KR20240048983A - Monopropellant thruster assembly for satellite using ADN-based monopropellant - Google Patents

Abstract

The present invention includes a thruster valve that opens and closes a supply passage that supplies propellant from a propellant storage tank to the reactor; A reactor including a thermal reactor that heats the supplied propellant and a catalytic reactor that causes a decomposition reaction with a catalyst material; A heater that surrounds and heats the outer peripheral surface of the thermal reactor; a nozzle for discharging gas generated by decomposition in the reactor; a thermal barrier that structurally connects the thrust valve and the thermal reactor and blocks heat generated from the thermal reactor; It is characterized in that a capillary tube-shaped injector is arranged in the thermal barrier, metal spheres are arranged in the thermal reactor, and a catalyst carrier is arranged in the catalytic reactor.

Description

Monopropellant thruster assembly for satellite using ADN-based monopropellant}

The present invention relates to a single-propellant thruster assembly for a satellite using an ADN (Ammonium Dinitramide)-based single propellant, and more specifically, to a thruster assembly in which the reactor is separated into a thermal reactor for vaporizing the propellant and a catalytic reactor for decomposition.

A single-propellant thruster used for satellite attitude control causes a reaction in which liquid propellant is decomposed into gas within a reactor and generates thrust using the high-temperature, high-pressure gas generated after decomposition of the propellant.

Hydrazine (N ₂ H ₄ ) propellant, which is widely used in satellite attitude control thrusters to date, is a molecular compound that has a relatively high evaporation pressure and is highly toxic, making it difficult to handle on the ground. To improve these problems, the ADN-based monopropellant, which is being studied as an alternative propellant, is an ionic compound and has the advantage of lower vapor pressure and lower toxicity than hydrazine, so research continues as an alternative propellant that is easy to handle on the ground.

Ionic propellants such as ADN-based monopropellants have a stronger binding force between components than hydrazine, so relatively high activation energy is required to cause a decomposition reaction within the thruster, which requires preheating of the reactor, so a heater must be installed outside the thruster. .

Considering the operating temperature of the satellite and the characteristics of the propellant, ADN-based single propellant must be stored in an environment within 0℃ to 50℃. This means that the temperature of the propellant injected into the reactor is also a similar temperature.

The reaction temperature between a platinum-based catalyst and an ADN-based single propellant is known to be approximately 130°C to 150°C. This means that the propellant stored in the propellant tank can be decomposed only after the temperature is raised to around 100℃ when sprayed inside the thruster. Therefore, the propellant and the materials inside the reactor undergo heat exchange before the decomposition reaction begins, and after the heat exchange, when the temperature of the propellant reaches the decomposition temperature, the decomposition reaction begins. Therefore, if heat exchange does not occur very quickly, undecomposed propellant accumulates in the reactor, which causes adverse effects such as explosion of the thruster.

Generally, the catalyst material used in a single propellant thruster uses a ceramic-based support such as Al ₂ O ₃ to increase reactivity, but since the ceramic-based support has weak strength and low specific heat, it is stable after the propellant at room temperature is sprayed on the catalyst. In order for the decomposition reaction to continue, a very high preheating temperature is required. Since the satellite's attitude control thruster receives power from the satellite, there is a problem that the preheating temperature is too high and is not appropriate for operation in the limited environment of the satellite.

In addition, when a liquid propellant is injected directly onto the catalyst and heat exchange occurs, part of the catalyst is cooled, a slow decomposition reaction occurs in the locally cooled low temperature region, and the propellant accumulates in the catalyst. As the reaction continues, the catalyst in which the propellant accumulated in the cooled, slow decomposition process gains propellant decomposition energy, causing the temperature to rise again, and when energy to decompose the accumulated propellant is obtained at once, a decomposition reaction occurs. At this time, the catalyst is damaged due to excessive pressure in the process of decomposing a large amount of propellant from liquid to gas. For this reason, a thruster using an ADN-based single propellant requires a structure to protect the catalyst in the reactor that is different from the existing general single propellant thruster.

Korean Patent No. 10-1885829

In order to solve the above problems, the present invention seeks to provide a thruster assembly for stably reacting an ionic propellant such as an ADN-based single propellant.

The single-propellant thruster assembly for a satellite using an ADN-based single propellant according to the present invention for the above-described problem is a thruster valve that opens and closes the supply passage that supplies the liquid ADN propellant to the reactor from the liquid ADN (Ammonium Dinitramide) propellant storage tank. ; A reactor consisting of a thermal reactor for vaporization and thermal decomposition of the supplied liquid ADN propellant and a catalytic reactor for subsequent decomposition of the vaporized propellant into a catalytic material; A heater that surrounds and heats the outer peripheral surface of the reactor; a nozzle for discharging gas generated by decomposition in the reactor; a thermal barrier that structurally connects the thrust valve and the thermal reactor and blocks heat generated from the thermal reactor; Inside the thermal barrier, a propellant supply passage injector supplies propellant from the thruster valve to the thermal reactor, metal spheres are arranged within the thermal reactor, and a catalyst carrier is arranged in the catalytic reactor.

The thermal reactor is characterized in that it is preheated to 300°C to 400°C with a heater before operation.

The thermal reactor and the catalytic reactor are characterized by being distinguished using a thermal reactor retainer.

The catalytic reactor and the nozzle are characterized by being distinguished using a catalyst retainer.

The thrust valve is characterized in that it has a dual structure.

The injector is characterized in that a portion rotates circularly.

The metal sphere is characterized in that it is made of stainless steel.

The catalyst is characterized in that it is formed by supporting a reinforcing agent containing transition metals such as La, Ba, and Si and a platinum group active material such as Ir, Pt, and Ru on a pulverized porous ceramic support containing Al ₂ O ₃ .

The single-propellant thruster assembly for a satellite using an ADN-based single propellant according to the present invention is composed of a thermal reactor for vaporization and thermal decomposition of the propellant and a catalytic reactor for subsequent decomposition of the vaporized propellant, so that undecomposed propellant accumulates in the reactor. Accidents leading to explosions can be prevented.

In addition, by using a thermal reactor made of a heat transfer material with high heat resistance, oxidation resistance, and high specific heat, such as a stainless steel sphere, the conventional ceramic-based support has weak strength and low specific heat, so that the decomposition reaction is stable after the propellant at room temperature is sprayed onto the catalyst. In order to last, a very high preheating temperature is required, which solves the problem of high power consumption and can be used as a single-propellant thruster for satellites.

Figure 1 is a block diagram of a single propellant thruster assembly for a satellite using an ADN-based single propellant according to the present invention.
Figure 2 is a cross-sectional view of a single propellant thruster assembly for a satellite using an ADN-based single propellant according to the present invention.
Figure 3 is a cross-sectional view of the thermal barrier and the rotating circular supply channel blower in the thruster assembly according to the present invention.
Figure 4 is a perspective view of an injector with a rotating circular supply flow path according to the present invention.
Figure 5 shows a top view (a), side view (b), mesh (c) and retainer frame (d) of a retainer according to the present invention.
Figure 6 is a photograph of the highly heat-resistant catalyst material used in the present invention.
Figure 7 is a graph of temperature change in the reactor when 1.6 mm STS spheres are used in the thermal reactor.
Figure 8 is a pressure graph when 3 mm STS spheres were used in the thermal reactor.
Figure 9 is a pressure graph when 1.6 mm STS spheres are used in the thermal reactor.
Figure 10 is a pressure graph when 1.0 mm STS spheres are used in the thermal reactor.

Hereinafter, the present invention will be described with reference to specific examples and drawings. The embodiments of the present invention are intended to explain one invention, and the scope of rights is not limited to the illustrated embodiments, and the illustrated drawings are limited to the drawings because only the core contents are enlarged and incidental details are omitted for clarity of the invention. It should not be interpreted as such.

Figure 1 is a block diagram of a single propellant thruster assembly for a satellite using an ADN-based single propellant according to the present invention, and Figure 2 is a cross-sectional view of the thruster assembly of Figure 1.

The present invention includes a thruster valve (10) that opens and closes a supply passage that supplies liquid ADN propellant from a propellant storage tank to the reactor; A reactor (40) consisting of a thermal reactor for vaporization and thermal decomposition of the supplied liquid ADN propellant and a catalytic reactor for subsequent decomposition of the vaporized propellant into a catalytic material; A heater 60 that surrounds and heats the outer peripheral surface of the thermal reactor; A nozzle 50 that discharges gas generated by decomposition in the reactor; A thermal barrier 30 that structurally connects the thrust valve and the thermal reactor and blocks heat generated from the thermal reactor; Metal spheres are arranged in the propellant supply passage and injector 20 and the thermal reactor 41 within the heat barrier, and a catalyst carrier is arranged in the catalytic reactor 42.

The thruster assembly in the present invention supplies liquid propellant to the reactor 40 through the operation of the thruster valve. The reactor is preheated to below 300°C to 400°C before operation with the heater 60, and is exposed to a temperature of around 1400°C during operation. Therefore, the two structures are equipped with a heat barrier (30) to prevent the temperature of the reactor from affecting the thruster valve. ), and the propellant supply channel and injector (20) connects the thruster valve (10) and the thermal reactor (41) to supply propellant.

The main object of preheating is the thermal reactor 41, but in reality, the entire thruster assembly connected to the reactor, including the metal sphere and catalyst, is heated, so a thermal barrier is used to protect the valve from heat.

The injector 20 forms the propellant supply passage in the form of a capillary tube to minimize the heat transfer area, and rotates the supply passage 360° as shown in Figures 3 and 4 to take thermal expansion into account and increase the heat transfer length. Form into a circle. In the present invention, the end of the capillary tube serves as an injector without using a separate injector.

When the capillary tube-type supply passage and sprayer are used separately, they are usually welded, so a space is created between the capillary tube and the sprayer. The inside of the reactor is preheated to 300°C to 400°C before operation, but depending on the structure, the injector is connected to the outside of the reactor, so the temperature rises to around 400°C. In this case, when the propellant is injected, there is a very high possibility that an explosion due to thermal decomposition will occur between the propellant supply passage and the injector before it enters the reactor through the injector from the propellant supply passage.

In the case of a typical 1N thruster used in space vehicles, the outer diameter of the capillary tube is about 1.6 mm and the inner diameter is about 0.25 mm, so a separate injector is not necessary.

The generally used single propellant injector thrust range is generally around 1N, around 50N, and around 200N. At 50N and 200N, the propellant supply flow path is 1/4 inch (6.35 mm) and 3/8 inch (9.53 mm) pipe. However, the diameter of the injector plate suitable for this is 30 to 50 mm, so the injector cannot be mounted at the end of the propellant supply channel.

Due to the structure of the thruster assembly, if the distance between the valve and the reactor is long, it becomes vulnerable to external forces such as vibration, so the heat barrier, which is a structural support, must be stronger. In the case of a thruster that does not require preheating, a straight, short propellant supply channel can be used because there are few problems due to the temperature rise of the propellant supply channel. However, in the case of ADN-based liquid propellant or similar structures that require preheating, or if the distance between the valve and the reactor is very short, the reactor It minimizes the heat transfer that occurs in the , and when the length of the propellant flow path suitable for the design flow rate is required, it can be rolled round or bent into a triangle as in the present invention.

Pressure drop is proportional to the channel length and inversely proportional to the channel cross-sectional area. In terms of flow rate and pressure drop, the same effect can be achieved by shortening the capillary tube and applying a crimping part, but a certain length of the capillary tube is required to prevent heat transfer due to preheating, and the physical distance between the valve and the reactor is minimized. To do this, a form (21) in which the tube is wound one turn is preferable.

The reason why there are holes in the heat barrier (30) is to minimize the heat transfer area by drilling holes in the middle because all the annular tubes have an area where heat can be transferred. In some cases, a cylindrical tube-shaped heat barrier is formed without making holes. In some cases, it is applied as is. If a hole is drilled in the heat barrier, the mechanical rigidity becomes weaker.

The injector is connected to a thermal reactor, and it is desirable to use a metal sphere with high heat resistance, oxidation resistance, and high specific heat for the thermal reactor. To preheat the thermal reactor, wrap a heater around it to ensure that the heat from the heater is well supplied to the thermal reactor.

The thermal reactor in which metal spheres are arranged and the catalytic reactor in which high heat-resistant catalysts are arranged are distinguished using a thermal reactor retainer (43), and the housing of the reactor is provided with a step (45) to secure the thermal reactor retainer (43). Process.

The highly heat-resistant catalytic reactor (42) preferably has a large active point and a large specific surface area so that the propellant vaporized through the thermal reactor (41) can be completely decomposed.

The catalyst retainer 44 fixes the position of the catalyst, and as the thruster operates for a long period of time, there is a possibility that part of the catalyst may be lost, and the particles lost from the catalyst may block the nozzle throat, so the size of the catalyst and the nozzle throat are affected. Considering the size, it is desirable that the mesh 47 constituting the retainer is sufficiently small.

Figure 5 (a) is a top view of the thermal reactor retainer 43, (b) is a side cross-sectional view, (c) shows the mesh, and (d) shows the retainer 43 frame. The mesh placed on the retainer frame is a retainer. The retainers of the thermal reactor and the catalytic reactor are similar in shape with only a different size.

The highly heat-resistant catalyst in the present invention is made by heat-treating gamma-state alumina with a reinforcing agent to create a hexaaluminate structure, and is characterized by a larger specific surface area than the alpha-phase alumina of the prior art and higher strength than the gamma-phase. In the case of the alpha phase, it is produced by heat-treating alumina at over 1,200°C, and the strength is very strong, but the specific surface area is greatly reduced and the activity is reduced. The gamma phase is alumina below 750℃ and has a very high specific surface area, but it is not suitable for ADN-based liquid propellants because it undergoes a phase transformation as it sinters into the alpha phase above 1,200℃. Catalysts using gamma-phase alumina are mainly applied to hydrogen peroxide and hydrazine thrusters.

The theoretical decomposition temperature of ADN-based liquid propellant is in the late 1,600°C, and when actually operated in a 1N thruster, the temperature is measured at the level of 1,400°C to 1,500°C. The high heat resistance catalyst in the present invention is characterized by heat-treating the La reinforcer with alumina at over 1,400°C to form a hexaaluminate structure. The hexaaluminate structure is a structure created when heat-treated with a transition metal reinforcing agent at high temperature. Although the hexaaluminate structure has a lower specific surface area than the gamma phase, it has a specific surface area necessary for catalytic reaction and is characterized by a strength comparable to that of the alpha phase.

The carrier is a transition metal such as La, Ba, and Si in the case of the reinforcing agent, and the active metal is a platinum-based noble metal such as Pt, Ir, and Ru. In the present invention, La is used as a reinforcing agent and Pt is used as an active metal. ADN-based liquid propellant generates oxygen during the decomposition reaction, and the catalyst is exposed to strong acids in a high-temperature environment, so not only the support is damaged, but also phase transformation such as sintering must not occur, and the active material must not be oxidized. The propellant itself contains a lot of ammonia and nitrogen, so there are many substances that simply react when heated, but substances that can withstand long-term conditions of high temperature, strong acid, and oxidation are limited to platinum group precious metals.

The thruster assembly of the present invention is structurally connected to a thruster valve, a heat barrier, a housing, a reactor, and a nozzle, and internally includes thruster valves 11 and 12, an injector 20, a thermal reactor 41, and a catalytic reactor ( 42). The structural connection is to maintain the shape of the thruster assembly under vibration/shock applied from outside the thruster and the high temperature/pressure generated inside the thruster, and the internal connection is to stably decompose the propellant and ensure that the thruster operates normally for its intended purpose. am.

Since ADN-based liquid propellants are produced by dissolving solid ADN salts, when exposed to vacuum, the solvent may evaporate and the ADN salts may precipitate again. If ADN salts precipitate within the valve, the moving parts of the valve may become stuck and the valve may not operate normally. Therefore, it is desirable that the thruster valve be manufactured in a redundancy structure (11, 12) considering that the airtightness performance may deteriorate during repeated operation.

In addition, since the propellant is continuously stored in the thruster valve, when the thruster valve is manufactured as an in-line type, it is desirable to manufacture it at low power so that heat generated in the coil does not affect the propellant. The propellant is very sensitive to heat, so if the coil heats up or external heat is transferred, there is a possibility that the propellant will explode within the valve.

The heat barrier 30, which structurally connects the thruster valve and the reactor, has a thickness that is sufficient to prevent damage from vibration/impact applied to the outside of the thruster, and the heat transfer area is considered to prevent heat generated in the reactor from being transferred to the thruster valve. It must be produced.

The thickness, diameter, and length of the injector 20 that connects the thruster valve and the reactor must be set in consideration of the injector pressure loss and heat transfer area. For a 1N class ADN-based liquid propellant thruster, a capillary tube with a length of approximately 50 mm and an outer diameter of 1/16″ or less is appropriate in terms of heat transfer. The injector inner diameter can vary depending on the propellant flow rate and combustion chamber pressure. Depending on the operation of the reactor, the temperature of the injector may rise to the level of 200°C to 300°C. Therefore, considering thermal expansion due to temperature increase, it is preferable that the injector rotates in a circular manner rather than in a straight line.

For normal operation of the thruster, operate the heater before operating the thruster valve to increase the temperature inside the thermal reactor to 300℃ ~ 400℃. The preheating temperature varies depending on the type and amount of catalyst and the type and amount of the thermal reactor, but in terms of the response speed of the thruster, the higher the temperature of the thermal reactor, the more preferable. However, because the temperature of the thermal reactor is related to the heater's operation time and energy consumption, an appropriate preheating temperature must be set according to the satellite's mission.

The preheating temperature of the thermal reactor depends on the temperature maintained when the thermal reactor is cooled after the propellant is injected. A temperature above 150°C, at which the ADN-based liquid propellant is judged to be sufficiently vaporized, is the minimum temperature standard for the thermal reactor. If the temperature of the thermal reactor is below 150℃, the decomposition reaction may slow down and adverse effects may occur.

The thermal reactor can use spheres made of various metal materials, including stainless steel spheres. Considering that the reactor's internal temperature will be exposed to 1000°C or lower, a metal with good heat resistance and oxidation resistance and high heat transfer and specific heat can be used. The stainless steel sphere constituting the thermal reactor in the present invention has a size of 1 to 3 mm. Smaller spheres pack the thermal reactor more densely and provide more surface area in contact with the propellant for heat exchange. However, unnecessary pressure loss occurs because it reduces the flow area of the vaporized propellant. The larger spheres contact the propellant at a smaller surface area, which results in a relative reduction in heat exchange efficiency, but also reduces the pressure losses occurring in the thermal reactor. In addition, the smaller sphere showed more desirable test results in terms of pressure perturbation, which can be expressed as combustion stability.

Generally, in a single propellant thruster, a decomposition reaction occurs when the propellant is sprayed onto the catalyst. If appropriate decomposition conditions, such as a temperature that satisfies the activation energy, are met, the propellant is decomposed in a conventional reactor and the monopropellant thruster operates. For example, a thruster that uses more than 90% hydrogen peroxide as a single propellant can operate because the catalytic reaction occurs at room temperature (300 K).

Although it carries out a catalytic reaction at room temperature, it is a slightly slow reaction at room temperature, and the hydrazine thruster, which can react quickly when the temperature rises slightly, is mainly used as a single reactor structure.

However, because ADN-based liquid propellant requires high activation energy for operation, it is necessary to raise the temperature of the propellant by heat exchange before the catalytic reaction. The specific heat and preheating temperature of alumina used as a catalyst support determine the amount of heat required to raise the temperature of the propellant. can't afford it. In addition, in the case of conventional metal monoliths, because they are metals, they can handle the pressure caused by the phase change of the propellant, but because they have low density, they cannot provide the amount of heat necessary to raise the temperature of the propellant during preheating.

The decomposition reaction of ADN is relatively complex as shown below, and hydrogen peroxide (H ₂ O ₂ -> H ₂ O + O ₂ ) or hydrazine (N ₂ H ₄ -> NH ₃ + N ₂ -> N ₂ + H ₂ ) Unlike other products, it does not decompose easily due to catalytic reactions at room temperature because it has a multi-step decomposition path.

It is the platinum group active material that can stably decompose ADN-based liquid propellant for a long time, and its activation temperature is around 130℃ to 150℃. The decomposition reaction below of ADN is an example, and much research is still in progress.

For the above reasons, the present invention proposes a thruster structure in which the propellant is first decomposed through thermal decomposition in a thermal reactor rather than directly catalytically decomposing the propellant.

The highly heat-resistant catalyst of the present invention is formed by supporting a reinforcing agent such as La, Ba, etc. and a noble metal active material such as Ir, Pt, etc. on a pulverized porous ceramic support such as Al ₂ O ₃ .

Figure 6 is a photograph of the catalyst Pt/La-Al ₂ O ₃ used in an example of the present invention.

Figure 7 is a graph of temperature change in the reactor when 1.6 mm STS spheres are used in the thermal reactor. Figure 7 shows temperature changes in the thermal reactor and catalytic reactor when ADN-based liquid propellant is injected. The temperature of the thermal reactor is the black line, and the temperature of the catalytic reactor is the red line. The thermal reactor is cooled because the low-temperature propellant is supplied. The components of ADN-based liquid propellant are ADN, methanol, ammonia, and water, which all vaporize or thermally decompose at temperatures above 150°C. Therefore, in Figure 7, since the temperature of the thermal reactor is maintained above 150°C, it can be determined that all of the injected liquid propellant has been vaporized or thermally decomposed within the thermal reactor. This means that all of the primary decomposed propellant introduced into the catalytic reactor is in a gaseous state.

Because the constituents of ADN-based liquid propellants used for space vehicles contain water, it is very difficult for catalytic reactions to occur below 100°C without raising the temperature of the propellant through heat exchange. One of the main characteristics of ADN liquid propellant is that it is stable and does not easily decompose when in an aqueous solution state. Strictly speaking, it is thought that the vaporization temperature of the solvent will increase further as the pressure inside the reactor increases, but in-depth research on the thermal characteristics of ADN-based liquid propellants has not yet been conducted.

Since energy is released when ADN thermally decomposes, the temperature of the thermal reactor is maintained when the energy cooled by the propellant and the energy released as the propellant decomposes are balanced. When more energy is released, the temperature rises. If the amount of heat cooled by the propellant is greater, the conditions for vaporization, thermal decomposition, and catalytic decomposition are not met, so the propellant accumulates in a liquid state. If the valve is closed and the propellant is no longer supplied, heat from outside the reactor flows into the reactor. This increases the temperature of the metal sphere and catalyst, and when the temperature reaches the decomposition temperature of the propellant, an adverse reaction occurs where the accumulated propellant reacts at once, and the thruster may explode.

Figure 7 shows that the temperature of the thermal reactor does not decrease and remains below 150°C even though the propellant is continuously supplied. This means that energy exchange within the thermal reactor is occurring appropriately.

Figure 8 is a pressure graph when 3 mm STS spheres were used in the thermal reactor. The main content of the present invention is not only to use metal spheres in a thermal reactor, but also to suggest appropriate sizes of metal spheres. In Figures 8-10, it is most desirable for the pressure graph to be a straight line under the same test conditions with different sizes of the metal spheres. However, in Figure 8, the pressure fluctuation is greater than in Figures 9 and 10.

Fluctuations in pressure indicate combustion instability, and this is presumed to be related to the decomposition process of ADN-based liquid propellant and the exposure of metal spheres and catalysts to extreme conditions such as high temperature and strong acid during decomposition. However, this is not certain because very limited combustion test data for ADN-based liquid propellants has been released. Therefore, there is still no standard for what level of pressure fluctuation is appropriate during combustion in ADN-based monopropellant thrusters. However, based on the inventor's research experience, the pressure fluctuation at the level of Figure 9 is acceptable, and Figure 10 shows a more improved pressure fluctuation. Therefore, looking at the pressure fluctuation data, smaller metal spheres are more desirable, but when filling the reactor with smaller spheres, the pressure drop that occurs as the propellant passes through the thermal reactor and the catalytic reactor increases, so an appropriate compromise is necessary.

The decomposed propellant passes through a thermal reactor and a highly heat-resistant catalyst and is accelerated through a nozzle. Since the propellant is all decomposed before reaching the nozzle, a separate space is unnecessary after the high heat resistance catalyst. The propellant gasified and accelerated through the nozzle generates thrust.

When an ionic propellant is vaporized by applying a thermal reactor made of metal as in the present invention, the thruster can be operated stably. The size and type of thermal reactor must be applied differently depending on the composition of the propellant and the preheating temperature, but minimizing the empty space inside the reactor is desirable for combustion stability.

10,11,12: thruster valve 20: injector
21: Circular rotating supply channel sprayer 30: Heat barrier
40: reactor housing 41: thermal reactor
42: catalytic reactor 43: thermal reactor retainer
44: catalyst retainer 45,46: step
47: mesh 50: nozzle
60: heater

Claims (8)

A thruster valve that opens and closes a supply passage that supplies liquid ADN propellant to the reactor from the liquid ADN (Ammonium Dinitramide) propellant storage tank;
A reactor consisting of a thermal reactor for vaporization and thermal decomposition of the supplied liquid ADN propellant and a catalytic reactor for subsequent decomposition of the vaporized propellant into a catalytic material;
A heater that surrounds and heats the outer peripheral surface of the reactor;
a nozzle for discharging gas generated by decomposition in the reactor;
a thermal barrier that structurally connects the thrust valve and the thermal reactor and blocks heat generated from the thermal reactor;
A propellant supply channel injector that supplies propellant from the thruster valve inside the heat barrier to the thermal reactor, and
A single propellant thruster assembly for a satellite using an ADN-based single propellant, characterized in that metal spheres are arranged in the thermal reactor and a catalyst carrier is arranged in the catalytic reactor. According to paragraph 1,
A monopropellant thruster assembly for a satellite using an ADN-based monopropellant, characterized in that the thermal reactor is preheated to 300 ℃ ~ 400 ℃ with a heater before operation. According to paragraph 1,
A monopropellant thruster assembly for a satellite using an ADN-based monopropellant, characterized in that the thermal reactor and the catalytic reactor are distinguished using a thermal reactor retainer. According to paragraph 1,
A single propellant thruster assembly for a satellite using an ADN-based single propellant, characterized in that the catalytic reactor and the nozzle are separated using a catalyst retainer. According to paragraph 1,
A single-propellant thruster assembly for a satellite using an ADN-based single propellant, characterized in that the thruster valve has a dual structure. According to paragraph 1,
A monopropellant thruster assembly for a satellite using an ADN-based monopropellant, wherein a portion of the injector rotates in a circular manner. According to paragraph 1,
A single propellant thruster assembly for a satellite using an ADN-based single propellant, characterized in that the metal sphere is made of stainless steel. According to paragraph 1,
The catalyst is an ADN-based single propellant, which is formed by supporting a transition metal reinforcement including La, Ba, and Si and a platinum group active material including Ir, Pt, and Ru on a pulverized porous ceramic support containing Al ₂ O ₃ . A monopropellant thruster assembly for a satellite.