JP2024050975A - Hybrid Rocket - Google Patents

Abstract

To provide a small-sized and light-weight hybrid rocket.SOLUTION: A hybrid rocket comprises: an oxidant storage chamber that stores liquid or gaseous oxidant; a fuel gas generation region/combustion region in which a solid fuel containing a tetra-olglycidyl azide polymer is stored for end-burning, and which generates fuel gas by self-exothermic decomposition of the tetra-olglycidyl azide polymer, and mixes and burns the oxidant and the fuel gas; and an oxidant supply pipe that connects the oxidant storage chamber and the fuel gas generation region/combustion region through the outside of the fuel gas generation region/combustion region.SELECTED DRAWING: Figure 1

Description

Application for application of Article 30, Paragraph 2 of the Patent Act has been filed. ▲1▼AJCPP 2018 distribution materials AJCPP2018-001, Asian Joint Conference on Propulsion and Power Distribution date: March 14, 2018 Distribution location: City Hotel Xiamen (No. 16, Huyin Road, Siming District, Xiamen City, Fujian Province, China) ▲2▼AJCPP 2018 AJCPP2018-017 Development of GAP-based propellant thrusters for small satellites Date held: March 15, 2018 Location: City Hotel Xiamen (No. 16, Huyin Road, Siming District, Xiamen City, Fujian Province, China)

The present invention relates to a hybrid rocket.

Solid rockets and liquid rockets are known as propulsion devices for space transport vehicles and satellite thrusters. Solid rockets use solid fuel and solid oxidizer as propellants. Because the fuel and oxidizer are in the same phase, they have excellent combustion properties. In addition, the rocket structure is relatively simple. On the other hand, because the fuel used in solid rockets is what is known as gunpowder, handling during storage and transportation is complicated. For this reason, careful management is required during storage and transportation of solid fuel.

Liquid rockets use liquid fuel and liquid oxidizer as propellants. The liquid fuel and liquid oxidizer are mixed in a spray state and burned to generate thrust. Like solid rockets, the propellant in liquid rockets is of the same phase, making it highly combustible. However, the fuel and oxidizer must be kept stable at extremely low temperatures. In addition, because the fuel and oxidizer tend to mix, handling during storage and transportation is cumbersome. Furthermore, the structure of the rocket is more complex than that of a solid rocket, making it difficult to control the combustion of the propellant.

Hybrid rockets have therefore attracted attention. Hybrid rockets generally use solid fuel and liquid or gaseous oxidizer as propellants. Because the fuel and oxidizer are in different phases, they do not easily mix and burn, making it easier to manage the propellant than solid and liquid rockets. In particular, hybrid rockets that use polymers as solid fuel are easier to manage during storage and transportation. For example, Patent Document 1 discloses a hybrid rocket that has a liquid oxidizer storage chamber, a solid fuel combustion chamber, and a mixing combustion chamber that mixes and burns excess fuel component gas generated from the solid fuel with the liquid oxidizer.

Japanese Patent Application Laid-Open No. 7-19120

When using the above hybrid rocket as a thrust source for a space transport vehicle or a thruster for an artificial satellite, there is a demand for further miniaturization and weight reduction of the hybrid rocket from the viewpoint of reducing the weight of the entire space transport vehicle or artificial satellite device and improving energy loss.

The hybrid rocket using the polymer disclosed in Patent Document 1 as solid fuel is easy to manage because the propellants do not easily mix, but compared to solid and liquid rockets, the amount of fuel gas generated per unit mass of solid fuel is small. In order to increase the amount of fuel gas generated, the surface area of the solid fuel needs to be increased, making it difficult to miniaturize the solid fuel combustion chamber.

The present invention was made in consideration of the above circumstances, and aims to provide a hybrid rocket that is even smaller and lighter.

The inventors have discovered that it is possible to miniaturize hybrid rockets by using a fuel containing tetra-ol glycidyl azide polymer (also called tetra-ol GAP), which has self-heating decomposition properties, as a solid fuel.

The present invention has the following aspects.
[1] A hybrid rocket comprising: an oxidizer storage chamber for storing a liquid or gaseous oxidizer; a fuel gas generation chamber for storing a solid fuel containing a tetra-ol glycidyl azide polymer and generating a fuel gas by the self-exothermic decomposition of the tetra-ol glycidyl azide polymer; and a combustion chamber for mixing and combusting the oxidizer and the fuel gas.
[2] The hybrid rocket described in [1], further comprising a pressure regulating means for reducing the pressure inside the fuel gas production chamber.
[3] The hybrid rocket described in [2], wherein the pressure regulating means is a valve that reduces the amount of oxidizer supplied to the combustion chamber.
[4] The hybrid rocket described in [2], wherein the pressure regulating means is a valve for discharging the oxidizer and the fuel gas from the combustion chamber.
[5] The hybrid rocket according to any one of claims [1] to [4] further comprising a pressurizing device for pressurizing the oxidizer and supplying it to the combustion chamber to maintain the pressure in the combustion chamber at 700 kPa or more.
[6] The hybrid rocket described in [5], wherein the pressurizing device includes a fuel gas supply pipe for extracting a portion of the fuel gas from the fuel gas generation chamber, and a pressurizing section connected to the fuel gas supply pipe and pressurizing the oxidizer storage chamber with the fuel gas.
[7] A hybrid rocket according to any one of [1] to [3] and [5], wherein the fuel gas generation chamber and the combustion chamber are integrally formed.

The hybrid rocket of the present invention can be made even smaller and lighter.

FIG. 1 is a schematic diagram of a hybrid rocket according to one embodiment of the present invention. FIG. 2 is a schematic diagram of a hybrid rocket according to another embodiment of the present invention. 1 is a graph showing the relationship between pressure in a combustor and combustion time in Examples 2 to 5. 13 is a graph showing the relationship between pressure in a fuel gas production chamber/combustion chamber and combustion time in Example 6. 13 is a graph showing the relationship between combustion efficiency and pressure in a fuel gas production chamber/combustion chamber in Example 6.

The following describes a hybrid rocket according to one aspect of the present invention, with reference to the drawings. Preferred examples and conditions may be shared among the following multiple embodiments. Furthermore, the numbers, quantities, positions, shapes, etc. may be changed, omitted, or replaced without departing from the spirit of the present invention. In all of the following drawings, the dimensions and ratios of each component may be appropriately changed in order to make the drawings easier to read.

One aspect of the present invention is a hybrid rocket that includes an oxidizer storage chamber for storing a liquid or gaseous oxidizer, a fuel gas generation chamber for storing a solid fuel containing tetra-ol glycidyl azide polymer and generating fuel gas by the self-heating decomposition of the tetra-ol glycidyl azide polymer, and a combustion chamber for mixing and burning the oxidizer and the fuel gas.

First Embodiment
The hybrid rocket of this embodiment will be described below with reference to Fig. 1. Fig. 1 is a schematic diagram of the hybrid rocket of this embodiment. The hybrid rocket 1 has a body 2, a fuel gas generation chamber 3, a solid fuel 4 containing tetra-ol glycidyl azide polymer as fuel, an oxidizer storage chamber 5 in which a liquid or gaseous oxidizer 6 is stored, an oxidizer supply pipe 7, a valve 8, a fuel gas generation region and combustion region 9, a nozzle 10, an ignition device 11, and an oxidizer storage chamber pressurization device 12.

A fuel gas generation chamber 3 is formed inside the aircraft body 2. The fuel gas generation chamber 3 stores solid fuel 4 and has a fuel gas generation area and combustion area 9. The space surrounded by the end face of the solid fuel 4 and the inner wall of the fuel gas generation chamber 3 is the fuel gas generation area and combustion area 9. The volume of the solid fuel 4 decreases as the solid fuel 4 self-heating decomposes, and the volume of the fuel gas generation area and combustion area 9 expands. Furthermore, in this embodiment, an oxidizer 6 is supplied into the fuel gas generation area and combustion area 9, and the fuel gas and the oxidizer 6 are mixed and burned. In other words, the space in which the fuel gas is generated and the space in which the oxidizer 6 and the fuel gas are mixed and burned are shared.

The aircraft 2 is equipped with an ignition device 11 for igniting the solid fuel 4. The fuel gas generation and combustion area 9 is connected to a nozzle 10. In the fuel gas generation and combustion area 9, the fuel gas and the oxidizer 6 are mixed and burned, generating high-temperature, high-pressure gas, which is ejected from the nozzle 10 to generate the thrust of the hybrid rocket.

The oxidizer storage chamber 5 is connected to the fuel gas generation area/combustion area 9 via an oxidizer supply pipe 7. The oxidizer supply pipe 7 is provided with a valve 8 for controlling the supply of oxidizer 6 from the oxidizer storage chamber 5. The oxidizer storage chamber 5 has an oxidizer storage chamber pressurizing device 12 for pressurizing the inside of the oxidizer storage chamber 5. By pressurizing the inside of the oxidizer storage chamber 5 with the oxidizer storage chamber pressurizing device 12, the supply amount of oxidizer 6 can be adjusted.

1, the oxidant storage chamber 5 has an oxidant storage chamber pressurizing device 12 for pressurizing the inside of the oxidant storage chamber 5, but this embodiment is not limited to this. When a liquid oxidant having a high vapor pressure (e.g., nitrous oxide (N ₂ O) or the like) is used as the oxidant 6, the oxidant storage chamber pressurizing device 12 does not need to be provided.

Next, each component will be described in detail.
The airframe 2 is configured such that at least an outer wall, an intermediate layer, and an inner wall are laminated in this order. The outer wall of the airframe 2 is formed of a metal such as aluminum in order to maintain mechanical strength and to withstand pressure from the fuel gas generation region/combustion region 9. The intermediate layer is formed of glass fiber reinforced plastics (also called Glass Fiber Reinforcing Plastics, GFRP), phenolic resin, or a composite material thereof, in order to provide a fuel cartridge capable of insulating from a combustion flame. The inner wall is formed of a polymeric compound having heat resistance, and examples of such polymeric materials include ethylene propylene diene rubber (also called Ethylene Propylene Diene Rubber, EPDM rubber). When ethylene propylene diene rubber is used as the material of the inner wall, the terminal hydroxyl groups of the tetra-ol glycidyl azide polymer are bonded to a primer applied on the ethylene propylene diene rubber, thereby allowing the solid fuel 4 and the inner wall to be sufficiently bonded to each other. The primer may be a primer having an isocyanate group.

The fuel gas generation chamber 3 is a space surrounded by the inner wall of the aircraft body 2, and contains solid fuel 4 containing tetra-ol glycidyl azide polymer. The solid fuel 4 contains tetra-ol glycidyl azide polymer and a hardener. The tetra-ol glycidyl azide polymer undergoes self-exothermic decomposition under conditions of 250°C or higher and 700 kPa or higher (7 atm or higher).

Tetra-ol glycidyl azide polymer is obtained by urethane polymerization of the prepolymer represented by chemical formula (1).

Since the tetra-ol glycidyl azide polymer has four hydroxyl groups, urethane bonds are more likely to be formed three-dimensionally than the di-ol glycidyl azide polymer and the tri-ol glycidyl azide polymer. Therefore, in the synthesis of the tetra-ol glycidyl azide polymer, a crosslinking agent used to improve moldability during the synthesis of the di-ol glycidyl azide polymer and the tri-ol glycidyl azide polymer is not required. Furthermore, the amount of the curing agent used during the synthesis of the tetra-ol glycidyl azide polymer may be smaller than that of the glycidyl azide polymer and the tri-ol glycidyl azide polymer. In other words, the proportion of the tetra-ol glycidyl azide polymer contained in the solid fuel 4 is larger than that of the solid fuel containing the di-ol glycidyl azide polymer and the tri-ol glycidyl azide polymer. Specifically, the proportion of the tetra-ol glycidyl azide polymer contained in the solid fuel 4 is 85 to 95 mass% with respect to the total mass of the solid fuel 4.

The hardener contained in the solid fuel 4 may be hexamethylene diisocyanate (also called HMDI) or isophorone diisocyanate (also called IPDI). The ratio of the hardener contained in the solid fuel 4 is preferably 5 to 15 mass % of the total mass of the solid fuel 4, and more preferably 9 to 12 mass %.

The solid fuel 4 may contain a catalyst used in the synthesis of the tetra-ol glycidyl azide polymer. Examples of the catalyst include dibutyltin dilaurate (also called DBTDL). The proportion of the catalyst contained in the solid fuel 4 is an extremely small amount relative to the total mass of the solid fuel 4.

Since the proportion of tetra-ol glycidyl azide polymer contained in the solid fuel 4 is 85 to 95% by mass, the amount of fuel gas generated by the self-heating decomposition of the tetra-ol glycidyl azide polymer per unit mass of the solid fuel 4 increases. In other words, it is possible to reduce the size of the fuel gas generation chamber 3 and the solid fuel 4. In addition, the amount of residue generated after the combustion of the solid fuel 4 is small.

It is preferable that the solid fuel 4 containing the tetra-ol glycidyl azide polymer has mechanical properties sufficient to accommodate the expansion of the fuel gas generation chamber 3 that may occur during combustion of the solid fuel 4.

Hardness is one index that indicates mechanical properties. When the size of the solid fuel 4 is 60 mm in diameter and 10 to 80 mm in length, it is preferable that the value measured in accordance with JIS K 6253-3 using a Type A durometer (GS-706N Type A, manufactured by TECLOCK) is in the range of 30 to 40. When the solid fuel 4 is 60 mm in diameter and 10 to 80 mm in length, if the hardness is 30 to 40, it can follow the expansion of the fuel gas generation chamber 3 while maintaining its shape as a solid fuel.

In particular, when attempting to increase the size of the hybrid rocket 1, it is necessary to reduce the weight of the entire hybrid rocket, and therefore it is necessary to reduce the wall thickness of the fuselage 2. If the wall thickness of the fuselage 2 is thin, the fuel gas generation chamber 3 is more likely to expand when the solid fuel 4 is burned. If the solid fuel 4 has appropriate mechanical properties and excellent adhesive properties, the solid fuel 4 can follow the expansion of the fuel gas generation chamber 3, and close contact between the fuel gas generation chamber 3 and the solid fuel 4 can be maintained, which is preferable.

Furthermore, the solid fuel 4 is required to have mechanical properties according to its size, since it may deform due to gravity during storage and acceleration during the flight of the hybrid rocket. The solid fuel 4 containing the tetra-ol glycidyl azide polymer of this embodiment has an appropriate hardness without containing a crosslinking agent, as described above. Therefore, by slightly adjusting the amount of hardener, it is possible to adjust the mechanical properties of the solid fuel to a value suitable for the size of the device in which the solid fuel is filled, without compromising its performance as a solid fuel. In addition, since the solid fuel 4 containing the tetra-ol glycidyl azide polymer of this embodiment has hydroxyl groups, it is possible to achieve strong adhesion to the inner wall of the fuel gas generation chamber 3 using urethane bonds.

The hardness of solid fuel can be measured using the following method. The size of the solid fuel measurement sample is 2 cm long, 10 cm wide, and 1 cm thick, and measurements are made using a Type A durometer hardness tester. First, the indenter of the Type A durometer hardness tester is pressed against the surface of the solid fuel as quickly as possible without causing any impact, and the pressure reference surface is brought into close contact with the surface of the measurement sample. The maximum reading of the pointer of the indicating device is read within three seconds, and this is taken as the hardness value of the measurement sample. Measurements are made at any five points on the measurement sample, and the average value is taken as the hardness of the solid fuel.

The method of filling the fuel gas generation chamber 3 with the solid fuel 4 is, for example, to degas a prepolymer of tetra-ol glycidyl azide polymer while heating it at 50-60°C, and then mix it by adding a curing agent. After that, a catalyst is dripped and further mixed at 50-60°C, and before curing, the fuel gas generation chamber 3 is filled and cured in the fuel gas generation chamber 3. By filling the fuel gas generation chamber 3 with the solid fuel in this way, when ethylene propylene diene rubber is used as the material for the inner wall of the fuel gas generation chamber 3, the terminal hydroxyl groups of the tetra-ol glycidyl azide polymer bond with the primer applied to the ethylene propylene diene rubber, and the solid fuel 4 can be sufficiently bonded to the inner wall.

The oxidizer storage chamber 5 is made of a corrosion-resistant material, such as aluminum or stainless steel.

After opening the valve 8, the oxidizer storage chamber pressurizing device 12 pressurizes the oxidizer storage chamber 5, thereby supplying the oxidizer 6 to the fuel gas generation region and combustion region 9 via the oxidizer supply pipe 7 and pressurizing the fuel gas generation region and combustion region 9. Examples of the oxidizer storage chamber pressurizing device 12 include a piston with a movable partition, an accumulator, etc.

The oxidizer 6 contained in the oxidizer storage chamber 5 may be a liquid oxidizer or a gaseous oxidizer, but a liquid oxidizer is preferable because of its small storage volume. Examples of liquid oxidizers include nitrous oxide, liquid oxygen, hydrogen peroxide, and an aqueous solution of hydroxylamine nitrate. Examples of gaseous oxidizers include oxygen.

The oxidizer supply pipe 7 connects the oxidizer storage chamber 5 with the fuel gas generation area/combustion area 9. The oxidizer supply pipe 7 is made of a corrosion-resistant material that is listed as a material for the oxidizer storage chamber 5. The valve 8 is a pressure adjustment means for the fuel gas generation chamber 3, and opens and closes the oxidizer supply pipe 7 to control the supply of the oxidizer 6.

The fuel gas generation and combustion area 9 is formed integrally with the fuel gas generation chamber 3 that contains the fuel as described above. The fuel gas generation and combustion area 9 is connected to the nozzle 10 through a small diameter hole 20.

The nozzle 10 may be a Laval nozzle or the like, but may have other shapes. The nozzle 10 reaches a temperature of 2000°C or higher because gas generated by the combustion of the fuel gas and the oxidizer 6 is discharged from the nozzle 10. For this reason, the nozzle 10 is made of a heat-resistant material such as graphite or a carbon composite material.

The ignition device 11 ignites the solid fuel 4 by controlling the pressure and temperature in the fuel gas generation and combustion region 9. In this embodiment, the tetra-ol glycidyl azide polymer contained in the solid fuel 4 is ignited by setting the temperature at or above 250°C and the pressure at or above 700 kPa. Therefore, the ignition device 11 has a mechanism for setting the temperature at or above 250°C and the pressure at or above 700 kPa in the fuel gas generation and combustion region 9. The configuration of the ignition device is not particularly limited, and for example, an ignition device that uses an aluminum-less solid propellant can be used.

The ignition device 11 has a mechanism that can re-ignite solid fuel 4 whose combustion has stopped. Specifically, the ignition mechanism does not operate during the first combustion, and ignites the tetra-ol glycidyl azide polymer by pressurizing and heating the fuel gas generation and combustion area 9 to 250°C or higher and 700 kPa or higher, which are the ignition conditions for tetra-ol glycidyl azide polymer. When re-igniting solid fuel 4 whose combustion has stopped, it is ignited using the ignition mechanism. Note that the pressurization and heating configuration is not particularly limited, but a gas torch or the like can be used, for example.

Next, the propulsion system of the hybrid rocket 1 in this embodiment will be described. In the hybrid rocket 1, the ignition device 11 first uses the fuel gas generation and combustion region 9 to raise the temperature to 250°C or higher and the pressure to 700 kPa or higher, thereby igniting the solid fuel 4. The ignited solid fuel 4 burns, and fuel gas is generated in the fuel gas generation and combustion region 9 due to the self-heating decomposition of the tetra-ol glycidyl azide polymer.

Then, the oxidizer storage chamber 5 is pressurized by the oxidizer storage chamber pressurizer 12, and the oxidizer 6 is supplied from the oxidizer storage chamber 5 to the fuel gas generation region and combustion region 9 via the oxidizer supply pipe 7. As the oxidizer 6 continues to be supplied from the oxidizer storage chamber 5, the pressure in the fuel gas generation region and combustion region 9 is maintained at 700 kPa or higher, and the self-heating decomposition of the tetra-ol glycidyl azide polymer continues. In the fuel gas generation region and combustion region 9, the fuel gas and the oxidizer 6 are mixed and burned, and the resulting high-temperature, high-pressure gas is ejected from the nozzle 10 to obtain the thrust of the hybrid rocket 1.

The thrust of the hybrid rocket 1 can be stopped by stopping the operation of the oxidizer storage chamber pressurizer 12 and closing the valve 8 to reduce or stop the supply of oxidizer 6 to the fuel gas production area and combustion area 9. As a result, the pressure in the fuel gas production area and combustion area 9 is reduced to 700 kPa or less, the self-heating decomposition of the tetra-ol glycidyl azide polymer stops, and the mixing and combustion of the fuel gas and oxidizer 6 stops.

When a liquid oxidizer with a high vapor pressure is used as the oxidizer 6, the supply of the oxidizer 6 to the fuel gas generation and combustion area 9 can be reduced or stopped by closing the valve 8.

To regain the thrust of the hybrid rocket 1, the following operations are performed. First, the ignition device 11 is used to again raise the temperature in the fuel gas generation and combustion area 9 to 250°C or higher and the pressure to 700 kPa or higher, and reignite the solid fuel 4. The subsequent steps are the same as those for regaining the thrust of the hybrid rocket 1 described above.

In the hybrid rocket 1 of this embodiment, since the amount of fuel gas generated per unit time from the solid fuel 4 is large, a sufficient amount of fuel gas can be generated without increasing the surface area of the solid fuel 4, in other words, the burning surface area of the solid fuel 4. In other words, the fuel filling rate can be improved. This makes it possible to reduce the volume of the solid fuel 4 compared to solid fuels that use conventional polymers.

Since the amount of fuel gas generated per unit mass of the solid fuel 4 is large, the fuel gas generation area and the combustion area can be formed as one unit, as in this embodiment. In other words, the space in which fuel gas is generated by the self-heating decomposition of the tetra-ol glycidyl azide polymer contained in the solid fuel 4 in the fuel gas generation chamber 3 and the space in which the oxidizer 6 and fuel gas mix and burn are shared. This makes it possible to further miniaturize the hybrid rocket 1.

Since the amount of fuel gas generated per unit mass of the solid fuel 4 is large, it is possible to generate a sufficient amount of fuel gas by end-burning the solid fuel 4. End-burning refers to burning only the surface of the solid fuel 4 exposed to the fuel gas generation and combustion region 9. By end-burning the solid fuel 4, the amount of fuel gas generated over time can be kept roughly constant. In other words, the output of high-temperature, high-pressure gas generated by mixing and burning the fuel gas and oxidizer 6 is stabilized, so the thrust of the hybrid rocket 1 can be maintained constant.

The solid fuel 4 contains tetra-ol glycidyl azide polymer, so it has sufficient hardness, i.e., structural strength, even without containing a crosslinking agent. Furthermore, the adhesive strength between the inner wall of the fuel gas generation chamber 3 and the solid fuel 4 is high. As a result, while the solid fuel 4 is being burned and while the hybrid rocket 1 is receiving its thrust, the solid fuel 4 does not separate from the fuel gas generation chamber 3, and the surface area from which fuel gas is generated in the solid fuel 4 can be kept constant. This makes it possible to keep the thrust constant. If the adhesive strength between the inner wall of the fuel gas generation chamber and the solid fuel 4 is weak, a gap will be generated at the interface between the inner wall of the fuel storage unit and the solid fuel while the solid fuel is being burned and while the hybrid rocket is receiving its thrust, and the amount of fuel gas generated may increase suddenly due to the increased combustion area.

The advantages of using the hybrid rocket 1 of this embodiment as a space probe propulsion device or a thruster for an artificial satellite can be explained as follows. For example, in the case of a small artificial satellite, it is often launched together with the rocket that launches the large satellite that is the main satellite. Therefore, it will rotate in the same orbit as the main satellite, but in order to move to another target orbit, a rocket that acts as a thruster must be provided. In order to place a rocket in a limited space, it is required to be small. When using a solid rocket, the fuel is explosive, so the stability of the fuel is low and its handling is complicated, which may affect the main satellite, and it may be necessary to avoid launching both. When using a liquid rocket, it is difficult to miniaturize it because the fuel and oxidizer may leak and the combustion mechanism is complicated.

On the other hand, the hybrid rocket of this embodiment is easy to handle because it uses tetra-ol glycidyl azide polymer as solid fuel, and is small because of its high fuel packing density. Therefore, the hybrid rocket of this embodiment can be used as a thruster for small artificial satellites, even when launched together with large satellites.

In addition, for purposes such as changing the orbit of a space transport vehicle or artificial satellite, it is necessary to temporarily stop the rocket's thrust and regain thrust by burning fuel and oxidizer again.

In rockets that use typical solid fuel, once the solid fuel has ignited, it cannot be stopped from burning until it has burned out. In order to temporarily stop the rocket's thrust, it is necessary to provide multiple independent solid fuel storage chambers. However, as mentioned above, there is a demand for rockets to be compact and lightweight, so providing multiple solid fuel storage chambers is not desirable.

The hybrid rocket 1 of this embodiment can stop the combustion of the solid fuel 4 with a relatively simple configuration of only stopping the oxidizer storage chamber pressurizer 12 and controlling the valve 8. Furthermore, when a liquid oxidizer with a high vapor pressure is used as the oxidizer 6, the combustion of the solid fuel 4 can be stopped with a relatively simple configuration of only controlling the valve 8. The ability to stop and restart the combustion makes it possible to control changes in the orbit of a space transport vehicle or artificial satellite.

Furthermore, since the hybrid rocket of this embodiment can repeatedly stop and restart combustion, it generates fuel gas from one fuel and one fuel gas generation chamber, and by combusting the oxidizer and fuel gas, it is possible to repeatedly eject the same amount of thrust from the same location. Therefore, compared to a case where multiple fuels and fuel gas generation chambers are provided in multiple locations, it is easier to control the course to reach the desired orbit.

Second Embodiment
The hybrid rocket of this embodiment will be described below with reference to Fig. 2. Note that the description of the configuration common to the first embodiment will be omitted.

Figure 2 is a schematic diagram of a hybrid rocket of the second embodiment. The hybrid rocket 1' of this embodiment has an aircraft 2, a fuel gas generation chamber 3, a solid fuel 4 containing tetra-ol glycidyl azide polymer as fuel, an oxidizer supply pipe 7, a nozzle 10, an ignition device 11, a fuel gas generation region 13, a combustion chamber 14, a first fuel gas supply pipe 15, a second fuel gas supply pipe 16, an oxidizer storage chamber 17 containing a liquid or gaseous oxidizer 6, a pressurizing section 18, and a combustion stop valve 19.

Inside the aircraft body 2, a fuel gas generation chamber 3 and a combustion chamber 14 are formed. The fuel gas generation chamber 3 stores solid fuel 4 and has a fuel gas generation area 13. The space surrounded by the end face of the solid fuel 4 and the inner wall of the fuel gas generation chamber 3 is the fuel gas generation area 13. The volume of the solid fuel 4 decreases as the solid fuel 4 self-heating decomposes, and the volume of the fuel gas generation area 13 increases. The fuel gas generation chamber 3 has a combustion stop valve 19, which is a pressure adjustment means for reducing the pressure inside the fuel gas generation chamber 3. The combustion chamber 14 is connected to the fuel gas generation area 13 via a first fuel gas supply pipe 15. Fuel gas and oxidizer 6 are supplied to the combustion chamber 14, and the fuel gas and oxidizer 6 are mixed and burned.

The aircraft 2 is equipped with an ignition device 11 for igniting the solid fuel 4. The combustion chamber 14 is connected to the nozzle 10. The high-temperature, high-pressure gas generated by mixing and burning the fuel gas and the oxidizer 6 in the combustion chamber 14 is ejected from the nozzle 10 to generate thrust for the hybrid rocket 1'.

The oxidizer storage chamber 17 is connected to the combustion chamber 14 via the oxidizer supply pipe 7. The oxidizer storage chamber 17 is equipped with an oxidizer storage chamber pressurizing device for pressurizing the inside of the oxidizer storage chamber 17. The oxidizer storage chamber pressurizing device is composed of the fuel gas generation chamber 3, the second fuel gas supply pipe 16, and the pressurizing unit 18.

Next, each component will be described in detail.
The configuration of the aircraft 2 is the same as that of the aircraft described in the first embodiment.
The fuel gas production chamber 3 is the same as the fuel gas production chamber 3 described in the first embodiment, except that it is formed integrally with the fuel gas production region 13 .
The solid fuel 4 is the same as the solid fuel described in the first embodiment.

The oxidizer storage chamber 17 is the same as the oxidizer storage chamber 5 described in the first embodiment, except that it is equipped with an oxidizer storage chamber pressurizing device.

The oxidizer storage chamber pressurizing device has a mechanism for pressurizing the oxidizer storage chamber 17 by operating the pressurizing unit 18 when a portion of the combustion gas generated in the fuel gas generation chamber 3 passes through the second fuel gas supply pipe 16. By pressurizing the oxidizer storage chamber 17 with the pressurizing unit 18, the oxidizer 6 is supplied to the combustion chamber 14 via the oxidizer supply pipe 7, and the combustion chamber 14 is pressurized. As in the first embodiment, the pressurizing unit 18 can be, for example, a piston with a movable partition, an accumulator, or the like.

By providing the oxidizer storage chamber pressurizing device of this embodiment, it is not necessary to provide a separate driving means for the oxidizer storage chamber pressurizing means, so the hybrid rocket 1' can be made smaller.

The oxidizing agent 6 may be the same as the oxidizing agent described in the first embodiment.
The oxidizer supply pipe 7 is the same as the oxidizer supply pipe described in the first embodiment, except that the valve 8 is not provided.

The fuel gas generation region 13 is connected to the combustion chamber 14 via the first fuel gas supply pipe 15. In the fuel gas generation region 13, a portion of the fuel gas generated by the self-heating decomposition of the tetra-ol glycidyl azide polymer is supplied to the combustion chamber 14 through the first fuel gas supply pipe 15. At the same time, another portion of the fuel gas operates the pressurizing unit 18 through the second fuel gas supply pipe 16. This causes the oxidizer 6 to be supplied to the combustion chamber 14. In other words, the self-heating decomposition of the tetra-ol glycidyl azide polymer in the fuel gas generation region 13 serves as a trigger to start mixing and burning the fuel gas and the oxidizer 6 in the combustion chamber 14.

The fuel gas generation chamber 3 is provided with a combustion stop valve 19 for stopping the self-heating decomposition of the tetra-ol glycidyl azide polymer. The combustion stop valve 19 is a pressure adjustment means for the fuel gas generation chamber 3. By opening the combustion stop valve 19, the fuel gas in the fuel gas generation chamber 3 is discharged all at once, the pressure in the fuel gas generation chamber 3 is reduced, and the self-heating decomposition of the tetra-ol glycidyl azide polymer can be stopped.

The combustion chamber 14 is connected to the nozzle 10 through a small diameter hole 20. The nozzle 10 may be the same as that in the first embodiment.
The ignition device 11 may be the same as that in the first embodiment.

Next, the propulsion system of the hybrid rocket 1' in this embodiment will be described. In the hybrid rocket 1', the solid fuel 4 is ignited by first using the ignition device 11 to raise the temperature in the fuel gas generation region 13 to 250°C or higher and the pressure to 700 kPa or higher. The ignited solid fuel 4 burns, and fuel gas is generated in the fuel gas generation region 13 by the self-heating decomposition of the tetra-ol glycidyl azide polymer.

As a result, a portion of the fuel gas is supplied to the combustion chamber 14 through the first fuel gas supply pipe 15. At the same time, the other portion of the fuel gas operates the pressurizing unit 18 through the second fuel gas supply pipe 16. This causes the oxidizer 6 to be supplied to the combustion chamber 14. In other words, the self-heating decomposition of the tetra-ol glycidyl azide polymer in the fuel gas generation region 13 is used as a trigger to start the mixing and combustion of the fuel gas and the oxidizer 6 in the combustion chamber 14. The high-temperature and high-pressure gas generated by the mixing and combustion of the fuel gas and the oxidizer 6 in the combustion chamber 14 is ejected from the nozzle 10, and the thrust of the hybrid rocket 1' can be obtained.

The propulsive force of the hybrid rocket 1' can be stopped by opening the combustion stop valve 19 and discharging the fuel gas in the fuel gas generation chamber 3 all at once. As a result, the pressure in the fuel gas generation region 13 falls to 700 kPa or less, the self-heating decomposition of the tetra-ol glycidyl azide polymer stops, and the supply of fuel gas to the combustion chamber 14 stops. At the same time, the operation of the oxidizer storage chamber pressurizer also stops, so the supply of oxidizer 6 to the combustion chamber 14 also stops, and the mixing and combustion of the fuel gas and oxidizer 6 stops.

To regain thrust for the hybrid rocket 1', the following operations are performed. First, the ignition device 11 is used to again raise the temperature in the fuel gas generation and combustion area 9 to 250°C or higher and the pressure to 700 kPa or higher, and reignite the solid fuel 4. The subsequent steps are the same as those for regaining thrust for the hybrid rocket 1' described above.

In the hybrid rocket 1' of this embodiment, the effect of using solid fuel containing tetra-ol glycidyl azide polymer is the same as that described in the first embodiment.

One of the differences between the hybrid rocket 1' in this embodiment and the hybrid rocket 1 in the first embodiment is the mechanism of the oxidizer storage chamber pressurizing device. The oxidizer storage chamber pressurizing device can trigger the self-heating decomposition of the tetra-ol glycidyl azide polymer in the fuel gas generation region 13 to start mixing and combustion of the fuel gas and oxidizer 6 in the combustion chamber 14. Therefore, there is no need to provide a separate driving means for the oxidizer storage chamber pressurizing means, and the hybrid rocket 1' can be made smaller.

Another difference between the hybrid rocket 1' in this embodiment and the hybrid rocket 1 in the first embodiment is the combustion stop valve 19 provided in the fuel gas generation chamber 3. The combustion of the solid fuel 4 can be stopped by a relatively simple configuration that only requires control of the combustion stop valve 19. The ability to stop combustion and restart combustion makes it possible to control changes in the orbit of a space transport vehicle or artificial satellite.

The present invention will be described in detail below with reference to examples, but the present invention is not limited to the following description.

(Evaluation method)
<Measurement of hardness of solid fuel>
The hardness of the solid fuel was measured in accordance with JIS K 6253-3 using a type A durometer hardness tester (manufactured by TECLOCK, GS-706N Type A). The size of the measurement sample of the solid fuel was 2 cm long, 10 cm wide, and 1 cm thick. The indenter of the type A durometer hardness tester was pressed against the surface of the measurement sample as quickly as possible without shock, and the pressure reference surface was brought into close contact with the surface of the measurement sample. The maximum indicated value of the pointer of the indicating device was read within 3 seconds, and this was taken as the hardness value of the measurement sample. Measurements were performed at any five points on the measurement sample, and the average value was calculated to be the hardness of the solid fuel.

Example 1
Preparation of solid fuel containing tetra-ol glycidyl azide polymer
90 parts by mass of a prepolymer of tetra-ol glycidyl azide polymer was degassed while being heated at 50° C., and then mixed with 9.1 parts by mass of hexamethylene diisocyanate as a curing agent. A small amount of dibutyltin dilaurate was dropped as a catalyst to cure the mixture, thereby obtaining a solid fuel.

The hardness of the solid fuel of Example 1 was measured by the method described above. Measurements were performed at five random locations on the solid fuel measurement sample, and the hardnesses at the five locations were 38, 38, 36, 38, and 37, respectively, with the average being 37.4.

From the above, the solid fuel of Example 1 had sufficient hardness even without the inclusion of a crosslinking agent. As a result, it was found that the content of tetra-ol glycidyl azide polymer in the solid fuel could be increased to 90 mass%.

Example 2
<Solid fuel combustion experiment in a combustor>
A 10 mm long solid fuel was prepared and loaded into a 60 mm diameter combustor by the same method as in Example 1. The combustor was sealed, and the solid fuel was ignited using an ignition device using an aluminum-less solid propellant, with the ignition device pressure set to about 1 MPa and the theoretical adiabatic flame temperature set to 2000°C, and the solid fuel was end-burned in the fuel gas generation chamber in the combustor. The design pressure of the fuel gas generation chamber during combustion was 5 MPa.

(Examples 3 to 5)
The same operation as in Example 2 was carried out, except that the solid fuel was packed in a combustor having a diameter of 60 mm so that the lengths thereof were 30 mm, 50 mm, and 80 mm. These were designated as Examples 3 to 5, respectively.
A graph showing the relationship between the pressure inside the combustor (hereinafter also referred to as the combustion pressure) and the combustion time is shown in FIG.

As shown in Figure 3, good ignition characteristics and stable combustion pressure curves were obtained in all of Examples 3 to 5. An increase in combustion time was observed with the extension of the solid fuel length in all of Examples 3 to 5. Furthermore, as shown in Table 1, the average combustion pressure P _c,ave in Examples 1 to 4 was below the design pressure of 5 MPa. The temperature coefficient of the burning rate obtained by the following formula (1) was 0.05 K ^-1 . π _k is defined as the temperature coefficient at equilibrium pressure.
σ _p =(1-n)π _k ... (1)

Example 6
<Hybrid rocket combustion test>
Of the hybrid rockets shown in FIG. 1, a test hybrid rocket with the nozzle removed was produced by the following method. A solid fuel with a length of 80 mm was produced and filled in a stainless steel fuselage with an inner diameter of 60 mm and a length of 80 mm by the same method as in Example 1. The part of the fuselage where the fuel was filled was the fuel storage section, and the space surrounded by the end face of the fuel and the inner wall of the storage vessel was the fuel gas generation chamber/combustion chamber. The solid fuel was ignited by using an ignition device using an aluminum-less solid propellant, with the ignition device pressure set to about 1 MPa and the theoretical adiabatic flame temperature set to 2000°C. After fuel gas was generated by the self-heating decomposition of the solid fuel, the pressure in the fuel gas generation chamber/combustion chamber was controlled to 1 MPa by introducing gaseous oxygen into the fuel gas generation chamber/combustion chamber.

The results of the pressure and combustion time in the fuel gas generation chamber/combustion chamber during solid fuel combustion are shown in Figure 4. In Figure 4, _Pc indicates the pressure in the fuel gas generation chamber/combustion chamber, and _Pox indicates the pressure in the oxidizer storage chamber. As shown in Figure 4, good ignition characteristics and a stable pressure curve during combustion were obtained. Figure 5 is a graph showing the relationship between the combustion efficiency C ^* and the pressure in the fuel gas generation chamber/combustion chamber in this embodiment. In Figure 5, ◯ indicates the combustion efficiency with only solid fuel, and ● indicates the combustion efficiency with a mixture of solid fuel and gaseous oxygen.

The combustion efficiency C ^* is calculated by the following formula (2).
C ^* (combustion efficiency) = C ^* (experimental value) / C ^* (theoretical value) (2)

C ^* (theoretical value) can be obtained by using NASA's CEA program (thermochemical equilibrium calculation program). C ^* (experimental value) can be obtained by the following formula (3).
C ^* (experimental value)=A _t (nozzle cross-sectional area)×P _c (combustion pressure)/m (combustion gas mass per unit time exiting from the nozzle) (3)

In formula (3), _Pc is the average value of the combustion pressure from 1 second to 5 seconds after the start of combustion, and m is the sum of the mass difference before and after the combustion of the solid fuel divided by the combustion time and the amount of oxygen supplied per unit time.
The combustion efficiency C ^* in the mixed state of solid fuel and gaseous oxygen was 93%.

1, 1'... hybrid rocket, 2... aircraft body, 3... fuel gas generation chamber, 4... solid fuel, 5, 17... oxidizer storage chamber, 6... oxidizer, 7... oxidizer supply pipe, 8... valve, 9... fuel gas generation area and combustion area, 10... nozzle, 11... ignition device, 12... oxidizer storage chamber pressurization device, 13... fuel gas generation area, 14... combustion chamber, 15... first fuel gas supply pipe, 16... second fuel gas supply pipe, 18... pressurization section, 19... combustion stop valve, 20... hole.

Claims (4)

an oxidizer storage chamber for storing a liquid or gaseous oxidizer;
a fuel gas generation region and combustion region in which a solid fuel containing a tetra-ol glycidyl azide polymer is accommodated so as to be burned at its end surface, in which a fuel gas is generated by the self-exothermic decomposition of the tetra-ol glycidyl azide polymer, and in which the oxidizer and the fuel gas are mixed and burned;
A hybrid rocket including an oxidizer supply pipe that connects the oxidizer storage chamber and the fuel gas generation region and combustion region through an outside of the fuel gas generation region and combustion region. The hybrid rocket of claim 1 further includes a pressure adjustment means for reducing the pressure in the fuel gas generation and combustion region. The hybrid rocket according to claim 2, wherein the pressure adjustment means is a valve that reduces the amount of oxidizer supplied to the fuel gas generation and combustion region. The hybrid rocket according to any one of claims 1 to 3, wherein the oxidizer supply pipe is made of aluminum or stainless steel.

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