. a b' NASA Technical Memorandum 105263 Technical Prospects for Utilizing Extraterrestrial Propellants for Space Exploration Prepared for the 42nd International Astronautical Congress sponsored by the International Astronautical Federation Montreal, Canada, October 5- 11, 1991 TECHNICAL PROSPECTS FOR UTILIZING EXTRATERRESTRIAL PROPELLANTS FOR SPACE EXPLORATION Diane L. Linne and Michael L. Meyer National Aeronautics and Space Administration Lewis Research Center Cleveland, OH 44135 ABSTRACT materials are used (ref. 6). Such a reduction in launch mass would alleviate the economic burden of NASA's Lewis Research Center has space exploration. supported several efforts to understand how lunar and martian produced propellants can be used to their best Unfortunately, the exploration mission advantage for space exploration propulsion. A transportation studies are limited by incomplete discussion of these efforts and their results is information for technologies proposed for the future presented. A manned Mars mission analysis study missions (ref. 7). Many of the proposed production identified that a more thorough technology base for processes for extracting useful elements from the propellant production is required before the net lunar regolith have not been developed beyond the economic benefits of in situ propellants can be concept level. Mars' atmospheric processing determined. Evaluation of the materials available on concepts have been demonstrated only at a laboratory the moon indicated metal/oxygen combinations are level. Similarly, the operation of rocket engines the most promising lunar propellants. A hazatds using the sometimes unconventional propellants has analysis determined that several lunar metal/LOX not been demonstrated. Because of the scarce base monopropellants could be safely worked with in small for these proposed technologies, the accuracy and quantities, and a characterization study was initiated confidence level for the space transportation study to determine the physical and chemical properties of results is uncertain. potential lunar monopropellant formulations. A bipropellant metal/oxygen subscale test engine which To eliminate some of the uncertainties in the utilizes pneumatic injection of powdered metal is propulsion technology, efforts at the NASA Lewis being pursued as an alternative to the monopropellant Research Center continue to evaluate the benefits and systems. The technology for utilizing carbon technical prospects of utilizing in situ propellants. monoxide/oxygen, a potential martian propellant, has The work has focussed on two areas: expanding and been studied in subscale ignition and rocket perfor- updating the space transportation studies with the mance experiments. most recent data for technology assumptions, and building a data base for the unconventional INTRODUCTION Propellants which may be produced from indigenous lunar and martian resources. This report discusses The Space Exploration Initiative, as the results and status of these efforts. proposed by U.S. President George Bush and developed in studies for NASA and the National Space Council (refs. 1-3), outlines an ambitious plan BACKGROUND for establishment of a permanent lunar base and for manned exploration of Mars. Other countries are The makeup of the moon and Mars are also planning ambitious space exploration missions known to a different extent. The US Apollo and (ref. 4,s). The leading restraint on exploration USSR Luna missions returned lunar samples to Earth possibilities is the cost of launching large masses into that allowed for a detailed analysis of the lunar orbit. Space transportation studies have shown that regolith (ref. 8). The US Viking landers to Mars significant reductions in launch mass may be realized provided an analysis of the martian atmosphere and, if propellants produced from indigenous space to a lesser extent, the composition of the martian soil 1 (ref. 9). It is from these banks of data that one first Figure 1 shows a qualitative comparison looks to determine what is available at the moon and between several propellant options for the moon. Mars that has potential to be used as a rocket Although rocket engine specific impulse is usually a propellant. Once the potential propellant leading factor in the selection of a propulsion system, combinations have been identified, mission analyses several other factors play an important role in the can be performed to quantify the expected benefits. selection of an in situ propellant combination. Abundance in the lunar soil, ease of manufacturing, Available Resources simplicity of the engine system, and technical background and experience with the propellant The lunar regolith is comprised of minerals, combination will also be important criteria. the most common of which are olivine Currently, the database for production and propulsion ((Mg,Fe),SiOJ, pyroxene ((Ca,Fe,Mg),Si,Od, technology is incomplete, and a quantitative selection ilmenite ((Fe,Mg)Ti03), and anorthite (CaAlSi,08). of the "best" propulsion system to use for future These minerals c a ~b~e found in varying lunar activities cannot yet be made. concentrations in the highlands and the maria. Table I shows the amounts of some elements on the moon The results of the Viking missions show that based on analysis of the mineral samples returned to the martian soil is rich in magnesium, iron, and Earth. From the table, it is apparent that oxygen is calcium, and that the polar caps may contain water plentiful, and can be used as the oxidizer. Hydrogen ice or dry (COJ ice. In addition, the regolith and carbon, the two elements that are most potentially contains a permafrost of water ice at some commonly used in rocket fuels, are not present on the depth below the surface, but the depth, quantities, moon to any appreciable extent. Some of the metals and distribution have not been ascertained. Another that are present on the moon, however, have been important resource is the thin martian atmosphere. It used in solid propellant fuels and therefore have has an average pressure of 0.07 - 0.10 psia and is potential as lunar in situ fuels. approximately 95 percent carbon dioxide. It has been postulated that this source of carbon dioxide can be Table I. - Elemental Breakdown of Lunar dissociated into oxygen and carbon monoxide to use Regolith as rocket propellants. An alternative option is to chemically react hydrogen or water (brought from Earth, obtained from the martian permafrost, or II Element Weight obtained from the martian moons) with the Percent atmosphere directly to obtain oxygen and methane propellants. oxygen, 0, 42 Because the carbon monoxide and methane Silicon, Si 20 fuels are higher performing than the magnesium, Aluminum, Al 9 iron, or calcium fuels, and because the processing of the gaseous atmosphere would be a much less Iron, Fe 9 complex task than mining and processing minerals from the regolith, the fuel selection at Mars appears Calcium, Ca 8 to be between the carbon monoxide and methane. A Magnesium, Mg 4 trade-off between the lower performance of carbon monoxide and the necessity to either supply hydrogen Titanium, Ti 3 from Earth or mine it at or near Mars for the Phosphorous, P, Sulfur, S, -1 methane appears to be a key factor in the selection of Sodium, Na, Potassium, K, the "best" martian propellant system. Chromium, Cr Benefits of In Situ ProDellants The utilization of indigenous space materials for propulsion offers several tangible benefits such as 2 L reduction in initial launch mass, increase in payload exploration, and in situ propellants can also help delivered, or reduction in Mars trip time. The achieve this goal. Based on a specified mass in intangible benefits include the establishment of self- LEO, every kilogram of propellant that does not need sufficiency, reduction in mission complexity, and to be saved for use at Mars can be used to increase development of technologies with potential terrestrial the departure energy from the Earth. An increase in applications. the available Earth departure energy will allow the spacecraft to travel a more direct path to Mars, The most important benefit of in situ thereby decreasing the trip time. Maximum reduction propellants is the potential to significantly reduce the in trip time is achieved with no decrease in initial mass required in low Earth orbit (LEO). When mass or increased payload. However, two or all launch costs to orbit are counted in thousands of three benefits can be achieved together to a lesser dollars per pound delivered to orbit, a reduction in degree. the mass required from Earth can be translated to a cost savings for the overall mission. Figure 2 shows Other benefits can be obtained with the two figures taken from the literature (refs. 10,ll) that utilization of in situ propellants. If the ultimate goal depict the magnitude of the mass-in-LEO reduction is to expand human presence into the solar system for a lunar and a Mars mission. Figure 2a shows and beyond, exploration must evolve into settlement. that the use of lunar produced aluminumloxygen The ability to utilize resources available at the new propellants for near-lunar operations can reduce settlement and to reduce or eliminate the dependency mass-in-LEO requirements by 63 percent over the all- on the homeland is the ultimate measure of an Earth produced hydrogedoxygen baseline. Figure 2b independent establishment. At Mars, in situ shows that the use of Mars produced methaneloxygen propellant production may provide the ability to propellants for Earth return of a 5 kg sample can perform direct return missions, thereby eliminating reduce Mars injected mass by 67 percent compared the need to perform autonomous rendezvous in to the baseline case that used solid propellants for martian orbit. Finally, the development of the Earth return. Even more importantly, the figure necessary technology to mine and beneficiate the shows that producing the return propellants at Mars lunar soil and to autonomously dissociate the martian allows the mission to be accomplished with only one atmosphere has potential terrestrial applications. The shuttle flight instead of two. The assumptions used lunar technologies may help to produce metals more in these two mission analyses can be obtained from efficiently on Earth, and the martian technologies references 10 and 11, or from reference 6, which may help control the levels of carbon dioxide in the contains a summary of these two analyses and five Earth’s atmosphere. others taken from the literature that show significant launch mass reduction with in situ propellants. TECHNOLOGY PROGRAM AT LEWIS Another potential benefit of in situ propellant The Space Propulsion Technology Division utilization is increased payload capability. This at the NASA Lewis Research Center began benefit is a direct relationship: for every kilogram of investigating the utilization of lunar and Mars return propellant that does not need to be delivered to propellants in 1989. The objectives of the activity the destination planet, a kilogram of payload can be are to identify potential uses of in situ propellants and added (for a specified initial mass). Additional quantify the benefits, to determine likely propellant payload capability can reduce the total number of combinations and the status of their technology, and missions or can allow for the delivery of larger to establish the technology database that will be payloads. Maximum payload capability is achieved needed to develop engines that use in situ propellants. when there is no reduction in initial mass in orbit. All work to date has focused on chemical propulsion Alternatively, the two benefits can be combined, with options. To accomplish these goals, efforts have some decrease in initial mass and some increase in been concentrated in three areas, mission analysis, payload capability. propellant characterization, and propellant performance. Decreasing the trip time to Mars has emerged as a key requirement for manned 3 Mission Analvsis makes the success of the maneuver at Mars less certain. In situ propellants therefore offer a potential A contracted study (ref. 7) was initiated to to replace aerobraking at Mars. If aerobraking at assess the benefits of in situ propellants for a series Mars becomes a reliable option, it can be combined of manued Mars missions. The hvo main objectives with in situ propellant utilization to further reduce of the analysis were to determine the reduction in launch mass requirements. Figure 3b compares the initial mass in low Earth orbit (IMLEO) for various same baselines with options that travel to the moon to propellant combinations using a consistent set of pick up lunar propellants for the trip to Mars and groundrules and assumptions, and to determine the back. Although these options further reduce the cost of delivering and maintaining the in situ IMLEO, these results do not include the propellant production infrastructure. All in situ propellant production requirements. This further emphasizes the options were compared to a baseline scenario in need for accurate determination of the infrastructure which Earth-supplied hydrogen and oxygen was used. requirements before the actual economic benefits of in situ propellants can be determined. The first objective of the study was to determine the mass of the necessary infrastructure, ProDellant Characterization and the cost of delivery and maintenance. Unfortunately, the concepts for various production The presence of metals on the moon that methods on the moon have not been developed may be used in the solid phase poses unique completely. It was therefore difficult to establish technology challenges in the development of a rocket accurate estimates of the mass, power, hardware engine utilizing completely lunar indigenous resupply, and reagent resupply requirements for the resources. Several options exist to inject the metal lunar production plants. Although some Mars pilot fuel into the combustion chamber. One option is to plants have been tested (refs. 12, 13), mass and suspend powdered metal particles in the liquid oxygen power estimates for full-scale models are still by means of a gellant. This is similar to metallized uncertain. The results of the first phase of the study, fuel technology that has been performed in the past therefore, merely emphasized the need for further (refs. 14,15). The resulting mixture would be a technology definition of the plant mass, power monopropellant with non-newtonian flow requirements, and production plant reagent resupply characteristics. Safety issues and rheological requirements. properties must be resolved before the monopropellant can be tested in a rocket engine. An Figure 3 (ref. 7) shows the results from the alternative method to inject the metal fuel into the study for steady state missions (i.e. after the combustion chamber is to entrain the particles in an production plant has been delivered). The mission inert or fuel gas stream that would carry the particles assumptions were a 2016 opposition-class flight into the chamber. The small amount of gas stream profile and the delivery of 25 metric tons of useful needed for this method would either be obtained at payload to the surface of Mars. Figure 3a compares the moon or brought from Earth. A third method is the initial mass in LEO for the baseline case with and to use a solidhiquid hybrid where the metal fuel is without aerobraking and several Mars in situ formed into rods or cylinders through which the propellant options. (Note that the in situ propellant liquid oxygen flows. Research into this alternative options do not use aerobraking at Mars.) The results will not be discussed in this paper. show that in steady-state operation, the in situ propellant options reduce the initial mass in LEO by Monoprowllant Hazards Assessment. The approximately 50 percent over the all Earth-supplied monopropellant concept has the potential to be a option. More significantly, the in situ propellant hazardous material, and hazards assessment and options without aerobraking are comparable to the all propellant formulation must be completed before any Earth-supplied propellant option that uses combustion experimentation can begin. The ultimate aerobraking. While the dynamics of aerobraking at objective of the hazards assessment activity is to the Earth are relatively well understood, questions assign an explosive classification to the regarding high entry angles and velocities and monopropellant so that the required safe handling variable atmospheric densities during dust stom procedures will be known. A preliminary goal of the 4 hazards assessment is to test small, laboratory-scale tests are only an indication that small scale quantities for explosive hazards such that formulation formulations can be safely mixed. research can begin with assurances of safety. The second phase of the hazards assessment consisted of mechanical impact tests, where a weight The first phase, conducted at NASA White was dropped into a small sample of the Sands Test Facility, consisted of mixing tests, where monopropellant from various heights to determine the small amounts of aluminum, titanium, silicon, and necessary energy to cause a reaction (ref. 17). iron powder were combined with liquid oxygen and Because impact test results can vary due to then stirred at low speeds (approximately 600 rpm) differences in the test apparatus, two well- characterized materials were - used to act as Table 2. Summary of Results of Mechanical ImDact Tests points for comparison. Sample Approx. 50% Height Impact Impact Energy The first test O/F Ratio (cm) Energy Density material was (Joules) (Joules/cm2) pentaerythrit o1 ~~ NA 51.0 45.4 19.4 tetranitrate ( P E T N ) , Nitromethane NA > 123.0' > 109.4 >46.8 which is a Titanium 2.33 < 15.2b C 13.6 C 5.8 solid Class A explosive 2.33 67.6 3.1 25.5 known to be i m p a c t Aluminum 2.33 > 123.0' > 109.4 >46.8 sensitive. Silicon 2.03 > 123.0' > 109.4 >46.8 Nitromethane, which is a 2.03 > 123.0' > 109.4 >46.8 flammable at highest height available o, apparatus liquid, was bReaction occurred 100 96 at -lowest height available on apparatus also used because the ox ygenlmetal mixtures were while being monitored for any signs of chemical more liquid than solid and nitromethane is known to reaction. A total of 63 tests were conducted that detonate under certain shock conditions. The varied the metal, metal to oxygen ratio, and presence materials tested in liquid oxygen for the and type of gelling agent (ref. 16). Figure 4 shows monopropellant impact tests were aluminum, a typical temperature versus time trace taken from an titanium, silicon, iron, 80 96 aluminud209 6 iron/oxygen test. The absence of any large magnesium alloy, aluminumlgellant, and temperature spikes indicates that no chemical reaction titaniudgellant. occurred. Metal particles were also analyzed chemically before and after mixing to verify that no The results were reported in terms of a 50 metal had been oxidized. There were no reactions percent height, which is the height at which a observed in any of the tests. It must be stressed that reaction occurred 50 percent of the time. Table 2 the results of these tests can be applied only to the (ref. 17) lists the results of the mechanical impact qwtntities and conditions that were tested. These tests. For all of the powders except titanium, the results can not be extrapolated to larger scale batches, results of the impact tests indicate that it is safe to nor can they be used to assess the hazards of mixing handle the monopropellants in the quantities and performed at higher shear rates. Therefore, these manners necessary to begin formulation and characterization of the monopropellant. The Al/Mg 5 alloy monopropellant should be treated as a secondary the monopropellant bum rate when it is contained in class A explosive (the classification of PETN) until an insulated line expected in an operational engine. more assessment tests can be completed. Pneumatic Powder Fuel Feed System. An Some of the tests that still need to be alternative approach to the metal/LOX performed to fully classify the monopropellants monopropellant is also being pursued. This option is include high-speed stirring and rotary friction, pneumatic injection of the powder into the thrust electrostatic discharge, water hammer, and detonation chamber. A camer gas is used to transport powder sensitivity. In addition, many of the tests would need from the fuel storage device to the injector. Injector to be repeated with larger quantities to determine the designs (impinging or coaxial), which are commonly scaleability of the results. used for gas mixing, can be used to disperse the powder and mix it with gaseous oxygen. The camer MonoDroDellant Formulation and gas may be either minimally reactive (e.g. nitrogen, Characterization. With the encouraging results from helium) for system safety or an additional fuel (e.g. the first phases of the hazards assessment program, hydrogen, methane) to enhance performance. Since a contracted effort (ref. 18) was initiated to prepare gases other than oxygen are only available in limited small formulations of the metal/oxygen amounts on the moon, the camer gas will probably monopropellant and to test some of the basic be brought from earth. Thus, the amount of camer physical, chemical, and rheological properties. The gas used must be minimized to reduce the system’s objective of the contract is to determine the minimum dependence on earth-based propellant. Related amount of gellant needed to prevent settling of the studies suggest the flow rate of gas could be limited metal powder but still allow for acceptable flow to as little as 1.0% of the metal mass flow (ref. 19). properties. The formulation experiments indicate that However, an empirical data base to predict the aluminum and silicon can be stably suspended in physical characteristics of a two-phase gas-solid fuel liquid oxygen with only one to two weight percent flow must be developed. gellant (amorphous fumed silica). With this amount of gellant, very little settling of the metal powder was Pneumatic conveying of dust or granular observed during the first 24 hour period. In general, materials is common in commercial industries, but the majority of the settling in such a gelled mixture these pneumatic systems cannot be directly applied to will occur in the first 24 hours after formulation. a rocket fuel feed system. A novel system will be required to operate at high pressures, minimize the A second objective of the formulation and amount of camer gas used, and accurately control the characterization task was to determine the ambient solids flow rate. Four types of pneumatic feed and pressurized bum rates of the monopropellants. systems were considered. The simplest device was a If the monopropellant bums faster than the injection fluidized bed. This is a vertical column with a bed of velocity into the chamber, then burning could powder supported by a gas distribution plate. Gas is propagate into the feed lines and the propellant tank, forced up through the powder, mobilizing the causing catastrophic failure. The ambient bum rate particles and giving the powder the appearance of tests were conducted with the monopropellant being fluid. Some of the powder is entrained into the submerged in a liquid nitrogen bath to prevent boil- flow and camed away from the bed. Another device off of the liquid oxygen before the start of the test. considered was a screw or worm feed system which During the test, this liquid nitrogen acted as a strong would use a helical blade to carry powder from a heat sink and absorbed the energy created by the hopper to either a camer gas line or the thrust combustion of the monopropellant. Because of this chamber. A third feeder option was a fluidized rapid transfer of heat, the monopropellant combustion hopper. Although this is similar to a fluidized bed, was unable to sustain itself after the solid propellant it differs in that the powder is drawn from the core of ignitioncharge was removed. Therefore, the ambient the fluidized particles. In this manner, a more dense pressure bum rate of the monopropellants in the stream of particles is extracted. The last entrainment presence of a sufficient heat sink approaches zero, device considered was a fluidized piston. The assuring that the flame will not propagate into the powder would be placed in a cylinder and pressurized feed lines. Further tests will be needed to determine with the piston. Gas injected near the piston exit would help expel the powder and carry it to the thrust Lunar Metal Propellant. The potential chamber. performance of a lunar metalloxygen rocket engine is the subject of much debate. The Chemical A pneumatic feed system for powder metal Equilibrium Composition computer program (ref. 20) fuels was evaluated to define powder entrainment has been used to calculate the ideal performance of issues in a fluidized bed. In a cold flow experiment, metal/oxygen propellants, but there are many non- the operation of a fluidized bed was studied with ideal losses which cannot be accounted for by this nitrogen fluidizing powdered metal oxides. The program. From aluminized solid propellant results indicated that a fluidized bed would require experience, it is known that the metal oxide large amounts of gas to entrain a powder fuel, and combustion products will condense into solid particles bubbling and slugging would occur in the particle which do not maintain velocity and thermal bed. Bubbling is the formation of small voids in the equilibrium in the exhaust. In a solid propellant particle bed which rise to the surface. As the voids rocket motor, the lack of velocity and them1 sporadically break through the surface, particles are equilibrium between the gases and particles results in spouted well above the surface, If the voids are large specific impulse losses of 2 - 5% (ref. 21). In enough to fill most of the column cross section, the addition, some of the metals of interest are slow behavior is termed slugging. Slugging causes burning, a characteristic which may cause incomplete periodic rising and collapsing of the particle bed combustion of the fuel and further degrade surface. Both of these behaviors would cause the performance (ref. 21). But, it is not possible to flowrate of metal to be inconsistent. The fluidized accurately estimate the total effect these and other bed was thus eliminated as a candidate for the fuel losses will have on metalloxygen propellant feed system. However, the cold flow tests did performance. Therefore, research has been initiated demonstrate like-on-like impingement of two into experimentally determining the performance powdedgas streams as a viable propellant injection losses and evaluating techniques to minimize these method. losses. Because the undesirable attributes of the Researchers at the University of Illinois are fluidized bed were largely due to free-surface studying the fundamentals of lunar metal particle behavior (bubbling and slugging), current powder combustion. Several parameters are being varied to feed system work is focussed on fabricating a determine the effect on ignition delay, combustion fluidized piston feed system. A fluidized piston has time, and exhaust particle size. The results of this no free surface, requires very little gas for project will provide a basis for choosing a metal fuel. entrainment, and operates independently from gravity. The use of a pneumatic feed system for the powder fuel provides another tool to positively affect Propellant Performance the lunar metal rocket performance. The carrier gas can dramatically affect performance even though only Because the fuels available at the moon and small quantities are used. The equilibrium Mars are nonconventional (i.e. containing no calculation results presented in Figure 5 show that hydrogen or nitrogen), little technology base exists. using hydrogen gas provides not only the added Under the task of in situ propellant performance, performance as expected but also lowers the several subtasks have been performed to determine combustion temperature. Lower combustion the performance of these nonconventional propellants temperatures would reduce cooling requirements--a in a rocket engine. For the lunar propellants, the significant issue which must still be addressed. subtasks include evaluation of theoretical specific impulse performance, metal particle ignition and Carbon Monoxide/Oxvgen. Carbon combustion, and experimental subscale engine monoxide and oxygen has been identified as a performance. For the Mars propellants, the subtasks propellant combination that can be obtained include evaluation of theoretical specific impulse completely from the martian atmosphere. Methane is performance, ignition characteristics, and engine another fuel which has been advocated for use at performance and combustion characteristics. Mars, although the hydrogen for the methane will either need to be delivered from the Earth or obtained boundary layer growth will further reduce the from the potential martian permafrost. The predicted specific impulse. Reference 23 contains a advantage of methane over carbon monoxide is a more complete discussion on the theoretical results. higher specific impulse, which means less propellant to produce. The advantage of carbon monoxide over The theoretical analysis indicated slow methane is the ability to obtain all of the fuel and kinetic reaction rates may affect the performance of oxygen from a single process (dissociation of carbon monoxide and oxygen. The slow kinetics may atmospheric carbon dioxide) and the elimination of also affect the ignition of the C010, mixture. The the need to mine hydrogen from the permafrost or equation for carbon monoxide oxidation is written as transport it from Earth, Phobos or Deimos, or nearby asteroids. To evaluate the performance potential of co + 112 0, -> co, (1) carbon monoxide and oxygen, a theoretical parametric evaluation was performed. This reaction, however, has a high activation energy. Onedimensional kinetics simulation indicates that at A onedimensional equilibrium computer the high temperatures and pressures typical in a code (ref. 20) was used to calculate vacuum specific rocket engine chamber, the energy bamer would be impulse as a function of mixture ratio, chamber overcome, and the reaction would be self-sustaining. pressure, and area ratio. Figure 6 shows the results Therefore, to determine if CO and 0, can be an of this parametric study for a mixture ratio range of effective propellant combination, an experimental 0.25 to 2.0, chamber pressures of 1.4 and 20.7 MPa program was conducted to investigate what ignition (200 and 3000 psia), and area ratios of 10, 60, 100, methods are required to initiate and nurture the 200, and 500. As expected, chamber pressure has a reaction until it produces sufficient energy to small effect on ideal specific impulse, with only a 5 overcome the high activation energy bamer and or 6 sec increase in specific impulse gained with an becomes self-sustaining. increase in chamber pressure from 1.4 to 20.7 MPa. One method of lowering the activation Figure 6 shows that theoretical specific energy of a reaction is by the introduction of a impulses as high as 313 sec are predicted for an catalyst to the system. Some transition metal and engine with a node expansion ratio of 500. This noble metal catalysts are known to promote this value, however, is an ideal theoretical prediction, and particular reaction, and are used in the automotive an actual engine would not be expected to deliver this catalytic converter (ref. 24). The presence of small performance. Another computer code (ref. 22) was amounts of hydrogen in the system will also act as a used to predict performance losses associated with catalyst. The key reactions in the mechanism are finite-rate kinetics, twodimensional flow, and listed below. boundary layer growth. A specific impulse efficiency was calculated by dividing the I, values predicted by 1/2 H, + 112 0,- > OH the various loss mechanisms by the ideal one CO + OH-> CO,+H (2) + dimensional equilibrium values. Figure 7 shows the H H-> H, specific impulse efficiencies obtained at the two different chamber pressures (1.4 and 20.7 MPa) The tests in this experimental program concentrated when finite-rate kinetics are included in the analysis. on the use of small amounts of hydrogen as the The figure shows that while the predicted kinetic catalyst for the reaction. Once ignition was initiated, losses at the stoichiometric mixture ratio are as much the hydrogen was no longer needed, and the reaction as 8 percent at low chamber pressure, the kinetic was allowed to proceed as a dry system. losses are only a little more than 3 percent for the high pressure. These efficiencies would reduce the The experiments were conducted in a spark- ideal predicted impulse of 313 seconds to 304 torch igniter, with small amounts of hydrogen added seconds at high pressure and 288 seconds at low to the carbon monoxide gas stream (.15 to 1.0 pressure. Although the kinetic losses were the most percent of the CO by weight). The following significant losses in the theoretical analysis, other conclusions were drawn from the experiments. losses caused by twodimensional effects and Gaseous oxygen and carbon monoxide will not 8 light in a spark-torch igniter. Ignition was achieved, conducted (ref. 25). The results indicate that, in however, with as little as 0.0062 weight percent general, liquid carbon monoxide acts as a better hydrogen in the carbon monoxide at a mixture ratio coolant than liquid oxygen for a CO/O, engine. The of 0.35 with ambient temperature oxygen. At higher differences in cooling performance are small enough, mixture ratios and lower oxygen temperatures, more however, that experimental testing will be needed to hydrogen was needed to initiate ignition (figure 8). confirm the proper selection of coolant. In addition, A definite mixture ratio range exists where the engine cycle analyses have been performed to identify carbon monoxide and oxygen will ignite and sustain most likely engine cycles and operating characteristics combustion even after both the hydrogen and the of a carbon monoxideloxygen engine system (ref. spark are shut off. These mixture ratio boundaries 26). are also dependent on the inlet temperature of the oxygen. CONCLUDING REMARKS With the success of the ignition tests, The resources available in the lunar soil and another experimental program was started to in the martian atmosphere contain large quantities of investigate actual engine performance. Two oxygen that can be used as the oxidizer in a rocket measures of engine performance were taken during engine. Although very little hydrogen exists at either the experimental tests. The first was characteristic site, other resources, such as lunar metals and velocity, C*, which was calculated based on the martian carbon monoxide, have great potential for measured chamber pressure and propellant flow rates. use as fuels. Mission analyses have shown that the The second measure of performance was the vacuum utilization of these in situ resources for near-planet specific impulse, which was calculated based on the and Earth-return propulsion can provide significant measured propellant flow rates and measured thrust benefits for space exploration. These advantages corrected to vacuum conditions by adding the nozzle include reduced launch mass, increased payload exit pressure force. Both of these performance capability, reduced trip time to Mars, and measurements were compared to theoretical values establishment of base self-sufficiency. predicted by the computer code. Figure 9 shows the experimental and theoretical vacuum specific impulse Because the in situ propellant combinations efficiencies as a function of mixture ratio. Again, the are not commonly used, a technical database must be theoretical ideal specific impulse was used as the established to support the development of rocket basis for the efficiency calculation. Because the engines that use these propellants. Recent activities expansion area ratio of the test hardware was only have already produced some results towards this goal. 2.36, and becaw the kinetic losses discussed Preliminary hazards assessment and formulation previously become more significant as expansion research have given strong indications that a liquid continues, the theoretical specific impulse efficiency oxygen/powdered metal monopropellant is a safe and is higher than that shown in figure 7. The viable candidate for the moon. Single particle theoretically predicted efficiencies are about 93 to 95 ignition research is focused on methods to reduce the percent, while the experimental efficiencies are 85 to performance losses anticipated with such a propellant. 89 percent. It was noted from the results for C* Similarly, technology work in the form of ignition efficiency (ref. 23) that the difference in theoretical and combustion performance evaluation has been and experimental efficiency was most likely caused conducted to build the necessary technology base to by incomplete energy release in the chamber that the develop an in situ propellant rocket engine for Mars. computer code does not predict. The difference This work has indicated that carbon monoxide and between theoretical and experimental specific impulse oxygen make a viable propellant combination for efficiency shown in figure 9, therefore, is likely Mars. caused by the same incomplete energy release. These preliminary investigations into the use In coordination with the experimental and of in situ propellants indicate that the rocket engine theoretical engine performance research, analytical technology should not be an impediment to achieving evaluations of the potential of carbon monoxide to be the potential benefits of in situ propellants. The used as a regenerative coolant are also being magnitude of these benefits, however, is dependent 9