An electrically controlled solid thruster afterburning grain and thruster system

Abstract

This invention belongs to the field of thrusters, specifically relating to an electrically controlled solid propellant afterburning propellant and a thruster system. The electrically controlled solid propellant afterburning propellant, by weight, comprises 65-95 parts oxidizer, 5-25 parts binder, and 0.3-5 parts curing agent. When the main combustion chamber is energized, the high-temperature, fuel-rich gas flow generated in the main combustion chamber washes over the surface of the afterburning propellant, rapidly reaching its pyrolysis temperature and instantaneously releasing oxygen. When applied to an electrically stimulated thruster combustion enhancement device, the afterburning propellant, upon heating, decomposes to produce oxygen-rich gas, which can mix and burn with the high-temperature, fuel-rich gas generated from incomplete combustion in the main combustion chamber. This fully utilizes the electrically controlled solid propellant, and the afterburning propellant offers controllable combustion, thereby significantly improving propellant combustion efficiency and thrust performance without affecting the thruster's electrically controlled combustion characteristics.

Description

Technical Field

[0001] This invention belongs to the field of thrusters, specifically relating to an electrically controlled solid rocket motor for afterburning propellant and a thruster system. Background Technology

[0002] Electro-controlled solid propellant propulsion technology utilizes a solid propellant (ECSP) with unique electrochemical properties. ECSP does not burn under high-temperature open flame, but it can be ignited by an electric current. By adjusting the applied voltage, the burning rate of the propellant can be controlled. When power is cut off, the propellant extinguishes; when power is restored, the propellant reignites, thus achieving re-ignition of the thruster. This technology has broad application prospects in the field of microsatellite propulsion. Electro-controlled solid propellants have become a research hotspot in the propulsion field due to their simple and reliable structure, re-start capability, and controllable thrust.

[0003] However, the non-flammability of the binder in electrically controlled solid propellants and the short residence time of oxidizer decomposition products in the combustion chamber result in incomplete combustion of the propellant within a limited time and space, leading to a large amount of incomplete combustion gases. This causes the actual mass flow rate of the working propellant and the gas temperature discharged during the operation of the electrically controlled solid thruster to be far lower than the theoretical values. More than 90% of the thrust of a solid thruster comes from its momentum thrust term (the product of the discharged working propellant mass flow rate and the exhaust velocity). To achieve sufficient thrust, it is necessary to increase the mass flow rate of the exhaust gases, meaning that the propellant combustion products must be converted into gases as completely as possible before being discharged; and to increase the exhaust velocity at the thruster nozzle exit, as the exit exhaust velocity is directly proportional to the propellant gas temperature, requiring an increase in the propellant gas temperature. Due to the incomplete combustion of the propellant during the operation of the electrically controlled solid thruster, the mass flow rate of the exhaust gases and the combustion chamber gas temperature are far lower than the theoretical values, resulting in the thrust performance of the thruster being far below the design value. To increase the combustion temperature of the propellant and achieve higher exhaust velocities, high-energy powder fuels, such as aluminum powder and boron powder, are typically added to electrically controlled solid propellants. However, the addition of high-energy powder fuels significantly reduces the controllability of the combustion of electrically controlled solid propellants, transforming it from non-self-sustaining combustion to self-sustaining combustion. Therefore, the current demand for improved thrust performance of thrusters, coupled with the incomplete combustion of propellants and the inability to significantly increase the exhaust gas temperature within a limited time / space range, is one of the main contradictions in the development of electrically controlled solid propulsion technology and a bottleneck problem that urgently needs to be solved in the engineering application and development of this technology. Summary of the Invention

[0004] The technical problem to be solved by the present invention is to provide an electronically controlled solid rocket motor afterburning charge and thruster system that can improve thrust performance and control combustion.

[0005] This invention provides an electronically controlled solid propellant booster combustion charge, which, by mass, comprises 65-95 parts of oxidant, 5-25 parts of binder, and 0.3-5 parts of curing agent.

[0006] Furthermore, the oxidant is one or more of the following oxidants: ammonium perchlorate, lithium perchlorate, magnesium perchlorate, barium perchlorate, calcium perchlorate, sodium perchlorate, silver perchlorate, rubidium perchlorate, cesium perchlorate, sodium nitrate, potassium nitrate, ammonium nitrate, and calcium nitrate.

[0007] Furthermore, the particle size of the oxidant is 50μm-500μm; Furthermore, the adhesive is one or more of the following: hydroxyl-terminated polybutadiene, polyvinyl alcohol, polyethylene oxide, polyurethane, polybutadiene, polysulfide rubber, and carboxyl-terminated polybutadiene.

[0008] Furthermore, the curing agent is one or more of toluene diisocyanate, boric acid, sodium tripolyphosphate, 3,3'-dichloro-4,4'-diaminodiphenylmethane, dicumyl peroxide, manganese dioxide, and bisphenol A epoxy.

[0009] Furthermore, when the adhesive is hydroxyl-terminated polybutadiene, the curing agent is toluene diisocyanate; When the adhesive is polyvinyl alcohol, the curing agent is boric acid; When the adhesive is polyethylene oxide, the curing agent is sodium tripolyphosphate; When the adhesive is polyurethane, the curing agent is 3,3'-dichloro-4,4'-diaminodiphenylmethane; When the adhesive is polybutadiene, the curing agent is dicumyl peroxide; When the adhesive is polysulfide rubber, the curing agent is manganese dioxide; When the adhesive is carboxyl-terminated polybutadiene, the curing agent is bisphenol A epoxy.

[0010] Furthermore, by weight, it also includes 0-5 parts of fuel, 0-5 parts of additives, 0-5 parts of plasticizer, and 0-2 parts of catalyst.

[0011] Furthermore, the fuel is one or more of a high-energy fuel element, a metal alloy, or a metal hydride; The catalyst is one or more of the following: ferrocene, tert-butylferrocene, octylferrocene, ethylferrocene, n-octylferrocene, iron oxide, copper oxide, and copper chromite. The additive is one or more of 1-methyl-1,2,4-triazole, 3-amino-1,2,4-triazole, 3-cyano-1,2,4-triazole, tetrazolium, 5-methyltetrazolium, 4-amino-1,2,4-tetrazolium, and 5-amino-1H-tetrazolium. The plasticizer is one or more of the following: diisooctyl sebacate, N-butylnitroethyl nitrate, glycerol, and polyethylene glycol.

[0012] The present invention also provides a thruster system, including an electrically stimulated responsive thruster combustion enhancement device. The electrically stimulated responsive thruster combustion enhancement device includes the aforementioned electrically controlled solid thruster afterburner propellant column disposed between the main combustion chamber outlet and the nozzle of the electrically controlled solid thruster. The electrically controlled solid thruster afterburner propellant column has a through hole in the middle connecting the main combustion chamber and the nozzle. The electrically controlled solid thruster afterburner propellant column is used to generate oxygen-rich gas through thermal pyrolysis, and mixes and burns with the high-temperature fuel-rich gas generated by incomplete combustion in the main combustion chamber. Moreover, the electrically controlled solid thruster afterburner propellant column has controllable combustion.

[0013] The beneficial effects of this invention are that in the afterburning propellant of this electrically controlled solid rocket motor, the oxidizer decomposes upon heating to generate oxygen-rich fuel gas. Furthermore, more than 65 parts of the afterburning propellant in this electrically controlled solid rocket motor are oxidizers, making the propellant inherently stable, yet releasing a massive amount of oxygen-rich fuel gas instantaneously upon pyrolysis. This also provides controllable combustion characteristics. A binder of 5-25 parts is used to bind the other components together to form a solid propellant. A curing agent of 0.3-5 parts reacts with the binder to cure it.

[0014] In this electrically controlled solid propellant thruster, the afterburning propellant, when the main combustion chamber is energized, generates a high-temperature, fuel-rich gas flow that washes over the surface of the afterburning propellant, rapidly reaching its pyrolysis temperature and releasing oxygen instantaneously. When applied to combustion enhancement devices for electrically stimulated thrusters, the afterburning propellant undergoes pyrolysis to produce oxygen-rich gas, which mixes and combusts with the high-temperature, fuel-rich gas generated from incomplete combustion in the main combustion chamber. This fully utilizes the secondary combustion achieved through gas-solid coupling of the electrically controlled solid propellant propellant. Furthermore, the afterburning propellant exhibits controllable combustion, meaning it possesses non-self-sustaining combustion characteristics (self-sustaining combustion occurs upon ignition and continues until the gas is completely burned, while non-self-sustaining combustion ceases when no high-temperature, fuel-rich gas is input). This significantly improves propellant combustion efficiency and thrust performance without affecting the thruster's electrically controlled combustion characteristics. Attached Figure Description

[0015] Appendix Figure 1 This is a schematic diagram of the structure of the present invention; Appendix Figure 2 This is a front sectional view of the present invention; Appendix Figure 3 This is a schematic diagram of the structure of the electronically controlled solid rocket motor for afterburning propellant in this invention.

[0016] In the diagram, 1-thruster housing; 2-cathode wiring structure; 21-terminal; 22-terminal plate; 3-spring; 4-electrode plate; 5-end plate; 6-electrode orifice plate; 7-main combustion chamber; 8-afterburning combustion chamber; 9-nozzle; 91-connecting ring; 10-electronically controlled solid propellant grain; 11-electronically controlled solid thruster afterburning grain; 111-through hole; 112-groove; 12-convex ring. Detailed Implementation

[0017] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.

[0018] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relative positional relationship and movement of each component in a certain specific posture (as shown in the figure). If the specific posture changes, the directional indication will also change accordingly.

[0019] Furthermore, in this invention, descriptions involving "first," "second," etc., are for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0020] In this invention, unless otherwise explicitly specified and limited, the terms "connection," "fixed," etc., should be interpreted broadly. For example, "fixed" can mean a fixed connection, a detachable connection, or an integral part; it can mean a mechanical connection, an electrical connection, a physical connection, or a wireless communication connection; it can mean a direct connection or an indirect connection through an intermediate medium; it can mean the internal communication of two elements or the interaction between two elements, unless otherwise explicitly limited. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0021] Furthermore, the technical solutions of the various embodiments of the present invention can be combined with each other, but only if they are feasible for those skilled in the art. If the combination of technical solutions is contradictory or cannot be implemented, it should be considered that such combination of technical solutions does not exist and is not within the scope of protection claimed by the present invention.

[0022] As attached Figure 1 - Appendix Figure 3 As shown, this invention provides an electrically stimulated response type thruster combustion enhancement device, including an electrically controlled solid propellant post 11 disposed between the outlet of the main combustion chamber 7 and the nozzle 9 of an electrically controlled solid propellant thruster. The electrically controlled solid propellant post 11 has a through hole 111 in the middle connecting the main combustion chamber 7 and the nozzle 9. The electrically controlled solid propellant post 11 is used to generate oxygen-rich gas through thermal pyrolysis, and mixes and burns with the high-temperature fuel-rich gas generated by incomplete combustion in the main combustion chamber 7. In this way, the electrically controlled solid propellant post 10 can be fully utilized to achieve secondary combustion through gas-solid coupling. Moreover, the electrically controlled solid propellant post 11 has controllable combustion, which means it has non-self-sustaining combustion performance (self-sustaining combustion performance means that self-sustaining combustion occurs after one ignition until the gas is completely burned, while non-self-sustaining combustion performance means that the reaction will stop when there is no high-temperature fuel-rich gas). Thus, without affecting the electrically controlled combustion characteristics of the thruster, the combustion efficiency of the propellant and the thrust performance of the thruster are greatly improved.

[0023] This invention utilizes the propellant chemical energy that was not completely consumed by setting an electronically controlled solid propellant supplementary combustion column 11 with controllable combustion between the outlet of the main combustion chamber 7 and the nozzle 9, thereby maximizing the release of the residual energy of the primary combustion and significantly improving the overall combustion efficiency of the electronically controlled solid propellant and the specific impulse and thrust of the thruster.

[0024] Simultaneously, the afterburning propellant 11 of the electronically controlled solid thruster is designed to have controllable combustion, maintaining the core characteristics of electronically controlled combustion in an electrically stimulated response thruster. In this case, the combustion of the afterburning propellant 11 relies on the high-temperature, fuel-rich gas discharged from the main combustion chamber 7 for heating and to provide the necessary gaseous environment. This means that when the thruster initiates combustion via an electronic control signal, the gas from the main combustion chamber 7 flows out and pyrolyzes the afterburning propellant 11. The unburned high-temperature, fuel-rich gas from the main combustion chamber 7 couples with the afterburning propellant 11 for secondary combustion, thus initiating and enhancing performance. When the thruster shuts off combustion via an electronic control signal (power off), the high-temperature, fuel-rich gas from the main combustion chamber 7 stops flowing out, losing its pyrolysis source and continuous supply of gaseous reactants. The combustion reaction of the afterburning propellant 11 immediately and automatically extinguishes (not self-sustaining) and will not continue to burn independently. Ultimately, it seamlessly integrates with electronic control capabilities. This "gas-driven, start-stop-with-the-main-chamber" characteristic means that the secondary combustion process is completely subject to the original electronic control switching logic. The thruster still maintains millisecond-level precise start-stop control and electrical regulation capability of combustion rate, without weakening the core controllability advantage of the electronically controlled solid rocket thruster.

[0025] Furthermore, this configuration ensures that the booster does not consume the electronically controlled solid rocket booster afterburner propellant 11 when not in operation. The electronically controlled solid rocket booster afterburner propellant 11 is only activated and consumed during the operation of the main combustion chamber 7, and its energy release is entirely dedicated to the thrust output phase, avoiding any ineffective energy waste.

[0026] The electrically stimulated thruster combustion enhancement device provided by this invention introduces an electrically controlled solid rocket motor propellant charge 11 with controllable non-self-sustaining combustion characteristics. This achieves efficient secondary utilization of the main propellant energy without sacrificing the inherent switching accuracy and responsiveness of the electrically controlled solid rocket motor, thereby significantly improving the overall energy conversion efficiency and thrust performance of the thruster. Its characteristic of "starting and stopping with the main chamber and having no self-sustaining risk" ensures the integrity of the electrical control capabilities and the safety of the system, representing a significant technological breakthrough. In one embodiment, the present invention further includes an end plate 5, a thruster housing 1, and a nozzle 9 arranged sequentially from upstream to downstream; The thruster housing 1 is hollow inside, and the end plate 5 and the thruster housing 1 enclose the main combustion chamber 7. This arrangement allows the main combustion chamber 7 to be connected to the nozzle 9, so that all the gas in the main combustion chamber 7 can flow to the nozzle 9, while ensuring a compact structure.

[0027] In one embodiment, a spring 3, an electrode plate 4, an electronically controlled solid propellant grain 10, and an electrode orifice plate 6 are sequentially arranged in the main combustion chamber 7 from upstream to downstream, and the holes on the electrode orifice plate 6 form the outlet of the main combustion chamber 7.

[0028] The electrode plate 4 and the electrode orifice plate 6 are used to apply voltage to the electrically controlled solid propellant grain 10, thereby controlling the start-up and shutdown of combustion. The holes on the electrode orifice plate 6 allow the high-temperature, fuel-rich combustion gas to enter the nozzle 9. The spring 3 ensures that the electrode plate 4, the electrically controlled solid propellant grain 10, and the electrode orifice plate 6 are in close contact with each other, guaranteeing control stability and effectiveness. Simultaneously, it ensures that the electrically controlled solid propellant grain 10 is in close contact with the outlet of the main combustion chamber 7 (the holes on the electrode orifice plate 6).

[0029] In one embodiment, a protruding ring 12 is provided on the inner wall of the thruster housing 1 on the outlet side of the main combustion chamber 7; The convex ring 12 is used to abut the downstream end of the electrode orifice plate 6 near its upstream end, and to abut the upstream section of the electronically controlled solid thruster afterburner propellant 11 near its downstream end.

[0030] In this embodiment, the inner diameter of the convex ring 12 is smaller than the outer diameter of the electrode orifice plate 6. By setting the convex ring 12, it is convenient to limit the installation of the electrode orifice plate 6, while ensuring the compact structure inside the cavity.

[0031] In one embodiment, a connecting ring 91 is provided on the upstream end face of the nozzle 9, and the connecting ring 91 is fitted into the hollow downstream end of the thruster housing 1. The upstream end of the electronically controlled solid rocket motor's afterburner charge 11 abuts against the downstream end of the convex ring 12, and the upstream end of the connecting ring 91 abuts against the downstream end of the convex ring 12.

[0032] In this embodiment, an axial positioning chain is formed consisting of "nozzle 9 - convex ring 12 - electronically controlled solid rocket motor afterburning propellant 11". The installation and positioning of the nozzle 9 and the electronically controlled solid rocket motor afterburning propellant 11 can be achieved by installing the nozzle 9. The convex ring 12 provides an axial hard positioning reference and support for the electronically controlled solid rocket motor afterburning propellant 11, making the position of the propellant in the electronically controlled solid rocket motor afterburning propellant 11 more accurate.

[0033] Furthermore, at this point, the inner wall of the convex ring 12 serves as a connecting section (annular channel) between the main combustion chamber 7 and the through-hole 111 on the electronically controlled solid rocket booster's afterburner propellant 11. This allows the main combustion gas flow to be forced across the surface of the electronically controlled solid rocket booster's afterburner propellant 11, maximizing heat transfer. This causes the high-temperature, fuel-rich gas to adhere tightly to and scour the upstream end face of the electronically controlled solid rocket booster's afterburner propellant 11. This forced close contact and scouring significantly improves the heat transfer efficiency (forced convection heat transfer) from the high-temperature, fuel-rich gas to the electronically controlled solid rocket booster's afterburner propellant 11. Simultaneously, a large amount of heat is transferred to the surface material of the electronically controlled solid rocket booster's afterburner propellant 11 within a very short distance (the axial length of the convex ring 12), rapidly and effectively driving the propellant material to undergo a pyrolysis reaction, continuously generating the required oxygen-rich gas. This is the core guarantee for achieving high efficiency in the first step of combustion enhancement (pyrolysis gas production).

[0034] In addition, at this time, before the high-speed, high-temperature fuel-rich gas flow enters the through hole 111 of the supplementary combustion propellant 11 of the electronically controlled solid thruster, it has already undergone intense turbulent mixing with the initial oxygen-rich gas released from the thermal decomposition of the supplementary combustion propellant 11 upstream of the electronically controlled solid thruster. In the high-temperature, high-turbulence region at the end of the convex ring 12 (near the inlet of the through hole 111), some of the unburned fuel in the main gas has the opportunity to undergo intense partial premixed combustion (or partial homogeneous combustion) with the active components in the oxygen-rich gas, igniting the entire secondary combustion process. That is, this structure can preheat and activate the surface of the supplementary combustion propellant 11 of the electronically controlled solid thruster, improving the speed of the secondary combustion response and its synchronization with the main combustion.

[0035] In addition, the inner wall of the convex ring 12 forms a pre-distribution cavity for the through hole 111 on the supplementary combustion column 11 of the electronically controlled solid thruster. This helps to temporarily buffer and initially equalize the pressure and flow of the gas (which is initially unevenly distributed) from the outlet of the main combustion chamber 7 in this annular space, and then distribute it relatively evenly to the circumferential points of the annular inlet of the through hole 111. This ensures that the pyrolysis of the inner surface of the supplementary combustion column 11 of the electronically controlled solid thruster is more uniform along the circumference, and the oxygen-rich gas produced is more evenly mixed with the main gas when it enters the main flow channel. This avoids local over-rich / over-lean situations and improves the uniformity and efficiency of the mixed combustion.

[0036] In this embodiment, the convex ring 12 is not only used to assist in the installation of the electrode orifice plate 6, the electronically controlled solid thruster afterburning charge 11 and nozzle 9, but also to assist in the entire combustion enhancement process.

[0037] In this embodiment, the section from the downstream end of the convex ring 12 to the upstream section of the connecting ring 91 forms a combustion chamber 8 to enhance combustion.

[0038] In one embodiment, the outer wall of the connecting ring 91 is threaded to the hollow downstream end of the thruster housing 1. This configuration simplifies the installation of the connecting ring 91 while ensuring the sealing of the connection between the connecting ring 91 and the thruster housing 1.

[0039] In one embodiment, the electrode plate 4 is a cathode plate, and the electrode orifice plate 6 is an anode plate; It also includes a cathode wiring structure 2, which includes a terminal 21 that passes through the end plate 5 and a terminal block 22 that is connected to the terminal 21 on the end plate 5. The upstream end of the spring 3 abuts against the terminal block 22 and the downstream end abuts against the electrode plate 4. The end plate 5 is an insulating end plate. The outer end of the cathode wiring structure 2 is connected to the negative terminal of the high voltage power supply. The electrode orifice plate 6 is connected to the positive terminal of the high-voltage power supply through the thruster housing 1.

[0040] In this embodiment, the electrode plate 4 is the cathode plate, the electrode hole plate 6 is the anode plate, and the terminal block 22 is placed at the cathode plate. The cathode has less reaction and a relatively low temperature, so the possibility of gas leakage and fastener failure is relatively small. Therefore, the overall structure has higher reliability.

[0041] In one embodiment, the through hole 111 of the electronically controlled solid rocket motor afterburner propellant 11 is star-shaped; The upstream section of the through hole 111 is provided with a groove 112. In this embodiment, the star-shaped design and the groove 112 can enhance the mixing and combustion of the front-end gas with the wall of the electronically controlled solid thruster afterburner propellant 11, thereby increasing the enhancement effect.

[0042] In one embodiment, the electronically controlled solid propellant charge 11, by weight, comprises 65-95 parts of oxidant, 5-25 parts of binder, and 0.3-5 parts of curing agent.

[0043] In this embodiment, the oxidizer is used to decompose upon heating. More than 65 parts of the afterburner propellant 11 in this electrically controlled solid rocket motor are oxidizers, serving as the source of oxygen-rich fuel gas. 5-25 parts of binder are used to bond the other components together to form a solid propellant stalk. 0.3-5 parts of curing agent are used to react with the binder to cure it.

[0044] The present invention also provides a method for enhancing combustion in an electrically stimulated thruster, using the above-mentioned electrically stimulated thruster combustion enhancement device, comprising the following steps: During combustion in the main combustion chamber 7, the supplementary combustion propellant 11 of the electronically controlled solid thruster is pyrolyzed to produce oxygen-rich gas, which mixes and burns with the high-temperature fuel-rich gas produced by incomplete combustion in the main combustion chamber 7. When combustion stops in the main combustion chamber 7, the electronically controlled solid propellant charge 11 stops pyrolysis to produce oxygen-rich gas.

[0045] This invention provides an electronically controlled solid propellant booster combustion charge, which, by mass, comprises 65-95 parts of oxidant, 5-25 parts of binder, and 0.3-5 parts of curing agent.

[0046] In this electrically controlled solid rocket motor, the oxidizer in the afterburning propellant 11 decomposes upon heating to produce oxygen-rich fuel gas. Furthermore, over 65 parts of the afterburning propellant 11 constitute the oxidizer, making the propellant inherently stable, yet releasing a massive amount of oxygen-rich fuel gas instantaneously upon pyrolysis. This also contributes to the controllability of combustion. A binder of 5-25 parts is used to bind the other components together, forming a solid propellant 11. A curing agent of 0.3-5 parts reacts with the binder to cure the propellant.

[0047] In this electrically controlled solid propellant thruster, when the main combustion chamber 7 is energized, a high-temperature, fuel-rich gas flow is generated in the main combustion chamber 7, which washes over the surface of the electrically controlled solid propellant thruster's afterburning propellant 11 and rapidly reaches its pyrolysis temperature, releasing oxygen instantaneously. When applied to an electrically stimulated thruster combustion enhancement device, the electrically controlled solid propellant thruster's afterburning propellant 11 undergoes pyrolysis to produce oxygen-rich gas, which can mix and burn with the high-temperature, fuel-rich gas generated from incomplete combustion in the main combustion chamber 7. This fully utilizes the secondary combustion achieved through gas-solid coupling of the electrically controlled solid propellant 10. Furthermore, the electrically controlled solid propellant thruster's afterburning propellant 11 has controllable combustion, meaning it exhibits non-self-sustaining combustion performance (self-sustaining combustion means that combustion begins immediately after ignition and continues until the gas is completely burned, while non-self-sustaining combustion means that the reaction stops when there is no high-temperature, fuel-rich gas). Thus, without affecting the electrically controlled combustion characteristics of the thruster, the combustion efficiency of the propellant and the thrust performance of the thruster are significantly improved.

[0048] In one specific embodiment, the mixture comprises, by weight, 85 parts of oxidant, 10.68 parts of binder, and 4.32 parts of curing agent. In this embodiment, when applied to an electrically stimulated thruster combustion enhancement device, a 50% increase in thrust can be achieved. In specific experiments conducted on this embodiment, the electrically controlled solid thruster afterburning propellant 11 can achieve three repeated ignition tests.

[0049] The afterburning propellant 11 in the electronically controlled solid rocket propulsion unit has an extremely high oxidizer content and will not burn in the absence of combustion gases. However, if the reaction between the afterburning propellant 11 and the combustion gases is too active, releasing a large amount of heat, this heat reacts on the electronically controlled solid propellant 10, causing its temperature to rise excessively and reach the critical temperature for self-sustaining combustion. In this case, the electronically controlled solid propellant 10 will continue to burn, and the thruster will lose its controllability. Conversely, if the reaction between the afterburning propellant 11 and the combustion gases is not active enough, the combustion gases cannot react in time when passing through the afterburning propellant 11, thus failing to produce a boosting effect.

[0050] In one embodiment, the oxidant is one or more of the following: ammonium perchlorate, lithium perchlorate, magnesium perchlorate, barium perchlorate, calcium perchlorate, sodium perchlorate, silver perchlorate, rubidium perchlorate, cesium perchlorate, sodium nitrate, potassium nitrate, ammonium nitrate, and calcium nitrate. Ammonium perchlorate, potassium nitrate, and lithium perchlorate are preferred. According to research, the oxidants suitable for use in the afterburning propellant grain 11 of an electronically controlled solid propellant, in descending order of reactivity, are ammonium perchlorate, potassium nitrate, and lithium perchlorate. Highly reactive oxidants can react rapidly with high-temperature, fuel-rich combustion gases, resulting in a significant enhancement effect. However, if the oxidant is too reactive, the small amount of fuel-rich combustion gases produced by the thermal decomposition of the electronically controlled solid propellant after power failure can still activate the reaction of the afterburning propellant grain, causing the thruster to self-sustain and lose controllability. In this embodiment, the above-mentioned oxidant has suitable reactivity, ensuring combustion controllability while preventing the electronically controlled solid propellant grain 10 from losing controllability.

[0051] In one embodiment, the particle size of the oxidant is 50 μm-500 μm.

[0052] The particle size of the oxidizer also has a significant impact on its performance. For example, using large-particle-size raw materials can reduce the reaction sensitivity of the afterburning propellant 11 in the electronically controlled solid rocket motor, while small-particle-size raw materials can improve the reaction sensitivity of the afterburning propellant 11. In this embodiment, by using the aforementioned particle size of the oxidizer, the afterburning propellant 11 of the electronically controlled solid rocket motor can be made to have a suitable reaction sensitivity.

[0053] In one embodiment, the adhesive is one or more of hydroxyl-terminated polybutadiene, polyvinyl alcohol, polyethylene oxide, polyurethane, polybutadiene, polysulfide rubber, and carboxyl-terminated polybutadiene.

[0054] In one embodiment, the curing agent is one or more of toluene diisocyanate, boric acid, sodium tripolyphosphate, 3,3'-dichloro-4,4'-diaminodiphenylmethane, dicumyl peroxide, manganese dioxide, and bisphenol A epoxy.

[0055] In one embodiment, when the adhesive is hydroxyl-terminated polybutadiene, the curing agent is toluene diisocyanate; When the adhesive is polyvinyl alcohol, the curing agent is boric acid; When the adhesive is polyethylene oxide, the curing agent is sodium tripolyphosphate; When the adhesive is polyurethane, the curing agent is 3,3'-dichloro-4,4'-diaminodiphenylmethane; When the adhesive is polybutadiene, the curing agent is dicumyl peroxide; When the adhesive is polysulfide rubber, the curing agent is manganese dioxide; When the adhesive is carboxyl-terminated polybutadiene, the curing agent is bisphenol A epoxy.

[0056] The combination of adhesive and curing agent in this embodiment is cleaner and can improve safety.

[0057] In one embodiment, the mixture, by weight, further includes 0-5 parts fuel, 0-5 parts additives, 0-5 parts plasticizer, and 0-2 parts catalyst. In this embodiment, adding fuel helps to increase the gas combustion temperature. Adding a catalyst can alter the oxidant decomposition reaction pathway or lower the reaction activation energy, thereby improving its pyrolysis effect and combustion controllability. Adding a plasticizer can improve casting fluidity and propellant uniformity, while adding additives can increase gas production and provide energy to promote combustion.

[0058] In one embodiment, the fuel is one or more of a high-energy fuel element, a metal alloy, or a metal hydride; wherein the high-energy fuel element is, for example, aluminum, magnesium, or boron; the metal alloy is, for example, an aluminum-magnesium alloy; and the metal hydride is, for example, aluminum trihydride.

[0059] The catalyst is one or more of the following: ferrocene, tert-butylferrocene, octylferrocene, ethylferrocene, n-octylferrocene, iron oxide, copper oxide, copper chromite, etc. The additive is one or more of the following: 1-methyl-1,2,4-triazole, 3-amino-1,2,4-triazole, 3-cyano-1,2,4-triazole, tetrazolium, 5-methyltetrazolium, 4-amino-1,2,4-tetrazolium, and 5-amino-1H-tetrazolium. The plasticizer is one or more of the following: diisooctyl sebacate, N-butylnitroethyl nitrate, glycerol, polyethylene glycol, etc.

[0060] This invention also provides a thruster system, including an electrically stimulated thruster combustion enhancement device. The device includes a supplementary combustion propellant column 11 disposed between the outlet of the main combustion chamber 7 and the nozzle 9 of an electrically controlled solid thruster. The supplementary combustion propellant column 11 has a through-hole 111 connecting the main combustion chamber 7 and the nozzle 9. The supplementary combustion propellant column 11 is used to generate oxygen-rich gas through pyrolysis, which mixes and combusts with the high-temperature, fuel-rich gas generated from incomplete combustion in the main combustion chamber 7. The supplementary combustion propellant column 11 provides controllable combustion. The specific effects of this thruster system are similar to those of the electrically stimulated thruster combustion enhancement device described above.

[0061] The above description is merely an embodiment and does not constitute any limitation on the present invention. Any person skilled in the art can make many possible variations, modifications, or alterations to the technical solutions of the present invention without departing from the scope of the present invention. Therefore, any simple modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the scope of the present invention, should fall within the protection scope of the present invention.

Claims

1. An electrically controlled solid rocket motor for afterburning propellant, characterized in that, By weight, it includes 65-95 parts of oxidant, 5-25 parts of adhesive and 0.3-5 parts of curing agent.

2. The electrically controlled solid rocket motor afterburning propellant as described in claim 1, characterized in that, The oxidant is one or more of the following: ammonium perchlorate, lithium perchlorate, magnesium perchlorate, barium perchlorate, calcium perchlorate, sodium perchlorate, silver perchlorate, rubidium perchlorate, cesium perchlorate, sodium nitrate, potassium nitrate, ammonium nitrate, and calcium nitrate.

3. The electrically controlled solid rocket motor afterburning propellant as described in claim 1, characterized in that, The particle size of the oxidant is 50μm-500μm.

4. The electrically controlled solid rocket motor afterburning propellant as described in claim 1, characterized in that, The adhesive is one or more of the following: hydroxyl-terminated polybutadiene, polyvinyl alcohol, polyethylene oxide, polyurethane, polybutadiene, polysulfide rubber, and carboxyl-terminated polybutadiene.

5. The electrically controlled solid rocket motor afterburning propellant as described in claim 1, characterized in that, The curing agent is one or more of toluene diisocyanate, boric acid, sodium tripolyphosphate, 3,3'-dichloro-4,4'-diaminodiphenylmethane, dicumyl peroxide, manganese dioxide, and bisphenol A epoxy.

6. The electrically controlled solid rocket motor afterburning propellant as described in claim 1, characterized in that, in When the adhesive is hydroxyl-terminated polybutadiene, the curing agent is toluene diisocyanate; When the adhesive is polyvinyl alcohol, the curing agent is boric acid; When the adhesive is polyethylene oxide, the curing agent is sodium tripolyphosphate; When the adhesive is polyurethane, the curing agent is 3,3'-dichloro-4,4'-diaminodiphenylmethane; When the adhesive is polybutadiene, the curing agent is dicumyl peroxide; When the adhesive is polysulfide rubber, the curing agent is manganese dioxide; When the adhesive is carboxyl-terminated polybutadiene, the curing agent is bisphenol A epoxy.

7. The electrically controlled solid rocket motor afterburning propellant as described in any one of claims 1-6, characterized in that, By weight, it also includes 0-5 parts fuel, 0-5 parts additives, 0-5 parts plasticizers, and 0-2 parts catalyst.

8. The electrically controlled solid rocket motor afterburning propellant as described in claim 7, characterized in that, The fuel is one or more of a high-energy fuel element, a metal alloy, or a metal hydride; The catalyst is one or more of the following: ferrocene, tert-butylferrocene, octylferrocene, ethylferrocene, n-octylferrocene, iron oxide, copper oxide, and copper chromite. The additive is one or more of 1-methyl-1,2,4-triazole, 3-amino-1,2,4-triazole, 3-cyano-1,2,4-triazole, tetrazolium, 5-methyltetrazolium, 4-amino-1,2,4-tetrazolium, and 5-amino-1H-tetrazolium. The plasticizer is one or more of the following: diisooctyl sebacate, N-butylnitroethyl nitrate, glycerol, and polyethylene glycol.

9. A thruster system, characterized in that, The device includes an electrically stimulated thruster combustion enhancement device, which includes an electrically controlled solid thruster afterburner propellant column (11) as described in any one of claims 1-8, disposed between the outlet of the main combustion chamber (7) and the nozzle (9) of the electrically controlled solid thruster. The electrically controlled solid thruster afterburner propellant column (11) has a through hole (111) in the middle connecting the main combustion chamber (7) and the nozzle (9). The electrically controlled solid thruster afterburner propellant column (11) is used to generate oxygen-rich gas by thermal pyrolysis and mix and burn with the high-temperature fuel-rich gas generated by incomplete combustion in the main combustion chamber (7). The electrically controlled solid thruster afterburner propellant column (11) has controllable combustion.

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