TECHNICAL ABSTRACT ( Limit 2000 characters, approximately 200 words)
The Mars Gas Hopper, or “gashopper” is a novel concept for propulsion of a robust Mars flight and surface exploration vehicle that utilizes indigenous CO2 propellant to enable greatly enhanced mobility. The gashopper will first retrieve CO2 gas from the Martian environment to store it in liquid form at a pressure of about 10 bar. When enough CO2 is stored to make a substantial flight to another Mars site, a thermal storage bed is heated to ~1000 K and the CO2 propellant is warmed to ~300 K to pressurize the tank to ~65 bar. A valve is then opened, allowing the liquid CO2 to pass through the hot thermal storage bed that heats and gasifies the CO2 for propulsion. Gashopper can be designed to function as either ballistic flight vehicles or winged airplanes, with the former offering simplicity and the latter greater range. The advantage of the gashopper is that it provides Mars exploration with a fully controllable aerial reconnaissance vehicle that can repeatedly land and explore surface sites as well.
The key technical issue that determines the potential performance of a gashopper is the overall specific heat of the thermal storage bed. In previous work, Pioneer Astronautics has demonstrated working gashopper airplanes and ballistic flight vehicles that utilized magnesium oxide pellets for thermal storage. While convenient for test purposes, MgO has a specific heat that is only roughly equal to CO2. This severely limits the attainable mass ratio and thus vehicle range. In contrast, lithium has four times the specific heat of CO2, so its use as a gashopper thermal bed material would greatly improve vehicle performance. The low density and liquid nature of high temperature lithium makes its utilization for gashopper engines a challenge. In the proposed program, Pioneer will resolve this challenge by designing, building, and testing high specific heat gashopper engines using liquid lithium for thermal storage.

POTENTIAL NASA COMMERCIAL APPLICATIONS ( Limit 1500 characters, approximately 150 words)
The gashopper concept is primarily designed to enable greatly enhanced mobility for robotic Mars exploration vehicles. However the advanced gashopper engine system (AGE) has many potential important commercial applications in space. Small AGE thrusters could be used for stationkeeping and reaction control system (RCS) propulsion for satellites. Currently the propellant of choice for such applications is monopropellant hydrazine, which is extremely toxic, dangerous, expensive, and offers a rather low performance (Isp = 220 s). Small AGE based propulsion systems with ammonia propellant will be much cheaper, safer, and easier to integrate than hydrazine, while offering comparable or even superior specific impulse performance. If the AGE operates above 1200 K, there can also be an extension of satellite life. With their ability to store thermal energy over time and release it suddenly, advanced gashopper engines offer all the flexibility, reliability and safety of resistojets with much higher thrust. Thus, gashopper engine technology could find a major commercial market in the satellite industry. Advanced gashopper engines could also be used to great advantage to provide station-keeping propulsion for the International Space Station employing waste CO2 from the life support system as propellant and the AGE to provide thrust at a much higher level than would be possible using resistojets.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS ( Limit 1500 characters, approximately 150 words)
Another possible commercial application for advanced gashopper engine (AGE) technology is for rocket assisted takeoff (RATO) units for small aircraft. The AGE can easily be made to deliver large amounts of thrust, and because it can use water as propellant it poses no danger of fire to the surrounding area. Therefore an AGE RATO system offers interesting practical possibilities, which in fact were demonstrated by the short takeoffs achieved by the aircraft flown in the LaRC Phase 1 gashopper airplane program. One concept would be to build droppable integrated tank/engine units that would be heated on the airfield using locally available electric power. It sometimes happens that under emergency conditions an airplane might be forced to land at an airfield that is to short for it to takeoff from. Using a detachable RATO unit to augment thrust, the airplane could be made airborne again. Propellant could be obtained from local water sources, and autogenously pressurized by heating. Since the use of a RATO engine in such a case would avoid loss of an entire aircraft, the sale of such a RATO assist could be priced high enough to justify using such a system, even if it had to be expended after being dropped from a single flight. A suitable parachute system attached to the RATO unit might make expending such systems unnecessary.
AGE systems could also be used to store solar thermal power acquired during the day and then generate electricity at very high efficiency at night.

TOPICS 
Rocket Propulsion, Space Technology, Mechanical Engineering, Vehicles, Thermal Engineering, Systems Engineering, In-Situ Resource Utilization (ISRU), Spacecraft Systems, Mars Analog Field Exploration

Advanced Mars Water Acquisition System

NASA SBIR 2017 Solicitation

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Advanced Mars Water Acquisition System (AMWAS) recovers and purifies water from Mars soils for oxygen and fuel production, life support, food production, and radiation shielding in support of human exploration missions. The AMWAS removes water from Mars soils using hot, recirculating carbon dioxide gas to provide rapid heat transfer. The AMWAS evaporates water from ice and salt hydrates, leaving dissolved contaminants in the soil residue. The water distilled from the extraction vessel is condensed, treated with activated carbon to remove residual volatiles and organic material, filtered to remove suspended solids, and subjected to deionization in preparation for proton exchange membrane electrolysis. Recuperative heat exchange is employed to minimize heat losses from recirculating carbon dioxide gas. Cold temperatures of the Mars atmosphere are used to facilitate condensation and separation of water from recycled carbon dioxide gas. A vacuum jacket is used to minimize heat losses from the extraction vessel. Much of the net heat input to the AMWAS can be supplied by solar concentrators or waste heat from radioisotope thermoelectric generators. The AMWAS vessel is equipped with a single, stationary seal that facilitates materials handling automation and minimizes potential leakage over the nominal operating period of up to 480 days.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary application of AMWAS is for production of clean water from Mars soils for electrolysis, fuel and oxygen production, food production, and radiation shielding. The AMWAS can provide a reliable, low-cost, low-mass technology to produce water, hydrogen, and liquid oxygen on the surface of Mars out of indigenous materials at low power. The ability to extract water from Mars could also serve to supply the crew of Mars missions with water, which is the second most massive logistic component of a Mars mission. Smaller versions of the AMWAS could be used to help make the return propellant for a Mars sample return mission on the Martian surface, thereby making such a mission both cheaper to launch and much easier to land.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The AMWAS could be implemented in arid terrestrial climates for recovery of water from soils. Even in the driest regions of Earth, the regolith is several times wetter than on Mars, and the AMWAS can operate efficiently under those conditions. Regions that are too far from the coastline to economically pipe water may be potential markets. Units sized for vehicles traveling in desert regions could reduce logistical requirements for the military and civilians operating in remote areas, since it is very lightweight, cheap, and portable. By enabling agriculture in arid areas the AMWAS could also support the production of renewable energy in the form of biofuels.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 3
End: 4

TOPICS
Space Technology, Mechanical Engineering, Thermal Engineering, Systems Engineering, In-Situ Resource Utilization (ISRU), Spacecraft Systems, Life Support, Water Treatment

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
Pioneer Astronautics proposes an advanced photocatalytic oxidation reactor for enhancing the reliability and performance of Water Recovery Post Processing systems aboard crewed spacecraft. This novel technology, called the Advanced Wastewater Photooxidation System (AWPS) is designed to oxidize and remove recalcitrant aqueous organic constituents in the water recovery post processing system under ambient temperature and pressure conditions. The basis of the innovation is the combination of high brightness and long lifetime UV LED light sources with efficient geometric illumination of a highly active photocatalyst immobilized on a high surface area support. This combined approach leads to numerous performance benefits including high conversion efficiency, low temperature and pressure operation, compact footprint, high reliability and low crew maintenance, and decreased equivalent system mass (ESM). The Phase I effort will clearly demonstrate the feasibility of these concepts by mineralization of polar water soluble organics and organosilanol constituents under long duration testing. Data from the Phase I will lead to a prototype scale-up of the device in Phase II. Development strategies for the Phase II device include component design verification testing and determining optimum reaction conditions. Long duration performance tests will validate the reactor design, and establish the technology applications in space and commercial markets.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
This technology applies to any system where removal of toxic industrial chemicals and pathogens are present in water, and where robust, reliable operation with minimal supply logistics is required. The primary application of the AWPS is to provide a compact, high performance wastewater purification device for spacecraft environmental control and life support system (ECLSS) for extending NASA’s mission beyond low earth orbit to include long-duration space habitation, Lunar, and Mars colonization missions. This is accomplished by the advanced AWPS device which combines energy efficient UV LED-based source illumination, high photon utilization, highly active photocatalysts, high reactor surface-to volume, in situ reactant use, and low temperature and pressure operation.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The technology has numerous applications in municipal water treatment and commercial sectors including pharmaceuticals, food processing, cosmetics, electronics, oil and gas industries, and hospitals. Other government agencies such as the Department of Defense and First Responders would benefit from this technology as a portable water treatment technology against chemical and biological pathogens. The AWPS technology is well suited for the growing market demand of wastewater equipment and is poised to solve a variety of industrial water treatment challenges with minimal environmental impact. The AWPS technology is composed of entirely green chemistry- just light, iron oxide, and environmentally friendly on-site generation of hydrogen peroxide. This technology leads to the facile generation of hydroxyl radical chemistry, which is vital to the complete destruction of industrial chemicals in water. No addition of chlorine is needed and therefore deleterious chlorine byproducts are not generated.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 2
End: 3

TOPICS
Rocket Chemical Processes, Space Technology, Materials, Systems Engineering, In-Situ Resource Utilization (ISRU), Life Support, Waste Recycling, Spacecraft Systems, Water Treatment

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Alternate Propellant Thermal Rocket (APTR) is a novel concept for propulsion of space exploration or orbit transfer vehicles. APTR propulsion is provided by utilizing a nuclear thermal reactor or solar thermal engine to heat a space storable propellant, preferably consisting of a volatile indigenous to the destination world, to form a high thrust rocket exhaust. Candidate propellants whose performance, materials compatibility, and ease of acquisition make them worthy of examination for APTR propulsion of exploration vehicles include carbon dioxide, water, methane, and methanol. An APTR utilizing indigenous CO2 propellant potentially offers high payoff to a robotic or manned Mars mission, both by sharply reducing the initial mission mass required in low Earth orbit, and by providing Mars exploration with unlimited mobility and global access. Additionally, an APTR could give nearly unlimited mobility to asteroid or outer solar system probes, while one using methane or nitrogen propellant could enable a Titan sample return mission. The APTR can also be used as the propulsion system for a high performance space storable orbit transfer system moving payloads from LEO to GEO or other orbits of commercial interest. In this case, leading candidate propellants include methane, ammonia, and methanol.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
APTRs have important commercial applications for satellites. A reusable orbit transfer vehicle using a space storable APTR system with comparable performance to the 450 s available from the cryogenic H2/O2 Centaur would represent a major cost saving to commercial satellite delivery. Small APTR engines powered by electrical heaters could be used for stationkeeping and RCS propulsion for satellites. APTR propellants are much cheaper, safer, and easier to integrate than toxic hydrazine, and could offer twice hydrazine’s specific impulse. For a given RCS propellant allocation, this could double a satellite’s useful life, resulting in a major saving to the satellite industry.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
The commercial potential of APTR technology goes way beyond the APTR thruster market itself. The use of water or CO2 as propellant in a high temperature thermal rocket requires the development of high temperature oxidation-resistant coatings. The development of such ultra-high temperature protective coatings for APTRs would also create a technology that could play a vital role in the development of numerous types of high performance chemical rockets, and more than that, find abundant use in a wide range of high-temperature oxidizing-environment industrial applications on Earth as well. The markets for commercial applications of such material technology are vast.

TOPICS
Rocket Rocket Propulsion, Chemical Processes, Space Technology, Mechanical Engineering, Thermal Engineering, Systems Engineering, Spacecraft Systems

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Carbon dioxide Oxidizer Rocket (COR) is a novel propulsion system concept that can enable global mobility for Mars exploration. In this concept a fuel, such as B2H6, B5H9, SiH4, LiBH4, Al(BH4)3, H2, Mg, NH3 or N2H4 is transported to Mars and burned in the COR engine using native Martian CO2 as the oxidizer. Specific impulses for these propellant combinations range from 200 to 320 s. Because the majority of the propellant mass is CO2, which can be replenished from Martian air with a simple pump, the effective specific impulse of the fuel transported from Earth can exceed 1600 s. The COR could be used to power a Mars hopper vehicle, which replenishes itself with CO2 oxidizer each times it lands, allowing one Mars mission to explore a large number of widely dispersed sites. COR rockets could also be used to support NASA’s planned Mars Sample Return (MSR) mission by eliminating the need to transport to Mars most of the propellant needed by the Mars ascent vehicle. As a result the MSR mission could both reduce the size of its required launch vehicle and increase the sample size returned to Earth, greatly increasing mission cost-effectiveness.

POTENTIAL COMMERCIAL APPLICATIONS
The primary application of the COR system is to enable Mars global mobility systems and to greatly increase the cost effectiveness of NASA’s planned Mars Sample Return mission. However, other commercial applications potential of the COR are important and manifest. CO2 is a safe, non-cryogenic, non-toxic oxidizer. If rocket engines can be developed employing it, they will find broad application for use in space storable upper stages, sounding rockets, and easy to integrate satellite RCS systems. Current spacecraft RCS systems employ hydrazine, which is dangerous, toxic, explosive, expensive to integrate into a spacecraft, and low-performing (220 s Isp). COR rockets would be a cheap, safe, easy-to-integrate space storable alternative with superior performance. COR propulsion for sounding rockets for use in the university environment are particularly attractive, as alternative oxidizers in current use are either cryogenic (LOX) or toxic (NTO) or explosive (H2O2), all of which faults are of significant concern to educators who desire student involvement in sounding rocket launch or related propulsion projects.

TOPICS
Spacecraft Systems, Rocket Propulsion, Chemical Processes, Mars Analog Field Exploration, Space Technology, Mechanical Engineering, Systems Engineering, In-Situ Resource Utilization (ISRU)

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Carbon Monoxide Metal Oxide Reduction System (COMORS) is a method of producing useful oxygen by reducing oxides found in Lunar and Martian surface material. The COMORS heats the oxides to temperatures where the oxygen atoms combine with carbon monoxide gas acting as a reducing agent, producing carbon dioxide. The CO2 is reduced with hydrogen in a reverse water gas shift (RWGS) reactor, and the CO is recycled to the metal oxide reactor. The water produced in the RWGS reaction is electrolyzed; the oxygen is stored and the hydrogen recycled to the RWGS reactor. The COMORS is similar to some hydrogen/ilmenite reduction methods proposed for Lunar missions. Unlike the hydrogen based methods, however, the carbon monoxide/RWGS cycle can operate with higher energy efficiencies, lower losses, and lower temperatures. The concept is very applicable to reduction of Mars metal oxides since the greater content of iron oxides in Mars minerals will result in greater oxygen production for a given unit of oxide feed. This proposal demonstrates CO reduction methods for reducing metallic minerals on the Moon and Mars for oxygen production for human exploration.

POTENTIAL COMMERCIAL APPLICATIONS
The potential commercial applications of the COMORS are numerous. High-quality iron will be produced at lower cost using pure carbon monoxide reductant rather than coal-based reductant that contains impurities that must be removed using consumable additives, more oxygen, and longer heating times. The replacement of coal-based reductant will eliminate coke-oven emissions and production of hazardous byproducts. The economics of steelmaking will be further improved since the COMORS provides an on-site source of high-purity oxygen than can be used for decarburization, desiliconization, and reduction of heating times (resulting in increased productivity). The primary commercial advantage of a combined metal oxide reduction/RWGS system is that it can cut CO2 emissions dramatically since it consumes only an electrical energy input and produces a clean oxygen effluent. As nations strive to reduce CO2 emissions, the iron oxide reduction energy requirement can be shifted from dirty carbon-based sources to clean renewable sources to greatly reduce CO2 emissions currently associated with iron and steel production.

TOPICS
TIn-Situ Resource Utilization (ISRU), Life Support, Spacecraft Systems, Chemical Processes, Mineral/Ore Processing, Mechanical Engineering, Systems Engineering, Materials, Thermal Engineering, Space Technology

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Carbon Monoxide Silicate Reduction System (COSRS) is an innovative method that for the first time uses the strong reductant carbon monoxide to both reduce iron and to evenly deposit carbon. This enables high temperature carbothermal reduction of silicon oxide yielding five times as much oxygen recovery from planetary regolith compared to hydrogen-based reduction systems. COSRS is an in situ planetary resource utilization process that yields useful oxygen and metals by reducing the majority of metal oxides in undifferentiated lunar, asteroidal, and Martian surface materials. The COSRS initially heats the materials to temperatures where the iron-bound oxygen combines with carbon monoxide, a strong reducing agent (reductant). Simultaneously, the produced iron metal catalyzes the disproportionation of carbon monoxide to carbon and carbon dioxide. The temperature is then raised for carbothermal reduction of the silicates, producing carbon monoxide, which is recycled back to the first stage process, and silicon metal. The carbon dioxide created in the iron reduction/disproportionation step is processed with hydrogen in a Reverse Water Gas Shift (RWGS) unit to make carbon monoxide and water. After electrolysis, the oxygen is stored while the CO is recycled to the reactor.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
The COSRS is a potentially enabling technology for human Lunar exploration because it can produce the majority of the oxygen available in undifferentiated Lunar soil, or roughly five times the yield of hydrogen reduction technologies. This increased productivity eliminates the need to beneficiate the soil, thereby enabling automated lunar oxygen facilities that could produce return propellant prior to the arrival of the crew. This will greatly decrease the launch costs required to support the lunar base, and also enable long range exploration using ballistic hoppers employing Lunar oxygen. The COSRS will also work on asteroids, Mars, and Jupiter’s moons.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
The integrated COSRS/RWGS/Carbothermal reduction system has an application to the production of pure silicon metal for terrestrial manufacturing of photovoltaics and electronics components. The integrated COSRS/RWGS/Carbothermal reduction system also has future applications to the production of large quantities of oxygen, iron metal, and silicon metal from random lunar and Martian regolith for lunar and Martian bases, and could be used in the same way to allow useful metal production from very low grade ores on Earth. Furthermore, the closed COSRS/RWGS system would enable the terrestrial production of iron and other metals without generating carbon dioxide greenhouse gas.

TOPICS
TLife Support, Spacecraft Systems, Chemical Processes, Space Technology, Mechanical Engineering, Mineral/Ore Processing, Thermal Engineering, Materials, Systems Engineering, In-Situ Resource Utilization (ISRU)

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Carbonaceous Asteroid Volatile Recovery (CAVoR) system produces water and hydrogen-rich syngas for propellant production, life support consumables, and manufacturing from in-situ resources in support of advanced space exploration. The CAVoR thermally extracts ice and water bound to clay minerals, which is then combined with small amounts of oxygen to gasify organic matter contained in carbonaceous chondrite asteroids. In addition to water, CAVoR produces hydrogen, carbon monoxide, and carbon dioxide that comprise precursors to produce oxygen for propellant and breathing gas and to produce organic compounds including fuels such as methane when integrated with a downstream methanation-electrolysis. Thermochemical production of hydrogen by CAVoR results in substantial reductions in electrolysis mass and power requirements compared to combustion-based volatile recovery methods. A conceptual Phase II continuous flow auger reactor design was based on successful Phase I batch reactor operations. Phase II advancements will include reactor seal designs to accommodate regolith simulant feeding and discharging while collaborations will be developed to aid the infusion of the CAVoR system into a conceptual asteroid resource utilization mission plan.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary application of the Carbonaceous Asteroid Volatile Recovery (CAVoR) system is to provide a compact, high performance apparatus for the extraction and recovery of water and organic matter in support of propellant production, breathing gas, and life support. The in-space production of these mission critical items results in substantial launch cost savings and can help to enable the extension of NASA’s mission beyond low earth orbit to include long-duration space habitation, lunar, and Mars colonization missions.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The autothermal steam reforming technology proposed for the CAVoR has applications in the recovery of water and energy values from terrestrial wastes and resources. Steam reforming technology has mostly been applied to feed matter containing only small amounts of inorganic matter. The efficient use and recovery of process heat to be established during the CAVoR program will enable non-catalytic autothermal steam reforming technology to be applied to feeds such as contaminated soils, low-grade hydrocarbon feeds, oil shale, un-sorted municipal waste, and other organic materials, including renewable resources. By so doing, many otherwise refractory, hazardous compounds can potentially be broken into syngas constituents for use as fuels rather than being incinerated with no economic gain. The CAVoR technology will be poised for entry into the growing market demand for waste volume reduction and low-grade fuels resources. The device solves a variety of industrial and municipal waste challenges with minimal environmental impact.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 4
End: 5

NASA SBIR 2015 Solicitation – Phase I

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)

The Carbonaceous Asteroid Volatile Recovery (CAVoR) system extracts water and volatile organic compounds for propellant production, life support consumables, and manufacturing from in-situ resources in support of advanced space exploration. The CAVoR thermally extracts ice and water bound to clays minerals, which is then combined with small amounts of oxygen to gasify organic matter contained in carbonaceous chondrite asteroids. In addition to water, CAVoR produces hydrogen, carbon monoxide, and carbon dioxide that comprise precursors to produce oxygen for propellant and breathing gas and to produce organic compounds including fuels and plastics. Additional CAVoR byproducts include nitrogen, sulfur, and phosphorus compounds that have potential uses as buffer gas for life support and reagents for more-advanced asteroid materials processing. Process residues are thermally stabilized by the CAVoR system, which renders them suitable as bulk shielding, as feed to mineral separation and concentration, or as raw material for manufacture of structural components.

The CAVoR is a low-pressure, non-catalytic, batch process aimed toward maximum recovery of valuable constituents in a difficult operating environment using steel or other light-weight reactor alloys. Key elements of the CAVoR will be systematically developed and demonstrated through a progression from an Earth laboratory environment to experiments in zero-g flights and ISS with appropriate scale up and performance validations leading to implementation on a Near Earth Asteroid (NEA). Hardware designs to achieve required sealing and operating performance over long durations in microgravity will be derived in part from Pioneer’s heritage in lunar and Mars ISRU. A combination of modeling and experimentation will be used to overcome the lack of current well-established state-of-the-art process methods and conditions for resource recovery from Near Earth Asteroids.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary application of the Carbonaceous Asteroid Volatile Recovery (CAVoR) system is to provide a compact, high performance apparatus for the extraction and recovery of water and organic matter in support of propellant production, breathing gas, and life support. These capabilities are key to extending NASA’s mission beyond low earth orbit to include long-duration space habitation, lunar, and Mars colonization missions.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The autothermal steam reforming technology proposed for the CAVoR has applications in the recovery of water and energy values from terrestrial wastes and resources. Steam reforming technology has mostly been applied to feed matter containing only small amounts of inorganic matter. The efficient use and recovery of process heat to be established during the CAVoR program will enable non-catalytic autothermal steam reforming technology to be applied to feeds such as contaminated soils, low-grade hydrocarbon feeds, oil shale, un-sorted municipal waste, and other organic materials. By so doing, many otherwise refractory, hazardous compounds can potentially be broken into syngas constituents for use as fuels rather than being incinerated with no economic gain. The relatively low-temperature residue from autothermal steam reforming will be de-agglomerated, rendered sterile, and made suitable for down stream physical separations and byproduct recovery.
The CAVoR technology will be poised for entry into the growing market demand for waste volume reduction and low-grade fuels resources. The device solves a variety of industrial and municipal waste challenges with minimal environmental impact. The primary CAVoR steam reforming technology does not require exotic chemicals or catalysts for the production of water and syngas. Only small amounts of catalysts or sorbents are required for contaminant removal and conventional downstream fuels synthesis.

TOPICS
Life Support, Rocket Propulsion, Spacecraft Systems, Systems Engineering, In-Situ Resource Utilization (ISRU), Thermal Engineering, Chemical Processes, Space Technology, Mechanical Engineering, Mineral/Ore Processing, Vehicles

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The counterflow regolith heat exchanger (CoRHE) is a device that transfers heat from hot regolith to cold regolith. The CoRHE is essentially a tube-in-tube heat exchanger with internal and external augers attached to the inner, rotating tube to move the regolith. Hot regolith in the outer tube is moved in one direction by a right-handed auger and the cool regolith in the inner tube is moved in the opposite by a left-handed auger attached to the inside of the rotating tube. In this counterflow arrangement a large fraction of the heat from the expended regolith is transferred to the new regolith. The spent regolith leaves the heat exchanger close to the temperature of the cold new regolith and the new regolith is pre-heated close to the initial temperature of the spent regolith. Using the CoRHE can reduce the heating requirement of a lunar ISRU system by 80%, reducing the total power consumption by a factor of two.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The counterflow regolith heat exchanger (CoRHE) provides an efficient means to transfer heat from hot regolith to cold regolith. The ability to conserve the heat from the expended regolith can lead to significant energy savings for a lunar oxygen production system. If, for example, oxygen is produced at a rate of 1 metric ton (MT) per year with an oxygen content of 2% in the soil, then 50 metric tons of regolith must be processed per year. With oxygen production occurring 50% of the time (only during daylight) then the heating load is an average of 2.8 kW. In comparison, the electrolysis power required to produce 1 MT of oxygen per year at 50% duty cycle is about 1.1 kW. Thus, heating the regolith is one of the major power consumers of a lunar oxygen production system. The counterflow regolith heat exchanger is intended to reduce the heating requirement for the lunar oxygen production system by 80% with minimal hardware and power requirements. This reduces the total power requirement of the oxygen production system from 3.9 kW to 1.7 kW, a power savings of 55%.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
There are many chemical processes where powders or granular materials are processed at high temperatures. In each of these processes energy is spent heating and cooling the chemicals. The CoRHE can be used to simultaneously heat and cool the chemicals for a significant energy savings.

NASA’s technology taxonomy has been developed by the SBIR-STTR program to disseminate awareness of proposed and awarded R/R&D in the agency. It is a listing of over 100 technologies, sorted into broad categories, of interest to NASA.

Expected Technology Readiness Level (TRL) upon completion of contract: 3 to 4

TOPICS
Energy, Spacecraft Systems, Space Technology, Mechanical Engineering, Thermal Engineering, Materials, Systems Engineering, In-Situ Resource Utilization (ISRU)

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Durable Dust Repellent Coating (DDRC) consists of nano-phase silica, titania, or other oxide coatings to repel dust in a vacuum environment over a wide range of temperatures. The coatings are engineered with dielectric properties to strongly repel particles from surfaces. Durability is attained by application methods such as sol-gel coating or physical vapor deposition onto stationary and rotating surfaces of EVA equipment, hatches and seals, lunar modules, ISRU hardware, and habitats prior to assembly. The application of the coating is followed by annealing at elevated temperatures. Initial development is planned for stainless steel, followed later by other metals and plastics. In addition to dust repellency, the DDRC provides abrasion resistance to lunar hardware. Some of the DDRC coatings also impart UV resistance to the substrate. Unlike convential dust removal methods such as brushing or blowing that may result in deep infiltration of particles, dust can be readily removed from DDRC surfaces by tilting or mild vibration.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
DDRC for stainless steel developed during Phase 1 will be useful for lunar operations that involve exposed stainless steel surfaces on the moon. Phase 2 coatings would have much wider application to activities on the lunar surface, including space suit fabric, other metals, and flexible materials.

Pioneer would work with the Contracting Officer’s Technical Representative (COTR) to establish Technology Infusion Advisors to help guide the Phase 1 work. Commercialization could be achieved in cooperation with targeted industries or government agencies willing to invest in adaptation of the DDRC to specific industrial needs and for use in specified environments. This would include developing efficient, economic methods of application to the surface that DDRC is needed for and a course of field testing to prove effectiveness prior to marketing.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
DDRC coatings would also find wide applications for the Department of Defense during military deployment to dusty regions where similar equipment breakdowns occur as a result of dust contamination.

Industries that create processing dust (such as mining and excavation activity) would be likely beneficiaries for the DDRC.

Expected Technology Readiness Level (TRL) upon completion of contract: 4

TOPICS
Spacecraft Systems, Space Technology, Mechanical Engineering, Vehicles, Materials, Systems Engineering

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Extraterrestrial Metals Processing (EMP) system produces ferrosilicon, silicon monoxide, a glassy mixed oxide slag, and smaller amounts of alkali earth compounds, phosphorus, sulfur, and halogens from Mars, Moon, and asteroid regolith by carbothermal reduction. These materials, in some cases after further processing with other in-situ resources, are used for production of high-purity iron and magnesium metals (for structural applications), high purity silicon (for photovoltaics and semiconductors), high purity silica (for clear glass), refractory ceramics (for insulation, thermal processing consumables, and construction materials), and fertilizer (from phosphorus recovered from carbothermal reduction exhaust gases). Carbothermal reduction also co-produces oxygen at yields on the order of 20 percent of regolith feed mass when integrating downstream processes to recover and recycle carbon. Many of the EMP products can be prepared in a fashion suitable for casting or additive manufacture methods and have broad application in support of advanced human space exploration. The EMP methods are based on minimal reliance on Earth-based consumables; nearly all of the gases and reagents required for processing can be manufactured from Mars in-situ resources or can be recovered and recycled for applications using Moon or asteroid resources.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary application of EMP is for production of iron, silicon, and magnesium metals as well as refractory metal oxides and byproducts including phosphors and oxygen from Mars, Moon, or asteroid in-situ resources for manufacturing in support of advanced human space exploration. The EMP product suite includes many useful materials that will expand exploration and colonization capabilities while substantially reducing the costs and risks of bringing supplies from Earth. Many EMP product streams are suitable for use in advanced casting or additive manufacturing methods to allow for efficient use of resources.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
One potential terrestrial EMP application is the production of high-grade silicon metal or ferrosilicon. The hydrogen-enhanced carbon monoxide disproportionation method employed in the EMP system enables high rates of carbon deposition onto pure silica in the absence of a metal catalyst. Direct carbon deposition from CO generated during carbothermal reduction integrated with RWGS-electrolysis modules would reduce the purchase of carbon for the process while significantly reducing overall carbon emissions compared to current practice. In a closed-loop system including reverse water gas shift-electrolysis, silicon or ferrosilicon manufacturing could be accomplished with virtually no carbon emissions.

The EMP techniques have additional potential for the processing of lower-grade ores and feed stocks including residues and wastes. As higher-grade ores on Earth are more-difficult to find and mine, feed costs for existing technologies rise. The EMP can help to reduce overall processing costs by enabling the use of non-conventional feed stocks.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 3
End: 4

TOPICS
In-Situ Resource Utilization (ISRU), Mineral/Ore Processing, Chemical Processes, Space Technology, Systems Engineering

Mars Gashopper Airplane

Press Release [Jul 25, 2005 ]

Pioneer Astronautics has demonstrated a new technology for flying around Mars.

The new flight system is called a gashopper. The vehicle system works by acquiring CO2 from the Martian atmosphere with a pump (Mars atmosphere is 95% CO2) , storing it in liquid form, then sending it through a preheated pellet bed to turn it into hot rocket exhaust to produce thrust for a flight vehicle.

The flight vehicle could either be a ballistic vehicle similar to the DCX vertical takeoff rocket, or a winged airplane that would take off and land like a Harrier, then transition to horizontal flight.

On Mars, a ballistic gashopper would be capable of flights of tens of kilometers per hop. A winged aircraft would be capable of hundreds of kilometers per flight.

After each landing, a small rover could be deployed for local exploration. While it is doing this, the gashopper would refuel from the atmosphere, using power from the solar panels on its wings to drive its CO2 acquisition pump. This procedure would take about a month, then the rover would be recalled, the pellet bed reheated, and the gashopper flown to a distant landing site to explore again.

The net result is a system that can fly repeatedly on Mars, conducting numerous aerial surveys and surface exploration at many diverse sites with a single spacecraft. Furthermore, unlike surface rovers, the gashopper would not be blocked by terrain obstacles. Also, since its exhaust is CO2, it would not contaminate landing sites with organics from a conventional rocket exhaust (which might confuse sensors looking for indigenous organics).

In a series of tests conducted during the final weeks of July, 2005, Pioneer Astronautics demonstrated the gashopper concept in flight at the Platte Valley airport near Brighton, Colorado. The test vehicle, named Mars Ship One, was run through fast taxi tests, then flown at speeds between 60 and 100 mph.

Mars Ship One has a wingspan of 14 ft and a dry mass of 118 lb, making it a full scale representative in mass and size of a gashopper airplane that might be used on a robotic mars exploration . During the late July tests, flight ranges of about 1660 ft were obtained, with the pellet bed preheated to 800 C and 13 lbs of propellant in the tank. On Mars, with a hotter pellet bed, high rocket nozzle expansion ratios, 1/3 Earth gravity, lighter aerospace grade materials, larger propellant loads, and higher flight speeds, such a system could be expected to travel about 100 kilometers per flight.

The Gashopper airplane program was funded by NASA Langley Research Center with an SBIR Phase 1 contract to Pioneer Astronautics. Robert Zubrin was the Principal Investigator at Pioneer Astronautics, while Chris Kuhl was the program Technical Monitor at NASA Langley. Other members of the Pioneer Astronautics team included: Gary Snyder, Electronics Lead; Dan Harber, Aerodynamics Lead; Nick Jameson, Mechanisms design; Mike Hurley, Pilot; Kyle Johnson, Intern, and James Kilgore, Machinist.

We call her Mars Ship One, Dr. Zubrin said, because the desert skies of Mars are its oceans, and she is the first craft designed to navigate them. A 1600 ft flight is a humble beginning for Martian aviation, but then so was the 700 ft achieved at Kitty Hawk. All great things start out small. Someday vehicles descended from her with give us the freedom to travel at will across the Red Planet.

Mars Ship one was displayed for public viewing at the 8th International Mars Society Convention, university of Colorado, Boulder, August 11-14, 2005.

Photos of Mars Ship one during takeoff and flight, and the Pioneer gashopper team are shown below. To see a video clip of Mars Ship one in flight, click the link below.

Mars Ship One takes off.

Mars Ship One in flight.

The Pioneer Astronautics Mars Gashopper Airplane Team. From left: Gary Snyder, Dan Harber, Nick Jameson, James Kilgore, Robert Zubrin, and pilot Mike Hurley.

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Mars Gas Hopper Airplane, or “gashopper” is a novel concept for propulsion of a robust Mars flight and surface exploration vehicle that utilizes indigenous CO2 propellant to enable greatly enhanced mobility. The gashopper will first retrieve CO2 gas from the Martian environment to store it in liquid form at a pressure of about 10 bar. When enough CO2 is stored to make a substantial flight to another Mars site, a hot pellet bed is heated to ~1000 K and the CO2 propellant is warmed to ~300 K to pressurize the tank to ~65 bar. A valve is then opened, allowing the liquid CO2 to pass through the hot pellet bed that heats and gasifies the CO2 for propulsion. The hot gas is piped to a set of thrusters beneath the aircraft, allowing vertical takeoff, after which the gas is shunted off to a primary rearward pointing thruster to generate forward flight speed. The hot gas system is also used for attitude control and main propulsion during landing. The advantage of the gashopper is that it provides Mars exploration with a fully controllable aerial reconnaissance vehicle that can repeatedly land and explore surface sites as well.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
The gashopper enables greatly enhanced mobility for robotic Mars exploration vehicles. However the gashopper pellet bed rocket system (PBRS) has many potential important commercial applications in space. Small PBRS thrusters using ammonia could be used for stationkeeping and reaction control system (RCS) propulsion for satellites, as they would provide a non-toxic alternative with comparable performance to hydrazine. PBRS engines could also be used to great advantage to provide stationkeeping propulsion for the International Space Station employing waste CO2 from the life support system as propellant and the PBRS to provide thrust at a much higher level than possible using resistojets.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
Another possible commercial application for PBRS technology is for rocket assisted takeoff (RATO) units for small aircraft. The PBRS engine can deliver large amounts of thrust, and because its propellant is CO2 it poses no danger of fire. A PBRS RATO system could be designed as a droppable integrated tank/engine units that would be heated on the airfield using local electric power. Such systems could enable the takeoff of aircraft that are forced to land on short airstrips in emergencies. Since such use of a RATO would avoid loss of an entire aircraft, it could be priced high.

TOPICS
Systems Engineering, In-Situ Resource Utilization (ISRU), Mars Analog Field Exploration, Spacecraft Systems, Rocket Propulsion, Space Technology, Mechanical Engineering, Vehicles, Thermal Engineering

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The application of the HPPORS is to provide a compact, high performance air purification device for spacecraft environmental control and life support system (ECLSS) for extending NASA’s mission beyond low earth orbit to include long-duration space habitation, Lunar, and Mars colonization missions. This is accomplished by the innovative HPPORS device which combines energy efficient visible LED-based source illumination, high photon utilization, visible light activated photocatalysts, high reactor surface-to volume, and low pressure drop.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
This technology applies to any system where removal of toxic industrial chemicals and pathogens are present in air, and where robust operation with minimal supply logistics is required. Other government agencies such as the Department of Defense would benefit from this technology in similar applications such as collective protection of the warfighter from chemical-biological weapons attack. The technology would have numerous commercial and private customers in sectors including homes, schools, commercial offices, hospitals, and public transportation. The potential public benefits of an effective air purification system include fewer lost school and workdays due to sick-building syndrome and communicable illness.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 2
End: 3

TOPICS
Life Support, Waste Recycling, Spacecraft Systems, Chemical Processes, Space Technology, Materials, Systems Engineering, Water Treatment

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
Lightweight Auto-Inflating Self-Rigidizing Booms (LAISRB) is a key technology to enable large space ‘gossamer’ structures such as radiometer or radar antennas or solar sails. In Pioneer Astronautics LAISRB concept, inflatable tubes containing a small amount of methanol and a UV-curing self-hardening resin are deployed in a canister attached to the gossamer structure. Upon reaching deployment orbit, the canister is opened. The methanol contained in the booms will then vaporize deploying the booms. Once the balloons are deployed, they will be exposed to the ultraviolet radiation environment of space, which will cause the booms to harden. Once hard, the booms will remain rigid even after the methanol fluid used to inflate them leaks out to the vacuum of space. The advantage of the LAISRB is that the inflation system requires no high pressure gas storage devices, pyro valves, or other heavy, expensive, and failure prone equipment, and the hardening of the boom is automatic once exposed to the ultraviolet environment of space. The system thus promises to be lightweight, cheap, and very reliable.

POTENTIAL COMMERCIAL APPLICATIONS
The market for lightweight auto-inflating self-rigidizing booms (LAISRBs) is potentially large. In addition to providing the necessary deployment system for large orbiting communication systems, radars, passive microwave antennas, and numerous other Earth and space observation systems, such boom systems could also be used to enable solar sail spacecraft. LAISRBs could also be used to deploy large photovoltaic power arrays for both robotic and manned spacecraft. In addition, they could be used to deploy solar concentrators to enable high-power solar thermal electricity generation on orbit. Solar thermal propulsion systems capable of generating specific impulses over 800s could also be deployed using LAISRBs, thereby enabling cheap reusable orbit transfer vehicles. The economic return from all such applications taken together runs into billions of dollars, and by enabling them LAISRBs could do much to help advance not only the space program, but life on Earth as well.

TOPICS
Spacecraft Systems, High Altitude Balloon Experiments, Space Technology, Materials, Systems Engineering

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The LOX Olefin engine (LOXO or LOX/LC2H4 or LOX/LC3H6) is a proposed technology designed to provide interplanetary spacecraft with high specific impulse, space storable propulsion. With the LOXO engine, the combination of liquid oxygen with sub-cooled liquid ethylene or liquid propylene as a rocket propellant enables development of compact lightweight high performance stages with isothermal common bulkhead propellant tanks. This could enable energetic deep space delta-V maneuvers. Currently, space missions requiring energetic delta V maneuvers weeks or months after launch are limited to use of moderate Isp conventional stored hypergolic propellants, typically achieving performance of 310-320 seconds. The much more energetic LOX/LH2 combination has the decided disadvantage of low density, and is not considered suitable for long term storage in space because of thermal management issues with the extremely cold (~20K) LH2. The LOXO propulsion system represents an important midway step that combines the advantages and avoids the major flaws of both of these extremes. Theoretical Isp for this propellant combination over 390 seconds has been calculated using AFALS chemical equilibrium code. With an achievable in-space thermal management system, LOXO is a good candidate for deep space propulsion, with performance greatly surpassing current options for mission planners.

POTENTIAL COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
Although primarily conceived of as an enabling technology for governmentally sponsored deep space exploration missions, LOXO space propulsion would have many other applications as well. A high specific impulse space storable stage would find many customers among commercial, military, and scientific satellites who would value it as a very cost-effective alternative to the current choices of low performance toxic hypergols or hard cryogenic LOX/H2 propulsion. Commercial applications would include the delivery of large geostationary and medium altitude satellites into orbit, apogee kick, and also for high efficiency re-assignment maneuvers and end of life superboosting for geostationary satellites. The LOXO could be used for onboard spacecraft propulsion or for dedicated upper and transfer stages supporting every type of mission.

POTENTIAL NASA APPLICATIONS (LIMIT 150 WORDS)
LOXO could have applications to interplanetary NASA spacecraft that require large delta-V maneuvers well into deep space by providing high Isp storable fuels. Such fuels could allow faster transit times to the outer planets by avoiding gravity-assist flybys of inner planets or allow larger payloads.

TOPICS
Rocket Propulsion, Spacecraft Systems, Chemical Processes, Space Technology, Systems Engineering

Lunar Exploration Gas Spectrometer

NASA SBIR 2019 Solicitation

TECHNICAL ABSTRACT ( Limit 2000 characters, approximately 200 words)
The Lunar Exploration Gas Spectrometer (LEGS) is an instrument for studying the gas composition of lunar regolith. In the LEGS a 2.5 GHz solid state microwave transmitter positioned on a downward pointing horn is deployed by a lunar lander or rover using a long boom (e.g. 1-2 m) to set it down on the lunar surface, and then beams power into the regolith using its microwave transmitter. The microwaves directed down onto and into the ground contained under the horn, heating regolith to depths of several tens of centimeters. As a result, gases will be evolved from the cold subsurface regolith into the horn, where their composition will be analyzed by a near-infrared ~1 to 2.4-micron spectrometer mounted horn, and looking through a sapphire window into the interior of the horn illuminated by a tungsten lamp, enabling transmission spectra of evolved gases to be obtained. These instruments will provide qualitative and quantitative data on volatiles, potentially including water, hydrogen, helium, CO₂, CO, ammonia hydrocarbons, and other species as they evolve from the subsurface over time. Since gases released by upper layers of regolith will reach the horn first, this procedure will also provide composition as a function of depth. Once gas emission ceases, the horn is lifted by the rod and placed on a new location, where the process is repeated. The LEGS deployment will involve very little disturbance to lunar soils prior to analysis, thereby preventing the accidental release of lightly-bound volatiles that is thought to be significant even following gentle handling. In the proposed program, a full scale working model of the LEGS, including horn, microwave transmitter, and spectrometer, will be built and tested in Pioneer Astronautics

POTENTIAL NASA COMMERCIAL APPLICATIONS ( Limit 1500 characters, approximately 150 words)
The LEGS program will provide NASA with a key technology finding volatiles on the Moon, which represent a tremendous resource for human exploration. The data produced by the LEGS would be invaluable for lunar science itself, providing essential information for understanding the origin and history of the Moon and similar bodies no doubt present in orbit around numerous planets in other solar systems. LEGS could also be used on Mars, Phobos, Deimos, asteroids, moons of the outer planets, Mercury, Pluto and even comets

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS ( Limit 1500 characters, approximately 150 words)
The LEGS be used on Earth without major modification employing its IR spectrometer to determine amounts of volatiles, including trade contaminants, in the soil. It thus represents an instrument with broad potential utility for geology, resource exploration, and environmental remediation.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 1
End: 4

TOPICS
Space Technology, Spacecraft Systems, In-Situ Resource Utilization (ISRU)

Lunar Flow Battery

NASA SBIR 2019 Solicitation

TECHNICAL ABSTRACT ( Limit 2000 characters, approximately 200 words)
The Lunar Flow Battery (LFB) is a scalable, long-duration energy storage solution featuring minimum capacity fade over many cycles that uses electrolytes derived from lunar regolith to minimize launch mass. The LFB operates by storing two separate solutions of redox-active species which are pumped past the cathode and anode respectively to produce a current. By pumping the redox-active fluids across the electrodes, the energy and power can be scaled independently. What makes the LFB distinct from other flow batteries is its use of locally available resources to produce the electrolyte solutions, thereby reducing the launch mass. Lunar-sourced iron, titanium, sulfur, oxygen, and water provide the bulk electrolyte solutions while the more sophisticated components such as membranes, pumps, and electrodes are transported from Earth. Compared to alternatives such as Li-ion batteries, the LFB has vastly superior cycle life and the energy storage is readily scalable, making it an ideal solution for long-term, stationary storage over the lunar day/night cycle. By using locally available materials to produce the redox-active species and with no need for replacement cells for dozens of years, the energy storage capacity is high relative to the total launched mass. The Phase I program will investigate the selective dissolution of ilmenite (FeTiO₃), an ore available in high concentrations in lunar mare basalts, using sulfuric acid to produce iron and titanium sulfate electrolyte solutions and incorporate these solutions into a functional redox flow cell. This cell will be cycle tested to quantify its performance with regards to capacity fade and specific energy while any operational issues or degradation pathways will be addressed.

POTENTIAL NASA COMMERCIAL APPLICATIONS ( Limit 1500 characters, approximately 150 words)
The principal future application of the LFB is to provide long-duration energy storage for a permanent lunar base. The LFB is ideally suited for such a remote outpost with long day/night cycles where locally available resources can provide the basic materials to produce a large-scale energy storage system with a lower launch mass than alternatives. This system could be scaled up or multiplied to provide power to any number of long-duration scientific platforms, human habitats, and ISRU processing systems.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS ( Limit 1500 characters, approximately 150 words)
The LFB technology could provide an alternative solution for large-scale remote storage where access to resources is limited. Alternatively, the development of sulfuric acid processing of mixed metal oxides could provide an improved method for the production and recycling of titanium dioxide as a pigment for the coatings industry, opening up ilmenite deposits for cheaper TiO₂ production.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 2
End: 4

TOPICS
Space Technology, Spacecraft Systems, Energy

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Lunar Materials Handling System (LMHS) is a method for transfer of bulk materials and products into and out of process equipment in support of lunar and Mars in situ resource utilization (ISRU). The LMHS conveys solids to the ISRU vessel, provides a gas-tight pressure/vacuum seal, and minimizes wear related to abrasive particles. Lunar and Mars ISRU scenarios require that equipment be operated over many cycles with minimal consumption of expendables and with minimal leakage in order to maintain high overall process leverage. ISRU processes can be demonstrated in the laboratory to establish basic feasibility with respect to reagent leverage. Reagent leverage is defined as the mass of commodity produced divided by the mass of reagents consumed. However, the process leverage component related to equipment wear and loss of gasses, reagents, or product through seals and valves is more difficult to establish from laboratory testing. The LMHS increases equipment life and minimizes process losses, thereby increasing overall leverage and reducing uncertainties in ISRU process evaluation. The LMHS is based on a seal arrangement by which lunar regolith can be introduced into and removed from reaction chambers operating under a wide range of batch operating conditions.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
The primary application of the Lunar Materials Handling System is to advance in situ resource utilization on the Moon and then Mars. A number of regolith processes to extract oxygen, metals, and metal oxides require a reliable, robust, reusable materials feed and discharge system to achieve high overall process leverage. A tight sealing system reduces process reagent losses. A highly reliable, reusable materials handling system reduces consumables requirements and improves process reliability.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
Potential terrestrial applications of LMHS include improved materials handling and seal systems for hazardous materials or waste processing and nuclear materials processing. Improved containment of process materials would reduce costs and reduce operating risks.

TOPICS
In-Situ Resource Utilization (ISRU), Materials, Systems Engineering, Mineral/Ore Processing, Mechanical Engineering, Space Technology

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Lunar Organic Waste Reformer (LOWR) utilizes high temperature steam reformation to convert all plastic, paper, and human waste materials into useful gases. In the LOWR, solar thermal concentrators are used to heat steam directly to 900 C, after which the steam is injected into a reactor which is being fed with waste materials via a lock hopper. At the high temperatures, the steam will react with all organic materials to produce a gas mixture largely composed of hydrogen and carbon dioxide, with small fractions of methane and CO. After removing the remaining steam from the product stream via condensation, the gases are dusulfurized and then fed through a polysulfone membrane which separates CO and CH4 in the retentate from CO2 and H2 in the permeate. The retentate CO/CH4 gas stream can be used to reduce regolith, while the CO2/H2 permeate stream is sent to a Reverse Water Gas Shift (RWGS) reactor which transforms the CO2/H2 gas into CO and H2O. The CO can then be used for regolith reduction, while the H2O can be used as is, or electrolyzed to make oxygen and hydrogen. With effective recycling of the steam, no consumables are lost in the process. All products are liquids or gases, making the system highly reliable and subject to automation. In the proposed Phase 1 program, Pioneer Astronautics will build on its extensive heritage with development of RWGS and regolith reduction systems developed for Lunar and Mars in-situ propellant production to build and demonstrate a LOWR unit.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The LOWR would provide NASA with a technology capable of completely recycling the metabolic and plastic wastes of the crew of a lunar base to produce pure breathing oxygen, water, as well as useful reductants or fuels including CO, hydrogen, methane, and/or methanol, thereby significantly reducing lunar base logistic support costs. Mass savings for a 4 person base could be as much as 6 tons per year in lunar payload delivery, which translates into a reduction of 30 tons per year launched to orbit. Using electrical heat in place of solar thermal concentrators to superheat steam, the LOWR could also be used to recycle wastes on the International Space Station, the Orion spacecraft, or at a Mars base. In addition, LOWR technology can also be used to turn Martian atmospheric CO2 into useful methane and oxygen bipropellant. The ability to make such propellant on Mars is potentially a huge cost saver for both robotic Mars sample return (MSR) missions and well as human Mars exploration. Indeed, currently a major show stopper for the Mars sample return mission is the inadequacy of existing aerobrake technology to deliver a payload as massive as a fully-fueled Mars ascent vehicle to the Martian surface. By sharply reducing the mass that needs to be delivered to the surface, LOWR technology could not only reduce the cost of the MSR mission, but potentially enable it.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
On Earth, the LOWR could be used as a means of recycling plastics and other wastes to produce such useful clean burning fuels as methane, which is a prime product for generating electricity, and hydrogen and methanol, both of which are of great interest for use in fuel cells. Manufacture of such fuels from wastes could help achieve a reduction in total emission of greenhouse gases, since if disposed of otherwise or left to decay on their own, the carbon in the waste products would eventually turn into CO2 without displacing other fuel use. Currently, there is much public discussion over the possibility of converting cars to run on natural gas or methanol. If such programs move forward, LOWR technology could also be used to produce fuel for the automotive transportation market as well, thereby contributing significantly to liberating the nation from its dependence on foreign oil.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 3
End: 5

TOPICS
Life Support, Waste Recycling, Spacecraft Systems, Chemical Processes, Space Technology, Mechanical Engineering, Thermal Engineering, Systems Engineering

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Lunar Soil Particle Separator (LSPS) is an innovative method to beneficiate soil prior to in-situ resource utilization (ISRU). The LSPS improves ISRU oxygen yield by boosting the concentration of ilmenite or other iron-oxide bearing materials found in lunar soils. LSPS particle size separations can be performed to improve gas-solid interactions and reactor flow dynamics. LSPS mineral separations can be used to alter the sintering characteristics of lunar soil. The LSPS can eventually be used to separate and concentrate lunar minerals useful for manufacture of structural materials, glass, and chemicals.
The LSPS integrates an initial centrifugal particle size separation with magnetic, gravity, and/or electrostatic separations. The LSPS centrifugal separation method overcomes the reduced efficiency of conventional particle sieving in reduced gravity. Feed conditioning, such as charge neutralization, can be incorporated into the LSPS to release and disperse surface fines prior to particle separations. The conceptual LSPS hardware design integrates many individual unit operations to reduce system mass and power requirements. The LSPS is applicable to ISRU feed processing as well as robotic prospecting to characterize soils over a wide region on the Moon. The LSPS is scalable and is amenable to testing and development under simulated lunar environmental conditions.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary initial application of the LSPS is for lunar particle separations in support of improving the feed to hydrogen reduction ISRU. The LSPS has direct use to improve the overall efficiency of hydrogen reduction ISRU by boosting the iron-oxide content of feeds. In addition, the LSPS has uses for optimizing particle size distribution to improve material flow properties and gas-particle interactions in fluidized bed and other reactors as well as adjusting mineral composition to minimize sintering during reduction. The LSPS can also serve as a component of a robotic lunar prospector to characterize soils and their potential for ISRU applications.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
One non-NASA commercialization application is directed toward small-scale terrestrial mineral processing. In particular, the LSPS is useful in remote locations where a compact, low-power device is needed to perform dry separations for production of mineral concentrates. A device such as the LSPS can be tailored to dry separation prospecting or small-scale minerals production to reduce the transportation of large amounts of un-beneficiated samples or ore to laboratories or downstream processing facilities. Applications may include prospecting or small-scale production of gold ores and heavy mineral sands.

NASA’s technology taxonomy has been developed by the SBIR-STTR program to disseminate awareness of proposed and awarded R/R&D in the agency. It is a listing of over 100 technologies, sorted into broad categories, of interest to NASA.

Expected Technology Readiness Level (TRL) upon completion of contract: 4 to 5

TOPICS
In-Situ Resource Utilization (ISRU), Mineral/Ore Processing, Materials, Space Technology, Systems Engineering, Mechanical Engineering

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Lunar Sulfur Capture System (LSCS) is an innovative method to recover sulfur compounds from lunar soil using sorbents derived primarily from in-situ resources. Most of the sulfur released from lunar soil during higher-temperature thermal treatment is trapped by the LSCS at lower temperatures on iron oxides present in lunar soil. As needed, small amounts of polishing sorbents are used to reduce equilibrium sulfur concentrations to the low ppm level. After sorbents become saturated, sulfur compounds are desorbed and converted to useful sulfur products. Sulfur is present in concentrations of about 0.1 percent in lunar soils and can be recovered by the LSCS as a useful product from in-situ resource utilization (ISRU). The LSCS can capture and recover sulfur from lunar soil as a primary product during thermal desorption of volatile compounds or during thermal reduction ISRU processes used for oxygen production. Removal of sulfur compounds is required during ISRU to prevent electrolyzer damage, catalyst poisoning, and equipment corrosion. The LSCS is applicable to thermal ISRU reduction processes in which sulfur is released in forms such as hydrogen sulfide (H2S), carbonyl sulfide (COS), or carbon disulfide (CS2).

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary initial application of the LSCS is for lunar sulfur capture and recovery. The LSCS has direct use to both protect ISRU hardware and catalysts while producing useful amounts of sulfur for other lunar ISRU applications. Implementation of the LSCS will proceed through the SBIR Phase 1, 2, and 3 programs, with increasing levels of development achieved through each step. During each phase, Pioneer will identify the LSCS requirements and will establish the commercial relationships needed to provide materials, fabrication, and implementation strategies for NASA lunar application.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
One non-NASA commercial application of LSCS is the large-market target of integrated gasification combined cycle power generation from coal. Other smaller-market opportunities are at least as likely to benefit from the LSCS applied to reduce emissions and waste disposal requirements from a variety of industrial applications. These opportunities are likely to grow as de-centralized fuel preparation technologies are advanced for conversion of biomass and other potentially contaminated feeds to alcohol and other fuels. Pioneer will be actively monitoring these activities and the potential applications and the use of the LSCS technologies. As appropriate, Pioneer will make contact with those involved in these markets to establish the business feasibility of the LSCS to terrestrial applications.

NASA’s technology taxonomy has been developed by the SBIR-STTR program to disseminate awareness of proposed and awarded R/R&D in the agency. It is a listing of over 100 technologies, sorted into broad categories, of interest to NASA.

Expected Technology Readiness Level (TRL) upon completion of contract: 4 to 6

NASA SBIR 2007 Solicitation | Phase I

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Lunar Sulfur Capture System (LSCS) is an innovative method to capture greater than 90 percent of sulfur gases evolved during thermal treatment of lunar soils. LSCS sorbents are based on lunar soil iron compounds that trap sulfur contained in hot in-situ resource utilization (ISRU) product gases. Small amounts of polishing sorbents are used as needed to reduce equilibrium sulfur concentrations to the ppm or sub-ppm level. The LSCS is an effective technology for protecting in-situ resource utilization (ISRU) hardware from damage caused by the corrosive effects of hydrogen sulfide (H2S) and other sulfur-containing gases. Saturated sorbents can be regenerated for reuse, and desorbed sulfur can be converted to elemental sulfur. Key process steps include bulk H2S capture on lunar soil, further capture of H2S on polishing sorbent, regeneration of soil sorbent for re-use, recovery of high-purity H2S, and conversion of H2S to elemental sulfur. The LSCS reduces the risk of using Earth-based sorbents for primary sulfur capture by ensuring a ready supply of sorbent in the event of poor regeneration performance or process upset. The LSCS primary sulfur sorbent can be used as a non-regenerable sorbent if necessary without significant consequence to the ISRU process.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary initial application of the LSCS is for lunar sulfur capture and recovery. The LSCS has direct use to both protect ISRU hardware and catalysts and to produce useful amounts of sulfur for other lunar ISRU applications. The LSCS is directed at reducing the mass penalty and risk associated with the use of Earth-based sorbents for hydrogen reduction and other lunar oxygen production processes.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The LSCS has direct commercialization potential for thermal-catalytic synthesis processes that require sulfur removal from reducing gases derived from coal, biomass, or other impure hydrocarbon feeds in support of fuels and chemicals manufacture. The closed-loop aspect of the LSCS sulfur recovery system is particularly attractive as a potential zero-emission sulfur recovery system for terrestrial applications.

Expected Technology Readiness Level (TRL) upon completion of contract: 3 to 4

TOPICS
Thermal Engineering, Materials, Systems Engineering, In-Situ Resource Utilization (ISRU), Mineral/Ore Processing, Chemical Processes, Space Technology, Spacecraft Systems

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)

The MACDOF project involves design and construction of a demonstration unit that will freeze carbon dioxide from the Martian atmosphere. The MACDOF is much less massive than a sorption pump sized for the same production rate and can significantly reduce the mass of the unit required to obtain carbon dioxide from the Martian atmosphere for ISRU processing. Joseph Trezathan was the JSC CTOR, and Robert Zubrin was the principal investigator at Pioneer Astronautics.

During the MACDOF project, Pioneer conducted a series of systems studies which showed a decisive advantage in CO2 acquisition system mass if a refrigerator was used in place of a passive sorption pump. We then built three prototype carbon dioxide freezers and successfully operated one of them in several different configurations. The freezers were sized to produce 700 grams of carbon dioxide per day. During a Mars mission using in situ resources, a spacecraft can use carbon dioxide frozen in the equipment demonstrated during this project to supply a fuel or oxygen production process.

While by no means demonstrating the full potential of refrigeration systems over sorption pumps, the data acquired in the Pioneer Astronautics refrigeration experiments nevertheless made a conclusive case. In operation, the refrigerator employed was able to demonstrate twice the daily CO2 acquisition capacity of a passive sorption pump massing 1.4 times as much. The freezer’s mass advantage is understated, however, because it was built out of steel while the sorption pump was built out of aluminum. Had the same materials been employed in construction, the freezer mass would have been about a fourth of the sorption pump, while still outproducing it daily by a factor of 2.

The MACDOF Project was supported by SBIR Funding from NASA Johnson Space Center.

TOPICS
Spacecraft Systems, Carbon Capture, Systems Engineering, In-Situ Resource Utilization (ISRU), Thermal Engineering, Space Technology

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)

Propellant production on Mars (in situ propellant production, ISPP) benefits from the easy availability of carbon in the form of atmospheric carbon dioxide. However, the arid nature of the Martian surface means that there is a low availability of hydrogen. Thus, most hydrocarbon or oxygenated hydrocarbon based ISPP plants require liquid hydrogen feedstock imported from Earth. For example, the current Sabatier/Electrolysis (S/E) ISPP scheme combines revaporized liquid hydrogen and Martian carbon dioxide in a catalytic reactor to create methane fuel and water with a fuel H/C ratio of four. The water is condensed and electrolyzed to produce oxidizer (half as much oxidizer as ideal methane stoichiometry requires) and hydrogen for recycle to the reactor. The imported liquid hydrogen is only a small portion of the mass of the total fuel/oxidizer mix (8% for the S/E system), but it has other problems. Since it is an extreme cryogen, it requires extensive insulation, which adds to launch mass. Despite the insulation, a significant fraction of the hydrogen launch mass could be expected to boil off before processing begins on the Martian surface. Finally, the low density of liquid hydrogen (approximately 70 kg/m³) means that the storage tanks take up a large volume. If sufficient hydrogen for the ISPP requirements can be obtained on Mars in an energy efficient manner, this will inevitably be preferred to liquid hydrogen importation. However, the methods for collecting hydrogen on Mars are energy intensive relative to the ISPP processes which use the hydrogen to make hydrocarbon fuels and/or will require substantial industrial infrastructure. Thus, it is clear that, all else being equal, an ISPP technology that uses less hydrogen will be greatly preferable to one which uses more hydrogen.

The Mars Aromatic Hydrocarbon and Olefin Synthesis System (MAHOSS) is a method for producing storable low hydrogen/carbon ratio fuel and oxygen on the surface of Mars with 98% of the required raw material mass derived from the Martian atmosphere. In the MAHOSS system, a reverse water gas shift (RWGS) reactor is integrated with a catalytic fuel production reactor. The reactors combine imported hydrogen with Martian atmospheric CO₂ to produce aromatic or olefin fuels and water, with the latter product subsequently electrolyzed to produce oxygen and return hydrogen feedstock to the system. In this system, approximately 40 kg of fuel/oxygen bipropellant are produced for every kilogram of hydrogen imported to Mars, an attractively high leverage ratio. Another significant advantage of the MAHOSS system is its low power consumption, about half the power of the S/E Mars ISPP system.

Product properties

The MAHOSS project demonstrated production of four primary hydrocarbon products: ethylene, propylene, benzene, and toluene. These chemicals are four of the most common commodity hydrocarbons produced today, with a total annual market of more than 50 million tonnes. On Earth, the principal feedstocks for making ethylene and propylene are natural gas or catalytic cracker offgas. The principal feedstocks for benzene and toluene are heavy crude oils.

Table 1: Low H/C ratio fuel properties

Note that all of these are endothermic molecules, which is expected given the large amount of carbon relative to hydrogen. They are also all stored relatively easily at Mars ambient temperatures and about one bar of pressure. Ethylene will require refrigeration and/or additional pressurization. Benzene will require some heating at night to maintain it as a liquid.

As fuels, olefins and aromatics burn considerably hotter than fuels such as methane or hydrogen because of the extra carbon. However, their performance is reasonably high. Ethylene ideally gets about 380 seconds specific impulse burning with oxygen, while benzene and toluene are some of the primary ingredients in gasoline and kerosene type fuels, and should be able to achieve on order of 360 seconds specific impulse.

Although all the compounds listed in Table 1 are usable as high quality fuels, their potential as feedstock to plastic production processes should be briefly mentioned. Ethylene has been called the key to the plastics industry. In terrestrial applications, it is the precursor for many different varieties and grades of polyethylene, which is the most widely utilized plastic in the world. Propylene is the precursor for polypropylene, which is the second largest volume commodity terrestrial plastic at this time and is widely used for textiles. Toluene is not directly used as a plastic monomer, but is usually converted to benzene and/or directly combined with ethylene or propylene. With various further processing steps, these chemicals end up in products as diverse as polystyrene, styrene-butadiene rubber, Nylon, polyurethane, epoxies, aspirin, and explosives. While these applications will be irrelevant to robotic and early human missions, they are potentially very valuable when it is time to build a permanent outpost on Mars.

Production of Aromatics and Olefins on Mars

Aromatics and olefins can be produced on Mars by running two reactions in series. The Reverse Water Gas Shift (RWGS) is conducted in the first reactor. The primary products of the RWGS system are water and carbon monoxide. The water is condensed and electrolyzed to produce the oxygen portion of the propellant mixture. The carbon monoxide from the RWGS reactor is mixed with hydrogen to form a mixture called synthesis gas (syngas). In the chemical industry, syngas is a widely marketed and very valuable commodity (terrestrially, it is formed via steam reforming of coal or methane rather than with a RWGS). The syngas is piped off to a second reactor, where it reacts catalytically to produce hydrocarbon products.

The Reverse Water Gas Shift

The reverse water gas shift (RWGS) reaction has been known to chemistry since the mid 1800’s. While it has been discussed as a potential technique for Mars propellant manufacture in the literature, until the initiation of the 1997 Phase I MMISPP program by Pioneer Astronautics, there had been no experimental work done to demonstrate its viability for such application. The RWGS reaction is given by equation (1).

CO₂ + H₂ → CO + ΔH₂O H = +9 kcal/mole (1)

This reaction is mildly endothermic and occurs rapidly in the presence of an iron-chrome or copper-on-alumina catalyst at temperatures of 300 C or greater. Unfortunately, at 400 C the equilibrium constant Kp driving it to the right is only about 0.1, and even at much higher temperatures Kp remains of order unity. There is thus a significant problem in driving the RWGS reaction to completion.

However, assuming that reaction (1) can be completely driven to water and CO products, an “infinite leverage oxygen machine” can be created by simply tying reaction (1) in tandem with the water electrolysis reaction. That is, the CO produced by reaction (1) is removed from the RWGS while the water is electrolyzed to produce oxygen (the net product), and hydrogen which can be recycled to reduce more CO₂. Since all the hydrogen, barring leakage, is recycled, this process can go on forever, allowing the system to produce as much oxygen as desired. The only imported reagent needed is a small amount of water, which is endlessly recycled.

The RWGS/electrolysis oxygen machine has many advantages over the alternative zirconia/electrolysis system which is sometimes identified for the same purpose. In contrast to the zirconia system, which is composed of thousands of small brittle ceramic tubes manifolded together, the RWGS reactor itself is just a simple steel pipe filled with catalyst, much like a Sabatier reactor, except that the catalyst is different. A similar condenser and identical water electrolysis system to that used in the well demonstrated Sabatier/Electrolysis is also employed. Because the RWGS reaction is only mildly endothermic (9 kcal/mole for RWGS compared to 57 kcal/mole for water electrolysis), system power requirements are dominated by the water electrolysis step, for which the available technology is highly efficient. Moreover, since the thermal power required by the RWGS is less than that produced by a fuel production reactor and their operating temperatures are comparable, such catalytic fuel-making reactors can be used to provide some or all of the heat required to drive the RWGS reactor.

The trick, however, is to find a practical way to drive the RWGS reaction to completion. Pioneer Astronautics demonstrated a way to accomplish this in 1997 during the MMISPP Phase I SBIR. Briefly, in the Pioneer Astronautics Phase I MMISPP device, effluent gas from the RWGS is run through a 5 C water condenser to remove the water vapor, and then the remaining gas is sent to a membrane separator device. The membrane separator returns the hydrogen and CO₂ components of the effluent gas to the reactor while the CO is either vented as waste or sent on to be used by the fuel reactor. In operation of this device, simultaneous conversions of over 98% of hydrogen and CO₂ were accomplished, or 100% of either when run as the lean reactant. While some hydrogen is lost from the system during recycle, propellant leverages (the ratio of CO/O2 produced to the hydrogen feedstock) as high as 288 were produced. As a result, the practicality of driving the RWGS to completion can now be considered demonstrated.

The discussion so far has shown how a RWGS reactor can be used as the sole component in a loop with an electrolyzer as an “infinite-leverage oxygen machine” on Mars. In addition, a RWGS reactor operating without an electrolyzer can be used to turn imported hydrogen into water on Mars (for crew consumables) with a mass leverage ratio of 9/1. However the RWGS reactor opens up additional remarkable possibilities.

For example, we can operate the RWGS reactor with an excess of hydrogen, but without recycling all the waste hydrogen effluent. As a simplified case, assume that the H2/CO₂ molar input ratio is 3/1, and that the CO₂ conversion rate is close to 100%. Thus, we have 3 units of H2 and 1 unit of CO₂ going into the reactor, 1 unit of H2O collected in the condenser, and 1 unit of CO and 2 units of H2 leaving the reactor. The water is electrolyzed to produce product 0.5 units of O2 for the propellant tanks and one H2 unit for recycle into the RWGS. More importantly, the CO and H2 effluent from the process forms a high quality syngas mixture that can be fed as input into a hydrocarbon fuel production reactor.

Use of RWGS Reactor Outlet to Produce Fuels

Thus, there is no doubt that a RWGS system can produce a feed consisting of carbon monoxide and hydrogen for the MAHOSS fuel production unit. The next step in the MAHOSS is to produce higher hydrocarbons from the syngas, which is done via the Fischer-Tropsch (F-T) synthesis reactions, shown as reactions (2) – (6):

(2n + 1) H₂ + n CO → C_nH_{(2n + 2)} + n H₂O (2)

2n H₂ + n CO → C_nH_2n + n H₂O (3)

(n + 1) H₂ + 2n CO → C_nH_{(2n + 2)} + n CO₂ (4)

n H₂ + 2n CO → C_nH_2n + n CO₂ (5)

2n H₂ + n CO → C_nH_{(2n + 1)}OH + (n – 1) H₂O (6)

In these reactions, n is any integer. Reactions (2) and (4) produce paraffinic hydrocarbons and are the dominant reactions for most F-T catalysts. Reactions (3) and (5) produce olefins and reaction (6) produces alcohols, although these reactions usually are not as prevalent as those that produce the paraffins. Note that for any n larger than 4, the production of even paraffinic hydrocarbon fuels already has a H/C ratio of less than 2.5, so this reaction by itself is approaching the desired H/C ratio for an ISPP plant. Note also that these reactions are all exothermic with heats of reaction on the order of 30 to 40 kcal/mole.

The F-T reaction was discovered in 1923 in Germany but, even today, has not been fully characterized. The reaction generally uses an iron-based catalyst, which may be mixed with some cobalt, copper, or manganese for hydrogen rich feeds. Actual activity of the catalyst depends on the presence not only of metallic iron, but also of iron oxide and iron carbide. The reaction proceeds at moderate temperatures of about 200 – 280 degrees C and moderate pressures anywhere from 1 to 25 bars. It produces a hugely varying range of product compositions and weights depending on the exact formulation of the catalyst used. The predominate products in most Fischer-Tropsch reactors are middle weight straight chain paraffinic hydrocarbons from ethane (C₂H₆) up through decane (C₁₀H₂₂), although specific types of catalysts will produce considerable amounts of unsaturated hydrocarbons or alcohols via reactions (3), (5), or (6). In general, the lighter fractions, such as ethane and propane, dominate in molar terms, but in terms of mass fraction, it is not unusual for the heavier hydrocarbons to form the majority of product. In the Phase I MAHOSS program, Pioneer used iron Fischer-Tropsch catalysts doped with potassium to produce high olefin yields, primarily via reaction (5).

Production of Benzene and Toluene

Aromatic components such as benzene and toluene can be produced by subjecting F-T reaction products, particularly olefins, to catalytic reforming using a shape specific zeolite. The zeolite most often mentioned for olefin aromatization in the literature is ZSM-5, a silica/alumina zeolite developed by Mobil in the 1960’s. The zeolite pores are just the proper size to piece together multiple olefin molecules via a reaction that looks something like (7):

3 C₂H₄ → C₆H₆ + 3 H₂ ΔH = -17.7 kcal (7)

Note that if one of the ethylene molecules is a propylene molecule, you will get toluene; if two molecules are propylene, you will get xylene. Since aromatization occurs directly to the F-T products, there are two possibilities for a reactor to create these aromatics. First, the reactor could be a separate vessel immediately following the F-T reactor. Second, you could put a mixture of both catalysts (Fischer-Tropsch and ZSM-5) in the same reactor. Various articles in the literature reported success with the second technique (particularly Caesar, et. al., 1979), so Pioneer pursued this option during the Phase I portion of the MAHOSS program and achieved excellent production of aromatic liquids.

Product Recovery and Separation

The aromatic and heavy olefin products from the reactors can be easily separated in a condenser, since they are liquid at reactor pressure and ice water temperatures, while CO, CO₂, hydrogen, and other components are gases at these conditions. A second membrane, similar to the one used in the RWGS system can be used to separate F-T reactants from CO₂ effluent; the CO₂y and remaining hydrogen can be sent back to the RWGS while the unreacted CO and uncondensed hydrocarbon products can be returned to the fuel production reactor. This process configuration was demonstrated during the Phase I portion of the MAHOSS project.

TOPICS
In-Situ Resource Utilization (ISRU), Rocket Propulsion, Chemical Processes, Spacecraft Systems, Synthetic Fuels, Space Technology, Thermal Engineering, Systems Engineering

The Mars Aqueous Processing System (MAPS) is a novel technology for recovering oxygen, iron, and other constituents from lunar and Mars soils. The closed-loop process selectively extracts and then recovers constituents from soils using sulfuric acid and bases. The emphasis on Mars is production of useful materials such as iron, silica, alumina, magnesia, and concrete with recovery of oxygen as a byproduct. On the Moon, similar chemistry is applied with initial emphasis on oxygen production.

During a NASA SBIR Phase I project, Pioneer Astronautics achieved program objectives by demonstrating the major MAPS unit operations in the laboratory. Magnesium sulfate was extracted from a Mars duricrust simulant and then recovered by crystallization from solution. Magnesium sulfate was decomposed to sulfur dioxide and oxygen gas while generating magnesium oxide. Sulfur dioxide and oxygen were combined with water to form sulfuric acid using a low-temperature, liquid-phase catalytic process. Acid produced by this method was used to selectively extract iron and other constituents from simulant. Iron was recovered from solution as a high-grade oxide concentrate (80% Fe2O3). The iron oxide was reduced to iron at temperatures less than 750oC. Other byproducts, such as alumina and silica, were also recovered from solution as high-grade precipitates by controlling time, temperature, and acidity. Samples of structural materials formed from spent simulant and extracted magnesium compounds exhibited a compressive strength of over 800 psi.

A subset of MAPS was demonstrated to be equally useful for lunar applications. Iron oxide was successfully extracted from lunar simulant in the laboratory. High-grade iron oxide concentrate recovered using MAPS technology could reduce thermal power requirements for lunar oxygen production by an order of magnitude compared to heating of bulk soil for hydrogen reduction.

MAPS can be commissioned first for lunar applications to generate oxygen with reduction in thermal power compared to bulk soil treatments. Later modular improvements can add recovery of iron, other metals, and metal oxides (including silicon dioxide, a key precursor for photovoltaic panel production). On Mars, all of these products plus concrete can be produced. The process can be developed in parallel for lunar and Martian purposes, resulting in mission cost savings and risk reduction.

Designs for lunar and Mars applications of MAPS are being developed under a Phase II NASA SBIR program. Kris Romig is the Contracting Officer’s Technical Representative for NASA Johnson Space Center. Mark Berggren is the Principal Investigator for Pioneer on the SBIR Phase I program.

NASA SBIR 2003 Solicitation | Phase I

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
Mars Aqueous Processing System (MAPS) is an innovative method to produce useful building materials from Martian regolith. Acids and bases produced from the regolith are used to aid the preparation of metals (such as iron) and cement ingredients (such as lime and aggregate) for construction of habitats and infrastructure needed for early human colonization. As more regolith is processed, more acids and bases will be produced for use in manufacture of plastics, metals, polymers, and reagents useful for later, larger-scale human habitation. With the apparent abundance of water in certain locations on Mars, the proposed technology will enable the manufacture and fabrication of a variety of materials using only Mars indigenous materials with the use of processing equipment and catalysts brought from Earth.

The proposed processing methods are capable of extracting and separating regolith constituents via aqueous extraction followed by selective precipitation based on solution pH and oxidation potential. Thermal treatments such as drying (to remove moisture), roasting (to remove volatile sulfate and chloride acid precursors), and oxide reduction (using hydrogen, carbon monoxide, or carbon derived from Martian water and atmosphere) are integrated with the aqueous extraction methods to manufacture the basic building materials required to facilitate human habitation.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
The construction of Mars bases will be enabled by technologies that are capable of producing needed structural materials from Mars resources. The proposed MAPS concept uses available Mars resources to produce strong components under less-severe manufacturing conditions than alternative methods. Therefore, MAPS provides the additional benefit of reducing the size and mass of the required manufacturing equipment delivered to Mars from Earth.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
Manufacturing technologies aimed at optimizing utilization of raw materials, minimizing equipment mass and complexity, and minimizing operating power have potential payoffs for terrestrial applications. In particular, developing regions with limited resources could produce useful products if strict composition and manufacturing requirements for export are not needed.

TOPICS
Life Support, Chemical Processes, Space Technology, Mineral/Ore Processing, Materials, Systems Engineering, In-Situ Resource Utilization (ISRU)

Firing the hot graphite rod engine on the test stand. The visible CO2 exhaust shows we have utilized nearly all the heat capacity of the bed hot exhaust is completely invisible

The protoflight version of the gashopper chained to the floor for tethered static tests.

The protoflight version of the gas hopper undergoing flight/hover test. The initial mass of the hopper is 51 lbs, and the engine thrust is rated 40 lbf. A 20 lbf. assist from a counter balance allowed flight and provided upward guidance to avoid upsetting the gashopper.

The balloon assisted Gashopper flight.

Left: CO2 gashopper before liftoff. Fully loaded, the gashopper weighs 51 lbs, and requires a little help from a helium weather balloon providing 20lbs of lift.

Center: Liftoff! The gashopper ascends rapidly on about 50 lbs thrust.

Right: The CO2 gashopper reached an estimated 40 meter apogee, then started downward towards a planned DC-X style landing.

TOPICS
Spacecraft Systems, Mars Analog Field Exploration, Rocket Propulsion, In-Situ Resource Utilization (ISRU), Vehicles, Space Technology

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Integrated Mars In-Situ Propellant Production System (IMISPPS) is an end-to-end system that will produce rocket propellant on Mars from CO2 in the Martian atmosphere. The IMISPPS combines the RWGS and Sabatier reactions in a single reactor to produce a useful high-specific impulse fuel (methane plus carbon monoxide) and water, which is condensed and electrolyzed to produce oxygen and hydrogen. The hydrogen is recycled back to the Sabatier/RWGS reactor to react with Martian CO2 to produce more fuel, while the oxygen is cryogenically stored to provide oxidizer. Some of the carbon monoxide is removed by cryogenic separation to increase propellant specific impulse. Carbon dioxide acquisition to feed the fuel reactor is accomplished using a lightweight freezer. Use of the IMISPPS has the advantage of producing all the oxygen needed to burn the methane with only in a single catalytic reactor required. In the proposed work, we will build a brassboard core of the IMSIPPS and demonstrate its performance and reliability.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
IMISPPS provides a technology for producing methane/oxygen rocket propellant on Mars with the correct oxidizer to fuel ratio in a single reactor using the CO2 atmosphere of Mars as the primary raw material. The leverage of the hydrogen imported from Earth would be at least 20, greatly reducing the cost and difficulty of sample return and human missions to Mars. Operation of a single reactor would reduce the system weight and complexity compared to systems based on the Sabatier process and an oxygen production process. Further processing of the methane and CO to benzene would increase the leverage to 53.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 100 WORDS)
On Earth, the IMISPPS has applications in the area of carbon dioxide sequestration and processing to reduce the greenhouse effect. For example, cement kilns emit CO2 in high concentrations amenable to separation and processing. In conjunction with a renewable or nuclear power supply that generates hydrogen, the IMISPPS could be used to combine the hydrogen and CO2 to convert the hydrogen into a readily transportable form (methane) that easily fits into the existing energy infrastructure. The carbon monoxide could be combined with hydrogen using the Fischer-Tropsch process to produce valuable hydrocarbon products, such as alcohols, olefins and waxes.

TOPICS
Spacecraft Systems, Rocket Propulsion, Chemical Processes, Space Technology, Systems Engineering, In-Situ Resource Utilization (ISRU), Vehicles, Carbon Capture

Mars Micro Balloon Probe

Press Release [29 August 2000]

The Mars Micro Balloon Probe (MMBP) is a project to create a low-cost airborne Mars photographic probe with a trans-Mars injection (TMI) mass an order of magnitude less than that of any Mars balloon probe designed to date. This can be done by approaching the gondola design in a spirit of ruthless minimalism, reducing it to a single instrument coupled with a computer, UHF radio transmitter/receiver, and a primary battery power supply suitable for a 1 day flight. In addition, large mass savings and greater simplicity and reliability can be achieved by replacing the traditional complex high pressure hydrogen or helium inflation gear coupled with super-pressure or over-pressure balloons with a novel self inflating zero-pressure balloon using ammonia or water vapor for lift. Combining these innovations, it is possible to create MMBP units with a total TMI mass, including entry system, on the order of 10 kg. Such light weight systems would be prime candidates to fly as hitch-hiker payloads on any of the numerous Mars missions planned for the near future. This would allow high resolution aerial photography to be performed on Mars without the loss of any of the surface or orbital science currently planned.

Burst and Tumble at 107k’ 6sec

Longer Version 11 sec

Test balloon inflation from different angle at 97k’ 67 sec

During the Phase 1 MMBP program, Pioneer demonstrated a working prototype of a 1 kg Mars aerial photographic gondola. During the Phase 2, we have concentrated on developing our innovative system for autonomous inflation of Mars balloons. On Aug, 26, this system was successfully demonstrated over Byers Colorado at an altitude of 100,000 ft.

Pioneer’s work on MMBP systems is supported by SBIR funding from Jet Propulsion Lab.

Test Balloon Deployed from above

Test Balloon inflating from below

Jack Jones(JPL), Mike Manes (EOSS), and Robert Zubrin (PA) watching the downlink

TOPICS
Spacecraft Systems, High Altitude Balloon Experiments, Radar Systems, Space Technology

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Methane and Carbon Monoxide/LOX rocket (MCLOX) is a technology for accomplishing ascent from Mars. Current Mars in-situ propellant production (ISPP) technologies produce methane and carbon monoxide in various combinations, but with neither generally produced in pure form. While separation and purification of methane fuel is possible, it adds complexity to the propellant production process and discards an otherwise useful fuel product. The MCLOX makes such complex and wasteful processes unnecessary by burning the methane/CO mixtures produced by Mars ISPP systems without further refinement. Despite the decrease in rocket specific impulse caused by CO admixture, the improvement offered by concomitant increased propellant density provides a net improvement in stage performance, and this mission advantage is amplified further by the increase of the total amount of propellant produced and the decrease in mass and complexity of the required ISPP plant. For these reasons the development of the MCLOX rocket is important to achieve maximum benefit from Mars ISPP systems.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
MCLOX propulsion would be uniquely suitable for a Mars sample return or a manned Mars mission using indigenous propellant. Because it can use the mixtures of methane and CO that are produced by the most readily available ISPP systems without further distillation, it minimizes the mass and complexity of such systems and maximizes their useful propellant leverage. MCLOX engines could also be used to take off from the Moon or asteroids, and can be used to fulfill any function required for high-energy space storable propulsion.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
The MCLOX has many commercial applications, since it can also work with pure methane fuel. An upper stage driven by such LOX/methane propulsion would be non-toxic, space storable, and offer the highest specific impulse of any chemical stage other than the much more expensive and bulkier LOX/H2 systems. As such, the MCLOX upper stage would find many customers among commercial, military, and scientific satellites, upper stages, and launch vehicles who would value it as a cost-effective alternative to the current choice of toxic hypergol, lower performing LOX/RP, or cryogenic LOX/H2 propulsion.

TOPICS
Spacecraft Systems, Rocket Propulsion, Chemical Processes, Space Technology, Thermal Engineering, Systems Engineering, In-Situ Resource Utilization (ISRU)

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Mars Solar Balloon Lander (MSBL) is a novel concept which utilizes the capability of solar-heated hot air balloons to perform soft landings of scientific payloads on the Martian surface. In the MSBL concept, a dark colored or metalized zero pressure balloon is inflated with Martian atmospheric CO2 during initial descent suspended by a parachute.. As a result of the favorable optical qualities of the balloon?s coloration, the gas inside the balloon is warmed to temperatures considerably exceeding the surrounding ambient atmosphere, thereby providing buoyancy. The MSBL can thus achieve stable level flight during daylight, or can be used to deliver payloads to the ground with arbitrarily low rates of descent. After the payload is landed, the balloon can be released for a free flight remote sensing mission, or can be retained as a tethered asset by the lander serving many useful functions, including local aerial imaging, communications, or lander towing. Key technical challenges to the MSBL concern dealing with horizontal velocity during terminal descent. However the MSBL is competitive on a mass basis when compared to alternative landing technologies such as airbags, and offers many novel additional capabilities for combined surface and aerial operations.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
The primary purpose of the MBSL is to provide a low cost, low mass means of soft landing payloads on Mars. The system can also be used to tow the lander at considerable speed across the Martian landscape, simultaneously providing aerial context imaging and other remote sensing data from its own gondola carried aloft. The solar balloon can increase its lift and fly the payload across chasms, as necessary. Such a mission could travel long distances across the Martian pole in summer, producing a science bonanza as it images and samples the surface across a wide swath of sites.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
The MBSL perform surveys across the polar regions of the Earth, mapping the circulation and sounding the ozone layer at many altitudes through long periods of polar summer. The MSBL landing system is also an attractive means of delivering payloads to the ground. Currently, the primary means of air-delivering payloads to areas where aircraft landing is impossible (and helicopter capacity is inadequate or too expensive) involves paradrop, which can be very violent and damaging to the payload. Using a solar balloon lander, much gentler delivery of equipment from air to ground could be achieved.

TOPICS
Spacecraft Systems, Mars Analog Field Exploration, Space Technology, Vehicles, In-Situ Resource Utilization (ISRU)

Methane to Aromatics on Mars

NASA SBIR 00-1 SOLICITATION

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Methane to Aromatics on Mars (METAMARS) project will design and build a machine for converting methane produced by the Sabatier reaction into a low hydrogen content, low vapor pressure, high density, aromatic fuel. Its primary advantage is a factor of four reduction of hydrogen feedstock importation requirements for production of rocket fuel compared to the standard S/E process for an equivalent mass of fuel. In addition, since all oxygen produced by the Sabatier system comes from the carbon dioxide feed, by reducing the hydrogen in the fuel the METAMARS process will simultaneously improve the stoichiometry of the fuel/oxidizer combination and will reduce the power required by the process. The benzene fuel is also considerably denser than methane fuel. A final advantage of the METAMARS process is that it operates at low pressure, in contrast to synthesis reactions for other higher hydrocarbons. These advantages make the METAMARS process a prime technology to improve the applicability of the Sabatier process for small scale unmanned Mars missions, such as the Mars Sample Return (MSR) mission, as well as a key technology for manned Mars missions.

POTENTIAL COMMERCIAL APPLICATIONS

There is already considerable terrestrial interest in the conversion of methane into transportable fuels and chemical feedstocks, with the majority of the work focusing on methanol or Fischer-Tropsch hydrocarbon production. Remote gas fields currently flare or reinject more than a billion dollars worth of natural gas every day, and economical means for liquefying this gas have tremendous commercial potential. Pioneer’s work will draw heavily on research into catalytic processes for methane dehydrogenation, but the METAMARS process itself is, to Pioneer’s best knowledge, new and unique. Potential products include benzene and toluene, with total current markets of more than 10 million tonnes per year.

The METAMARS system is an in situ resource utilization (ISRU) technique that can convert methane produced from the carbon dioxide in the Martian atmosphere to low hydrogen content liquid aromatic fuels for an Earth Return Vehicle, thus greatly increasing the leverage of the hydrogen imported from the Earth. More importantly, the METAMARS system reduces the amount of hydrogen imported from Earth by a factor of four, leading to dramatic reductions in mission cost. This project involved design and construction of an oxygen/aromatic hydrocarbon production facility sized to produce 1 kg of bipropellant per day. Because aromatic fuels contain only about one hydrogen atom per carbon atom, such a system would give extremely high leverages on the order of 53 in the production of fuel and oxidizer for a Mars Sample Return mission and human Mars missions. Project work was carried out from February through August 2001.

Clyde Parrish was the Kennedy Space Center (KSC) program manager and Anthony Muscatello was the principal investigator at Pioneer Astronautics.

TOPICS
In-Situ Resource Utilization (ISRU), Carbon Capture, Chemical Processes, Systems Engineering, Space Technology

Methanol Ejector Ramjet

NASA SBIR 1996/1997 Solicitation

The Ejector Ramjet is the simplest of all Rocket-Based combined Cycle (RBCC) propulsion systems. In the past the two main propellants of interest for such systems have been hydrogen and kerosene. For many applications, the ease of handling of kerosene makes it a more desirable fuel than hydrogen, despite the latter’s’ much higher energy-to-mass ratio. However, at flight speeds beyond Mach 4 to 4.5, kerosene fails because the high temperatures generated in the engine causes it to break down and coke up the system. Methanol’s energy/mass ratio is even lower than kerosene, causing it to not be considered in the past for Ejector Ramjet propulsion. However, because methanol, CH3OH, contains one oxygen atom for every carbon atom, high temperatures will cause it to decompose not into carbon and hydrogen, but carbon monoxide and hydrogen. Thus, when used as an engine coolant, methanol should not cause coking. Because the maximum flight velocity of a hydrocarbon fueled Ejector Ramjet is limited not by fuel energy, but by cooling capacity, methanol may be a better option for fueling such systems than kerosene. We therefore chose to examine the potential performance of a methanol fueled ejector ramjet system.

In the course of the Methanol Ejector Ramjet (MER) program, 18 methanol, burning rocket engines were built, and tested inside of five different ramjet engines in static tests. Over 1000 seconds of hot engine firings were achieved. A flight vehicle was built, and is currently available for testing.

TOPICS
Spacecraft Systems, Rocket Propulsion, Chemical Processes, Space Technology, Mechanical Engineering, Vehicles, Systems Engineering

Methanol Lift Gas Cracker

NASA SBIR 2003/2004 Solicitation

View from Lift Gas Cracker Test Balloon

(100,000 feet over Denver, Colorado; Photo courtesy of K. Mark Caviezel)

The Lift Gas Cracker (LGC) is a new method for extending the duration of high-altitude scientific balloon flights and for enabling the launch of balloons from remote locations. The LGC produces balloon lift gas by catalytic steam reforming of methanol to generate hydrogen and carbon dioxide.

A portable LGC methanol reforming system was prototyped by Pioneer Astronautics during a NASA SBIR Phase I project. The LGC demonstrated lift gas generation for launching meteorological balloons from remote locations where heavy helium cylinders are not available. The LGC can also generate hydrogen for fuel cells by incorporating a gas separation step.

During Phase II, the LGC is being developed for extending high-altitude balloon flight duration. For this application, an on-board LGC produces lift gas at night. During the day, some of the lift gas (which is predominantly hydrogen) is burned at low pressure with atmospheric air to produce water ballast. The water ballast can be dropped or can be recycled to the LGC for steam reforming of methanol at night. These techniques can extend the duration of high-altitude flights by a factor of three or more compared to conventional methods of dropping ballast at night and venting gas during the day.

Phase I Lift Gas Cracker Technical Accomplishments

• Demonstrated a portable, methanol-heated LGC producing 100 standard liters per minute of lift gas,
• Demonstrated membrane separations to produce lift gas with an average molecular weight of about 6,
• Demonstrated catalytic hydrogen/air combustion and trapping of water ballast at about 40 millibar, thus enabling a significant technology for extension of flight duration.
• Filled and launched a 1500-gram latex balloon (~560m3 at burst). The carrier balloon lifted smaller zero-pressure helium and lift gas balloons along with a camera plus telemetry, altitude, and temperature sensors to 100,000 feet to collect valuable information regarding the performance of the lift gas cracker system as well as the effect of carbon dioxide on balloon temperatures.
• Conducted mission analyses demonstrating significant extension of high-altitude flight duration using LGC.

Lift Gas Cracker Demonstration Flight (May 30, 2003)

Phase II Lift Gas Cracker Tasks

• Stratospheric flight test of lift gas cracker, and
• Stratospheric flight test of low-pressure combustor.

NASA Applications

NASA applications of the Lift Gas Cracker include greatly extending scientific balloon flight duration. Use of the LGC to provide nighttime makeup gas during flight can nearly double the duration of a stratospheric zero-pressure balloon flight over what is now possible. If, in addition, the excess lift gas in the balloon during daytime is reacted with air to produce water ballast instead of simply being vented, the flight duration can be tripled. This is an extraordinary benefit for all types of stratospheric scientific ballooning. The LGC is also applicable to balloon flight on Mars, Venus, and Titan.

Non-NASA Commercial Applications

The LGC can produce lift gas in remote areas, such as the Arctic, Antarctic, and remote regions where conventional gas cylinders are difficult and costly to obtain. Such gas is needed in large quantities to support meteorological campaigns and flight service stations that provide winds aloft data to pilots. In addition, winds-aloft information is needed by military units, such as artillery, which fire projectiles through high altitudes. In remote regions, the provision of compressed helium bottles to field units could prove difficult, and the LGC offers corresponding logistic advantages.

Debora Fairbrother was the NASA Contracting Officer’s Technical Representative at the Goddard Space Flight Center during Phase I. Dr. Robert Zubrin is the Principal Investigator for Pioneer. Mark Berggren is Pioneer’s lead engineer for the LGC project. Jerry Sterling is the COTR at Goddard for the Phase II Lift Gas Cracker program.

NASA SBIR 02-1 Solicitation | Phase I

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Lift Gas Cracker (LGC) is new method for producing lift gas for balloons for both terrestrial and extra-terrestrial applications, eliminating the need for heavy, bulky gas bottles. The LGC produces a lift gas from the cracking of methanol, an easily obtained, easily stored liquid, to make either the hydrogen/carbon monoxide cracking product or a hydrogen-enriched product by using membrane separation. Combustion of about 10% of the methanol provides sufficient energy to crack the remainder. The LGC product gas can be converted back into methanol, allowing a balloon to maintain altitude during the daytime without having to vent lift gas. The methanol can then be converted back into lift gas at night to maintain altitude without dropping ballast. Thus, long duration balloon flights could be extended a factor of 20 or more compared to conventional means.

POTENTIAL COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
In addition to supplying lift gas in remote locations, such as during military operations or worldwide weather balloon launch campaigns, LGC in the membrane separation mode could supply cheap, easily generated hydrogen for multiple fuel cell applications, including automobiles and remote power supply stations.

POTENTIAL NASA APPLICATIONS (LIMIT 150 WORDS)
NASA applications would include ultra-long duration balloon flights for earth observation, greatly extending the time aloft. Other applications would be the production of lift gas in remote areas, such as the Arctic and the Antarctic, where conventional gas bottles are difficult and costly to obtain. LGC could also be used for long-duration balloon flights on Mars and Venus, enabling much more time for data acquisition compared to conventional technologies.

TOPICS
High Altitude Balloon Experiments, Systems Engineering, Chemical Processes

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The multi-cell thermal battery (MCTB) is a device that can recover a large fraction of the thermal energy from heated regolith and subsequently apply this energy to heat up cool regolith. The individual cells of the MCTB contain a thermal storage media that is specifically designed for optimal performance at a given temperature range. Each of these cells is charged with thermal energy from hot regolith that has been used in a lunar ISRU application. Once the MCTB is charged, the heat is transferred from the battery to newly harvested regolith. In this manner over 85% of the heat can be transferred from the expended to the new regolith. This is a large improvement especially considering that this reduces the heating requirement to produce 1000 kg of O2 from lunar regolith from an average of 1 kW to only 0.15 kW (assuming 3% O2 recovery by weight). The other irreducible power consumption of lunar ISRU O2 production is electrolysis which consumes at least 0.3 kW. Hence, using the MCTB decreases the irreducible power consumption of lunar ISRU by 65 %.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The primary initial application of the multi-cell thermal battery is for heat recovery from expended lunar regolith to new regolith during O¬2 production via hydrogen reduction. The MCTB reduces the heating requirement for lunar O2 production by 85% and reduces the overall power requirement by 65% given an O2 production rate of 1000 kg/year and 3% O2 recovery by weight. During the phase I the superior performance of the MCTB will be demonstrated by transferring over 60% of the heat from hot regolith at 800 oC to cold regolith initially at 25 oC. During the subsequent Phase II and Phase III programs, the MCTB will be further optimized and refined to integrate seamlessly with state-of-the-art lunar regolith reduction systems. The MCTB will be light-weight and provide significant energy savings for lunar ISRU applications.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
The multi-cell thermal battery will be a valuable tool wherever heat needs to be stored or transferred from one solid substance to the next. Any chemical process that requires high temperatures and operates in batch mode would greatly benefit from using the MCTB to conserve energy. Batch furnaces for hardening and annealing metals, for example, require high temperatures and operate in batch mode. Using the MCTB in this application could greatly reduce the operational cost of such facilities. In such a process the expended product can be used to fill the thermal battery with heat and this heat can then be used to pre-heat the following batch. As shown before, this can lead an energy savings of over 85%.

Expected Technology Readiness Level (TRL) upon completion of contract: 4 to 6

TOPICS
Energy, Thermal Engineering, In-Situ Resource Utilization (ISRU), Mechanical Engineering

TECHNICAL ABSTRACT (Limit 2000 characters, approximately 200 words)
The Nitrous Ethane-Ethylene Rocket with Hypergolic Ignition (NEERHI) engine is a proposed technology designed to provide small spacecraft with non-toxic, non-cryogenic, high performance, hypergolic propulsion. When passed over a warm ruthenium catalyst bed, gaseous nitrous oxide and an ethylene-ethane gaseous blend combust instantly. A small 1 N thruster can be designed to provide small satellite propulsion systems with a specific impulse of approximately 300 seconds. Both propellants are self-pressurizing, capable of delivering feed line pressures in excess of 800 psi at room temperature, and 400 psi if cooled to 0?C. For longer duration missions, both nitrous oxide and an ethane-ethylene fuel blend do not require thermal heating to maintain a liquid state, and as such, can be stored on Earth or in space for in-definite periods of time with no parasitic power drain required to maintain a liquid propellant. Compared to other available chemical propulsion systems, a NEERHI system offers a cost effective solution as other hypergolic engines use hydrazine and nitrogen tetroxide which are toxic and dangerous to handle, increasing ground costs. As an added capability, the NEERHI engine has the ability to operate as a monopropellant engine if the ruthenium catalyst be is heated with a bipropellant reaction, increasing the lifetime of the catalyst bed and reducing heating loads on the engine.

POTENTIAL NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
A NEERHI system is capable of replacing any monopropellant or bipropellant space propulsion system currently used by NASA with a green propellant, self-pressurizing, cold-storable, hypergolic rocket system. The recent MAVEN mission, which uses a propulsion system based off of the Mars Reconnaissance Orbiter, uses a total of 20 hydrazine monopropellant thrusters. A NEERHI system could be adapted to future missions to provide a greater specific impulse with a much lower ground cost due to the low toxicity of the propellants. Future lunar missions, which have historically used an NTO and MMH propellant engine, could use a NEERHI system to not only provide RCS thrust, but the nitrous oxide can also be used to produce a breathable atmosphere for any manned mission. The current technology roadmap for NASA also features a main propulsion unit for the micro-satellite, which could employ a NEERHI engine to provide delta-V maneuvers, station keeping, and even Earth-escape missions. Almost all satellite systems that don’t have ion RCS systems could greatly benefit from the integration of a NEERHI unit to reduce the launch cost of the system.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (Limit 1500 characters, approximately 150 words)
A NEERHI system can be used on any commercial satellite system that requires a simple, hypergolic, RCS propulsion unit but wishes to avoid the difficulties encountered when working with a nitrogen tetroxide and hydrazine system. The NEERHI can be used in the emerging cubesat industry, were the primary development teams are university students designing their first space system. A NEERHI engine would provide a safe and affordable system for universities that often have rigorous safety standards, and as such, avoid current hydrazine-based propulsion. In the new field of commercial crew development efforts, the SpaceX capsule currently uses the Draco rocket engine to provide attitude control. The Draco uses an MMH and NTO propellant combination. A NEERHI system could be built to replace these thrusters, and with a supply of Nitrous oxide onboard, future Dragon spacecraft could use the nitrous to produce breathing air instead of bringing along an additional system, taking up mass and space on the craft. A hypergolic and green propellant is the solution sought by all companies to phasing out the use of the dangerous hydrazine-based thrusters, and the NEERHI program could revolutionize the market.

Estimated Technology Readiness Level (TRL) at beginning and end of contract: 
Begin: 3
End: 4

TOPICS
Spacecraft Systems, Synthetic Fuels, Rocket Propulsion, Chemical Processes, Space Technology, Vehicles, Systems Engineering

Nitrous Oxide Based Oxygen Supply System

Update: Pioneer Astronautics was issued United States

Patent Number 6,347,627 for the NOBOSS technology

NASA SBIR 1999 Solicitation | Phase I

The Nitrous Oxide Based Oxygen Supply System (NOBOSS) is a method for storing oxygen and nitrogen for EVA breathing and mobility. Nitrous oxide, N2O, is a common storable chemical that potentially could be used as a convenient, low cost, lightweight, safe and reliable source of oxygen and nitrogen. A nitrous oxide atmosphere supply would have the following advantages: easier storage with greater density at much lower pressure, no cryogenic insulation problems, longer duration oxygen supplies with less weight, and much lower fire hazard than compressed or liquid oxygen. In the NOBOSS, nitrous oxide is decomposed via heating in a catalytic reactor, and as the decomposition is exothermic, the N2O will continue to decompose without further energy input. The NOBOSS can be used to supply breathing gas to existing 4.3 psi space suits by employing an air separation membrane or molecular sieve to eliminate the nitrogen. The waste nitrogen could then provide propellant gas for an MMU. Alternatively, the 2/3 nitrogen, 1/3 oxygen gas mixture produced by N2O dissociation is the ideal gas supply for an 8 psi space suit. In addition, expansion of stored liquid N2O produces cold gas that can be used to help cool a spacesuit.

Project summary

The Nitrous Oxide Based Oxygen Supply System (NOBOSS) is a method for efficient generation of breathable gases for space, terrestrial, and marine applications. Nitrous oxide (N₂O) is a common storable liquid that can be used as a convenient, low-cost, light-weight, safe, and reliable source of breathing gas. NOBOSS offers non-cryogenic storage with greater density at much lower pressure and much lower fire hazard than supply systems containing 100% oxygen. The primary advantage of nitrous oxide is the small volume and low mass of storage tanks to support long-duration breathing needs. Nitrous oxide liquid at 750 psi is about 3 times denser than compressed air or oxygen at 3,000 psi. Compared to high-pressure breathing gas cylinders, the same mass of N₂O can be stored in one-third the volume with one-twelfth the storage cylinder mass.

Nitrous oxide can be decomposed thermally in the presence of a catalyst, and the exothermic dissociation continues without further energy input. The resulting product gas, following polish filtration of parts-per-million byproducts, is 33.3 percent oxygen (by volume) with balance nitrogen, suitable for breathing. The following reaction describes the dissociation of nitrous oxide to oxygen and nitrogen.

N₂O → N₂ + ½O₂ + 81.8 kJoules

After initial in-house proof-of-concept testing, Pioneer developed the NOBOSS technology during NASA SBIR Phase I and II programs. A novel, light-weight, efficient catalytic reactor system incorporating simple pneumatic flow and thermal controls was built. Pioneer’s reactor design achieves virtually complete nitrous oxide dissociation with very low byproduct gas formation. The gas production system requires no external power after initial start up. The nitrous oxide feed pressure (about 750 psi at 20oC) is used to deliver gas to the system and provides motive force for cooling and downstream gas cleanup systems. Product gas is delivered to the user through a self-pressurized surge tank. A fully operational prototype backpack was built and demonstrated during the SBIR Phase II program.

Michael Rouen was the NASA Contracting Officer’s Technical Representative at Johnson Space Center for the NOBOSS project. Dr. Robert Zubrin was the Principal Investigator for Pioneer Astronautics. Mark Berggren was the lead engineer for Pioneer on the Phase II project.

TOPICS
Life Support, Spacecraft Systems, Chemical Processes, Space Technology, Systems Engineering

Nitrous Oxide Emergency Power Unit

NASA SBIR 2003 Solicitation | Phase II

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Nitrous Oxide Monopropellant Rocket (NOMR) is a new Extravehicular Mobility Unit (EMU) thruster concept using nitrous oxide as a monopropellant. Liquid monopropellants are often used in propulsion systems where simplicity of design, restartable/control on demand, and repeatability is desired. Unfortunately, many monopropellants are toxic and dangerous, ruling them out for EMU thruster application. Thus, EMU thrusters have relied upon cold gaseous nitrogen, which offers very low specific impulse and propellant mass fraction. A NOMR, however, uses nitrous oxide, a readily available safe and storable propellant which is not toxic, has performance comparable to hydrazine, and does not decompose spontaneously like hydrogen peroxide. Furthermore, Pioneer Astronautics has demonstrated a system that decomposes N2O into a breathable mix of oxygen and nitrogen. Thus, for example, an EMU propelled by a NOMR system would provide an astronaut with a large emergency backup supply of oxygen. Such a dual use system could also have great utility as the propulsion system for manned spacecraft, such as the International Space Station or the Space Shuttle, where safety is paramount and breathing gas reserves are desired. Replacing current liquid monopropellant thrusters with NOMRs would greatly reduce ground processing time and costs, while providing comparable performance.

POTENTIAL COMMERCIAL APPLICATIONS
The Nitrous Oxide Monopropellant Rocket has the potential of promoting use of liquid monopropellant rockets. Currently, the difficulty and cost in safely handling and storing conventional liquid monopropellants has hindered their use. Conversely, less toxic monopropellants tend to have limited specific impulse. The NOMR would not have these limitations. The market for NOMRs is potentially large, and goes well beyond EMU propulsion. Hydrazine is commonly used as a monopropellant in attitude control thrusters in spacecraft. Unfortunately, hydrazine is extremely toxic. NOMRs, however, use a benign monopropellant, and so much less care and expense is necessary, while still achieving similar performance. NOMRs would be much more profitable and attractive to use than current common monopropellant rockets systems, and should be able to dominate the market for liquid monopropellant propulsion systems, such as are used in attitude control of spacecraft. Currently, some 1700 satellites are planned for launch in the next 10 years, and all will need attitude control systems. These satellites and their necessary replacements thus guarantee the NOMR a large and highly lucrative commercial market. Very low cost, safe, sounding rockets could also be enabled, with many applications in the educational and amateur rocket communities.

TOPICS
Spacecraft Systems, Rocket Propulsion, Space Technology, Systems Engineering

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Nitrous Oxide-Organic Liquid Monopropellant Rocket (NOOLMR) is a new rocket concept using nitrous oxide mixed with a small amount of a liquid organic fuel, such as an alcohol or hydrocarbon, as a monopropellant. Liquid monopropellants are often used in propulsion systems where simplicity of design, restartability/control on demand, and repeatability is desired, such as in spacecraft reaction control systems (RCS), which exhibit modest performance. Unfortunately, even the best performing monopropellants offer only modest specific impulse (Isp), are extremely toxic, and difficult and expensive to store and handle. Other monopropellants, such as hydrogen peroxide, are safer, but tend to be associated with even lower Isp. A NOOLMR system, however, would have a much higher specific impulse than any current monopropellant rocket systems, and would use a much less toxic propellant than other high specific impulse monopropellants. Replacing current liquid monopropellant propulsion systems with NOOLMR-based systems would provide much greater performance than current systems, while greatly reducing ground processing time and costs.

POTENTIAL COMMERCIAL APPLICATIONS
The Nitrous Oxide-Organic Liquid Monopropellant Rocket has the potential of promoting use of liquid monopropellant rockets. Currently, even the best performing monopellants offer only limited specific impulse and are difficult and costly to safely store and handle for extended periods. For example, hydrazine is commonly used as a monopropellant in attitude control thrusters in spacecraft. Unfortunately, hydrazine only offers an Isp of about 230 s, and is extremely toxic, so great care and expense must be incurred in its handling and containment for only modest performance. NOOLMRs, however, use a relatively benign high-performance monopropellant, and so are cheaper and easier to use. NOOLMRs, with an Isp of over 300 s, would be much more profitable and attractive to use than current common monopropellant rockets systems, and should be able to dominate the market for liquid monopropellant propulsion systems, such as are used in attitude control of spacecraft. Very low cost, safe, sounding rockets could also be available, with many potential applications in the educational and amateur rocket communities.

Rocket Propulsion, Chemical Processes, Space Technology, Systems Engineering

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The Nitrous Oxide Paraffin Hybrid engine (N2OP) is a proposed technology designed to provide small launch vehicles with high specific impulse, indefinitely storable propulsion. In the N2OP engine, the combination of liquid nitrous oxide on solid paraffin as a rocket propellant allows for the development of compact lightweight high performance stages using densely packed propellant tankage. This is because N2O/paraffin hybrids have a very high oxidizer/fuel mixture ratio and because paraffin has a much higher regression rate than typical hybrid hydrocarbon fuels. Propellant slumping can be prevented by molding the paraffin into a 3% by volume graphite sponge matrix. Currently, space launch missions require cryogenic or extremely toxic propellants which are limited in their storage times, reducing their capability for rapid response launch. The much more storable solid propellants have higher cost, and lower performance while still being a large explosive hazard. The N2OP propulsion system also is compatible with ocean temperatures, allowing launch by floating in water. The achievable Isp for this propellant combination using autogenous pressurization is about 235 seconds at sea level and over 310 s in vacuum, making its performance fully adequate to support operation of a safe, fully storable, highly-responsive multi-stage launch vehicle.

POTENTIAL NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
The N2OP is an extremely attractive technology for enabling fast response space launch missions, it would have many other applications as well. Commercial applications would include the delivery of small geostationary, medium, and low altitude satellites into orbit, apogee kick, and also for high efficiency re-assignment maneuvers and end of life superboosting for geostationary satellites. The N2OP engine could be used for onboard spacecraft propulsion or for dedicated upper and transfer stages. As an additional advantage, the same N2O that powers the N2OP can be used in N2O monopropellant thrusters for spacecraft RCS, thereby eliminating hydrazine and reducing costs.

POTENTIAL NON-NASA COMMERCIAL APPLICATIONS (LIMIT 150 WORDS)
A high specific impulse safe, nontoxic long term storable stage would find many customers among commercial, military, and scientific satellites, launch vehicles, or sounding rockets who would value it as a very cost-effective alternative to the current choices of low performance toxic hypergols, cryogenic LOX, or solid propulsion. Smaller N2OP propulsion systems could meet the needs of the amateur experimental rocketry market. The use of N2OP technology for military missile propulsion is also highly attractive, as such systems would be much safer than solids during manufacture, transport, and front line storage, and potentially be much cheaper as well.

TOPICS
Chemical Processes, Rocket Propulsion, Space Technology, Mechanical Engineering

Rare Earth Materials Recovery System

DOD SBIR 2016 Solicitation | Phase I

Rocketplane/Bantam Upper Stage

Private Design Study

An upper stage for a small rocketplane launch vehicle. This project was a private design study program, for more information please contact Pioneer Astronautics.

TOPICS
Rocket Propulsion, Vehicles

RPSEA Green Oil^™

Private Contract

CO₂-EOR is a very promising technique for recovering oil from abandoned small producer wells. However, many small producers cannot apply EOR technology due to high capital costs resulting from site remoteness, limited pipeline access, and long construction lead-times. To serve the small producer needs, Pioneer is developing a modular and truck-portable biomass-based steam reforming technology to create CO₂ for EOR and H₂ for electricity generation. The Pioneer technology will allow small producers to have access to EOR for domestic oil recovery in a timely, scalable, locally field deployable, and cost effective manner. The technology will also produce clean electrical power to offset producer capital and operating costs.

The Pioneer device is unique from other gasifiers in that it is tailored to produce clean, high pressure, high concentrations of CO₂ for on-site well flooding, and H₂ for electrical power for local or grid service. In parallel with this technology development effort, studies are being done to identify and characterize suitable fields for field testing the unit.

To date, Pioneer has designed, fabricated, and performance tested a fully automated 200,000 SCFD demonstrator (Figure 1). The performance of the device with respect to operating flow rate, pressure, and gas composition have been demonstrated, and is sufficient for meeting CO₂-EOR operational requirements. Future plans include completing the scale-up to 1,000,000 CFD output followed by field testing.

Figure 1. Pioneer’s 200,000 SCFD demonstrator for CO2-EOR

TOPICS
Chemical Processes, Energy, Thermal Engineering, Materials, Systems Engineering

Reverse Water Gas Shift

Private Design Study

The Reverse Water Gas Shift (RWGS) reacts hydrogen with CO2 to produce CO and water. After the water is electrolyzed, oxygen produce it stored while hydrogen is recycled to produce more CO and water. Positioned on Mars, it can make an infinite amount of CO and O2 from Martian CO2, using only a few kilograms of water to provide the recycled hydrogen reactant. The CO2 can be used for rocket propellant or life support, while the CO can be used for rocket fuel, as feedstock to make other fuels such as methanol or dimethyl ether (DME), or to reduce Martian iron oxide to make steel. Pioneer Astronautics was the first company to demonstrate an efficient RWGS system and the use of its products to make methanol, DME, and steel. This project was a private design study program, for more information please contact Pioneer Astronautics.

TOPICS
Synthetic Fuels, Chemical Processes, In-Situ Resource Utilization (ISRU), Systems Engineering, Life Support

Pioneer Sea Glider

 DARPA SBIR 2003 Solicitation | Phase II

 DARPA SBIR 2003 Solicitation | Phase II

Solar Sail Microspacecraft

Department of Defense – Missile Defense Agency – SBIR 1999 Solicitation |

The Solar Sail Microspacecraft (SSM) is a low-cost concept for implementing solar sail propulsion on a practical spacecraft with present-day technology. In the SSM, a simple micro-spacecraft derived vehicle is employed which could cheaply investigate multiple targets and simultaneously demonstrate the utility of small solar sails. The SSM reduces technology risk by using off-the-shelf aluminized mylar. A very small core vehicle with short range communication systems drastically reduces the size of the sails, allowing the spacecraft to be launched as a hitchhiker payload. Because the spacecraft is small, the sail is small, allowing it to be self-deployed using either a rolled spring-steel or inflatable self-deploying boom system. Because of its maneuverability, the SSM could visit multiple targets, engaging in photographic inspection of friendly or adversarial satellites. A SSM could be used to disable or destroy other satellites by parking itself in a position where it blocked the target spacecraft’s solar arrays. It could also be used to interfere with the operation of an opponent’s remote sensing vehicle by using its sails to block the view.

Pioneer’s work on the SSM was supported by SBIR funding from the Ballistic Missile Defense Organization.

Abstract:The Solar Sail Microspacecraft (SSM) is a low-cost concept for implementing solar sail propulsion on a practical spacecraft with present-day technology. In the SSM, a simple micro-spacecraft derived vehicle is employed which could cheaply investigate multiple targets and simultaneously demonstrate the utility of small solar sails. The SSM reduces technology risk by using off-the-shelf aluminized mylar. A very small core vehicle with short range communication systems drastically reduces the size of the sails, allowing the spacecraft to be launched as a hitchhiker payload. Because the spacecraft is small, the sail is small, allowing it to be self-deployed using either a rolled spring-steel or inflatable self-deploying boom system. Because of its maneuverability, the SSM could visit multiple targets, engaging in photographic inspection of friendly or adversarial satellites. A SSM could be used to disable or destroy other satellites by parking itself in a position where it blocked the target spacecraft’s solar arrays. It could also be used to interfere with the operation of an opponent’s remote sensing vehicle by using its sails to block the view. This proposed study shall examine the design and construction of a low-cost, near-term SSM vehicle for immediate use in near Earth space

TOPICS
Space Technology, Spacecraft Systems, Mechanical Engineering, Vehicles

TECHNICAL ABSTRACT (LIMIT 200 WORDS)
The magnetic sail, or “magsail” is a concept which propels spacecraft by using the magnetic field generated by a loop of superconducting cable to deflect interplanetary or interstellar plasma winds. Assuming high temperature superconductors with the same current/mass ratio as existing low temperature superconductors, a magsail sailing on the solar wind at a radius of one AU can attain accelerations on the order of 0.01 m/s2, much greater than that available from a conventional solar lightsail. A net tangential force, or “lift” can also be generated. Using these forces, a magsail can transfer payloads to and from any two circular orbits in the solar system in a flight time slightly larger than the Hohmann ballistic transfer time without the expenditure of propellant. A magsail operating within the magnetosphere could interact with the Earth’s magnetic poles to generate a series of perigee kicks to drive a payload of several times the magsail’s e mass to interplanetary space in times scales of a few months. Magsails can also be made to interact with planetary ionospheres to lower orbits. Magsails could be used to create drag against the interstellar medium to decelerate ultrafast interstellar spacecraft, thereby enabling flights to the stars.

POTENTIAL COMMERCIAL APPLICATIONS
Magsails have important potential commercial applications. Many satellites need to be transported from Low Earth orbit (LEO) to higher orbits such as geosynchronous orbit (GEO), and the fact that the magsail can be used to enable such transfers without requiring the expenditure of propellant offers the prospect of extraordinary cost savings. Current launch systems can transport at least four times as much to LEO as to GEO. Thus, if a magnetic sail with a mass less than or equal to the payload it is boosting can be used to move a payload from LEO to GEO, a considerable saving in launch requirement can be realized. Magsails could also be used to deliver NASA space probes to interplanetary destinations. Magsails could also be used to create friction against the Earth?s ionosphere, thereby allowing reusable orbit transfer vehicles to return to LEO without requiring either the expenditure of propellant or putting the system through the trauma of the thermal and gravitational loads associated with aerobraking. By creating drag against the interstellar medium, magsails could be used to decelerate very fast interstellar spacecraft without the expenditure of propellant, thereby making a substantial contribution towards enabling interstellar missions. Finally, magsails can provide shielding for interplanetary spacecraft against solar flares. The multitude of potential applications of magsail technology make it a very promising commercial development.

TOPICS
Space Technology, Systems Engineering

The Magnetic Sail

NASA Institute for Advanced Concepts (NIAC)

The magnetic sail or magsail, is a device which can be used to accelerate or decelerate a spacecraft by using a magnetic field to accelerate/deflect the plasma naturally found in the solar wind and interstellar medium. Its principle of operation is as follows:

A loop of superconducting cable perhaps tens of kilometers in diameter is stored on a drum attached to a payload spacecraft. When the time comes for operation, the cable is played out and a current is initiated in the loop. This current once initiated, will be maintained indefinitely in the superconductor without further power. The magnetic field created by the current will impart a hoop stress to the loop aiding the deployment and eventually forcing it to a rigid circular shape. The loop operates at low field strengths, typically 0.0001 Tesla, so little structural strengthening is required. The loop can be positioned with its dipole axis at any angle with respect to the plasma wind, with the two extreme cases examined for analytical purposes being the axial configuration, in which the dipole axis is parallel to the wind, and the normal configuration, in which the dipole axis is perpendicular to the wind. A magsail with payload is depicted in Fig. 1.

Figure 1. Magsail deployed with payload

In operation, charged particles entering the field are deflected according to the B-field they experience, thus imparting momentum to the loop. If a net plasma wind, such as the solar wind, exists relative to the spacecraft, the magsail loop will always create drag, and thus accelerate the spacecraft in the direction of the relative wind. The solar wind in the vicinity of the Earth is a flux of several million protons and electrons per cubic meter at a velocity of 400 to 600 km/s. This can be used to accelerate a spacecraft radially away from the sun and the maximum speed available would be that of the solar wind itself. While inadequate for interstellar missions, these velocities are certainly more than adequate for interplanetary missions.

However if the magsail spacecraft has somehow been accelerated to a relevant interstellar velocity, for example by a fusion rocket or a laser-pushed lightsail, the magsail can be used to create drag against the static interstellar medium, and thus act as an effective braking device. The ability to slow spacecraft from relativistic to interplanetary velocities without the use of rocket propellant results in a dramatic lowering of both rocket mass ratio and the mission time.

If the magsail is utilized in a non-axial configuration, symmetry is destroyed and it becomes possible for the magsail to generate a force perpendicular to the wind, i.e. lift. Lift can be used to alter the magsail spacecraft’s angular momentum about the sun, thus greatly increasing the repertoire of possible maneuvers. In addition, lift can be used to provide steering ability to a decelerating relativistic interstellar spacecraft.

The magsail as currently conceived depends on operating the superconducting loop at high current densities at ambient temperatures. In interstellar space, ambient is 2.7 degrees K, where current low temperature superconductors NbTi and Nb3Sn have critical currents (depending upon temperature and local magnetic field) of approximately 1.0×1010 and 2.0×1010 Amps/m2 respectively. In interplanetary space, where ambient temperatures are above the critical temperatures of low temperature superconductors, these materials would require expensive and heavy refrigeration systems However the new high temperature superconductors such as YBa2Cu3O7 have demonstrated comparable critical currents in microscopic samples at temperatures of 77 K or more, which would make them maintainable in interplanetary space using simple multi-layer insulation and highly reflective coatings. Assuming that this performance will someday be realizable in bulk cable, we can parameterize the problem of estimating potential magsail performance by assuming the availability of a high temperature superconducting cable with a critical current of 1010 Amps/m2, i.e. equal to that of NbTi. Because the magnets are operating in an ambient environment below their critical temperature, no substrate material beyond that required for mechanical support is needed. Assuming a fixed magnet density of 5000 kg/m3 (copper oxide), such a magnet would have a current to mass density (j/pm) of 2.0×106 Amp-m/kg.

By interacting with the Earth’s magnetic poles, the magsail can generate sufficient force to allow it to drive both itself and a substantial payload up to escape velocity via a series of perigee kicks. Once escape has been reached, the magsail will find itself in interplanetary space where the solar wind is available to enable further propulsion. Magsail operations can enable both maneuvering in heliocentric space and deceleration of ultra-high velocity interstellar spacecraft without the use of propellant. It also provides an option for lowering the orbit of a spacecraft by creating drag against a planetary ionosphere.

Pioneer’s work on Magnetic Sails was supported by funding from the NASA Institute for Advanced Concepts (NIAC.)

TOPICS
Space Technology, Systems Engineering

Ultralight Solar Sail for Interstellar Travel

NASA Institute for Advanced Concepts (NIAC)

Solar sails offer unique promise for interstellar travel because their lack of need for rocket propellant frees them from the velocity limits imposed by the rocket equation. Conventional solar sails made of aluminized plastic films cannot reach velocities relevant for interstellar flight on the power of sunlight alone, however, as their thrust to mass ratio is too low to allow the attainment of high velocities before leaving the solar system. For this reason, alternative concepts have been advanced which involve pushing light sails with very high-powered lasers or other transmitting devices positioned within the solar system. Such schemes, which involve generating and accurately projecting tens to hundreds of terawatts of power across interstellar distances involve a host of very formidable technical challenges.

In this proposal, an alternative approach is advanced, that of manufacturing ultralight perforated solar light sails which operate without plastic backing. Based upon a preliminary analysis of the fundamental physics of such systems, we have found that they can achieve thrust to mass ratios in sunlight at 1 AU on the order of 10 m/s², and that if solar system departure is initiated at 0.1 AU, that terminal velocities on the order of 1% the speed of light can be achieved with no other source of energy than sunlight. The ability to achieve such high velocities with such a simple and relatively near-term system define it as a technology of extreme interest, enabling routine ultra-high speed interplanetary travel, and offering a first-generation capability for interstellar flights.

Pioneer’s work on Ultralight Solar Sails was supported by funding from the NASA Institute for Advanced Concepts (NIAC.)

TOPICS
Space Technology, Spacecraft Systems, Mechanical Engineering, Vehicles