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Hydrocarbon-Seeded Ignition System for Small Spacecraft Thrusters Using Ionic Liquid Propellants

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Hydrocarbon-Seeded Ignition System for Small Spacecraft Thrusters Using Ionic Liquid Propellants SSC13-VII-6 Hydrocarbon-Seeded Ignition System for Small Spacecraft Thrusters Using Ionic Liquid Propellants Stephen A. Whitmore, Daniel P. Merkley, and Shannon D. Eilers Mechanical and Aerospace Engineering Department, Utah State University 4130 Old Main Hill, UMC 4130; (435)-797-2951 [email protected] Terry L. Taylor Technology Development and Transfer Office, NASA Marshall Spaceflight Center, ZP-30, MSFC AL 35812; (256)-544-5916 [email protected] ABSTRACT "Green" propellants based on Ionic-liquids (ILs) like Ammonium DiNitramide and Hydroxyl Ammonium Nitrate have recently been developed as reduced-hazard replacements for hydrazine. Compared to hydrazine, ILs offer up to a 50% improvement in available density-specific impulse. These materials present minimal vapor hazard at room temperature, and this property makes IL's potentially advantageous for "ride-share" launch opportunities where hazards introduced by hydrazine servicing are cost-prohibitive. Even though ILs present a reduced hazard compared to hydrazine, in crystalline form they are potentially explosive and are mixed in aqueous solutions to buffer against explosion. Unfortunately, the high water content makes IL-propellants difficult to ignite and currently a reliable “cold-start” capability does not exist. For reliable ignition, IL-propellants catalyst beds must be pre-heated to greater than 350 C before firing. The required preheat power source is substantial and presents a significant disadvantage for SmallSats where power budgets are extremely limited. Design and development of a "micro-hybrid" igniter designed to act as a "drop-in" replacement for existing IL catalyst beds is presented. The design requires significantly lower input energy and offers a smaller overall form factor. Unlike single-use "squib" pyrotechnic igniters, the system allows the gas generation cycle to be terminated and reinitiated on demand. INTRODUCTION The proposed ignition technology is offered as a This paper details the development and testing of a "drop-in" replacement for existing IL catalyst beds, and novel ignition system that is ideally suited for eliminates the need for a large pre-heat power source. propellants based on Ionic-liquids (ILs). IL-based The technology requires significantly lower input propellants have recently emerged as likely "reduced- energy and offers a smaller overall form factor. Unlike hazard" candidates to replace hydrazine as a primary single-use "squib" pyrotechnic igniters, the system propellant for space propulsions systems. Compared to allows the gas generation cycle to be terminated and hydrazine, ILs offer up to a 50% improvement in reinitiated on demand. When fully developed the available density-specific impulse. These materials system will allow reliable IL-propellant ignition present minimal vapor hazard at room temperature, and without a high wattage power source, toxic pyrophoric this property makes IL's potentially advantageous for ignition fluids, or a bi-propellant spark ignitor. "ride-shares" launch opportunities where hazards The approach is fundamentally different from all introduced by hydrazine servicing are cost-prohibitive. other “green propellant” solutions in the aerospace in Even though ILs present a reduced hazard compared the industry. Although the proposed system is more to hydrazine, in crystalline form they are potentially correctly a “hybrid” rocket technology, since only a explosive and are mixed in aqueous solutions to buffer single propellant feed path is required, it retains all the against explosion. Unfortunately, the high water content simple features of a monopropellant system. The technology is based on the principle of "seeding" an makes IL-propellants difficult to ignite and currently a 1 reliable “cold-start” capability does not exist. For oxidizing flow with a small amount of hydrocarbon. reliable ignition, existing catalyst beds for IL- The ignition is initiated electrostatically with a low- propellants must be pre-heated to greater than 350 C wattage inductive spark. Combustion gas byproducts before firing. The required preheat power source is from the hydrocarbon-seeding ignition process can substantial and presents a significant disadvantage for exceed 2400 C and the high exhaust temperature SmallSats where power budgets are extremely limited. ensures reliable propellant ignition. The system design is described in detail in this report. Whitmore 1 27th Annual AIAA/USU Conference on Small Satellites BACKGROUND AND LITERATURE REVIEW On the Need for Reduced Toxicity, Reduced Hazard This section will discuss the current state of the art Replacements for Hydrazine for hydrazine-based monopropellant systems, and Although procedures are in place to allow hydrazine motivate the need for “green” alternatives. Current to be managed safely at tightly controlled government state-of-the-art options will be considered, and the or defense-contractor operated test reservations and operational readiness for these options will be launch facilities, the toxicity and explosion potential of described. A thorough literature review of several hydrazine requires extreme handling precautions that promising “green” monopropellant options will be drive up operating costs. Increasingly, with a growing presented. Issues associated with operational readiness regulatory burden, infrastructure requirements of the most promising options will be discussed. associated with hydrazine transport, storage, servicing, Current State of the Art for Space Monopropellants and clean up of accidental releases are becoming cost prohibitive. As SmallSat operations continue to shift Hydrazine (N2H4) is by far the most commonly used propellant for primary spacecraft propulsion and from government–run organizations to private companies and universities operating away from attitude control thrusters. Hydrazine thrusters, which government-owned test reservations, servicing payloads typically consist of an electric solenoid valve, a requiring hydrazine as a propellant becomes pressurant tank, and a catalyst bed of alumina pellets ® 2 operationally infeasible. Additionally, these extreme impregnated with iridium (Shell 405 ) , feature simple handling precautions generally do not favor hydrazine design architectures, are highly reliable, and offer as a propellant for secondary payloads. vacuum specific impulse (Isp) exceeding 220 seconds. In a typical design, the catalyst initiates an exothermic A recent study by the European Space Agency’s decomposition of the hydrazine to produce ammonia, European Space Research and Technology Center nitrogen, and hydrogen gases with approximately 1600 (ESA/ESTEC) has identified two essential design J/g of heat released. Although hydrazine decomposition elements to achieving low cost space access; 1) using the Shell 405 catalyst can be performed without Reduced production, operational, and transport costs additional heat input to the catalyst, typical designs pre- due to lower propellant toxicity and explosion hazards, heat the catalyst bed to insure reliable ignition and a and 2) Reduced costs due to an overall reduction in consistent burn profile. subsystems complexity and overall systems interface complexity.8,9 This study showed the potential for Unfortunately, hydrazine is a powerful reducing considerable operational cost savings by simplifying agent that poses serious environmental concerns. propellant ground handing and servicing procedures. A Hydrazine is extremely destructive to living tissues and non-toxic, stable “green” alternative for hydrazine is is a probable human carcinogen. Exposure produces a variety of adverse systemic effects including damage to recommended by the study. liver, kidneys, nervous system, and red blood cells.3 In The ESA/ESTEC recommendations are directly aligned with NASA's In-Space Propulsion Systems addition to these biological and toxicological impacts, 10 hydrazine presents significant environmental dangers Roadmap (ISPSR) TA02.1.1.1, Monopropellants. A for the spacecraft and launch vehicle.4,5 When heated key element of ISPSR, identifies difficulties associated rapidly or exposed to extreme shock, such as the nearby with hydrazine as a spacecraft propellant and explosion of a linear charge or blasting cap, hydrazine recommends development of "less hazardous, less detonation is possible. Linear charges and explosive toxic" alternatives. NASA recently awarded a $40M bolts are common on launch vehicles and spacecraft, contract to a team lead by Ball Aerospace Corporation for a Technology Demonstration Mission (TDM) to thus the presence of hydrazine presents a tangible risk. 11 Solid hydrazine buildup, possibly augmented by frozen space-test a thruster system based on AF-M315E. AF- M315E is a USAF-proprietary Hydroxylammonium oxygen, detonated by a linear charge in an adapter ring 12 is suspected in the failure of an Atlas-Centaur upper Nitrate (HAN)-based IL propellant blend. stage.6 Hydrazine Replacement Monopropellant Options Most significantly, hydrazine has a high vapor A useful monopropellant replacement for hydrazine pressure at room temperature, approximately 1000 kPa must be sufficiently energetic to easily decompose and (145 psia). Because of this higher vapor pressure, produce good combustion properties. Non-storable hydrazine fumes significantly at room temperature and cryogenic or high freezing point propellants requiring presents a high risk as a respiratory hazard. All temperature control are not appropriate for space hydrazine-servicing operations must be
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