NASA Contractor Report 198531 rJ

Chemical Microthruster Options

Wim de Groot and Steve O1eson NYMA, Inc. Brook Park, Ohio

October 1996

Prepared for Lewis Research Center Under Contract NAS3-27186

National Aeronautics and Space Administration Trade names or manufacmre_" names are used in this relx_ for identification only. This usage does not constitute an official endorsement, either expressed or implied, by the National Aeronautics and Space Admini_ation. Chemical Microthruster Options

Wim de Groot° and Steve Oleson_ NYMA, Inc., Engineering Services Division Brookpark, OH

Abstract gas jet. It typically consists of nitrogen pressurized to approximately 21 MPa and, Chemical propulsion systems with potential depending on the thrust level, provides a application to microsatellites are classified maximum specific impulse (Isp)of 76 s in a blowdown mode. by propellant phase, i.e. gas, liquid, or solid. Four promising concepts are selected based An excellent review of the status of on performance, weight, size, cost, and reliability. The selected concepts, in varying microthrust technology in the mid 60s was stages of development, are advanced given by Suthedand and Maes. 1 Some of the technology items have since fallen out monopropellants, tridyne TM, electrolysis, and of favor, such as radio isotope heating of solid gas generator propulsion. Tridyne TM the working fluid to enhance thruster and electrolysis propulsion are compared vs. existing cold gas and monopropellant performance. Other technologies, not sufficiently developed for serious systems for selected microsatellite consideration at the time, could present missions. Electrolysis is shown to provide a attractive choices for some missions. significant weight advantage over Among such thrusters are gas-solid hybrid monopropellant propulsion for an orbit thrusters and gas generator type thrusters. transfer and plane change mission. Tridyne TM is shown to provide a significant The term microthrust refers to the thrust advantage over cold gas thrusters for orbit level and not to the physical size of the trimming and spacecraft separation. thruster. Typically,. rockets with a thrust between 10" and 10 N fall within this class. Introduction Depending on the type of energy that is supplied to provide the thrust, The desire to reduce the size and weight of microthrusters can be classified as satellites, especially those planned for chemical, electrical, or electro-chemical. future NASA missions, creates a renewed interest in microthrusters for trajectory Recent developments in electronics and corrections and attitude control. Previously, energy storage have made the application the most important parameters in the choice of electric thrusters, with their high specific of microthruster type were the performance impulse, more attractive in the microthrust and reliability. With the current trend toward region.2 In many applications where high small inexpensive satellites, however, cost, impulse bits are required or where electric size, and weight have been added as dominant considerations. Both reliability power is scarce or not available (missions to the outer solar system), chemical systems improvement and reduction in size, weight, are an attractive option. This paper only and cost can be accomplished by means of covers the chemical class. Electrolysis is reduced system complexity. These included because the nature of the energy enhancements, however, often occur at the provided to the working fluid to generate expense of performance. A standard simple thrust is chemical. system for microthrust propulsion is the cold

"Senior Propulsion Specialist, Senior Member AIAA # ResearchEngineer,Member AIAA

TMTfidyne is a registeredTrademark of Rockwell Intemational/Rocketdyn¢ Division. Chemical microrockets can be classified and thrust level are less important for such variously. The clearest distinction, missions. In a system designed for attitude analogous to that of Suthedand et al., 1is by control, response time becomes more the phase in which the propellant is stored. important. If attitude control needs to be These are referred to as solid, liquid, and provided over a period of several years, a gaseous rockets, as well as some hybrid high total impulse is required favoring a high systems, where the fuel and oxidizer appear specific impulse thruster. This often implies, in different phases. however, that system complexity increases as well as the weight and size, and that In general, solid propellant systems reliability issues become important. represent a minimum system dry weight and complexity as the result of the absence of Propellant phase has been chosen as the high pressure tanks, propellant lines and primary classification for this paper. valves. The need for a thermal source to Problems associated with propellant initiate combustion or decomposition will management systems, such as phase slightly reduce this weight advantage. A separation in the absence of gravity, slosh, disadvantage of solid propellant thrusters is and vaporization, are largely similar across the difficulty in accurately controlling the each class of propellants. A review is given thrust and cycling. of the currently available propulsion options including propellant properties and thruster Vaporizing liquid propellant rockets also do performance characteristics. From the not require high pressure tankage and available systems, four promising concepts propellant lines, but are more complex than for microthrust propulsion are selected for solid propellant rockets because of the need further analysis. Example mission analyses for phase separation between liquids and for two of these concepts, tridyne TM and gas, and, for applications with a lack of electrolysis, has been done. A substantial inertia, the need to restrain the liquid improvement in weight, size, cost, and (slosh). Gaseous systems carry the highest performance over currently available penalty in system weight, but will sometimes commercial propulsion systems is shown. be needed to deliver the response and/or thrust level that the solid and liquid systems Types of microthrust systems lack with great system simplicity. Each propellant class (gas, liquid, solid) All three propellant types produce thrust represents a number of different propulsion based on the same principle. A gas is concepts. In this section, concepts in each generated inside a chamber with sufficient class are summarized, roughly in order of pressure to choke a nozzle. To provide an complexity. The advantages and acceptable Isp, the molecular weight of the disadvantages of each type are listed, gas needs to be sufficiently low for a without details on specific mission significant acceleration in the nozzle. As an applications. Pertinent properties of addition to each of these three systems, an propellants, such as chemical composition, electrical heater can be provided (resistojet) specific density, and product composition to heat up the gas in order to provide a are listed by phase in Table 1. Projected higher acceleration and higher specific performance characteristics for the impulse while adding only slightly to system described propulsion systems are listed in complexity. The heater system requires a Table 2. power supply, which might be prohibitive for specific applications. In order to obtain a more complete comparison of the different concepts, other The propulsion system selection requires a criteria were taken into consideration and detailed analysis of each mission. A mission will be discussed. Some are quantitative, that requires a large AV for trajectory such as duty cycle, pulse response, cost, corrections, for example, will require a and development status. Others are more system that can provide a large total subjective, e.g. safety and reliability and impulse. This clearly favors a higher will not be addressed here. specific impulse thruster. Response time GASEOUSPROPELLANTS the absence of propellant slosh, liquid/gas Cold Gas. A cold gas (nitrogen or helium) is separation problems, and potentially simple stored at high pressure (~21 MPa). A injector design. Disadvantages of the use of regulator brings the pressure down to the gaseous bipropellants in low thrust desired pressure (0.1-1 MPa) after which applications include the poor mixing the gas is fed by means of check valves to efficiency due to short residence time in the a nozzle. The advantages are safety, low combustor, material incompatibility with cost, rapid response, and a simple system oxidizer, and high temperature material with high reliability. Although the thrusters issues,s Mixing problems limit the lower are low in weight, the overall system weight level of thrust that can reliably be obtained will be moderate due to the need for high to 0.1 N. The low propellant density requires pressure tanks and a distributed feed relatively large tankage. As a result, system. The system is volumetrically gaseous bipropellants are limited to inefficient due to the large tankage needed moderate total impulse missions. A higher for gaseous storage. Typical thrust levels degree of complexity than the cold gas range from 0.005 to 250 N. 3 Specific thrusters is inherent and this reduces impulse depends on thrust level and varies reliability. Poor mixing can cause a pulse to from 45 to 75 sec at 295 K for a nozzle area pulse variation in both magnitude and time ratio of 100, due to increasing nozzle losses history of thrust pulses, which is undesirable for decreasing thruster size. This system for many applications. has been used extensively to provide spacecraft and launch vehicle attitude LIQUID PROPELLANTS control (ACS) and divert propulsion, which Vapodzin.q liquids. These thrusters use is used to change a vehicle's direction." liquid propellants that can be stored over a pedod of years. The liquid vapodzes into a Tddyne TM propellants. A tridyneTM system consists of a nondetonable mixture of low molecular weight vapor (water, hydrogen, oxygen, helium, and/or nitrogen, ammonia or propane) with a high enough which is passed over a catalyst bed to vapor pressure to provide acceptable release energy. 4 The released energy heats performance. Heat of vaporization is up the bulk helium or nitrogen gas which is supplied by the bulk liquid, whose expelled through a nozzle creating thrust. temperature can be maintained from Typical chamber operating temperature and spacecraft structure or using electric pressure are 1100 K and 0.07-0.35 MPa, heaters. A theoretical I_o for ammonia of respectively. Thrust levels of 0.04 N have 100 sec can be projected at 293 K with a been obtained with a specific impulse of 138 nozzle area ratio of 50. Thrust levels vary sec for an area ratio of 100. The from 10"4-0.5 N.z Liquid systems are more advantages are a higher specific impulse complex as the result of a liquid control than cold gas systems and that the system mechanism, a vaporizer, propellant lines, weight compares favorably with cold gas and valves. The separation of gas and liquid thrusters for a given total impulse. At the in the absence of gravity or satellite spin is same time, all the simplicity of a cold gas an additional concern. In some cases, a separate plenum chamber is used for vapor thruster is retained. Tddyne TM systems have not been used as flight thrusters, although storage. Thrust decay will occur for liquid they would be applicable to spacecraft using vapor thrusters because vaporization cold gas propulsion. typically does not occur as rapidly as the Gaseous bipropellants. Gaseous fuel and propellant is expelled. Because the heat of oxidizer are injected, mixed, and reacted in vaporization is obtained from the liquid, the the combustion chamber and exit through pressure decay during long periods of thrust the nozzle. Advantages of these systems depends on the ratio of thrust to propellant are the high performance (with oxidizers mass and to the power input if an extemal such as oxygen or fluorine (O2/F2), and fuels source is used to provide heat. such as methane, ethylene, or hydrogen Liquid Monopropellants. A liquid propellant is passed through a catalyst bed. It (CH4]C2H4/H2)) with a theoretical Isp of 350- 400 sec for an area ratio of 100. There is no decomposes into a high pressure, high need for an external pressurization system temperature gas, which is expelled through and there are no freezing or boiling point a converging-diverging nozzle. The majority difficulties. Additional advantages include of monopropellant systems use hydrazine (N2H4), although hydrogen peroxide (H202) these systems is the volatility and toxicity of was used extensively in the past. Products the propellants. CIF3 is extremely damaging species are hydrogen, nitrogen, water, and to skin and eyes and reacts violently with ammonia. Monopropellant performance at organic materials. Cleanliness requirements microthrust levels is slightly lower than for storage are stringent. Safety problems bipropellants but with greater reliability and and the increased cost as the result of at lower complexity and cost. Thrust levels handling and servicing make these type of are upwards of 0.1 N with an Ispof 206-235 propellants unattractive for microthrust sec for area ratio of 100. 7 Other advantages applications. are a fast response (<10 msec) and unlimited restart capability. Because SOLID PROPELLANT catalyst deterioration is a function of Solid gas generator compounds that propellant flowrate, an added benefit at very decompose into low molecular weight gases low thrust levels is the excellent catalyst are used in microthrust applications. Gas durability. A disadvantage is the toxicity generator compounds can be classified into (hydrazine) and/or volatility (hydrogen subliming, photo-decomposing, and peroxide) of the propellants, which exothermally decomposing materials. adversely affect the cost by an increase in Subliming compounds vaporize by required safety procedures in handling the absorbing heat from the bulk propellant or propellants and servicing the propulsion from the spacecraft. Photo decomposing system. Alternative propellants are being materials react under exposure of ultraviolet investigated that provide equal or better (UV) radiation by absorbing light from the performance which are much more benign.8 sun or UV source. Exothermic compounds Liquid Bipropellants. Bipropellant engines decompose by absorbing energy provided with liquid phase injection operate at thrust by the breakup of the propellant itself. levels above 1 N. Propellant combinations are nitrogen tetroxide/hydrazine (N204/N2H4) Sublimin,q thrusters. Subliming thrusters use or nitrogen tetroxide/monomethyl hydrazine solid propellants that can be stored over a (N20_VIMH). Specific impulse of flight type period of years. Two types of subliming thrusters is typically 295 sec for a 100:1 solid thrusters have been considered, area ratio.9 At lower thrust levels, injector valveless and valved. In both types, problems occur as the result of the low flow propellants sublime into low molecular rates. Mixing is typically not well controlled weight vapor with a high enough vapor and orifices tend to plug. Other problems at pressure to provide acceptable performance microthrust levels are low reliability and at workable propellant feed system poor pulse performance. Liquid bipropellant conditions and nozzle size. Types of rockets cover the range of higher thrust and propellants are usually carbonates (X(HCO3) higher total impulse missions. or carbamates (X(COz,NHz)) subliming into NH3, COz, and H20. '_ Subliming thrusters A different type of liquid bipropellant are the simplest microthrusters from a thruster employs vaporizing liquid systems perspective. As the result of the propellant. Injector and orifice problems low storage pressure and lack of propellant associated with the liquid phase are absent. lines and valves, the overall system weight Some vapor control is needed to obtain is very small and the reliability is very high. desired mixture ratios, and gas-liquid Additional advantages are its overall separation technology is needed in the simplicity and a high propellant bulk density, absence of satellite spin. These types of resulting in a reduced system size. thrusters have a high degree of complexity Disadvantages are the low thrust level, low and a low reliability. Some propellants are performance and limited duty cycle. not hypergolic in the liquid phase but are Additional problems occur due to hypergolic in the vapor phase, such as recondensation of the gaseous products chlorine trifluoride (CIF3) as oxidizer and inside the propellant lines unless propellant methyl amine (CNHs) as fuel. Both lines are kept at sufficiently high propellants have acceptable vapor pressure temperature. for injection. Thrust levels for these types are from 1 - 50 N. Theoretical specific In valved versions, 11 solid propellants are impulse is 336 sec. The major drawback of stored in solid-vapor equilibrium conditions in a closed light-weight tank. A valve in the propellants are the hazards involved. propellant feed line from the tank to the Ammonium azide is a strong bloodpressure nozzle controls the thrust. Opening the depressant, and needs to be handled valve produces thrust by expulsion of the carefully. All azides need to be stored away vapor through the nozzle. The solid from water and carbon dioxide, as these sublimes to make up for the lower vapor substances tend to decompose the azide, pressure. Heat of sublimation is provided by forming hydrazoic acid, a strong explosive, the bulk propellant, which needs to be as a byproduct. thermally controlled by heat from the spacecraft structure or external heater. Performance of subliming thrusters can be Theoretical Ispof 80-85 sec for an area ratio improved by heating the gases just of 100 are possible with simple system upstream of the nozzle (resistojet). This design. Because the vapor pressure requires a more complex system. A determines the thrust, the thrust will decay theoretical performance improvement from dudng a pulse depending on how fast the 130 sec to 245 sec I_ can be obtained for a solid will replace the vapor. Thrust levels photochemical microthruster using are low, but pulse response is high. ammonium azide as propellant, when the products are heated to 1100 K. This option A typical valveless thruster consists of a is not viable for carbon containing vapors solid propellant inside a chamber with an because carbon deposits tend to open connection to the nozzle. Between accumulate inside the heating element and uses, the propellant is self cooled and the cause arcing to occur. vapor pressure is negligible. When needed, sunlight or an external heat source sublimes HYBRID PROPELLANT the solid, creating a vapor pressure that A hybrid thruster typically consists of a solid provides thrust. Thrust levels from 103-10 1 fuel grain (hydroxyl-terminated N and specific impulse from 50-70 sec are polybutadiene (HTPB) or polyethylene) low, system response is slow, and the power inside a combustion chamber through which needed for heating the propellant is 2-3 kW a liquid oxidizer (LOX, H202) flows. 1_14 per N thrust. Advantages are that these devices are throtUeable, restartable, and have a specific Photochemical rocket. UV light (<2537 A) impulse close to storable liquid from the sun or from an artificial source bipropellants. They are also relatively safe causes the propellant to decompose, from an energetic point of view (virtually providing thrust. 12The propellant must have nondetonable). Other advantages are the a useful decomposition rate. Some such small volume (due to high propellant propellants are the azides (ammonium density), low cost, and the simple system azide NH4N3), oxalates, permanganates, design (i.e. a single liquid). Projected perchlorates, and organic solids such as specific impulse is 290 sec for a 400 N 1,3,5,7 tetranitro-l,3,5,7 tetrazacyclooctane thrust engine. This type of device has not (HMX). Typical products for the ammonium been considered for smaller applications to azide are ammonia (NH3) and nitrogen (N:) date. with molecular weight of 22.5. Advantages are a very low system weight, reliability and ELECTROCHEMICAL long life. Theoretical performance Electrolysis. In electrolysis systems, a liquid calculations give an Ispof 130 sec at a thrust (usually water) is decomposed by means of level of 10-4N. Depending on the propellant an electric current into its elemental gases. type, the response can be in the order of These gases are used as propellant in a msec. standard gas/gas chemical thruster. Advantages include high theoretical specific A major disadvantage of all subliming impulse (390-425 sec), low weight as the thrusters is the tendency of the vapor to result of absence of pressurization system, condense or solidify inside the propellant and a high propellant storage density. Rapid lines or nozzle. This is especially true for pulse response and a wide range of thrust carbon biproducts. It can cause a levels and impulse bits can be achieved. deterioration of performance and limit life. Additional advantages lie in the dual use Disadvantages of the azide based applicability of the electrolysis unit. As the resultof the reversibilityof the electrolysis general guideline for applicability. For pdnciple,thesystemcanbe usedasa fuel typical missions, important parameters are cellto providepowerfor cases where solar duty cycle (percentage of time the thruster arrays are not sufficient. Solar energy can can be operated), minimum impulse bit size be stored in the electrolysis products and repeatability, and steady state whenever there is excess power available. performance. This energy can be recovered when needed by reversing the electrolysis process. By A simple schematic for each of the systems replacing the batteries with an integrated described in the previous section is shown electrolysis/fuel cell unit, large weight in Figure 2. An attempt was made to savings can be obtained for specific account for all the components required for missions,is operation as well as to suggest a potential propulsion system configuration. Pressure Disadvantages are the limited duty cycle, transducers and thermocouples have been caused by the slow electrolysis process, the lef_ out of the schematic for clarity but are need to dry the electrolysis gases, and the included in the weight analysis that follows. need for a specialized electrolysis cell. The overall system complexity is high relative to Advanced Monopropellants are targeted to other candidates. Reliability of the fill the high thrust, high total impulse electrolysis system is high, however, as the requirements of microthrust applications. result of a long history of use in submarine Propellants currently under investigation and commercial applications. were originally developed as gun propellants." The most common oxidizer in Promising Concepts for Microthrust the proposed monopropellant mixture is Application HAN (hydroxyl amine nitrate), a salt which is dissolved, together with a fuel, in water. Each of the above described chemical One of the fuels investigated is TEAN propulsion concepts can be adapted for (triethyl amine nitrate). The theoretical microsatellite propulsion. The selection specific impulse of a HAN/TEAN mixture process of an optimum propulsion system with 20% water is nominally 253 sec. for specific applications requires mission Projected thrust levels are from 0.4 N to 500 analysis. Some simple guidelines can be N. Below this lower limit, problems are used, however, to downselect to a few expected with the low flowrate over the concepts, without considering specific catalyst bed. Problems with combustion at missions. Major evaluation criteda are rocket chamber pressures below 3 MPa weight, size, cost, reliability and forced the search for new fuels such as performance. Other considerations include DEHAN (diethyihydroxyammonium-nitrate). propellant charactedstics (toxicity, volatility, storability), system complexity, Water in the advanced monopropellants development status, operable thrust levels, acts as a diluter to prevent explosive dual mode adaptability, need for thermal hazards. The lowest acceptable water control of stored propellants, and spacecraft percentage is determined by safety, contamination by exhaust products. typically 10%. A higher water percentage implies a lower performance. For the For the purpose of this paper, advanced propellant combinations currently under

(liquid) monopropellants, tridyne TM investigation, a typical water percentage is propulsion, electrolysis, and exothermic 20%. This higher percentage will reduce decomposing solid (gas generator) performance somewhat but will also cause propulsion are described in more detail. the decomposition in the catalyst to occur at Cold gas thrusters and state-of-the-art lower temperatures, avoiding critical high monopropellant thrusters are used as a temperature catalyst and combustor baseline for comparison, because they are material issues. simple, reliable and are in common use. Potential advantages over hydmzine Projected ranges of thrust versus specific monopropellants are the higher anticipated impulse for the chosen systems are shown performance, higher density, and the in Figure 1. The ranges indicated are a reduction of thermal control of the propellants (HAN/TEAN freezes at 200 K, controls the propellant flow rate to the hydrazine freezes at 273 K). More thruster. The propellant passes over the importantly, the non-toxic nature of the catalyst bed where the exothermal reaction propellants eliminates costly loading and heats up the gas. The products are expelled handling procedures required for hydrazine. through a diverging nozzle, providing thrust. The range of thrust levels and total impulse are projected to be similar to state of the art Advantages of the tridyne TM system have monopropellants. Monopropellant rockets in been described in the propulsion concept general perform better at higher duty cycles review section. A potential disadvantage is and at higher thrust levels. As a result of the that the impulse bit response is not as rapid good specific impulse, higher total impulse as for cold flow thrusters due to the missions usually require monopropellants. residence time of the propellant in the catalyst bed. Potential problems associated A potential single stdng system schematic is with catalyst deterioration as the result of shown in Figure 2. A blowdown system is cracking of the catalyst material during used for its simplicity and minimum weight. thermal cycling were not considered in this The liquid monopropellant is stored in a high analysis. These problems depend on the pressure tank, separated from the propellant flow rate over the catalyst. Lower pressurizing gas by means of a diaphragm. flow rates are beneficial to catalyst life. The pressurizing gas feeds the Problems can also be reduced by thermal monopropellant to the thruster via propellant management of the propellant. This may lines. A pyrovalve is used to prevent require some power and a slightly more propellant loss for long storage times and to complex system. activate the propulsion system when needed. A filter is used to prevent particles Minimum impulse bit available for the from clogging the injector or contaminating Tddyne TM is projected to be larger than for the catalyst bed. A latch valve controls the cold gas thrusters. Tddyne TM thrusters are propellant flow to the catalyst bed. The expected to operate over the full range of propellant passes through the catalyst bed duty cycles, depending upon the thrust where it decomposes into high temperature level. The applicable range of total impulse products. The gas is expelled through the is limited as a result of the lower specific nozzle, providing thrust. impulse as compared to mono- or bipropellant thrusters. The propellant Further details on the status of advanced volume required for a large total impulse monopropellant development are given in mission is prohibitive. Reference 8.

Tddyne TM systems can also be used as

Tddyne TM Propellants. Tddyne TM propellants pressurization systems for blowdown of are targeted to compete with cold gas jets liquid monopropellants, bipropellants, or for a wide range of applications, increasing hybrid rockets.TM the envelope of total impulse that can be covered with a simple, single gas. Hydrogen Electrolysis was used in the Gemini and and oxygen are mixed with nitrogen and/or Biosatellite space programs. It is a mature helium as diluent. This gives a concept that has been refined to provide nondetonable mixture that, if passed over a high performance with the safety inherent in catalyst bed, heats up by means of the water as propellant and the reliability hydrogen-oxygen reaction. The choice of proved with years of commercial nitrogen or helium depends on the behavior applications._7 Measurements on a 0.44 N of the catalyst bed. thruster showed an experimental specific impulse of 331 sec with a molybdenum A simple single string system schematic is combustor. TM High temperature combustion shown in Figure 2. The propellant is stored chamber materials (such as refractory at a high pressure (21 MPa - 2 MPa) in a metals) could improve that number by 20 propellant tank. A pyrovalve prevents sec by reduced film cooling needs and propellant loss and is used to activate the enhanced mixing.19 system. A filter installed in the propellant line clears particles and a latch valve A simple electrolysis system schematic is Exothermic decomoosing solid oropellants. shown in Figure 2. Water stored in a light Subliming compounds have been weight tank is fed to the electrolysis unit. exhaustively investigated for utilization in Hydrogen and oxygen products from the gas generators. Thrust levels obtained electrolysis are stored, either in separate depend on the vapor pressure of the tanks, or as pressurizing gas in water tanks. subliming compound, usually less than 0.1 For simplicity and low weight, a blowdown N over a short period of time. The duty system is used for the propulsion section. cycle depends on the thrust level, but is Check valves prevent the gases from rather limited because sublimation replaces flowing back into the electrolysis unit. the expelled gas slowly. As the result of the Filters clean the gases from particles and low specific impulse, only low total impulse latch valves control propellant flow to the missions can be served with subliming thrusters. solids. Response is only limited by valve opening time. Minimum impulse bit and Electrolysis/fuel cell round-trip electrical repeatability are excellent. efficiencies (electrolysis followed by fuel cell) currently are between 47% and A compound has been developed for gas 66%. 17'2° Such a system can replace generator applications 21 that has the currently used batteries at a great reduction potential to be used in decomposing solid in weight. A carefully designed single string rockets with the simplicity of subliming system can provide the energy storage and rockets but with a higher projected thrust fulfill many of the on-board propulsion tasks, level, higher I=p, higher propellant density, This unified approach can save substantial and higher impulse bits. The chemical weight and reduce overall spacecraft composition of this heterocyclic compound complexity. While recently developed with ring structure is C2H4NsO2.The heat of Lithium-Ion batteries have better energy formation is 39.2 Kcal/mole. Theoretical densities than SOA NiH2 batteries, they are performance of this monopropellant is 265.2 still in an early stage of development. sec for area ratio of 100. The propellant density is 1.86 g/cc, promising compact Whether the dual propulsion and power packaging. The compound is non-toxic, non- capacity will actually be used depends on carcinogenic, and does not form fragments the application. For very short discharge or particles during decomposition, and times (as in LEO orbit, where the cycle therefore does not present a health hazard. between dark- and sun- cycles is relatively It is also not detonable on impact and not shod), batteries have a higher energy electrostatically sensitive. Pdma_j products density (25 Whr/kg for NiCd vs. 12 Whr/kg are hydrogen, nitrogen, carbon dioxide and for fuel cell). Because the electrolysis/fuel carbon monoxide, with an average cell is available for propulsion, a weight molecular weight of 21.2 g/mole. Projected savings can still result if the battery is thrust levels vary from 0.001 to 10 N. replaced. Ignition can be accomplished with a diode For missions with increasing discharge laser. After ignition, the process is self- times, the weight savings become sustaining above 0.7 MPa by heat of increasingly large. The energy density of a decomposition. Decomposition can be fuel cell for a solar probe, for example, terminated by a sudden depressurization. could exceed 1000 Whr/kg, as opposed to Several scenarios can be designed for a 200 Whr/kg for the best available Lithium restartable solid monopropellant. Ion batteries. A" pill-dispenser" type will place a pill of the The duty cycle of electrolysis units varies compound at the head end of the thruster. A depending on the thrust level, the size of the diode laser initiates decomposition. The electrolysis cells, and the total power input solid decomposes at 1900 K and is expelled available. Response is rapid and impulse through a nozzle, providing a fixed impulse bits are repeatable. For the blowdown type bit, determined by the pill size. A second system that is considered for its simplicity, design, shown schematically in Figure 2, steady state thrust will decay over time due consists of a small diameter rod of solid to the drop in upstream pressure. propellant which is pushed with a spring mechanisminto the chamberthroughthe Sun-synchronous Picostar Mission centerof theheadend.Thediodeinitiates Many satellite manufacturers are offering decomposition.Initially, the decomposing extremely small spacecraft platforms for surfaceareaislargeandthepressurebuilds customer defined missions. With a dry bus uprapidly,sustainingdecomposition.As the mass of 12.5 kg and a typical payload surface area decreases and recesses into capability of 5 kg, the Orbital Sciences the head end of the thruster, the mass flow Corporation' s (OSC) Picostar is an example through the nozzle exceeds the mass of such a small spacecraft. The spacecraft addition due to decomposition. When the can provide approximately 10 watts of chamber pressure drops below a cdtical power for the payload from its 20 watt solar limit, decomposition is terminated. The array and 50 Whr NiCd battery system. motor can be restarted by pushing the rod Attitude control is provided by active further. Impulse bits are determined by the magnetics during its three year lifetime. amount of propellant decomposed. Picostar is proposed as a secondary Projected response times are rapid. payload for Pegasus. Consequently, the Preliminary tests show that a 170 mg size Picostar customer must be satisfied with the pellet of the compound decomposes in 7 primary payload's final orbit, or use an on- msec and does not leave any particles. Gas board propulsion system to attain the analysis shows that the pdmary products are desired orbit. For the mission assumed 46.5% N2, 29% CO2, and 0.5% 02, with the here, the Landsat spacecraft's zz 705 km remaining species H20, H2, CO, and CH4 circular, 98.2 = sun-synchronous operating not measured. orbit is arbitrarily assumed as the final orbit. This must be achieved after insertion into a Mission Trade Studies 550 km, 97.5 ° circular orbit by the Pegasus launch vehicle planned for 1997. Secondary Performance parameters, such as thrust payload mass and volume are available for level and specific impulse, often do not fully this launch. Assuming a mission starting as describe the benefits of a specific a secondary payload launched to this orbit, propulsion system as applied to a given the Picostar is required to change its orbital spacecraft mission. To fully understand the altitude and inclination to reach its benefit of a new propulsion system, a hypothetical 705 km, 98.2 ° circular comparative mission analysis must be operating orbit. Assuming impulsive performed. As a start in this direction, Hohman transfers for the monopropellant preliminary evaluations of electrolysis and system, the required mission AVs are 62

tddyne TM for selected missions were m/s transfer from 550 km circular @ 97.5 ° performed. Advanced monopropellant to 550 km x 705 km @ 97.84 °, and 63 m/s thrusters fall in the domain of another paper transfer from 550 km x 705 km @ 97.84 ° to (Ref. 8), and the solid gas generator 705 km circular @ 98.2 °. Impulsive multiple concept is in too early a stage to warrant bum transfers are used for the electrolysis inclusion in mission trade studies at this system. This requires approximately the point. same AV. Added to these AV requirements is an assumed 20 m/s of orbit correction per Based on the simple schematics of Figure year for three years. Thus the total mission 2, a weight analysis was obtained for AV requirement is 184 rrds for each system. specific mission descriptions. For the missions described in this section, tridyne TM Propulsion systems considered for this and electrolysis propulsion are compared to mission are state of the art monopropellant baseline cold gas and state of the art and electrolysis systems. Each of the hydrazine monopropellant propulsion potential propulsion configurations assumes systems, respectively. The examples a single thruster mounted along the demonstrate that either a decrease in spacecraft axis. Picostar is controlled using launch mass or an increase in payload mass torque coils to contro| its spin stabilization. It is achievable with either the electrolysis or is assumed that the Picostar will be able to tridyne TM concepts, within each of the point its thrust axis in the appropriate mission definitions. direction to apply the orbit transfer and maintenancebums.Theimpactonthe bus which 0.45 kg is already accounted for in subsystems(e.g. thermal and attitude the propulsion system trade off. Replacing control)of addingapropulsionsystemto the the battery, therefore, provides another 1.6 Picostarshouldbe consideredsimilar for kg in weight savings, for a total of 3.4 kg both propulsion system choices and, thus, is weight savings (or 16%). not considered here. Multiple bums for the transfer may be necessary depending on Reliability of electrolysis is projected to be the thrust level chosen for the propulsion high as the result of long term commercial system. use of electrolysis devices. 17 Whether the attractive option to unify the propulsion and The required propulsion system wet masses power system is used depends mostly on for the SOA monopropellant and electrolysis engineering decisions. The integrated unit are shown in Table 3. The payload and bus will be more complex, and will have a mass were set to 5 kg and 12.5 kg, conversion efficiency of only 47%. Also, respectively. Table 3 shows that the integration of propulsion and power might electrolysis based propulsion system be seen as a liability for some missions. provides a weight advantage of 1.8 kg (or 8%) over the monopropellant system. The ORBCOMM Mission required fuel volume for the electrolysis Another small satellite launched by the concept is greatly reduced as compared to Pegasus is the ORBCOMM spacecraft. monopropellants, although this may be Based on the Microstar bus, this vehicle slightly offset by the size of the electrolysis provides low orbit communication service. system. Additionally, electrolysis can Two Microstar satellites are already in provide a significant cost savings as the operation. ORBCOMM is larger than the result of less expensive propellant loading Picostar both in size and mass (see Table and handling procedures. 3). A small gaseous nitrogen (GN2) cold gas system is used for initial orbit trimming and By using a commercially available 4.48 N spacecraft separation for multiple launches. thruster for the monopropellant system, 7the The required &V capability is reported to be 23 bum time is short enough to allow for the 11 m/s. Table 3 shows that replacing the orbit transfer of less than a day. This is not GN2 cold gas system with the tridyne TM the case for the electrolysis system which is system cuts the fuel mass and volume by limited by the H2 and 02 production rate by 42% and 37%, respectively. Tankage for the the available electrical power. Assuming an tridyne TM system is lighter as the result of input power to the electrolysis unit of 12 W the lower fuel mass. This will partly be (7 cells), the fuel production rate is 0.65 offset, however, by the presence of a mg/s during the sunlit portion of the orbit. catalyst bed in the tridyne TM concept. A Consequently, the total transfer time for the reduction of the tddyne TM system dry mass electrolysis system from 550 km x 550 km over cold gas for this mission is still @ 97.5 ° to 705 km x 705 km @ 98.2 ° is 22 realized. A comparison of both systems for days. this mission gives an overall weight of 1.3

kg for the tridyne TM vs. 1.7 kg for the cold The possibility exists to replace the standard gas, with similar complexity and fuel NiCd battery for this platform with the same handling problems. An added advantage of electrolysis system working as a fuel cell, in the tridyneTM propulsion system, especially an integrated system. The baseline battery for small satellites, is the smaller required has a battery capacity of 50 Whr. _ With the volume. current SOA battery power density of 25 Whr/kg 24this means that the battery weight Summary is approximately 2 kg. The mission analyzed here consists of approximately 33 minute An overview is given of propulsion concepts dark cycle and 65 minutes sun-cycle. A that have potential application for micro battery replacing fuel cell that provides a satellite propulsion, with thrust levels capacity of 10 Whrs during the 33 minute ranging from 10"s to 10 N. Major concepts dark cycle could provide all housekeeping classification was by the propellant storage and payload requirements. The weight of phase, i.e. gas, liquid, solid, or hybdd, any such a fuel cell is approximately 0.8 kg, of combination of the above.

10 Fourpropulsionconceptswereanalyzedfor 2. Myers, R.M., Oleson, S.R., Curran, F.M., their applicability to microthrust propulsion: and Schneider, S.J.: " Chemical and Electrical Propulsion Options for Small advanced monopropellants, tridyne TM, electrolysis, and solid gas generator Satellites," Proceedings of the 8th AIAA Utah State University Conference on Small propulsion. Major factors in the selection process were performance, weight, size, Satellites, Aug. 29-Sept. 1, 1994. cost, and reliability. Non-toxicity of the propellants used in each of the selected 3. Bzibziak, R.:" Miniature Cold Gas concepts provided an additional benefit in Thrusters," AIAA 92-3256, Nashville, July 6- terms of cost savings due to reduced 8, 1992. loading and handling problems. 4. Barber, H.E., Falkenstein, G.L., Buell, Two of the selected propulsion concepts, C.A., and Gumitz, R.N.:" Microthrusters Employing Catalytically Reacted Nz-Oz-H2 tridyne TM and electrolysis, were compared Gas Mixtures, Tr_dyne," J. Spacecraft and against currently used propulsion systems, cold gas and state of the art Rockets, Vol.8, No. 2, February 1971 monopropellants. As a potential mission for comparing electrolysis versus state of the 5. Schneider, S.J.: " High Temperature Thruster Technology for Spacecraft art monopropellant systems, an orbital transfer and plane change of an existing Propulsion," Acta Astronautica, Vol. 28, pp. microsatellite with subsequent orbit 115-125, 1992. maintenance for three years was analyzed. A tridyne system was compared against a 6. Genthe, D. : "Attitude Control Systems for cold gas system for orbit tdmming and Satellites," (In German) Deutsche For- spacecraft separation of an existing schungs- und Versuchsanstalt fuer Lufl- und commercial satellite. Raumfahrt, RPT#: DLR-MITT-70-14, 1970

Electrolysis propulsion provided a significant 7. Hydrazine Handbook, Rocket Research weight advantage over state of the art Company, Olin Defense Systems Group. monopropellants for the full mission, 1.6 vs. 2.7 kg wet mass. Most of the weight 8. Jankovsky, R.S. : " HAN-Based Mono- advantage was obtained because of higher propellant Assessment for Spacecraft," Isp, 350 vs. 223, and reduced propellant AIAA Paper 96-2863, July, 1996. storage weight. The non-toxicity of the propellants provided an additional benefit. A 9. Kaiser Marquardt data sheets KM Model potential added advantage, which was not 53 (9 N thrust), 1995. fully investigated, was the replacement of the baseline battery system with an 10. Hardt, A.P., Foley, W.M., and Brandon, integrated electrolysis/fuel cell system for R.L.:" The Chemistry of Subliming Solids for spacecraft power. This could lead to Micro Thrust Engines," Astronautica Acta, additional weight savings but would Vol. 11, No. 5, 1965. significantly increase complexity. 11. " Development of the Subliming Solid Control Rocket', NASA CR-712, Rocket For the mission analyzed for the tridyneTM Research Corporation, Seattle, WA, 1967. and cold gas systems, tridyneTM provided a significant weight and volume advantage over cold gas thrusters, while retaining the 12. Maycock, N., and Pal Vemeker, V.R.: " A Photochemical Microrocket for Attitude system simplicity. Problems associated with propellant service and handling for both Control," J. Spacecraft and Rockets, Vol. 6, systems are similar. No. 3, March 1969.

References 13. Sellers, J.J., Paul, M., Meerman, M., and Wood, R.: " Investigation into Low-Cost 1. Suthedand, G. S., and Maes, M.E., : "A Propulsion Systems for Small Satellite Review of Microrocket Techno/ogy:lG 6 to 1 Missions." presented at the 9th Annual AIAA/USU Small Satellite Conference, Ibf Thrust," J. Spacecraft and Rockets, Vol. Logan, UT, September 1995. 3, No. 8, August 1966.

1] 14. Wemimont,E.J., and Meyer, S.E.: 31st Joint Propulsion Conference, AIAA " HydrogenPeroxideHybridRocketEngine Paper 95-2401, July, 1995. PerformanceInvestigation,"AIAA94-3147, Indianapolis,IN,June27-29,1994. 20. Hoberecht, M.A., Miller, T.B., Rieker, L.L., and Gonzalez-Sanabria, O.D.:" Design 15. Mueller,J.M., and McFarlane,J.S., Considerations for a 10 kW Integrated " Design of Tridyne Pressudzation Systems Hydrogen-Oxygen Regenerative FUEL Cell for Liquid Oxygen/Polybutadiene Hybrid System," 19th Intersociety Energy Rocket Motors," AIAA-91-2406, Conversion Engineering Conference, San Sacramento, CA, June 24-26, 1991. Francisco, CA, Aug. 19-24, 1984.

16. Thaller, L.H.:" Electrochemistry and 21. Manser, G., GenCorp. Aerojet, Private Storage," Space Power, NASA Conference Communication. Publication 2352, April, 1984. 22. Wilson, A. (Ed.) :Jane's Space 17. " Unitized Regenerative Fuel Cell Directory, 10th Ed., 1994-1995. Energy Storage Systems for Aircraft and Orbital Applications," Hamilton Standard 23. Meurer, B.: " First Class Science on a Space and Sea Systems. Coach Class Ticket," 9th Annual Conf. on Small Satellites, Utah State Univ., Sept. 18. Stechman, R.C., Campbell, J.G. "Water 1995. Electrolysis Satellite Propulsion System," The Marquardt Company, Technical Report 24. Tuck, C.D.S. (Ed.) Modem Battery AFRPL-TR-72-132, January, 1973. Technoloav. EUis-Horwood, New York, pp 279. 19. Reed, B.D. :" Long Life Testing of Oxide-Coated Iridium/Rhenium Rockets,"

12 Propulsion Concept Propellants Products (MW) I densityIRefl3 l/cm

Cold Gas N2 (3000 psia) 0.225 1,3 N2(28)

Gas Bipropell. 02 (3000 psia) 0.257 1 H20(18), CO2 (44), CH4 (C2H4/H2) 0.129 co(28)

0.206 1 H2(2), 02(32), He(2) Tddyne TM H2, 02, He,N2 (3000 psia) 14

Vapodz. water (H20) 1.00 4,5 H20(18),NH3(17), ammonia (NH3) 0.611 C3H8(44) Pmpane(C3Hs) 0.585

12 Liquid Monopm- peroxide (H202) 1.39 H20(28),O2(32), pellants hydrazine (N2H4) 1.01 6 N2(28),NH3(17)

gaseous inj CIF3 1.77 1 C(12),HF(19), Bipropell. monomethylamine 0.699 HCI(35),H2(2), liquid inj. N;zO,dMMH 1.10 7 N_(28),CH4(16)

Subliming Carbonates (X(HCO3)), 1.38 1,8 NH3 (17),N2 (28) Carbamates, (X(CO2NH2)), SUBLEX-A (RRC) 0.69 9

Valveless Photochem Azides (X(HN3)), 1.346 10 NH3(17), N2(28) Solid Oxalates, 1.39 COs (44), OrganicSolids, H20(18) Permanganates, Perchlorates

Decompos. C2H4N602 1.86 H2(2),N2(28),CO2(44), CO(28), H20(18)

Water (H20), 1.00 H20(18) Camphor (CloH160), 0.990 CloH160(152), Valved Acetamide 1.159 CH3CONH2(59), (CH3CONH2), Naphtalene(CloHs), 1.025 CloH8(128), Biphenyl(C12Hlo) 0.8660 C1;_Hlo(154)

Hybdd GOX (3000 psia), 0.257, 11 H20(18), CO2 (44) polybutadiene 0.97 12 ((C4H6)n)

Ele_ro- Electrol. water (H20) 1.00 15 H20(18) chem. 16

Table 1: Propellant properties of chemical microthruster candidates.

]3 Propulsion Concept Performance F(N) I_ Resp Req. N N.sec msec I Ref. I kW/NEnerg.

Cold Gas 76e (50) 0.1-5 25OO 1, 3 0.03

Gas Bipropell. 420t (50) 10"2.10 250000 1 0.01

10_-10 5000 Tridyne TM 138e (100) 4,13 0.01

Vaporiz. H20, NH3, 100e (50) 105-0.5 30000 1,5 C3H8

Liquid Monopro- H202 165t (40) >2 50000 10 1,5 pellants N2H4 227= (50) >1 300000 10 6,7

gaseousi_ 310t (40) >1o-2 250000 10 1 Bipropell. liquid inj. 295e (100) >0.5 500000 10 1,7 315e(300)

Subliming 75e (100) 2500 1,8,9 2

Photochem 130t (60) 2500 10 Valveless

Solid Decompos. 262t (100) 0.1

Valved 75e (100) 10"s-10.2 2500 8,9

Hybrid 280e (80) 500000 11,12

Electro- Electrol. H20 331e(100) 0.01-15 150000 10 15,16 6 chem.

..==experimental,..r--theoreUcal.

Table 2: Performance characteristics of potential chemical microthrusters

14 Mission: Picostar to Landsat ORBCOMM Orbit Orbit and Trimming Maintenance

Cold Gas PropulsionSystem N2H4 Electrolysis Tndyne TM Monoprop

Fuel Density (g/cc) 1.00 1.00 0.225 0.206 11 11 Thruster AV (m/s) 184 184 138 Primary Thruster Isp(s) 223 350 80 # of Thrusters 1 1 1 1 43 43 Initial Mass (kg) 22.0 20.2 0.0 0.0 0.0 Cant Angle (°) 0.0 1.15 0.6 0.35 Fuel Mass (kg) 1.8

Propulsion System Dry Mass 2.68 1.6 1.08 0.94

Total Fuel Mass 1.8 1.1 0.6 0.35 Total Fuel Volume (m3) 0.00178 0.00107 0.0027 0.0017 Total Propulsion Syst. Wet Mass 4.5 2.75 1.7 1.3

Net Mass (Initial-Wet Prop.) 17.5 17.5 41.3 41.65 Spacecraft Internal Vol. 0.0063 0.0063 0.0014 0.0014

Primary Engine Thrust (N) 4.45 045 5.56 N 5.56 N

# Engine Thrusting 2 2

15 137 0.7 0.7 Primary Thruster Bum Time (min.)

Table 3: Picostar Potential Propulsion System Performance Comparison

I5 lO,_ _'_ ---- Advanced T

...... -r--_. _ Monopropellant /

..... Solids : 10,2 "" "" "" • m "e jI I

_--- 10"3 ...... -.---J...... 4 I ..... Subliming Solid I I -- Electrolysis I I

10"s _ -- r I 101 10 2 lO_ Total Impulse (N-sec)

Figure 1: Operating Range of different microthruster concepts

ActvQncect Tr_ctyne Decomposing Etectrotysis Monopropelton_ Solid

Pressur;z;ng Gos

pet!.on± _

Legenc_ r'l Regulator /-_ Nozzle

l_ Hanuot Votve I Propellon± Line

Pyro Votve [] Decomposition Chanl_er l_ Check Votve [] Comlous_:ion Ch(_mloer

In Ret;e¢ Volve _ Filter

R Lotch Volve I Sotid Propetlon± Rod I Q Tank

Figure 2: Simple Schematic of four chemical micropropulsion concepts.

]6

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4. TITLE AND SUBTITLE 5. FUNDING NUMBERS Chemical MJcrod.-uscr Opdons

WU-233-1B-1B 6. AUTHOR(S) C-NAS3-27186 Wim de Groot and Steve Olcson

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11. SUPPLEMENTARYNOTES Prepared for the 32nd Joint Propulsion Conference, cosponsored by A/hA, ASME, SAE and ASEE, Lake Buena Vista, Florida, July 1-3, 1996. Project Manager, Steven J. Schneider, Space Propulsion Technology Division, organization code 5300, (216) 433-7484. 12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

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13. ABSTRACT (Maximum 200 words)

Chemical propulsion systems with potential application to microsatellites are classified by propellant phase, i.e. gas, liquid, or solid. Four promising concepts are selected based on performance, weight, size, cost, and reliability. The selected

concepts, in varying stages of development, are advanced monopropellants, tridyne TM, electrolysis, and solid gas generator

propulsion. Tridyne TM and electrolysis propulsion are compared vs. existing cold gas and monopropellant systems for selected microsatellite missions. Electrolysis is shown to provide a significant weight advantage over monopropellant

propulsion for an orbit transfer and plane change mission. Tridyne TM is shown to provide a significant advantage over cold gas thrusters for orbit trimming and spacecraft separation.

14. SUBJECT TERMS 15. NUMBER OF PAGES ]9 Microthruster; Chemical 16. PRICE CODE A03

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