Solid rocket fuel mixture and method of use

Abstract

Solid rocket fuel mixtures and methods for propelling a rocket using the mixture are described. The solid rocket fuel mixtures include a first mixture of propellant oxidizer particles and metal fuel particles and a second mixture of water and, optionally, magnesium hydroxide particles. The first mixture is combusted to produce a first stage propulsion, and the byproducts of the combustion are combined with the second mixture for a second combustion and a second stage propulsion.

Description

ANTONIJO LICITAR ET AL. FEBRUARY 3, 2026P5892PC00SOLID ROCKET FUEL MIXTURE AND METHOD OF USE

[0001] The present application claims the benefit of US provisional application 63 / 753,665 filed on February 4th, 2025.BACKGROUND

[0002] Solid propellants have been used since the 13thcentury invention of gunpowder. Solid propellants were continually developed and were commonly used as a rocket fuel including in aerospace applications. Solid rocket fuels provide the benefits of a high propellant density and allow for easy storage and handling. Additionally, solid rocket fuels also require a simpler engine design as compared to liquid rocket fuels. However, solid rocket fuels also present problems in propellant efficiency and in controlled use. The solid fuels produce high levels of pollutants, have a low specific impulse, and are not able to be throttled once in use. As a result, liquid rocket fuels were developed to address the deficiencies in solid rocket fuels.

[0003] Liquid rocket fuels have a higher specific impulse than solid rocket fuels allowing for more efficient fuel use. Furthermore, the flow of the liquid fuel is controllable allowing for the throttling, shutting down, and restarting of the rocket engine. Liquid rocket fuels must be stored at a low temperature which allows for the use of the fuels as a coolant for the rocket engines. These advantages led to the widespread use of liquid rocket propellants, including liquid hydrogen and liquid methane, in main rocket engines. However, liquid rocket fuels are more reactive than solid rocket fuels. Liquid rocket fuels have a lower density than solid rocket fuels and require more storage space while increasing the mass of the rocket. Liquid fuels additionally may leak and require a more complex storage system and rocket engine to prevent this potential issue.

[0004] There exists a need for an improvement in rocket fuel technology to address the deficiencies in both solid rocket fuels and liquid rocket fuels.SUMMARY

[0005] In some aspects, the techniques described herein relate to a solid rocket fuel mixture including: a first mixture including propellant oxidizer nanoparticles having an average particle size of about 1 nm to about 900 nm and magnesium nanoparticles having an average particle size of about 1 nm to about 900 nm configured to react in a first stage propulsion andproduce a propellant oxidizer byproduct, and a second mixture including water configured to react with the propellant oxidizer byproduct in a second stage propulsion.

[0006] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the propellant oxidizer includes one of silica, ammonium nitrate, N2O, KNO3, NH4CIO4, H2O2, MgO, lithium perchlorate (l_iCI04) , sodium perchlorate (NaCI04) , potassium perchlorate (KCI04), barium peroxide (BaO2), ozone (O3), liquid oxygen (LOX), tetranitromethane (C(NO2)4), nitronium perchlorate (NO2CIO4), Fe2O3, CuO, Co3O4, MnO2and CaO.

[0007] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the propellant oxidizer byproduct includes silicon.

[0008] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the second mixture further includes magnesium hydroxide configured to react with the propellant oxidizer and the water.

[0009] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the magnesium hydroxide includes magnesium hydroxide nanoparticles with an average particle size of about 1 nm to about 900 nm.

[0010] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the propellant oxidizer nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

[0011] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the water and the magnesium nanoparticles are provided in a molar ratio of about 1:1.

[0012] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the propellant oxidizer nanoparticles and the magnesium nanoparticles are provided in a molar ratio of about 1:3 to about 3:1.

[0013] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the propellant oxidizer nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

[0014] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the magnesium nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

[0015] In some aspects, the techniques described herein relate to a solid rocket fuel mixture, wherein the second mixture is configured to react with the propellant oxidizer byproduct to form hydrogen gas.

[0016] In some aspects, the techniques described herein relate to a method of propelling a rocket, the method including: providing a first mixture including propellant oxidizer nanoparticles having an average particle size of about 1 nm to about 900 nm and magnesiumnanoparticles having an average particle size of about 1 nm to about 900 nm; combusting the first mixture to produce a first stage propulsion and to produce a propellant oxidizer byproduct; providing a second mixture including water; and combusting the second mixture and the propellant oxidizer byproduct to produce a second stage propulsion.

[0017] In some aspects, the techniques described herein relate to a method, wherein the propellant oxidizer includes one of silica, ammonium nitrate, N2O, KNO3, NH4CIO4, H2O2, MgO, lithium perchlorate (l_iCI04) , sodium perchlorate (NaCI04) , potassium perchlorate (KCI04), barium peroxide (BaO2), ozone (O3), liquid oxygen (LOX), tetranitromethane (C(NO2)4), nitronium perchlorate (NO2CIO4), Fe2O3, CuO, Co3O4, MnO2and CaO.

[0018] In some aspects, the techniques described herein relate to a method, wherein the propellant oxidizer byproduct includes silicon.

[0019] In some aspects, the techniques described herein relate to a method, wherein the second mixture further includes magnesium hydroxide configured to react with the propellant oxidizer and the water.

[0020] In some aspects, the techniques described herein relate to a method, wherein the magnesium hydroxide includes magnesium hydroxide nanoparticles with an average particle size of about 1 nm to about 900 nm.

[0021] In some aspects, the techniques described herein relate to a method, wherein the propellant oxidizer nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

[0022] In some aspects, the techniques described herein relate to a method, wherein the water and the magnesium nanoparticles are provided in a molar ratio of about 1:1.

[0023] In some aspects, the techniques described herein relate to a method, wherein the propellant oxidizer nanoparticles and the magnesium nanoparticles are provided in a molar ratio of about 1 :3 to about 3: 1.

[0024] In some aspects, the techniques described herein relate to a method, wherein the propellant oxidizer nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

[0025] In some aspects, the techniques described herein relate to a method, wherein the magnesium nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

[0026] In some aspects, the techniques described herein relate to a method, wherein combusting the second mixture and the propellant oxidizer byproduct forms hydrogen gas.

[0027] In some aspects, the techniques described herein relate to a method, further including burning the hydrogen gas to produce additional propulsion.

[0028] In some aspects, the techniques described herein relate to a method, wherein combusting the first mixture is produces a temperature of at least about 2,000 °C.BRIEF DESCRIPTION OF THE FIGURESFigure 1 schematically illustrates an example of a method of propelling a rocket.DETAILED DESCRIPTION

[0029] Embodiments are directed towards a solid rocket fuel mixture including a first mixture including propellant oxidizer nanoparticles and metal fuel nanoparticles. In some embodiments, the metal fuel and the propellant oxidizer can be configured to combust to produce a first stage propulsion. In some embodiments, the metal fuel includes one of magnesium or aluminum. The use of the metal fuel with the propellant oxidizer provides a lighter and higher energy density reaction as compared to traditional rocket fuels. The combustion produces byproducts that may be used in further reactions, including a propellant oxidizer byproduct. In some embodiments, the propellant oxidizer nanoparticles include silica nanoparticles, and the propellant oxidizer byproduct includes silicon. The solid rocket fuel mixture can further includes a second mixture including water. The water can be configured to react with the propellant oxidizer byproduct in a second exothermic reaction to produce a second stage propulsion. The reaction of the water and the propellant oxidizer byproduct may produce hydrogen gas that may be burned to provide additional propulsion. The high surface area to volume ratio provided by the metal fuel and propellant oxidizer nanoparticles can increase the reaction speed during ignition, thereby increasing the propulsion of a rocket. The dry rocket fuel composition may be adaptable to function efficiently in both atmospheric and vacuum conditions. This versatility may make the propulsion system suitable for various mission profiles, from launch to deep space exploration. In some embodiments, the use of metal fuel, such as magnesium, and silicon dioxide nano powders may enable in-situ resource utilization on celestial bodies such as the Moon or Mars. The abundance of these materials on planetary surfaces may allow for potential refueling of rockets using locally sourced materials, reducing the need to carry all required fuel from Earth.

[0030] The propellant oxidizer nanoparticles may include any material known to one of skill in the art effective to enhance the burning of the solid rocket fuel mixture and to provide a source for oxygen. In some embodiments, the propellant oxidizer may be selected to reduce the amount of waste after the second stage propulsion. In some embodiments, the propellant oxidizer may be selected to form a propellant oxidizer byproduct effecting for reacting with water to form hydrogen gas. In some embodiments, the propellant oxidizer can include one or more of silica, ammonium nitrate, N2O, KNO3, NH4CIO4, H2O2, MgO, lithium perchlorate (LiCI04) , sodium perchlorate (NaCI04) , potassium perchlorate (KCI04), barium peroxide (BaO2), ozone (O3), liquid oxygen (LOX), tetranitromethane (C(NO2)4), nitronium perchlorate (NO2CIO4), Fe2O3, CuO, Co3O4, MnO2and CaO. In some embodiments, the propellant oxidizer includes silica nanoparticles.

[0031] The propellant oxidizer may be present in the rocket fuel mixture at any amount of the rocket fuel mixture. In some embodiments, the amount of the propellant oxidizer is selected to achieve a stoichiometric ratio with the metal fuel nanoparticles. In some embodiments, thepropellant oxidizer nanoparticles and the metal fuel nanoparticles are provided in a molar ratio of about 1:5, about 1:4, about 1:3, about 1:2, about 1:1, about 2:1, about 3:1, about 4:1, about 5:1, or any value or range of values between any two of these values. In some embodiments, the propellant oxidizer nanoparticles and the metal fuel nanoparticles are provided in a molar ratio of about 1 :3 to about 3: 1.

[0032] The propellant oxidizer nanoparticles may have any average particle size effective for combustion as a rocket fuel. In some embodiments, the average particle size may be the average particle diameter. In some embodiments, the propellant oxidizer nanoparticles have an average particle size of about 1 nm, about 2 nm, about 3 nm, about 4 nm, about 5 nm, about 6 nm, about 7 nm, about 8 nm, about 9 nm, about 10 nm, about 15 nm, about 20 nm, about 25 nm, about 30 nm, about 35 nm, about 40 nm, about 45 nm, about 50 nm, about 100 nm, about 150 nm, about 200 nm, about 250 nm, about 300 nm, about 350 nm, about 400 nm, about 450 nm, about 500 nm, about 550 nm, about 600 nm, about 650 nm, about 700 nm, about 750 nm, about 800 nm, about 850 nm, about 900 nm, about 950 nm, about 1000 nm, or any value or range of values between any two of these values. In some embodiments, the propellant oxidizer nanoparticles have an average particle size of about 1 nm to about 900 nm, about 50 nm to about 800 nm, or about 100 nm to about 500 nm.

[0033] The propellant oxidizer nanoparticles may have any surface area effective for combustion as a rocket fuel. In some embodiments, the propellant oxidizer nanoparticles have an average particle size of about 1 m2 / g, about 2 m2 / g, about 3 m2 / g, about 4 m2 / g, about 5 m2 / g, about 6 m2 / g, about 7 m2 / g, about 8 m2 / g, about 9 m2 / g, about 10 m2 / g, about 15 m2 / g, about 20 m2 / g, about 25 m2 / g, about 30 m2 / g, about 35 m2 / g, about 40 m2 / g, about 45 m2 / g, about 50 m2 / g, about 100 m2 / g, about 150 m2 / g, about 200 m2 / g, about 250 m2 / g, about 300 m2 / g, about 350 m2 / g, about 400 m2 / g, about 450 m2 / g, about 500 m2 / g, about 600 m2 / g, about 700 m2 / g, about 800 m2 / g, about 900 m2 / g, about 1,000 m2 / g, or any value or range of values between any two of these values. In some embodiments, the propellant oxidizer nanoparticles have an average surface area of about 1 m2 / g to about 500 m2 / g, about 50 m2 / g to about 400 m2 / g, or about 100 m2 / g to about 300 m2 / g.

[0034] In some embodiments, the metal fuel nanoparticles comprise one of aluminum nanoparticles or magnesium nanoparticles. The metal fuel nanoparticles may have any average particle size effective for combustion as a rocket fuel. In some embodiments, the average particle size may be the average particle diameter. In some embodiments, the metal fuel nanoparticles have an average particle size of about 1 nm, about 2 nm, about 3 nm, about 4 nm, about 5 nm, about 6 nm, about 7 nm, about 8 nm, about 9 nm, about 10 nm, about 15 nm, about 20 nm, about 25 nm, about 30 nm, about 35 nm, about 40 nm, about 45 nm, about 50 nm, about 100 nm, about 150 nm, about 200 nm, about 250 nm, about 300 nm, about 350 nm, about 400 nm, about 450 nm, about 500 nm, about 550 nm, about 600 nm, about 650 nm,about 700 nm, about 750 nm, about 800 nm, about 850 nm, about 900 nm, about 950 nm, about 1000 nm, or any value or range of values between any two of these values. In some embodiments, the metal fuel nanoparticles have an average particle size of about 1 nm to about 900 nm, about 50 nm to about 600 nm, about 10 nm to about 500 nm, or about 100 nm to about 500 nm.

[0035] The metal fuel nanoparticles may have any surface area effective for combustion as a rocket fuel. In some embodiments, the metal fuel nanoparticles have an average particle size of about 1 m2 / g, about 2 m2 / g, about 3 m2 / g, about 4 m2 / g, about 5 m2 / g, about 6 m2 / g, about 7 m2 / g, about 8 m2 / g, about 9 m2 / g, about 10 m2 / g, about 15 m2 / g, about 20 m2 / g, about 25 m2 / g, about 30 m2 / g, about 35 m2 / g, about 40 m2 / g, about 45 m2 / g, about 50 m2 / g, about 100 m2 / g, about 150 m2 / g, about 200 m2 / g, about 250 m2 / g, about 300 m2 / g, about 350 m2 / g, about 400 m2 / g, about 450 m2 / g, about 500 m2 / g, about 600 m2 / g, about 700 m2 / g, about 800 m2 / g, about 900 m2 / g, about 1,000 m2 / g, or any value or range of values between any two of these values. In some embodiments, the metal fuel nanoparticles have an average surface area of about 1 m2 / g to about 500 m2 / g, about 50 m2 / g to about 400 m2 / g, or about 100 m2 / g to about 300 m2 / g.

[0036] In some embodiments, the first mixture further includes one or more burn rate modifiers. The one or more burn rate modifiers may include any material effective for controlling a combustion of the first mixture. In some embodiments, the one or more burn rate modifiers include one or more of carbon nanotubes, transitional metal oxides, CuO, Fe2Oa, or graphene oxide.

[0037] In some embodiments, the first mixture further includes one or more combustion catalysts. The one or more combustion catalysts may include any material effective for increasing a reaction speed of a combustion of the first mixture. In some embodiments, the one or more combustion catalysts include one or more of MoOa or V2O3.

[0038] In some embodiments, the first mixture may include a binding agent to enhance the structural integrity of the first mixture. The binding agent may include any material effective for improving the structural integrity of a powder mixture known to one of skill in the art. In some embodiments, the binding agent may include one or more of polymeric binders and inorganic binders. In some embodiments, the binding agent includes an energetic binder configured to increase a second combustion of the second mixture. In some embodiments, the binding agent includes one or more of polyvinylidene fluoride (PVDF), polytetrafluoroethylene (PTFE), or hydroxyl-terminated polybutadiene (HTPB).

[0039] In some embodiments, the first mixture further includes one or stabilizers. The one or more stabilizers may include any material effective for preventing a premature reaction of the first mixture. In some embodiments, the one or more stabilizers include one or more of dicyandiamide (DCD) or a metal chelate.

[0040] The water may be provided in any amount effective for combustion as a rocket fuel. In some embodiments, the amount of the water is selected to achieve a stoichiometric ratio with the propellant oxidizer byproduct. In some embodiments, the water can be provided in a molar ratio as compared to the propellant oxidizer byproduct of about 1 :5, about 1 :4, about 1 :3, about 1 :2, about 1:1, about 2: 1 , about 3: 1 , about 4: 1 , about 5:1, or any value or range of values between any two of these values. In some embodiments, the water can be provided in a molar ratio as compared to the propellant oxidizer byproduct of about 1:3 to about 3:1.

[0041] In some embodiments, the second mixture may further include magnesium hydroxide. The magnesium hydroxide may be configured to react with the water and the propellant oxidizer byproduct to produce an exothermic reaction and the second stage propulsion. The magnesium hydroxide may be present in the second mixture at any amount effective for combustion as a rocket fuel. In some embodiments, the amount of the magnesium hydroxide is selected to achieve a stoichiometric ratio with the propellant oxidizer byproduct and the water. In some embodiments, the magnesium hydroxide and the water are provided in a molar ratio of about 1 :5, about 1 :4, about 1 :3, about 1 :2, about 1:1, about 2: 1 , about 3: 1 , about 4:1, about 5:1, or any value or range of values between any two of these values. In some embodiments, the magnesium hydroxide and the water are provided in a molar ratio of about 1:3 to about 3:1.

[0042] In some embodiments, the magnesium hydroxide includes magnesium hydroxide nanoparticles. In some embodiments, the average particle size may be the average particle diameter. In some embodiments, the magnesium hydroxide nanoparticles have an average particle size of about 1 nm, about 2 nm, about 3 nm, about 4 nm, about 5 nm, about 6 nm, about 7 nm, about 8 nm, about 9 nm, about 10 nm, about 15 nm, about 20 nm, about 25 nm, about 30 nm, about 35 nm, about 40 nm, about 45 nm, about 50 nm, about 100 nm, about 150 nm, about 200 nm, about 250 nm, about 300 nm, about 350 nm, about 400 nm, about 450 nm, about 500 nm, about 550 nm, about 600 nm, about 650 nm, about 700 nm, about 750 nm, about 800 nm, about 850 nm, about 900 nm, about 950 nm, about 1000 nm, or any value or range of values between any two of these values. In some embodiments, the magnesium hydroxide nanoparticles have an average particle size of about 1 nm to about 900 nm, about 20 nm to about 900 nm, about 50 nm to about 600 nm, or about 100 nm to about 500 nm.

[0043] The magnesium hydroxide nanoparticles may have any surface area effective for combustion as a rocket fuel. In some embodiments, the magnesium hydroxide nanoparticles have an average particle size of about 1 m2 / g, about 2 m2 / g, about 3 m2 / g, about 4 m2 / g, about 5 m2 / g, about 6 m2 / g, about 7 m2 / g, about 8 m2 / g, about 9 m2 / g, about 10 m2 / g, about 15 m2 / g, about 20 m2 / g, about 25 m2 / g, about 30 m2 / g, about 35 m2 / g, about 40 m2 / g, about 45 m2 / g, about 50 m2 / g, about 100 m2 / g, about 150 m2 / g, about 200 m2 / g, about 250 m2 / g, about 300 m2 / g, about 350 m2 / g, about 400 m2 / g, about 450 m2 / g, about 500 m2 / g, about 600 m2 / g, about700 m2 / g, about 800 m2 / g, about 900 m2 / g, about 1,000 m2 / g, or any value or range of values between any two of these values. In some embodiments, the magnesium hydroxide nanoparticles have an average surface area of about 1 m2 / g to about 500 m2 / g, about 50 m2 / g to about 400 m2 / g, or about 100 m2 / g to about 300 m2 / g.

[0044] The second mixture may have any pH effective for combustion as a rocket fuel. In some embodiments, the second mixture has a pH level of about 7.0, 7.1, 7.2, 7.3, 7.4, 7.5, 7.6, 7.7, 7.8, 7.9, 8.0, 8.1, 8.2, 8.3, 8.4, 8.5, 8.6, 8.7, 8.8, 8.9, 9.0, or any value between any two of these values. In some embodiments, the rocket fuel mixture has a pH level of about 7.0 to about 9.0.

[0045] In some embodiments, the second mixture further includes one or more burn rate modifiers. The one or more burn rate modifiers may include any material effective for controlling a combustion of the second mixture. In some embodiments, the one or more burn rate modifiers include one or more of carbon nanotubes, transitional metal oxides, CuO, Fe2O3, or graphene oxide.

[0046] In some embodiments, the second mixture further includes one or more combustion catalysts. The one or more combustion catalysts may include any material effective for increasing a reaction speed of a combustion of the second mixture. In some embodiments, the one or more combustion catalysts include one or more of MoOa or V2O3.

[0047] In some embodiments, the second mixture may include a binding agent to enhance the structural integrity of the second mixture. The binding agent may include any material effective for improving the structural integrity of a powder mixture known to one of skill in the art. In some embodiments, the binding agent may include one or more of polymeric binders and inorganic binders. In some embodiments, the binding agent includes an energetic binder configured to increase a second combustion of the second mixture. In some embodiments, the binding agent includes one or more of polyvinylidene fluoride (PVDF), polytetrafluoroethylene (PTFE), or hydroxyl-terminated polybutadiene (HTPB).

[0048] In some embodiments, the second mixture further includes one or stabilizers. The one or more stabilizers may include any material effective for preventing a premature reaction of the second mixture. In some embodiments, the one or more stabilizers include one or more of dicyandiamide (DCD) or a metal chelate.

[0049] In some embodiments, the second mixture further includes one or surfactants. The one or more surfactants may include any material effective for dispersing nanoparticles in the second mixture. In some embodiments, the one or more surfactants include sodium dodecyl sulfate.

[0050] In some embodiments, the second mixture further includes one or dispersants. The one or more dispersants may include any material effective for reducing agglomeration inthe second mixture. In some embodiments, the one or more dispersants include polyethylene glycol.

[0051] Embodiments can be directed towards a method of propelling a rocket using the above-described solid rocket fuel mixture fuel mixture. An example is illustrated in figure 1.

[0052] A method of propelling a rocket includes providing, at block 100, a first mixture including propellant oxidizer nanoparticles and metal fuel nanoparticles. In some embodiments, the metal fuel nanoparticles comprise one of aluminum nanoparticles or magnesium nanoparticles. In some embodiments, the propellant oxidizer includes one of silica, ammonium nitrate, N2O, KNO3, NH4CIO4, H2O2, MgO, lithium perchlorate (l_iCI04) , sodium perchlorate (NaCI04) , potassium perchlorate (KCI04), barium peroxide (BaO2), ozone (O3), liquid oxygen (LOX), tetranitromethane (C(NO2)4), nitronium perchlorate (NO2CIO4), Fe2O3, CuO, Co3O4, MnO2, and CaO.

[0053] The method may further include, at block 120, combusting the first mixture to produce a first stage propulsion. The first combustion can be an exothermic reaction, and the heat that is released that provides the first stage propulsion. In some embodiments, the first combustion is self-sustaining due to the heat generated by the exothermic process. In some embodiments, the first combustion produces a temperature of at least about 1 ,500 °C, at least about 1,550 °C, at least about 1,600 °C, at least about 1,650 °C, at least about 1,700 °C, at least about 1 ,750 °C, at least about 1 ,800°C, at least about 1 ,850°C, at least about 1 ,900°C, at least about 1 ,950 °C, at least about 2,000 °C, at least about 2,050 °C, at least about 2, 100 °C, at least about 2,150 °C, at least about 2,200 °C, at least about 2,250 °C, at least about 2,300 °C, at least about 2,350 °C, at least about 2,400 °C, at least about 2,450 °C, at least about 2,500 °C, or any value or range of values between any two of these values. In some embodiments, the first combustion has a reaction enthalpy of about 2,500 kJ / mol, about 2,600 kJ / mol, about 2,700 kJ / mol, about 2,800 kJ / mol, about 2,900 kJ / mol, about 3,000 kJ / mol, about 3,100 kJ / mol, about 3,200 kJ / mol, about 3,300 kJ / mol, about 3,400 kJ / mol, about 3,500 kJ / mol, about 3,600 kJ / mol, about 3,700 kJ / mol, about 3,800 kJ / mol, about 3,900 kJ / mol, about 4,000 kJ / mol, or any value or range of values between any two of these values.

[0054] The first combustion further can produce a metal oxide and a propellant oxidizer byproduct. In some embodiments, the propellant oxidizer is silica, and the propellant oxidizer byproduct includes silicon. In some embodiments, the metal fuel is magnesium and the metal oxide is MgO. The use of silica provides the advantage of use as an oxidizer in the first stage propulsion and providing the byproduct of silica for later stage reactions. Furthermore, the use of metal fuel nanoparticles and the propellant oxidizers provides the benefit of controlling the amount of each reactant. This can assist in managing the temperature of the reaction, which reduces the thermal stress on the rocket engine. The silica and magnesium may react as shown in Equation 1.

[0055] Equation 1: 2Mg + SiC>2 = 2MgO + Si

[0056] The method may further include, at block 130, providing a second mixture including water. The method may further include, at block 140, combusting the second mixture with the propellant oxidizer byproduct to produce a second stage propulsion. The second combustion can be an exothermic reaction, and the heat can provide the second stage propulsion. In some embodiments, the high temperature produced by the first combustion can assist in achieving the second combustion reaction. In some embodiments, the propellant oxidizer byproduct can include silica, and the second combustion can react as shown in Equation 2.

[0057] Equation 2: Si + H2O = SiC>2 + 2H2

[0058] The hydrogen gas produced by the second combustion reaction may provide additional propulsion. In some embodiments, the method may further include burning the hydrogen gas to produce additional heat and propulsion.

[0059] In some embodiments, the second mixture further includes magnesium hydroxide. In some embodiments, the magnesium hydroxide can include magnesium hydroxide nanoparticles. At the high temperatures produced by the first combustion can result in the decomposition of the magnesium hydroxide to produce magnesium oxide and water vapor. The released water vapor may increase the propellant mass for expulsion and increase the propulsion of the rocket, the propellant oxidizer byproduct includes silica, and the second combustion reacts as shown in Equation 3. Furthermore, the water vapor produced in the second combustion may be used as a coolant to assist in the thermal regulation of the rocket engine.

[0060] Equation 3: Mg(OH)2 + H2O + Si = MgO + H2 + SiC>2

[0061] The final byproducts of MgO and silica may be recycled and used to provide additional oxidizer for the reaction. This can reduce the environmental impact of the solid rocket fuel mixture as compared to traditional rocket fuels.

[0062] In some embodiments, the method may include monitoring the consumption of the metal fuel and propellant oxidizer nanoparticles. The method may further include adjusting the mixture ratio of the metal fuel and the propellant oxidizer nanoparticles to optimize the performance of the solid rocket fuel. In some embodiments, the consumption of the metal fuel and propellant oxidizer nanoparticles may be performed by a processing device.

[0063] The use of the multi-staged propulsion can allow for an improved control over propulsion during flight as well as improved fuel utilization. In some embodiments, the first combustion can be used as an initial thrust, and the second combustion can be used for sustained propulsion of the rocket. The use of the silica and metal fuel nanoparticles further allows for a precise control over the fuel consumption during flight as well as a high burn rate. In some embodiments, the burn rate of the nanoparticles is about 10 mm / s, about 11 mm / s,12 mm / s, about 13 mm / s, about 14 mm / s, about 15 mm / s, about 16 mm / s, about 17 mm / s, about 18 mm / s, about 19 mm / s, about 20 mm / s, about 21 mm / s, about 22 mm / s, about 23 mm / s, about 24 mm / s, about 25 mm / s, about 26 mm / s, about 27 mm / s, about 28 mm / s, about 29 mm / s, about 30 m / s, or any value or range of values between any two of these values. In some embodiments, the burn rate of the nanoparticles is greater than about 30 mm / s.

[0064] The multi-staged propulsion may provide a high energy density resulting in an increase in thrust-to-weight ratio as compared to traditional solid rocket fuels. In some embodiments, the multi-staged propulsion may provide an increase in thrust-to-weight ratio as compared to traditional solid rocket fuels of about 11 %, about 12 %, about 13 %, about 14 %, about 15 %, about 16 %, about 17 %, about 18 %, about 19 %, about 20 %, or any value or range of values between any two of these values.

[0065] The multi-staged propulsion pay provide a higher specific impulse as compared to traditional solid rocket fuels. In some embodiments, the multi-staged propulsion may provide a specific impulse of about 300 seconds, about 310 seconds, about 320 seconds, about 330 seconds, about 340 seconds, about 350 seconds, about 360 seconds, about 370 seconds, about 380 seconds, about 390 seconds, about 400 seconds, or any value or range of values between any two of these values.

[0066] The use of magnesium provides additional benefits as compared to traditional rocket fuels. The magnesium-based fuels may have a specific impulse of about 30 MJ / kg, about 31 MJ / kg, about 32 MJ / kg, about 33 MJ / kg, about 34 MJ / kg, about 35 MJ / kg, about 36 MJ / kg, about 37 MJ / kg, about 38 MJ / kg, about 39 MJ / kg, about 40 MJ / kg, about 41 MJ / kg, about 42 MJ / kg, about 43 MJ / kg, about 44 MJ / kg, about 45 MJ / kg, or any value or range of values between any two of these values. The magnesium-based fuels may have a specific impulse of about 30 MJ / kg to about 45 MJ / kg.

[0067] The present disclosure is not to be limited in terms of the particular embodiments described in this application, which are intended as illustrations of various aspects. Many modifications and variations can be made without departing from its spirit and scope, as will be apparent to those skilled in the art. Functionally equivalent methods and apparatuses within the scope of the disclosure, in addition to those enumerated herein, will be apparent to those skilled in the art from the foregoing descriptions. Such modifications and variations are intended to fall within the scope of the appended claims. The present disclosure is to be limited only by the terms of the appended claims, along with the full scope of equivalents to which such claims are entitled. It is to be understood that this disclosure is not limited to particular methods, reagents, compounds, compositions or biological systems, which can, of course, vary. It is also to be understood that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting.

[0068] It is to be understood that this invention is not limited to the particular processes, compositions, or methodologies described, as these may vary. It is also to be understood that the terminology used in the description is for the purpose of describing the particular versions or embodiments only and is not intended to limit the scope of the present invention, which will be limited only by the appended claims. Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art. Although any methods and materials similar or equivalent to those described herein can be used in the practice or testing of embodiments of the present invention, the preferred methods, devices, and materials are now described. All publications mentioned herein are incorporated by reference in their entirety. Nothing herein is to be construed as an admission that the invention is not entitled to antedate such disclosure by virtue of prior invention.

[0069] It must also be noted that as used herein and in the appended claims, the singular forms “a,” “an,” and “the” include plural reference unless the context clearly dictates otherwise. Thus, for example, reference to “a sandwich composite panel” is a reference to “one or more sandwich composite panels” and equivalents thereof known to those skilled in the art, and so forth.

[0070] As used herein, the term “about” means plus or minus 10% of the numerical value of the number with which it is being used. Therefore, about 50% means in the range of 45%-55%.

[0071] With respect to the use of substantially any plural and / or singular terms herein, those having skill in the art can translate from the plural to the singular and / or from the singular to the plural as is appropriate to the context and / or application. The various singular / plural permutations may be expressly set forth herein for sake of clarity.

[0072] It will be understood by those within the art that, in general, terms used herein, and especially in the appended claims (for example, bodies of the appended claims) are generally intended as “open” terms (for example, the term “including” should be interpreted as “including but not limited to,” the term “having” should be interpreted as “having at least,” the term “includes” should be interpreted as “includes but is not limited to,” et cetera). While various compositions, methods, and devices are described in terms of “comprising” various components or steps (interpreted as meaning “including, but not limited to”), the compositions, methods, and devices can also “consist essentially of” or “consist of” the various components and steps, and such terminology should be interpreted as defining essentially closed-member groups. It will be further understood by those within the art that if a specific number of an introduced claim recitation is intended, such an intent will be explicitly recited in the claim, and in the absence of such recitation no such intent is present.

[0073] For example, as an aid to understanding, the following appended claims may contain usage of the introductory phrases "at least one" and "one or more" to introduce claimrecitations. However, the use of such phrases should not be construed to imply that the introduction of a claim recitation by the indefinite articles "a" or "an" limits any particular claim containing such introduced claim recitation to embodiments containing only one such recitation, even when the same claim includes the introductory phrases "one or more" or "at least one" and indefinite articles such as "a" or "an" (for example, “a” and / or “an” should be interpreted to mean “at least one” or “one or more”); the same holds true for the use of definite articles used to introduce claim recitations.

[0074] In addition, even if a specific number of an introduced claim recitation is explicitly recited, those skilled in the art will recognize that such recitation should be interpreted to mean at least the recited number (for example, the bare recitation of "two recitations," without other modifiers, means at least two recitations, or two or more recitations). Furthermore, in those instances where a convention analogous to “at least one of A, B, and C, et cetera” is used, in general such a construction is intended in the sense one having skill in the art would understand the convention (for example, “a system having at least one of A, B, and C” would include but not be limited to systems that have A alone, B alone, C alone, A and B together, A and C together, B and C together, and / or A, B, and C together, et cetera). In those instances where a convention analogous to “at least one of A, B, or C, et cetera” is used, in general such a construction is intended in the sense one having skill in the art would understand the convention (for example, “a system having at least one of A, B, or C” would include but not be limited to systems that have A alone, B alone, C alone, A and B together, A and C together, B and C together, and / or A, B, and C together, etcetera). It will be further understood by those within the art that virtually any disjunctive word and / or phrase presenting two or more alternative terms, whether in the description, claims, or drawings, should be understood to contemplate the possibilities of including one of the terms, either of the terms, or both terms. For example, the phrase “A or B” will be understood to include the possibilities of “A” or “B” or “A and B.”

[0075] In addition, where features or aspects of the disclosure are described in terms of Markush groups, those skilled in the art will recognize that the disclosure is also thereby described in terms of any individual member or subgroup of members of the Markush group.

[0076] As will be understood by one skilled in the art, for any and all purposes, such as in terms of providing a written description, all ranges disclosed herein also encompass any and all possible subranges and combinations of subranges thereof. Any listed range can be easily recognized as sufficiently describing and enabling the same range being broken down into at least equal halves, thirds, quarters, fifths, tenths, et cetera. As a non-limiting example, each range discussed herein can be readily broken down into a lower third, middle third and upper third, et cetera. As will also be understood by one skilled in the art all language such as “up to,” “at least,” and the like include the number recited and refer to ranges that can besubsequently broken down into subranges as discussed above. Finally, as will be understood by one skilled in the art, a range includes each individual member. Thus, for example, a group having 1-3 cells refers to groups having 1, 2, or 3 cells. Similarly, a group having 1-5 cells refers to groups having 1, 2, 3, 4, or 5 cells, and so forth.

[0077] Various of the above-disclosed and other features and functions, or alternatives thereof, may be combined into many other different systems or applications. Various presently unforeseen or unanticipated alternatives, modifications, variations or improvements therein may be subsequently made by those skilled in the art, each of which is also intended to be encompassed by the disclosed embodiments.

Claims

CLAIMS1. A solid rocket fuel mixture comprising:a first mixture comprising propellant oxidizer nanoparticles having an average particle size of about 1 nm to about 900 nm and magnesium nanoparticles having an average particle size of about 1 nm to about 900 nm configured to react in a first stage propulsion and produce a propellant oxidizer byproduct, anda second mixture comprising water configured to react with the propellant oxidizer byproduct in a second stage propulsion.

2. The solid rocket fuel mixture of claim 1 , wherein the propellant oxidizer comprises one of silica, ammonium nitrate, N2O, KNO3, NH4CIO4, H2O2, MgO, lithium perchlorate (LiCI04) , sodium perchlorate (NaCI04) , potassium perchlorate (KCI04), barium peroxide (BaO2), ozone (O3), liquid oxygen (LOX), tetranitromethane (C(NO2)4), nitronium perchlorate (NO2CIO4), Fe2O3, CuO, Co3O4, MnO2and CaO.

3. The solid rocket fuel mixture of claim 1 or 2, wherein the propellant oxidizer byproduct comprises silicon.

4. The solid rocket fuel mixture of any of claims 1 - 3, wherein the second mixture further comprises magnesium hydroxide configured to react with the propellant oxidizer and the water.

5. The solid rocket fuel mixture of claim 4, wherein the magnesium hydroxide comprises magnesium hydroxide nanoparticles with an average particle size of about 1 nm to about 900 nm.

6. The solid rocket fuel mixture of claim 5 or 5, wherein the water and the magnesium nanoparticles are provided in a molar ratio of about 1:1.

7. The solid rocket fuel mixture of any of claims 1 - 6, wherein the propellant oxidizer nanoparticles and the magnesium nanoparticles are provided in a molar ratio of about 1:3 to about 3:1.

8. The solid rocket fuel mixture of any of claims 1 - 7, wherein the propellant oxidizer nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

9. The solid rocket fuel mixture of claim 1 - 8, wherein the magnesium nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

10. The solid rocket fuel mixture of any of claims 1 - 9, wherein the second mixture is configured to react with the propellant oxidizer byproduct to form hydrogen gas.

11. A method of propelling a rocket, the method comprising:providing a first mixture comprising propellant oxidizer nanoparticles having an average particle size of about 1 nm to about 900 nm and magnesium nanoparticles having an average particle size of about 1 nm to about 900 nm;combusting the first mixture to produce a first stage propulsion and to produce a propellant oxidizer byproduct;providing a second mixture comprising water; andcombusting the second mixture and the propellant oxidizer byproduct to produce a second stage propulsion.

12. The method of claim 11, wherein the propellant oxidizer comprises one of silica, ammonium nitrate, N2O, KNO3, NH4CIO4, H2O2, MgO, lithium perchlorate (l_iCI04) , sodium perchlorate (NaCI04) , potassium perchlorate (KCI04), barium peroxide (BaO2), ozone (O3), liquid oxygen (LOX), tetranitromethane (C(NO2)4), nitronium perchlorate (NO2CIO4), Fe2O3, CuO, Co3O4, MnO2and CaO.

13. The method of claim 11 or 12, wherein the propellant oxidizer byproduct comprises silicon.

14. The method of any of claims 11 - 13, wherein the second mixture further comprises magnesium hydroxide configured to react with the propellant oxidizer and the water.

15. The method of claim 14, wherein the magnesium hydroxide comprises magnesium hydroxide nanoparticles with an average particle size of about 1 nm to about 900 nm.

16. The method of claim 13, wherein the water and the magnesium nanoparticles are provided in a molar ratio of about 1:1.

17. The method of any of claims 11 - 16, wherein the propellant oxidizer nanoparticles and the magnesium nanoparticles are provided in a molar ratio of about 1:3 to about 3:1.

18. The method of any of claims 11 - 17, wherein the propellant oxidizer nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

19. The method of any of claims 11 - 18, wherein the magnesium nanoparticles have a surface area of about 1 m2 / g to about 500 m2 / g.

20. The method of any of claims 11 - 19, wherein combusting the second mixture and the propellant oxidizer byproduct forms hydrogen gas.

21. The method of claim 20, further comprising burning the hydrogen gas to produce additional propulsion.

22. The method of any of claims 11 - 21 , wherein combusting the first mixture produces a temperature of at least about 2,000 °C.

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