Functionally graded, in situ manufacturable energetics and heating methods

Abstract

A functionally graded regolith-based energy fuel and method are disclosed for driving heating systems to help personnel and equipment operate in harsh lunar conditions. The functionally graded material includes energetic particles with a metallic fuel and a regolith-based oxidizer. The energetic particles and regolith-based oxidizer are mixed to form an energetic material. A regolith-based combustible material is disclosed that includes 20%, 30%, or 40% w / w micromagnesium and a regolith-based oxidizer, and the material is ball milled for up to 5 hours.

Description

[Technical Field]

[0001] FIELD OF THE INVENTION The embodiments disclosed herein relate to energetic particles, and in particular to functionally graded, in-situ produced energetic particles. [Background technology]

[0002] Energetic particles, such as microthermites, nanothermites, and / or metallic materials and / or fuels, can be used for heating and combustion as energy sources to meet energy demands on Earth and in space. Thermite materials have been used in railroad construction, mining, and defense industry applications. Generally, thermite materials consist of two distinct components: a fuel and an oxidizer. Reduced metals, metal alloys, and metalloids, such as aluminum (Al), magnesium (Mg), and silicon (Si), are usually chosen as fuel components due to their large enthalpy changes during combustion. Meanwhile, the oxidizers are typically metal oxides, halides, and oxyanion salts. Al-based thermites are the most commonly used thermites in this family due to their abundance, ease of preparation, and very high theoretical enthalpy of combustion of 31 kJ / g. Large heat is released during the exothermic oxidation-reduction reaction between the fuel, oxidizer, and intermediates, leading to numerous applications in propulsion, pyrotechnics, welding, and more.

[0003] Microthermites exhibit significantly improved reactivity compared to conventional thermite mixtures with larger particle sizes. This increased reactivity is due to the increased surface area available for reactants to contact each other. Small particle sizes allow for tight control of the thermite reaction, which is important in applications where the timing and intensity of the reaction must be finely tuned. Microthermites often have lower ignition temperatures than their larger counterparts. This means they can be ignited more easily, even with lower heat or energy inputs. Microthermites can be engineered to have enhanced energetic properties, such as higher energy density and faster reaction rates, making them attractive for a wide range of applications, including propulsion and heating devices. Their small size and high energy output make microthermites particularly well-suited for miniaturized devices and applications. They are used in microthrusters for small satellites and micropropulsion systems for point-to-point transport of materials, spacecraft, and / or manned and unmanned cargo, including station-keeping and orbit raising for space systems. Microthermites can be used for precise heating of small areas or objects. This is useful for applications such as joining, welding, and / or sintering use cases where traditional heating methods are impractical. In some applications, microthermites can be designed and operated to be functionally graded, with safety standards to prevent unintentional ignition. This is particularly important when using them in small-scale devices to fit user-defined combustion and / or heating applications or use cases.

[0004] Furthermore, nanothermites, also known as metastable intermolecular composites, consist of nanoscale fuel or oxidizer components, or both. Nanothermites exhibit significantly enhanced combustion rates due to greater surface contact between the fuel and oxidizer particles, and an amplified role for reactive interfaces with large surface-to-volume ratios. Intensive research efforts have focused on developing nanothermites with various compositions and micro / mesostructures to further improve their reaction rates and result in optimized combustion and combustion chamber pressurization rates. However, real-world applications of nanothermit materials remain very limited due to several significant drawbacks, including difficulties in reactivity control, the degree of combustion completion, and issues related to safe fabrication, handling, and transportation. The lack of controllability of reactivity is caused by multiscale physics related to phase separation during synthesis, reactive sintering of energetic nanoparticles during combustion and / or heating, and the formation and fabrication of functionally graded energetic particles. Furthermore, nanothermites exhibit extreme sensitivity to external stimuli such as electrostatic discharge, friction, and mechanical shock.

[0005] In a typical solid rocket motor (SRM) design, fuel is cast into the rocket case using a mandrel. This is a highly dangerous process and prone to explosions. The conceptual design phase of a rocket propulsion system aims to match mission requirements with pre-design considerations of the engine within the specific constraints of the problem. Parametric design space is particularly important for solid or hybrid rocket engines because propellant properties, particle structure, and combustion chamber geometry can be independently varied for a specific mission. Furthermore, as combustion progresses and the fuel retracts, the thermodynamic properties of the engine change simultaneously as the total combustion chamber volume increases and the normalized surface area changes. Fuels typically have a hollow geometry, with combustion occurring radially until all fuel is consumed. As the fuel burns, the combustion chamber area increases (reducing pressure), the combustion area increases with increasing radius, and increasing the combustion area increases the amount of fuel burned. The highly multidimensional nature of this problem has motivated the development of a framework for matching mission profiles to propulsion system design.

[0006] Depending on the focus of the conceptual design, some frameworks propose comprehensive considerations for the optimization of propulsion systems for ascending trajectories, including the consideration of nonlinear aerodynamic forces (Federici, L., et al., "Integrated optimization of first stage SRM and ascent trajectories of multistage launch vehicles", Journal of Spacecraft and Rockets, vol. 58, no. 3, pp. 786-797, 2021). Other studies aim to optimize the geometry of fuel grains to meet mission-specific desired thrust profiles (Oh, S.H., et al., "Study of hybrid optimization technique for grain optimum design," International Journal of Aeronautical and Space Sciences, vol. 18, no. 4, pp. 780-787, 2017), provide system and structural modeling (Adami, A., et al., "A new approach in multidiscriminal design optimization of upper-stages using a combined framework," Acta Astronautica, vol. 114, pp. 174-183, 2015), or perform performance matching optimization (Zeping, W., et al., "Solid rocket motor design having an efficient performance matching approach," Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, vol. 233, no. 11, pp. 4052-4065, 2019). In many of these studies, propellant properties such as burn rate are not considered independent variables given the strong interdependence of propellant regression rate, heat release, and corrosion characteristics.

[0007] Existing methodologies often rely on variational optimization approaches to determine the optimal geometric parameters of fuel particles. These optimizations seek to minimize overall mass and / or total cost, constrained by the physics of the problem. However, for complex fuel particles, manufacturability is often the biggest constraint on the optimization problem.

[0008] Existing research has explored the application of optimization techniques to more complex geometric cases with additional considerations (Johannsson, M., "Optimization of Solid Rocket Grain Geometry," 2012). Methodologies such as burnback analysis based on level set methods use a level set approach to track changes in the topology of fuel-driven SRMs (Wang, D.H., et al., "An integrated framework for solid rocket motor grain design optimization," Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, vol. 228, no. 7, pp. 1156-1170, 2014). Most of these optimization techniques rely on continuous bounded ranges for each parameter being optimized. These techniques also require designers to appropriately assign bounds to the problem so that the program can reach the optimal solution in a timely manner. Some recent studies have proposed machine learning approaches (Oh, SH, et al., "New design method of solid propellant grain using machine learning," Processes, vol. 9, no. 6, 2021) or two-component propellant grain optimization (Alazeezi, M., et al., "Two-component propellant grain for rocket motor: Combustion analysis and geometric optimization," Thermal Science, vol. 26, no. 2 Part B, pp. 1567-1578, 2022), which may be more beneficial for optimizing fuel grains in solid rocket motors. Ultimately, these constraints limit the ability to optimize more complex thrust profiles. Summary of the Invention [Problem to be solved by the invention]

[0009] Therefore, there is a need for new and improved functionally graded regolith-based energy fuel particles for fuels and materials and methods to drive heating systems to help people and equipment operate in harsh conditions and enable long-term sustainable operation. [Means for solving the problem]

[0010] overview Described herein is a multi-functionally graded material that includes energetic particles that include a metal fuel and a regolith-based oxidizer. The energetic particles and the regolith-based oxidizer are mixed to form the energetic material.

[0011] According to one embodiment, the regolith-based combustible material is formed into pellets.

[0012] According to one embodiment, the regolith-based oxidizer is JSC-1A.

[0013] Described herein is a method of heating using a regolith-based combustible material, the method including adding micro-magnesium and a regolith-based oxidizer to a ball mill at a micro-magnesium concentration of 20%, 30% or 40% w / w (mass % concentration), forming a combustible material by ball milling the micro-magnesium and regolith-based oxidizer for up to 5 hours, and heating the combustible material until it ignites.

[0014] According to one embodiment, igniting the flammable material comprises igniting the flammable material in a vacuum.

[0015] According to one embodiment, the method further comprises forming the combustible material into pellets.

[0016] According to one embodiment, heating the combustible material comprises exposing the combustible material to a laser beam.

[0017] According to one embodiment, the heating of the combustible material comprises electromagnetic heating.

[0018] According to one embodiment, the heating of the combustible material comprises non-radiative heating.

[0019] According to one embodiment, the regolith-based oxidizer is JSC-1A.

[0020] According to one embodiment, the method further comprises heating, including collecting by-products after combustion.

[0021] According to one embodiment, the method further comprises harvesting in situ a regolith-based oxidizer of lunar origin.

[0022] Other aspects and features will become apparent to those of ordinary skill in the art upon review of the following description of several exemplary embodiments.

[0023] The drawings included herein are intended to illustrate various examples of the articles, methods, and apparatus herein. [Brief explanation of the drawings]

[0024] [Figure 1A] , [Figure 1B] , [Figure 1C] , [Figure 1D] , [Figure 1E] , [Figure 1F] 1A-1C are diagrams of various fuel particle structures and corresponding thrust profiles.

[0025] [Figure 2] FIG. 1 is a diagram of functionally layered fuels.

[0026] [Figure 3A] An exemplary receding (simple) thrust profile.

[0027] [Figure 3B] An exemplary two-peak (complex) thrust profile.

[0028] [Figure 3C] 1 is an exemplary pressure versus altitude curve for optimal thrust during climb.

[0029] [Figure 4] 1 is a flowchart of a method for optimizing a conceptual design of a graded function solid rocket motor.

[0030] [Figure 5A] This is an ideal burn rate profile for a simple thrust.

[0031] [Figure 5B] This is an ideal burn rate profile for complex thrust.

[0032] [Figure 5C] This is the ideal burning rate profile for pressure matching.

[0033] [Figure 6A] , [Figure 6B] , [Figure 6C] These are thrust profiles optimized based on the ideal combustion rates shown in FIGS. 5A to 5C.

[0034] [Figure 7A] , [Figure 7B] 1 is a scanning electron microscope (SEM) image of the sample structure before combustion.

[0035] [Figure 8A] , [Figure 8B] , [Figure 8C]1 is a scanning electron microscope (SEM) image of the sample structure before combustion.

[0036] [Figure 9] 1 is a differential scanning calorimetry (DSC) and thermogravimetric analysis (TGA) showing the energy release of regolith-based combustion materials with different ball milling times, according to some embodiments.

[0037] [Figure 10] 1 is a high-speed video frame capturing the burning of a sample structure during an open-air ignition, according to one embodiment.

[0038] [Figure 11] 1 is a high-speed video frame capturing the burning of a sample structure during an open-air ignition, according to one embodiment.

[0039] [Figure 12] 1 is a high-speed video frame capturing the burning of a sample structure during an open-air ignition, according to one embodiment.

[0040] [Figure 13] 1 is a high-speed video frame capturing the combustion of a sample structure upon vacuum ignition, according to one embodiment.

[0041] [Figure 14] 1 is a pyrometric thermal video frame showing combustion temperature and emissivity of a pellet, according to one embodiment.

[0042] [Figure 15] 1 is a pyrometric thermal video frame showing the burning temperature and emissivity of pellets in an open combustion according to one embodiment.

[0043] [Figure 16]1 is a pyrometric thermal video frame showing the burning temperature and emissivity of pellets in an open combustion according to one embodiment.

[0044] [Figure 17] 1 is a pyrometric thermal video frame showing the burning temperature and emissivity of pellets in an open combustion according to one embodiment.

[0045] [Figure 18] 1 is a pyrometric thermal video frame showing the burning temperature and emissivity of pellets in an open combustion according to one embodiment.

[0046] [Figure 19A] and [Figure 19B] 1A-1C are scanning electron microscope images of a sample structure before and after combustion, respectively, according to one embodiment.

[0047] [Figure 20A] and [Figure 20B] 1A-1C are scanning electron microscope images of a sample structure before and after combustion, respectively, according to one embodiment.

[0048] [Figure 21] FIG. 1 is a block diagram of multi-functionally graded structures, according to one embodiment.

[0049] [Figure 22] FIG. 1 is a block diagram of a multifunctional gradient material combined with a detachable electromagnetic system, according to one embodiment.

[0050] [Figure 23] FIG. 1 is a block diagram of a multi-source heating assembly, according to one embodiment.

[0051] [Figure 24]FIG. 1 is a block diagram of a multi-source heating assembly, according to one embodiment. DETAILED DESCRIPTION OF THE INVENTION

[0052] Various devices or processes are described below to provide examples of each claimed embodiment. None of the embodiments described below limit the claimed embodiment, and any claimed embodiment may cover a process or device different from the one described below. A claimed embodiment is not limited to a device or process having all the features of any one device or process described below, or to features common to some or all of the devices described below.

[0053] A description of an embodiment having several components in communication with each other does not imply that all such components are required, to the contrary, various optional components are described to illustrate the wide variety of possible embodiments of the present invention.

[0054] Furthermore, although process steps, method steps, algorithms, etc. may be described (in this disclosure and / or claims) in a sequential order, such processes, methods, and algorithms may be configured to function in alternative orders. In other words, any sequence or order of steps that may be described does not necessarily indicate a requirement that the steps be performed in that order. Steps of processes described herein may be performed in any order practical. Moreover, some steps may be performed simultaneously.

[0055] Where a single device or article is described herein, it will be readily apparent that two or more devices / articles (whether or not they cooperate) may be used in place of the single device / article. Similarly, where two or more devices or articles are described herein (whether or not they cooperate), it will be readily apparent that a single device / article may be used in place of the two or more devices or articles.

[0056] In situ resource utilization to prepare combustion materials offers important advantages for sustained human presence in space and potential industrial applications. In situ resource utilization refers to the preparation of desired items using resources found in situ. For example, lunar regolith contains fragmented materials composed of minerals such as anorthite and ilmenite, as well as various metal oxides. These minerals include elements such as aluminum and iron, which are the main components of thermite, an effective combustion material with important applications. For long-duration space missions, combustion materials prepared using in situ resources can be used as fuel for propulsion or infrastructure activities such as construction and metal extraction. This can result in a reduction in the overall payload for space missions, leading to cost savings, reduced size and cargo weight, and reduced risk of supply shortages. Regolith-based thermite can be used to provide heating to improve the survivability of personnel and equipment during lunar nights or for ambitious lunar polar missions.

[0057] In another example, Martian regolith contains fragmented materials composed of minerals, rocks, dust, sand, soil, carbonates, silicates, sulfates, phyllosilicates, perchlorates, metals, and various metal oxides. These metals include elements such as iron, aluminum, magnesium, sulfur, silicon, sodium, potassium, chromium, nickel, cobalt, copper, titanium, gold, and platinum, which are the main components of thermite, an effective combustion material with important applications. For long-duration space missions, heating and / or combustion materials prepared using in situ resources can be used for propulsion or power generation, or as fuel for infrastructure activities such as construction, welding, and metal extraction. In other embodiments, metals, metal alloys, rocks, and other minerals can be sourced from asteroids, other moons, planets, asteroids, and other celestial bodies. As a result, the overall payload for space missions can be reduced, leading to cost savings, reduced size and cargo weight, and reduced risk of supply shortages. In situ regolith-based thermite could be used to provide heating and combustion to improve survivability of people and equipment on the lunar night, or to support lunar polar missions, Martian nights, and / or other ambitious space missions.

[0058] Recent years have seen dramatic changes in solid and hybrid propulsion technologies as new propulsion paradigms take hold. Recent studies have proposed the integration of pyrophoric additives based on metal-organic frameworks (MOFs) (Jobin, O., et al., "Metal-organic frameworks as hypergolic additives for hybrid rockets," Chemical Science, vol. 13, no. 12, pp. 3424-3436, 2022). At the same time, new manufacturing processes for solid-state fuels through additive manufacturing (AM) techniques are opening up new design opportunities for novel classes of solid and hybrid propulsion systems.

[0059] Historically, the selection of fuel grain structure represented an optimum approach to achieve a desired thrust-duration curve.

[0060] Various fuel grain structures and corresponding thrust profiles (100) are shown in Figures 1A-1F. For known propellant burn and regression rates, the imposed grain structure results in a change in area, and therefore combustion heat release, over time. For example, a relatively simple tubular grain design (Figure 1A) results in a progressive (increasing) thrust curve because the burn area increases as the grain retreats toward the outer wall. A more complex star-shaped fuel grain (see Figure 6A, right) is required for a neutral (constant) thrust curve. The design of these grain structures utilizes burnback models. Furthermore, these complex structures are prone to significant erosive burn, structural integrity issues, and manufacturability constraints.

[0061] As discussed above, the new opportunities offered by AM of energetic fuels mean that solid and hybrid engine design considerations can be shifted away from complex geometric fuel particles to new design considerations by modifying thrust-duration curves and functionally grading single or even multi-propellant engines. By layering various fuels with spatially varying binder and propellant compositions and / or densities, it is possible to effectively construct matching thrust-duration profiles without the need for complex fuel particles, thus opening up new opportunities for novel engine design and optimization considerations.

[0062] Various energetic materials and nanothermit aerogels are summarized in Table 1, which includes Al as the fuel, Bi2O3 as the oxidizer, and UV or other polymer binders. Other metal fuel and oxidizer combinations may also be possible. For example, the metal fuel may be Mg, Si, Fe, etc., and the oxidizer may be a fluoropolymer, iodine oxide, or metal oxide (e.g., Fe2O3, SiO2, MgO, etc.). By varying the fuel composition, the burn rate can be controlled as shown in Table 1. The nanothermit aerogels summarized in Table 1 exhibit extremely fast burn rates, demonstrating the potential to adapt this technology to fuels with lower burn rates. The aerogels described herein can be formed into fuel particles for a variety of Earth, lunar, and space applications, including tunable heating and combustion implementations in solid rocket motors (SRMs) and / or hybrid rocket motors.

[0063] [Table 1]

[0064] A computational framework for optimizing fuel grain structures to match a desired thrust curve profile is provided. The framework includes two solvers with varying levels of fidelity to efficiently optimize over a large parameter space. A system-level code (zero-dimensional) coupling the combustor and nozzle system was first developed to evaluate the behavior of the entire system. A quasi-one-dimensional code was then developed to incorporate spatial variations and acoustic modes in the combustion chamber and nozzle of a given functionally graded engine. Given the complex combustion kinetics of solid fuels, which remain largely unknown, simplified combustion and regression models were used, as summarized below.

[0065] A system-level solver was first developed using the entropy nozzle relationship to estimate the engine's vacuum thrust characteristics. The total thrust of the engine can be estimated based on the total pressure and temperature generated in the combustor for a given nozzle geometry. The vacuum thrust equation can be reformulated as follows:

number

[0066] The 0D model uses isotropic flow equations to relate combustion chamber conditions (total pressure and temperature) to nozzle exit conditions. In a supersonic nozzle, the mass flow rate is fixed for a given geometry and thermodynamic state of the engine. For a known nozzle geometry (Ae / At), the exit temperature and pressure can be calculated knowing the thermodynamic conditions in the engine.

number

[0067] The 0D model also utilizes the ideal rocket equation to find the relationship between altitude and time, which becomes important for one of the test cases: where v is the rocket speed, u is the exit velocity of the combustion gases, and m is the total mass leaving the rocket.

number

[0068] A quasi-one-dimensional solver was simultaneously developed to account for spatial variations in the combustion chamber and nozzle, as well as to investigate acoustic coupling in the functionally graded engine. The quasi-one-dimensional code solves the one-dimensional Navier-Stokes equations (conservation of mass, momentum, and energy). Spatial fluxes were calculated by a fifth-order WENO scheme, and the equations were integrated in time by Strang splitting for robustness. Similar to the system-level framework, a constant linear burn rate was used to characterize each propellant.

[0069] From the above framework, we can proceed with equations for a given combustion chamber and nozzle geometry and calculate the thrust generated. The thrust of a rocket is given by F th =mu e +(P e -P atm )A e When the equation is integrated over time and the fuel is retracted, the thrust profile curve can be estimated.

[0070] Functionally graded (200) is achieved by stacking different fuels (202, 204, and 206), as shown in Figure 2. Several parameters are known to affect the performance of an SRM. The primary parameters are heat release, gas release, and fuel burn rate. Burn rate can be controlled by manipulating several fuel properties, such as chamber pressure, fuel density, fuel porosity, chemical composition, and physical composition. In the case of additively manufactured energetic materials, several aspects of the physical composition, such as propellant loading, additional polymer additives, and binder materials, can affect the burn rate. Varying these parameters can vary the burn rate of solid fuels from a few millimeters per second to hundreds of meters per second. Therefore, functionally graded engines can be designed by stacking different fuels with diverse burn characteristics to alter the burn characteristics during combustion. Therefore, a mission-specific thrust profile can be achieved with optimal fuel stacking.

[0071] To illustrate how this framework can be used, three well-defined test cases were considered: a simple, receding thrust profile, a complex thrust profile with multiple peaks, as proposed by Federici et al., and a conceptual case in which the total pressure conditions in the engine are adjusted for full expansion in the nozzle during ascent. Test case 300 is shown in Figures 3A-3C.

[0072] Test Case 1 (Figure 3A) corresponds to a receding profile typically achieved by using complex particle geometries, such as a double-anchor shape. Test Case 2 (Figure 3B) is a more complex, two-peak thrust profile presented by Federici et al. Finally, the third test case (Figure 3C) is a conceptual case in which the engine's combustion characteristics are modified to match a fully expanded nozzle flow during ascent. The ideal pressure versus altitude curve (302) shown in Figure 3C is converted to a pressure versus time curve in the fuel particle optimization. As previously mentioned, the ideal rocket equation relates rocket speed to rocket mass and the exit velocity of the burning gases. Using this information, rocket mass, rocket speed, and height are calculated as combustion occurs, and therefore atmospheric pressure changes with time. From this, an ideal exit pressure (perfectly matching atmospheric pressure) as a function of time curve is created and optimized.

[0073] Figure 4 shows the optimization framework for the conceptual design of a functionally graded solid rocket engine to match thrust-time and pressure-time profiles for a given mission. Given the ability to create a functionally graded engine, fuel burn rate and layer thickness were selected as independent parameters of the optimization problem, and tubular grains (Figure 1A) were assumed for simplicity. Given a moderate parametric space, a brute-force profile matching optimization approach was implemented, but the modularity of the code allows for constrained optimization algorithms to be implemented. In such cases, constraints are applied to the burn characteristics.

[0074] At 12, a thrust-pressure curve for a given mission is selected. At 14, the optimization process begins in the 0D code to time-advance a set of equations to determine the burn rate that will most closely match the thrust or pressure profile to the sample profile at each time step. At each time step, the ideal burn rate and the radius at which this occurs are tracked, allowing the ideal radial distance vs. burn rate curve to be plotted.

[0075] Once the optimal burn rate at each radial value is known, the fuel field is divided into four approximate sections that act as layers within the fuel at 16. The burn rate and thickness of each section are bounded after analyzing the ideal burn rate curve. To aid in bounding the burn rate and layer thickness, coarser velocity samples are used to create a new ideal velocity vs. burn rate curve. Four to five burn rate samples are used within the minimum and maximum values ​​of the ideal curve. Although less accurate, this can be used to determine the fuel layer characteristics. As the velocity varies between the coarser values, bins are created within which the ideal velocity resides. At 18, the range of velocity inputs for optimization is bounded by the minimum and maximum bin values. The radius input for each layer is bounded by the radial location where the bin value varies by ±10% without overlapping with the boundaries of adjacent layers. The layer characteristics are optimized within these defined boundaries. The size of these layers and their burn rates are then optimized to best fit the desired thrust or pressure profile.

[0076] In 20, optimization of each layer thickness and burn rate is completed using a heuristic approach, using a range of random inputs of layer thickness and burn rate to determine the most accurate solution. The model preserves the layer thickness and burn rate of the model with the lowest error. The error between the desired profile and the profile generated by the model is the sum of the absolute differences between the two models over all time steps. Finally, the optimized layered solution is passed to a 1D code to evaluate spatial variations and acoustics within the engine.

[0077] Referring to Figure 5A, an exemplary ideal burn rate profile (501) for a simple thrust is shown, plotting the ideal burn rate at each radial distance based on time step. Figure 6A shows an optimized solution (601) based on a recursive method, selecting an approximated random burn rate and layer thickness based on the ideal rate shown in Figure 5A. Curve 602 represents the ideal profile. Curve 603 represents the optimization. It can be seen that there are four distinct layers of fuel before the combustion chamber depressurizes at the end of the launch. The optimal layer thicknesses, outer radii, and burn rates for the simple case are shown in Table 2 below.

[0078] [Table 2]

[0079] In the simple case, we can see that the fuel layer thicknesses vary as the fuel burn progresses: during the initial thrust peak at the start of the orbit, the thickness of the first layer accounts for 43% of the total fuel radius, while the second, third, and fourth fuel layers account for 19%, 16%, and 22% of the thickness, respectively.

[0080] Referring to Figure 5B, an ideal burn rate profile (511) for the complex thrust is shown, plotting the ideal burn rate at each radial distance based on the time step. Figure 6B shows the optimized solution (611) based on a recursive method, selecting estimated random burn rates and layer thicknesses based on the ideal rates shown in Figure 5A. Curve 612 represents the ideal profile. Curve 613 represents the optimization. Note that the range of inputs is different from the simple case (Figures 5A and 6A) because the range of values ​​is based on the optimal burn rate at each radius for both models. For this reason, a wider range of inputs was used for the second test case when setting up the optimization test. It can be seen that with four simple fuel layers with different burn rates, the modeled SRM thrust profile can closely match the desired thrust profile of the rocket. The optimal radii and burn rates for the complex case are shown in Table 3 below.

[0081] [Table 3]

[0082] In the complex case, the fuel layer thickness can be seen to vary as the fuel burn progresses. During the first peak of thrust at the start of the trajectory, the thickness accounts for 46% of the fuel thickness, while the second, third, and fourth fuel layers account for 14%, 31%, and 9% of the thickness, respectively. This is a larger variation in layer thickness than in the simple case. We can also see that there is a larger variation in the burn rate in the complex case.

[0083] In the third test case, the goal is to match the exhaust pressure of the combustion gases to atmospheric pressure. Referring to Figure 5C, an ideal burn rate profile (521) for pressure matching is shown, plotting the ideal burn rate at each radial position. Figure 6C shows the optimal pressure profile (621) over time using four separate layers. Curve 622 represents the ideal profile. Curve 623 represents the optimization. As mentioned in the test case subsection above, the ideal pressure profile over time is found by using the ideal rocket equation and determining the rocket height over time based on mass and gas exit velocity. In this case, the rocket exit pressure is seen to follow the ideal case test quite well. The optimal burn rates and layer thicknesses are shown in Table 4 below.

[0084] [Table 4]

[0085] For all three test cases, using more layers could improve the agreement produced by the model. In the future, when fuels can be functionally graded, the burn rate could be controlled throughout the fuel to produce a more accurate SRM. This would result in an optimized profile similar to the ideal profile modeled here. Additive manufacturing processes (e.g., 3D printing) could be used to use the extrusion width of the material as the minimum layer thickness, resulting in a tunable profile.

[0086] The above framework and methods for functionally graded rocket engines can be extended to the development of functionally graded energetics for various Earth and space applications, particularly for propulsion and power generation, construction, storage systems, and controlled heating of equipment. For example, the lunar environment is one of the most hostile environments in which humans have operated. Temperatures near the equator can drop to 140 K, and temperatures near the poles can drop to as low as 25 K in areas of permanent shadow. Such low temperatures can damage equipment and personnel.

[0087] Existing methods for heating in the lunar or space environment have limitations. Radioactive heating elements are typically controlled commodities inaccessible to civilian workers. Batteries are often large and heavy and may need to be custom-built for the lunar climate. Fuels such as liquid hydrogen / liquid oxygen (LH / LOX) have not yet been developed to the point where water can be extracted from the lunar surface, requiring the transport of other fuels as cargo. Therefore, in-situ fabrication of multifunctional graded fuels and materials is desirable.

[0088] In situ resource utilization (ISRU) is the process of using locally sourced resources to achieve a desired goal. In space applications, ISRU reduces mission costs by reducing cargo size and weight. Materials in the lunar / Martian regolith and / or asteroids consist of a series of metal oxides that, when fuel is added to generate heat, can be used as thermite. ISRU operations are typically limited to the production of construction materials, ignoring heat release.

[0089] In one embodiment, the additive manufacturing process described above involves the application of polymer-free nanothermit aerogels to 3D printing, which is used to prepare regolith-based nanothermites.

[0090] The regolith-based combustion material uses micromagnesium as a fuel and JSC-1A as an oxidizer. Micromagnesium, when milled to the micro- or nanoscale, increases its surface area, providing higher reactivity and a faster exothermic response than conventional thermite mixtures. Furthermore, combustion materials using micromagnesium as a fuel ignite at lower temperatures, making them effective on the lunar surface. JSC-1A is a lunar regolith stimulant with mineralogical and chemical properties similar to those of lunar regolith. In another embodiment, regolith from an in situ environment is used as a source of fuel and / or oxidizer.

[0091] Test Cases

[0092] A framework for sampling regolith-based combustion materials is provided using lunar regolith stimulators as oxidizer sources. Test cases were selected with various magnesium compositions in the combustion material, namely, 2-5%, 30%, and 40% magnesium. The combustion materials were then subjected to a ball milling process for 0, 2, 5, and 10 hours, respectively, to increase their surface area.

[0093] Overall, 12 samples were prepared for combustion testing with different magnesium composition percentages and ball milling times. For combustion testing, 30 mg pellets of the samples were prepared for combustion testing under 250 psi pressure for 1 minute. The pressure can be calibrated or modified to suit the needs of the experiment.

[0094] Referring to Figures 7A-7B, scanning electron microscope (SEM) images of the pre-combustion sample structures are shown. Figure 7A provides a general view of the sample structure (701) without magnesium and processed in a ball mill for 2 hours. Figure 7B provides a general view of the sample structure (711) with magnesium and processed in a ball mill for 2 hours. Ball milling can produce smaller fragments of the sample structure (712). As can be seen from the images, no significant differences were observed between the samples. Many large particles are visible in both samples. Furthermore, as can be seen in Figure 7C, the magnesium (722) is not uniformly distributed.

[0095] Referring to Figures 8A-8C, scanning electron microscope (SEM) images of the pre-combustion sample structure are shown. Figure 8A provides a general view of the sample structure (801) without magnesium and processed in a ball mill for 10 hours. Figure 8B provides a general view of the sample structure (811) with magnesium and processed in a ball mill for 10 hours. Ball milling can produce smaller fragments of the sample structure (812). As can be seen from the images, smaller particles are observed with increased ball milling time. However, larger magnesium particles are observed in Figure 8B. Furthermore, as can be seen in Figure 8C, the magnesium (822) is not uniformly distributed.

[0096] Referring to FIG. 9, there is shown differential scanning calorimetry (DSC) and thermogravimetric analysis (TGA) (900) illustrating the energy release of regolith-based combustion materials with different ball milling times, according to some embodiments.

[0097] Differential scanning calorimetry (DSC) and thermogravimetric analysis (TGA) were used to calculate the energy release of sample regolith-based combustion materials, focusing on the thermal properties of the sample structure. Test cases confirmed that both the composition percentage of magnesium in the sample structure and the milling time affect the energy release.

[0098] Based on the DSC-TGA energy release values, it was observed that the sample structures with mill times over 5 hours (40% for 902, 30% for 904, and 20% for 906) showed little difference in energy release. [Table 5]

[0099] The energy release can provide a heat source for a variety of applications. The composition and ball milling time can be optimized and adjusted to meet specific objectives.

[0100] Referring to FIG. 10, there is shown high speed video frames (1000) capturing the burning of a sample structure in an open fire, according to one embodiment.

[0101] The sample structure used for combustion consisted of 20% magnesium, and the ball milling time was 5 hours. The video frames show the combustion in various stages, including flame initiation (1002), ignition (1004), combustion propagation (1006), and termination / exhaustion (1008). The combustion propagation was not violent, and no explosions or eruptions were observed. The sample structure initiated combustion upon exposure to the laser beam and continued to burn after the laser was turned off. Small flames were observed as the reaction propagated. Because the combustion was conducted in the open, oxygen was part of the reaction.

[0102] Referring to FIG. 11, there is shown high speed video frames (1100) capturing the burning of a sample structure in an open fire, according to one embodiment.

[0103] The sample structure used for combustion consisted of 30% magnesium, and the ball milling time was 5 hours. The video frames show the combustion in various stages, including flame initiation (1102), ignition (1104), combustion propagation (1106), and termination / exhaustion (1108). The combustion propagation was violent, with an explosion and several combustion eruptions observed. The explosive nature was likely due to the gasification of the magnesium. The sample structure initiated combustion upon exposure to the laser beam and continued to burn even after the laser was turned off.

[0104] Referring to FIG. 12, there is shown high speed video frames (1200) capturing the burning of a sample structure in an open fire, according to one embodiment.

[0105] The sample structure used for combustion consisted of 40% magnesium, and the ball milling time was 5 hours. The video frames show combustion in various stages, including flame initiation (1202), ignition (1204), combustion propagation (1206), and termination / exhaustion (1208). The combustion propagation was most intense in the sample structure, causing the sample to explode into multiple smaller fragments. The explosive nature is likely due to the more rapid gasification of magnesium. Furthermore, the addition of magnesium results in the formation of larger amounts of gas.

[0106] Referring to FIG. 13, there is shown high speed video frames (1300) capturing the combustion of a sample structure under vacuum ignition, according to one embodiment.

[0107] The sample structure used for combustion consisted of 20% magnesium, and the ball milling time was 5 hours. Video frames show combustion in various stages, including flame initiation (1302), ignition (1304), combustion propagation (1306), and termination / exhaustion (1308). The combustion initiation demonstrated that atmospheric oxygen was not required for the reaction to occur. The combustion propagation involved one small initial explosion followed by controlled combustion. Sample structures with higher magnesium content demonstrated an explosion, which only extinguished if the explosion occurred and the laser was not activated.

[0108] Referring to FIG. 14, a pyrometric thermal video frame (1400) showing combustion temperature and emissivity of a pellet is shown according to one embodiment.

[0109] Pyrometry was used to estimate the combustion temperature and emissivity of the sample structure pellet. The emissivity of the sample structure has a significant effect on the temperature. The temperature of the sample structure pellet during combustion, near the flame, and after the flame was found to be approximately 1700 K. This temperature significantly exceeds the boiling point of magnesium (1365 K). This indicated an emissivity value of approximately 35%.

[0110] Referring to FIG. 15, a pyrometric thermal video frame (1500) showing the burning temperature and emissivity of pellets in an open fire is shown, according to one embodiment.

[0111] In Figure 15, the sample structure used for combustion consisted of 20% magnesium and the ball milling time was 2 hours. The video frames show combustion at various stages, including flame initiation (1502), ignition (1504), combustion propagation (1506), and termination / exhaustion (1508).

[0112] Referring to FIG. 16, a pyrometric thermal video frame (1600) showing the burning temperature and emissivity of pellets in an open combustion is shown, according to one embodiment.

[0113] In Figure 16, the sample structure used for combustion consisted of 20% magnesium and had a ball milling time of 5 hours. The video frames show combustion at various stages, including flame initiation (1602), ignition (1604), combustion propagation (1606), and termination / exhaustion (1608). Significant deformation of the sample after combustion was observed.

[0114] Referring to FIG. 17, a pyrometric thermal video frame (1700) showing the burning temperature and emissivity of pellets in an open combustion is shown, according to one embodiment.

[0115] The sample structure used for combustion consisted of 20% magnesium, and the ball milling time was 10 hours. The reaction involved combustion, and after the reaction propagated, the sample remained at elevated temperature for approximately 10 seconds. Video frames show combustion at various stages, including flame initiation (1702), ignition (1704), combustion propagation (1706), and termination / exhaustion (1708). Less post-combustion deformation was observed compared to the sample structures in Figures 9 and 10.

[0116] Referring to FIG. 18, a pyrometric thermal video frame (1800) showing the burning temperature and emissivity of pellets in an open-air combustion is shown, according to one embodiment.

[0117] The sample structure used for combustion consisted of 30% magnesium, and the ball milling time was 5 hours. Video frames show combustion at various stages, including flame initiation (1802), ignition (1804), combustion propagation (1806), and termination / exhaustion (1808). The reaction involved more rapid combustion propagation and violent reactions compared to the other sample structures. Solid ejections were observed as they left the sample. The burning sample structure reached and maintained high temperatures for longer periods.

[0118] Referring to Figures 19A and 19B, scanning electron microscope images of the sample structure before (1900) and after (1950) combustion, respectively, are shown according to one embodiment.

[0119] The sample structure used for combustion consisted of 20% magnesium, and the ball milling time was 2 hours. The post-combustion material exhibited a homogeneous final structure (1952). Once combustion was complete, the magnesium was observed to be clearly distributed (1954).

[0120] Referring to Figures 20A and 20B, scanning electron microscope images of the sample structure before (2000) and after (2050) combustion, respectively, are shown according to one embodiment.

[0121] The sample structure used for combustion consisted of 20% magnesium, and the ball milling time was 10 hours. The post-combustion material exhibited a homogeneous final structure (2052). Once combustion was complete, the magnesium was observed to be clearly distributed. The sample structure showed smaller differences compared to its pre-combustion profile.

[0122] The combination of magnesium with regolith-based thermite has been observed to be a viable solution for providing combustion and heating solutions for solar surface devices. In particular, a mixture of 20% magnesium as fuel and JSC-1A as an oxidizer, with tailored properties such as ball milling time, can offer broad applicability due to its limited explosive potential. The production conditions can be tailored to suit the energy objectives and reaction intensity. The combustion by-products can be used as construction materials, thereby providing in-situ resource utilization.

[0123] Functions can be designed into the energetic particles and use cases can be tailored in the manufacturing process to optimize for heating and / or combustion applications.

[0124] According to one embodiment, heating may be achieved using electromagnetic radiation and / or non-radiative methods of magnetic particles using magnetic and / or electromagnetic heating.

[0125] According to one embodiment, in situ resource utilization can be performed and materials can be sourced from objects on Earth or in outer space.

[0126] According to one embodiment, functional particle implementations are provided that use engineering and manufacturing processes to induce precise and specific behavior. An example is engineering a specific catalyst to drive a chemical reaction. When the optimal amount of energy to drive the reaction is known, the catalyst is created and added to the reaction.

[0127] According to one embodiment, propulsion applications include in-situ resource utilization in space (ISRU / ISRP). On Earth, propulsion applications include nanothermites and core-shell nanothermites, for example, the inner and outer shells of the core-shell can be designed for various heating and / or combustion applications.

[0128] According to one embodiment, power and heating applications include in-situ resource utilization in space (ISRU / ISRP). On Earth, power and heating applications include cogeneration of electricity and heat for industrial and commercial applications. According to one embodiment, energy storage applications include the production of functional materials for fuels and battery components. For example, additive printing of battery components including Li-ion, sodium-based, aluminum-based, iron-based, etc. In other embodiments, multifunctional materials may be incorporated to fabricate components for thermophotovoltaic systems and / or thermal batteries.

[0129] According to one embodiment, the space construction process may include the use of aerogel, a lightweight and strong material for space construction, and welding for joining applications.

[0130] According to one embodiment, the construction process on Earth may involve the use of locally sourced materials.

[0131] According to one embodiment, applications and use cases include propulsion, energy generation, and space construction. Additionally, the material can be additively manufactured into user-defined shapes and multi-dimensional configurations. In other implementations, other configurations and types of thermite can use the in-situ material as a source of fuel and / or oxidizer.

[0132] According to one embodiment, as shown in FIG. 21, the multi-functional tilting structure may be a kinematic system, a deployable system, and / or an inflatable system.

[0133] According to one embodiment, as shown in FIG. 22, the multifunctional gradient material can be combined with a detachable electromagnetic system, where an array of magnets and / or electromagnets can be used to combust, heat by sintering the material, and / or induce motility using magnetohydrodynamics.

[0134] According to one embodiment, as shown in FIG. 23 , a multi-source heating assembly can heat and combust an energetic material to heat a working fluid, which can be circulated through a connected network of multiple channels for circulating the working fluid (gas, liquid, and / or phase-change material) containing the energetic material. In other implementations, the network of channels is embedded in the structure or outputs a feed to other chambers to maintain a user-defined temperature. In other implementations, magnets and / or electromagnets can be used to move magnetic energy particles from one chamber to another. In other implementations, the multi-source heating assembly uses one or more of the following radiation sources: laser, maser, MW, millimeter wave, and other electromagnetic radiation. In other implementations, a wireless power transmission transceiver using electromagnetic radiation may be combined with the multi-source heating assembly.

[0135] According to one embodiment, as shown in FIG. 24 , the multi-source heating assembly uses induction for heating and / or combustion applications. Additionally, energetic materials may be heated and / or combusted with output to a by-product capture / recycling system or directed to an exhaust nozzle. In other implementations, other non-radiative sources using one or more of inductive heating, inductive coupling, and / or magnetic coupling systems may be used in the multi-source heating assembly. In other implementations, a wireless power transmission transceiver using non-radiative transmission may be combined with the multi-source heating assembly using one or more of inductive heating, inductive coupling, and / or magnetic coupling systems.

[0136] According to one embodiment, the energetic particles may be optimized for other applications and the following use cases:

[0137] Metal ion implantation: In semiconductor manufacturing, metal ions can be used in ion implantation processes to modify the electrical properties of semiconductors, which is important in microelectronics for creating transistors and integrated circuits.

[0138] Surface Coatings and Treatments: Energetic metallic particles, such as metal ions or clusters, can be used in physical vapor deposition (PVD) and chemical vapor deposition (CVD) processes to deposit thin metallic coatings on a variety of substrates. These coatings can provide corrosion resistance, wear resistance, and improved aesthetics in applications including automotive parts, architectural finishes, and cutting tools.

[0139] Metal nanoparticle synthesis: Energetic particles can be used to create metal nanoparticles of controlled size and composition. These nanoparticles have applications in catalysis, electronics, sensors, and medical imaging.

[0140] Materials testing and characterization: Energetic metallic particles, such as high-energy electrons or ions, can be used in materials characterization techniques, including Auger electron spectroscopy and Rutherford backscattering spectroscopy, to analyze the composition and structure of materials, including metals and alloys.

[0141] Ion Beam Analysis (IBA): IBA techniques, including Rutherford Backscattering and Nuclear Reaction Analysis, use metal ions to determine the elemental composition and depth profiling of materials. These techniques may be present in materials science and archaeology, among other fields.

[0142] Radiation Shielding: High density metallic materials and alloys, including lead and tungsten, can shield against energetic particles such as x-rays and gamma rays to protect workers and equipment from radiation exposure in medical and industrial applications.

[0143] Metallurgical Research: Energetic metal particles can be used in metallurgical research to study the phase transitions, crystallography, and mechanical properties of metals and alloys. This research informs the development of new materials and processes for a variety of industries.

[0144] Nuclear Fusion: In the pursuit of controlled nuclear fusion as a future energy source, metal particles containing deuterium and tritium isotopes could serve as fuel for the fusion reaction. Energetic metal ions are utilized to heat and confine a plasma in an experimental fusion reactor.

[0145] Surface modification and hardening: Metal energetic particles can be used to change the surface properties of metals, for example, hardening, alloying, or introducing desirable surface features. These techniques find applications in the aerospace, automotive, and manufacturing industries.

[0146] Metallurgy in space: Metal particles are used in various ways in space exploration, for example to shield spacecraft against cosmic rays and micrometeorites, or for propulsion in electric ion thrusters.

[0147] Although the above description provides one or more example devices, methods, or systems, it will be recognized that other devices, methods, or systems may be within the scope of the claims, as interpreted by one of ordinary skill in the art.

Claims

1. Energetic particles containing metallic fuel A multifunctional gradient material comprising a regolith-based oxidizer, The energetic particles and the regolith-based oxidizer are mixed to form an energetic material.

2. The energetic particles are containing 20% ​​to 40% w / w micromagnesium, The material is ball milled for up to 5 hours. The multifunctional gradient material according to claim 1 .

3. 10. The multifunctional gradient material of claim 1, formed into at least one of a pellet or an independent measuring unit of energetic material.

4. The multifunctional gradient material of claim 1 , formed into a multidimensional structure.

5. The multifunctional gradient material according to claim 1 , wherein the regolith-based oxidizer is a regolith analog.

6. 1. A method of heating using a multifunctional energy gradient material, comprising: providing energetic particles including a metal fuel and a regolith-based oxidizer mixed with the energetic material; and providing energy to the energetic material using an energy source; method.

7. The method of claim 6 , wherein the energy source is a multi-source heating assembly.

8. 1. A method for manufacturing a multifunctional energy gradient material, comprising: adding micromagnesium and a regolith-based oxidizer to a ball mill at 20% to 40% w / w of said micromagnesium; forming the energetic material by ball milling the micromagnesium and the regolith-based oxidizer for up to 5 hours; and heating the energetic material until it ignites. method.

9. 10. The method of claim 8, wherein igniting the energetic material comprises igniting the energetic material in a vacuum.

10. The method of claim 8 further comprising forming the energetic material into pellets.

11. The method of claim 8 , wherein the energetic material comprises one or more of a microthermite and a nanothermite.

12. 7. The method of claim 6, further comprising exposing the energetic material to at least one of laser, maser, microwave, millimeter wave, infrared or other electromagnetic radiation.

13. 10. The method of claim 8, wherein heating the energetic material comprises inductive heating using magnets and / or electromagnets.

14. The method of claim 8 , wherein heating the energetic material comprises non-radiative heating.

15. The method of claim 8 , wherein the regolith-based oxidizer is a regolith analog.

16. The method of claim 8 further comprising collecting post-combustion by-products.

17. The method of claim 8 , further comprising harvesting the lunar-origin regolith-based oxidizer in situ.

18. 10. The method of claim 8, further comprising in situ harvesting the regolith-based oxidizer from at least one of a Martian, an asteroid, and a space source.

19. 10. The method of claim 8, further comprising recovering energetic material from space debris and other material from human activity.

20. 10. The method of claim 8, further comprising passing the energetic material through a connected network of multiple channels for circulating the energetic material.

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