FUNCTIONALLY GRADED, IN-SITU MANUFACTURABLE ENERGETICS AND METHODS FOR HEATING

Abstract

A functionally-graded regolith-based energetic fuel and methods for powering heating systems to help humans and equipment operate in harsh lunar conditions are disclosed. A muti-functionally graded material includes energetic particles comprising a metallic fuel and a regolith-based oxidizer. The energetic particles and the regolith-based oxidizer are mixed to form an energetic material. A regolith-based combustible material is disclosed comprising: micro-magnesium in 20%, 30% or 40% w/w; and a regolith-based oxidizer, wherein the material is ball milled for up to 5 hours.

Claims

1. A muti-functionally graded material comprising: energetic particles comprising a metallic fuel; and a regolith-based oxidizer; wherein the energetic particles and the regolith-based oxidizer are mixed to form an energetic material.

2. The muti-functionally graded material claim 1, wherein the energetic particles comprise: micro-magnesium in 20% to 40% w/w, wherein the material is ball milled for up to 5 hours.

3. The muti-functionally graded material of claim 1, shaped into at least one of: pellets or independent measured units of energetic materials.

4. The muti-functionally graded material of claim 1, shaped into multi-dimensional structures.

5. The muti-functionally graded material of claim 1, wherein the regolith-based oxidizer is a regolith simulant.

6. A method for heating using multi-functionally graded energetic material comprising: providing energetic particles comprising metallic fuels and a regolith-based oxidizer mixed into an energetic material; and energizing the energetic material using an energy source.

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

8. A method for manufacturing a multi-functional graded energetic material comprising: adding micro-magnesium and a regolith-based oxidizer to a ball mill, wherein the micro-magnesium is in 20% to 40% w/w; forming the energetic material by ball milling the micro-magnesium and the regolith-based oxidizer for up to 5 hours; and heating the energetic material until ignition.

9. 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 shaping the energetic material into pellets.

11. The method of claim 8 wherein the energetic material comprises one or more of: microthermites and nanothermites.

12. The method of claim 6, further comprising exposing the energetic material to at least one of the following: lasers, masers, microwaves, mm-wave, infrared or other electromagnetic radiation.

13. 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 simulant.

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

17. The method of claim 8 further comprising in-situ harvesting the regolith-based oxidizer from a lunar source.

18. The method of claim 8 further comprising in-situ harvesting the regolith-based oxidizer from at least one of the following: mars, asteroids, and a source in space.

19. The method of claim 8 further comprising harvesting energetic materials from space debris, and other materials from human activity.

20. The method of claim 8 further comprising passing the energetic material through a connected network of a plurality of channels for cycling energetic materials.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0023] The drawings included herewith are for illustrating various examples of articles, methods, and apparatuses of the present specification. In the drawings:

[0024] FIG. 1A-1F are diagrams of various fuel grain structures and corresponding thrust profiles;

[0025] FIG. 2 is diagram of a functionally layered fual;

[0026] FIG. 3A is an exemplary regressive (simple) thrust profile;

[0027] FIG. 3B is an exemplary two peaks (complex) thrust profile;

[0028] FIG. 3C is an exemplary pressure vs. altitude curve for optimal thrust during ascent;

[0029] FIG. 4 is a flow chart of a method for optimizing the conceptual design of functionally-graded solid rocket engines;

[0030] FIG. 5A is an ideal burn velocity profile for simple thrust;

[0031] FIG. 5B is an ideal burn velocity profile for complex thrust;

[0032] FIG. 5C is an ideal burn velocity profile for pressure matching;

[0033] FIGS. 6A-6C are optimized thrust profiles, based on the ideal burn velocities shown in FIGS. 5A-5C, respectively;

[0034] FIGS. 7A to 7B are scanning electronic microscope (SEM) images of the pre-combustion sample structure;

[0035] FIGS. 8A to 8C are scanning electronic microscope (SEM) images of the pre-combustion sample structure;

[0036] FIG. 9 are differentiative scanning calorimetry (DSC) and

[0037] thermogravimetric analysis (TGA) showing energy release for regolith-based combustion material with different ball mill times, according to several embodiments;

[0038] FIG. 10 is a high-speed video frame capturing combustion of sample structure in an open-air ignition, according to an embodiment;

[0039] FIG. 11 is a high-speed video frame capturing combustion of sample structure in an open-air ignition, according to an embodiment;

[0040] FIG. 12 is a high-speed video frame capturing combustion of sample structure in an open-air ignition, according to an embodiment;

[0041] FIG. 13 is a high-speed video frame capturing combustion of sample structure in vacuum ignition, according to an embodiment;

[0042] FIG. 14 are pyrometric thermal video frames indicating temperature of combustion and the emissivity of the pellets, according to an embodiment;

[0043] FIG. 15 are pyrometric thermal video frames indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment;

[0044] FIG. 16 are pyrometric thermal video frames indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment;

[0045] FIG. 17 is a pyrometric thermal video frame indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment;

[0046] FIG. 18 is a pyrometric thermal video frame indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment;

[0047] FIGS. 19A and 19B are scanning electron microscope images of a sample pre-combustion and post-combustion respectively, according to an structure embodiment; and

[0048] FIGS. 20A and 20B are scanning electron microscope images of a sample structure pre-combustion and post-combustion respectively, according to an embodiment;

[0049] FIG. 21 is a block diagram of multi-functionally graded structures, in accordance with an embodiment;

[0050] FIG. 22 is a block diagram of multi-functionally graded materials coupled with a detachable electromagnetic system, in accordance with an embodiment;

[0051] FIG. 23 is a block diagram of a multi-source heating assembly, in accordance with an embodiment; and

[0052] FIG. 24 is a block diagram of a multi-source heating assembly, in accordance with an embodiment.

DETAILED DESCRIPTION

[0053] Various apparatuses or processes will be described below to provide an example of each claimed embodiment. No embodiment described below limits any claimed embodiment and any claimed embodiment may cover processes or apparatuses that differ from those described below. The claimed embodiments are not limited to apparatuses or processes having all of the features of any one apparatus or process described below or to features common to multiple or all of the apparatuses described below.

[0054] A description of an embodiment with several components in communication with each other does not imply that all such components are required. On the contrary, a variety of optional components are described to illustrate the wide variety of possible embodiments of the present invention.

[0055] Further, although process steps, method steps, algorithms or the like may be described (in the disclosure and/or in the claims) in a sequential order, such processes, methods and algorithms may be configured to work in alternate 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. The steps of processes described herein may be performed in any order that is practical. Further, some steps may be performed simultaneously.

[0056] When a single device or article is described herein, it will be readily apparent that more than one device/article (whether or not they cooperate) may be used in place of a single device/article. Similarly, where more than one device or article is described herein (whether or not they cooperate), it will be readily apparent that a single device/article may be used in place of the more than one device or article.

[0057] In-situ resource utilization for preparing combustion materials provides key advantages for sustained human presence and potential industrial applications in outer space. In-situ resource utilization refers to the use of locally found resources to prepare desired articles. For example, the moon's regolith includes fragmented materials composed of minerals such as anorthite and ilmenite, and a variety of metal oxides. These minerals include elements such as aluminum and iron, which are chief components of thermite, an effective combustion material with significant applications. For long-duration space missions, combustion materials prepared using in-situ resources may be used as a fuel for propulsion or infrastructure activities such as construction and metal extraction. Consequently, the overall payload for space missions may be reduced, resulting in cost savings, size and cargo weight reduction, and reducing risks of supply shortages. The regolith-based thermite may be used to provide heating to improve survivability of humans and equipment in the lunar night or for ambitious lunar-polar missions.

[0058] In other examples, Mars' regolith includes fragmented materials composed of minerals, rocks, dust, sand, soil, carbonates, silicates, sulfates, phyllosilicates, perchlorates, metals and a variety of 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 chief components of thermite, an effective combustion material with significant applications. For long-duration space missions, heating and/or combustion materials prepared using in-situ resources may be used as a fuel for propulsion or power generation or infrastructure activities such as construction, welding and metal extraction. In other implementations, metals, metal alloys, rock and other minerals may come from asteroids, other moons, planets, planetoids, and other celestial bodies. Consequently, the overall payload for space missions may be reduced, resulting in cost savings, size and cargo weight reduction, and reducing risks of supply shortages. The in-situ regolith-based thermite may be used to provide heating and combustion to improve survivability of humans and equipment in the Lunar night or support lunar-polar missions, Martian night, and or for other ambitious space missions.

[0059] Recent years have seen drastic changes in solid and hybrid propulsion technology as new propulsion paradigms are taking hold. Recent works have proposed the integration of hypergolic additives based on a metal-organic framework (MOF) (Jobin, O., et al., Metal-organic frameworks as hypergolic additives for hybrid rockets, Chemical Science, vol. 13, no. 12, pp. 3424-3436, 2022). Concurrently, new manufacturing processes of solid-state fuels, via additive manufacturing (AM) techniques, are opening up new design opportunities for a novel class of solid and hybrid state propulsion systems.

[0060] Historically, the selection of the fuel grain structure represented the optimal approach to achieving a desired thrust-time curve.

[0061] Various fuel grain structures and corresponding thrust profiles (100) are shown in FIGS. 1A-1F. For a known propellant burn- and regression rate, the imposed grain structure causes a change in area, concomitantly combustion heat release, with time. For example, a relatively simple tubular grain design (FIG. 1A), results in a progressive (increasing) thrust curve as the burn area increases as the grain regresses towards the outer wall; a more complex star-type fuel grain (see FIG. 6A, right), is needed for a neutral (constant) thrust curve. The design of these grain structures make the use of burn-back models. Furthermore, these complex structures are prone to significant erosive burning, structural integrity issues, and manufacturability constraints.

[0062] The new opportunities afforded by AM of energetic fuels, as discussed above, means that the solid and hybrid engine design considerations can shift away from complex geometrical fuel grains to modify the thrust-time curve and move towards new design considerations by functionally grading single- or even multi-fuel propellant engines. By layering various fuels with spatially-varying binder and propellant compositions and/or density, it is possible to effectively construct a matching thrust-time profile without the need for complex fuel grains, thus opening new opportunities for novel engine design and optimization considerations.

[0063] Various energetic materials and nanothermite aerogels are summarized in Table 1 include Al as the fuel and Bi.sub.2O.sub.3 as the oxidizer and a UV or other polymer binder. Other metallic fuel and oxidizer combinations may be possible. For example, the metallic fuel may be Mg, Si, Fe, etc. and the oxidizer may be a fluoropolymer, iodine oxide or a metal oxide (e.g., Fe.sub.2O.sub.3, SiO.sub.2, MgO, etc.). Varying the composition of the fuel can allow for the burn rate to be controlled as shown in Table 1. While the nano thermite aerogels summarized in Table 1 show extremely fast burn rates, it demonstrates the potential to adapt this technology for lower burn rate fuels. The aerogels described herein may be formed into fuel grains for various Earth, Luna and in-space applications to tunable heating and combustion implementations in solid rocket motors (SRMs) and/or hybrid rocket motors.

TABLE-US-00001 TABLE 1 Energetic data of rGO/Al/Bi.sub.2O.sub.3 products from DSC and combustion results. Onset Peak Energy release Linear Aerogel Name Temp. Temp. before Al melt burning rate rGO(20%)/ 513 C. 552 C. 248 0.39 Al/Bi.sub.2O.sub.3 42 J/g 0.02 m/s rGO(10%)/ 518 C. 554 C. 377 8.0 Al/Bi.sub.2O.sub.3 27 J/g 0.4 m/s rGO(5%)/ 516 C. 558 C. 415 10 Al/Bi.sub.2O.sub.3 48 J/g 0.5 m/s

[0064] A computational framework is provided to optimize the fuel grain structure to match a desired thrust curve profile. The framework includes two solvers with varying levels of fidelity, to efficiently optimize over a large parameter space. A system-level code (zero-dimension) linking combustor and nozzle systems was first developed to assess the overall behavior of the system. A quasi-one-dimensional code was then developed to incorporate the spatial variation and acoustic modes in the combustion chamber and nozzle for a given functionally-graded engine. Given the complex combustion kinetics of the solid fuel, which remains to a large extent poorly understood, simplified combustion and regression models were used as summarized below.

[0065] A system-level solver was first-developed using isentropic nozzle relations to estimate the vacuum thrust characteristics of the engine. The total thrust of an engine can be estimated based on the total pressure and temperature generated within the combustor for a given nozzle geometry. The equation for vacuum thrust can be recast as:

[00001]$\begin{array}{cc}{F}_{th,vac}={\frac{A_t{P}_e\left(M_e^2+1\right)}{M_e^2}\left[\left(2+\frac{-1}{+1}{M}_e^2\right)\right]}^{\frac{+1}{2\left(-1\right)}}& \left[1\right]\end{array}$

Where P and M are the pressure and Mach number, the subscript e indicates the nozzle exit; At and are respectively the throat area and the specific heat ratio.

[0066] The 0D model uses isotropic flow equations to relate the combustion chamber state (total pressure and temperature) to the nozzle exit state. 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 knowing the thermodynamic conditions in the engine can be computed:

[00002]$\begin{array}{cc}\frac{T_{\mathrm{comb}}}{T_e}=\left[1+\frac{-1}{2}{M}_e^2\right]& \left[4\right]\end{array}$$\begin{array}{cc}\frac{P_{\mathrm{comb}}}{P_e}={\left[1+\frac{-1}{2}{M}_e^2\right]}^{\frac{}{-1}}& \left[5\right]\end{array}$

[0067] The 0D model also makes use of the ideal rocket equation to find the relationship between altitude and time which will be important for one of the tests cases. Where v is the velocity of the rocket, u is the exit velocity of the combustion gas and m is the total mass off the rocket:

[00003]$\begin{array}{cc}{}_{v_i}^{v}dv=-u{}_{m_i}^m\frac{1}{m}dm& \left[9\right]\end{array}$

[0068] A quasi-one-dimensional solver was concurrently developed to account for spatial variations in the combustion chamber and nozzle, as well as investigate the acoustic coupling in a functionally-graded engine. The quasi-one-dimensional code solves the one-dimensional Navier-Stokes equations (conservation of mass, momentum, and energy). The spatial fluxes were computed via a fifth order WENO scheme and the equations were integrated in time with 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-described framework, for a given combustion chamber and nozzle geometry, the equations can be advanced and generated thrust can be computed. The thrust of the rocket can be directly computed, by assuming a perfect expansion in the nozzle, as: F.sub.th=mu.sub.e+(P.sub.eP.sub.atm) A.sub.e. As the equations are integrated in time, and the fuel regresses, the thrust profile curve can be estimated.

[0070] Functional grading (200) is achieved through the layering of different fuels (202, 204, and 206), as shown in FIG. 2. There are several parameters that are known to affect the performance of a SRM. The main parameters being the heat release, gas release, and burn rate of the fuel. The burn rate can be controlled by manipulating several fuel properties such as the chamber pressure, fuel density, fuel porosity, chemical composition, and physical composition. For additively manufactured energetic materials, several aspects of the physical composition may affect the burn rate such propellant loading, extra polymer additives, and binder material. The variation of these parameters can cause the burn rates of solid fuels to range from the order of millimeters per second to hundreds of meters per second. Thus, by layering the different fuel with varying combustion characteristics, a functionally-graded engine can be designed by varying the combustion properties during the burn. Thus, through an optimal fuel layering, a mission-specific thrust profile can be achieved.

[0071] To illustrate how this framework can be used, three well-defined test cases were considered: a simple, regressive, thrust profile; a complex thrust profile with multiple peaks, as proposed by Federici et al.; and a conceptual case where the total pressure conditions in the engine are tuned for a perfect expansion in the nozzle during ascent. The test cases (300) are illustrated in FIGS. 3A-3C.

[0072] Test case 1 (FIG. 3A), corresponds to a regressive profile which is usually achieved by using a complex grain geometry such as a double anchor geometry. Test case 2 (FIG. 3B) is a more complex thrust profile two peaks, which was presented by Federici et al. Finally, the third test case (FIG. 3C) is a conceptual case wherein the burn characteristics of the engine are modified to match a perfectly expanding nozzle flow during ascent. The ideal pressure vs. altitude curve (302), shown in FIG. 3C, will be converted into a pressure vs. time curve in the optimization of the fuel grain. The ideal rocket equation, as mentioned previously, relates the velocity of the rocket to the mass of the rocket and exit velocity of the combustion gases. Using this information, as the combustion occurs the mass of the rocket, the velocity and the height of the rocket will be calculatedthe atmospheric pressure will thus change in time. From this, an ideal exit pressure (that perfectly matches the atmospheric pressure) as a function of time curve is created and optimized for.

[0073] FIG. 4 shows an optimization framework, for the conceptual design of functionally-graded solid rocket engines that matches the thrust-time, and pressure-time profiles for a given mission. Given the ability to create a functionally-graded engine, the burn rate of the fuel and layer thickness were selected as the independent parameters for the optimization problem and a tubular grain (FIG. 1A) was assumed for simplicity. Given the modest parametric space, a brute-force profile matching optimization approach was implemented, although the modularity of the code allows for any constrained optimization algorithm to be implemented. In such a case the 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 starts, in the 0D code, advancing the equation sets in time to determine the burn velocity 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 velocity and the radius at which this occurs is tracked, allowing for an ideal radial distance vs burn rate curve to be plotted.

[0075] At 16, once the optimal burn rate at each radial value is known, the fuel regime is broken in to 4 approximate sections which act as the layers within the fuel. The burn rate and thickness of each section are bounded after analyzing the ideal burn rate curve. To aid in the bounding of the burn rates and layer thicknesses, a coarser sample of velocities are used to create a new ideal velocity vs. burn rate curve. A sample of 4 to 5 burn rates are used within the minimum and maximum of the ideal curve. Although less accurate, this can be used to determine the layer properties of the fuel. As the velocity fluctuates between the coarser values, bins are created where the ideal velocity lies. At 18, the range of velocity inputs for optimization are bounded by the bin minimum and maximum. The radius input for each layer is bound by the radial position at which the bin value changes 10% without overlapping the bounds of neighboring layers. The layer properties are optimized within these defined bounds. The size of these layers and their burn rates are then be optimized to best suit the desired thrust or pressure profile.

[0076] At 20, the optimization of each layer thickness and burn rate is completed using a heuristic approach using a range of random inputs for the layer thickness and burn rate to determine the most accurate solution. The model will save the layer thicknesses and burn rates of the model with the lowest error. The error between the desired profile and the profile generated by the model as the sum of the absolute difference between the two models across all time steps. Finally, the optimized layered solution is passed to the 1D code to assess the spatial variations and acoustics in the engine.

[0077] Referring to FIG. 5A, shown therein is an exemplary ideal burn velocity profile (501) for simple thrust plotting the ideal burn velocity at each radial distance based on the time step. FIG. 6A shows the optimized solution (601) based on the heuristic method of selecting random burn velocities and layer thicknesses approximated based on the ideal velocities shown in FIG. 5A. The curve 602 represents the ideal profile. The curve 603 represents the optimized. 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 thickness, outer radii, and burn velocities for the simple case are shown in Table 2 below.

TABLE-US-00002 TABLE 2 Simple Profile Optimal Results Layer Thickness (mm) Radius (mm) r.sub.b (mm/s) 1 160 360 8.5 2 70 430 6.5 3 60 490 5.5 4 80 570 4.5

[0078] For the simple case it can be seen that the layer thickness of the fuel differs as combustion of the fuel progresses. During the initial peak in thrust at the start of the trajectory the thickness of the first layer accounts for 43% of the total fuel radius. Whereas the second, third, and fourth layers of fuel account for 19%, 16%, and 22% of the thickness respectively.

[0079] Referring to FIG. 5B, shown therein is an ideal burn velocity profile (511) for complex thrust plotting the ideal burn velocity at each radial distance based on the time step. FIG. 6B shows the optimized solution (611) based on the heuristic method of selecting random burn velocities and layer thicknesses approximated based on the ideal velocities shown in FIG. 5A. The curve 612 represents the ideal profile. The curve 613 represents the optimized. It should be noted that the range of inputs is different from that of the simple case (FIGS. 5A and 6A) as the ranges of values are based on the optimal burn rate at each radius for both models. Due to this, a wider range of inputs were used for the second test case when setting up the optimization trial. It can be seen that with 4 simple layers of fuel with different burn rates, the thrust profile of the modeled SRM can closely match the desired thrust profile of the rocket. The optimal radii and burn velocities for the complex case are shown in Table 3 below.

TABLE-US-00003 TABLE 3 Complex Profile Optimal Results Layer Thickness (mm) Radius (mm) r.sub.b (mm/s) 1 160 360 11 2 50 410 6.5 3 110 520 4.5 4 30 550 3.5

[0080] For the complex case it can be seen that the layer thickness of the fuel differs as combustion of the fuel progresses. During the initial peak in thrust at the start of the trajectory the thickness accounts for 46% of the fuel thickness. Whereas the second, third, and fourth layers of fuel account for 14%, 31%, and 9% of the thickness respectively. This is a greater variation in layer thicknesses than the simple case. It is also seen that there is a greater variation in the burn rates for the complex case.

[0081] In the third test case, the goal is to match the exhaust pressure of the combustion gas to the atmospheric pressure. Referring to FIG. 5C, shown therein is an ideal burn velocity profile (521) for pressure matching plotting the ideal burn velocity at each radial position. FIG. 6C shows the optimal pressure profile (621) over time using four distinct layers. The curve 622 represents the ideal profile. The curve 623 represents the optimized. As mentioned in the test case subsection above, the ideal pressure profile over time is found using the ideal rocket equation and determining the height of the rocket over time based on the mass and gas exit velocity. It can be seen in this case that the exit pressure of the rocket follows the ideal case test fairly well. The optimal burn rates and layer thickness are shown in Table 4 below.

TABLE-US-00004 TABLE 4 Pressure Matching Profile Optimal results: Layer Thickness (mm) Radius (mm) r.sub.b (mm/s) 1 5 205 8.5 2 30 235 5 3 15 250 3 4 9 259 1.5

[0082] For all three test cases if more layers were to be used, a closer match could be generated by the model. In the future as fuel may be functionally graded, the burn rate could be controlled throughout the fuel to create a more exact SRM. This would cause the optimized profiles to resemble the ideal profiles modeled herein. Using an additive manufacturing process (e.g., a 3D printing method) the material extrusion width may be used as the minimum layer thickness which will lead to more tunable profiles.

[0083] The frameworks and methods described above for functionally graded rocket engines can be extended to the development of functionally graded energetics for various Earth and in-space applications, in particular for propulsion and power generation, construction, storage systems and controlled heating of equipment. For example, the lunar environment is one of the harshest that humans have operated in. Temperatures near the equator can drop to 140 K; temperatures near the poles can drop as low as 25K in permanently shadowed areas. Such low temperatures can be damaging to equipment and humans.

[0084] Existing methods of heating in Lunar or in-space environments have limitations. Radioactive heat generators are controlled goods that are typically not accessible to private operators. Batteries hare often large and heavy and may need to be tailor-made for the Lunar climate. Fuel such as liquid hydrogen/liquid oxygen LH/LOX has not yet been developed to the point where water can be extracted from the lunar surface, or otherwise entails other fuel to be transported as cargo. Thus, in-situ production of multi-functionally graded fuels and materials are desirable.

[0085] In-situ resource utilization (ISRU) is the process of using locally sourced resources to achieve a desired goal. For in-space applications, ISRU reduces mission cost by reducing size and weight of cargo. Lunar/Martian regolith and/or materials in asteroids are composed of a series of metal oxides and can be used as thermite if a fuel is added to generate heat. Work in ISRU has typically been limited to creating construction materials but neglects heat release.

[0086] In an embodiment, the additive manufacturing process as described above, including as applied on 3D printing of polymer-free nanothermite aerogels, is used for regolith based nanothermite preparation.

[0087] The regolith-based combustion material uses micro-magnesium as a fuel, and JSC-1A as the oxidizer. Micro-magnesium, due to its increased surface area when ground to micro or nano scales, provides high reactivity and a faster exothermic response compared to traditional thermite mixtures. Further, combustion materials using micro-magnesium as fuel ignites at lower temperatures, thus making it effective for lunar surface. JSC-1A is a lunar regolith stimulant with minerology and chemical properties similar to the lunar regolith. In other implementation, regolith from in-situ environments are used as a source for the fuel and/or oxidizer

Test Cases

[0088] A framework for sampling a regolith-based combustion material is provided, using a lunar regolith stimulant as a source of the oxidizer. The test cases have a varied composition of magnesium in the combustion materials i.e., the materials with compositions of 2-5%, 30%, and 40% magnesium were chosen. Secondly, to increase the surface area, the combustion material was subjected to a ball milling process for 0 hour, 2 hours, 5 hours, and 10 hours respectively.

[0089] Overall, 12 samples were prepared for combustion testing in combinations of percentage composition of magnesium and ball milling times. For the combustion tests, 30 mg pellets of the samples were prepared under 250 psi pressure for 1 minute, for the combustion tests. The pressure may be calibrated or changed to suit the needs of the experiment.

[0090] Referring to FIGS. 7A to 7B, shown therein are scanning electronic microscope (SEM) images of the pre-combustion sample structure. FIG. 7A provides an overall view (701) of the sample structure without any magnesium and processed for 2 hours in the ball mill. FIG. 7B provides an overall view (711) of the sample structure with magnesium processed for 2 hours in the ball mill. The ball mill may result in the creation of smaller fragments (712) of the sample structure. As seen from the images, no significant difference between the samples was observed. In both samples, many large particles are visible. Further, as observed in FIG. 7C, the magnesium (722) is not evenly distributed.

[0091] Referring to FIGS. 8A to 8C, shown therein are scanning electronic microscope (SEM) images of the pre-combustion sample structure. FIG. 8A provides an overall view (801) of the sample structure without any magnesium and processed for 10 hours in the ball mill. FIG. 8B provides an overall view (811) of the sample structure with magnesium and processed for 10 hours in the ball mill. The ball mill may result in the creation of smaller fragments (812) of the sample structure. As seen from the images, few larger particles are observed due to the longer ball milling time. However, larger magnesium particles are observed in FIG. 8B. Further, as observed in FIG. 8C, the magnesium (822) is not evenly distributed.

[0092] Referring to FIG. 9, shown therein are differentiative scanning calorimetry (DSC) and thermogravimetric analysis (TGA) (900) showing energy release for regolith-based combustion material with different ball mill times, according to several embodiments.

[0093] The energy release for the sample regolith-based combustion material was calculated using the differentiative scanning calorimetry (DSC) and thermogravimetric analysis (TGA) to note the thermal properties of the sample structures. Test cases confirmed that both the percentage composition of the magnesium in the sample structure and the milling time impacted energy release.

[0094] Based on the DSC-TGA energy release values, it was observed that the sample structures with mill time over 5 hours (40% at 902, 30% at 904, and 20$ at 906) demonstrate little difference in energy release.

TABLE-US-00005 TABLE 5 Energetic data of the sample regolith-based combustion material from DSC and combustion results Sample Sample Sample structure structure structure Mill Time with 20% with 30% with 40% (hours) Magnesium Magnesium Magnesium 0 355 J/g 514 J/g 743 J/g 2 384 J/g 528 J/g 774 J/g 5 398 J/g 542 J/g 792 J/g 10 405 J/g 557 J/g 797 J/g

[0095] The energy release may provide a heat source for a variety of applications. The composition and ball mill time may be optimized and adjusted to meet specific objectives.

[0096] Referring to FIG. 10, shown therein are high-speed video frames (1000) capturing combustion of sample structure in an open-air ignition, according to an embodiment.

[0097] The sample structure used in the combustion was composed of 20% magnesium with a ball milling time of 5 hours. The video frames show burning at different stages such as triggering the flame (1002), ignition (1004), propagation of the combustion (1006), and finished/exhaustion (1008). The propagation of the combustion was non-violent, and no explosions or ejections were observed. The sample structure commenced burning with exposure to the laser beam, and the burning continued after the laser was turned off. A small flame was observed once the reaction had propagated. Since the combustion was performed in open air, oxygen became part of the reaction.

[0098] Referring to FIG. 11, shown therein are high-speed video frames (1100) capturing combustion of sample structure in an open-air ignition, according to an embodiment.

[0099] The sample structure used in the combustion was composed of 30% magnesium with a ball milling time of 5 hours. The video frames show burning at different stages such as triggering the flame (1102), ignition (1104), propagation of the combustion (1106), and finished/exhaustion (1108). The propagation of the combustion was violent, and explosions and some combustive ejections were observed. The explosiveness is likely attributed to the gasification of magnesium. The sample structure commenced burning with exposure to the laser beam, and the burning continued after the laser was turned off.

[0100] Referring to FIG. 12, shown therein are high-speed video frames (1200) capturing combustion of sample structure in an open-air ignition, according to an embodiment.

[0101] The sample structure used in the combustion was composed of 40% magnesium with a ball milling time of 5 hours. The video frames show burning at different stages such as triggering the flame (1202), ignition (1204), propagation of the combustion (1206), and finished/exhaustion (1208). The propagation of the combustion was most violent among the sample structures, and the sample exploded to multiple smaller pieces. The explosiveness is likely attributed to a more rapid gasification of magnesium. Further, additional magnesium causes formation of a larger volume of gas.

[0102] Referring to FIG. 13, shown therein are high-speed video frames (1300) capturing combustion of sample structure in vacuum ignition, according to an embodiment.

[0103] The sample structure used in the combustion was composed of 20% magnesium with a ball milling time of 5 hours. The video frames show burning at different stages such as triggering the flame (1302), ignition (1304), propagation of the combustion (1306), and finished/exhaustion (1308). The occurrence of the combustion demonstrated that atmospheric oxygen was not necessary for the reaction to take place. The propagation of the combustion included one small initial explosion, followed by a controlled combustion. The sample structures with higher magnesium quantity only demonstrated explosions, and the extinguished if an explosion occurred and the laser was not turned on.

[0104] Referring to FIG. 14, shown therein are pyrometric thermal video frames (1400) indicating temperature of combustion and the emissivity of the pellets, according to an embodiment.

[0105] Pyrometry was used to estimate the temperature of combustion and the emissivity of the sample structure pellets. The emissivity of the sample structure has a large effect on the temperature. The temperature of the sample structure pellet during the burn, near and behind flame from was found to be approximately 1700K. This temperature is considerably over the boiling point of magnesium (1365K). This demonstrated an emissivity value of approximately 35%.

[0106] Referring to FIG. 15, shown therein are pyrometric thermal video frames (1500) indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment.

[0107] In FIG. 15, the sample structure used in the combustion was composed of 20% magnesium with a ball milling time of 2 hours. The video frames show burning at different stages such as triggering the flame (1502), ignition (1504), propagation of the combustion (1506), and finished/exhaustion (1508).

[0108] Referring to FIG. 16, shown therein are pyrometric thermal video frames (1600) indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment.

[0109] In FIG. 16, the sample structure used in the combustion was composed of 20% magnesium with a ball milling time of 5 hours. The video frames show burning at different stages such as triggering the flame (1602), ignition (1604), propagation of the combustion (1606), and finished/exhaustion (1608). A large deformation of the samples was observed post-combustion.

[0110] Referring to FIG. 17, shown therein are pyrometric thermal video frames (1700) indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment.

[0111] The sample structure used in the combustion was composed of 20% magnesium with a ball milling time of 10 hours. The reaction included a combustion and the sample remained hot for approximately 10 seconds after the reaction had propagated. The video frames show burning at different stages such as triggering the flame (1702), ignition (1704), propagation of the combustion (1706), and finished/exhaustion (1708). Compared to the sample structures in FIGS. 9 and 10, a lesser post-combustion deformation of the same was observed.

[0112] Referring to FIG. 18, shown therein are pyrometric thermal video frames (1800) indicating temperature of combustion and the emissivity of the pellets in an open-air combustion, according to an embodiment.

[0113] The sample structure used in the combustion was composed of 30% magnesium with a ball milling time of 5 hours. The video frames show burning at different stages such as triggering the flame (1802), ignition (1804), propagation of the combustion (1806), and finished/exhaustion (1808). The reaction included a more rapid propagation of combustion and a violent reaction compared to other sample structures. Solid ejections were observed as leaving from the sample. The sample structure in combustion reached high temperatures and remained at high temperatures for a longer duration.

[0114] Referring to FIGS. 19A and 19B, shown therein are scanning electron microscope images of a sample structure pre-combustion (1900) and post-combustion (1950) respectively, according to an embodiment.

[0115] The sample structure used in the combustion was composed of 20% magnesium with a ball milling time of 2 hours. The post-combustion substance shows a homogenous final structure (1952). Magnesium was observed to be clearly distributed (1954) once the combustion was complete.

[0116] Referring to FIGS. 20A and 20B, shown therein are scanning electron microscope images of a sample structure pre-combustion (2000) and post-combustion (2050) respectively, according to an embodiment.

[0117] The sample structure used in the combustion was composed of 20% magnesium with a ball milling time of 10 hours. The post-combustion substance shows a homogenous final structure (2052). Magnesium was observed to be clearly distributed once the combustion was complete. The sample structure showed lesser difference compared to its pre-combustion profile.

[0118] The combination of magnesium and regolith-based thermite was observed to be a viable solution for providing combustion and heating solutions for equipment on the solar surface. Particularly, the mixture of 20% magnesium as a fuel and JSC-1A as an oxidizer with tailored properties such as ball milling time may provide wide applicability due to limited explosiveness. The manufacturing conditions may be tailored for energy objectives and reaction violence. The post-combustion by-products may be used as construction materials, thereby providing in-site resource use.

[0119] The function may be engineered into energetic particles, and the use cases may be tailored in the manufacturing process to optimize for heating and/or combustion application.

[0120] According to an embodiment, the heating may be achieved using electromagnetic radiation, and or non-radiative methods for magnetic particles using magnetic and/or electromagnetic heating.

[0121] According to an embodiment, the in-situ resource utilization may be performed, and materials may be sourced from the Earth or from objects in outer space.

[0122] According to an embodiment, the implementations of functional particles using engineering and manufacturing processes are provided, to induce precise and specific behaviors. Examples include engineering a specific catalyst to drive a chemical reaction. When the optimal amount of energy to drive the reaction is known, a catalyst is created to be added into the reaction.

[0123] According to an embodiment, the applications in propulsion include in-situ resource utilization (ISRU/ISRP) in outer space. On Earth, the applications in propulsion include Nanothermites and CoreShells Nanothermites, for example the inner and other shells of the coreshell may be engineered for various heating and/or combustion applications.

[0124] According to an embodiment, the applications in power and heating include in-situ resource utilization (ISRU/ISRP) in outer space. On Earth, the applications in power and heating include co-generation of electricity and heat generation for industrial and commercial applications. According to an embodiment, the application for stored energy includes manufacturing of the fuels and functional materials for parts of a battery. For example, additively printing parts of the battery including Li-ion, sodium-based, aluminum-based, iron-based or others. In other embodiments, multi-functional materials may be incorporated to manufacture parts of a thermophotovoltaic system and/or a thermal battery.

[0125] According to an embodiment, the construction process in space may involve aerogels i.e., using light and strong materials for space structures and welding for joining applications.

[0126] According to an embodiment, the construction process on Earth may involve use of materials sourced locally.

[0127] According to an embodiment, the applications and use cases include propulsion, energy generation, and construction in outer space. Furthermore, materials can be additively manufactured info user-defined shapes and multi-dimensional configurations. In other implementations, other configuration and types of thermites may be using in-situ materials as sources for fuel and/or oxidizer.

[0128] According to an embodiment, as shown in FIG. 21, multi-functionally graded structures may be kinematic, deployable, and or inflatable systems.

[0129] According to an embodiment, as shown in FIG. 22, multi-functionally graded materials may be coupled with a detachable electromagnetic system, where an array of magnets and/or electromagnets may be used for combustion, heating via sintering of the materials and or induce motility using magnetohydrodynamics.

[0130] According to an embodiment, as shown in FIG. 23, a multi-source heating assembly, where energetic materials can be heated and combusted to heat a working fluid, which can be cycled through a connected network of a plurality of channels for cycling a working fluid (gas, liquid, and/or phase change material) with energetic materials. In other implementations, the network of channels are embedded into a structure, and or output feed into other chambers, to maintain a user defined temperature. In other implementations, magnets and or electromagnets may be used to move magnetic energetic particles from one chamber to another. In other implementations, the multi-source heating assembly where radiative sources are used with one or more of the following; lasers, masers, MW, mm-wave, other electromagnetic radiation. In other implementations, wireless power transmission transceivers using electromagnetic radiation, may be coupled with the multi-source heating assembly.

[0131] According to an embodiment, as shown in FIG. 24, a multi-source heating assembly, where the multi-source heating assembly uses induction for heating and or combustion applications. Additionally, energetic materials may be heated and or combusted with outputs to the capture byproducts/a recycling system or directed into the exhaust nozzle. In other implementations, other non-radiative sources using one or more of the following: inductive heating, inductively-coupled and/or magnetically-coupled system, may be used in the multi-source heating assembly. In other implementations, wireless power transmission transceivers using non-radiative transmission, using one or more of the following: inductive heating, inductively-coupled and/or magnetically-coupled system, may be coupled with the multi-source heating assembly.

[0132] According to an embodiment, the energetic particles may be optimized for other application and use cases including:

[0133] Metallic Ion Implantation: In semiconductor manufacturing, metallic ions may be used in ion implantation processes to modify the electrical properties of semiconductors. This is crucial for creating transistors and integrated circuits in microelectronics.

[0134] Surface Coatings and Treatments: Metallic energetic particles, such as metal ions or clusters, may be used in physical vapor deposition (PVD) and chemical vapor deposition (CVD) processes to deposit thin metallic coatings on various substrates. These coatings may provide corrosion resistance, wear resistance, and improved aesthetics in applications including automotive parts, architectural finishes, and cutting tools.

[0135] Metallic Nanoparticle Synthesis: Energetic particles may be used to create metallic nanoparticles of controlled size and composition. These nanoparticles find applications in catalysis, electronics, sensors, and medical imaging.

[0136] Materials Testing and Characterization: Energetic metallic particles, such as high-energy electrons or ions, may be used for materials characterization techniques including Auger electron spectroscopy and Rutherford backscattering spectroscopy to analyze the composition and structure of materials, including metals and alloys.

[0137] Ion Beam Analysis (IBA): IBA techniques, including Rutherford backscattering and nuclear reaction analysis, use metallic ions to determine elemental composition and depth profiling of materials. These techniques may be in materials science and archaeology, among other fields.

[0138] Radiation Shielding: High-density metallic materials and alloys, including lead and tungsten, may be for shielding against energetic particles, such as X-rays and gamma rays, in medical and industrial applications to protect workers and equipment from radiation exposure.

[0139] Metallurgical Research: Energetic metallic particles may be used in metallurgical research to study phase transformations, crystallography, and mechanical properties of metals and alloys. This research informs the development of new materials and processes for various industries.

[0140] Nuclear Fusion: In the pursuit of controlled nuclear fusion as a future energy source, metallic particles including deuterium and tritium isotopes may be as fuel for fusion reactions. Energetic metallic ions are employed to heat and confine the plasma in experimental fusion reactors.

[0141] Surface Modification and Hardening: Metallic energetic particles may be used to modify the surface properties of metals, such as hardening, alloying, or introducing desirable surface features. These techniques find applications in aerospace, automotive, and manufacturing industries.

[0142] Metallurgy in Space: In space exploration, metallic particles are used in various ways, such as for spacecraft shielding against cosmic rays and micrometeoroids or for propulsion in electric ion thrusters.

[0143] While the above description provides examples of one or more apparatus, methods, or systems, it will be appreciated that other apparatus, methods, or systems may be within the scope of the claims as interpreted by one of skill in the art.