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Summary of the NASA Innovative Advanced Concepts study of new concepts for returning a sample from the surface of Venus using a high-temperature airplane and a balloon-borne platform carrying the return launch vehicle along with propellant-manufacturing capability.
In-situ produced propellants have been identified in many architecture studies as key to implementing feasible chemical propulsion missions to destinations beyond lunar orbit. Some of the more noteworthy ones include: launching from Mars to return to Earth (either direct from the surface, or via an orbital rendezvous); using the Earth-Moon Lagrange point as a place to refuel Mars transfer stages with Lunar surface produced propellants; and using Mars Moon Phobos as a place to produce propellants for descent and ascent stages bound for the Mars surface. However successful implementation of these strategies require an ability to successfully transfer propellants from the in-situ production equipment into the propellant tankage of the rocket stage used to move to the desired location. In many circumstances the most desirable location for this transfer to occur is in the low-gravity environment of space. In support of low earth orbit propellant depot concepts, extensive studies have been conducted on transferring propellants in-space. Most of these propellant transfer techniques will be applicable to low gravity operations in other locations. Even ground-based transfer operations on the Moon, Mars, and especially Phobos could benefit from the propellant conserving techniques used for depot refueling. This paper will review the literature of in-situ propellants and refueling to: assess the performance benefits of the use in-situ propellants for mission concepts; review the parallels with propellant depot efforts; assess the progress of the techniques required; and provide recommendations for future research.
The uses of in-situ derived propellants are discussed for lunar and Mars transportation systems and the impact on mission architectures. Transportation systems include surface rovers, ascent/descent vehicles and space transfer vehicles at both the Moon and Mars. A brief survey of available propellant types is included. A brief description of typical transportation systems proposed for use on the Moon and Mars is presented. Comparisons are made of overall mission parameters such as mass, costs, schedules and performance assuming the use of in-situ derived propellants versus Earth supplied propellants using 4 different mission scenarios to cover the range of missions being studied today.
Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.
Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.
Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for chemical rockets only a cryogenic propulsion system would provide high enough performance to be considered for a Mars ascent vehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situ propellants is highly attractive. This option would significantly reduce the overall mass of launch vehicles. Consequently, the cost of the mission would be greatly reduced because the number and size of the Earth launch vehicle(s) needed for the mission decrease. NASA/Johnson Space Center has initiated several concept studies of in-situ propellant production plants. Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane (CH4), ethylene (C2H4), and methanol (CH3OH). MSFC initiated a technology development program for a cryogenic propulsion system for the Mars human exploration mission in 1998. One part of this technology program is the effort described here: an evaluation of propellant injection concepts for a LOX/liquid methane Mars Ascent Engine (MAE) with an emphasis on light-weight, high efficiency, reliability, and thermal compatibility. In addition to the main objective, hot-fire tests of the subject injectors will be used to test other key technologies including light-weight combustion chamber materials and advanced ignition concepts. This state-of-the-art technology will then be applied to the development of a cryogenic propulsion system that will meet the requirements of the planned Mars sample return (MSR) mission. The current baseline propulsion system for the MSR mission uses a storable propellant combination [monomethyl hydrazine/mixed oxides of nitrogen-25. However, a mission option that incorporates in-situ propellant production and utilization for the ascent stage is being carefully considered as a subscale precursor to a future human mission to Mars.
Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for non-toxic chemical rockets only a cryogenic propulsion system would provide high enough performance to be considered for a Mars ascent vehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situ propellants is highly attractive. This option would significantly reduce the overall mass of the return vehicle. Consequently, the cost of the mission would be greatly reduced because the number and size of the Earth launch vehicle(s) needed for the mission decrease. NASA/Johnson Space Center has initiated several concept studies (2) of in-situ propellant production plants. Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane (CH4), ethylene (C2H4), and methanol (CH3OH). MSFC initiated a technology development program for a cryogenic propulsion system for the Mars human exploration mission in 1998. One part of this technology program is the effort described here: an evaluation of propellant injection concepts for a LOX/liquid methane Mars Ascent Engine (MAE) with an emphasis on light-weight, high efficiency, reliability, and thermal compatibility. In addition to the main objective, hot-fire tests of the subject injectors will be used to test other key technologies including light-weight combustion chamber materials and advanced ignition concepts. This state-of-the-art technology will then be applied to the development of a cryogenic propulsion system that will meet the requirements of the planned Mars sample return (MSR) mission. The current baseline propulsion system for the MSR mission uses a storable propellant combination [monomethyl hydrazine/mixed oxides of nitrogen-25(MMH/MON-25)]. However, a mission option that incorporates in-situ propellant production and utilization for the ascent stage is being carefully considered as a subscale precursor to a future human mission to Mars.
Both the Martian and lunar forms of implementation of the Mars Direct architecture are discussed. Candidate vehicle designs are presented and the means of performing the required in situ propellant production is explained. The in situ propellant process is also shown to present very high leverage for a Mars Rover Sample Return mission flown as a scaled down precursor version of the manned Mars Direct. Methods of coping with the radiation and zero gravity problems presented by a manned Mars mission are discussed. Prime objectives for surface exploration are outlined and the need for substantial surface mobility is made clear. Combustion powered vehicles utilizing the in situ produced methane/oxygen are proposed as a means for meeting the surface mobility requirement. Nuclear thermal rocket propulsion is suggested as a means to improve mission capability.
Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for chemical rockets only a cryogenic propulsion system would provide high enough performance to be considered for a Mars ascent vehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situ propellants is highly attractive. This option would significantly reduce the overall mass of launch vehicles. Consequently, the cost of the mission would be greatly reduced because the number and size of the Earth launch vehicle(s) needed for the mission would decrease. NASA/Johnson Space Center has initiated several concept studies of in-situ propellant production plants. Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane (CH4), ethylene (C2H4), and methanol (CH3OH). MSFC initiated a technology development program for a cryogenic propulsion system for the Mars human exploration mission in 1998. One part of this technology program is the effort described here: an evaluation of propellant injection concepts for a LOX/liquid methane Mars Ascent Engine (MAE) with an emphasis on light-weight, high efficiency, reliability, and thermal compatibility. In addition to the main objective, hot-fire tests of the subject injectors will be used to test other key technologies including light-weight combustion chamber materials and advanced ignition concepts. This paper will address the results of the liquid methane/LOX injector study conducted at MSFC. A total of four impinging injector configurations were tested under combustion conditions in a modular combustor test article (MCTA), equipped with optically accessible windows. A series of forty hot-fire tests, which covered a wide range of engine operating conditions with the chamber pressure varied from 320 to 510 and the mixture ratio from 1.5 to 3.5, were performed. The test matrix also included a variation in the combustion chamber length for the purpose of investigating its effects on the combustion performance and stability.
A small reusable 'hopper' vehicle, the Mars In-situ Propellants Rocket (MIPR), is proposed to fly autonomously on Mars, using in-situ propellant production to manufacture rocket propellant directly out of the Martian atmosphere. The MIPR explores the Martian surface under rocket power and can repeatedly takeoff and land, carrying a suite of science instruments over a range of hundreds of meters per hop. The flight demonstration will accomplish a range of technology objectives important to both unmanned probes and to future human missions, including: (1) demonstration of a sub-orbital Mars launch vehicle, (2) demonstration of a pressure-fed small propulsion system for Mars ascent vehicles, (3) demonstration of a lightweight space engine, and (4) use for the first time of propellants manufactured in-situ on another planetary body. In addition to these technology objectives, the MIPR vehicle can carry a science payload that will advance our understanding of the surface and atmosphere of Mars.
The use of propellants manufactured from indigenous space materials has the potential to significantly reduce the amount of mass required to be launched from the Earth's surface. The extent of the leverage, however, along with the cost for developing the infrastructure necessary to support such a process, is unclear. Many mission analyses have been performed that have attempted to quantify the potential benefits of in situ propellant utilization. Because the planning of future space missions includes many unknowns, the presentation of any single study on the use of in situ propellants is often met with critics' claims of the inaccuracy of assumptions or omission of infrastructure requirements. The results of many such mission analyses are presented in one format. Each summarized mission analysis used different assumptions and baseline mission scenarios. The conclusion from the studies is that the use of in situ produced propellants will provide significant reductions in Earth launch requirements. This result is consistent among all of the analyses regardless of the assumptions used to obtain the quantitative results. The determination of the best propellant combination and the amount of savings will become clearer and more apparent as the technology work progresses.
This paper describes an assessment of potential benefits offered through use of lunar/Mars in situ propellant production to support manned missions to the moon and Mars. Resources available at each location and processing options for their extraction are discussed. Key infrastructure and support systems needed to sustain the propellant production operation and to deliver the propellants from their point of origin to the point of application are defined. The analysis focuses on reductions of earth launched mass over the course of multiple missions to assess the potential savings offered. Initial results show the impact of requirements for sustaining in situ propellant production systems on benefits offered during steady-state operation.
Project Minerva is a low-cost manned Mars mission designed to deliver a crew of four to the Martian surface, using only two sets of two launches. Key concepts which make this mission realizable are the use of near-term technologies and in-situ propellant production, following the senario originally proposed by R. Zubrin of Martin Marietta. The first set of launches delivers two unmanned payloads into low earth orbit (LEO): one consists of an Earth Return Vehicle (ERV), a propellant production plant, and a set of robotic vehicles, and the second consists of the upper stage/trans-Mars injection (TMI) booster. In LEO, the two payloads are joined and inserted into a Mars transfer orbit. The landing on Mars is performed with the aid of multiple aerobraking maneuvers. On the Martian surface, the propellant production plant uses a Sabatier/electrolysis-type process to combine six tons of hydrogen brought from earth with carbon dioxide from the Martian atmosphere to produce 100 tons of liquid oxygen and methane, which are later used as the propellants for the rover expeditions and the manned return journey of the ERV. Once the in-situ propellant production is completed, approximately two years after the first set of launches, the manned portion of the mission leaves earth. This set of two launches is similar to that of the unmanned vehicles; the two payloads are the Manned Transfer Vehicle (MTV) and the upper stage/TMI booster. The MTV contains the manned rover and the habitat which houses the astronauts enroute to Mars and on the Martian surface. During the 180-day trip to Mars, artificial gravity is created by tethering the MTV to the TMI booster and inducing rotation. Upon arrival the MTV performs aerobraking maneuvers to land near the fully-fueled ERV, which will be used by the crew a year and a half later to return to earth. The mission entails moderate travel times with relatively low-energy conjunction-class trajectories and allows ample time for scientific exploration. This set of missions can be repeated every two years in order to continue exploration at a variety of sites and gradually establish the infrastructure for a permanent base on Mars.
The Sabatier Methanation Reactor technology is of vital importance to the success of the human and robotic exploration program. In order to achieve an affordable program, the logistics supply to support the mission must be minimized to the fullest extent possible. One area of potential reduction with high return on investment is the closure of life support loops, particularly oxygen and water. The Sabatier system accomplishes this by utilizing hydrogen and carbon dioxide, waste products from the life support system, to produce water and methane. The recovered water is then recycled back into the life support system to provide oxygen; while the methane can be used for propulsion, or can be broken down further to recover the hydrogen. This technology is applicable not only to transit phases of exploration, but surface habitats as well as in-situ propellant production. The Sabatier Reactor system has been developed for ground based demonstration experiments extensively over the past 30 years. Over the past three years, NASA has funded development of the Sabatier Carbon Dioxide Reduction Assembly (CRA) for use on the ISS. Currently this system is at TRL 5 and it is expected that the system will be flown on the ISS as a flight experiment, The purpose of the flight experiment is to integrate the Sabatier CRA into a synchronized system with the oxygen generation system and the carbon dioxide concentrator. The flight experiment will verify the integration of the different systems working together plus it will verify the capability of the system to operate, and effectively separate its products in a micro-gravity environment. Subsequent to design validation, the flight experiment can remain onboard the ISS providing valuable water to offset logistics re-supply requirements. Some of the challenges facing the development of the Sabatier system include handling vibration induced particulates, microgravity phase separation and containment of hazardous gases. Plans for adequately addressing these issues will be presented. The Sabatier carbon dioxide reduction process will greatly benefit any of the extended duration human exploration missions because of the tremendous savings of consumables realized. Any of these mission scenarios, be they transit or surface based, must consider closing the life support loops in order to make the mission achievable, let alone affordable. Carbon dioxide reduction technology will be necessary for future outpost habitats, and the technology needs to be proven viable in a space application. The Sabatier methanation reaction is also a desirable method for producing propellant from the Mars atmosphere. The common system could be designed to accept carbon dioxide from an indoor air revitalization loop concentrator, or from an outdoor atmosphere compressor. Carbon dioxide reduction validation is but one step in the spiral development of the in-situ propellant production system desired for future planetary exploration.
The Mars ISPP Precursor, “MIP”, was a flight experiment on the 2001 Mars Surveyor Lander that was designed to demonstrate In-situ Propellant Production-- ISPP-- for the first time on Mars. The experiment was designed to show that it was possible to produce oxygen from the carbon dioxide atmosphere, and to demonstrate the individual technology components that would go into a full-scale production plant. The aimpoint of the oxygen production was to showcase the possibility of producing oxygen for use as rocket fuel, ultimately as an enabling technology for a future human expedition to Mars. The MIP team built and qualified flight hardware for the experiment to fly on the Surveyor-2001 lander, but following the failure of the 1999 Mars Polar Lander spacecraft, the Surveyor-2001 lander mission was cancelled, and the MIP experiment was never flown. The experience in building and testing the hardware did show that the carbon dioxide electrolysis process was feasible, and led to incorporation of in-situ propellant production into the NASA reference plans for human Mars missions. Twenty years later, the technology is flying to Mars in the form of the MOXIE experiment on the Perseverance rover.
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The CO2 Acquisition Membrane (CAM) project will develop, test, and analyze membrane materials for separation and purification of carbon dioxide (CO2) from mixtures of gases, such as those found in the Martian atmosphere. The CAM technology will enable passive separation of these gases, allow energy efficient acquisition and purification of these important resources, and lay the foundation for future unmanned sample return and human space missions. The CAM membranes are targeted toward In Situ Resource Utilization (ISRU) applications, such as In Situ Propellant Production (ISPP) and In Situ Consumables Production (ISCP).
The CO2 Acquisition Membrane (CAM) project was performed to develop, test, and analyze thin film membrane materials for separation and purification of carbon dioxide (CO2) from mixtures of gases, such as those found in the Martian atmosphere. The membranes developed in this project are targeted toward In Situ Resource Utilization (ISRU) applications, such as In Situ Propellant Production (ISPP) and In Situ Consumables Production (ISCP). These membrane materials may be used in a variety of ISRU systems, for example as the atmospheric inlet filter for an ISPP process to enhance the concentration of CO2 for use as a reactant gas, to passively separate argon and nitrogen trace gases from CO2 for habitat pressurization, to provide a system for removal of CO2 from breathing gases in a closed environment, or within a process stream to selectively separate CO2 from other gaseous components. The membranes identified and developed for CAM were evaluated for use in candidate ISRU processes and other gas separation applications, and will help to lay the foundation for future unmanned sample return and human space missions. CAM is a cooperative project split among three institutions: Lockheed Martin Astronautics (LMA), the Colorado School of Mines (CSM), and Marshall Space Flight Center (MSFC).