JP2010229852A - Catalyst decomposition type thruster for space vehicle - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a catalyst decomposition type thruster for a space vehicle, for actualizing high safety and easy assembly even when using HAN (Hydroxylammonium Nitrate) based liquid propellant for one liquid propulsion system. <P>SOLUTION: The catalyst decomposition type thruster for the space vehicle includes a hollow tank for storing the liquid propellant, a combustor for decomposing the liquid propellant and injecting reactive gas, and a propellant valve for controlling the supply of the liquid propellant from the tank to the combustor. The combustor includes a catalyst layer having a catalyst to be used for decomposing the liquid propellant, a heating device for heating the catalyst layer, and an injector for supplying the liquid propellant to the catalyst layer. Between the injector and the propellant valve, a heat insulating member is provided which has a flow path for the liquid propellant from the propellant valve to the combustor. <P>COPYRIGHT: (C)2011,JPO&INPIT

Description

  The present invention relates to a catalytic decomposition type thruster for space vehicles. In particular, the present invention relates to a catalytic decomposition type thruster that can be suitably used for a one-component propulsion system used in a space vehicle.

As a propulsion system used to control the attitude and orbit of a spacecraft such as a spacecraft in outer space, a propelling force has been conventionally achieved by decomposing a liquid propellant with a catalyst and injecting the resulting reaction gas. Catalytic cracking thrusters, which are so-called one-part propulsion systems, have been used.
Hydrazine is often used as a fuel for a one-component propulsion system for space vehicles. Hydrazine is an excellent fuel because of its high reactivity, but it is dangerous to handle. Therefore, in recent years, hydroxylammonium nitrate (“HAN”)-based liquid propellants have attracted attention as low-toxic liquid propellants that can replace hydrazine.
For example, Japanese Patent Application Laid-Open No. 2007-023135 discloses hydroxylammonium nitrate as a highly safe liquid oxidizing agent used for a monopropellant (single liquid propellant) that is a propellant that functions only with one liquid. HAN) and hydrazinium nitrate (HN) are dissolved in water (H ₂ O), and a liquid oxidizer containing 10% by weight or less of a fuel component is disclosed. A hot gas generation method is disclosed in which liquid fuel is stored separately and mixed or brought into contact immediately before use to ignite and generate hot gas.
Japanese Patent Laid-Open No. 2004-340148 discloses that a HAN group propellant is introduced into a reactor, the propellant is decomposed so as to dissociate at least most of the HAN in the propellant, and the output of the reactor is burned. A method is disclosed in which the output of the reactor is combusted in the combustor such that the dissociated HAN dissociation product is combusted with unreacted fuel in the propellant.

As the catalytic decomposition type thruster when hydrazine is used as the fuel, for example, the one shown in FIG. 1 can be considered. In the catalytic decomposition type thruster 10, a liquid propellant such as hydrazine is stored in a propellant tank (not shown), and the opening and closing of the propellant valve 11 is controlled to make a metal capillary with a small diameter. Feed to combustor 13 through tube 12. The combustor 13 includes a catalyst layer 14 that catalytically decomposes a liquid propellant such as hydrazine. The liquid propellant such as hydrazine that has been conveyed to the injector 15 of the combustor 13 through the capillary tube 12 is jetted toward the catalyst layer 14 by the injector 15.
In the catalytic decomposition type thruster shown in FIG. 1, the propellant valve 11 and the combustor 13 are connected by the capillary tube 12 having a small diameter for the following reason. That is, when the liquid propellant is catalytically decomposed in the combustor 13, high-temperature combustion gas is generated, so that the temperature of the combustor 13 rises. When this heat is conducted from the combustor 13 to the propellant valve 11 through the pipe for supplying the liquid propellant, there is a risk that the liquid propellant self-decomposes and explodes at a high temperature. For this reason, in the catalytic decomposition type thruster, the propellant valve 11 and the combustor 13 are connected using a small-diameter metal capillary tube 12 having a small heat capacity, so that heat is not easily transmitted to the propellant valve 11. is there.

JP 2007-023135 A JP 2004-340148 A

The conventional catalytic decomposition type thruster 10 as shown in FIG. 1 has various problems because it employs the metal capillary tube 12 having a small diameter for the reasons described above.
That is, first, both ends of a metal capillary tube having a small diameter (outer diameter of 2 mm or less) are welded to a metal plate while preventing the tube from being crushed, and then the metal plate is attached to the propellant valve and the combustor 13. Each connection requires an extremely complicated assembly process. In order to prevent heat conduction from the combustor to the propellant valve, when processing the metal capillary tube in a spiral shape, an adapter for holding the spiral metal capillary tube is required. The score will increase.
Further, impurities may be present in the liquid propellant. Further, when the propellant valve 11 is a solenoid valve, foreign matters such as burrs generated by friction between metal parts constituting the solenoid valve may be mixed when the liquid propellant passes through the propellant valve 11. is there. These impurities and foreign matters can cause the liquid propellant to block the flow path when passing through the small-diameter metal capillary tube 12. Further, when the catalytic decomposition type thruster 10 is in a weightless state such as a space environment, or is in an assembly process on the ground, the catalyst having a small particle size detached from the catalyst layer 14 flows backward, There is a possibility of closing the small-diameter metal capillary tube 12.
Further, in the catalytic decomposition type thruster as shown in FIG. 1, when a HAN liquid propellant is used instead of hydrazine, a new problem occurs. That is, although the HAN liquid propellant is superior in terms of low toxicity as compared with hydrazine, the reactivity is low. For this reason, in order to catalytically decompose the HAN-based liquid propellant, it is necessary to heat the catalyst in the catalyst layer 14 in advance to increase the activity of the catalyst at a high temperature. In such a case, if a heater for preheating the catalyst layer 14 is provided, heat from the heater is conducted from the combustor 13 to the metal capillary tube 12. In addition, since the HAN-based liquid propellant has high performance, the temperature of the combustor 13 rises greatly due to the generation of combustion gas, and the amount of heat flowing from the combustor 13 into the metal capillary tube 12 is much larger than before. It becomes. As a result, in the system using the conventional metal capillary tube 12, the propellant valve 11 may be heated to a high temperature that deviates from the temperature limit as a component. On the other hand, there is a limit to solve this problem by further reducing the diameter of the metal capillary tube 12, and it becomes difficult to supply sufficient liquid propellant to the combustor 13. Further, it is obvious that the problem of the blockage of the capillary tube 12 as described above is exacerbated as the diameter is reduced.

  Therefore, an object of the present invention is to provide a catalytic decomposition type thruster that can obtain high safety and easy assembly even when a HAN liquid propellant is used in a single liquid propulsion system. Is.

  Instead of the conventional metal capillary tube, the present inventors connect the injector and the propellant valve with a heat insulating member, and by providing a liquid propellant flow path in the heat insulating member, While preventing heat conduction from the combustor to the propellant valve, it is possible to provide a liquid propellant flow path with a size that is difficult to block, and based on the knowledge that the above problems can be solved, the present invention has been achieved. It is a thing.

  That is, the present invention is a catalytic decomposition type thruster for a spacecraft, which obtains a propulsive force by injecting a reaction gas obtained by decomposing a liquid propellant with a catalyst, and is a hollow for containing a liquid propellant. And a propellant valve for controlling the supply of liquid propellant from the tank to the combustor, the combustor further comprising a liquid propellant. A catalyst layer having a catalyst for causing catalyst decomposition, a heating device for heating the catalyst layer, and an injector for supplying a liquid propellant to the catalyst layer are provided, and a thrust is provided between the injector and the propellant valve. There is provided a catalytic decomposition type thruster provided with a thermally insulating member having a liquid propellant flow path from a chemical valve to a combustor.

  According to the present invention, since the injector and the propellant valve are connected by the heat insulating member, heat conduction from the combustor to the propellant valve can be prevented, and the explosion caused by the self-decomposition of the liquid propellant And the possibility of failure due to deviation from the temperature limit of the propellant valve can be effectively avoided. Further, when the heat-insulating member is processed to provide the liquid propellant flow path, the flow path can have a size that is difficult to be blocked by impurities in the liquid propellant, mixed foreign matter, or the like. Furthermore, since the heat insulating member is easy to handle and process in both size and material, it is easy to assemble a catalytic decomposition type thruster.

It is a cross-sectional schematic diagram of a conventional catalytic decomposition type thruster. It is a cross-sectional schematic diagram of the catalytic decomposition type thruster by this invention.

Hereinafter, embodiments for carrying out the present invention will be described with reference to the drawings.
FIG. 2 is a schematic cross-sectional view of the catalytic decomposition type thruster according to the present invention. The catalytic decomposition type thruster 1 is a catalytic decomposition type thruster that obtains a driving force by injecting a reaction gas obtained by decomposing a liquid propellant with a catalyst. This catalytic decomposition type thruster 1 includes a hollow tank (not shown) for containing a liquid propellant, a combustor 2 that decomposes the liquid propellant and injects a reaction gas, and a liquid from the tank to the combustor. And a propellant valve 3 for controlling the supply of the propellant.

a. Tank The tank used in the present invention is not particularly limited. For example, a cylindrical tank having an inner diameter of 50 mm can be used. In view of the strong acidity of the HAN liquid propellant, it is desirable that the material of the tank be a stainless steel or titanium material that has been confirmed to be compatible by a material compatibility test.
The tank includes a piston as an example of a gas-liquid separation mechanism. The piston forms a compartment filled with liquid propellant inside the tank, and the liquid propellant opens and closes the propellant valve 3 when the piston operates. Accordingly, the fuel can be sent to the combustor 2. The tank having such a gas-liquid separation mechanism is important when the catalytic decomposition type thruster of the present invention is used for a space vehicle. In other words, if a gas-liquid separation mechanism is provided inside the tank to form a section filled with liquid propellant, the liquid propellant is unevenly distributed near the inner wall of the tank under zero gravity when used in a space environment. Therefore, it is possible to prevent the gas from being discharged when the liquid propellant is discharged from the tank. From the viewpoint of resistance to the HAN liquid propellant, it is desirable that the gas-liquid separation mechanism such as the piston is also made of stainless steel or a polymer material.
The inner surface of the tank has a mirror finish. In addition, the tank has a safety factor of 2 times or more (4 times or more in the ground test), and can sufficiently withstand the pressure assumed on the orbit of the spacecraft with an internal pressure of about 3 MPa. It is desirable to have a breakdown voltage structure. When the liquid propellant is filled in the tank, a so-called vacuum filling method generally used as a liquid propellant filling method in the spacecraft can be adopted as follows. Since the HAN propellant can be stored for a long time, the liquid propellant can be stored in a state where the tank is filled.

b. Combustor The combustor 2 is preferably a stainless alloy system. The material of the combustor 2 may be selected as long as it is compatible with the high-temperature (1000 ° C. or higher) reaction gas and the strongly acidic unreacted liquid propellant. The shape of the combustor 2 may be a general pressure-resistant shape used in a space vehicle, and may be, for example, a cylindrical shape. The structure of the combustor 2 is such that the pressure of the reaction gas is less than 1 MPa, so that it can sufficiently withstand this, and more than twice the safety factor (more than four times the safety factor in ground tests) It is desirable to ensure.
As the thrust force obtained by the combustor 2, it is desirable to obtain about 15N. The magnitude of the obtained thrust force can be changed by adjusting the thruster shape.
The nozzle 4 of the combustor 2 may be of the type normally used in space vehicles. As for the nozzle 4, the reaction gas is choked at the throat of the nozzle 4, whereby the pressure of the reaction gas can be measured, and the thruster performance can be evaluated from the measurement result. A nozzle system that can be choked can be selected based on the evaluated thruster performance.
The process of obtaining the propulsive force by discharging the reactive gas from the nozzle 4 is that the reactive gas is choked at the throat of the nozzle 4, passes through the throat at the speed of sound, and then expands properly, thereby obtaining the propulsive force. Is.

c. Propellant Valve As the propellant valve 3 used in the present invention, it is desirable to select a propellant valve that has a sufficient track record for use in space vehicles. For example, a sliding fit type or a suspended armature type solenoid valve that has a stainless alloy structure and drives and seals the valve body by a solenoid can be preferably used. The propellant valve 3 has a role of supplying a liquid propellant to the combustor 2 in the space vehicle, and controls combustion by controlling the supply of the liquid propellant.
Specifically, the propellant valve 3 functions as follows. That is, the liquid propellant pressurized and pumped from the upstream tank is not supplied to the combustor 2 including the catalyst layer 5 because the propellant valve 3 is closed. Only when the generation of thrust by the catalytic decomposition thruster 1 is required, the propellant valve 3 is opened, the liquid propellant is supplied to the catalyst layer 5, and thrust is generated by catalytic decomposition. Thus, the propellant valve 3 is an element having a function of controlling combustion.
The combustor 2 used in the present invention further includes a catalyst layer 5 having a catalyst for catalytically decomposing the liquid propellant, a heating device 6 for heating the catalyst layer 5, and a liquid propellant for the catalyst layer 5. And an injector 7 for supplying the same.

d. Catalyst Layer The catalyst layer 5 is preferably configured so that the liquid propellant ejected from the injector 7 is applied to the entire catalyst layer 5 without excess or deficiency inside the combustor 2, and from the tip of the injector 7 to the catalyst layer 5. It is important to optimize and arrange the distance. For example, the catalyst layer 5 can be disposed at a position of about 50 mm from the tip of the injector 7.
The reaction gas of the liquid propellant decomposed in the catalyst layer 5 is completely decomposed in a region downstream of the catalyst layer 5 in the combustion chamber, and is discharged from the nozzle 4. Graphite is used as the heat-resistant material on the inner surface of the part involved in this process in the combustor 2.
As a catalyst used for the catalyst layer 5, for example, a catalyst of an alumina carrier supporting iridium such as S405 can be used. The type of the catalyst is not particularly limited as long as it is a catalyst usually used for a one-part propulsion system of a space vehicle. As the size and shape of the catalyst, it is preferable to use particles having a particle size of about 0.5 mm or less. The catalyst layer 5 can be configured by sandwiching the catalyst with a stainless steel wire mesh with a fine mesh to such an extent that such a catalyst does not spill. Further, for holding the catalyst layer 5, it is desirable to use a stainless alloy wire mesh as a heat-resistant and pressure-resistant material and to use carbon fiber and a ceramic material as a reinforcing material.
As a liquid propellant decomposed | disassembled by the catalyst of the catalyst layer 5, what mixed ammonium nitrate, water, and methanol with the mass ratio 95: 5: 8: 21 can be used for hydroxyammonium nitrate, for example. The liquid propellant having such a composition has a specific gravity of 1.4 and a freezing point of −68 ° C.
The mechanism of obtaining a driving force by decomposing such a HAN liquid propellant with the catalyst of the catalyst layer 5 is that the HAN liquid propellant is decomposed by the catalyst and decomposed into nitrogen, hydrogen, etc. It is explained that the fuel component methanol burns in the combustor 2 and is discharged from the nozzle during the reaction process.

e. Heating device There is no restriction | limiting in particular as the heating device 6, For example, the tape heater which processed the heating wire in tape shape can be used. Such a heater, for example, generates heat by applying a voltage of 24V to a 10Ω heating wire. When the tape heater is used, the heating device 6 can be applied by winding the tape heater around the outside of the combustor 2 near the catalyst layer 5. And the catalyst layer 5 is heated using the heat transfer of the combustor 2 made of stainless steel.

f. Injector It is desirable that the injector 7 is also made of a material having material compatibility and high temperature resistance such as stainless steel for the same reason as other elements. The injector 7 is not particularly limited as long as it can supply the liquid propellant to the catalyst layer 5, but is preferably an atomizer disposed upstream of the catalyst layer 5 in the combustor 2. In order for the liquid propellant sprayed from the injector 7 to be sprayed uniformly on the catalyst of the catalyst layer 5, it is considered that the spray diffusion angle is preferably 50 ° from the measurement.
The particle size of the liquid propellant droplet sprayed by the injector is preferably about 0.1 mm to 0.7 mm, and more preferably about 0.3 mm to 0.5 mm. In order to hold the powdered catalyst in the catalyst layer 5, the catalyst is usually sandwiched from both sides using a metal net. It is extremely difficult to spray the liquid propellant directly on the catalyst of the catalyst layer 5 having such a structure. On the other hand, a metal mesh for holding a catalyst used to fix a granular catalyst having a particle size of about 0.5 mm or less as described above is usually about 0.2 mm in wire diameter and about 40 mesh. Therefore, if the droplet is too larger than this size, it will collide with the metal wire constituting the wire mesh, making it difficult for the droplet to reach the catalyst. On the other hand, a droplet that is too small has a small inertial force and a moving speed is lowered, so that it becomes difficult for the liquid propellant to reach the catalyst layer 5 efficiently and the surface tension becomes very large. It becomes difficult to pass through and cause a large amount of unreacted liquid propellant to be formed. Large quantities of unreacted liquid propellant droplets become large lumps, resulting in a decrease in catalyst temperature, while also having the risk of causing an explosion by reacting suddenly. Furthermore, droplets that are too fine exhibit a sensitive reaction, and when the catalyst holding wire mesh that is heated at the same time as the catalyst is heated, it also reacts and melts the holding wire mesh. . Therefore, when the catalyst layer 5 of the type using a metal mesh for holding the catalyst is adopted, the droplet diameter of the sprayed liquid propellant is 70 to 120% of the mesh interval of the metal mesh. It is desirable to make it.
In the catalytic decomposition type thruster according to the present invention, the injector 7 can spray the liquid propellant onto the catalyst layer 5 by applying a predetermined pressure to the piston. At that time, it is desirable that the pressure applied to the piston be set to 0.75 MPa or more in consideration of the viscosity of the liquid propellant and the like. If the pressing pressure is too low, the liquid propellant is not sprayed and becomes a water stream. Further, the upper limit of the pressing pressure is considered to be about 4 MPa from the viewpoint of use in the space environment.
Furthermore, the catalytic decomposition type thruster of the present invention is provided with a heat insulating member 8 having a liquid propellant flow path from the propellant valve 3 to the combustor 2 between the injector 7 and the propellant valve 3. It is characterized by this.

g. Thermal insulating member As a material used for the thermal insulating member 8 used in the present invention, it is preferable to use a polymer having excellent thermal insulation and high heat resistance. Examples of such a resin include fluororesin, polyimide resin, polysulfone, polyphenylene sulfide, polyether ether ketone, and polyarylate. Among these, it is particularly preferable to use a fluororesin such as Teflon (registered trademark) or a polyimide resin such as Upilex (registered trademark) from the viewpoint of processability.
The shape of the heat insulating member 8 can be a cylindrical shape, and the outer diameter thereof is the outer diameter of the injector 7 and the propellant valve 3 to which the heat insulating member 8 is connected, or when these have flanges, the flanges. What is necessary is just to be the same as the outer diameter of. The thickness of the heat insulating member 8 is preferably as thick as possible from the viewpoint of heat insulation, but if it is too thick, the flow path of the liquid propellant provided inside the heat insulating member 8 becomes long, and the heat insulating member It should be noted that the weight of 8 increases. For example, as the heat insulating member 8 used for the combustor 2 having an outer diameter of about 45 mm, a member having an outer diameter of about 40 mm and a thickness of about 5 mm can be suitably used.
The diameter of the flow path of the liquid propellant provided inside the heat insulating member 8 is not particularly limited, but is preferably about 1 mm. The larger the channel diameter, the less likely the problem of blockage of the channel occurs, but it should be noted that if the channel diameter is too large, the liquid propellant may be dried in the middle of the channel.
When the heat insulating member 8 is installed, for example, a screw hole is cut on the injector 7 side, a hole through which a screw passes is provided on the heat insulating member 8, and screwed from the flange side of the propellant valve 3. A way to do this is conceivable.

A catalytic decomposition type thruster according to the present invention having a structure as shown in FIG. 2 was produced, and an experiment for confirming the effect of thermal insulation by the thermal insulation member 8 was conducted. As the heat insulating member 8, a Teflon (registered trademark) member having an outer diameter of 40 mm and a thickness of 5 mm was used. A liquid propellant flow channel having a diameter of 1 mm was provided in the central portion of the heat insulating member 8, and holes for screwing were formed at three locations near the outer periphery. A screw hole was cut in the stainless steel injector 7 side and screwed through the heat insulating member 8 from the flange side of the stainless steel propellant valve 3.
The performance of the heat insulating member 8 is obtained by heating the catalyst layer 5 to 250 ° C. with the heating device 6 and monitoring the change over time in the flange temperature of the propellant valve 3 while maintaining the temperature. evaluated. For comparison, a catalytic decomposition type thruster in which the injector 7 and the propellant valve 3 were connected without using the heat insulating member 8 was produced, and the same experiment was performed. The results are shown in Table 1.

  From Table 1, in the case of the catalytic decomposition type thruster of the present invention that employs a heat insulating member, the temperature of the propellant valve is about 50 ° C. even when a long time elapses, whereas the heat insulating member is not used. It can be seen that the temperature of the propellant valve far exceeds 80 ° C. in a short time. The heat resistance of the internal structure of the propellant valve is usually designed at about 70 ° C., confirming the effectiveness of the present invention.

1 Thruster 2 Combustor 3 Propellant valve 4 Nozzle 5 Catalyst layer 6 Heating device 7 Injector 8 Thermal insulating member

Claims (1)

A catalyst-decomposable thruster for space vehicles, which obtains propulsive force by injecting a reaction gas obtained by decomposing a liquid propellant with a catalyst,
A hollow tank for containing the liquid propellant;
A combustor that decomposes the liquid propellant and injects a reactive gas;
A propellant valve for controlling the supply of the liquid propellant from the tank to the combustor;
With
The combustor further includes
A catalyst layer having a catalyst for catalytically decomposing the liquid propellant;
A heating device for heating the catalyst layer;
An injector for supplying the liquid propellant to the catalyst layer;
The catalyst is characterized in that a thermal insulating member having a flow path of the liquid propellant from the propellant valve to the combustor is provided between the injector and the propellant valve. Decomposable thruster.

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