EP1700024A1 - Procede permettant d'allumer un moteur de fusee et moteur de fusee - Google Patents

Procede permettant d'allumer un moteur de fusee et moteur de fusee

Info

Publication number
EP1700024A1
EP1700024A1 EP04808706A EP04808706A EP1700024A1 EP 1700024 A1 EP1700024 A1 EP 1700024A1 EP 04808706 A EP04808706 A EP 04808706A EP 04808706 A EP04808706 A EP 04808706A EP 1700024 A1 EP1700024 A1 EP 1700024A1
Authority
EP
European Patent Office
Prior art keywords
fuel
ignition
propellant
rocket
liquid
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP04808706A
Other languages
German (de)
English (en)
Inventor
Herman Fedde Rein SCHÖYER
Anton Gerhardus Maria Maree
Hubertus Marie Sanders
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Nederlandse Organisatie voor Toegepast Natuurwetenschappelijk Onderzoek TNO
Original Assignee
Nederlandse Organisatie voor Toegepast Natuurwetenschappelijk Onderzoek TNO
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Nederlandse Organisatie voor Toegepast Natuurwetenschappelijk Onderzoek TNO filed Critical Nederlandse Organisatie voor Toegepast Natuurwetenschappelijk Onderzoek TNO
Priority to EP04808706A priority Critical patent/EP1700024A1/fr
Publication of EP1700024A1 publication Critical patent/EP1700024A1/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/96—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by specially adapted arrangements for testing or measuring
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements

Definitions

  • the invention is directed to a process for igniting a rocket motor or rocket engine, which process comprises igniting a fuel rich propellant mixture, or streams of fuel rich propellant.
  • Many rocket motors or rocket engines need an ignition system to start the engine (it is more or less common to indicate rocket motors that use solid propellants as a rocket motor and those that use liquid propellants as a rocket engine).
  • An ignition system is necessary for rocket engines that do not use hypergolic propellants (i.e. propellants that ignite spontaneously when brought into contact with each other) or that use catalysts such as the well- known Shell 405 ® catalyst (30% iridium on alumina) that is commonly used in conjunction with hydrazine monopropellant thrusters.
  • Ignition systems are varied: solid propellant ignition systems, pyrotechnic ignition systems(a pyrotechnic ignition system mainly generates hot particles), pyrogen ignition systems (a pyrogen ignition system generates hot gas and resembles a small solid rocket motor), ignition systems based on a spark plug, ignition systems with a spark plug and torch, ignition systems with special chemicals that are hypergolic with one of the main propellants or ignition systems based on pyrophoric chemicals.
  • Multi propellant (i.e. bi-propellant or tri-propellant) rocket engines often use a fuel rich propellant mixture as the maximum performance of a rocket engine or motor depends primarily on two parameters:
  • the molar mass of the combustion products should be low (this is the reason that hydrogen is often used as fuel) and the combustion temperature should be high. In fact, the ratio of the combustion temperature and molar mass (T/M) should be maximized.
  • Oxidizers have a high molar mass; most fuels a low one.
  • Rocket motors or engines often use a fuel-rich mixture. Rocket motors or engines that use hypergolic propellants do not need an ignition system. The two propellants ignite spontaneously upon contact.
  • An example of hypergolic ignition is given in US patent 3,040,521. This concerns a liquid rocket engine using hydrogen peroxide as oxidizer and kerosene or a similar hydrocarbon as a fuel. The hydrogen peroxide is decomposed catalytically and upon contact with the kerosene (or other hydrocarbon) spontaneous combustion takes place.
  • the engine therefore does not need a separate ignition system.
  • this design has been used amongst others in the 'Black Knight'.
  • Many rocket engines and especially high performance liquid rocket engines use non-hypergolic liquid propellants while all solid rocket motors use solid propellants.
  • Examples of non-hypergolic liquid rocket engines are: - The HM-7 (used as an upper-stage engine on the Ariane-4 and to be used on Ariane 5), The HM-60 Vulcain that is the main engine on Ariane 5, The American RL- 10, - The American RL-60, - The Japanese LE 5, The Japanese LE 7, The foreseen VINCI engine, The Fl engine of the Saturn 5.
  • Typical examples of such motors are the Ariane 5 solid boosters, the solid boosters of the US Space Shuttle, the solid boosters of the Ariane 4, and many military missiles.
  • the mixture ratio is often fuel rich.
  • the Fl used a pyrophoric mixture (85% triethylborane +15% triethylaluminium) that spontaneously ignites when coming into contact with air or (liquid) oxygen, as is the case in the Fl engines of the Saturn 5.
  • Electric spark plugs that ignite a mixture of gaseous oxygen and gaseous hydrogen are also used.
  • the mixture ratio of such ignition systems is fuel rich to cool the ignition system and prevent over heating. Other engines are ignited by pyrogen ignition systems.
  • the ignition failure was caused by a fuel rich flame trying to ignite a fuel rich mixture.
  • the cost of these two failed launches amounted to about G €l: split into about M € 200 for the loss of two Ariane 4 launchers, and about M € 800, the cost of the lost satellites itself.
  • the problem was rectified without changing the composition of the ignition system propellant, it illustrates the sensitiveness of the ignition and the large costs associated with an ignition failure. Re-ignition is sometimes a requirement for rocket motors. In that case, pyrotechnic and pyrogen ignition systems are less suitable as these can only be used once.
  • spark plug ignition allows multiple ignitions and has demonstrated this on numerous occasions.
  • a problem of existing electrical ignition systems is that they require many components: spark plug ignition requires a high-tension electrical circuit, a reliable spark plug and two valves to admit the oxidizer and fuel in the correct sequence and proportion. This leads to relatively complex systems and is one of the reasons that also for liquid engines often pyrotechnic, pyrogen or pyrophoric ignition systems are being used.
  • the present invention solves these problems for igniting fuel rich propellant combinations: it allows multiple ignitions, provides hot oxidizer rich ignition gases, does not require high-tension electrical circuitry and is relatively simple. Moreover, the invention lends itself to a generic design that has mainly to be adapted in size for the various applications. More in particular this means that the principle of the system is independent of the specific type and size of rocket motor or engine.
  • the invention is directed to a process to ignite a rocket motor or rocket engine, that contains a fuel rich propellant mixture, or wherein fuel rich streams of propellant are being injected, which propellant or streams of propellant are based on a combination of fuel and oxidizer, said fuel being selected from the group of hydrogen, methane, kerosene, aluminium, HTBP, GAP, Poly Nimmo and BAMO; the fuel rich propellant is ignited with a free- molecular oxygen containing hot gas obtained by decomposing at least one ignition system liquid.
  • the invention is directed to a rocket motor or rocket engine assembly comprising at least one rocket motor or rocket engine to be operated using a fuel rich propellant mixture, or fuel rich propellant streams, which propellant is based on a combination of fuel and oxidizer, said fuel being selected from the group of hydrogen, methane, kerosene, aluminium, HTBP, GAP, Poly Nimmo and BAMO, means for containing the said propellant mixture in the said rocket motor, or feeding the propellant streams to the said rocket engine and means for igniting said fuel rich propellant mixture, or fuel rich propellant streams, said means comprising a storage tank for an ignition system liquid containing at least one of the group of H2 ⁇ 2, solutions of hydrazinium nitroformate (HNF) and/or ammonium dinitramide (ADN) in water, means for decomposing the ignition system liquid to produce a free-molecular oxygen containing hot gas and means for feeding said hot gas into the rocket motor or engine to ignite the fuel rich propellant mixture, or fuel rich propellant streams, said
  • the principle of the invention is based on the use of an oxygen rich hot gas for igniting a fuel rich propellant mixture in a rocket motor, or streams of fuel rich propellant in a rocket engine. It has been found that this method will overcome, most or all of the drawbacks of the known systems as discussed above.
  • the invention comprises an ignition system for rocket motors or rocket engines that provides oxidizer-rich hot ignition gas, i.e. a gas that contains free-molecular oxygen, to specific fuel-rich streams of propellants or mixtures of propellants that have to be ignited. This is preferably achieved by using a storable liquid oxidizer that is decomposed catalytically, thereby releasing heat. This heat of decomposition raises the temperature of the combustion products to well over 1000 K.
  • the storable liquid oxidizers can be stored for prolonged periods at room temperature, under pressure, are safe and require no specific safety measures beyond those that are common for chemicals.
  • the operating pressure exceeds the pressure at which the rocket motor or engine, or a pre-combustion chamber or gas generator has to be ignited.
  • the ignition pressures usually are substantially less than the steady state operation pressure of the engines, but nevertheless may easily exceed 4 MPa.
  • the oxidizers can operate at these high pressures without problems. Suitable oxidisers, i.e. materials (preferably liqnid) that are energetic materials able to provide oxygen rich hot gas (i.e.
  • T > 750 K) are ADN, HNF, H2O2, Ammonium perchlorate, ammonium nitrate, hy ⁇ droxyl- ammonium nitrate and combinations thereof, optionally in the form of an aqueous solution thereof.
  • Ammonium perchlorate ammonium nitrate, hy ⁇ droxyl- ammonium nitrate and combinations thereof, optionally in the form of an aqueous solution thereof.
  • the following oxidizers that can be decomposed catalytically are of interest for the ignition systems of this invention:
  • Hydrogen peroxide H2O2.
  • a solution of -70% ADN in water -70% NH 4 N(NO 2 )2 + -30% H 2 0
  • the free-oxygen from the ignition system liquid is used to ignite the fuel rich propellant, which is based on the combination of a fuel and an oxidizer.
  • the fuel to be ignited is selected from the group of hydrogen, methane, kerosene, aluminium, HTBP (hydroxyl terminated polybutadiene), GAP (Glycidylazide Polymer), Poly Nimmo (Poly3- nitromethoxy-3-methyloxetane) and BAMO (Bis-azido Methyl Oxetane). With these fuels the ignition with hot free -molecular oxygen proceeds rapidly and safely.
  • oxidizer use is preferably made of O2, N2O4, HNO3, H2O2, NH4CIO4, hydrazinium nitroformate and ammonium dinitramide.
  • the table below gives the temperature of the decomposition products of the ignition system liquid and their main composition at 4 and 8 MPa.
  • liquid oxidizers produce hot oxygen rich gases ranging from over 5 wt.% free-oxygen to over 42 wt.% free-oxygen and the gas temperatures all exceed 1000 K. This temperature is more than sufficient to ignite the fuel rich fuel/oxidizer combination in the rocket engine or motor.
  • a catalyst where it decomposes.
  • catalysts of different decomposition can be used. In general the catalysts will be solid heterogeneous catalysts, wherein the catalytically active material has been applied to a suitable support or is present in a suitable shape.
  • Examples are fixed bed catalysts based on supported catalytically active material, the supports being porous granules or extrudate of inert material, granules or extrudate of catalytically active material and meshes or screens of (supported) catalytically active material.
  • the supports being porous granules or extrudate of inert material, granules or extrudate of catalytically active material and meshes or screens of (supported) catalytically active material.
  • one of the following combinations of oxidisers and catalysts may be used.
  • Catalyst beds on an alumina support are commercially available (e.g. from Shell, Aerojet, Degussa and Rhone Poulenc).
  • the other supports mentioned above may also be used.
  • the process of the present invention uses specific combinations of ignition liquid system on the one hand and combinations of fuel and oxidizer on the other hand.
  • the mixture ratio of fuel and oxidizer is such that the amount of fuel is in excess of the amount of oxidizer based on complete combustion of the fuel. More in particular the solid systems, based on aluminium, HTBP
  • the invention is suitable for combinations of concentrated H2O2, or aqueous solutions of HNF or ADN as igniter lipjuid system with one or more of H2, CH4 and Kerosene as fuels and one or more of O2, N2O4, HNO3 as oxidizers in rocket engines and one or more of HTBP, GAP, Poly Nimmo, BAMO possibly together with aluminium powder as fuel and one or more of NH4CIO4, HNF, ADN as oxidizer in rocket motors while the propellant mixture ratio fuel/oxidizer is always fuel rich, i.e. after consumption of all the oxidizer the combustion products still contain combustible species.
  • Another extremely interesting application is for test facilities where repeatedly rocket motors or other combustors have to be ignited for test purposes.
  • the ignition system of this invention avoids the use of expensive or hazardous ignition systems or ignition systems that presently often have to be used.
  • Figure 1 shows a schematic pressure regulated ignition system
  • Figure 2 a blow-down ignition system
  • the ignition liquid is stored in a tank (3) that is pressurized to the required pressure by gas stored in a high-pressure gas tank (1).
  • the pressurant gas can be any inert gas that is commonly used to pressurize rocket propellants, e.g. N2 or He.
  • a pressure regulator (2) controls the pressure in the tank with ignition liquid.
  • latching valve (4) that can be opened and closed on command. This provision allows for multiple ignitions at a constant ignition pressure.
  • a valve pyro valve or latching valve
  • a filter will be placed downstream of any pyro valve and upstream of the pressure regulator. This is not shown in Figure 1, but is common practice in rocket propulsion systems and known to the specialists in the field.
  • the ignition system tube may have any suitable length or shape to fit into the combustion chamber and have a special shape to enhance mixing of the hot oxygen rich ignition system gases with the propellants to be ignited or impinge on the propellant to be ignited. To this end the ignition system tube may introduce a swirl to the ignition system gases, inject them sideward with respect to the main axis of the rocket motor or engine. By closing valve (4), the ignition sequence can be terminated. Contrary to pyrogen or pyrotechnic ignition systems, the ignition system according to this invention allows determining experimentally during development tests the optimum duration of the ignition process as this only depends on the duration that valve (4) is opened. To do this with a pyrotechnic or pyrogen ignition system, one would have to develop different ignition systems for every ignition duration.
  • the ignition system As there is a high- pressure gas tank (1) and a pressure regulator (2) the ignition system, according to the schematic of Figure 1, operates at a constant pressure. This allows also multiple ignitions at the same pressure, which pressure can be selected to be the most suitable ignition pressure for that particular rocket motor or rocket engine.
  • the dimensions of the ignition system depend on the total amount of ignition liquid to be stored in the tank (3), the pressure at which the system has to operate and the required mass flow rate of the ignition liquid.
  • the pressure regulator (2) can be set to a lock-up pressure in accordance with the required operating pressure. This will affect the mass flow rate of the ignition liquid. To control the mass flow rate the system allows mounting a cavitating venturi between the tank (3) and the valve (4), or between the valve (4) and the catalyst bed or screens (5).
  • the tank with the ignition liquid only needs a small ullage volume as the system operates in the pressure-regulated mode.
  • a simpler schematic for an ignition system according to the present invention is given in Figure 2.
  • the ignition system operates in blow-down mode.
  • the difference with the ignition system from Figure 1 is that the pressurant gas tank and pressure regulator are absent, while the tank with ignition liquid (3) is only partially filled with ignition liquid while the remainder of the tank volume is filled by a pressurant gas, such as nitrogen, helium or other suitable pressurant gas.
  • the large initial ullage volume avoids large pressure drops during the ignition sequence or for successive ignitions.
  • Using a large ullage volume for tanks containing liquids is common practice in rocket propulsion and known to specialists in the field.
  • the ignition fluid tank is pressurized to a pressure level that is adequate for the number of ignitions that have to be performed and in accordance with the specified ignition conditions in the rocket motor or rocket engine.
  • the latching valve (4) can be replaced by a pyro valve. In that case a filter may be placed downstream of the pyro valve (4) and upstream of the catalyst bed or screens (5). This is common practice in rocket propulsion systems and known to the specialists in the field.
  • the dimensions of the ignition system depend on the total amount of ignition liquid to be stored in the tank (3), the pressure at which the system has to operate and the required mass flow rate of the ignition liquid. For systems that require a different operating pressure the ullage pressure in the tank (3) may be adjusted. This will change the mass flow rate of the ignition liquid.
  • To control the mass flow rate the system allows mounting a cavitating venturi between the tank (3) and the valve (4), or between the valve (4) and the catalyst bed or screens (5). By selecting the proper dimension of the cavitating venturi for every application, it is possible to use the same universal ignition system for a wide range of requirements without changing major system components. Other layouts are not excluded, the simplified schematics are provided as typical examples but do not limit or exclude other arrangements.
  • the advantages of this invention compared to the existing ignition systems are the following:
  • pyrotechnic, pyrogen and pyrophoric ignition systems are a safety hazard. Pyrotechnic and pyrogen ignition systems use an explosive mixture and have to be initiated by a pyrotechnic initiator. To prevent accidental ignition a Safe & Arm device is usually installed. All these safety measures contribute to the mass and complexity and introduce failure modes. It also requires a reliable electrical ignition circuit and an electrical circuit to arm/disarm the Safe & Arm device. 6. The transport and handling of pyrotechnic devices requires specific safety measures.
  • Electro-gaseous ignition systems require a high-tension (voltage) current and a spark plug. This requires an elaborate electrical system that is not required for a catalytic ignition system.
  • the ignition system according to the present invention can be easily adjusted for different ignition pressures and ignition flow rates by adjusting the pressure regulator (2), adjusting the ullage pressure in the ignition liquid tank (3) and by incorporating a cavitating venturi between the ignition liquid tank (3) and the valve (4) or between the valve (4) and the catalyst bed or screens (5).
  • the ignition system according to the present invention allows for a generic design that only has to be scaled for the specific application. 12. Because of the simplicity and safety of the ignition system, this system is also ideally suited for test facilities if ignition devices are required there.

Abstract

L'invention concerne un procédé permettant d'allumer un moteur de fusée ou un moteur de fusée utilisant un gaz chaud contenant un oxygène à molécules libres obtenu par décomposition d'au moins un liquide (1) de système d'allumage. Ledit moteur de fusée comprend un mélange de propergol riche en carburant à base de carburant et de comburant, le carburant dudit mélange étant sélectionné dans le groupe constitué par aluminium, HTBP, GAP, Poly Nimmo et BAMO. Ce moteur de fusée utilise des flux de comburant riches en carburant, lesdits flux de comburant étant à base d'une combinaison de carburant et de comburant, et le carburant étant sélectionné dans le groupe constitué par hydrogène, méthane et kérosène.
EP04808706A 2003-11-18 2004-11-11 Procede permettant d'allumer un moteur de fusee et moteur de fusee Withdrawn EP1700024A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP04808706A EP1700024A1 (fr) 2003-11-18 2004-11-11 Procede permettant d'allumer un moteur de fusee et moteur de fusee

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP03078614A EP1533511A1 (fr) 2003-11-18 2003-11-18 Procédé d'allumage de moteur fusée et moteur de fusée
PCT/NL2004/000790 WO2005049999A1 (fr) 2003-11-18 2004-11-11 Procede permettant d'allumer un moteur de fusee et moteur de fusee
EP04808706A EP1700024A1 (fr) 2003-11-18 2004-11-11 Procede permettant d'allumer un moteur de fusee et moteur de fusee

Publications (1)

Publication Number Publication Date
EP1700024A1 true EP1700024A1 (fr) 2006-09-13

Family

ID=34429451

Family Applications (2)

Application Number Title Priority Date Filing Date
EP03078614A Withdrawn EP1533511A1 (fr) 2003-11-18 2003-11-18 Procédé d'allumage de moteur fusée et moteur de fusée
EP04808706A Withdrawn EP1700024A1 (fr) 2003-11-18 2004-11-11 Procede permettant d'allumer un moteur de fusee et moteur de fusee

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP03078614A Withdrawn EP1533511A1 (fr) 2003-11-18 2003-11-18 Procédé d'allumage de moteur fusée et moteur de fusée

Country Status (3)

Country Link
EP (2) EP1533511A1 (fr)
CA (1) CA2546371A1 (fr)
WO (1) WO2005049999A1 (fr)

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FR2984307B1 (fr) * 2011-12-20 2014-01-10 Sme Procede de generation de gaz de combustion a partir d'un materiau solide precurseur d'oxygene et d'un materiau reducteur solide dissocies et dispositif associe
US11408376B2 (en) 2012-07-30 2022-08-09 Utah State University Thrust augmentation of an additively manufactured hybrid rocket system using secondary oxidizer injection
US10527004B2 (en) 2012-07-30 2020-01-07 Utah State University Restartable ignition devices, systems, and methods thereof
US10774789B2 (en) 2012-07-30 2020-09-15 Utah State University Methods and systems for restartable, hybrid-rockets
US11407531B2 (en) 2012-07-30 2022-08-09 Utah State University Space Dynamics Laboratory Miniaturized green end-burning hybrid propulsion system for CubeSats
CN103993984B (zh) * 2014-05-23 2016-07-13 清华大学深圳研究生院 一种过氧化氢辅助点火装置
CN109057996A (zh) * 2018-09-27 2018-12-21 北京航天动力研究所 一种液体火箭发动机四机并联热试验装置
US20200325822A1 (en) * 2019-04-14 2020-10-15 Hamilton Sundstrand Corporation Power modules with blow down fuel and propellant delivery systems
CN112206782B (zh) * 2019-07-12 2023-11-14 南京理工大学 含Ni/MnO2复合镀层的催化剂芯片及其制备方法
CN112610363B (zh) * 2020-12-18 2021-11-16 西安航天动力研究所 全流量补燃循环发动机富氧半系统热试装置及热试方法
CN114436722B (zh) * 2021-12-30 2023-04-18 西安近代化学研究所 一种催化高氯酸铵的催化剂、制备方法及应用
CN114352440A (zh) * 2022-01-07 2022-04-15 北京理工大学 一种模块化固体火箭冲压发动机地面直连试验装置
CN116929159B (zh) * 2023-09-18 2024-01-09 北京星河动力装备科技有限公司 固液混合动力的运载火箭及其发射方法

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Also Published As

Publication number Publication date
WO2005049999A1 (fr) 2005-06-02
EP1533511A1 (fr) 2005-05-25
CA2546371A1 (fr) 2005-06-02

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