Systems and methods for producing energetic particles by additive manufacturing

Abstract

A system and method for producing polymer-free nanothermitic aerogel-like energetic particles by additive manufacturing is provided. An ink containing graphene oxide (GO), Al, and Bi2O3 nanoparticles in propylene carbonate is prepared. The method involves in situ reduction of graphene oxide by ethylenediamine, and printing the ink on a substrate during extrusion with a simple printing system. The printed aerogel contains reduced GO as a porous scaffold for the aerogel, within which Al and Bi2O3 clusters are embedded. The linear burning velocity of the printed aerogel reached a higher velocity (10 m / s) than reported for typical polymer-assisted 3D printed nanothermitic products. A framework is also provided for optimizing nanothermitic fuel particle structures to match desired burning profiles.

Description

[Technical field]

[0001] FIELD OF THE DISCLOSURE Embodiments disclosed herein relate to energetic particles, and in particular to systems and methods for additively manufacturing energetic particles, such as polymer-free nanothermitic aerogels. [Background technology]

[0002] Introduction Energetic particles such as thermite and / or metallic materials and / or fuels can be used for heating and burning as an energy source 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 different components: fuel and oxidizer. Reducible metals and metal salts such as aluminum (Al), magnesium (Mg) and silicon (Si) are usually chosen as the fuel component due to their large enthalpy change upon combustion. On the other hand, the oxidizer is generally a metal oxide, halide or oxyanion salt. Al-based thermite is the most commonly used thermite in this system due to its abundance, ease of manufacture and very high theoretical enthalpy of combustion of 31 kJ / g. Since considerable heat is released during the redox reaction between the fuel, oxidizer and intermediates, it is used in many applications such as propellants, pyrotechnics and welding.

[0003] Nanothermites, also known as metastable intermolecular composites, consist of nanoscale fuel or oxidizer components, or both. Nanothermites have a large surface contact between fuel and oxidizer particles, and their large surface area to volume ratio amplifies the role of the reactive interface, resulting in a significant increase in the burning rate. Intensive research efforts have been made to develop nanothermites with various compositions and micro / meso structures to optimize the burning rate and the rate of pressurization of the combustion chamber. However, the real-world application of nanothermitic materials is still very limited, as they suffer from several significant drawbacks, including the difficulty in controlling the reactivity, the degree of combustion completion, and issues related to safe manufacture, handling, and transportation. The lack of controllability of the reactivity is caused by multiscale physics, including phase separation during synthesis, reactive sintering of Al nanoparticles during combustion, and the formation and growth of inert alumina shells on the surfaces of Al nanoparticles. In addition, nanothermites exhibit extreme sensitivity to external stimuli, such as electrostatic discharge, friction, and mechanical shock.

[0004] Graphene and functionalized graphene have been recently utilized as additives in nanothermitic composites. Although the role of graphene in the thermite reaction is still controversial, the exfoliated 2D sheets provide a surface area for the assembly of Al and metal oxide nanoparticles, and conductive graphene and reduced graphene oxide (rGO) may reduce the risk of accidental ignition due to electrostatic discharge.

[0005] Additive manufacturing has become a key technological innovation in many industries in recent decades due to its versatility in prototyping, rapid tooling, and instant manufacturing. As one of the most important methods to facilitate additive manufacturing, extrusion-based printing, or direct ink writing, has proven promising in fabricating multiple component structures and developing controlled mechanical and physical properties. It is also considered as an important tool to develop novel and flexible thermite-based architect reactive interfaces and structures. Although polymers are commonly used as additives in additive manufacturing and nanothermites printing to obtain good “inks” with desirable viscosity, distribution, and structural integrity after drying, the burn rates of the printed solid products are usually relatively low.

[0006] Shen et al. used a polymer combination of hydroxypropylmethylcellulose, nitrocellulose, and polystyrene at a total of 10 wt% to print Al-CuO nanothermites with a maximum linear burning velocity of 25 cm / s, a common parameter that represents the burning rate of nanothermites (Non-Patent Document 1). Mao et al. used fluororubber F2311 at 5-25 wt% to obtain Al-CuO printed nanothermites that reached a maximum linear burning velocity of 1.5 m / s (Non-Patent Document 2).

[0007] Additive manufacturing (e.g., 3D printing) is essential for safe and customizable production in future applications of nanothermitic materials. Currently, polymers are typically used to tune the rheology of the ink and support the final structure of the printed nanothermitic product, which significantly limits the energy performance of the printed material. To increase the burning rate of printed nanothermites to over 10 m / s, specific designs, such as hollow structures, need to be considered and specially fabricated. Therefore, novel additive manufacturing systems and methods for polymer-free nanothermitic aerogel fuel particles are needed.

[0008] The conceptual design phase of a rocket propulsion system aims to match mission requirements to preliminary engine design considerations, within the specified constraints of the problem. Parametric design space is particularly important for solid and hybrid rocket engines, where propellant properties, fuel grain structure, and combustion chamber geometry can be uniquely varied to suit a specific mission. Furthermore, the thermodynamic properties of the engine change over time, as combustion chamber volume increases and normalized surface area changes as combustion progresses and fuel retreats. The high dimensionality of this problem has motivated the development of a framework to match mission profiles to propulsion system design.

[0009] Depending on the conceptual design focus, some frameworks propose a comprehensive consideration for the optimization of propulsion systems for ascending trajectories, including the consideration of nonlinear aerodynamics (Non-Patent Document 3). Other works aim to optimize the shape of the fuel grain to meet the desired thrust profile (Non-Patent Document 4), feed system and structural modeling (Non-Patent Document 5), or perform performance-matching optimization (Non-Patent Document 6). In many of these works, propellant properties such as burn rate are not considered as independent variables, given the strong interdependence of propellant regression rate, heat release, and erosion characteristics.

[0010] Existing methodologies often rely on variational optimization approaches to determine the optimal geometric parameters of the fuel particles. These optimizations are constrained by the physics of the problem while seeking to minimize the overall mass and / or total cost. However, for complex fuel particles, it is often manufacturability that imposes the greatest constraint on the optimization problem.

[0011] Existing studies have applied optimization techniques to cases with more complex geometries and further explored them (Non-Patent Document 7). Methodologies such as level set-based burnback analysis use a level set approach to follow the changes in the topology of a burning SRM (Non-Patent Document 8). Most of these optimization techniques rely on continuous and bounded ranges for each parameter to be optimized. These methods also require the designer to properly assign boundaries to the problem so that the program reaches the optimal solution in a timely manner. More recent studies have proposed machine learning approaches (Non-Patent Document 9) or optimization of two-component propellant grains (Non-Patent Document 10) to benefit from the best utilization of fuel grains in solid rocket motors. Ultimately, these constraints limit the ability to optimize more complex thrust profiles.

[0012] In recent years, a new propulsion paradigm is taking hold, which is bringing about dramatic changes in solid and hybrid propulsion technologies. Recent studies have proposed the integration of pyrophoric additives based on metal-organic frameworks (MOFs) (Non-Patent Document 11). At the same time, new manufacturing processes for solid fuels by additive manufacturing (AM) techniques are opening up new design opportunities for a new class of solid and hybrid state propulsion systems.

[0013] Historically, the selection of fuel grain structure was the best-fit approach to achieve a desired thrust-time curve. For a known propellant burn rate and regression rate, imposed fuel grain structures cause a change in area with time, and the associated release of heat of combustion. For example, a tubular grain design (see Figure 6A, left) results in an advancing (increasing) thrust curve because the burn area increases as the grains retreat toward the outer wall. A star-shaped fuel grain (see Figure 6A, right) is required for a neutral (constant) thrust curve. Designing such grain structures requires the use of burn-back models. Furthermore, these complex structures are subject to significant erosive burn, structural integrity issues, and manufacturing constraints.

[0014] As discussed above, the new opportunities afforded by AM of energetic fuels mean that design considerations for solid and hybrid engines can shift away from complex geometric fuel particles to modify thrust-time curves and towards functionally grading single-fuel or multi-fuel propellant engines (see Figure 6B). By layering various fuels with spatially varying binder and propellant composition, it is possible to effectively build matching thrust-time profiles without the need for complex fuel particles, thereby opening up new opportunities for novel engine design and optimization studies. [Prior art documents] [Non-patent literature]

[0015] [Non-Patent Document 1] Shen, H. et al., “Combustion of 3D printed 90 wt% loading reinforced nanothermite,” Combust. Flame, 215 (2020) 86-92. [Non-Patent Document 2] Mao, L., et al., “3D Printing of Micro-Architected Al / CuO-Based Nanothermite for Enhanced Combustion Performance,” Adv. Eng. Mater., 21 (2019) 1900825. [Non-Patent Document 3] Federici, L., et al., “Integrated optimization of first stage SRM and ascent trajectory of multistage launch vehicles,” Journal of Spacecraft and Rockets, vol. 58, no. 3, pp. 786-797, 2021. [Non-Patent Document 4] Oh, SH, 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. [Non-Patent Document 5] Adami, A., et al., “A new approach in multidisciplinary design optimization of upper-stages using combined framework,” Acta Astronautica, vol. 114, pp. 174-183, 2015. [Non-Patent Document 6] Zeping, W., et al., “Solid rocket motor design employing 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. [Non-Patent Document 7] Johannsson, M., “Optimization of Solid Rocket Grain Geometries,” 2012. [Non-Patent Document 8] Wang, DH, 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. [Non-Patent Document 9] Oh, SH, et al., “New design method of solid propellant grain using machine learning,” Processes, vol. 9, no. 6, 2021. [Non-Patent Document 10] 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. [Non-Patent Document 11] Jobin, O., et al., “Metal-organic frameworks as hypergolic additives for hybrid rockets,” Chemical Science, vol. 13, no. 12, pp. 3424-3436, 2022. Summary of the Invention

[0016] We describe a method and additive manufacturing system for producing energetic particles by additive manufacturing, including 3D printing of polymer-free nanothermitic aerogels. Inks were prepared by dispersing graphene oxide, Al, and Bi2O3 nanoparticles in propylene carbonate. Graphene oxide comprised 5% to 20% (by mass) of the ink, and the nanoparticle loading was varied between 80% and 95%. Graphene oxide was reduced in situ by ethylenediamine during the printing process. The in situ reduction and gelation of graphene oxide (GO) to reduced graphene oxide (rGO) by ethylenediamine allowed nanothermitic aerogels with flexible nanoparticle loading to be easily 3D printed at room temperature using a relatively simple printing system with a syringe pump and linear actuator.

[0017] The printed rGO / Al / Bi2O3 aerogel contained a porous structure consisting of interconnected rGO scaffolds and wrapped Al / Bi2O3 nanothermitic clusters. The nanoparticles showed no signs of phase separation and were highly homogeneous. The densely packed fuel and oxidizer nanoparticles significantly reduced the reaction temperature and promoted the occurrence of condensed-phase reactions. With nanoparticle loadings above 90%, the material exhibited linear burning velocities up to 10 m / s, much higher than existing polymer-assisted 3D printed nanothermitic composites.

[0018] According to some embodiments, there is a method for 3D printing a nanothermitic aerogel, which includes mixing a printable ink including nanothermitic and graphene oxide with ethylenediamine in situ in an extrusion tube to form a gel, extruding the gel onto a substrate, immersing the substrate and the gel thereon in stirred tert-butanol, and freeze-drying the gel to form a nanothermitic aerogel.

[0019] The method may include preparing a printable ink by separately dispersing each of graphene oxide, fuel nanoparticles, and oxidizer nanoparticles in propylene carbonate under sonication, mixing the fuel nanoparticle dispersion and the oxidizer nanoparticle dispersion with the graphene oxide dispersion in various concentrations and combinations, and allowing the resulting mixture to stand.

[0020] According to one embodiment, there is an additive manufacturing printing system. The system includes a first syringe containing a printable ink including graphene oxide and a second syringe containing an additive for reducing the graphene oxide. An adjustable syringe pump drives the first and second syringes. An extrusion tube having an outlet is connected to the first tube for extruding the ink onto the substrate. A needle is connected to the second syringe for injecting the reducing additive into the extrusion tube at a location between the first syringe and the outlet. The printing system includes a stage for mounting the substrate. The stage is movable in a horizontal plane by two linear actuators. In some examples, a third actuator may be added to allow movement in a vertical axis. In other embodiments, butanediamine may be used to reduce the graphene oxide.

[0021] According to some embodiments, there is a nanothermitic aerogel comprising a porous cross-linked scaffold of reduced graphene oxide and a plurality of nanothermitic clusters embedded in the porous scaffold. The nanothermitic clusters include fuel nanoparticles and oxidizer nanoparticles. The nanothermitic aerogel can be formed into an energetic material for combustion for power generation, propulsion, and / or construction applications.

[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. [Brief description of the drawings]

[0023] The drawings included herein are intended to illustrate various examples of the articles, methods, and apparatus herein. [Figure 1A] FIG. 1A is a flow diagram of a method for direct ink printing of polymer-free nanothermitic aerogels according to one embodiment. [Figure 1B] FIG. 1B is a diagram of an additive manufacturing system according to one embodiment. [Figure 2A] FIG. 2A is an image of nanothermitic ink before printing according to one embodiment. [Figure 2B] FIG. 2B is an image of a nanothermitic gel just after printing according to one embodiment. [Figure 2C] FIG. 2C is an image of the nanothermitic gel of FIG. 2B during solvent exchange. [Figure 2D] Figure 2D is an image of the nanothermitic gel of Figures 2B-2C after freeze-drying. [Figure 2E] FIG. 2E is an X-ray diffraction spectrum of the dried nanothermitic gel shown in FIG. 2D. [Figure 3A-3C] 3A-3C are scanning electron microscope images of printed rGO / Al / Bi2O3 aerogels according to one embodiment. [Fig. 3D-3F] 3D-3F are energy dispersive spectroscopy images mapping aluminum, bismuth, and carbon, respectively, in the image shown in FIG. 3B. [Figure 4A] FIG. 4A is a differential scanning calorimetry curve of various nanothermitic aerogels during a slow heating process according to some embodiments. [Figure 4B] FIG. 4B is a thermogravimetric analysis of nanothermitic aerogel during a slow heating process according to some embodiments. [Figure 5A-5C] 5A-5C are video frames showing the combustion of nanothermit aerogel, according to some embodiments. [Figure 6A] FIG. 6A is an illustration of a classic fuel particle shape. [Figure 6B] FIG. 6B is a diagram of functionally graded fuel particle shapes. [Figure 7A] FIG. 7A is a (simple) thrust profile for retreat. [Figure 7B] FIG. 7B is a two-peak (complex) thrust profile. [Figure 7C] FIG. 7C is a pressure versus altitude curve for optimum thrust during climb. [Figure 8A] FIG. 8A is an idealized burn rate profile for simple thrust according to one embodiment. [Figure 8B] FIG. 8B is a simple optimized thrust profile based on the ideal velocities shown in FIG. 8A. [Figure 9A] FIG. 9A is an ideal burn rate profile for complex thrust according to one embodiment. [Figure 9B] FIG. 9B is an optimized complex thrust profile based on the ideal velocities shown in FIG. 9A. [Figure 10A] FIG. 10A is an ideal burn rate profile for pressure matching according to one embodiment. [Figure 10B] FIG. 10B is an optimized pressure profile for pressure matching according to one embodiment. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0024] Detailed Description Various devices or processes are described below to provide examples of each claimed embodiment. The embodiments described below are not intended to limit the claimed embodiments, which may be directed to processes or devices different from those described below. The claimed embodiments are not limited to devices or processes having all of the features of any one device or process described below, or to features common to some or all of the devices described below.

[0025] 1A, shown therein is a flow diagram of a method 10 for direct ink printing of nanothermitic aerogels according to one embodiment. Advantageously, this method 10 does not use a polymer in the printable / extrudable ink, but rather reduces graphene oxide to form a 3D scaffold for the nanothermitic particles in the ink.

[0026] In 12, graphene oxide (GO) may be provided or synthesized according to a modified Hummers method as previously described (see, e.g., Thiruvengadathan, R., et al., “A Versatile Self-Assembly Approach toward High Performance Nanoenergetic Composite Using Functionalized Graphene,” Langmuir, 30 (2014) 6556-6564 and Wang, A., et al., “Reactive nanoenergetic graphene aerogel synthesized by one-step chemical reduction,” Combust. Flame, 196 (2018) 400-406).

[0027] The Hummers method is summarized as follows. One gram each of graphite (grade H-5) and sodium nitrate (NaNO3) was added to 46 mL of sulfuric acid (H2SO4, 95-98%) and stirred in an ice-water bath for 10 min. Then, 6 grams of potassium permanganate (KMnO4) was slowly added to the mixture. The water bath was then heated to 35 °C and kept for 1 h for the oxidation of graphene to occur. Then, 80 mL of deionized water was added dropwise and the water bath was heated to 90 °C and kept for 30 min. Then, the mixture was cooled to room temperature, after which 200 mL of deionized water and 6 mL of hydrogen peroxide (H2O2, 30 wt%) were added. After sonication and final centrifugation, the aqueous solution was dried at 65 °C overnight, finally yielding GO sheets.

[0028] At 14, a graphene nanothermitic (GO / Al / Bi2O3) ink may be prepared. To prepare the GO / Al / Bi2O3 ink, as-prepared GO (from step 12), fuel nanoparticles (e.g., Al nanoparticles, ∼100 nm diameter, 83% active), and oxidizer nanoparticles (e.g., Bi2O3, ∼120 nm diameter) are dispersed in propylene carbonate under ultrasonication. After 3 hours, the Al and Bi2O3 dispersions are mixed and ultrasonicated for another hour before mixing with the GO dispersion. The GO / Al / Bi2O3 mixture is then stirred for 5 minutes and allowed to sit overnight (at least 12 hours) before printing. The amounts of GO and fuel / oxidizer nanoparticles used in the different ink preparations according to various embodiments are shown in Table 1.

[0029] [Table 1]

[0030] It is noted in Table 1 that the mass and concentration of GO are kept constant. The final product rGO(x%) / Al / Bi2O3 represents a gel printed using a GO(x%) / Al / Bi2O3 ink, as reduction of GO occurs during the printing process. Percentages are calculated between the mass of GO and the total mass of GO and nanoparticles. 5-20% GO is preferred, but can be dropped to 2% or even lower. The equivalence ratio (ER) is calculated as follows:

number

[0031] Steps 12 and 14 can be performed prior to the remainder of method 10, and the GO and inks prepared by these steps can be stored for later use.

[0032] At 16, the GO / Al / Bi2O3 ink is mixed in situ with a reducing additive (e.g., ethylenediamine) to reduce GO to rGo at room temperature in an additive manufacturing system, e.g., the printing system 100 shown in FIG. 1B. The GO / Al / Bi2O3 ink and ethylenediamine are loaded into respective syringes 102, 104 driven by syringe pumps 106, 108. A 1.6 mm (1 / 16 inch) diameter polyvinyl chloride (PVC) extrusion tube 110 is connected to the syringe 102 of GO / Al / Bi2O3 ink and a needle 112 is connected to the syringe 104 of ethylenediamine to inject ethylenediamine into the ink in the extrusion tube 110 to initiate the gelation reaction.

[0033] Proper flow rate and reaction time in the extrusion tube 110 are essential for successful printing. The movement speed of the syringe pumps 106, 108 is adjustable. Preferably, the movement speed of the syringe pumps 106, 108 is set to a volume ratio of GO / Al / Bi2O3 ink to ethylenediamine of 20:1, and the bulk velocity of the material in the tube 110 is set to a maximum of 40 mm / s, such that the total reaction time between the location 114 where the ethylenediamine is injected into the tube 110 and the tube exit 116 is 6 seconds. No extra equipment was utilized to assist the flow of liquid in the tube or the extrusion of the gel. According to some embodiments, the ratio of ink to reducing additive may be greater than 20:1 (e.g., 30:1). In general, the amount of reducing additive added should be as small as possible to ensure complete reduction of Go to rGO without substantially diluting the ink.

[0034] At 18, the rGO / Al / Bi2O3 gel is extruded onto a substrate 118 at room temperature. Gelation and in situ reduction of GO is induced by injecting ethylenediamine into the extrusion tube 110. That is, steps 16 and 18 are generally performed simultaneously. According to one embodiment, the gel is printed onto a substrate 118, which is a 6.1 cm x 6.1 cm acrylic plate.

[0035] At 20, the substrate 118 is moved in the xy plane while the gel is extruded onto the substrate 118. The substrate 118 is attached to a stage 120 controlled by xy linear actuators 122, 124. The stage 120 is movable in the xy (horizontal) plane so that the gel can be printed in a curvilinear shape onto the substrate 118 (see FIGS. 2B-2D). A microcontroller 126 (e.g., of a control device) provides input signals to the linear actuators 122, 124 to move the stage 120 in the xy plane. The microcontroller 126 is also connected to the syringe pumps 106, 108 and adjusts the movement speed of the syringe pumps 106, 108, thereby controlling the flow rates of the ink and ethylenediamine, respectively.

[0036] After printing, the substrate 118 and the gel thereon are placed in a petri dish and immersed in stirred alcohol, preferably tert-butanol (99%), at 22. The gel is left for 3 days at 24 to allow the propylene carbonate and ethylenediamine to be exchanged with the alcohol, which is refreshed every 12 hours.

[0037] In 26, after the solvent exchange step, the gel is freeze-dried to obtain the rGO / Al / Bi2O3 aerogels.

[0038] At 28, the rGO / Al / Bi2O3 aerogel may be cut into pellets. According to other embodiments, the aerogel may be directly formed as an energetic material such as pellets by printing the pellets onto the substrate 118 in steps 18 and 20. In other embodiments, the energetic material may be printed into other desired shapes using a stage-mounted mold to mold the energetic material.

[0039] The nanothermit aerogels produced by method 10 can be used as a fuel source for propulsion or energy generation. Such fuel sources are particularly advantageous when used as propellants and combusted in rocket engines for propulsion. Additionally, such fuel sources can be particularly advantageous in the energy generation and / or combustion of these materials that act as heat sources for thermoelectric generation, as well as for thermophotovoltaic systems to convert heat to electricity. In other embodiments, these fuel sources can be used for combined cooling, heat and power applications.

[0040] Referring to Figures 2A-2D, optical images of nanothermite (rGO / Al / Bi2O3) ink / gel at different stages during the printing process are shown. The GO sheets, Al, and Bi2O3 nanoparticles all had negative surface charges in the propylene carbonate dispersion. As a result, after mixing and standing overnight, the GO / Al / Bi2O3 ink showed no precipitation and a uniform dispersion was obtained (Figure 2A). The as-printed rGO / Al / Bi2O3 gel with propylene carbonate (Figure 2B) went through a solvent exchange process (Figure 2C) and then freeze-dried to obtain the final dried rGO / Al / Bi2O3 aerogel (Figure 2D). After freeze-drying, the rGO / Al / Bi2O3 aerogel retained its shape.

[0041] Figure 2E shows the X-ray diffraction (XRD) spectrum of the dried aerogel shown in Figure 2D, along with standard spectra for Al and Bi2O3. XRD spectra were performed using a PANalytical X'pert Pro MRD in Grazing Incidence X-Ray Diffraction (GI-XRD) mode in the range of 10 to 80° with a step of 0.1°. Both Al and Bi2O3 were found in the final aerogel. However, the XRD peak of rGO at about 24° is indistinguishable in the XRD curve. This is partly due to the relatively low percentage of GO, but also due to the unique thin sheet structure of rGO sheets in the aerogel, as confirmed by SEM images.

[0042] Referring to Figures 3A-3C, scanning electron microscope (SEM) images of the printed rGO / Al / Bi2O3 aerogels are shown at different magnifications. The gel microstructure was observed using a Hitachi Su-5000 SEM and an Oxford EDS (Energy Dispersive Spectroscopy). Figure 3A is an overall view of the printed rGO / Al / Bi2O3 aerogel, showing a porous structure with nanoparticles embedded throughout the structure. Zooming in on Figure 3B reveals that the pore size is on the order of micrometers. The Al and Bi2O3 nanoparticles were unable to form and hold a free-standing three-dimensional structure by themselves, and it was the rGO sheets that constituted the framework of the three-dimensional aerogel structure. These are shown as translucent cloth-like sheets in Figures 3B-3C. Such an extremely thin rGO layer was formed because the GO was well dispersed in a very thin layer in the propylene carbonate in the ink before printing. After in-line mixing with ethylenediamine, the reduction and interconnection of GO occurred, forming the porous framework of the aerogel. Both the Al and Bi2O3 nanoparticles were covered and wrapped by the thin rGO sheets, forming small clusters of nanothermites with a size of a few microns (circles in Figure 3B).

[0043] Referring to Figures 3D-3F, the elemental mapping of aluminum, bismuth, and carbon in the image shown in Figure 3B are shown, respectively. Comparing Figures 3D and 3E, it can be seen that the two different types of Al and Bi2O3 nanoparticles are homogeneously distributed without phase separation. Unlike aluminum and bismuth, which are mostly present inside the clusters of assembled nanoparticles, rGO (carbon, Figure 3F) is uniformly distributed throughout the aerogel and played the most important role as the supporting structure / scaffold for the Al and Bi2O3 nanoparticles.

[0044] 4A-4B and Table 2, shown therein are Differential Scanning Calorimetry (DSC) and Thermogravimetric Analysis (TGA) showing the heat exchange and mass change, respectively, of various nanothermitic aerogels during the slow heating process, according to some embodiments. DSC-TGA was analyzed on a Netzsch STA 449 F1 from room temperature to 1000° C. under the protection of argon flow at 40 mL / min, with a heating rate of 20° C. / min. [Table 2]

[0045] The exothermic peaks at 160–300 °C and the subsequent endothermic peaks shown in the DSC curves in Figure 4A are attributed to the decomposition of oxygen-containing functional groups remaining in rGO, which corresponds to the weight loss at the same temperatures shown in Figure 4B. The main reaction between Al and Bi2O3 had an onset temperature near 515 °C and a peak temperature near 555 °C, which is much lower than the peak temperature of Al / Bi2O3 loose powder at 600 °C, due to the close proximity of the fuel (Al) and oxidizer (Bi2O3) nanoparticles in each cluster, as seen in the SEM images (Figures 3B–3C).

[0046] The exothermic peak after Al melting was caused by the reaction of Al with Bi2O3 after Al melted and flowed across the rGO structure. This decrease in onset temperature indicates improved reactivity of the printed nanothermitic aerogels. The formation of Al / Bi2O3 clusters in the 3D printing process is expected to promote agglomeration of reactive fuel and oxidizer nanoparticles, enhancing reactivity. The agglomerates as "unit cells" are ignited and react locally, with sintering of the fuel and oxidizer nanoparticles within their volume.

[0047] Referring to Figure 4B, a rapid weight loss occurred at almost the same temperature as the main reaction, which is probably due to the reaction of carbon in rGO with Bi2O3. The weight loss increased for aerogels with a higher percentage of rGO. It is noteworthy that the energy release numbers for rGO(5%) / Al / Bi2O3 and rGO(10%) / Al / Bi2O3 are very close, which resembles the difference in the percentage of nanothermite particles in the products. The shapes of the DSC curves (Figure 4A) for these two aerogels were also very similar, with the condensed phase reaction before Al melting dominating and the exothermic peak after melting being smaller. However, the energy release of the condensed phase reaction for rGO(20%) / Al / Bi2O3 was smaller. This decrease was significantly larger than the change in the nanoparticle percentage in the material, indicating some differences in the structure and reaction at relatively low nanoparticle loadings.

[0048] As seen in the SEM images (Figures 3A-3F), the rGO / Al / Bi2O3 aerogels were composed of cross-linked rGO cages and encased Al / Bi2O3 nanoparticle clusters. Since the GO concentration in the ink and the ratio of ink to ethylenediamine were constant among the imaged aerogels, it is reasonable to assume that the degree of cross-linking and the amount of clusters were close among aerogels with different nanoparticle loadings. Thus, aerogels with low nanoparticle loadings not only had a larger dead mass of rGO, but more importantly, smaller Al / Bi2O3 aggregates, which could lead to less interfacial contact with the fuel and oxidizer, resulting in a much reduced notable condensed phase reaction in the DSC. The total energy release of the aerogels was also lower compared to the Al / Bi2O3 loose powder (~700 J / g) as a result of the isolated nanoparticle clusters.

[0049] To further understand how the percentage of rGO changes the energetics, flame propagation speeds were measured using aerogels with 5–20% rGO. The density of the aerogels was determined by the GO concentration in the ink (10 mg / cm). 3 ) divided by the GO percentage (5-20%), giving 50-200 mg / cm 3 was estimated to be approximately 1-3% of the theoretical mass density (TMD). Combustion of the aerogel was induced by a nickel-chromium wire (~10 Ω, 4 cm) connected to a 30 V DC power supply. Combustion videos (512 × 320 resolution) were captured using a Phantom v2012 monochrome high-speed camera at 100,000 frames per second with an exposure of 5 μs and an Extended Dynamic Range (EDR) of 1 μs (20%). The camera and DC power supply were synchronized using a TTL signal output from a Tektronix AFG1022 arbitrary function generator.

[0050] 5A-5C, high-speed video frames capturing the combustion of nanothermitic aerogels according to some embodiments are shown. The combustion propagated from right to left for all aerogels. Similar to the DSC results (FIGS. 4A-4B, Table 2), the rGO(5%) / Al / Bi2O3 aerogel (FIG. 5A) and rGO(10%) / Al / Bi2O3 aerogel (FIG. 5B) showed similar results, while the rGO(20%) / Al / Bi2O3 aerogel (FIG. 5C) showed different behavior.

[0051] When the nanoparticle loading was 90% or higher (Figures 5A and 5B), the reaction was quite vigorous. Combustion propagated at faster velocities (8-10 m / s) than typically observed in polymer-assisted 3D printed nanothermite products (<1.5 m / s), and hot flames spread into the surrounding environment due to the unique porous structure and presence of rGO. However, when the nanoparticle loading was reduced to 80% (Figure 5C), the energy release from the reaction was just enough to maintain self-sustaining combustion, with a flame propagation speed of only 39 cm / s. Residual rGO was visible.

[0052] The ignition delay was about 9 ms for both the 90% and 95% nanoparticle-filled aerogels. Therefore, the temperature of the nickel wire can be estimated by the following equation:

number

[0053] This is consistent with the reaction temperatures obtained from the DSC results. However, for the aerogels filled with 80% nanoparticles (Figure 5C), the ignition delay time increased to about 26 ms, and the nickel-chromium wire became visibly hot before the aerogel ignited. It is expected that the large amount of rGO in the printed aerogel not only prevented the formation of larger Al / Bi2O3 clusters, but also transferred heat throughout the aerogel, delaying the onset of the energetic reaction.

[0054] Combining the TGA / DSC results and flame speed measurements, it can be concluded that rGO in the printed nanothermitic aerogels played multiple roles in the structure, promoting and simultaneously reducing the reaction. On the one hand, rGO constituted a three-dimensional framework for the nanoparticles to assemble, forming a unique porous structure, which helped to promote the reaction. In contrast, in the polymer-assisted 3D printed nanothermitic materials, the polymer in the structure forms a continuous phase, and the Al and metal oxide nanoparticles decorate the interior of the "polymer sea", thus greatly reducing the direct contact between the fuel and the oxidizer. Although some of the polymers used in the 3D printed nanothermitic materials are considered as oxidizers (e.g., fluoroplastics) and explosives (e.g., nitrocellulose), the slow decomposition step of the polymer is the rate-determining step for the combustion.

[0055] However, in the 3D printed rGO / Al / Bi2O3 aerogels, no polymer is present, and more importantly, the Al and Bi2O3 nanoparticles are tightly packed into clusters by the rGO sheets. As mentioned above, the sintering mechanism supports the local ignition and combustion of aggregates in the nanothermitic aerogels, and the formation of clusters consisting of both fuel and oxidizer nanoparticles during printing is expected to facilitate the thermite reaction. Meanwhile, the significantly improved surface contact of Al and Bi2O3 leads to a very fast condensed-phase reaction. Moreover, the presence of micron-sized pores inside the structure allows the hot combustion products to better propagate inside the structure, resulting in the localized formation of hot spots within the scaffold.

[0056] High thermal conductivity rGO also became an alternative for heat transfer in low TMD% nanothermitic aerogels. On the other hand, when the nanoparticle loading was too low (higher percentage of rGO), the mass-wise combustion enthalpy from the thermite reaction was smaller. The added rGO effectively reduced the intimacy between the Al and Bi2O3 nanoparticles, decreasing their reactivity. The highly thermally conductive rGO scaffold further increased the heat loss and hindered the propagation of combustion. Therefore, to guarantee the promotion effect of rGO in the final aerogel, the nanoparticle loading must be higher than a threshold value. This value depends on multiple parameters, including the topology of the printed aerogel, the properties of the nanothermitic material, and the amount of ethylenediamine used during the reaction. The results shown in Figures 4-5 suggest that increasing the particle loading to 95% or more could result in even faster and more energetic combustion, but in this case, the effect of rGO in localized cluster formation is reduced.

[0057] The energetic materials and nanothermitic aerogels described herein include Al as the fuel and Bi2O3 as the oxidizer, although one skilled in the art would understand that other metal fuel and oxidizer combinations are possible. For example, the metal fuel can be Mg, Si, Fe, etc., and the oxidizer can be a fluoropolymer, iodine oxide, or metal oxide (e.g., Fe2O3, SiO2, MgO, etc.). The aerogels described herein can be formed into fuel particles for burning in solid rocket motors (SRMs). Below, a computational framework is described for optimizing the fuel particle structure to match a desired thrust curve profile.

[0058] The framework includes two solvers with different fidelity levels to efficiently optimize a wide parameter space. First, a system-level code (zero-dimensional) is developed to evaluate the behavior of the entire system. Second, a quasi-one-dimensional code is developed to incorporate the spatial variations and acoustic modes of the combustion chamber and nozzle of a functionally upgraded engine. As the complex burning kinetics of such solid fuels are not yet fully understood, simplified combustion and regression models are used.

[0059] System Level Solvers First, a system level solver using isotropic nozzle relations is developed to estimate the engine vacuum thrust characteristics. The total thrust of the engine can be estimated based on the total pressure and temperature experienced in the combustor for a given nozzle geometry. The vacuum thrust equation is replaced by:

number

[0060] The corresponding specific impulse Isp is:

number

[0061] The 0D model uses isotropic flow equations to relate the combustion chamber conditions (total pressure and temperature) to the nozzle exit conditions. In a supersonic nozzle, the mass flow rate is fixed for a given geometry and engine thermodynamic conditions. This allows the area ratio (exit and throat) to be related to the exit Mach number of the system.

number

[0062] Known nozzle shape (A e / At ), knowing the thermodynamic state of the engine, the exit temperature and pressure can be calculated.

number

[0063] Using these exit conditions, the exit velocity can be calculated by knowing the exit sound speed.

number

[0064] In the combustion chamber, the thermodynamic state is directly linked to the combustion kinetics and regression of the fuel particles. The mass flow rate of the gases produced by combustion is calculated as follows:

number

[0065] The pressure in the combustion chamber is:

number

[0066] This simplified combustion model assumes a constant burning rate r for a given fuel. b and solid fuel density ρ propMore generally, Saint Robert's law can be used to explain the pressure dependence of the linear burning rate:

number

number

[0067] The change in mass of the rocket is m calculated above. out Related to.

[0068] Quasi-One-Dimensional Solver A quasi-one-dimensional solver was developed in parallel to account for spatial variations in the combustion chamber and nozzle and to investigate acoustic coupling in functionally graded engines. The quasi-one-dimensional code solves the one-dimensional Navier-Stokes equations (conservation of mass, momentum and energy).

number

number

[0069] The viscous flux terms and viscous flux terms are defined as follows:

number

[10] represents the source term.

number

[0070] Because the area of ​​the combustion chamber is constantly changing as the fuel retreats, the second source term on the right hand side of equation

[10] accounts for the time and spatial changes in area. The source term is defined as:

number

[0071] The above equations are closed with the ideal gas equations, which is a reasonable assumption given the high temperatures of combustion, despite the high pressures in the combustion chamber.

[0072] The governing equations are implemented in StanShock, a quasi-one-dimensional framework. Spatial fluxes are calculated with a fifth-order WENO scheme, and the equations are time integrated with Strang decomposition for robustness. Cantera, an open-source chemical kinetics solver, is used to calculate the thermodynamics of the system. As with the system-level framework, a constant linear burning rate is used to characterize each propellant.

[0073] For a given combustion chamber and nozzle geometry, the equations can be advanced and the thrust generated calculated. The thrust of the rocket can be calculated directly by assuming perfect expansion in the nozzle as follows: F th =mu e +(P e -P atm )×A e As the equation is integrated over time and the fuel is retracted, the thrust profile curve can be estimated.

[0074] For the purposes of the present framework, a functionally graded solid rocket engine design is considered. The functional grade is achieved by stacking different fuels, as shown in Figure 1B. The objective of the present framework is to match the fuel grain structure to a given thrust or pressure profile depending on the test case, via an optimization solution.

[0075] Functionally Graded Engines There are several known parameters that affect the performance of SRM. The main parameters are fuel heat release, gas release, and 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 energetic materials produced by additive manufacturing, several aspects of the physical composition can affect the burn rate, such as propellant loading, extra polymer additives, and binder materials. Due to the variation of these parameters, the burn rate of solid fuels can range from the order of millimeters per second to hundreds of meters per second. Thus, by layering fuels with different combustion properties, the combustion properties during combustion can be changed to design a functionally graded engine. Thus, by optimal fuel layering, a mission-specific thrust profile can be achieved.

[0076] Test Case We provide a framework that can be used for future conceptual designs of functionally graded solid rocket engines. Three well-defined test cases are presented to illustrate how this framework can be used: a simple retract 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 adjusted to fully expand at the nozzle during ascent. The test cases are shown in Figures 7A-7C.

[0077] Test case 1 (Figure 7A) corresponds to a retraction profile that is typically achieved by using a complex particle shape such as a double anchor shape. Test case 2 (Figure 7B) is a more complex thrust profile with two peaks proposed by Federici et al. Finally, the third test case (Figure 7C) is a conceptual case in which the engine combustion characteristics are modified to match a nozzle flow that is fully expanded during ascent. The ideal pressure vs. altitude curve shown in Figure 7C is converted to a pressure vs. time curve in the fuel particle optimization. As previously mentioned, the ideal rocket equation relates the rocket speed to the rocket mass and the exit velocity of the burning gases. With this information, the rocket mass, speed, and height are calculated as the combustion occurs. From here, an ideal exit pressure (perfectly matching atmospheric pressure) curve as a function of time is created and optimized.

[0078] Optimization Framework An optimization framework is presented for the conceptual design of functionally graded solid rocket engines that fit the thrust-time, pressure-time profiles of a given mission. Considering the ability to create functionally graded engines, the fuel burn rate and layer thickness are selected as independent parameters of the optimization problem. For simplicity, a tubular grain (Figure 6A, left) is assumed. Due to the limited parameter space, a brute force profile matching optimization approach is implemented, but the modularity of the code allows for the implementation of constrained optimization algorithms, where constraints are applied on the burn characteristics.

[0079] The objective function is defined as the minimization of the integral L2 norm over the thrust profile, subject to the geometric and thermodynamic constraints described below. The parameters are the thickness and burn rate of each layer. The burn rate limit of the fuel layers is defined as:

number

[0080] The optimization process begins by stepping through a set of equations in the 0D code in time, determining at each time step the burn rate that most closely approximates the thrust or pressure profile to the sample profile. At each time step, the ideal burn rate and the radius at which it occurs are tracked, and a curve of ideal radial distance vs. burn rate is plotted.

[0081] Once the optimal burn rate at each radius value is known, the fuel region is divided into four approximate sections that act as layers in the fuel. The burn rate and thickness of each section are bounded after analyzing the ideal burn rate curve. To determine the burn rate and layer thickness boundaries, coarser velocity samples are used to create a new ideal velocity vs. burn rate curve. Within the minimum and maximum values ​​of the ideal curve, 4-5 burn rate samples are used. Although less accurate, this can be used to determine the layer properties of the fuel. As the velocity varies between the coarser values, bins are created where the ideal velocity resides. The range of velocity inputs for optimization is bounded by the minimum and maximum values ​​of the bins. The radius input for each layer is bounded by the radial location where the bin value changes by ±10% without overlapping with the boundary of an adjacent layer. The layer properties are optimized within these defined ranges. The size and burn rate of these layers are optimized to best suit the desired thrust or pressure profile.

[0082] Optimization of each layer thickness and burn rate is completed using a heuristic approach that uses random input ranges of layer thickness and burn rate to determine the most accurate solution. The model saves the layer thickness and burn rate of the model with the least error. The error between the desired profile and the profile generated by the model is calculated as the sum of the absolute values ​​of the differences between the two models over all time steps. Finally, the optimized layered solution is passed to a one-dimensional code to evaluate the spatial variation and acoustics of the engine.

[0083] Details of the rocket shape and the fuel used in the model are shown below. Initial radius of fuel, r i =0.2m Final radius of fuel, r f ≦0.6m Length of combustion chamber, L=5.0m ·Exit area, A e =0.0314m 2 Throat area, A t =0.00314m 2 - density of the propellant, ρin =900kg / m 3 Combustion adiabatic temperature, T comb =2900K Initial mass of the system (fuel and rocket), m i =900kg

[0084] Test Case 1: Simple Thrust Case Referring to Figure 8A, an ideal burn rate profile for a simple thrust is shown, plotting the ideal burn rate at each radial distance based on time steps. Figure 8B shows the optimized solution based on a heuristic method of selecting approximate random burn rates and layer thicknesses based on the ideal rates shown in Figure 8A. It can be seen that there are four distinct layers of fuel before the combustion chamber depressurizes at the end of launch. The optimal layer thicknesses, outer radii, and burn rates for the simple case are shown in Table 3 below.

[0085] [Table 3]

[0086] In the simple case, we can see that the fuel layers have different thicknesses as the fuel burn progresses: at the initial thrust peak at the start of the orbit, the first layer occupies 43% of the total fuel radius, while the second, third and fourth layers occupy 19%, 16% and 22%, respectively.

[0087] Test Case 2: Complex Thrust Case Referring to FIG. 9A, the ideal burn rate profile for the complex thrust is shown plotting the ideal burn rate at each radial distance based on the time step. FIG. 9B shows the optimized solution based on a heuristic method of selecting approximated random burn rates and layer thicknesses based on the ideal rates shown in FIG. 9A. It should be noted that the range of input values ​​is different from the simple cases (FIG. 8A and FIG. 8B) since it is based on the optimal burn rate at each radius for both models. For this reason, a wider range of inputs was used in the second test case when setting up the optimization trial. It can be seen that the thrust profile of the modeled SRM closely matches the desired thrust profile of the rocket when using four layers of simple fuel with different burn rates. The optimal radii and burn rates for the complex case are shown in Table 4 below.

[0088] [Table 4]

[0089] In the complex case, we can see that the fuel layer thickness varies as the fuel burn progresses. At the first thrust peak at the beginning of the orbit, the fuel thickness accounts for 46% of the fuel thickness. Meanwhile, the fuel thicknesses of the second, third and fourth layers are 14%, 31% and 9%, respectively. This is a larger variation in layer thickness than the simple case. We can also see that the burn rate variation is larger in the complex case.

[0090] Test Case 3: Pressure Matching Case In the third test case, the goal is to match the exhaust pressure of the combustion gases to atmospheric pressure. Referring to Figure 10A, the ideal burn rate profile for pressure matching is shown plotting the ideal burn rate at each radial location. Figure 10B shows the optimal pressure profile over time using four different layers. 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 height of the rocket over time based on the mass and gas exit velocity. In this case, it can be seen that the rocket exit pressure follows the ideal case test quite well. The optimal burn rates and layer thicknesses are shown in Table 5 below.

[0091] [Table 5]

[0092] In all three test cases, using more layers would produce a closer match to the model. In the future, if fuels were functionally graded, it may be possible to control the burn rate across the fuel to create a more accurate SRM. This would result in an optimized profile similar to the ideal profile modeled here. With additive manufacturing processes (e.g., 3D printing), the extrusion width of the material could be used as the minimum layer thickness, resulting in a more tunable profile.

[0093] In some examples, heterogeneous layers of energetic materials (nano- and / or micro-thermites) may be utilized in various concentrations to create tunable profiles.

[0094] In some embodiments, novel flexible thermite-based architected reactive interfaces and structures can be utilized to create tunable two- and / or three-dimensional structures.

[0095] In some examples, magnetohydrodynamics using magnets and / or electromagnets can be used to mix and / or stir the materials.

[0096] In some examples, thermite-based architected reactive interfaces and structures may be utilized for heating applications, while in other examples, interfaces and surfaces may include phase change materials for heating applications and power generation use cases.

[0097] In some examples, thermite-based reactive interfaces and structures can be heated using wireless power transfer. In other examples, the wireless power transfer can use radiative (e.g., electromagnetic frequencies such as microwaves and / or lasers) and / or non-radiative (inductive and / or magnetic coupling) methods.

[0098] In some examples, fuels and oxides can be sourced from Earth. In other examples, sources can include recycling of space debris, reuse of satellites in orbit, or other materials transported from Earth to space. In other examples, metal powders can be sourced from space. Sources can also include in situ resource utilization, such as materials from the Moon (lunar regolith), Mars (Martian regolith), and / or asteroid sources.

[0099] In some examples, the metal powder may be sourced from the waste output of industrial processes, or the products of other combustion processes, or waste disposal from human or robotic activity.

[0100] In some instances, a spacecraft may transport multiple additive manufacturing systems for operations, logistics, maintenance, point-to-point transportation, or orbital raising.

[0101] In some examples, thermite-based interfaces and structures may be embedded in satellites and / or space systems to aid in space debris removal and / or remediation efforts.

[0102] In other instances, larger energetic particles or microthermites can be introduced into the bed to tailor the combustion profile.

[0103] Although the above description provides examples of one or more devices, methods, or systems, it will be understood that other devices, methods, or systems may be included within the scope of the claims, as interpreted by one of ordinary skill in the art.

Claims

1. 1. A method for producing energetic particles by additive manufacturing, comprising: mixing in situ a printable ink comprising energetic particles and graphene oxide with an additive for reducing said graphene oxide in an extrusion tube to form a gel; extruding the gel onto a substrate; immersing the substrate and the gel thereon in agitated alcohol; and freeze-drying the gel to form an aerogel. method.

2. further comprising moving the substrate in a horizontal plane while extruding the gel onto the substrate. The method of claim 1.

3. further comprising varying the type and concentration of the energetic particles and the graphene oxide within the printable ink. The method of claim 1.

4. further comprising cutting the aerogel into pellets. The method of claim 1.

5. dispersing each of the graphene oxide, fuel nanoparticles and oxidizer nanoparticles separately in propylene carbonate under ultrasonic treatment for 3 hours; sonicating the dispersion of fuel nanoparticles and the dispersion of oxidizer nanoparticles under ultrasonication for 1 hour; mixing the dispersion of the fuel nanoparticles and the dispersion of the oxidizer nanoparticles with the dispersion of the graphene oxide for 5 minutes under stirring; allowing the resulting mixture to stand for at least 12 hours, thereby preparing the printable ink.

2. The method of claim 1 ,

6. forming a solution of 1 gram each of grade H-5 graphite and sodium nitrate in 46 mL of sulfuric acid; Stirring the solution in an ice-water bath for 10 minutes; adding 6 grams of potassium permanganate to the solution; heating the solution to 35° C. for 1 hour; adding 80 mL of deionized water dropwise to the solution; heating the solution to 90° C. for 30 minutes; Cooling the solution to room temperature; adding deionized water and 30% hydrogen peroxide to the solution until a pH of 5 is reached; sonicating and centrifuging the solution; and drying the solution at 65° C. to synthesize the graphene oxide. The method of claim 1.

7. The in situ mixing comprises: injecting the additive into the extrusion tube containing the printable ink at room temperature; The method of claim 1.

8. further comprising adjusting the material flow rate in the extrusion tube to 40 mm / sec for a total in situ mixing time of 6 seconds. The method of claim 1.

9. further comprising refreshing the alcohol every 12 hours for three days. The method of claim 1.

10. The alcohol is tert-butanol. The method of claim 1.

11. The additive is ethylenediamine. The method of claim 1.

12. The additive is butanediamine. The method of claim 1.

13. a first syringe containing a printable ink comprising graphene oxide; A second syringe containing an additive for reducing the graphene oxide; an adjustable syringe pump for driving the first syringe and the second syringe; an extrusion tube connected to the first syringe, the extrusion tube having an outlet for extruding the ink onto a substrate; a needle connected to a second syringe for injecting the additive into the extrusion tube at a location between the first syringe and the outlet; a stage for mounting the substrate; and two linear actuators for moving the stage in a horizontal plane. Additive manufacturing system.

14. a microcontroller for regulating a speed of movement of the syringe pump and providing an input signal to the linear actuator; The additive manufacturing system of claim 13.

15. The substrate is an acrylic plate. The additive manufacturing system of claim 13.

16. The extrusion tube is made of a polyvinyl chloride tube having a diameter of 1.6 mm. The additive manufacturing system of claim 13.

17. a porous crosslinked scaffold of reduced graphene oxide; and a plurality of nanothermite clusters embedded within the porous scaffold. Nanothermitic aerogel.

18. The nanothermitic clusters include fuel nanoparticles and oxidizer nanoparticles.

18. The nanothermitic aerogel of claim 17.

19. The fuel nanoparticles are comprised of aluminum nanoparticles with diameters up to 100 nm.

20. The nanothermitic aerogel of claim 18.

20. The oxidizer nanoparticles are bismuth oxide (Bi) nanoparticles up to 120 nm in diameter. ２ O ３ ) nanoparticles, 20. The nanothermitic aerogel of claim 18.

21. The reduced graphene oxide is 5-20% w / w; The nanothermite clusters are 80-95% w / w; 18. The nanothermitic aerogel of claim 17.

22. The reduced graphene oxide is 5% w / w; The nanothermite clusters are 95% w / w; 22. The nanothermitic aerogel of claim 21.

23. The reduced graphene oxide is 10% w / w; The nanothermite clusters are 90% w / w.

22. The nanothermitic aerogel of claim 21.

24. 20. Use of the nanothermitic aerogel of claim 17 as a propellant in a rocket engine.

25. 20. Use of the nanothermitic aerogel of claim 17 as a fuel for energy production.

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