Ignition device for hybrid rocket engine, method for igniting hybrid rocket engine, and rocket propulsion system

Abstract

The present invention provides a hybrid rocket engine ignition device, a hybrid rocket engine ignition method, and a rocket propulsion system that can be reliably reignited, miniaturized and lightweight, and mounted on relatively small satellite systems. [Solution] A high-temperature gas generating means is provided to generate high-temperature gas using rocket propellant. The system is configured to ignite a hybrid rocket engine using the high-temperature gas generated by the high-temperature gas generating means.

Description

Technical Field

[0001] The present invention relates to an ignition device for a hybrid rocket engine, a method for igniting a hybrid rocket engine, and a rocket propulsion system.

Background Art

[0002] Rocket engines include solid-propellant rocket engines that use solid propellants (oxidizers and fuels), liquid-propellant rocket engines that use liquid propellants, and hybrid rocket engines that use liquid or gaseous oxidizers and solid fuels. Solid-propellant rocket engines have a relatively simple structure, are easy to miniaturize, and fuel storage is easy, but combustion control is difficult. Liquid-fuel rocket engines are easier to control combustion than solid-propellant rockets, but fuel storage is difficult, the risk during explosion is high, and the launch process is also complex. Hybrid rocket engines have the advantages of both, and in particular, they have the characteristic of high safety because the fuel can be stored in a solid phase while the oxidizer can be stored individually.

[0003] Among these, regarding hybrid rocket engines with high safety, in recent years, the development of ignition devices has been actively promoted. As ignition devices for hybrid rocket engines, for example, those using arc discharge (see, for example, Non-Patent Document 1), those using a diode laser for a PMMA (polymethyl methacrylate resin) / GOX (gaseous oxygen) hybrid rocket (see, for example, Non-Patent Document 2), those heating a glow plug to burn the vaporized fuel and using the combustion for ignition (see, for example, Patent Document 1 or Non-Patent Document 3), etc. have been developed.

[0004] Furthermore, a 0.5N class propulsion system utilizing a catalytic reaction with a monopropellant composed of low-toxicity HNP has been developed for use in small satellite systems and the like (see, for example, Non-Patent Document 4). The HNP propellant is composed of HAN (Hydroxylammonium nitrate), HN (Hydrazine nitrate), methanol, and water. In addition, HN is less toxic than existing hydrazine (anhydrous hydrazine). [Prior art documents] [Non-patent literature]

[0005] [Non-Patent Document 1] SA Whitmore and MA Bulcher, “Vacuum Test of a Novel Green-Propellant Thruster for Small Spacecraft”, 53rd AIAA / SAE / ASEE Joint Propulsion Conference, 2017, <http: / / arc.aiaa.org | DOI: 10.2514 / 6.2017-5044> [Non-Patent Document 2] David M. Dyrda et al., “Diode Laser Ignition Testing for PMMA / GOX Hybrid Motors”, AIAA Propulsion and Energy Forum, 2019, 〈http: / / arc.aiaa.org | DOI: 10.2514 / 6.2019-4095〉 [Non-Patent Document 3] S. Hirai et al., “Development of safe, low-cost, re-ignitable rocket ignition system”, AIAA Propulsion and Energy Forum, 2021, 〈http: / / arc.aiaa.org | DOI: 10.2514 / 6.2021-3509〉 [Non-Patent Document 4] S. Igarashi and Y. Matsuura, “0.5N Thruster for Small Satellite Using HAN / HN-Based “Safety-Focused” Green Monopropellant”, AIAA Propulsion and Energy Forum, 2021, 〈http: / / doi.org / 10.2514 / 6.2021-3563〉 [Patent Documents]

[0006] [Patent Document 1] Japanese Patent Publication No. 2022-83773 [Overview of the project] [Problems that the invention aims to solve]

[0007] The ignition device for a hybrid rocket engine utilizing arc discharge, as described in Non-Patent Document 1, requires a voltage of 1000 V to generate the arc discharge for ignition, resulting in a relatively large and heavy overall device. Therefore, while it can be mounted on large satellite systems, it presents a challenge in mounting it on relatively small satellite systems, such as those weighing several hundred kg or less.

[0008] Furthermore, the diode laser ignition device described in Non-Patent Literature 2 also faced challenges in its suitability for use in small satellite systems, as it was complex, large, and heavy due to limitations on the placement of the laser source, the transparent window through which the laser passes, and the fuel. Additionally, the window would fog up due to combustion gases, hindering laser penetration and sometimes preventing re-ignition. Similarly, the glow plug ignition device described in Non-Patent Literature 3 required heating the glow plug to over 600°C, resulting in a large and heavy overall device due to the high-voltage power supply required, making it unsuitable for use in small satellite systems. Thus, conventional hybrid rocket engine ignition devices lacked sufficient size and lightness for use in small satellite systems.

[0009] This invention addresses these challenges and aims to provide a hybrid rocket engine ignition device, a hybrid rocket engine ignition method, and a rocket propulsion system that can be reliably reignited, miniaturized and lightweight, and mounted on relatively small satellite systems. [Means for solving the problem]

[0010] To achieve the above objective, the ignition device for a hybrid rocket engine according to the present invention is characterized by having a high-temperature gas generating means that generates high-temperature gas using rocket propellant, and is configured to ignite a hybrid rocket engine using the high-temperature gas generated by the high-temperature gas generating means.

[0011] The hybrid rocket engine ignition device according to the present invention ignites the hybrid rocket engine using high-temperature gas generated by the rocket propellant, and therefore does not require high voltage or high power, resulting in a relatively simple configuration. This allows for a relatively small high-temperature gas generation mechanism, enabling miniaturization and weight reduction. Furthermore, this makes it possible to mount the device on relatively small satellite systems, for example, weighing 500 kg or less.

[0012] The hybrid rocket engine ignition device according to the present invention generates high-temperature gas using rocket propellant, allowing for stable ignition of the hybrid rocket engine even in outer space, thus providing high reliability in ignition performance. Furthermore, repeated ignition (re-ignition) can be reliably performed. Therefore, it can be mounted on relatively small satellite systems and used for atmospheric re-entry or interplanetary flight.

[0013] In the ignition device for a hybrid rocket engine according to the present invention, the high-temperature gas generating means may be any device that generates high-temperature gas using rocket propellant, such as an existing thruster.

[0014] The ignition device for a hybrid rocket engine according to the present invention may ignite the hybrid rocket engine directly from the high-temperature gas generating means, or it may ignite the hybrid rocket engine via another ignition means from the high-temperature gas generating means. When using another ignition means, for example, the device may have a combustion ignition means that generates combustion gas by burning fuel with the high-temperature gas generated by the high-temperature gas generating means, and is configured to ignite the hybrid rocket engine with the combustion gas. In this case, the high-temperature gas generated by the high-temperature gas generating means can be kept to a low energy level that cannot ignite the hybrid rocket engine, and the high-temperature gas generating means can be made smaller and lighter. For example, as the high-temperature gas generating means, a thruster with a thrust of about 1 / 500 to 1 / 100 of the thrust of the hybrid rocket engine to be ignited can be used.

[0015] The hybrid rocket engine ignition method according to the present invention is characterized by generating high-temperature gas using rocket propellant and using that high-temperature gas to ignite the hybrid rocket engine.

[0016] The hybrid rocket engine ignition method according to the present invention can be suitably implemented using the hybrid rocket engine ignition device according to the present invention. Since the hybrid rocket engine ignition method according to the present invention utilizes high-temperature gas generated by the rocket propellant to ignite the hybrid rocket engine, it can be implemented with a relatively simple configuration that does not require high voltage or high power. Therefore, it can be used even in relatively small satellite systems. Furthermore, because the hybrid rocket engine ignition method according to the present invention generates high-temperature gas using the rocket propellant, it can reliably ignite the hybrid rocket engine even in outer space. Repeated ignition can also be reliably performed.

[0017] The hybrid rocket engine ignition method according to the present invention may involve directly igniting the hybrid rocket engine with the generated high-temperature gas, or it may involve igniting the hybrid rocket engine via other ignition means. When using other ignition means, for example, the high-temperature gas may be used to heat the fuel to generate combustion gas, and this combustion gas may be used to ignite the hybrid rocket engine.

[0018] In the ignition device and hybrid rocket engine ignition method according to the present invention, the propellant may be in any state, such as liquid, solid, or gas. When the propellant is a liquid, it is preferable that the device is configured to generate the high-temperature gas by decomposing the propellant with a catalyst. In this case, the high-temperature gas can be generated relatively easily and with low power. Furthermore, it is preferable that the propellant is less toxic than hydrazine. In this case, it is safer than using existing hydrazine (anhydrous hydrazine).

[0019] The rocket propulsion system according to the present invention is characterized by having a hybrid rocket engine ignition device according to the present invention and a hybrid rocket engine provided so as to be ignitable by the hybrid rocket engine ignition device.

[0020] Since the rocket propulsion system according to the present invention has a hybrid rocket engine ignition device according to the present invention, not only the ignition device but also the miniaturization and weight reduction of the rocket propulsion system itself can be achieved. Therefore, it can be used as a relatively small satellite system or the like. Further, the rocket propulsion system according to the present invention can stably ignite the hybrid rocket engine even in outer space by the hybrid rocket engine ignition device according to the present invention, and can also surely perform repeated ignition. Therefore, reentry into the atmosphere and transportation between planetary orbits where re-ignition is required can be efficiently performed.

Effects of the Invention

[0021] According to the present invention, it is possible to provide a hybrid rocket engine ignition device, a hybrid rocket engine ignition method, and a rocket propulsion system that can surely be re-ignited, can be miniaturized and weight-reduced, and can be mounted on a relatively small satellite system or the like.

Brief Description of the Drawings

[0022] [Figure 1] It is a (a) perspective view and (b) side view of a hybrid rocket engine ignition device and a hybrid rocket engine showing an outline of a rocket propulsion system according to an embodiment of the present invention. [Figure 2] It is a block diagram showing a configuration in a ground ignition test of a hybrid rocket engine ignition device according to an embodiment of the present invention. [Figure 3] It is a graph showing the time change of the pressure measured at each position in a ground ignition test performed with the configuration of the ignition device shown in FIG. 2. [Figure 4] This graph shows the time variation of the pressure measured at each position during a vacuum ignition test conducted using the ignition device configuration shown in Figure 2. [Figure 5] This is a block diagram showing the configuration of a rocket propulsion system according to an embodiment of the present invention during a vacuum engine ignition test. [Figure 6] Figure 5 shows a graph illustrating the time variation of pressure measured at various positions during a vacuum engine ignition test conducted with the rocket propulsion system configuration shown. [Figure 7] This graph shows the time evolution of ζ, which represents the ignition characteristics of a hybrid rocket engine, calculated based on the graph shown in Figure 6. [Modes for carrying out the invention]

[0023] Embodiments of the present invention will be described below with reference to the drawings. Figures 1 to 7 show an ignition device for a hybrid rocket engine, a hybrid rocket engine ignition method, and a rocket propulsion system according to embodiments of the present invention. As shown in Figure 1, the rocket propulsion system 10 includes a hybrid rocket engine ignition device 11, an oxidizer tank 12, and a hybrid rocket engine 13.

[0024] As shown in Figure 1(b), the ignition device 11 for the hybrid rocket engine includes a high-temperature gas generating means 21 and a combustion ignition means 22. In one example, the high-temperature gas generating means 21 holds a catalyst and is configured to generate high-temperature gas by supplying liquid propellant for the rocket and decomposing the propellant with the catalyst. In this case, the high-temperature gas generating means 21 may consist of any component that generates high-temperature gas by decomposing liquid propellant with a catalyst, for example, it may consist of a 0.5N class propulsion system as described in Non-Patent Literature 4. Furthermore, it is preferable that the propellant is less toxic than existing hydrazine (anhydrous hydrazine). The propellant is not limited to a liquid, but may be in any state, such as a solid or gas, as long as it can generate high-temperature gas. In addition, the high-temperature gas generating means 21 may have a means for supplying a catalyst to the propellant.

[0025] The combustion ignition means 22 has fuel and is configured to generate combustion gas by heating the fuel with high-temperature gas generated by the high-temperature gas generating means 21. Preferably, the combustion ignition means 22 is configured to combust by supplying an oxidizer from the oxidizer tank 12 to the generated combustion gas. The combustion ignition means 22 is also configured to ignite the hybrid rocket engine 13 with its combustion gas. The combustion ignition means 22 can be anything as long as it is capable of generating combustion gas by heating the fuel with high-temperature gas generated by the high-temperature gas generating means 21. The fuel is preferably a solid fuel considering miniaturization and safety, but it can be in any state, such as liquid or gas.

[0026] The ignition device 11 for the hybrid rocket engine only needs to be capable of igniting the combustion ignition means 22 with the high-temperature gas generating means 21, and capable of igniting the hybrid rocket engine 13 with the combustion ignition means 22. The high-temperature gas generated by the high-temperature gas generating means 21 may be low-energy and not capable of directly igniting the hybrid rocket engine 13.

[0027] As shown in Figure 1(a), the oxidizer tank 12 is provided to store the oxidizer and to supply the oxidizer to the combustion ignition means 22 and the hybrid rocket engine 13. In the specific example shown in Figure 1(a), the oxidizer tank 12 consists of two tanks. The tanks other than the oxidizer tank 12 are used for attitude control thrusters.

[0028] The hybrid rocket engine 13 is configured to propel the onboard rocket propulsion system 10 in space. The hybrid rocket engine 13 has a solid-phase fuel as propellant 13a and is configured to ignite the propellant 13a with combustion gas from the combustion ignition means 22 of the hybrid rocket engine ignition device 11. Preferably, the hybrid rocket engine 13 is configured so that the propellant 13a burns when an oxidizer is supplied from the oxidizer tank 12 to the combustion gas generated by heating. The hybrid rocket engine 13 has a nozzle 13b that ejects the combustion gas generated by the combustion of the propellant 13a. The rocket propulsion system 10 is configured to be propelled by the ejection from the nozzle 13b.

[0029] The hybrid rocket engine ignition method according to an embodiment of the present invention can be suitably carried out by a hybrid rocket engine ignition device 11. In the hybrid rocket engine ignition method according to an embodiment of the present invention, a high-temperature gas generating means 21 generates high-temperature gas using rocket propellant, the fuel in the combustion ignition means 22 is heated by the high-temperature gas and an oxidizer is supplied to generate combustion gas, and the hybrid rocket engine 13 is ignited by the combustion gas.

[0030] Next, I will explain the mechanism of action. The rocket propulsion system 10 ignites the hybrid rocket engine 13 using high-temperature gas generated by the rocket propellant, as dictated by the hybrid rocket engine ignition device 11 and hybrid rocket engine ignition method according to the embodiment of the present invention. Therefore, ignition can be performed with a relatively simple configuration that does not require high voltage or high power. As a result, the high-temperature gas generating means 21 can be made relatively small, and the hybrid rocket engine ignition device 11 and the rocket propulsion system 10 itself can be made smaller and lighter. For this reason, the rocket propulsion system 10 can be used, for example, as a relatively small satellite system weighing 500 kg to 100 kg or less.

[0031] The rocket propulsion system 10 generates high-temperature gas using rocket propellant via a high-temperature gas generation means 21, enabling stable ignition of the hybrid rocket engine 13 even in outer space, resulting in highly reliable ignition performance. Furthermore, repeated ignition is also reliably achieved. Therefore, it can efficiently perform tasks such as re-entry into the atmosphere and inter-planetary transport requiring re-ignition.

[0032] The rocket propulsion system 10 can enhance the safety of the hybrid rocket engine 13 by using a less toxic propellant than existing hydrazine (anhydrous hydrazine) as the propellant for the high-temperature gas generating means 21. In this case, for example, by using the 0.5N class propulsion system described in Non-Patent Literature 4 as the high-temperature gas generating means 21, a less toxic propellant than existing hydrazine (anhydrous hydrazine) can be used.

[0033] Furthermore, since the rocket propulsion system 10 ignites the hybrid rocket engine 13 via the combustion ignition means 22, the high-temperature gas generated by the high-temperature gas generating means 21 can be kept to a low energy level that cannot ignite the hybrid rocket engine 13, and the high-temperature gas generating means 21 can be made smaller and lighter. In one specific example, a thruster with a thrust of about 1 / 500 to 1 / 100 of the thrust of the hybrid rocket engine 13 can be used as the high-temperature gas generating means 21.

[0034] Furthermore, the rocket propulsion system 10 may be configured to ignite the hybrid rocket engine 13 directly from the high-temperature gas generating means 21, without having a combustion ignition means 22. In this case, the high-temperature gas generating means 21 must be capable of igniting the hybrid rocket engine 13, which may increase the size of the hybrid rocket engine ignition device 11. However, since the combustion ignition means 22 can be omitted, the increase in size can be kept to a minimum. [Examples]

[0035] First, a ground ignition test was conducted under atmospheric pressure using the ignition device configuration shown in Figure 2. As shown in Figure 2, in the test configuration, a propellant tank 31 storing low-toxicity HNP225 (propellant) is connected to a high-temperature gas generating means 21 via a solenoid valve, and the high-temperature gas generating means 21 is further connected to a combustion ignition means 22. In addition, an oxidizer tank 12 and a nitrogen (GN2) tank 32 are connected to the combustion ignition means 22 via solenoid valves, non-return valves, needle valves, etc. The high-temperature gas generating means 21 consists of a 0.5N class propulsion system described in Non-Patent Literature 4, which uses low-toxicity HNP225 (propellant) as fuel. The combustion ignition means 22 consists of a system that burns solid fuel.

[0036] In the test, first, the fuel HNP225 was supplied to the high-temperature gas generating means 21 by pressure, with the timing adjusted by a solenoid valve. Next, the supplied HNP225 was supplied to a catalyst, where it was decomposed by a catalytic reaction, generating high-temperature gas in the high-temperature gas generating means 21, and this high-temperature gas was supplied to the combustion ignition means 22. After the fuel in the combustion ignition means 22 was sufficiently heated by the high-temperature gas, an oxidizer was supplied from the oxidizer tank 12 to the combustion ignition means 22, and the solid fuel was burned to generate combustion gas.

[0037] As shown in Figure 2, during the test, pressure sensors were installed at the following locations to measure the pressure during these processes: inside the propellant tank 31 (PNO in the figure), upstream of the flow path from the oxidizer tank 12 (PLO in the figure), downstream of the flow path from the oxidizer tank 12 (before the combustion ignition means 22; PBO in the figure), inside the high-temperature gas generating means 21 (high-temperature gas pressure measurement; PTH in the figure), upstream of the chamber of the combustion ignition means 22 (PPF in the figure), and downstream of the chamber of the combustion ignition means 22 (PPB in the figure).

[0038] In the test, the high-temperature gas from the high-temperature gas generating means 21 preheated the solid fuel in the combustion ignition means 22 before supplying the oxidizer. Furthermore, to achieve the target combustion pressure in the chamber of the combustion ignition means 22, the mass flow rate of the oxidizer from the oxidizer tank 12 was pre-set based on chemical equilibrium calculations and adjusted using a needle valve. Nitrogen from the nitrogen tank 32 was used as a purge gas to assist in stopping combustion in the combustion ignition means 22.

[0039] Figure 3 shows the time variation of the pressure measured at each position during the test. As shown in Figure 3, it was confirmed that the value of PTH, which indicates the pressure of the high-temperature gas generated in the high-temperature gas generation means 21, increased due to the catalytic reaction in the high-temperature gas generation means 21 (in the range of 0 seconds to 20 seconds in the figure). At the same time, it was confirmed that the value of PBO, which indicates the pressure downstream of the oxidizer tank 12, and the value of PPF, which indicates the pressure upstream of the chamber of the combustion ignition means 22, also gradually increased. This increase in PBO and PPF pressures is thought to be due to the vaporization of the solid fuel in the combustion ignition means 22 by the generated high-temperature gas. The temperature of the high-temperature gas generated in the high-temperature gas generation means 21 was approximately 900°C.

[0040] Next, when the oxidizer was supplied to the combustion ignition means 22, combustion began in the combustion ignition means 22 (at 20 seconds in the figure), and it was confirmed that the values ​​of PPF, which indicates the pressure in the upstream part of the chamber of the combustion ignition means 22, and PPB, which indicates the pressure in the downstream part, rose sharply. Furthermore, during combustion in the combustion ignition means 22 (in the range of 20 to 30 seconds in the figure), the pressures in PPF and PPB were fluctuating, suggesting that combustion oscillations were occurring, but these combustion oscillations are not considered to be severe enough to destroy the ignition device in this test. After combustion for 10 seconds (from 30 seconds onward in the figure), the supply of the oxidizer was stopped, and nitrogen purge gas was supplied to the combustion ignition means 22 to stop the combustion. [Examples]

[0041] Next, a vacuum ignition test was conducted by placing all components of the ground ignition test configuration shown in Figure 2, except for the oxidizer tank 12 and nitrogen tank 32, inside a vacuum tank. The test was carried out in the same manner as the ground ignition test, with the inside of the vacuum tank under vacuum. Figure 4 shows the time change of the pressure measured at each position during the test. The catalytic reaction time in the high-temperature gas generation means 21 is set to 20 seconds in Figure 3, while it is set to 15 seconds in Figure 4. Otherwise, it was confirmed that the vacuum ignition test shown in Figure 4 yielded almost the same results as the ground ignition test shown in Figure 3. [Examples]

[0042] Next, a vacuum engine ignition test was conducted using the rocket propulsion system 10 configuration shown in Figure 5. As shown in Figure 5, in the vacuum engine ignition test, in addition to the configuration of the ground ignition test shown in Figure 2, a hybrid rocket engine 13 was connected downstream of the combustion ignition means 22, and the oxidizer tank 12 and nitrogen tank 32 were also connected to the hybrid rocket engine 13 via solenoid valves, check valves, etc., respectively. Also, as shown in Figure 5, everything except the oxidizer tank 12 and nitrogen tank 32 was installed inside the vacuum tank 33. In the test, the inside of the vacuum tank 33 was evacuated, and the solid fuel of the combustion ignition means 22 was heated and burned in the same way as in the ground ignition test. The combustion gas was then supplied to the hybrid rocket engine 13, and at the same time, oxidizer was supplied from the oxidizer tank 12 to the hybrid rocket engine 13, burning the solid fuel of the hybrid rocket engine 13. The hybrid rocket engine 13 has a thrust of 200 N class.

[0043] As shown in Figure 5, during the test, pressure sensors were installed at the following locations during these processes: upstream of the flow path from the oxidizer tank 12 to the hybrid rocket engine 13 (PLO in the figure), downstream of the flow path from the oxidizer tank 12 to the hybrid rocket engine 13 (PIO in the figure), upstream of the flow path from the nitrogen tank 32 to the hybrid rocket engine 13 (PNL in the figure), downstream of the flow path from the oxidizer tank 12 to the combustion ignition means 22 (before the combustion ignition means 22; PBO in the figure), inside the high-temperature gas generating means 21 (high-temperature gas pressure measurement; PTH in the figure), upstream of the chamber of the combustion ignition means 22 (PPF in the figure), downstream of the chamber of the combustion ignition means 22 (PPB in the figure), upstream of the fuel injection section of the hybrid rocket engine 13 (PFP ​​in the figure), upstream of the chamber of the hybrid rocket engine 13 (PTF in the figure), and inside the chamber of the hybrid rocket engine 13 (PTB in the figure), and the pressure was measured.

[0044] In the test, the oxidizer was supplied to the hybrid rocket engine 13 only after the solid fuel of the hybrid rocket engine 13 had been sufficiently heated by the combustion gas from the combustion ignition means 22. In addition, nitrogen from the nitrogen tank 32 was used not only for combustion in the combustion ignition means 22 but also as a purge gas to assist in stopping combustion in the hybrid rocket engine 13.

[0045] Figure 6 shows the time variation of the pressure measured at each position during the test. As shown in Figure 6, it was confirmed that the pressure changes in the high-temperature gas generation means 21 and the combustion ignition means 22 were almost the same as those in the vacuum ignition test shown in Figure 4. Also, as shown in Figure 6, it was confirmed that supplying an oxidizer to the hybrid rocket engine 13 5 seconds after the start of combustion in the combustion ignition means 22 (at 25 seconds) caused the hybrid rocket engine 13 to start burning smoothly, and the PTF, which indicates the pressure in the upstream part of the chamber of the hybrid rocket engine 13, and the PTB, which indicates the pressure in the downstream part, rose sharply. Furthermore, 5 seconds after the start of combustion in the hybrid rocket engine 13 (at 30 seconds), the supply of oxidizer to the combustion ignition means 22 was stopped to halt combustion, but even after the combustion of the combustion ignition means 22 stopped, the pressures of PTF and PTB did not decrease, confirming that the hybrid rocket engine 13 continued to burn.

[0046] From the results in Figure 6, in order to confirm the ignition status of the hybrid rocket engine 13, the time change of ζ calculated by equation (1) is obtained and shown in Figure 7. Here, P c This is the pressure inside the chamber of the hybrid rocket engine 13 (PTB in Figure 5), and m o Time derivative (m o The value with a dot above it represents the mass flow rate of the oxidizer to the hybrid rocket engine 13. The subscript "st" indicates a steady state. In Figure 7, the start time of oxidizer supply to the hybrid rocket engine 13 is set to 0 seconds.

[0047]

number

[0048] The hybrid rocket engine 13 normally exhibits an O / F (mass flow rate ratio of oxidizer to fuel) shift during combustion, and even if the mass flow rate of the oxidizer during combustion is constant, the mass flow rate of the fuel changes due to the fuel retreat characteristics. However, as shown in Figure 6, in the vacuum engine ignition test, it was confirmed that when the mass flow rate of the oxidizer was constant, the pressure (PTB) inside the chamber of the hybrid rocket engine 13 was also nearly constant. Therefore, in the vacuum engine ignition test, there is no O / F shift, and it is considered that the time required to reach a steady state can be expressed by equation (1).

[0049] As shown in Figure 7, the value of ζ increased with time, and it was confirmed that a steady state was reached in approximately 1.5 seconds. It was also confirmed that the time required to reach 80% of the steady state was approximately 0.46 seconds. From these results, it can be said that the rocket propulsion system 10 shown in Figure 5 reaches a steady state in a relatively short time. [Explanation of Symbols]

[0050] 10 Rocket propulsion systems 11. Ignition system for hybrid rocket engines 21 High-temperature gas generating means 22 Combustion ignition means 12 Oxidizer Tank 13 Hybrid rocket engine 13a Propellant 13b Nozzle 31 Propellant Tank 32 Nitrogen Tanks 33 Vacuum Tank

Claims

1. It has a means for generating high-temperature gas using rocket propellant, The system is configured to be able to ignite a hybrid rocket engine using the high-temperature gas generated by the high-temperature gas generating means. A distinctive feature is the ignition system for hybrid rocket engines.

2. The propellant is a liquid, The high-temperature gas generating means is configured to generate the high-temperature gas by decomposing the propellant with a catalyst. The ignition device for a hybrid rocket engine as described in claim 1, characterized by its features.

3. The ignition device for a hybrid rocket engine according to claim 2, characterized in that the propellant consists of a substance with lower toxicity than hydrazine.

4. The ignition device for a hybrid rocket engine according to claim 1, characterized in that it has a combustion ignition means provided to heat fuel with the high-temperature gas generated by the high-temperature gas generating means to generate combustion gas, and to ignite the hybrid rocket engine with the combustion gas.

5. A hybrid rocket engine ignition method characterized by generating high-temperature gas using rocket propellant and using that high-temperature gas to ignite a hybrid rocket engine.

6. The hybrid rocket engine ignition method according to claim 5, characterized in that the fuel is heated by the high-temperature gas to generate combustion gas, and the hybrid rocket engine is ignited by the combustion gas.

7. An ignition device for a hybrid rocket engine according to any one of claims 1 to 4, A hybrid rocket engine that can be ignited by the aforementioned ignition device for hybrid rocket engines, A rocket propulsion system characterized by having [a certain feature].

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