinitrogen tetroxide is also toxicologically questionable, but replacing it with hydrogen peroxide reduces the effectiveness of the fuel.

Ammonium dinitramide (ADN) and hydroxylammonium nitrate (HAN) are used as less toxic alternatives to hydrazines. These monergolic substances are highly explosive, so they can only be used in the form of aqueous solutions. gene are manageable, optionally in combination with methanol and/or ammonia as additional fuels. The water content in turn leads to reduced ignitability, so that engines based on ADN or HAN, in contrast to hydrazine engines, cannot be cold-started, but have to be heated up beforehand. In addition, ADN and HAN are comparatively expensive.

The invention is based on the object of proposing a propellant, in particular for rocket engines, with 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 liquid or gelatinous propellant which comprises the following: an inorganic salt as an oxidizer, the inorganic salt having an oxygen content of at least 60% by weight; and a solvent as a fuel, comprising a monohydric or dihydric alcohol, a nitroalkane and/or an ionic liquid with nitrate as the anion, the solvent having an oxygen content of 30 to 55% by weight, the inorganic salt being dissolved in the solvent present.

The propellant according to the invention is liquid or gel-like, and in the case of a gel-like propellant it is also pumpable (see below). This avoids the typical disadvantages of solid propellants. The propellant of the present invention is non-cryogenic, which fundamentally simplifies the storage, handling, and delivery of the propellant to the engine. Finally, the components of the fuel according to the invention are toxicologically and ecologically relatively harmless, at least in comparison to hydrazine and its derivatives. Compared to ADN or HAN, the fuel according to the invention offers a clear cost advantage.

Although the fuel according to the invention is a diergolic system, since oxidizer and fuel are present as separate chemical compounds, it is but advantageously a homogeneous mixture that is fed to the engine. The invention makes use of the fact that the inorganic salts used as oxidizers have relatively good solubility in various solvents that are suitable as fuels, so that the required mixing ratios of oxidizer and fuel can be set. In this context, it is of particular advantage that both low-energy fuels (monohydric or dihydric alcohols) and high-energy fuels (nitroalkanes/nitrates) can be used as solvents in order to optimize the specific properties of the fuel such as oxygen balance, energy density, specific impulse etc. to be able to adapt to different areas of application and requirements.

The propellant according to the invention and its components are generally not explosive, which also simplifies handling and production and increases the safety of the propellant. On the other hand, the fuel according to the invention has good ignition properties, with electrical, thermal, catalytic-thermal or catalytic ignition being possible in principle.

The inorganic salt used as the oxidizer is preferably selected from lithium nitrate, lithium dinitramide, lithium perchlorate or a mixture thereof. These salts have relatively good solubility in several suitable fuels, particularly alcohols. For example, the solubility of lithium nitrate in methanol is about 58 g/100 g and the solubility of lithium perchlorate in methanol is about 182 g/100 g. The solubilities are lower in nitroalkanes and ionic liquids with nitrate as the anion, and in this case it is possible to increase the solubility of the salts by adding alcohols or other solvents.

The proportion of the inorganic salt in the fuel can be varied over a wide range, it is usually in the range from 15 to 65% by weight. The proportion depends on the one hand on the chosen solvent and the previously discussed solubility of the inorganic salt, on the other hand it can be varied depending on the desired oxygen balance during combustion of the fuel.

In a preferred embodiment of the invention, the solvent comprises ethanol, methanol or n-butanol, in particular as the sole component. Due to the good solubility in these monohydric alcohols, a relatively high proportion of the inorganic salt in the fuel can be selected in this case, preferably from 50 to 65% by weight.

In a further advantageous embodiment of the invention, the solvent comprises nitromethane or nitroethane, in particular as the sole component. Because of the somewhat lower solubility, the proportion of the organic salt in this case is preferably from 10 to 40% by weight, more preferably from 20 to 30% by weight.

As previously mentioned, the fuel of the present invention may comprise a single alcohol or a single nitroalkane as the solvent and fuel, respectively. According to a further embodiment of the invention, the solvent additionally comprises another alcohol, in particular n-butanol (if this is not the primary solvent) and/or ethylene glycol and/or a carbonic acid ester, in particular dimethyl carbonate, diethyl carbonate, ethyl methyl carbonate and/or propylene carbonate. Carbonic acid 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 the inorganic salt in the fuel is preferably from 10 to 40% by weight, more preferably from 15 to 25% by weight.

In addition to the ionic liquid, the solvent preferably includes an alcohol, in particular ethylene glycol and/or ethanol. Such a mixture can affect the solubility of the inorganic salt (oxidant). a single solvent can be increased. In the case of ethylammonium nitrate, the mixing ratio of ethylammonium nitrate to alcohol is preferably in the range from 6:1 to 1:3.

The fuel according to the invention preferably does not contain any water, and as a result differs in particular from the known fuels based on ADN or HAN. Due to the absence of water, the fuels according to the invention have good ignitability.

According to a further preferred embodiment, the fuel according to the invention is a gel-type fuel. In order to achieve a gel-like consistency, the propellant comprises a thickener, which is preferably selected from polyacrylic acids, pyrogenic silicon dioxides, microscale to nanoscale metal powders, titanium dioxide nanoparticles and/or carbon nanotubes. In the case of metal powder, this is preferably selected from aluminum, magnesium, aluminum-magnesium alloys, boron, iron and zirconium.

Compared to liquid fuels, gel-like propellants have the advantage that they tend to be safer, on the one hand because the vapor pressure of the liquid components is reduced, and on the other hand because the exit velocity in the event of a leak is lower due to the higher viscosity. A further advantage is that in a gel-like propellant, insoluble components can also be kept in a suspension, which would settle out in a liquid propellant. This applies in particular to the metal powders optionally contained in the propellant, which, in addition to their function as a thickening agent, also represent an additional fuel and can be used to increase the energy density of the propellant.

The proportion of thickener in the fuel according to the invention is preferably up to 10% by weight, more preferably from 1 to 5% by weight. The type and quantity of the thickening agent can advantageously be selected in such a way that the gel-like propellant according to the invention essentially has all the advantages of a liquid propellant, ie in particular a good one Pumpability and flexibility (easy scalability, throttle controllability and reignitability). As a result, the gel-like propellants according to the present invention differ significantly from known gel propellants from the prior art, which have a rheology with a structurally viscous (shear-thinning) behavior and a pronounced yield point, so that they can be used with the delivery systems customary for liquid propellants are not pumpable.

The fuel according to the invention can also comprise one or more hydrides of light metals, which are preferably selected from AlH3, NaBH4 and/or AlUH4. The metal hydrides are additional fuels that can be used to modify the energy content and performance of the fuel.

According to a further embodiment of the invention, the propellant also comprises a further oxidizer which is preferably selected from the nitrates and perchlorates of ammonium, sodium and potassium. This is particularly true in the case of gel fuels in which these oxidizers are in suspended form due to their lower solubility.

The fuel according to the invention typically has a density in the range from 900 to 1700 kg/m ³ , preferably in the range from 1100 to 1400 kg/m ³ .

When combusted, the fuel according to the invention advantageously has an oxygen balance of 0 to -50%, more preferably of -20 to -40%. With an oxygen balance of 0, combustion is completely stoichiometric, so that the energy content of the fuel is fully utilized. However, a negative oxygen balance, ie an excess of fuel compared to the oxidizer, is preferred in most cases in order to avoid the fuel spontaneously igniting (explosively) too easily. Due to the possibilities described above of varying the qualitative and quantitative composition of the liquid or gel fuel within the scope of the present invention, the specific impulse of the fuel can also lie within a wide range (e.g. in the range 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 can be used in various types of rocket engines in the aerospace tech technology, both for main drives and for auxiliary drives, especially for launchers, booster rockets, rocket stages or orbital engines. In addition, propellants with a specific impulse at the lower end of the above range can also be used to operate gas generators in aerospace systems.

Outside of space travel, the fuel according to the invention can also be used to power aircraft (e.g. for auxiliary starting units) or civil or military missiles.

Another advantageous area of application for the propellant according to the invention is mining, where the propellant can be used, for example, for cutting torches or drills. But the fuel could also be used outside of mining, e.g. to drive machine tools for joining and cutting metals.

These and other advantages of the invention are explained in more detail using the following examples.

Examples of liquid fuels

Table 1 below shows the percentage composition, specific impulse, density, adiabatic combustion temperature, ture and the oxygen balance are given. Conventional fuels based on hydrazine (VI) or ammonium dinitramide (V2 and V3) serve as comparative examples.

Table 1

The values for the specific impulse (at a combustion pressure of 5 MPa and an expansion ratio of 50:1) are comparable to those of conventional fuels or in some cases even higher. The density of the fuels according to the invention is also in a similar range. The adiabatic combustion temperature, like the specific impulse, was calculated using the NASA-CEA program (McBride & Gordon, 1996).

In Examples 2 to 4, this value is also in a similar range to that in the comparative examples. This means that very similar materials can be used to build the engines, which simplifies the technical implementation of the new fuels. Example 1 is an exception. Here, both the combustion temperature and the power (specific impulse) are significantly higher, so that it may be necessary to adapt the existing engine technologies (high-temperature-resistant construction materials, in particular of catalytic converter devices), which, however, is due to the realization of the performance potential seems worthwhile.

The oxygen balance of the fuels according to the invention is lower than with the conventional "Green Propellants" based on ADN. This means on the one hand that the combustion is less stoichiometric, which does not lead to a complete conversion of the chemical energy into propulsion energy; on the other hand, this can be an indication of this be that the newly developed propellants are more difficult to detonate under mechanical and thermal stresses.

FIG. 1 shows a diagram with the performance potential (Delta v) of the fuels listed in Table 1, expressed as a percentage deviation compared to the reference fuel hydrazine (VI), in different space vehicle configurations. These are represented by the burnout mass ratio (BMV, z), which is plotted on the right-hand axis. For example, a BMV of 0.55 corresponds to a typical spacecraft fueled with hydrazine (i.e., the fuel occupies 45% of the total spacecraft mass); a BMV of 0.92 would correspond to an earth observation satellite fueled with hydrazine.

By using denser fuel, more fuel can be carried with the same tank volume, ie the BMV decreases and the delta v des spacecraft, ie the amount of orbital velocity adjustments required for orbital maneuvers, is increasing. The lines in the graph represent how much more delta v can be gained using propellants other than hydrazine in the same spacecraft. With conventional propellants based on ADN (V2 and V3), between 30 and 50% more delta v can be achieved, thus extending the service life of probes and satellites by a factor of up to 1.5. The liquid propellants of the invention are in a similar range or above (Ex. 1) and are therefore competitive with the conventional hydrazine substitutes.

Examples of aelförmiae propellants

Table 2 below gives the percentage composition, specific impulse, density and C* combustion efficiency for three examples of gel-like propellants according to the invention (Ex. 5 to 7). Various conventional fuels (C4 to C7) serve as comparative examples.

Table 2

The specific impulse values here are lower than for the cryogenic or semi-cryogenic bipropellant propellants and more comparable to the energy characteristics of solid propellants. However, the densities of the propellants according to the invention are higher, but not as high as the density of solid propellants.

FIG. 2 shows a diagram with the performance potential (Delta v) of the propellants listed in Table 2, expressed as a percentage deviation compared to comparative example C4, in different spacecraft configurations. The BMV values shown are typical for missiles and sounding rockets (0.15 to 0.35), as well as for booster stages (0.3 to 0.45) or upper stages (0.4 to 0.65). With the same stage size, all the propellants shown in the diagram would provide less delta v than the solid propellant reference (V4), which is used, for example, in the P-80 booster stage of the VEGA carrier system. In the case of cryogenic or partially cryogenic fuels, the delta v is between 1 and 20% less. In the case of the fuels according to the invention, the delta v is between 12 and 25% less. Example 7 is an exception: the delta-v performance of this fuel is in the range of conventional cryogenic and partially cryogenic bipropellants.

However, the fact that all of the fuels examined had a lower delta v than the reference fuel does not mean that their use is disadvantageous. What is not taken into account in this comparison are the differences between the individual drive systems. For example, it must be taken into account that the "tank" of a solid propellant engine is also the combustion chamber and for this reason must withstand high pressures and at the same time high thermal loads, which leads to higher structural masses, especially in the case of very large stages with small BMV In addition, due to the cavity in the center of the fuel block along the longitudinal axis, solid fuel engines have a lower maximum fuel filling rate than liquid fuel stages.

Another important system aspect comes into play when comparing liquid or gel mono- and bipropellants: In the first case, only one substance has to be stored and conveyed within a rocket stage, in the second case two substances. Monopropellant systems are therefore significantly less complex than bipropellant systems. In addition, none of the monopropellants according to the invention is cryogenic. This also simplifies handling considerably. Unlike other storable propellants such as MON or hydrazine, the high energy monopropellants of the present invention are not toxic, carcinogenic or environmentally hazardous. If these system aspects are considered, propulsion systems that are operated with the propellants according to the invention have the potential to be superior to many conventional propulsion systems. The propellant according to Example 7, which has the performance of bipropellants with simultaneous system advantages, deserves special mention in this context: Thrust control and re-ignition of the engine can be easily implemented. As described above, this is far more complex for solid propellant engines.

Claims

P atentClaims

1. Liquid or gel fuel, in particular for rocket engines, comprehensive

- An inorganic salt as an oxidizer, the inorganic salt having an oxygen content of at least 60% by weight; and

- A solvent as a fuel, comprising a monohydric or dihydric alcohol, a nitroalkane and/or an ionic liquid with nitrate as the anion, the solvent having an oxygen content of 30 to 55% by weight, the inorganic salt being present dissolved in the solvent .

2. Propellant according to claim 1, wherein the inorganic salt is selected from lithium nitrate, lithium dinitramide, lithium perchlorate or a mixture thereof.

3. Propellant according to claim 1 or 2, wherein the proportion of the inorganic salt in the propellant is from 15 to 65% by weight.

4. Fuel according to one of the preceding claims, in which the solvent comprises ethanol, methanol or n-butanol, in particular as the sole component, and in which the proportion of the inorganic salt in the fuel is preferably from 50 to 65% by weight.

5. Fuel according to one of the preceding claims, wherein the solvent comprises nitromethane or nitroethane, in particular as the only component, and wherein the proportion of the inorganic salt in the fuel is preferably from 10 to 40% by weight, more preferably from 20 to 30% by weight .%.

6. Fuel according to claim 4 or 5, wherein the solvent additionally comprises another alcohol, in particular n-butanol and/or ethylene glycol and/or a carbonic acid ester, in particular dimethyl carbonate, diethyl carbonate, ethyl methyl carbonate and/or propylene carbonate.

7. Propellant according to one of claims 1 to 3, wherein the solvent comprises an ionic liquid, in particular ethylammonium nitrate, and wherein the proportion of the inorganic salt in the propellant is preferably from 10 to 40% by weight, more preferably from 15 to 25% by weight .% amounts to.

8. Propellant according to claim 8, wherein 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 from 6:1 to 1:3.

9. Propellant according to one of the preceding claims, wherein the propellant contains no water.

10. Propellant according to one of the preceding claims, wherein the propellant further comprises a thickener, which is preferably selected from polyacrylic acids, pyrogenic silicon dioxides, micro- to nanoscale metal powders, titanium dioxide nanoparticles and/or carbon nanotubes.

11. Propellant according to claim 10, wherein the metal powder is selected from aluminum, magnesium, aluminum-magnesium alloys, boron, iron and zirconium.

12. Fuel according to claim 10 or 11, wherein the proportion of thickening agent is up to 10% by weight, preferably from 1 to 5% by weight.

13. Fuel according to one of the preceding claims, wherein the fuel further comprises one or more hydrides of light metals, which are preferably selected from AlH3, NaBH4 and/or AlUH4.

14. Fuel according to one of the preceding claims, further comprising a further oxidizer which is selected from the nitrates and perchlorates of ammonium, sodium and potassium.

15. Propellant according to one of the preceding claims, wherein the propellant has a density of 900 to 1700 kg/m ³ , preferably 1100 to 1400 kg/m ³ ; and/or wherein the fuel has an oxygen balance of 0 to -50% during combustion, preferably of -20 to -40%.

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