2001mars society convention part 3

Mars Environment Simulator,Environmental And Aerospace Physiology Laboratory

Simon Fraser University

Andrew P. Blaber, Ph.D.[1999]

AbstractEstablished in September of 1997, the Aerospace Physiology Laboratory in the School of Kinesiology at Simon FraserUniversity (Burnaby, British Columbia, Canada) is equipped for a wide range of human physiological testing. It containsequipment for measurement of non-invasive blood pressure, electrocardiograms, breath-by-breath respiratory gasexchange, blood gases, and Doppler ultrasound blood flow. The laboratory is also equipped with a respiratory feedbackcontrol system, computer-controlled tilt table and cycle ergometer. The Aerospace Physiology Laboratory is integratedvia computer (audio, video, and data) with the existing Environmental Physiology Unit (EPU) at Simon Fraser University.The main feature of the EPU is: a dive / altitude chamber complex with an altitude capability of 33.5 km (110,000 ft,equivalent to Mars atmospheric pressure). The dive / altitude chamber has living quarters for four with life support andcommunications systems for eight. The integration of the Aerospace Physiology Laboratory and the EPU provides aunique facility for Mars related research. Mars hardware, extra-vehicular activity (EVA) and life support systems as wellas human physiology and performance can be studied in a controlled simulated Mars environment.

We have embarked on an ambitious program to build a state-of-the-art aerospace physiology laboratory and to reshapethe Environmental Physiology Unit to meet the demands of the next century. This combined Environmental andAerospace Physiology Laboratory is the only university research facility in Canada with the capability to researchphysiological issues associated with diving, aviation, and space environments. The facility extends Canadian sciencecapabilities into research related to astronaut health and life support, decompression sickness, and EVA life supporttechnology development. In addition to physiological research and testing, the Environmental and Aerospace PhysiologyLaboratory provides a world-class scientific and technical training facility for both academic and industrial partners.

IntroductionThe Environmental Physiology Unit (EPU) was installed in the School of Kinesiology at Simon Fraser University in1981. The main features of the EPU (Figure 1), are a dive / altitude chamber complex (a life support and environmentalcontrol system with an original operating range of 305 meters sea water dive depth to 12,000 meters altitude. We haveupgraded the altitude capabilities to 33,530 meters.), a climatic chamber capable of simulating temperatures of -30ºC to50ºC, and hot and cold immersion tanks ranging from 5ºC to 50ºC. Now in its fifteenth year of operation, the EPU isundergoing a technological overhaul to upgrade the facility to allow it to meet the demands of innovative research in thenew millennium.

The author’s Aerospace Physiology Laboratory contains equipment for measurement of non-invasive blood pressure (BP),electrocardiograms (ECG), breath-by-breath respiratory gas exchange, blood gases, and Doppler ultrasound blood flow.The laboratory also contains a respiratory feedback control system, computer-controlled tilt table and cycle ergometer.

These devices are fully integrated to function within the climate and dive / altitude chambers in the EPU. TheEnvironmental and Aerospace Physiology Laboratory allows for innovative research related to human physiologicalresponses and adaptations to both terrestrial (including aquatic) and space environments

This “Environmental and Aerospace Physiology Laboratory” at SFU provides a unique research and teaching facility for thestudy of human physiology and performance (such as the effects of diving, altitude, temperature, humidity and environmentalgases) in extreme environments. I will focus on the altitude capabilities and research possibilities of this facility.

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Andrew P. Blaber, Ph.D.; Assistant Professor, Co-director of the Environmental Physiology Unit. School of Kinesiology, Faculty of AppliedSciences, Simon Fraser University, Burnaby, B.C. V5A 1S6; Tel: (604) 291-3276; Fax (604) 291-3040; email: [email protected]

The Facility

The Environmental Physiology Unit (EPU) at Simon Fraser University’s School of Kinesiology was installed in 1981.The main features of the EPU (Figure 1) are a dive / altitude chamber complex, a climatic chamber, and hot and coldimmersion tanks.

Dive / Altitude ChamberThe central feature of the EPU is the combinedhyper / hypobaric system. The chamber complexconstructed to PVHO-1 (Pressure Vessel for HumanOccupancy-1) standards was designed, fabricatedand installed (Figure 2) by Perry OceanEngineering of Florida. The main features areoutlined in Table 1. It consists of threeinterconnected chambers: entry lock, wet chamberand living chamber. The wet chamber is situatedbelow the entrance lock and connected by a 0.75metre diameter trunk. The design incorporates bothinternal and external doors. This accommodatesseparate pressurization of each chamber, andevacuation of either the living chamber or thecomplete complex.

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Mars Environment Simulator, Environmental And Aerospace Physiology Laboratory, Simon Fraser University

Figure 1. Diagram of Environmental Physiology Unit

Figure 2. Installation of dive / altitude chamber

We have expanded the altitude capabilities of the chamber so that we can achieve an atmospheric pressure equivalent tothat on Mars (~5 mm Hg: 33,530 meters or ~110,000 ft Earth altitude).

The main living chamber is outfitted to support a maximumof eight persons. It includes a 30 cm diameter medical lock,communications, fire detection and suppression system,four fold-up bunks, a table, and demand line-breathingmasks. Two independent breathing lines (BIBS) provideemergency and treatment gases: one line to supply specialmixtures and one dedicated to oxygen. The oxygen isexhausted from the masks to an external dump. High-pressure air is supplied to the system by means of two 17cfm Bauer compressors. Air storage is provided by four 143m3 (5,060 cubic feet) cylinders at 26,700 kPa (3,000 psi).This allows air pressurization of the chamber system to amaximum of 305 meters for wet or dry equipment testing.During human occupancy one storage cylinder is dedicatedto driving the fire suppression system. Sprinklers arecontrolled by a tracking pressure regulator and cover theentrance lock and main chambers with a deluge of 6.8 L·s-1

from a high-pressure reservoir.

Control of all three chambers is provided from a centralconsole with voice and television monitoring. The consoleprovides independent system control to each chamber. Incase of a line malfunction, the chamber controls have cross-connections to allow isolation of any component. Allthrough-hull penetrations have internal and external shut-offvalves. Atmospheric monitoring of each chamber isprovided at the control console. An Environmental ControlSystem developed by the Nova Scotia Research Foundation

controls the ambient conditions within the chambers. The system consists of two loops, serving the wet and drychambers respectively. These control ambient air purity and temperature in all units of the chamber complex and alsoprovide temperature control and filtration of water in the wet chamber.

A custom computer controlled hydraulic breathing machine is also available. This device is capable of simulatinghuman ventilatory function over a wide range of pressures.11 A range of tidal volumes, respiratory frequencies andgases can be programmed into the device to test a wide variety of commercial and experimental breathing apparatus.

For hardware testing, the bunks and other non-essential items can be removed from the chamber to provide a largervolume for the test hardware. Single components must be less than 0.76 m (30”) in diameter and assembled systemsmust be less than 2.0 m (78”) in diameter (See Table 1 for specs). Through the treatment gas console various gasmixtures can be introduced into the chamber and continuously monitored, allowing for the simulation of both Marspressure and atmospheric gas conditions.

Aerospace Physiology LaboratoryThe author’s Aerospace Physiology Laboratory is situated across the hall on the same level as the EPU, in the Schoolof Kinesiology. The researchers in the laboratory are experts in the field of G-physiology including the effects of spaceflight on cardiovascular control and the development of orthostatic intolerance in astronauts. This facility is equippedto allow for a wide range of human physiological monitoring and testing. The following are available:

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Transcranial Doppler (TCD) Ultrasound: TCD provides mean flow velocity (MFV) of the red blood cells. Thistechnique has been used extensively in studies of astronauts, bed rest subjects, and healthy or patient populations tomeasure MFV in the middle cerebral artery. The MultiFlow Doppler unit (DWL Elektronische Systeme GmbH,Germany) has the ability to collect two ultrasound signals simultaneously with analog data input and output ports formulti-signal data collection (A similar device was used on Neurolab, STS-90).

Non-invasive Arterial Blood Pressure: The NIBP 7000™ (Colin Medical, San Antonio, TX) using the oscillometricmethod (Semiconductor pressure sensor over the radial artery) can be used to collect continuous non-invasive bloodpressure and provide beat-to-beat estimates of arterial blood pressure. If required, gravitational correction can beapplied for estimates of heart or brain level arterial pressure. As well, the modular construction of this device makes itideal for blood pressure monitoring in the hyper / hypobaric chamber (see integration below).

Heart-rate monitor: The ECG wave form can be recorded from the analog output of a LifePak-8 Cardiac Monitor (Physio-Control,Redmond WA). This can be used to determine heart rates and RR-intervals (commonly used in cardiovascular research).

Ventilation and Gas Exchange: A breath-by-breath gas analysis system has been established by Dr. R. L. Hughson(University of Waterloo) for precise measurement of ventilation and gas exchange and has been the basis fordevelopment of software by Marquette Electronics Inc. (Milwaukee, WI) as part of the current GASMAP project forNASA. This software is being used in this research laboratory. The small, versatile RAMS M-100 Laboratory GasAnalyzer is used because of its size and compatibility with existing hardware and software. This system is also used byNASA for Mir, the Space Shuttles, and will be used on the International Space Station. This allows for smooth transferof experimental design from ground-based to space-based projects.

Respiratory feedback control system: Gas mixtures will be regulated and monitored using a computerized gas mixingsystem.14 Respiratory gases will be monitored from the RAMS unit and air of various concentrations of CO2, O2 andN2 can be mechanically mixed to produce specified gas concentrations. In the situation where expiratory values arebeing regulated by inspiratory gas mixtures expired gas concentrations will be monitored and inspired gas mixturesmodified using a computer algorithm.8

Blood Gas and Haemoglobin analysis: Blood gas and hemoglobin analyzers (AVL Scientific Corporation, Roswell, GA)with microsamplers are available. Research projects involving an integration of cerebral, cardiovascular and respiratoryphysiology often require the ability to obtain reliable blood gas and haemoglobin content values.

Computer-controlled cycle ergometer: The breath-by-breath system can also control the work rate on the cycleergometer. Many research protocols for testing human exercise capacity and performance require computer control ofwork rate on a cycle ergometer. This is essential for any program involving human cardiorespiratory assessment.

Computer-controlled tilt table: A major component of the research by the author involves the investigation andassessment of orthostatic intolerance. Clinical tests of orthostatic intolerance involve the use of a tilt table. The researchin this lab will involve investigations of transitions between head-down and head-up tilt (negative and positive “G,” withthe direction of gravity to or away from the head). The time spent in each position and the rate of transition may haveimportant implications for orthostatic intolerance. A custom, computer-controlled, tilt table that can be used in the laband in the altitude chamber has been designed to obtain the necessary range and rate of motion; ~90° (head down tilt)to +90° (head up tilt) (±70° in the chamber) at 45°s-1 maximum rotation.

Lower body negative pressure (LBNP): LBNP is also used as a cardiovascular challenge and has been used extensivelyto determine the effects of cardiovascular deconditioning (seen with space flight). Tilt tests are not always conduciveto many of the measurements that are needed to test research hypotheses. LBNP can be applied at low levels toprimarily stimulate the cardiopulmonary baroreceptors or at higher levels to also include the arterial baroreceptors. Aswell LBNP can be applied in conjunction with tilt to increase orthostatic loading in the head-up tilt position.

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Data collection apparatus: The analog signals from these devices are recorded simultaneously on a PC using aMetrabyte™ compatible A/D board and a sixteen-channel computer strip chart recorder (RUN Technologies, LagunaHills, CA). Beat-by-beat analysis of these data is performed off-line. All analog signals are transmitted to the A/Dboard via standard BNC connector cables and may be connected to any device that is compatible.

Integration: Environmental & Aerospace PhysiologyThe Aerospace Physiology Laboratory is integrated with the Environmental Physiology Unit. This involves completeintegration of the research devices in the EPU for on-line remote observation with network / internet access. All of thecomputers and the TCD are on an internal 100 base T network using a SUN UltraSPARC system as the network server.This is connected to the University network (currently at 10 base T). Large data files can be transferred from one deviceto another with great ease and speed. The SUN UltraSPARC also serves as a station for multi-signal data analysis, withsoftware for converting Doppler audio signals into velocity profiles used in beat-by-beat analysis. All of the equipmentcan be moved easily between the Aerospace Physiology Laboratory and EPU. Sufficient network ports exist to maintainnetwork connections in either or both rooms.

At present, the majority of monitoring devices cannot withstand exposure to the temperatures and pressures that mayexist in the EPU. Furthermore, for safety reasons, only low current, low voltage devices may be used in the dive /altitude chamber. Monitoring devices, and computer hardware necessary for conducting research in the EPU, have tobe outside with their sensors (e.g., Doppler ultrasound probes, blood pressure sensors) located inside. The dive / altitudechamber has been fitted with specialized data access ports (Table 1) in both the entry lock and the main chamber to linkoutside data collection devices with their respective sensors inside the hyper / hypobaric chamber. Along withspecialized connections for specific equipment residing in the facility, data ports contain wiring with internal andexternal BNC connectors for generic use. Some of these are being used to connect ECG and EEG electrodes to theirrespective monitoring devices.

A computer keyboard, flat screen monitor and pointing device (mouse) are being modified or protected to withstand theenvironments within the chamber; the remainder of the computer will remain outside the chamber. This will allowresearchers inside the facility access to data display as well as full network access. These can also be used in studiesinvolving human computer interactions.

At present we are working with Stephen Braham, also at Simon Fraser University (member M.A.R.S. experiment,Haughton Crater, Devon Island, Canada) to have the research devices in the facility accessible for on-line remoteobservation and interaction with network / internet access. The facility has video monitors and voice communicationdevices that will be integrated into the computer network system so that two-way communication with video, voice anddata will be possible via the Internet, as well as the usual e-mail and text communication. In long duration studies (e.g.,involving extended stays in the chamber6) the computer components will allow subjects greater access to and from theoutside world.

The School of KinesiologyThe School of Kinesiology is a diverse department with research in environmental physiology (thermal, altitude, anddiving) whole body exercise physiology, biomechanics, ergonomics, motor control behavioral neuroscience, and cellular/ biochemical mechanisms of disease (neural, heart and diabetes).

My interests cover the full range of environmental and aerospace physiology. I have interests in cardiovascular andcerebrovascular modeling1,2,4,5 with specific interest in orthostatic intolerance3 and space flight deconditioning.7 Mylab is currently investigating orthostatic cerebrovascular dysautoregulation (OCD). This condition may cause syncopeduring orthostatic stress. Persons with this condition have decreases in cerebral blood flow with orthostatic stresswithout apparent decrease in systemic blood pressure. This condition is thought to be due to an inappropriate cerebralblood flow autoregulation response to orthostatic stress.

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Figure 3. Altitude / dive chamber at SFU

Several other faculty members work in areas related to this facility. These areas include: the interface of human andmechanical systems in areas such as underwater work and industrial ergonomics12,13 (Dr. J Morrison); human motorcontrol, grasping and remote manipulation in human-computer interaction (Dr. C MacKenzie); the effects of prolongedexposure to altitude on brain stem function using EEG (Dr. H Weinberg); the relationship and markers of geneticabnormalities in heart proteins and cardiovascular responses to tilting and environmental stress (Dr. E Accili); and, nitrogengas kinetics during compression and decompression, specifically investigating nitrogen absorption, transport, saturationand elimination in real time9,10 (G Morariu, Adjunct Professor and Senior Research Engineer, Aerospace PhysiologyLaboratory; Dr. M Lepawsky, Adjunct Professor SFU, and Head, Hyperbaric Medicine, Vancouver General Hospital).

More information on the School of Kinesiology and its faculty can be found at “http://fas.sfu.ca/kin/”.

SummaryThis laboratory is the first Canadian University research facility that allows for a full range of aerospace physiologytesting. Not only does the integration of the Aerospace Physiology Laboratory and the Environmental Physiology Unitat SFU provide a unique research and teaching facility for the study of human physiology and performance, but it alsoprovides a state-of-the-art scientific and technical training facility for both academic and industrial partners.

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The main component is the dive / altitude chamber (Figure 3). This chamber is unique in Canada and most probablyone of a few chambers world wide with both diving and altitude capabilities. In the areas of space related research weare able to investigate the physiological effects related to astronaut EVA (30,000-ft altitude pressure, pure oxygen),including decompression sickness research, and human-machine interface. As well we can assess hardware under Marsconditions including Mars EVA equipment.

As a completely enclosed environmental system the hypo / hyperbaric chamber complex also provides a facility wherereal simulations of a Mars Habitat can be run. The participants are isolated and communicate via an audio, video anddata link (over which realistic time delays and signal problems can be simulated). Through computer control of theonboard systems, various scenarios such as air pressure leaks and EVA’s can be performed with continuous monitoringof hardware and astronaut physiology.

Current research projects include: the investigation of the effects of altitude and orthostatic cerebrovasculardysautoregulation on orthostatic intolerance (BCHRF); human eye tear film bubble nitrogen kinetics in hyperbaricenvironments (BCHRF); and effect of chronic elevations in environmental CO2 on cerebral autoregulation (NSERC /CSA). We also provide, on a service contract basis High Altitude Indoctrination (Physiologic Training) for private andprofessional pilots, and flight training schools.

These are only a few of the activities that are possible or presently underway in this facility. Any person, group, orcompany interested in using the facility is asked to contact the author.

Abbreviations:BCHRF – British Columbia Health Research FoundationNSERC – Natural Science and Engineering Research Council (Canada)CSA – Canadian Space Agency.

References1. Blaber AP, Bondar R, Stein F, Dunphy PT, Moradshahi P, Kassam M and Freeman R (1997) Complexity of middle cerebral artery blood flow

velocity: effects of tilt and autonomic failure. Am.J.Physiol. 273:H2209-H2216.2. Blaber AP, Bondar RL, Stein F, Dunphy PT, Moradshahi P, Kassam M and Freeman R (1997) Transfer function analysis of cerebral

autoregulation dynamics in autonomic failure patients. Stroke 28(9): 1686-1692.3. Blaber AP, Bondar RL, Moradshahi P, Dunphy PT, Serrador JM, and Hughson RL (1997) Inspiratory CO2 increases time to presyncope during

repeated 90 head-up tilt. Aviat. Space Environ. Med., 68:A266.4. Blaber AP, Bondar RL, and Freeman R (1996) Coarse grained spectral analysis of HR and BP variability in patients with autonomic failure.

Am.J.Physiol. 271:H1555-H1564.5. Blaber AP and Hughson RL (1996) Cardiorespiratory interactions during fixed pace resistive breathing. J.Appl.Physiol. 80:1618-1626.6. Goldberg SV, Schoene RB, Haynor D, Trimble B, Swenson ER, Morrison JB, Banister EJ. (1992) Brain tissue pH and ventilatory

acclimatization to high altitude. J. Appl. Physiol. 72:58-637. Hughson RL, Yamamoto Y, Blaber AP, Maillet A, Fortrat JO, Pavy-LeTroan A, Marine JF, Güell A, and Gharib C (1994) Effect of 28 day

continuous head down bedrest with countermeasures on heart rate variability during LBNP. Aviat. Space Environ. Med., 65:293-300.8. Modarreszadeh M, Kump KS, Chizeck JH, Hudgel DW, & Bruce EG: (1993) Adaptive buffering of breath-by-breath variations of end-tidal

CO2 of humans. J.Appl. Physiol. 75:2003-2012.9. Morariu GI, Strath RA, Lepawsky M, Dobrescu RF (1998) A quantitative study of post-decompression tear film bubble formation. Proceedings

of the European Underwater Biological Society. August 1998, Stockholm Sweden, Proceedings XXIVth Ann. Meeting EUBS: 212-215.10. Morariu GI, Strath RA, Lepawsky M, Longely CR. (1996) Exercise induced post - decompression ocular bubble development. Proceedings of

the International Joint Meeting on Hyperbaric & Underwater Medicine. Milano, Italy. Pp 509 - 512. 4 - 8.11. Morariu GI (1992) A volumetric-pump type respiratory simulator. Proceedings at the 5th International Conference on Environmental

Ergonomics. Nov. 2-7, Maastricht, Netherlands. Pp. 196-197.12. Morrison JB,. Taylor NAS and Voogt SL. (1992). The effect of hydrostatic imbalance on respiratory mechanics of the diver. In: Lung

physiology and diver’s breathing apparatus. Ed. V. Flook, A. Brubakk, Aberdeen University Press, Aberdeen, U.K. p 101-124.13. Morrison JB, Taylor NAS. (1990). Measurement of static and dynamic pulmonary work during pressure breathing. Undersea Biomed. Res.:

17(5) 453-467.14. Robbins PA, Micco AJ, Swanson GD, & Schubert WP. (1982) A fast gas mixing system for breath-to-breath respiratory control studies. J.

Appl. Physiol. 52: 1358-1362.

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Dry Reforming: A Unique Flowsheet for Fuel Production on Mars

Brian M. Frankie[2001]

ABSTRACTA new conceptual flowsheet is presented for Martian in situ fuel production. The dry reforming flowsheet incorporatesthe well-known Sabatier-Electrolysis process with a carbon dioxide / methane reforming step to consume some of theSabatier methane. By varying the ratio of effluent to reformed methane, any desired methane / oxygen ratio can beproduced by the dry reforming process. Such a machine will enable utilization of all imported hydrogen into an optimalmethane / oxygen fuel mixture, with copious quantities of surplus oxygen produced for crew consumables.

The reforming process is highly endothermic and requires temperatures above 650 centigrade on precious metalcatalysts. Appropriate feed / effluent heat exchange reduces the reformer power requirements, but an increased oxygen/ methane ratio increases the power requirements. In addition, the complexity introduced by the reformer and itsinteractions with the Sabatier system make the system relatively difficult to automate or control remotely. The energyusage and complexity imply that a dry reforming process will not be useful in the early stages of Mars exploration.However, the increased material usage efficiency and oxygen generation capability of the dry reforming technique willmake it an attractive technology to consider for second generation ISRU systems. In addition, the potential ease ofretrofitting Sabatier / Electrolysis units with a dry reformer provide an important advantage for early adoption of thetechnology. Minor preinvestment in the Sabatier system – essentially just provision for interconnections – will allowthe addition of a reformer, thus extending the useful lifetime of the Sabatier system, instead of replacing early Sabatiersystems with entirely new second generation systems. Thus, dry reforming will be an important technology to allowcost effective expansion of early Martian exploration and base building efforts.

IntroductionThe concept of in situ resource utilization (ISRU) has been advocated as a means to reduce the amount of mass launchedfrom Earth for Mars missions. Primary resources on Mars include the components of the atmosphere, especially themajority constituent, carbon dioxide. Primary near term utility for this resource includes oxygen for life support androcket fuel oxidizer, and carbon for hydrocarbon or oxygenated hydrocarbon fuels.

The advantages of an ISRU application are often described by its “mass leverage.” The mass leverage is the ratio of themass of usable product produced on Mars to the mass of required feedstock and equipment that needs to be flown toMars. For example, if 5 kg of equipment and feedstock can produce 100 kg of Mars product, then the mass leveragewould be 20. First order calculations of the mass leverage can ratio the product to hydrogen feedstock masses only(hydrogen is typically the only feedstock brought from Earth), but more detailed calculations require including allknock-on masses, such as required power systems, refrigeration and storage units, chemical reactors, etc. In turn, themass required for the complete ISRU system needs to be compared to completed designs of alternative systems thatwould accomplish the same mission. Such alternative systems should include a “traditional” type mission architecture(i.e., a mission using storable propellants and tankage launched from Earth), as well as suitable alternative ISRUconcepts. The comparison also needs to account for various secondary properties of the ISRU products, such as specificimpulse or energy density.

The SEDR concept is closely related to the Sabatier / Electrolysis (S/E) ISRU concept, but extends the concept to allowproduction of any desired fuel / oxygen ratio. This is done by introducing a carbon dioxide reformer, or “dry” reformer,which reforms methane produced by the Sabatier reactor into carbon monoxide and hydrogen. Carbon monoxide canbe vented while the hydrogen is recovered to the Sabatier system, thus potentially creating a closed hydrogen loop.Since hydrogen is the feedstock material brought from Earth, this concept allows a significant improvement of the massleverage of the S/E ISRU system, from a value of 12, to a value of 20 or more. In addition, the SEDR provides a

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responsive and flexible product distribution for variable crew demands. This improvement can be accomplished withnegligible preinvestment in the S/E system, and only modest marginal cost for the dry reforming system.

The SEDR Flowsheet – Process DescriptionThe SEDR flowsheet is shown in Figure 1. In this flowsheet, carbon dioxide is gathered by a freezer or other suitablecarbon dioxide acquisition system, and pumped at moderate pressure to the Sabatier reactor. The reforming reaction isfavored by low pressure, and the Sabatier reaction can achieve high conversion at moderate pressure, so the entiresystem can operate at pressures on order of one bar, thus reducing the required performance and power requirement ofthe carbon dioxide collection system. In the Sabatier reaction, CO2 reacts with hydrogen from the fresh feed and recyclestreams to produce methane and water. Water is condensed from the reactor outlet and electrolyzed. Oxygen from theelectrolyzer is condensed and stored, while hydrogen is recycled to the Sabatier reactor feed for further production.

Methane from the Sabatier water condenser is also sent to liquefaction and storage. However, a slipstream from themethane product, and methane boil-off from storage (and any unreacted hydrogen from the Sabatier reactor), can be sentto the reformer with a portion of the carbon dioxide feed gas. Feed gas to the reformer passes through a feed / reformerexchanger to recover heat from the reactor and reduce the heating requirements of the highly endothermic reactor. Thefeed CO2 reacts with the methane over an appropriate catalyst to form carbon monoxide and hydrogen. Cooled effluentgas from the reformer has water removed and sent to electrolysis, and then is compressed and separated in a highlyselective polymer membrane. Permeable gases, including hydrogen, any produced water, and unreacted carbon dioxide,flow through the membrane and back to the Sabatier reactor, where they can be captured in useful products.Impermeable gases, primarily carbon monoxide, are vented from the system. Such polymer membranes are used inindustrial gas separations, and have been demonstrated in Mars ISRU systems built by Pioneer Astronautics.

The methane slipstream ratio to the reformer is a primary control variable for the system operation. When the slipstreamto the reformer is set to zero, the system is simply a standard Sabatier / Electrolysis system, producing a LOX / LCH4bipropellant combination in a 1:1 molar ratio, which gives a mass ratio of 12 (first order calculation based on H2 import).When the slipstream to the reformer is set to 0.5 of the methane flow, the system recovers half the hydrogen in methaneproduct, allowing a 2:1 molar ratio LOX / LCH4 bipropellant, or a mass ratio of 4, which is the ideal stoichiometricratio. This provides a net system mass leverage of 20. If all of the methane product from the Sabatier reactor is sent tothe reformer, oxygen will be the only net system product and there will be no net consumption of hydrogen, thus givinga infinite ideal mass leverage.

Carbon Dioxide / Methane ReformingThe SEDR flowsheet seems to offer some definite benefits, but how easy is it to reform methane? Does this processflowsheet have the possibility of turning into a practical process? Fortunately, the CO2 / methane reforming reaction iswell known and widely utilized in the direct iron reduction industry. The basic reforming reaction is:

CO2 (g) + CH4 (g) —> 2 CO (g) + 2 H2 (g) DH = + 59.1 kcal/mole CH4

(Note: all reaction DH’s at 298ºK, throughout paper)

Since the reaction does not involve water, it is known as dry reforming, and reactor design is simplified by keeping waterout to the extent possible. Nevertheless, with three of the four constituents of the water-gas shift reaction present, thesystem is susceptible to water formation via that reaction pathway. In addition, the reaction is highly endothermic –significantly more so than a standard steam reforming reaction – which means that elevated temperatures are desirablefor reasonable equilibrium conversions. However, the elevated temperatures open the possibility of various cokingreactions, most predominately the methane dehydrogenation, the Beggs, and the Boudouard reactions. The variouscompeting reactions make dry reforming a complex reactor design problem, and catalyst selection and operatingparameters become absolutely critical. A summary of the previously mentioned primary competing side reactionsincludes:

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Figure 1. The proposed SEDR flowsheet.

CO2 (g) + H2 (g) <—> CO (g) + H2O (g) DH = + 9.85 kcal/mole

CH4 (g) —> C (s) + 2 H2 (g) DH = + 17.9 kcal/mole CH4

2 CO (g) <—> CO2 (g) + C (s) DH = - 41.2 kcal/mole CO2

CO (g) + H2 (g) <—> C (s) + H2O (g) DH = -31.4 kcal/mole

Water formed from the gas shift will also promote the steam reforming reaction. This will be a secondary reaction sincewater content of the reactor is fairly low, and will also help convert feed methane to product. The equation for steammethane reforming is:

H2O (g) + CH4 (g) —> CO (g) + 3 H2 (g) DH = +49.3 kcal/mole CH4

It’s clear that the desired reaction system is very complex, which makes it difficult to design a reactor that achieves highand selective yields. Fortunately, we have a long history of Earthly industrial dry reforming practice and research todraw upon to help design a functional reactor, and to narrow the possible desirable operating regimes. In fact, severalvendors offer commercial dry reformers for direct iron reduction industrial application. One of the leading vendors forthis process is Midrex, a subsidiary of Kobe Steel Co., which produces dry reformers of up to 500 tons/day capacity.

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These reformers use 510 tubes, each 250 mm diameter x 8 m length, filled with sulfur-passivated nickel-basedheterogeneous catalyst. At tube temperatures in excess of 1,020ºC (1,293ºK), low pressure, and in the presence of thecatalyst, the feed gases are reformed into hydrogen and carbon monoxide. These reducing gases are then sent to theshaft furnace for the reduction of the iron ore. A picture of the entire direct reduced iron process is shown in Figure 2.

Figure 2. Direct Reduced Iron Process from the Midrex web site. The SEDR process can use a reformer based on thatused in the Midrex process. The Midrex reformer is direct fired, which is not appropriate for the Mars ISRU system,

but since the thermal energy in the ISRU reformer outlet is not required for iron reduction (as is the case in the Midrexprocess), this energy can be recaptured in the feed, allowing electrical heaters to supply the modest trim heating duty.

Nickel CatalystsThe use of nickel catalyst is widespread for methane reforming reactions in industry, where the relatively low cost ofthe catalyst is a prime concern. However, nickel is very susceptible to coking reactions at modest temperatures.Typically, a small amount of sulfur is allowed to poison the nickel catalysts, which reduces the coking reactionssubstantially, while not interfering significantly with the reforming reactions. Nickel is also active for the water-gas shiftreaction, so this reaction can be assumed to be at equilibrium. Midrex notes that there are two separate mechanisms forcarbon deposition – thermal cracking of hydrocarbons (methane dehydrogenation) and dissociation of adsorbed carbonmonoxide (the Boudouard and Beggs reactions). These two mechanisms have very different characteristics. Thermalcracking reactions are very endothermic and slow. They are favored by high temperature, long residence times of thebase hydrocarbons at the elevated temperature, and acidic catalyst supports, such as alumina. Coke from these reactionstends to form along the wall of the tubes in fired reformers, where the temperature is highest. In contrast, the carbonmonoxide deposition reactions are exothermic, tend to occur quickly in low temperature regions near the center of thetubes, and are inhibited by acidic catalyst supports. A chart of the equilibrium constants for the coking reactions overthe typical reactor operating temperature range, along with the constants for the reforming and water gas shift reactions,is shown in Figure 3.

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Figure 3. The equilibrium constants for reactions expected to impact the reformer design. Clearly this is a complexreaction system, with all six reaction equilibria of the same order at ~900ºK. The two different types of potential cokingreactions are also apparent, with the exothermic Beggs and Boudouard reactions quickly falling to insignificance above900ºK, while the endothermic dehydrogenation reaction competes with the desired reforming reactions. The water gas

shift is only a weak function of temperature, barely varying over the temperature range of interest.

The characteristics of the reactions give strategies for eliminating the formation of coke in the reactor more than can beaccomplished with simple sulfur passivation. To eliminate the dehydrogenation reaction, the catalyst should bemaintained active so that hydrocarbons are reformed quickly and the production of cracked products isthermodynamically impossible. To eliminate the formation of carbon monoxide deposition products, the reactortemperature should be maintained high enough to make the thermodynamics unfavorable whenever there is carbonmonoxide present.

In Midrex’ commercial reactors, elimination of carbon deposition is accomplished by two reaction stages in the tubes.The inlet of the reactor tubes heats up the feed gas very quickly over a low activity catalyst on a magnesia (basic)substrate. Thus, the region where carbon monoxide dissociation can occur is traversed very quickly over a catalyst thatis not active enough for these reactions, and where there is a very low partial pressure of CO, since the reactants have nothad time to reform. When the reactants reach a temperature high enough for hydrocarbon cracking to occur, the basicsubstrate inhibits this reaction, and the low activity nickel catalyst produces enough reforming reaction products quicklyenough that cracking reactions become thermodynamically unfavorable – i.e., the partial pressure of methane falls andhydrogen increases, inhibiting this reaction pathway. However, the low catalyst activity prevents the endothermicreforming reactions from occurring fast enough to cool the reactor back into the CO deposition temperature range.

The second stage of the Midrex bed uses a high activity nickel-on-alumina catalyst. This allows the reforming reactions toproceed arbitrarily close to equilibrium, but there isn’t enough reactant left at the start of this zone for the reactions to coolinto the CO deposition temperature range. In addition, the acidic substrate inhibits carbon deposition from CO dissociation.

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Midrex’ methodology, in effect, allows a near constant reaction rate through the length of the reactor tube, balancing thereforming reactions in a narrow temperature and reaction rate region between the exothermic CO deposition andendothermic cracking reactions. One of the disadvantages of this technique is that very fast heating is required in theinlet zone of the reactor tubes. Midrex accomplishes this with a direct fired heater, which is obviously impractical fora Mars apparatus. Solutions for a Mars reformer using nickel catalyst would be an enhanced heat transfer feed / effluentexchanger, such as the microchannel devices being developed at Pacific Northwest National Laboratory. However, the1000+ degree C temperature required in the reformer makes the design of a microchannel exchanger a difficult technicalchallenge. A rigorous R&D effort would be required to solve the material concerns and verify that coking is preventedusing the designed reaction profile over a nickel catalyst.

Precious Metal CatalystsIf the technical challenges of a high temperature nickel based reformer prove too formidable, an alternative solution fora Mars system would be to scrap the nickel catalyst entirely in favor of a more selective catalyst. Numerous academiclaboratories have studied carbon dioxide / methane reforming reactions over noble metal catalysts such as rhodium,rhenium, iridium, ruthenium, and platinum. Many variants have been found to be very active and selective for thereforming reactions at modest temperature, while resisting any carbon deposition. For example, an iridium on aluminacatalyst allowed equilibrium reforming at 650 degrees C, without any carbon deposition. Iridium or ruthenium havealso been found to be active on europium oxide supports without coke formation. Rhodium on alumina has displayeda high activity without carbon formation at low methane / CO2 ratios. Platinum is also very active for the reformingreactions, but tends to coke readily. However, promotion of platinum with rhenium enhances the resistance to carbondeposition reactions without substantially impacting the reforming reactions, apparently by ensemble control of Pt sitegroupings that are required for carbon polymerization. In some respects, the Pt/Re combination resembles theperformance of sulfur passivated nickel, but, of course, it is active at a much lower temperature. Overall, rhodium andruthenium tend to display the highest selectivity for reforming reactions over carbon deposition reactions, although theyare also the most expensive catalysts.

The issue of expense is significant; the precious metal catalysts are much more expensive than the nickel catalyst.However, since relatively small volumes will be required for even a massive Mars exploration effort, and since hightransportation and development costs dominate the overall mission cost anyway, the catalyst cost may well be a secondaryconsideration. The lower temperatures from the precious metal catalysts will also benefit the Mars system, reducingmaintenance and increasing the lifetime of reformer tubes, electrical heaters, and instrumentation in the furnace.

Thus, the question of whether carbon dioxide / methane reforming is a practical solution for a Mars ISRU missionarchitecture seems to have a positive answer. It is clear that there are techniques on Earth that allow this reformingreaction to be used in industrial applications with standard nickel catalyst. In theory, these techniques can be adaptedto Mars ISRU applications. There are also experimental results that show promise for noble metal catalysts. A R&Deffort is required to develop an optimal reactor design for a Mars ISRU application. At this time, it does not appear thatany of the technical obstacles to a CO2 / methane reformer for Mars will be intractable.

Other Process Unit OperationsThe reformer provides the majority of the technical challenge in the dry reforming system. However, a couple of otherunit operations need to be mentioned.

MembraneA separation of the reformer products is required to allow produced carbon monoxide to exit the system. Numerousseparation technologies are available, but the one that is likely to be most suitable for a Mars ISRU is a polymericmembrane. In previous work done at Pioneer Astronautics and Kennedy Space Center, polymer membranes providedefficient separation of CO from more permeable gases including CO2 and hydrogen. The relative permeabilities (?) ofthese gases in commercially available polyimide membranes are order of 10 for CO2 / CO and better than 25 for H2 / CO.When the feed stream is compressed to a sufficient pressure to allow permeation, the desired species will preferentially

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pass through the membrane to the low pressure permeate side. The residual stream (retentate) will be almost entirelycarbon monoxide. Multiple membranes can be used in series to increase the purification of the permeate or retentate.

Note that methane typically also has a very low permeability in polymeric membranes, so any unreformed methane inthe reactor effluent will be likely to vent with the carbon monoxide. As methane is an extremely valuable product, thismakes it desirable to boost the reformer conversion as high as possible. A small amount of excess CO2 in the reformerfeed should effectively scavenge any methane in the reformer outlet, as well as reducing any tendency to coke.

Compression / ExpansionCompression work is required to increase the reformer effluent to a pressure that will allow permeation of desirableproducts at a pressure high enough to return to the Sabatier system. However, the residual stream from the membranewill still be at high pressure. The available energy in the membrane retentate can be recovered in an expansion turbineand used to power the compression of the reformer effluent. Since the membrane retentate is being released at Martianatmospheric pressure (approximately 10 mbar), there is plenty of expansion work available, which can provide all thepower necessary for the compression, with some excess generated power. This will reduce the amount of powerrequired for the system, at the cost of increasing the complexity and number of moving parts. In addition, the easiestway to achieve the combined compression / expansion is with a centrifugal machine, which may not be appropriate atlow ISRU flow rates.

The approximate power required to compress 1 kg/hr of reformer effluent is 69 Watts. To achieve the same power viaan expansion turbine, about 0.937 kg/hr of carbon monoxide would be expanded to an outlet pressure of 0.17 bar. Bothnumbers assume 75% polytropic efficiency.

Energy BalanceThe SEDR system is a relatively modest consumer of power. The major portion of power is consumed in the waterelectrolysis and product liquefaction steps. The dry reformer is an endothermic reactor and requires some heat input,but only consumes a fraction of the power of the Sabatier system. For a system that produces 1 kg/hr of stoichiometricLOX / LCH4 output (i.e., 4:1 mass ratio), the power required will be:

CO2 Acquisition:The CO2 acquisition system is difficult to estimate because there are many different techniques for acquiring CO2, which havewidely varying energy requirements. Nevertheless, assuming some sort of carbon dioxide freezer is used, approximately 350W will be used to collect 1.65 kg/hr of CO2, which will be used to manufacture 1.0 kg/hr of net LOX / LCH4 product.

Sabatier:The Sabatier will have some startup heating requirements, but since the Sabatier reaction is exothermic, there will benet thermal energy production at 670ºK from this unit. This thermal energy can be utilized elsewhere in the process, orcan be used to generate power via thermoelectric generators. The Sabatier reactor will produce a total of 1.13 kW perkg of net LOX / LCH4 product.

Electrolysis:Assuming 85% efficiency, the electrolyzer requires 6.58 kW electrical power per kg/hr of liquid oxygen production, orabout 5.26 kW per kg/hr of net LOX / LCH4 product.

Reformer:The reformer is an endothermic reactor that requires thermal energy at more than 920ºK. Assuming this is providedwith resistive heaters with near 100% efficiency, the reformer needs 1.14 kW electric power per kg/hr processed throughthe reformer. Since the reformer only needs to process 0.75 kg per kg of net LOX / LCH4 production at a 4:1 LOX /LCH4 mass ratio, only 0.85 kW electric power are required per kg/hr of net LOX / LCH4 product.

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Compression:As mentioned in the previous section, the compression power is about 69 Watts per kg/hr processed in the reformer. For thedesired mass ratio, this means about 52 Watts of shaft work per kg/hr of net LOX / LCH4 product. However, this amountof power will be recovered in the expansion turbine, leaving no net power consumption by the reformer compressor.

Refrigeration:A Stirling cycle refrigerator is used for both the LOX and LCH4 liquefaction, and the Stirling can achieve 20% of Carnotefficiency at the condensing temperatures of these two components. Assuming negligible heat leak, the power consumptionof the refrigeration system will be 0.95 kW of power per kg/hr LOX production and 1.30 kW of power per kg/hr of LCH4production. The overall net power consumption for one kg/hr of LOX / LCH4 in a 4:1 mass ratio is 1.02 kW.

The results of the energy balance are shown in Table 1.

Table 1. Power consumption of SEDR subsystems operating at 4:1 LOX / LCH4 mass ratio

Scaling of SEDR SystemThe entire SEDR system is highly scalable, by simply adjusting the volume of the reformer, the size of the flow passagesin the compressor, and the area of the membrane separator. The largest dry reformers currently in operation processmore than 500 tons/day. Considering that the first generation Mars ISRU systems being considered for robotic missionsare potentially going to generate on order of 1 kg production per day, it’s apparent that it will be a long time indeedbefore capacity constraints are a problem.

But is it possible to scale the system down? Large industrial units generally operate much less effectively at extremelysmall scales. However, the case is probably not as extreme as that suggested above for the robotic scale ISRU systems.It’s likely that the additional complexity of the SEDR system, relative to the S/E, will require human attention, at leaston occasion. Thus, the dry reformer will probably only be of use for the first ISRU systems that support a humanpresence on the Red Planet – second generation systems. The human support ISRU systems should produce on orderof 10’s of kg per hour, or 100 times the scale of the robotic precursor ISRU systems. This will be a perfectly suitablescale for the dry reformer system, particularly as one dry reformer might process the output of several Sabatier systems.

Retrofitting a Reformer in an Existing S/E SystemAs noted in the previous sections, the SEDR flowsheet is significantly more complex than a simple Sabatier /Electrolysis flowsheet, which is currently one of the leading candidates for a first generation Mars ISRU system.However, the SEDR overcomes many of the objections to the S/E system, most particularly the low mass leveragecaused by a less-than-optimum product stoichiometry, and the inability of the S/E to effectively produce oxygen. Thesefactors suggest that a working dry reformer is a likely candidate for an early second generation ISRU system.Retrofitting one or more S/E systems with a dry reformer offers an intriguing means to dramatically increase theflexibility and usefulness of the ISRU concept by adding lightweight modules as Mars efforts grow more ambitious. Ineffect, the SEDR can grow as the Mars program effort grows.

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The task of retrofitting a S/E unit with a dry reformer can accomplished with minimal preinvestment. Only five physicaltie ins are required:1) The CO2 feed from the acquisition unit to the reformer2) The methane slipstream from the S/E condenser to the reformer3) The methane overpressure from LCH4 storage to the reformer4) The recycle stream returning CO2 and H2 to the Sabatier unit5) Water from the reformer system condenser to the electrolyzer

The control systems for the Sabatier unit will require a small amount of prework to prepare for a reformer extensionunit. Extra connections on the S/E control system bus will allow the reformer to be tied in with minimal fuss. And thecontrol software should be written to allow operation of the fully integrated SEDR system from the start of operation.The software will still operate normally before the SEDR system is added, as this will simply be the special case of zeroflow to the methane slipstream.

A slight physical oversizing is desirable for the Sabatier / Electrolysis unit, relative to the base case withoutpreinvestment for the dry reformer. Such an oversizing of the electrolyzer and Sabatier reactor will allow the SEDRsystem to operate at the same overall mass throughput as the S/E system, despite the fact that the product slate isswitched to a higher LOX / LCH4 ratio. If the electrolyzer (in particular) is not oversized, then it will be the overallprocess bottleneck, and the addition of the dry reformer will merely decrease the methane production rate whilemaintaining constant LOX production. Note that the refrigerator does not have to be oversized, as the overall productionof the SEDR is approximately the same as the S/E system. However, some of the refrigeration duty will be switchedfrom the LCH4 to the LOX.

Process ExtensionsThe dry reformer is a logical extension to add to the base Sabatier / Electrolysis unit. However, it is not the final say inprocess extensions. The SEDR system allows extraction of oxygen from CO2, with venting of the resulting CO. ButCO will be a valuable product on Mars in the future. With an additional source of hydrogen, it can be used to produceFischer-Tropsch hydrocarbons or alcohols. It can also be used as a reducing agent for the direct reduction of iron. Thesetypes of units can be added to the SEDR system in a modular fashion as Mars base requirements and resources expand.These process additions will allow a SEDR system to be a flexible ISRU option far into the future.

The fundamental difference in the dry reformer is that it allows mission planners to start thinking of an integrated seriesof Mars missions, rather than a series of standalone Mars missions. This is enabled by thinking of ISRU units on theMartian surface as valuable, extendible assets, rather than as use-and-abandon hardware. With appropriate preplanning,these assets can be utilized far into the future, tremendously increasing the cost effectiveness of Mars exploration andbase building programs.

ConclusionThe SEDR provides a technology for application to second generation (first human support capacity) Martian ISRU fuelproduction systems based on the Sabatier reaction. The advantages of the dry reformer added to the Sabatier reactor are many,including increased mass leverage and the ability to vary the product slate on the fly to produce excess quantities of oxygen.

The dry reformer has a complex set of reactions occurring in it, including a troubling set of carbon deposition reactions.However, there is a well-proven heritage of reformer operation on Earth that provides an excellent basis for design ofthe reactor. Existing reactors utilize nickel based catalysts with numerous design features to reduce carbon deposition.These features include sulfur passivation, and division into different reaction zones to control the concentration ofreactants and the rate of reaction. In addition to the standard nickel-based industrial catalysts, a considerable amount ofacademic work has been performed on noble metal catalysts. Although too expensive for terrestrial use, a modest sizeMartian ISRU unit would be able to afford these catalysts, which are active for reforming at modest temperatures whileproviding excellent resistance to coking.

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Integration of the reformer into an existing Sabatier unit will be quite easy with minimal preinvestment. A number ofphysical tie ins and control system modifications will be required, but the total effort will be a small percentage of thecomplete Sabatier design. In addition, the reformer will prove to be easily scalable, and further process units can extendits usefulness.

Planning to incorporate a dry reformer into a Sabatier system should take place immediately. Provision for retrofittingthe Sabatier unit used in NASA’s Mars Design Reference Mission with a reformer would be a minor engineering effortand would provide significant returns. The engineering effort would involve revision of existing Sabatier unit drawingsand a change in the control software. A further R&D effort to determine the proper design of the reformer should beimplemented. With this modest effort, the usefulness of the Sabatier ISRU concept would be tremendously extended,and ISRU could finally start to realize its promise to open the Martian frontier.

Technical AppendixThermodynamic properties of reformer reactions.

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References1. Bissett, L. 1977. “Equilibrium Constants for Shift Reactions.” Chemical Engineering, 84 (23):1552. Dubois, J. -L., Sayama, K., Arakawa, H. 1992. “CO2 Hydrogenation over Carbide Catalysts.” Chem. Lett., 5-8.3. Kelley, Bruce. “Natural Gas and Reformer Catalyst.” 2000. 6 part document published by Midrex Direct Reduction Corporation, Charlotte,

NC, US. Available on the corporate website at http://www.midrex.com/info/resource_making.asp4. Kitayama, Y., Watanabe, Y., Muramatsu, K., Kodama, T. 1997. “Catalytic Reduction of Carbon Dioxide on Ni-Cu Alloys.” Energy, 22, 177-

182.5. Meyer, T. and C. McKay, “The Atmosphere of Mars - Resources for the Exploration and Settlement of Mars,” Proceedings of The Case for

Mars Conference, Boulder, CO 1981.6. Nagaoka, K., Seshan, K., Lercher, J. A., and Aika, K. 2001. “Mechanism of Carbon Deposit/Removal in Methane Dry Reforming on

Supported Metal Catalysts.” From: Iglesia, et. al., Proceedings of the Natural Gas Conversion VI Conference, June, 2001, Anchorage, AK,US, 129 – 134.

7. Nozaki, F., Sodesawa, T., Satoh, S., Kimura, K. 1987. “Hydrogenation of Carbon Dioxide into Light Hydrocarbons at Atmospheric Pressureover Rh/Nb2O5 or Cu/SiO2- Rh/Nb2O5 Catalyst.” J. Catal., 104, 339-346.

8. Pena, M. A., Gomez, J. P., and Fierro, J. L. G. 1996. “New Catalytic Routes for Syngas and Hydrogen Production.” App. Catal. A: General,144, 7 - 57.

9. Perera, J. S. H. Q., Couves, J. W., Sankar, G., and Thomas, J. M. Catal. Lett., 11 (1991), 219.10. Richardson, J. T., Jain-Kai Hung, and Zhao, J. “CO2-CH4 Reforming with Pt-Re/?-Al2O3 Catalysts.” From: Iglesia, et. al., Proceedings of

the Natural Gas Conversion VI Conference, June, 2001, Anchorage, AK, US, 203 - 208.11. Tonkovich, A. L. Y., Call, C. J., Jimenez, D. M., Wegeng, R. S., and Drost, M. K. 1996. “Microchannel Heat Exchangers for Chemical

Reactors.” AIChE Symposium Series, Heat Transfer, Houston, No. 310, 92, 119 – 125.12. Tonkovich, A. L. Y., Roberts, G. L., Call, C. J., Wegeng, R. S., and Wang, Y. 1999. “Active Microchannel Heat Exchanger.” International

Patent # WO 99/00186 assigned to Battelle Memorial Institute.13. R. Zubrin, B. Frankie, Tony Muscatello, and T. Kito-Borsa, “Progress in the Development of Mars in situ Propellant Production Systems,”

AIAA-99-0855, 37th AIAA Aerospace Sciences Meeting, Reno, NV, January 11 - 14, 1999.14. R. Zubrin, B. Frankie, and T. Kito, “Report on the Construction and Operation of a Mars in situ Propellant Production Unit Utilizing the

Reverse Water Gas Shift,” AIAA-98-3303, 34th AIAA/ASME Joint Propulsion Conference, Cleveland, OH, July 13 - 15, 1998.15. Zubrin, R., B. Frankie, and T. Kito. “Final Report for Mars Methanol in situ Propellant Production SBIR Phase 1 Study,” NASA Contract

Number NAS9-97082, Pioneer Astronautics, September 15, 1997.

The Case for Mars Funding – Making Exploration Pay

Thomas Andrew Olson; Paul Contursi; Beverly Woessner; Kevin Beary[1999]

AbstractThe estimated cost of a sustainable, long term human Mars exploration / settlement initiative is $50 billion. Raisingsuch sums in 10-30 years’ time is unlikely within the framework of “traditional” venture funding. Moreover, againstcompetitive, more experienced special interest groups, and in an unpredictable political climate, it is unlikely that wewill be able to win sufficient long-term government funding, either.

In response to this challenge, the authors propose establishing a new venture capital funding entity, offering 30-year“bond-shares” for sale at mass-market unit prices, in a public global appeal to individuals, corporations andgovernments. The fund would be a new class of investment vehicle, sharing characteristics with mutual funds, IPOs,junk bonds and limited partnerships. The low minimum investment would make the Fund “venture capital for themasses,” easily available to countless individuals and investment groups. The Fund’s managers would select a projectmanagement firm to be Prime Contractor for Mars missions, through which direct capital outlays would be made.

Capital not immediately allocated for Mars-specific purposes would be placed in more traditional short to medium-terminvestments. But the concept’s uniqueness comes from large corporations’ ability to invest in the fund by supplyinggoods and services required by the Fund’s Prime Contractor. The fund would issue a form of “bearer bond” to suchcorporate suppliers, redeemable by said contractor.

As a first step, capable of testing / proving the concept for both the engineering and the economic side, the authorsadditionally propose a “mini Mars Direct” project. This is an unmanned sample-return mission involving the ISPPconcept developed for the Mars Direct plan, funded via the funding / marketing mechanisms outlined below, at aprojected cost of $150-175 million. Publicity and financial returns from this project’s success would provide a positivetrack record and speed the capital funding process for manned missions.

Building upon the success of “Mini Mars Direct,” and meeting the initial capital goal of $10 billion, the Fund could thenmount the first human expeditions and begin reaping immediate returns. To the Fund would belong exclusively all data,samples, media coverage, and post-mission “tie-ins” – the marketing of which would more than cover the entire cost ofthe first human mission. Furthermore, any new mission hardware or infrastructure would become the joint property ofthe Contractor and the Fund.

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Thomas Andrew Olson, Cyberjox, Ltd. Consulting Services, email: [email protected] / Paul Contursi: [email protected] /Beverly Woessner, R.N.; [email protected] / Kevin Beary; [email protected]

In the spirit of the fundamental engineering paradigm shift represented by the “Mars Direct” concept, we wish to createa concomitant economic paradigm shift to open the Martian frontier.

IntroductionTo date, the Mars Society, in terms of its function and stated goals, is right on track. Like that of its model, The CousteauSociety, the Mars Society’s mission of raising people’s consciousness about the promise of Mars has been verysuccessful. To this end, soliciting corporate sponsorships is a worthy and necessary endeavor; selling informational orlogo-centric items at public events is also desirable – increasing membership rolls will be necessary to secure a corecadre of involved people to help handle colonization issues when it begins to grow to unmanageable size. My onlycriticism of the organization (if it can really be called a criticism) is that of thinking too small . . .

A long-term sustainable Mars exploration and colonization effort has been estimated by Zubrin et al. to cost around $50billion. We’re not going to raise that in any realistic time frame utilizing traditional fund-raising methods. In fact, thefunding structure this paper describes is undoubtedly beyond the scope of the Mars Society’s charter, and will be betterhandled by an independent organization (which I also propose to create). Nevertheless, I offer the concept here beforethe Mars Society’s membership, because to make it succeed will require the membership’s support and enthusiasm.

Pitfalls of “Traditional” Government Funding Methods“NASA’s obstacle is not a technology barrier – rather it is a barrier of financial abilities. Space activitiesrequire decades of planning. Short-term constraints of a political agenda do not address this necessity.”

— Sen. Conrad Burns (R-MT) member of the Commerce,Science and Transportation Committee, in “Roll Call”

To date, the only large-scale ($5+ Billion) funding idea proposed either within the Mars Society or elsewhere is what Icall the “Government Funding As Usual” model (GFAU – “guffaw”); namely, elbowing our way into the public troughand siphoning away tax dollars for Mars before they can be siphoned off by another special interest group. The onlypromise of “payback” offered for this generous public largesse would be the technological spin-offs promised, in duecourse, to filter their way back down into the economy, ideally creating new jobs and an improved tax base. The reputedlong term improvements to our standard of living may be debatable; but even granting that argument, if we use theApollo-moon experience as an historical perspective, this trickle-down approach can be such a convoluted and cloudedprocess that the average citizen often misses the connection between them.

This method of doing things was a great sell to Cold War America in the early 60’s, but after July 20th, 1969, when weofficially “won” the Space Race, the public, inundated daily with media reports of domestic unrest, political corruption andthe Vietnam body counts, rapidly lost interest. Funding was dropped for political expediency, and six more Americans whoshould have walked on the moon1 never got the chance, never mind mentioning the additional loss to science.

Today we have a generation growing up who has never known a Cold War, and we as a nation no longer perceive anyother single nation as being a genuine threat to our way of life. This effectively removes the “national security / nationalpride” impetus for public funding of manned interplanetary exploration and/or settlement.

Another downside of GFAU is that it is “non-inclusive”; GFAU perpetuates two very powerful myths: (1) the NASA-promoted myth of space as being the exclusive playground of a small cadre of highly trained techno-elitists withgovernment backing, and (2) the myth that only governments can “afford” to go to space. These are myths that I feelwe need to publicly discredit and rid ourselves of once and for all, if we are serious about doing space for real and trulymaking it available to all who wish to commit themselves to its exploration and economic development.

The biggest problem with GFAU, though, is that in seeking government funding we end up doing battle with all theother agenda-laden special interest groups who have a lot more experience at fighting those public trough battles thanwe do. Let’s face facts: most Mars Society members are professional people in the sciences, engineering, computing,

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or the arts, and are, effectively, “babes in the woods” when it comes to political wrangling. Compared to the entrenched,experienced social-agenda special interests, the pro-space activist constituency is relatively small, and to date, not veryvocal. All our opposition (and there is lots of it) need do to overwhelm our call for a renewal of a 60’s-era “NewFrontier” is to trot out statistics on falling U.S. education standards, photos of hungry kids and homeless people, and adecaying infrastructure. Combine this with a liberal dose of the word “deficit,” and it’s all over. The opposition getsthe funding . . . Mars does not. If events surrounding current space appropriations bills are any indication, the politicalwill to make Mars happen won’t be forthcoming either from this, or any future Congress in this generation. There aremany senators and congressmen who would NEVER appropriate a dime for a manned Mars mission, so long as there isa single hungry child (or potential swing vote) in their respective districts. To have even half a chance at the “D.C.game,” we would need to expend a significant amount of our valuable resources on lobbyists, PR firms, and otherconsultants – with no real predictable results to show for the expense.

“Special interests take care of the politicians who take care of them.

There’s one problem: most special interests thrive at the expense of taxpayers or the competitors thatthey’re protected from. Most special interests don’t want a level playing field. They want to fix thegame so they always win.”

— Jo Jorgensen

So What Do We Do Instead?As was true for Mars Direct for the engineering side, I contend that Mars funding also requires “a blank sheet of paper”;a different approach to effectively bypass the Washington, D.C. process, an approach in which those involved in makingMars happen are answerable directly and solely to investors, many of whom may themselves be directly involved in theeffort. The only need we may have for politicians and bureaucrats would be to create a regulatory framework allowingindependent, private, market-driven efforts to flourish.

For a privately held long-term investment fund to succeed, it must publicly state its purpose (other than overall profitsfor the investors, which is the goal of any investment vehicle). In this case, I think any investor prospectus writtenshould state the following goals:

• To provide investment / venture funding for the express purpose of establishing a permanent, and economically self-sustaining human community on Mars, developing and exporting Martian resources for Earth’s benefit,

• To invest in developing technologies that support those efforts, and• Provide careful management and due-diligence to ensure that said items provide the expected long-term return for

the Fund’s investors.

In any business funding proposal, the primary focus is generally on detailed business plans, organization, marketing,timelines / milestones, and, of course, the bottom-line (profit margins). These things DO matter greatly, but for now,for purposes of discussion, let’s put those issues aside temporarily and talk “high concept,” to borrow a Hollywoodexpression.

The biggest argument critics will have against what I’m about to propose is that “venture capitalists,” in their opinion,are concerned only with the above-mentioned economic issues. Well, yes and no. The people those critics are referringto are NOT really venture capitalists at all, but rather investment bankers, whose needs are far different, and whoseoutlook tends to be more conservative and near-term. Real venture capitalists are in it for the long haul, because theyknow that for something outrageously visionary to succeed, it’s going to take time, hard work, and a lot of guts. Andthey understand full well that in the venture-capital business, you win some and you lose some – but the wins are usuallypretty big, and make it all worth the occasional bust. The entire history of Silicon Valley was underwritten by the verysort of “financial cowboys” Mars colonization will require – a rare breed these days, but they’re still out there, everlooking for the next “insanely great” idea.

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By even suggesting that we go to Mars on our own, on a shoestring, and “living off the land” when we arrive, the MarsSociety has bitten off a great deal. Robert Zubrin’s The Case for Mars successfully challenged the old big-government“Battlestar Galactica” philosophy for reaching Mars. As stated previously, we now also need a fundamental shift inthinking on the economic side, providing a solid challenge to the GFAU model.

In an age of ever-increasing regulation and scrutiny in the financial markets, starting a venture capital fund remains arelatively straightforward procedure – the legal and regulatory processes are really not all that cumbersome orexpensive, as opposed to a bank or brokerage firm. Once the essential framework is in place, all that remains is toconvince people with dollars and dreams to invest in your particular formula.

And that is precisely what I suggest we do.

We craft and promote a unique, new funding concept, one with global appeal (as getting to Mars and staying there istruly a global effort). This concept, with proper marketing, offers what we think of as “the power of inclusion” for thegreatest number, worldwide...and here it is:

We organize a “populist” venture capital funding entity, “The Colony Fund,” selling “bond-shares,”at $1,000 per . . . a mass, global appeal for investment capital . . . at an affordable price . . . literally“Venture capital for the masses.”

To raise the necessary $50 billion for long-term Mars exploration leading to productive colonization (the real payback),we would need to sell 50 million bond-shares of this venture-cap-bond-fund over approximately the next decade, at$1000 per share. Sounds like a daunting prospect on the surface, if one still thinks traditionally. What I propose (toborrow from Apple Computer’s motto) is to “Think Different.”

We are not offering here just a lofty dream, but a stake in the future for our children and grandchildren. We’re offeringthe average citizen of the planet Earth an opportunity to have a personal stake in this adventure, something that has neverbeen offered before, and something that they and their posterity will literally profit from. This populist approach appealsto anyone with a serious interest in the future of space and/or our species. Granny might buy a couple of shares for hergrandkids to inherit. 10 people of more modest means could form an “investment club” and kick in $100 each for ashare. OR one Bill Gates, Steve Jobs, Ted Turner, Larry Ellison, or Donald Trump can buy – well, as many shares astheir hearts fancy. And if we do things right, they’ll fancy a lot.

Participation in the Colony Fund doesn’t necessarily have to involve cash, however. There may be other opportunitiesto leverage investor shares, specifically with entities such as large corporations.

Should a major aerospace firm, for example, wish to “invest” a $ billion worth of booster rocketry and logistics support. . . that could be their pathway to a million shares. The same would apply to any other technology company that wishesto provide some of what they do best directly, in exchange for a piece of the long-term “action.” The Fund would issue“corporate bearer-bonds” with a certain share-value, redeemable by the contracting company the Fund bankrolls (“theContractor”), at the supplier / shareholder’s in a sort of circular relationship. Once the bearer-bond is redeemed, thesupplier has met its in-kind obligation and is entitled to full value for its shares, like any pure cash investor.

By this method, governments around the world could also play a more appropriate role, by becoming equal venturepartners with private sector companies and individuals. Whether it’s the U.S., Europe, Japan, or the Sultan of Brunei,governments invest in market-based growth funds every day – some may want to invest and buy shares in this Fund,either in cash or in-kind. In the case of financially strapped Russia, “in kind” could be VERY helpful for both parties(remember Energia?). It would be a mass global funding appeal at all levels.

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This is the only funding concept offered to date that is as voluntary as one can make it, allows the entire global economyto participate, and even hints there will be a concrete, long-term payback. It’s like buying a 30-year bond. We want tomake SURE investors understand that they are committed for the long term. And, as in any other investment vehicle, thereare risks – but those risks are manageable. They’ve already been laid out and dealt with in detail in The Case for Mars.

This proposal can’t be pigeonholed into a specific slot. It’s not a pure bond fund, venture capital fund, mutual fund,limited partnership, IPO or Junk Bond, but it has elements of all those things. It will be riskier than, say a US SavingsBond, but not too much so, and marketed with the same sort of mass appeal. It is a “populist hybrid” investment vehicle.We’ll be asking the world to put their faith – and their dollars – into an incredible dream. To make it fly, we proactivelymarket it (“sell the sizzle”), get every major corporate and government player we can to come on board with us, and wekeep thumping the “Zubrin bible” and showing the world the hard numbers. We inundate the media, both traditional,and the “new media” of the Internet. New Media, properly done, is like a direct line into public consciousness, and ifanyone is capable of taking full advantage of it, it’s people like Mars Society members . . . technically savvy, energetic,and committed. It’s the one area where we truly have a distinct advantage over the forces of stasis.

Pathfinder’s web site got 800 million hits in 2 months! . . . you tap 1% of that at a $1000 per share, and you have yourfirst manned mission plus a nice hedge . . .

Bootstrapping the New ParadigmThe accumulation of $5-10 billion in investment capital for the first human missions within this new commercialframework will be a significant challenge. To lay the groundwork for human Mars exploration, an interim goal shouldbe chosen that will serve as a “proof of concept” in the broadest possible sense of the term. Ideally, the interim goalwould demonstrate the financial, technical and management feasibility of the enterprise in a dramatic and profitable way.It should be mounted for a relatively modest cost in a reasonable time frame and return some valuable scientific data.

The ideal candidate for such an interim goal would be an unmanned Mars sample return mission designed around theuse of In-Situ Propellant Production (ISPP) technology. R. Zubrin and S. Price of Lockheed Martin proposed such amission to NASA in March of 1995. This design is essentially a technical dress rehearsal for Dr. Zubrin’s Mars Directhuman mission architecture.

This “Mini Mars Direct” mission begins when a single booster is launched on a direct trajectory to Mars. The lander,rover and Earth Return Vehicle (ERV), containing a propellant plant, would have a total mass at liftoff of only about 540kilograms. This modest payload could be launched on a commercially available booster, such as the Boeing Delta 2.

After using a heat shield and a parachute to land on Mars, the rover is deployed to retrieve samples. Meanwhile, thechemical plant aboard the ERV pumps in carbon dioxide from the Martian atmosphere. The plant could producemethane / oxygen or carbon monoxide / oxygen propellant. Both types of systems have already been successfullydemonstrated in laboratory tests. By the time the launch window for the flight to Earth opens, the ERV’s tanks will befull. The ERV lifts off and flies directly back to Earth. After reentry, the ERV and its precious cargo of samples arerecovered via parachute.

As in the original Mars Direct mission architecture, it is the manufacture of propellant on the Martian surface that keepsthe scale (and cost) of Mini Mars Direct manageable. For example, since the ERV will be launched from Earth with itstanks essentially empty, it will only weigh about 70 kilograms on the pad. Only one launch is required and there is noneed for complex rendezvous and docking in Mars orbit. The original cost estimate for the mission was $302 million.However, in light of the same assumptions used in our human mission estimates, the actual cost might be reduced to aslittle as 50-60% of that figure.

The lead time required for the development of the spacecraft would be relatively short, since almost all the systemsrequired, with the exception of a flight-ready ISPP system, are already available. Using the new NASA “Faster, Better,

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Cheaper” management strategy as a guide, it should be possible to launch this mission about three years afterdevelopment begins.

From the perspective of commercial exploration, Mini Mars Direct has a great deal to offer. Virtually all of the profitpotential of a human mission would be available in a shorter time frame at a modest cost, albeit on a smaller scale. Asan added bonus, the mission would serve as a dramatic “all-up” demonstration of the Mars Direct strategy without anyrisk to human life.

Other aspects of Mini Mars Direct offer unprecedented public relations opportunities that will be instrumental inattracting more investors. NASA is currently planning a joint sample return mission in concert with the French SpaceAgency. NASA during the 2003 and 2005 windows will launch several rover-equipped landers. Each rover will collectsamples and load them aboard a Mars Ascent Vehicle (MAV) which will launch the samples into a low orbit about Mars.The French will launch a Mars orbiter aboard their new Ariane V booster in the 2007 window. The orbiter willrendezvous and dock with the various MAVs, take on the samples and bring them back to Earth. Many details of thejoint mission have yet to be defined. At this time, for example, it is not clear whether ISPP technology will play anyrole in the design. In any case, automated rendezvous, docking and sample transfer in Mars orbit will represent a seriesof significant engineering challenges and costly development tasks. Even if all these technological risks can be managedsuccessfully, the price tag of the joint mission will be very high due to the sheer number of launches and vehiclesrequired. With so many mission elements yet to be defined, cost estimates are nebulous but even the most optimisticguesses point to a price tag of at least $800 million. Less optimistic estimates put the cost of the mission in the billionsof dollars.

Against the backdrop of the cost, complexity and lengthy time horizon of the NASA / French project, the profitablecompletion of the first Mini Mars Direct mission would serve as incontrovertible evidence for the advantages ofcommercial Mars exploration that potential investors in human missions could not ignore.

Should Mini-Mars-Direct successfully prove the concept, our $10 billion initial funding goal may be accomplished morereadily, and our first manned mission will become a reality sooner than anticipated. As that mission, in turn, is beingsuccessfully achieved, the mass appeal for investment funds will be made again . . . and selling the other $40 billion-worth will be much easier . . . and mostly for cash. At that point . . . we stop. Having the restraint to stop selling shareswhen we reach our ultimate sales goal is not just worthwhile on its face, it will help leverage greater per-share returnsdown the road. One we’ve reached the our market-cap goal of $50 billion, there will be no need to ask for more.

Payback Is Not “A Bitch” . . .Not only will the first flights be profitable, but even before that time, the fund itself will be profitable to the bond-shareholders. Capital that’s unapplied toward Mars-specific purposes (short term) can be invested in more traditional,fully liquid cash management vehicles (like the Kaufman Fund) and earn a nice rate of return unless / until it is needed.We’re going to be under a lot of media (and some regulatory) scrutiny from the get-go. However, every public event isnot necessarily “public.” Remember, we can sell both media time and media space at a big rate . . . those proceeds goback into the Fund. All data returned from a “private” mission is property of the investors and can be rented / sold /leased / licensed. More details on that follow below.

American-style marketing techniques are second to none worldwide, judging from the plethora of highly visible,branded products and services available in the marketplace.

By utilizing similar marketing wizardry, and tapping the incredible creative talents of Madison Ave. in ways never usedbefore in the financial world, we stand to gain a new world.

The first $5-10 Billion will probably be a heavy mix of corporate entity in-kind offerings (as much as 30%), someprivate / federal grants or loan guarantees, plus a small percentage of individual investors worldwide. Once we’ve done

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it, however, we’ll probably raise the rest of it, again, in a fraction of the time, and mostly cash! At that point, our owngovernment and NASA will want to get into the act, if only to stop their own PR-bleed at being shown up on whatthey’ve always believed to be their private turf.

One Plausible Scenario for the Initial Breakdown:

If we accept Zubrin’s “$5 billion/per” estimate for one manned mission, we will have enough for two manned missions– missions that can begin paying for themselves immediately.

How do we make money?

HOW DON’T WE MAKE MONEY????

1. Mediaa. Media salesb. Documentariesc. 3rd party licensing

2. Dataa. Helmet VR datab. Resource datac. Hard scienced. Engineering

3. Post-missiona. Videosb. CD-ROMSc. Resource ‘datapaks’d. VR “arcade”e. Soil packets

MediaThe first manned mission to the Red Planet will be the first huge media event of the new century. And if it’s going tobe privately financed, then the investors can (and should) dictate what media access is going to cost! This event is not“free”! For exclusive coverage of the launch alone we can charge a staggering amount of money. And don’t feel sorryfor the network with the “exclusive.” They in turn will be charging their advertisers sums that will rival those chargedduring the Super Bowl . . . if not a lot more. We can also host subscriber pay-per-view exclusive event coverage of boththe launches and the landings and first footprints. Ten to twenty million subscribers worldwide could net a hefty sum,in the range of $500-$750 million. We offer inside views that the networks won’t have access to. This is an issue thatwould have to be negotiated carefully between the Fund, the network, the Pay-per-view organization and the Contractor

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to ensure the greatest good for the greatest number and that all parties are satisfied, but it is definitely something thatcan be worked out.

We can offer some free web sites containing information on the mission, but most importantly, exclusive subscription-only sites for live updates, streaming media and non-proprietary data downloads all during the mission. With properadvance marketing, we could conceivably reach ten to twenty million subscribers, perhaps more (rememberPathfinder?). If we only charged $100/ea for the service, the mission’s nearly half paid for.

Have you ever seen a NASCAR race? Look closely at the cars – they have ‘sponsor stickers’ all over them. Our boosterswill look the same way. An Ares-class booster has roughly the same general configuration as the current Shuttle, as itis comprised of many of the same components. The combined surface area of two solid boosters and the external tankalone is approximately 25,000 sq. ft. At $5,000/sq. ft, that adds up to $125 million in one-time ad revenue.

There will be a great number of science / engineering / human-interest stories taking place in the time leading up to thelaunch. We’ll need to line up TV time in advance to promote the Colony Fund during commercial breaks. CNN doesthis all the time. Its “News from Medicine” segment is sponsored by a pharmaceutical company, for example. In fact,TV ads should be very aggressive. Every sci-fi movie on the major networks should carry Colony Fund commercials,particularly if Red Mars (the popular novel dealing with the colonization of Mars) truly does become a miniseries,courtesy of Mars Society member James Cameron.

On-line trading: The Fund manager will have on-line trading capability, and Fund shares will be able to be sold directlyto individual and institutional investors. Internet marketing is going to do nothing but grow in huge proportions as thenew century dawns, and is something that will pay us immense dividends to take advantage of.

3rd party licensing: Sales of aftermarket “Mars Direct” products could be a nice lucrative sideline. We either take inthe licensing fees on all “official” products, allowing 3rd parties to produce and market them, or form a side-companywith Fund venture dollars, produce t-shirts, hats, etc. on our own – and distribute them wholesale. Every single profitopportunity pays dividends to the Colony Fund.

The first manned vehicle to return from the Red Planet becomes an instant touring attraction, and we can chargewhatever the traffic will bear for entry, at least as much as that charged by your major theme parks. Ernest Shackletonpaid off his debts between Antarctic expeditions by selling tickets to guided tours of his ship.

DataOpen-Source utopian dreams simply don’t apply here. If we’re spending $5-10 billion to go to another planet, all dataretrieved becomes VERY precious, and in fact has market value! It must have market value; else the initial investmentto obtain it can’t be justified. This is not an obstacle in any way: data from any / all unmanned probes we send belongsto us, exclusively.

When the first human explorers arrive on Mars, I propose that their helmets be mounted with 3-D-effect digital stereocameras. Every move they make, everything they say (while on duty) will be digitally recorded and uploaded daily.Just another aspect of the total mission plan. Back in Silicon Valley, our exclusive VR arcade programming mavens areusing that data to create the ultimate arcade experience, which we can charge a bundle for! Japanese businessmen payas much as $1000 a pop for greens fees, of all things. How much would you pay to spend an hour on Utopia Planitia,as though you were really there? We would own the rights to that as well, and that alone would be a real moneymakerfor the fund.

Keep in mind the long-term goal of all this isn’t just short-term feel-good glory. We are attempting to open up a worldto colonization – and colonies must pay their way. That’s how / why this country was colonized. The hard-science datawe find – specifically resource data, underground water sources, etc. – will be very valuable to the 3rd party

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organizations, largely mining combines and global science consortiums, who will want to develop those resources. Nowit becomes fun – for when we discover a particularly rich lode of a specifically desired resource, we in turn enter into afee-arrangement with the developer (better yet, take bids from several!). The developer is on his own after that, but longterm, as profits from development come in, the fund gets a “commission,” perhaps for 30 years or more. Again, payingthe bills from the get-go (this all also applies to medical data collected).

By that point, a legal structure for the sale of land / mineral rights will by necessity be in place. And again, theContractor, and by extension, Colony Fund bond-share holders, will get a cut of the deal – by virtue of the fact that wegot there first with the best. We’ll get paid either way, because future developers will need the hardware, facilities,boosters, landers, rovers, etc., and our venture partners are going to sell them!

Hard Science / EngineeringAny new mission hardware or infrastructure developed for both unmanned and manned missions becomes joint propertyof Mars Direct, Inc. and its primary capital investor, the Colony Fund, and by extension, the shareholders. Again, thathardware may ultimately be mass-produced for sale / lease to other independent business entities engaged in the actualdevelopment of Martian resources.

Post Mission – Back here on Earth

Other things we can market, via independent companies we fund and/or license to sell these products:

Survey “datapaks” – High profit potential, as we would be selling valuable resource-imaging-survey data both from theunmanned probes, and the manned missions;

VR arcade – detailed above – also high profit potential;

CD-ROMs – “low-end” version, based on the arcade (but then again, who can predict where such technologies will bein 10 years??? They may end up one and the same thing);

Home Media sales – DVDs / Videos of the mission highlights.

Soil – This is a new one recently brainstormed, and will probably only work for the first manned mission, but this alonecould make the mission pay for itself. In a 1999 special issue on space produced by Scientific American, the article“Making Money in Space” mentioned the efforts of the firm Applied Space Resources to fund its lunar probe bybringing back 10 kg of soil from an area no probe has ever been. The firm would give half of it away to researchers,and sell the other half to the public at $6,000 per gram. This is a neat trick, as (1) it sets a high market value forextraterrestrial soils, (2) raises $30 million gross, and, just as important, (3) allows for an equal $30 million charge-offon the books for tax purposes.

Now if that can be accomplished in circumlunar space, what might the return be for Martian soil from the first mannedmission? If we do the same thing, perhaps charging as much as $10,000 per gram, then a single metric ton of Martiansoil would literally pay for the entire cost of the mission, all by itself! Everything we do in addition is pure profit!

What’s Available Right Here, Right Now?Much research has been conducted highlighting the profitability of lunar exploration and LEO projects. No formal studieshave targeted potential economic returns for a Mars specific manned mission. Nevertheless, data collected for othercurrent and potential markets is helpful to identify possible economic returns and resources for Mars related missions.

A space market summary conducted in 1997 by the Commercial Space Industry of NASA projected the future in 17orbital marketing categories. Significant increase in payload demand was noted in communications (satellite launch

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services), remote sensing, outer space burial, entertainment, and advertising. No market projections were listed for 9categories; Movies, Novelties, Tourism, Settlements, Waste Disposal, Mining, Earth Transport or Utilities and, nochange was projected in; Military, Science / Technology and Manufacturing. The study maintained that viable privatecommercial adventures would require an internal rate of return 15-25% within the first 10 years of operation and wasbased on a hypothetical $5 billion investment.

Revenues from each flight were determined upon the payload capability and the price per flight, balanced against therecurring cost charged to that flight, repayment of the investment debt incurred in constructing the system, and returnto the commercial investors. The final analysis concentrated on developing a bounding set of parametric conditions withregard to the financial feasibility of any commercial system. Such offers gross baseline, albeit conservative, payloadfee guidelines to entrepreneurial space companies. Considering the restriction of a 10-year investment return and lowprojected development in the other 9 major potential markets, it is not surprising the study concluded that governmentassistance would be required for commercial space development.2

Conversely, the Colony Fund plan does not depend on the traditional guarantee of a 25% 10-year investment returnbased solely on commercialized space missions, nor does the initial “Mini Mars Direct” mission demand a minimum $5billion investment. Ironically, however, the Colony Fund model of funding, investment and IRR fall well within thegovernment projected economic viability criteria!

The entrepreneurial space community is undisputedly growing, in spite of governmental, legislative and legal obstacles.For example, the Houston Space Society that is known to support private non-government supported commercializedefforts follows the activities of over 43 smaller technologically skilled U.S. space oriented companies (some of whomare presenters at the Mars Society Convention). The Society also keeps tabs on the entry of new space orientedcompanies continuously entering the stock market. NASA generously provides on-line access to the listing to 151 U.S.privately owned “prime contractors” as well as a PDF index (also available in CD ROM) of over 7,000 equipment andmaterials suppliers.

[At publication time, research on European / Asian efforts, international dollars and the associated tech companieschasing those dollars was not complete, and hence could not be included in this portion of the study; nor were hardnumbers pertaining to international space legal issues. However initial work suggests that those additions in futureissues will only help to bolster our “case for Mars funding”]

In summary, estimates of international market dollar exchange for direct space related products range from $127-850billion to over a trillion plus when support supplies, materials and ancillary services are included. In short there ispresently no shortage of global “dollars” being spent on space related technologies, even prior to the development offuture potential markets such as celestial mining, space entertainment, commercial space travel markets, etc. There doesthere appear to be a lack of competent, experienced space technology contractors, legal consultants and potentialpartners, all of whom are most likely ready to supply goods and services necessary to meet mission specifications. TheColony Fund need only provide a focus point to funnel these existing resources.

Putting It All TogetherTo reiterate – our Mars-Colonization funding and implementation will be a multi-pronged attack: separate entities, butmutually interdependent.

Capitalization entity: “The Colony Fund,” which will raise $10-50 billion ($10BB short-term, $50BB long-term) byselling bond-shares either via actual cash ($1000/share) or via “cash-value bearer-bonds” from corporate entities (e.g.,an aerospace firm makes a “promissory note” for $1 BB worth of product and support in exchange for 1 MM shares).CF Board of Directors, including 3rd party Fund Manager, oversees short-term investment of cash in “conventional”manner, as well as effective usage of corporate bearer-bonds offered to Prime Contractor.

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Project Management / Engineering Prime Contractor: The people who actually “do” Mars. Contractor will be primarilyfunded by the Colony Fund, and will be the only organization allowed by the fund’s managers to cash in the corporatebearer-bonds. The Contractor is comprised of the Project Management group responsible for the overall effecting of theactual Mars missions, the people that “bring home the gold.” However, all data retrieved is owned by the investors –namely the Board of Directors of the Colony Fund, and by extension, all its shareholders.

There will also be many other side-business ventures, independently operated, and started up with Colony Fund money,the profits of which roll back into the Fund.

All media / new media access and products will be marketed towards the goal of maximum return for the shareholders.

Official licensing of consumer products and technological spin-offs, the sale / lease of mineral rights, etc., will all helpto “grow” the fund. As it grows, and colonization gets started, the Fund can begin to make direct investments in colonialefforts on a case-by-case basis.

Why are entities 1 and 2 separate? There are 5 reasons:

• If the Colony Fund were a more traditional, pure-cash operation, the Colony Fund and the Contractor could be oneand the same. Keep in mind that the Colony Fund is taking in more than just cash – they’re also accepting “as value”the equivalent of used cars, chickens, corn whiskey, and Queen Isabella’s jewels.

• The fund, accountable to the shareholders to increase fund value, may wish to make some traditional investments asa short-term hedge, or even make riskier venture investments in non-Contractor-related endeavors. It needs thefreedom to do that while also serving the long-term goal.

• Given that the Contractor is accepting a combination of both cash and corporate-bearer-bonds in lieu of pure capitalfrom the Colony Fund, the bearer-bonds redeemer may also become a strategic partner with the Contractor

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• If the Contractor wishes to make other strategic alliances with companies unwilling to invest in the Colony Fund ormake supplemental finance arrangements with other financial institutions or even with the U.S. government, as anindependent entity it has the freedom to do that, so long as its obligations to the Fund are met.

• The Contractor will actually be developing (or subcontracting development, for a fee, of) mission hardware andinfrastructure, using Colony Fund capital. Once developed, that hardware and any patents become joint property ofthe Colony Fund and said Contractor – property they can then in turn jointly market to both repay Fund obligationsand increase the Contractor’s market cap. Within 30 years the Contractor could earn enough on its own to becomea fully economically independent operation – perhaps even “going public,” the old fashioned way, at which point theColony Fund and the Contractor could opt to go their separate ways.

Results? If we put $50BB in a pot over the next decade and manage it in this manner, 30 years from inception, that potcould easily grow to $500BB+, making each original share now worth $10,000 or more! Your “30-year bond” just grewto 10 times its value.

Someone who buys 30 shares today may find that, 30 years from now, his investment has grown such that, if cashed in,he could grubstake 4 of his grandkids to one of the colonies on Mars – transportation included.

At that point, the original investors would have a choice: to cash-in their shares, or roll them over into a new, but moretraditional financial instrument, as the Fund evolves into the first true Martian industrial-development-banking entity –behaving in a far more “traditional” fashion, but hopefully never forgetting its roots. At that point it might get into thedirect loan business for individual colonists and their families. No doubt, the early Martian colonies will have their ownnon-traditional ideas about local banking regulations. The Fund could become the first true banking institution on theRed Planet.

Issues to Research FurtherIn developing this concept, I’ve worked closely with the people to whom I’ve given co-author credit on the title page.Their input has been highly valuable in that, in the first place, no one likes to work in a vacuum, and secondly, there area lot of legitimate issues we will need to research further in order to finalize this concept and develop it into a workingmodel for funding Mars colonization. Let’s look at them here.

A clearly defined unit of investment: the “bond-share”

As this is the first time we’ve ever created such a financial instrument, we appreciated the need to determine in concreteterms what shape the single investment unit will be. Is it a true “bond,” “convertible security,” or more of a “long-termaccrual security”? Is it a “sector fund,” “bond fund,” “bond trust,” “unit trust”? What we have described above seems,at first, to defy conventional description, but 30 years ago, who knew what a “junk bond” was? We need to clearlydefine, presumably with assistance from the pros who may manage this fund for us, what exactly this thing is, beforewe can mass market it.

In fact, there may be more than one unit of investment . . . there may be 10-year and 20-year units in addition to the 30-year, but different rules may apply, to encourage the longer-term commitment.

Tax and legal implications

Can this unit of investment, this $1000/per Mars investment-venture-30-year-bond-share of ours, be tax-deferred? Canthis unit be a welcome addition to individual IRA and 401k portfolios? Legal research remains to be done in this area,which will no doubt necessitate bringing a lawyer onboard who’s experienced at writing prospecti the SEC will approve.

Also, we’ll need to research further, with any fund manager we select, a way to structure the ‘barter implications’ of thehardware-for-shares aspect to allow it equal value with straight-up cash investors, in a manner regulatory bodies will approve.

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What if people just want to GIVE us money?

While the scope of this paper has its prime focus on Mars venture capital funding, it is entirely possible that, with someearly successes behind us, there may be people out there who would actually wish to include Mars colonial efforts intheir estates, charitable trusts, etc. If for some reason, the Mars Society itself did not become the direct beneficiary ofsuch philanthropy, a “sister” organization could easily be formed to handle it, “The Colony Fund Foundation.” It wouldserve a great purpose by helping to endow small Mars / space advocacy groups who would not qualify for Colony Fundmoney. It could also offer scholarships, “genius grants,” etc. More research will have to be done, of course, but theresults of such endeavors could benefit the long-term goal in ways we can’t predict on these pages.

Corporate structure, etc.

To start with, the Colony Fund will have to incorporate somewhere as a legal entity. Forms will have to be filed withthe SEC, a Board of Directors will have to be formed, and seed and working capital will have to be raised. The Boardwill be responsible for managing the Fund’s business affairs, using all due-diligence, under the terms of the proposedoverall relationship structure, and the Fund’s general charter. They will also select the Fund’s co-manager, the PrimeContractor, and the Marketing / Advertising firm. Strategically, initial funding will probably be limited to moretraditional investment vehicles until a certain minimum balance is attained ($300 MM, if we go with the “Mini-MarsDirect” model) then it can be split between a conventional fund and the Contractor.

Once Again, a Call to Action“The dream will continue. Men and women will still explore space. The frontier has been broken; likeany human frontier that has ever existed, it will never be abandoned. Mankind has never turned its backon any place it has visited at least once. We have an innate urge to explore new worlds; no matter howfar away they may be, or how hostile the environment. Like it or not, it’s part of our genetic makeup.”

— Allen Steele

The Tranquillity Alternative:How do we start? For now (should this concept fly with those who peruse it) the next logical step is to do as muchresearch as we can on our own, followed by finding advisors from the financial community for the tough questions wecan’t answer ourselves, then finally seeking grants or “business angel” funding to actually form the Fund’s businessstructure. Timing is crucial, and I believe the time is ripe for new initiatives. If we are to seriously move on this, itneeds to be moved on without delay. Mars won’t wait!

This is a first-draft version of concepts that have evolved over the last year. I suspect that, over the next few months,interested readers of this proposal will bring to my attention concepts I haven’t considered, as well as any fundamentalflaws herein (hopefully few) that need addressing.

This is meant to inspire, to get the creative juices flowing, and to help us THINK BIG – for getting to Mars in ourlifetimes requires that. Let us rise to the challenge.

We can do this. We can sell the bond-shares. We can sell the Congress into creating the regulatory climate that allowsthis to flourish. Our suppliers can build the machinery. So what are we waiting for?

Every economic construct or investment vehicle ever usedsuccessfully by humankind has been just that – an artificial construct of human thinking.

There is no reason we cannot go out and create as many more such constructs as we deem necessary in order to fulfillour future needs.

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Two bicycle mechanics from Dayton, Ohio succeeded in heavier-than-air flight against all the conventional wisdom ofthe era – perhaps precisely because they were not beset with trainloads of well-meaning “experts” constantly yammeringat them that it couldn’t be done.

Napoleon once summed up his military strategy very neatly: “If you want to take Vienna, take Vienna.”

Indeed. And if we want to take Mars, then we take Mars. Nothing else will do.

References1. Apollo 18, 19 and 20 were canceled in 1970. Originally Harrison Schmidt was to have been LM pilot on 18, but was bumped to 17 out of

NASA’s desire to have the last lunar crew include a geologist.2. Until now these markets have never been explored by NASA. Commercial civilian space travel alone was estimated to have a potential Gross

Revenue value per year of over $12 billion US in 1994.

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Mars Habitat: Putting The Pieces Together

Tom Hill[2000]

AbstractArguably, the most critical vehicles for any Mars mission are those which support the crew for long periods of time.Typically, in the design of such complex systems, only the very highest of functions (mission, etc.) are considered fromthe top down. Systems and sub-systems are split up in the race to produce the product. This method does not takeadvantage of some system interconnectedness, which in the end can save complexity and mass. A habitat design isdiscussed, emphasizing its co-dependent systems and how they augment each other. In the process, one possible layoutfor the craft is presented, along with useful references for anyone interested in the field.

AbbreviationsLCH4 – Liquid MethaneLOX – Liquid OxygenCO2 – Carbon DioxideH2 – Hydrogen

IntroductionThere are many issues which must be addressed before a serious Mars habitat design can start. History shows, however,that detailed discussions about the method to reach a goal start early, even before such a goal is officially mandated.1Analysis of various design options now will create a basis from which to carry on further design, and prevent a start at“square zero.”

No vehicle has been built, or even taken past the early design phase, which has the requirements facing a Mars habitat.The system will have to function for up to 220 days, providing everything for its crew with only their knowledge andskills, along with the spare parts and materials already aboard to sustain operations. At that point, depending on themission scenario chosen, it is possible that another biosphere could take over for the habitat in a catastrophic situation.No matter which of the current options are chosen, a maximum of three support systems (using Mars Direct terms: thehabitat, the Earth return vehicle, and the backup ERV) will support the crew for almost 1000 days.

While the current attention is focused on whether or not making a mission cheap causes it to fail, other research indicatesthat a mission’s complexity is the greatest factor2 in predicting success or failure. The complexity of a crewed missioncompared to a robot mission can not be understated, but any steps that can simplify the crew’s support system whilemaintaining or increasing margins should be carefully considered.

This paper assumes a Mars Direct / NASA reference mission-type approach to sending the first humans to Mars. Thedesigns are based on a four-member crew, although other sizes are considered. Calculations worked from a 220 day tripto Mars with a methane/oxygen propulsion system for a short final landing maneuver after a parachute descent.

“Interdependency”There are many examples of interdependency in current and past space flight. The fuels cells on board the space shuttleand the Apollo Command Module are one case. Hydrogen and oxygen are brought together, forming electricity andwater. The electricity is funneled off to power the ship, while the water is potable, and stored for crew consumption.Without this interdependency, two separate systems would be required for power generation and water storage. Themass savings for this particular example are debatable. Solar panels could have been used, thus only requiring a supplyof water equal to the mass of hydrogen and oxygen brought aboard to run the fuels cell. In the case of the space shuttle,solar panels are not really an option due to the need to retract them before atmospheric reentry.

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Tom Hill; 16819 Centerfield Way, Olney, MD 20832; [email protected]

This solution combined 2 systems to provide necessary services while saving mass. Because of technical advancements,it is possible to combine more systems on a Mars habitat, creating a greater savings in mass.

System 1 – Propellant Chilling:Cryogenic propellants, on average, provide a more efficient propulsive force per kilogram used than storablepropellants. While the propellants are “better” theoretically, there are handling and storage concerns that make themless attractive. Refrigeration systems, which can keep the propellants cold, are heavy. Another option of allowingpropellant to boil off wastes the liquid. Any system that can remove heat from a cryogenic fuel and oxidizer with aminimum weight penalty would be an improvement.

System 2 – Food Freezing / Refrigeration:Astronaut food as come a long way since John Glenn squeezed a snack out of a tube during his 3-orbit mission. Foodaboard the space shuttle allows a much greater freedom of choice by the crew and is much more palatable. Longermissions, especially those which have no resupply plans, must increase the food variety while improving as many otherattributes as possible (freshness, storability, etc.).3 A low-mass freezer and refrigerator would aid greatly on all fronts.

System 3 – Carbon Dioxide Scrubbing:Each astronaut on board produces approximately 1 kilogram of carbon dioxide a day.4 In early missions, lasting lessthan two weeks, a valid option was simply to scrub the offending gas out of the air, pulling it into canisters for disposal.Most systems used lithium hydroxide for this purpose. Skylab used a system that pulled the carbon dioxide out of theair without forcing it into a chemical solution, and dumped the gas overboard.

Long duration space missions cannot afford the luxury of disposing between four and seven kilograms of material fromtheir biosphere daily. This precious compound must be drawn out of the atmosphere and used within the system. Notdoing so is a waste of launch mass.

System 4 – Water Generation:Current Completely Enclosed Life Support Systems (CELSS) have a reuse efficiency of between 90 and 95% for potablewater. There appears to be agreement that a long-term mission requires a recycling percentage of 95% or greater to beviable. A source of potable water, outside the recycling process, would be useful to replenish supplies, perhaps relievingthe stringent 95% recycling requirement.

System 5 – Fuel Generation:Although this author knows of no attempt to use this technology in actual space flight, an abandoned concept formaneuvering fuel for the space station involved fuel generation. Wastewater would be electrolocized into its componentgases, which would be used to run maneuvering thrusters on board. This option was dropped due to development cost,in favor of transporting other fuel and oxidizer from the ground. The life-cycle cost savings of this decision are unclear,but marginal at best.

The fact is that on-board fuel generation can save mission mass at a critical time . . . liftoff. The less mass required forfuel, the more mass can be devoted to other endeavors.

Influences On Habitat DesignA robust system which merges all five systems is feasible. The design has no moving parts, and simply involves channelinga gas (hydrogen) throughout various ship systems. The system can be pressure-tested to prevent leaks, and the hydrogencan be topped off until just before launch, allowing design simplicity. The design is graphically depicted in figure 1.

Thermodynamics and safety requirements dictate that propellants should be stored as far away from the crew aspracticable. This allows cryogens to be stored at a vacuum, the best condition for passive storage. Simple shading cutsthe ambient temperature to approximately -150°C, not including heat radiating from the ship’s structure itself.

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For this particular design, the tanks are placed above the habitat module, describing ‘above’ as the direction oppositethe gravitational field the craft is exposed to in almost all conditions. A centered hydrogen tank surrounded by 5 others(2 LOX and 3 LCH4) cuts the amount of piping required to flow between them.

Boiloff from the hydrogen tank is channeled through cooling coils, first within the oxygen tanks, then within themethane tanks. This stepping process slowly raises the temperature of the hydrogen, while at the same time drawingheat out of the LOX and LCH4 tanks.

Next the hydrogen is routed into the habitat module. While this poses some threat due to leakage, there are no valvesin this system. While hydrogen is a notoriously difficult gas to contain, the lack of any moving parts within its pipingwill be a tremendous aid.

Within the habitat, the hydrogen, now warmed to at least the boiling point of methane (112K or -161°C) is routed aroundthe deep freezer. This freezer is primarily intended to store foodstuffs for the crew, although the temperature should below enough to store just about anything requiring freezing. A deep freezer will allow the crew to enjoy a wideassortment of foods throughout their trip, and food variety is important for a group facing long-term solitude. The onlyrequirement for this system to maintain an extremely low temperature is insulation from the rest of the habitat. The deepfreezer is designed to stay closed most of the time, as constant opening and closing of the door will pull water out of theatmosphere and cause frost. The mission design would have a crew member enter the deep freezer occasionally on asupply run, pulling foods out for the next period of time. A sub-chamber within or near this deep freezer will be setaside for common use. Stocked with supplies from the larger unit, it will allow easy access to the current foods.

The next loop for the ever-warming hydrogen is the refrigerator, which has an obvious use for leftover food and storageof recently thawed supplies.

While the hydrogen cools ship’s systems, the on-board carbon dioxide scrubber gathers the crew’s waste gases. As thegas is gathered, it is stored for a short time, then fed into the Sabatier chamber.

After passing through the cooling systems, the hydrogen is funneled to the same reaction chamber as the carbon dioxide.As described in The Case for Mars,5 the Sabatier reaction combines hydrogen and carbon dioxide in the presence of acatalyst to form water and methane.

The equation for this chemical reaction is as follows:

This mechanism, when given 1 kg of carbon dioxide and mixed with .182 kg of hydrogen, produces .364 kg of methaneand .818 kg of potable water.

Assuming a 220 day trip to Mars, at 1 kg of CO2 production per crew member, the numbers look like this:

Table 1. Masses required and produced using the Sabatier reaction with varying crew sizes.

(amounts in kg)

This reaction takes place at approximately 400°C, and is exothermic, producing its own heat. As long as the chamberis properly insulated, the reaction can sustain itself. The byproduct is water vapor and methane, which can be bubbled

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through an existing water supply (separate from the crew’s water supply in use, to allow sampling before mixing). Thisprocess condenses the water out of the mixture and allows the methane to move on to its cryogenic storage tank.

Figure 1. Flow Diagram

Mass SavingsIt is easy to look at Table 1, and believe this system, applied with a crew of six, will save over 1500 kg in launch massfor the habitat, but the picture is more complicated. Water recycling will be a way of life on board a Mars habitat, andthe current percent recycling rate for potable water is 80-99%.6 This means that for every 1 kg of water used by thecrew, .8-.99 kg is returned to the system.

Assuming that the water is recycled at 90% efficiency, and over the course of the journey to Mars water is recycled 10times, the picture changes. For analysis, this system’s water contributions were split into 10 parts, spaced evenly acrossthe mission. This fact yields the following numbers for useful water for the crew (units are kilograms):

Table 2. The effect of recycling on launch masses and water production systems’ numbers.

Launch mass calculations are based on the amount of water required at launch to provide as much water as the systemthroughout the cruise to Mars. The launch mass requirements are lower than the system production numbers becauseall of the water is cycled ten times.

The fuel mass savings are as straightforward as Table 1 indicates. The crew needs to bring 320 kg less fuel with themusing this system than they would without it.

This passive cooling system removes the requirements for refrigeration systems on board the spacecraft for as long asthere is hydrogen on board. A design decision will be required as to whether or not the habitat would continue to use

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an on-board hydrogen supply to cool the crew’s freezer and refrigerator on the surface of Mars. Without the “warming”provided by the methane and liquid oxygen (no longer necessary after landing) the hydrogen may be too cold forcommon-use freezing. Although on the surface, carbon dioxide is a common element, the additional water that thissystem provides will always be welcome.

Control SystemsIn order to maintain the simplicity of the system, valves and other control mechanisms should be kept out. One simplemethod to allow this is to over-insulate all cold components, and place heaters outside of the reactants, yet within theinsulation. This applies control with a minimum weight penalty. For example, if for some reason the crew is producingtoo much CO2, and the hydrogen is not boiling off fast enough to meet the demands of the Sabatier reactor, heaters nearthe hydrogen tank can be switched on, which will increase the flow of the reactant. Similar measures will be possiblewithin the freezer and refrigerator in case of over-cooling.

TradeoffsWhile the simplicity and elegance of this solution are marks in its favor, its usefulness on a future Mars mission will bebased on many factors.

Safety is one issue. Hydrogen is a difficult gas to contain, and there will always be a concern with storing it near a crew.By keeping the system as simple as possible, this risk is minimized.

Technical tradeoffs are a factor as well. For this analysis, and methane / oxygen landing system was assumed. Anotherpropellant which proves marginally better mass-wise is hydrogen/oxygen. A habitat with a dry mass of 25,200 kg thatneeds to create a 200 m/sec velocity change requires 1,210 kg of hydrogen and oxygen propellants (assuming specificimpulse of 435). The same mass lander using methane and oxygen requires 1429 kg of propellants. The H2/O2 systemis marginally better when considered alone, but when the savings of using this water and fuel system are factored in, thebalance tilts in favor of methane and oxygen.

External ConfigurationFigure 2 is a representational diagram of the spacecraft enroute to Mars. This incarnation uses its spent upper stage togenerate artificial gravity, and keeps the tether and upper stage “dead,” in that the habitat does not use them beyond theircounterbalance properties.

In this design, the rover and other items not required for the trip out are stored between the habitat and the aeroshell.After atmospheric entry and drifting to the landing site under canopy, the habitat separates from the shell and its cargo.The habitat proceeds to a precision landing under its own power, while the aeroshell and its cargo drop under their ownparachute to the surface. When this unit approaches the ground, a small rocket fires to slow their decent the final meters.This design gives the crew more useful space within their habitat.

The propellants and hydrogen reactant are stored on top of the habitat. Solar panels cover one side of the vehicle, andextend above the pressurized envelope to shade the tanks from the sun. This provides a lower ambient temperature forthe cryogens. The solar panels were not included in this diagram for clarity.

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Figure 2. Habitat in Cruise Mode Enroute to Mars

Internal ConfigurationCross Section:The habitat is split into two levels, with equipment bays above the 2nd floor ceiling and below the 1st level floor. Thestructure’s diameter is 8 meters. An airlock runs through the center of the craft, with a hatch (not illustrated) at thebottom. While the maximum diameter of the airlock is small, using its height to an advantage makes up for thedifference. The environmental suits are stored in the upper portion of the airlock, and are brought down as needed. Thisconfiguration is pictured in Figure 3.

Level 1:Refer to Figure 4 for a diagram of this floor. Habitat level 1 is designed for common use by the crew. They are providedwith a galley, common area, laboratory and a latrine/shower facility on this floor. A single ladder leads up to level 2.Although the plan is acknowledged as a violation of normal building fire codes, by this author’s estimation, a fire in thisfacility will lead to so many other problems that every effort must be made to prevent it from happening. If absolutelynecessary, a trap door from one of the wardrooms on level 2 would allow access to level 1.

The water storage tank partially surrounding the airlock contains all the mission’s water (save that which is beinggenerated through the system described in this paper). Potable and cleaning water are stored next to each other,separated by a baffle within the tank. This tank provides a protective ‘shadow’ for the crew in case of a solar storm,and that shadow includes the latrine, for use in case of a long episode. The galley can be found just outside of thisprotection, allowing a member to occasionally grab supplies for the crew during a long event.

The common area and laboratory are configurable to various missionphases. While the laboratory will not see much use during the outboundtrip, its usage will pick up once the crew lands on the surface. Thecommon area will also serve as the physical fitness area.

The latrine location was chosen due to astronaut complaints aboardSkylab, where the facilities were located right next to sleepingcompartments. Astronauts complained of being woken up if someoneused the facilities through the sleep period. It was thought that keepingthe latrine on a separate floor, and proper soundproofing of the crewcabins would remedy this issue. Another concern with the facilities nearthe common area is odor control, which will require careful design of theairflow system.

The galley is designed allowing a cook to do their best with the suppliesat hand. While its location is away from the long-term refrigeration andfreezing compartments, this was necessary to properly place the latrine.It was thought that the inconvenience of having to change levels to getcooking supplies would be better than the convenience of beingdisturbed through crew rest cycles and not having the latrine facilitiesavailable during a solar storm.

The hatches are provided to allow future expansion. Without them, an additional habitat will provide more living space,but not a sense of community that may become critical on Mars. In this image, the feature near the top hatch is acollapsed, inflatable garage for the rover that the crew deploys after landing.

The solar cells (thick black lines to the left of the hab structure in this representation) indicate the position of the sunrelative to the habitat. While the craft will rotate around its common center of mass with its upper stage, the panels mustalways face the sun for power. This same fact removes the need for a water shield the entire way around the vehicle.

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Figure 3. Hab Cross Section

Level 2:Pictured in Figure 5, this level is largely devoted to storage and personnel quarters. The layout chosen gives the optionof adding more staterooms in the same habitat design, should propulsive technology change and allow a larger crew tomake the trip. If this option were taken, the storage seen in this diagram would largely be moved outside and crewsupply runs to the storage space between the hab and aeroshell would become an occasional necessity.

Each crew member has their own stateroom. Effort should be made to make the staterooms as soundproof as possible.Each crew member has room to stand up and change, along with a small work area and bunk. It is recommended thatthe rooms be reconfigurable, either on the ground before flight or during a mission, so that each crew member canarrange and decorate the room to his/her tastes. In a pre-launch reconfiguration, it should be feasible to remove the wallbetween two rooms, allowing married crew members to share one larger cabin.

OTHER SYSTEMSThe following are short topics on other systems within the habitat. Most require more research to be consideredseriously for a mission.

Communications:Figure 2 shows the communications antennae on top of the habitat. While there has been much discussion about usinga despun platform, or an actively tracking antennae for the habitat en route to Mars, a simpler method involves placinga number of fixed parabolic antennae on the habitat, covering a portion of its rotational arc. By including anomnidirectional or “whip” antenna, the habitat can be in communications with Earth throughout its rotational period.

Once the period is established, ground antennae can communicate at low rates, for use by the whip antenna, untilcalculations show the parabolic dishes coming around. When the dishes have line-of-sight with Earth, the data rate canbe increased dramatically. Between the two methods of communications, a high average data rate can be achieved.

Spares:In order to minimize the number of spares required, it is recommended to use common parts where possible. Forexample, any circulating pump, be it for air, water, or waste, can be driven by the same type of motor. The onlyrequirement is that the motor be geared differently for different uses. Using this philosophy, bringing one or two sparemotors will provide replacements for any system that happens to fail.

For a second layer of spares, critical parts (brushes and bearings in the example of the motor) should be brought alongon the mission, as those parts are the most likely to fail.

Clothes Washing:The proposed system for washing clothing on board uses liquid carbon dioxide. While more research is required on thistopic, the chemical is known to have cleaning properties. Once a laundry cycle is complete, the washer need only pullthe carbon dioxide out of the chamber. The clothes will then be dry immediately.

Full Body Cleansing / Water Use:It is recommended that crew members get a rationed amount of water to use for cleaning per week, and allow them touse it as they see fit. For people who prefer longer showers, they have the option of taking a shower once a week, withsponge baths for the rest of the period. For those who prefer to shower every day, they are afforded a quick rinse off.This method allows the crew to choose their reaction to the environment, which has proven helpful to morale in the past.

All water should be heated (or chilled) to a preferred temperature before use. This method prevents the waste of waterrunning down the drain while awaiting cool down or warm up. In the case of the shower, a measured amount of watercan be stored just above the stall, then heated to the user’s temperature preference.

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ConclusionWhile many argue Mars Mission design is premature at this point, the fact remains that it was just such work whichbrought about the recent shift in thinking to in-situ propellant generation and living off the land. Continuing this effortand refining methods, missions to the red planet become more feasible.

The system presented here makes use of an existing resource (crew carbon dioxide) and converts it into usable waterand fuel for the mission. At the same time, the system passively cools propellants and foodstuffs for the crew.

Figure 4. Habitat Level 1

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Figure 5. Habitat Level 2

References1. Baker, David The History of Manned Space flight, chapter 6. ©1981, New Cavendish Books2. Complexity-Based Risk Assessment of Low-Cost Interplanetary Missions: When is a Mission too Fast and Cheap? David A. Bearden, The

Aerospace Corporation Fourth IAA International Conference on Low-Cost Planetary Missions, JHU/APL, Laurel, MD, May 2-5, 20003. Stuster, Jack Bold Endeavors, chapter 10. ©1996 Naval Institute Press4. Eckart, Peter Spacecraft Life Support and Biospherics, p. 92. ©1996, Microcosm Press5. Zubrin, Robert and Wagner, Richard The Case for Mars: The Plan to Settle the Red Planet and Why we Must, p. 150. ©1996 The Free Press6. Eckart, 226-233

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Mars Kites for Human Habitation

William Byron (Joe) Poston, Ph.D.[2001]

ProjectA study to determine the possibilities of melting the ice and snow at mars polar regions to create an atmosphere forhuman habituation. Utilizing parabolic reflectors to re-direct (concentrate) the sun’s energy to create warmth at the Marsnorth pole to melt the snow and ice for Earth-like features.

MethodologyUtilizing the most recent advances in communications (Parabolic Reflectors) and space explorations (Mars Odyssey) itis conceivable to conclude that the feasibility of this concept can become a reality. The reality of Earthlings on Mars isnow a very distinct possibility within our foreseeable future. This concept is easily proven with college math (separatepublication) and the dissertation is presented to familiarize the scientific (& political) communities to the feasibility ofthis concept and give credence for its implementation. This concept utilizes a well know paradigm (Man on Mars)…but which had no practicality except to send someone there in a clumsy space suit. This abstract, utilizing modern dayaccomplishments, illustrates a different methodology.

Exact Sized Solar System

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William Byron (Joe) Poston, Ph.D.; Phone / Fax: 727-343-3488; [email protected]; www.ij.net/personalized; P.O. Box 48974, St. Pete,Florida 33743

Inner Solar System

Sun and Designated Planets

Solar System Dimensional Aspects

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This data is shown to provide a comparison of existing and future study of the parameters.

Venus Kites: Providing Cooling Aspects

Sun and Designated Planets

Mars Kites: Providing Heating Aspects

The Basic premise is to have the kites redirect the Sun’s rays (energy) onto the surface of the Mars planet at the northpole region. The Kites can meander around Mars planet with the rays re-directing over a large area, or be stationary andlet planet revolve under the kites rays to heat a specific spot such as the explorers habitat.

As these areas are melted into streams of water, the evaporation process will be stimulated and an atmosphere will becreated. The solar winds will push the atmosphere on the back-side of the planet as it rotates daily and the cloudyatmosphere will exists in the solar winds eddy currents.

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As shown in the examples, the Sun’s rays (energy) are converged (1-km diameters) or (10 / 20-km diameters) directlyonto the surface of the Mars planet 24-hours daily, 7-days per week. As the polar ice caps begin to melt the water willflow into the existing rivers and basins for the Mars explorers at their (placed) Mars Habitats.

Yes! The Mars habitat is the most significant aspect of life on Mars. Without it Mars will never be explored, or laterpopulated. Also, without melted water there is no hope of Mars explorations.

Mars Kites: Providing Heating Aspects

This plane is flat – yes! flat. But as shown below, it is really at 45°, thus no shadows are placed onto the Mars planet.Heating occurs 24-hours daily, 7-days per week.

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Mars Habitat: (as shown at the NASA Kennedy Space Center)

The knowledge and wisdom that visualized this concept for Mars habitation is what’s required for future explorationson the Mars planet.

However, the Mars planet must be made suitable for humans to live without very bulky space suits, which limits theirability to explore.

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Mars Libration Point Mission Simulations

Jon D. Strizzi; Joshua M. Kutrieb; Paul E. Damphousse; John P. Carrico[2001]

AbstractThe equilibrium points of the Sun-Mars system bring some unique characteristics to the discussion of future inner solarsystem exploration missions, particularly an expedition to Mars itself. Existing research has identified potential utilityfor Sun-Mars libration point missions, particularly for satellites orbiting each of the co-linear, near-Mars, Sun-Marslibration points (the L1 and L2 points) serving as Earth-Mars communication relays. Regarding these Lissajous orbits,we address questions of “Why go there?” “How to get there?” and “How to stay there?” Namely, we address utility andusefulness, transfer and injection, and station-keeping. The restricted 3-body problem involving a spacecraft in thatsystem is reviewed; and past and present research and proposals involving the use of these orbits are summarized anddiscussed. We use commercial, desktop tools (Satellite Tool Kit (STK) / Astrogator) for simulation and analysis ofEarth-Mars transfers, Lissajous orbit insertions, and station-keeping trajectories. On-going, successful collaborationbetween military and industry researchers in a virtual environment is demonstrated. Much of this study focuses on 2016Earth-Mars transfers to these mission orbits with their trajectory characteristics and sensitivities. This includes analysisof using a mid-course correction as well as a braking maneuver at close approach to Mars to control Lissajous orbitinsertion and the critical parameter of the phasing of the two-vehicle relay system. Station-keeping sensitivities areinvestigated via a Monte Carlo technique. The resulting data provides confirmation and insight for existing researchand proposals, as well as new information on Mars transfer and Lissajous orbit insertion strategies, communicationscoverage, and station-keeping sensitivities. The data provides new information on these trajectories to futureresearchers and mission planners

Introduction“NASA’s vision is to . . . focus more of our energy on going to Mars and beyond.”

— Dan Goldin, AWST, January 2001

“All the questions we have about Mars could now be answered . . . if we could just walk around on theplanet for a few days.”

— Michael Malin, Malin Space Science Systems, National Geographic, February 2001

“NASA is seeking innovation to attack the diversity of Mars . . . to change the vantage point from whichwe explore . . .”

— CNN, June 2001

As NASA and the space community renew their focus on Mars exploration, student researchers find several topicsawaiting further study. From our work that originated in an advanced astrodynamics course at the US NavalPostgraduate School, we became interested in Mars, various aspects of the three-body problem, and the Lagrange orlibration points, and we were eager to team with industry to conduct mission simulations and analysis. We examinedseveral documented research efforts dealing with diverse aspects of these topics.6,9,11,14,19,20,24,25 The concept of usingcommunication relay vehicles in orbit about colinear Lagrange points to support exploration of the secondary body isnot entirely new, being first conceptualized in the case of the Earth-Moon system by R. Farquhar.24,25 An innovativeapproach on the concept that caught our interest was that introduced by H. Pernicka, et al, for a 2-satellitecommunications relay with one spacecraft in orbit about each of the co-linear, near Mars, Sun-Mars libration points, L1and L2.6 Further work by graduate researchers (Kok-Fai Tai and Danehy) refined this proposal and conductedinvestigations into the technical and fiscal aspects of such a mission, including trade studies on communication relayconstellation options.15,16 This analysis resulted in some favorable conclusions and rationale for a Mars communicationrelay system that utilizes 2-spacecraft in large amplitude Lissajous orbits, including system cost and performancemeasures comparable to a 3-spacecraft aerosynchronous system.

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Jon D. Strizzi, Joshua M. Kutrieb & Paul E. Damphousse; Department of Aeronautics & Astronautics, US Naval Postgraduate School,Monterey, CA <[email protected] / John P. Carrico; Senior Astrodynamics Specialist, Analytical Graphics, Inc., Malvern, Pennsylvania

A primary purpose of this work, then, was to re-examine the 2-vehicle system orbiting the Sun-Mars co-linear librationpoints, including transfer orbits and station-keeping, through desktop computer simulation using full-force models andthe interplanetary propagation / targeting techniques of the STK / Astrogator module. Essentially, we wished to see howpast studies and data using simplified models compared to our new full-force model targeting and propagation, and togenerate innovative scenarios and data for future missions.

An additional purpose of the project was to demonstrate successful collaboration between military graduate researchersand industry professionals. Timely, affordable results from specific research can be obtained when diverse groups suchas these can work, virtually and collaboratively, on pieces of a complex problem. These ideas flow into another purposeof the study: to show how commercial desktop computing can be used to easily create and analyze these types ofmissions and problems, again leading to faster and cheaper studies by more researchers. As far as we could determine,this area of study for Mars missions has not been investigated previously in this manner.

As output for this study, we expand upon discussions of the usefulness of these orbits for Mars missions as well as re-examine the 2003 Earth-Mars transfers and L1 libration orbit insertions presented in the original Pernicka study and thefollow-on work. We then expand that simulation and analysis to include the planning horizon of a 2016 transfer andmission orbit insertion, considering an L2 orbit insertion as well. We investigate the effects of orbit amplitude on insertion∆V requirements and show some innovative mission orbit insertion techniques that results in ∆V savings, namely usinga Mars swingby and braking maneuver to assist in the insertion. We also investigate trajectory design methods to achievethe necessary two-vehicle phasing for effective mission operations. We present figures of merit to assess thecommunications relay coverage and analyze station-keeping ∆V requirements via Monte Carlo simulation and analysis.

Please note that this report consolidates and expands upon some results presented at American Institute of Aeronauticsand Astronautics and American Astronautical Society conferences,22,23 along with additional explanatory material. Allof the simulation and analysis presented here represents a first effort of utilizing full force models and current desktopcomputing tools to generate some useful data on the Sun-Mars libration point transfer and communication relayproblem. The data does not represent optimized numerical solutions or proposed detailed mission designs, but ratherinformation and baseline data for follow-on researchers and mission planners.

Background: The Three-Body Problem and Libration Points 3,4,7

A standard simplified approach for initial investigations into interplanetary trajectories relies on reducing a complex“four-body” problem (Sun, Earth, Planet, and Spacecraft) down to three phases of simplified “two-body” problems(Primary body and Spacecraft). These are usually identified as the departure, cruise, and arrival phases, and thisapproximate solution approach is typically called the “patched conic approach.” A general picture of this is depicted inFigure 1 for a transfer from Earth to another planet (in this case, an inner planet and not Mars). The “spheres ofinfluence” mark the boundaries for the phases of the trajectories.

Figure 1. Spheres of Influence for Patched Conic Trajectories

To consider trajectories to and orbits about the libration points, one must address the “three body problem,” whichinvolves two primary masses and a much smaller third mass (the spacecraft). There is some history behind the search

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for analytic solutions to this problem involving some rather well known figures. Newton tackled some aspects of thisproblem when he computed the orbit of the moon to within 8% in 1687. Euler developed his problem of “two fixedforce centers” in 1760 and the rotating / synodic coordinate system in 1772. Jacobi created his integral solutions fromthe Euler “restricted” three-body system in 1772. That same year, Lagrange identified the “equilibrium” points of arestricted three-body system. These points (there are five in all) have since been labeled “Lagrange points” and are alsoknown as “libration points.” Orbits about these points in space are termed “libration orbits” or “Lissajous orbits.” Aspecific type of Lissajous orbit is often called a “halo orbit.”

The five libration points are defined in a rotating coordinate frame in which 2 large primary bodies rotate about theircommon center of mass. Looking “down” upon the plane of rotation of these two bodies, these points can be labeledas in Figure 2.

Figure 2. Geometry of the Lagrange Points of Two Primary Masses P1 and P27

The resulting motion and force balancing in this synodic frame produces the five equilibrium points, where a third bodyof small relative mass would theoretically remain once placed there. The points numbered 1, 2, and 3 are termed thecolinear points since they lie on the line connecting the two primary bodies, and are unstable in the sense that an objectplace there will eventually depart due to the unstable nature and disturbance forces. Points 4 and 5 are the triangularpoints since lines connecting them to the primaries form equilateral triangles, and are stable. A way to envision thissystem of points is that of energy balancing, where the colinear points represents balancing on the top of a “peak” andthe triangular points represent balancing in the trough of a “valley.” Fairly stable use of the colinear points can beachieved by placing an object in orbit about the points. It is also interesting to note that the existence of theseequilibrium points, as predicted by Lagrange in 1772, were finally confirmed 134 years later by the discovery of the so-called “Trojan” asteroids, which are objects that have collected near the Sun-Jupiter triangular points.

A Note on Historical Missions Using Libration PointsIt is important to note that there have been successful space missions utilizing libration point orbits. The first proposalfor such a mission was in 1966 when John Breakwell and Robert Farquhar proposed that a satellite orbiting about theEarth-Moon L2 point could provide a communications link to support exploration of the “dark side” of the moon (moreaccurately referred to as the “far side”). They developed a periodic, out of plane solution for the spacecraft orbit thatled to the development of halo orbits.

The first true mission into one of these obits was the International Sun-Earth Explorer-3 (ISEE-3), launched in 1978 foroperations about the Sun-Earth L1 point. Its trajectories are represented in Figure 3 and are indicative of some of thecomplex dynamics involved in these types of missions.

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Figure 3. ISEE-3 Trajectory Summary

Another successful mission of note was the Solar and Heliospheric Observatory (SOHO) launched in 1995, which providedan unobstructed view of the Sun from its orbit about the Sun-Earth L1 point. The schematic is shown in Figure 4.

Figure 4. SOHO Trajectory Summary

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There have of course been other missions since these times, but these represent some of the first successes and havepaved the way for expanding upon the use of these types of orbits. There have been no missions associated with theSun-Mars libration points to date.

Mars Communication ConstellationsCurrently, spacecraft missions to Mars rely on their on-board equipment to provide faint transmissions directly to Earth orto a relay in Martian orbit. The addition of on-board communication equipment capable of reaching out through theinterplanetary void between Mars and the Earth adds weight, cost, and risk to missions that operate within tight margins inthese areas. To exacerbate this problem, once a lander has made it safely to the surface, it can only relay information toEarth when it is in direct line of sight. With a Martian day of just over 24 hours, there exist well over 12 hours of “blackout”where no signal can be sent to the Earth. With the addition of an orbiting relay, these times are reduced substantially butsignificant blackouts will still exist. In order to provide continuous coverage for the entire Martian surface, a minimum offour satellites (in elliptical orbits) are required.6 This, once again, raises the issues of cost and risk.

Some of these problems can be solved by the use of a communication network around Mars that takes advantage of thegeometry provided by placement at the Sun-Mars Lagrange points. A minimum of two satellites located at the Sun-Mars L1 and L2 points could provide near continuous coverage for multiple vehicles on the surface and in orbit.6 Thesepoints are termed the co-linear, near-Mars, Sun-Mars Libration points and are defined as above in a three-body orbitalsystem where the two primaries are the Sun and Mars and the much smaller body is the orbiting communication relay.Note also that the Lagrange points remain at the same relative locations as the two primary bodies rotate about theircenter of mass. The communication satellites would be inserted into large amplitude orbits about the L1 and L2 points,circling their respective Lagrange points and the Sun-Mars line. These satellites could communicate with landersanywhere on the Martian surface, with any spacecraft in Martian orbit, and provide the critical communications linkbetween the Earth and Mars.6

Other constellations could be used for a Mars communications network, but each has disadvantages that outweigh theadvantages.16 A group (four to six) of low to medium orbiting relay satellites would ensure that every satellite wouldcover the entire planet at some point, but the cost and risk of inserting so many satellites and the limited instantaneousfield of view the satellites can offer do not make it an attractive option. Four satellites in common-period, inclinedorbits, or a Draim constellation, could cover the entire surface of Mars, but again require twice as many satellites as theLissajous orbit concept, as well as the added complexity of a ground station continuously switching from one satelliteto another. An aerosynchronous constellation (like Earth geosynchronous but at Mars, approximately 20,462 kmaltitude) requires three or four satellites, and works well with ground stations that can simply point to one spot in thesky. However, in addition to the fact more than two satellites are required, there is virtually no polar coverage. Anotherproposal places communication landers on the Martian moons of Phobos and/or Deimos, but this constellation has thesame inefficiencies as the aerosynchronous satellites with large gaps of polar coverage.

The L1 and L2 orbit constellation requires only two satellites for a fully operational constellation (each spacecraft seesalmost half of Mars at all times), thus making it the most attractive option. The Sun is always visible to both satellites,greatly simplifying power requirements for that spacecraft. Lander pointing requirements are simple, given thespacecraft is always the same relative distance from the Sun-Mars line, and the spacecraft station-keeping budget isrelatively small. Disadvantages are overcoming the approximate one million kilometer distance from the Lagrangepoints to the Martian surface. This distance likely requires a large, high frequency antenna, which could complicatesolar panel design to minimize antenna shadow and may require a more complex lander communication system tointeract with the high frequency signals. Interference from constant solar radiation along the Sun-Mars line and for aparticular Earth viewing geometry may also have to be considered, and the loss of one satellite means half the planetloses communications coverage for approximately 12 hours.

Please note that this study does not attempt to address any issues related to communication link specifications orperformance, but does focus on the trajectory designs to place communication relay vehicles into their mission orbits.

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Sun-Mars Libration Point Orbits

Figure 5. Sun-Mars L1 and L2 Halo Orbit Constellation6

The design of the halo orbit communication network around Mars is a simple but elegant one, as depicted in Figure 5.6In order to prevent an unneeded overlap of coverage, the orbits of each satellite at L1 and L2 would be opposed by 180degrees but moving in the same orbital direction. Herein lies a minor problem with this configuration: because L1 andL2 are at finite distances from Mars (1x106 km), the actual view of Mars is slightly less than hemispherical. Despitethis geometry, the original study predicted that the network would be able to view 99.81% of the planet at all times. The“down time” in this scenario would be minimal: a vehicle caught in this band would have to wait a mere 1.5 minutesbefore coverage would be switched over and reestablished with the other satellite.

In designing the proper orbits in which to place the two satellites, the most important consideration is that they permitefficient maintenance of the 180 degree offset.6 An additional consideration is that of avoiding having the satellite crosswhat is known as the “solar exclusion zone,” the line between Mars and the Sun. Passing through this zone,communications would be disrupted due to intense solar interference. To avoid this problem the orbit must be largeenough to avoid this crossing; an orbit of period greater than 0.9 years should suffice. Another obvious considerationis the choice of geometry and size of the orbit that reduces the required insertion maneuvers, and thus cost, from Earth.

As an aside, one might wonder if the L4 and L5 points could play some role in the design of a communication networkaround Mars. The L4 and L5 Lagrange points lead and trail Mars by 60 degrees in its orbit, thus forming equilateraltriangles with Mars and the Sun (see Figure 2). The distance from Mars to either of these two Lagrange points is thesame as the distance from Mars to the Sun (227.9 x 106 km). To communicate over these distances, currentinterplanetary missions use very large dishes, such as the Goldstone Deep Space Network (DSN) facility in California,in order to eliminate the need for large, powerful transceivers on the spacecraft itself. Links over this 230 millionkilometer range would require space borne communications elements whose size, weight, and power would be on theorder of a DSN ground station. The size and power of the needed equipment for these distances make the L4 and L5unrealistic as locations for the network.8

There is one other interesting aspect of the L4 and L5 points worth mentioning here. While their stability can beexploited for use in missions that require minimal station-keeping, this same stability also attracts a multitude ofinterplanetary bodies that populate this region of the solar system. These special bodies are known as “Trojans” becausethe first few such objects discovered were named for several heroes from the Trojan War. By convention established bythe International Astronomical Union, all similar objects must be named after Trojan War heroes, Greeks ahead of theplanet and Trojans trailing the planet. Two of the larger Martian Trojans (in the 1-2 km range), 5261 Eureka and 1998VF31, represent what could be thousands of other bodies that reside at the L4 and L5 points making these fairlydangerous places indeed. Consideration must be made as to whether the benefits of the inherent stability of the L4 andL5 points outweigh the risks of residing there.8

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Mars L1 and L2 Constellation AdvantagesProbably the most significant advantage to using the large amplitude Lissajous orbit constellation is minimum costassociated with only 2 spacecraft required, when compared to other constellation options.15,16 Additionally, spacecraftorbiting about L1 and L2 can readily see the Sun and Earth, potentially simplifying spacecraft solar cell placement andcommunication antenna design.

The vehicles circling the Sun-Mars L1 and L2 points will orbit the Sun-Mars line with periods on the order of 1 year.The long period dynamics of such orbits may make them attractive for interplanetary missions with significantcommunication time delays. Perhaps even more significantly, on a given day (or series of days) the relay vehicle willappear motionless and remain in a fixed position relative to the sun (or local midnight vector, for the L2 relay). Thus,a communication relay tracking system for explorers on the surface could be simplified and automated for tracking ofthis position in the sky. From the surface, there would be only one switch between relay vehicles each day, a distinctadvantage over Draim or low orbiting concepts. The Sun-line geometry may also allow for simplified and robust safe-modes for the vehicles based on the sun vector. Since both spacecraft orbit about the Sun-Mars line in large amplitudeLissajous orbits and avoid eclipsing, the relays would always have access to the sun for their solar cells, thus allowingreduced battery sizes.

An additional benefit of the unique geometry offered by Lissajous orbit missions is a secondary mission for thesecommunications relay vehicles as observation platforms. The L1 satellite is able to perform continuous solar activitymonitoring via a secondary payload on the vehicle, and thus provide advance warning of activity to Mars surfacemissions. This could be done using low power, simple instruments for simple early warning of solar storms / flares,perhaps derived from legacy missions. Regular monitoring of the sun is thus possible, and can additionally be comparedto solar data from sensors closer to the sun. This secondary mission for solar activity monitoring increases in importancewhen the Earth is on the opposite side of the solar system from Mars. Another observation mission for both spacecraftcould include Martian weather sensing and relay to Mars expeditions and Earth. The L2 vehicle offers the opportunityof a secondary payload for asteroid and outer solar system observations.

One of the fundamental advantages of Lissajous orbit is, of course, the relatively small ∆V maneuvers required forstation-keeping, when compared to other mission orbits. Annual ∆Vs for each vehicle could to be on the order of 2m/s14 and studies have produced data showing annual vehicle station-keeping estimates of 50 m/s for low orbits, almost200 m/s for aerosynchronous, and 30 m/s for the inclined common period missions.15 We provide more discussion ofstation-keeping in a later section of this report.

Mars Mission Simulations and AnalysisSatellite Tool Kit (STK) / Astrogator26

STK / Astrogator was the simulation and analysis tool used to generate the data for this study of utilizing the Sun-Marslibration point for orbiting communication relays. This is an interactive orbit maneuver and space mission planning toolthat is fully integrated within the larger STK tool and is widely used for Earth orbiting, Lunar, libration point, andinterplanetary missions. The key features of this tool are the use of user-defined gravity fields, propagators, and coordinatesystems. It also uses a powerful “targeted trajectory” design process and features a variety of on-line help features.

On a historical note, Astrogator has its roots in a tool called Swingby, which was developed by Computer SciencesCorporation (CSC) in 1989 for NASA Goddard Spaceflight Center (GSFS). This was commercialized as Navigator byCSC in 1994. In turn, the Navigator tool was purchased by Analytical Graphics, Inc. (AGI), the developers of STK. Inresponse to GSFS requests for COTS products, Astrogator was developed in 1997 using some heritage algorithms fromSwingby and Navigator.

For designing libration orbit missions, there are some specific tools within the Astrogator module which are employed.One of these is the rotating libration point (RLP) coordinate system. This is defined for a system of primary and less-massive secondary gravitating bodies as shown in Figure 6.

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Figure 6. RLP Coordinate System26

The origin is placed at the libration point of interest. The x-axis defines the line from the primary to the secondary body(in this specific case, from the Sun to Mars). The y-axis is orthogonal to the x-axis in the plane and direction of thesecondary body’s motion about the primary. The z-axis completes the orthogonal set to x and y.

For velocity change (∆V) computations, the velocity-normal-conormal (VNC) frame is used (see Figure 7). In thisframe, the x-axis is along the velocity vector of the spacecraft, the y-axis is along the orbit normal (radius cross velocity),and the z-axis completes the orthogonal triad.

Figure 7. VNC Coordinates

The real computational strength within Astrogator lies in the targeting sequence, which allows the definition of maneuversin terms of goals to achieve. The basic targeting problem is: given a set of orbital goals, how can the control parametersbe perturbed (and solved) to meet them? Astrogator uses a differential corrector process to iterate to a solution.Determining libration orbit insertion trajectories often requires multiple phased targeting segments where the goals definethe intended orbits and the control parameters are the velocity change components and other transfer orbit characteristics.

2003 Direct Insertion into L1 Large Amplitude Lissajous OrbitThe original study by Pernicka, et al, was a system level analysis with some simplifying assumptions. It used co-planertransfer trajectories and circular orbits to determine C3 energy and orbit insertion ∆V requirements for a 2003 Sun-MarsLissajous communications relay concept. (Note: C3 is defined as negative the gravitational parameter of the centralbody divided by the semi-major axis. For hyperbolic orbits this is the square of the hyperbolic excess velocity.) Thatoriginal data was generated with simplified force models and direct transfer to an orbit about L1. Our study used STK/ Astrogator and the targeting process described in detail in the subsequent section to reproduce a subset of this data forthree different times of flight (TOF) for closer analysis. No Z amplitude specifications were considered for thesescenarios. A screen capture of the trajectory from the STK output is shown in Figure 8, where the view is of the XYplane looking in the –Z direction, using a Sun-Mars rotating coordinate frame. We found some agreement between theresulting data, which is presented in the table that follows.

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Figure 8. L1 Orbit Direct Insertion

Table 1. Summary of Direct Transfer to L1 Lissajous; 13 Jun 03 Departure

The differences in C3 energy and ∆V are most likely due to the differences in the model parameters of the studies. Thisstudy used full force models, non-circular and non-coplaner transfers, and direct computation of the orbit insertion ∆V.Based on inspection of the Mars arrival trajectories, use of actual non-coplaner, eccentric planetary orbits seems to be amajor factor. The data trends are still evident, however: the 200 day TOF case provides the minimum C3 energy andorbit insertion values (for the cases studied) and the longer and shorter duration flights require more energy and velocitychange. Thus, for a 2003 mission (a baseline comparison year to be compatible with the original study) we provide thisrefined transfer data from our simulations with full-force models and non-coplaner, eccentric orbits. With these missionscenarios developed, more extensive simulation and analysis for various transfer parameters could be undertaken.

2003 Transfer with Braking Maneuver at Mars PeriapsesBased on a helpful suggestion by Chauncey Uphoff, and some historical approaches identified by Dave Dunham, weinvestigated the use of a braking maneuver at close approach to Mars to lower the ∆V required for the Lissajous orbitinsertion maneuver. This added a segment to the trajectory design and required some careful targeting for a closeswingby and braking maneuver around Mars (targeting details are discussed in the next section of the paper). Wemodeled and simulated the 200 day TOF case for the 2003 mission to L1 with the braking maneuver, shown in Figures9a (looking edge-on at the XZ plane) and 9b (looking down on the XY plane). The data and a comparison to the directinsertion case are presented in the table below.

Figures 9a and 9b. L1 Orbit Insertion with Braking Maneuver

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Table 2 Comparison of 2003 Transfers to L1 Orbit; 200 Day TOF

The braking maneuver resulted in a ∆V savings of 1.465 km/sec, which would lead to fuel mass savings and/or increasein payload capacity. This type of maneuver seems promising as a ∆V conserving technique, and so we adopted it forthe other mission simulations that follow. However, there is plainly room for future investigations into the applicabilityof this trajectory for various mission profiles.

2016 Transfer Braking ManeuverIn order to provide data from this study that may aid future mission planners or lead to further research, we modeled a2016 mission to place two vehicles in orbit about L1 and L2. We simulated a 200 day TOF as a baseline, as well as a181 day TOF which, along with the departure date of 20 Feb 2016, was inspired by a JPL Ballistic Earth-Mars Trajectorystudy.17 These trajectories are depicted in Figures 10a and 10b and relevant data is shown in the two tables which follow.

Figures 10a and 10b. L1 and L2 Orbit Insertionwith Braking Maneuver

Table 3. Comparison of 2016 Transfers to L1 Orbit for Different TOF

Table 4. 2016 Transfers to L1 & L2 Orbits for 200 Day TOF

Table 3 shows that a shorter TOF to Mars can be achieved with a lower C3 energy value, but that trajectory requires alarger braking maneuver than the longer transfer, to achieve the same mission orbit. This indicates that with these typesof missions the lower energy transfer may not yield a lower braking and insertion ∆V specification.

Table 4 shows how two vehicles could start on the same transfer trajectory initially (as with a simultaneous launch) andthe L2 vehicle targeted for it’s close approach via a small mid-course correction. The simulation method is explainedfurther in the next section. The total TOF to Lissajous orbit insertion is different for each vehicle, which would assist inthe phasing of the vehicles that is required for the communications relay system to maintain adequate coverage of Mars.

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Z Amplitude and ∆∆V AnalysisSince specific mission performance specifications will drive the Lissajous orbit shaping requirements, we performedsome basic investigations into the relationship between the orbit Z amplitudes and the ∆V needed for the braking andmission orbit insertion maneuvers. We examined the 2016 transfer to an L1 orbit for 200 day TOF, targeting variousamplitude values. The trajectories were very sensitive to small changes in the maneuvers (see Figure 11, lookingtowards the Sun), and the results are summarized in Table 5 below. The C3 energy for all cases was kept constant at10.377 km2/sec2.

Figure 11. Orbits about L1 with Different Z Amplitudes

Table 5. 2016 Transfer to L1 Orbit with Varying Z Amplitude

The data from Table 5 demonstrate a correlation of the geometry of periapsis with the Z-amplitude. As a measure ofthe geometry, the elevation angle of the periapsis measured with respect to Mars’ orbit plane was used. Very slightchanges in the elevation angle caused dramatic changes in the Z amplitude. (It was also noticed that the class of theLissajous orbit could be changed by large variation of elevation angle, however this was not thoroughly investigated forthis study.) As shown in the table, the mid-course correction ∆V to change the elevation angle at periapsis isinsignificant. Additionally, there is no significant change in the braking maneuver, leading to the conclusion that a widerange of Z amplitudes can be achieved with no fuel penalty.

Relative Phasing SelectionAs noted previously (see Figure 5), the key to obtaining sufficient communication coverage is to achieve 180 degreephasing of the two vehicles in their orbits. In other words, one is north of the Mars orbit plane while the other is south,and one is leading Mars while the other trails. The two-vehicle simulation from the sections above does not attempt toachieve the proper phasing for actual relay mission operations. This 180 degree phasing can be achieved by causingboth spacecraft to reach their appropriate libration orbit insertion (LOI) points at the same time. In the above resultswith no phasing control, the “baseline scenario” shown in Figure 10 had an LOI time difference of 56 days.

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Of course, each vehicle could be launched separately to achieve the proper phasing, but with very constrained launchwindows. The redundant launch costs may also make that approach cost-prohibitive. This paper focuses on scenarioswhere both spacecraft are launched on the same launch vehicle. As a result, the relative phasing of the spacecraft intheir Lissajous orbits is controlled by their onboard propulsion to affect LOI insertion time. Three possible methodswere investigated: using a midcourse maneuver (as above), adjustment of TOF from periapsis Mars to LOI, and the useof a Martian phasing loop. All are discussed in the next section.

Trajectories to Achieve Two-Vehicle PhasingThe first method investigated to control the relative phasing was to use the mid course correction (MCC) maneuver 30days after launch to change the time of arrival of the spacecraft at Mars. By adjusting the time of arrival, the time ofinsertion into the Lissajous orbit (the LOI maneuver) would also be changed.

The baseline scenario in Figure 10 shows that the trajectories arriving at Mars are not symmetric with respect to the Sun-Mars rotating coordinate system; the incoming trajectories arrive from the L1 side of Mars. A consequence of this isthat the L1 spacecraft inserts before the L2 spacecraft reaches its insertion point. A first step in getting both spacecraftto arrive at their LOI points simultaneously was to adjust the L2 spacecraft trajectory so that it would arrive at theperiapsis Mars point earlier then the L1 spacecraft. To do this, however, required a very large midcourse maneuver. Inaddition, the decreased time of flight caused the incoming velocity at Mars to increase, and changed the direction of theincoming asymptote, as shown in Figure 12 below. In that figure, the dashed line is the original baseline and the solidrepresents an earlier arrival at periapsis Mars.

Figure 12. Mid-Course Maneuver Allows Earlier Periapsis Mars (solid line)

The change in the asymptote angle in turn caused the time of flight from periapsis Mars to LOI to increase, and the epochof the LOI point did not vary the same as the change in periapsis Mars epoch. In fact, for the case examined, an earlierperiapsis epoch resulted in a two day delay in LOI. In addition, the retrograde braking maneuver and the LOI ∆V costsincreased, again because of the change of the transfer trajectories. Table 6 shows the comparison data and the increasein ∆V. This method proved unfeasible because of the large ∆V cost associated with moving the epoch of LOI even afew days, and the direction of that movement for this case.

Table 6. Using MCC to Change Periapsis Date and Effect LOI Epoch

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The second method investigated to control phasing was to vary the time of flight of both vehicles from periapsis Marsto LOI. The time of flight is correlated with the amplitude, which is controlled by targeted B-dot-R value.22 Thus, thisTOF must be controlled to cause the L1 vehicle to insert later and have the L2 vehicle insert earlier so that their LOItimes coincide. Our investigation results, detailed in the figures and Table 7 below, show that varying this TOF resultsin very large (and most likely unwanted) Lissajous orbit amplitudes (over 500,000 km) without achieving thesynchronicity of the LOI required for the system phasing. However, there may be some applicability of this method toaffect small changes in phasing if needed during operations.

Figure 13a. Amplitude Variations Due to TOF Changes

Figure 13c (above). Amplitudes About L1Figure 13b (left). Periapsis Mars

The third method to control phasing is the use of a phasing loop orbit about Mars prior to LOI. With this approach, theL2 vehicle performs its LOI maneuver after Mars swingby as in the baseline case and establishes the LOI time andphasing that the L1 vehicle must match. Using a retrograde capture maneuver at Mars periapsis, the L1 vehicle entersa phasing orbit about Mars. After one revolution in this orbit, a subsequent maneuver at periapsis transfers the vehicleout to the LOI point. The period of this phasing orbit summed with the time of flight of the transfer to LOI must besuch that the desired epoch at LOI is achieved.

One might think that the phasing orbit period would be equal to the difference between LOI times for the L1 and L2vehicles in the baseline configuration (56 days). However, since the phasing orbit periapsis point rotates in the Sun-Mars rotating frame, the transfer to LOI is longer than in the baseline scenario. Thus, to obtain the correct total TOFthe phasing orbit period is shorter than expected. In fact, using a two-body orbit period calculation right after the capturemaneuver set equal to the 56 day time difference actually causes a time delay far in excess of that required. The data

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and figures below show these results. The use of a phasing orbit also introduces some flexibility into the execution ofthe entire mission. The result is that the 180 degree phasing of the two vehicles can be obtained. The results are below.

Table 7. Z-axis amplitude and LOI date as a function of Time of Flight (TOF)

Table 8. Phasing Loop Used to Equate LOI Epochs

Figure 14a. Phasing Loop Trajectory(Solid) Figure 14b. Periapsis View

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Estimation of Communication Coverage AchievedPrevious work has indicated that the 180 degree offset phasing of this communication system will allow near continuouscoverage of the Martian surface so that exploration missions there would only experience communications loss for a fewminutes twice a day, at the most. In order to verify and uncover more details on this, we analyzed the baseline 180degree phased trajectories using the STK / Coverage module.

Two metrics were used to quantify the quality of coverage: “Maximum Revisit Time” and “Number of Gaps.” Gaps arethe times on the surface of Mars that are not within line-of-sight of either satellite. Maximum Revisit Time is definedas the maximum duration of the gap in coverage over the entire coverage interval, which starts at LOI and goes for 674Earth days (over one Martian year). “Number of Gaps” are the number of times in this same interval that contact cannotbe made with at least one of the satellites. This was measured along one longitude line at 10 degree intervals of latitude.

This analysis approach takes into account the rotation of Mars and its tilt, using fully integrated trajectories phased viathe method in the proceeding section. Table 9 shows tabular results for one particular epoch. For the mid-latituderegions, the gaps in coverage are about a half an hour, with one to two gaps per day. As a special note, the south polarregion had the longest revisit times (6 days), which occurred 4 times during the year. By inspection, it seems that theremay be a seasonal variation that should be investigated because of the effect near the poles. A complete investigationwould include LOI epochs at various times of the Martian year and consider all locations on the planet.

Table 9. Summary of Coverage Analysis

Targeting Methods Using STK / Astrogator1

In this section we provide more detail on using the simulation and analysistools which generated the results of the previous section. The transfer fromthe Earth to a Mars Lagrange orbit was targeted in a series of steps. Thepurpose of the targeting was to determine the control variables necessary toachieve this transfer. The initial orbit state represented the post launch Earth-centered hyperbolic trajectory. This was specified in target vector form, in theEarth-centered mean ecliptic and equinox of J2000 coordinate system. Theseven parameters of the target vector are: epoch, radius of periapsis, C3energy, right ascension (RA) and declination (Dec) of the outgoing hyperbolicasymptote, the velocity azimuth at periapsis, and the true anomaly.

For this study, the epoch was chosen to match previous work, the trueanomaly was set to zero, the velocity azimuth set to 90 degrees, and the radiusof periapsis set to 6678.0 km. This represents a satellite near the Earth atperigee. The remaining parameters, C3 energy and the direction of thetrajectory (RA and Dec of the asymptote) were used as control parameters.Two methods of insertion into Lagrange orbits were utilized and are discussedseparately below.

Direct Transfer to L1 Lagrange OrbitFor the direct transfer from Earth to the L1 Lagrange orbit, the controlparameters were adjusted using a differential corrector technique to achievethree constraints at the point the trajectory crossed the ZX plane of Sun-Marsrotating libration-point coordinate system (this is the plane containing theSun-Mars line and perpendicular to Mar’s orbit plane). The three constraints are the desired epoch of arrival, and theX and Z positions in the Sun-Mars rotating libration-point coordinates. This is shown in the figure below.

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Figure 15. Direct Transfer Targeting

Once the desired time and position was achieved, the three components of the Lagrange-orbit insertion maneuver (LOI)was targeted as an impulsive ∆V maneuver in four steps. First, LOI was targeted to achieve somewhat ideal velocitycomponents for the Lissajous orbit at this point. Velocity in the X and Z rotating libration point directions were targetedto zero. Velocity in the Y direction was targeted to -0.16 km/sec (a representative value from previous analysis).Second, the LOI maneuver was corrected so that after propagating the trajectory a half revolution to the first ZX planecrossing, the X component of velocity (Vx) would be zero (this represents a perpendicular plane crossing whenprojected into the XY plane, and is the same energy balancing technique mentioned by Dunham and Roberts14). Afterachieving the first ZX plane crossing, the third and fourth steps were to correct the LOI maneuver to achieve Vx of zeroat the second, and then the third ZX plane crossings. The figure below illustrates this process.

Figure 16. Direct LOI Targeting

Transfer Using Braking Maneuver at Mars PeriapsisThe transfer to a Mars Lagrange point orbit using a braking maneuver at the close approach at Mars before the LOImaneuver was also targeted in stages. First, the target vector control parameters were adjusted by the differentialcorrector to achieve an epoch at periapsis Mars, and B-Plane components to place the trajectory on the anti-Sun side ofMars. Since this stage is just a first guess, the values used were B-dot-T of -10,000 km, and B-dot-R of 0.0 km. Theentire sequence is shown in the figure and explained further below.

The second step refined this to the desired close approach conditions. Using the same control parameters, the radius ofclose approach was used instead of B-dot-T, and was targeted to a radius of 3,600 km (about 200 kilometers altitude).

After the constraints at periapsis were met, the magnitude of a retrograde braking maneuver (anti-velocity direction) wasused at periapsis to shape the trajectory until the trajectory crossed the XZ plane at the desired X distance in the Sun-Mars rotating libration-point coordinate system (XRLP). After the retrograde maneuver was calculated, the LOImaneuver was planned using the same 4-step method previously described for the direct transfer.

The transfer to the L2 Lagrange orbit was planned in a similar manner, except that the trajectory must pass on the Sun-ward side of Mars at the close approach. This was done using a mid-course correction (MCC) maneuver as a control

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parameter, which also allowed the initial transfer parameters to be the same for both the L1 and L2 vehicles. Someadditional graphic explanation of B-plane targeting is shown in the figures below, and further detail can be found on theSTK web site.26

Figure 17. LOI Using a Braking Maneuver

Figure 18. B-Plane Targeting27

Transfer Using Braking Maneuver to Achieve Desired Z AmplitudeThe concept for targeting a desired Z amplitude for the Lagrange orbit is analogous to the technique using Earth’s moonas described by Sharer, et al.21 The Z amplitude can be directly controlled as a function of the position of the trajectoryas it passes through its close approach to Mars. The process is outlined in the figure and described below.

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Figure 19. The B-Plane27

Figure 20. Targeting Z-Amplitude Variations

The initial targeting is the same B-Plane targeting described above: first B-dot-T and B-dot-R, and then B-dot-R andRadius of periapsis. The initial target vector parameters were not used at this step because the corrections to theparameters were too small, being on the order of a double precision number. Instead, a MCC maneuver was used 30days after Earth departure.

After the epoch, B-plane, and radius of periapsis constraints were achieved, the retrograde braking maneuver wastargeted to achieve as described above. Then the Z distance (amplitude) was checked, and if it was significantly farfrom the desired value, the previous step was repeated with a different B-dot-R value. (B-dot-R is directly related to theelevation of periapsis with respect to the Mars’ orbit plane.)

The next stage involved targeting the four constraints that must be simultaneously met: the epoch at periapsis, the radiusof periapsis (to prevent the trajectory from hitting Mars), the X position at LOI, and the Z amplitude. In addition to thethree components of the MCC maneuver, the magnitude of the braking maneuver was also used as a control.

Once this step converged, the LOI maneuver was targeted using the same 4-step method described above for the direct insertion.

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Transfer to Achieve Relay PhasingThe first two methods attempted to control phasing used the approaches described above. The 3rd method, whichutilized a phasing loop for the L1 satellite, required a modified targeting procedure. The capture maneuver was targetedsuch that after one phasing loop and then subsequent transfer to LOI, the epoch of LOI would match the epoch of theL2 satellite LOI. In order to achieve this, two differential corrector targeting schemes were employed simultaneously.An inner targeter calculated the maneuver magnitude in the velocity direction needed to transfer from the phasing orbitto the Lissajous orbit. This inner targeter was “wrapped” by an outer targeter which calculated the retrograde maneuverat the first Mars periapsis. This targeter was set up to adjust the maneuver magnitude to achieve the epoch at LOI whichoccurs after the inner targeter has converged on a solution. Therefore, each time the outer targeter iterated and searchedfor the capture maneuver, the inner targeter was re-run to calculate the transfer from the phasing orbit to LOI.23

Station-keepingDue to the precarious nature of the Lissajous orbit, precise and continuous station-keeping (SK) techniques must beemployed. Additionally, the precision required for certain missions located around the Sun-Mars Lagrange pointsrequires the fidelity of such SK maneuvers to be extremely high. Historical data and studies indicate that SK ∆Vs aslittle as 1 mm/sec could be required.

Station-keeping techniques fall into two major categories.14 The first, referred to as a “tight” control technique, attemptsto target the vehicle back to a nominal three-dimensional path. The second is the “loose” control technique that uses asimpler “orbital energy balancing” strategy to closely mirror a Lissajous orbit. The two control techniques differ onlyin the number of ∆V components that are varied. The loose technique will simply vary one component of ∆V while thetight technique varies two or more to achieve a nominal Lissajous orbit.

History14,18

The third International Sun-Earth Explorer (ISEE-3) flown to the Sun-Earth L1 point in 1978 used the tight controltechnique in an attempt to maintain its trajectory as close to a nominal halo orbit as possible. This mission, being thefirst to orbit a Sun-Earth libration point, had the luxury of a large supply of fuel to allow for uncertainties in the insertionto and maintenance of the new orbit. The relatively small errors encountered during insertion into the halo orbit left alarge amount of fuel that could be used specifically for station-keeping. Over the four years that ISEE-3 was establishedat the L1 point, 15 SK maneuvers were performed totaling 30.06 m/sec at an average of 2.00 m/sec per maneuver. Thetime between the maneuvers averaged 82 days.

While the large amount of fuel planned for the ISEE-3 mission allowed for very tight control of its halo orbit, a moreoptimal SK method was planned for the Solar Heliospheric Observatory (SOHO). Prior to its establishment at the Sun-Earth L1 point in 1996, SOHO mission planners sought ways in which to decrease its SK costs. If the complexity ofSK maneuvers for SOHO could be dramatically reduced, or “loosely” controlled, the fuel load, and therefore costs,could be also be reduced. In the “orbital energy balancing” technique that evolved, only one component of ∆V, in thiscase the x-component, would be varied. The result of this simplification achieved a threefold reduction in SK costs fromroughly 7.5 m/sec per year for ISEE-3 to less than 2.3 m/sec per year for SOHO.

The major drawback with the loosely controlled technique used on SOHO was that it did not maintain a periodic haloorbit precisely. The resultant orbit was, however, a Lissajous path that mirrored the nominal halo orbit so closely thatfor all practical purposes it could be considered equivalent. The loose control technique was therefore proven as aneffective means of achieving lower SK costs when precise orbit mapping was not necessary.

Station-keeping for the Sun-Mars Lissajous OrbitsFor future missions to Mars using the Sun-Mars Lagrange points, mission planners will have to consider several factorsprior to making a decision on the SK technique to be used. Obviously mission requirements will dictate whether theloose control technique can be used to optimize SK costs or if the higher precision of the tight technique is necessary.

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In our example of a communication system in orbits about the tions system example, the timing of SK burns is criticalso as to prevent unexpected and inconvenient losses in communications coverage to the users on the Martian surface.One solution to such a problem is to overlap the SK maneuvers with the spacecraft’s preplanned attitude and momentumadjustments. This allows the attitude control, momentum management, and SK maneuvers to complement one anotherand minimizes the down time of the system. Dunham and Roberts have shown that for any ∆V error, the SK that followsis also minimized. If, however, the insertion ∆V error is greater than that expected the magnitude of the SK maneuverswill increase and the SK costs will rise.

The second factor in SK frequency determination is the effect that subsequent SK burns have on the overall orbit error.This can be further broken down to the actual magnitude of the orbit error at the end of the last burn, the time since thatburn, and the accuracy of the burn itself as executed. Obviously, this orbit error will increase with time and the larger theerror, the sooner a subsequent burn will need to be performed. The key here is to minimize the magnitudes of the burns.14

Once the frequency and magnitude of the required SK maneuvers are determined, the optimal timing of such maneuverswill need to be considered. For the communications system example, the timing of SK burns is critical so as to preventunexpected and inconvenient losses in communications coverage to the users on the Martian surface. One solution tosuch a problem is to overlap the SK maneuvers with the spacecraft’s preplanned attitude and momentum adjustments.This allows the attitude control, momentum management, and SK maneuvers to complement one another and minimizesthe down time of the system.

Dunham and Roberts have shown that small ∆V errors on order of 0.1 mm/sec for the Sun-Earth / Moon system causenoticeable deviation from the nominal after about three revolutions in the Lissajous orbit.14 In our study, the same errorwas applied to the Sun-Mars L1 Lissajous orbit as shown in Figure 21. This error caused noticeable deviations afteronly one and a half revolutions. However, because the period of the Mars Lissajous is about twice that of the EarthLissajous, the deviations occur approximately after the same duration in time. This is an indicator that the station-keeping requirements for the Mars Lissajous will be on the same order as seen for the Earth missions, in terms of fuelused per year. Of course, a thorough error analysis study could be made later to prove this, accounting for the errorsand uncertainties.

Figure 21. Effect of small errors on Lissajous orbit

Further research for this study explored the station-keeping sensitivities of spacecraft in these orbits via Monte Carloanalysis. The uncertainties included those due to the orbit determination process, possible change in the effective area of thespacecraft affecting the Solar radiation pressure acceleration, and possible errors in the station-keeping maneuver execution.

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Monte Carlo Simulation ApproachThe uncertainties were modeled as uncorrelated errors. The uncertainty magnitudes were: 100 meters in position, 10cm/second in velocity, 10% uncertainty in the area of the spacecraft (which could represent attitude changes), and a ∆Verror of 10 cm/second (These could be attributed to both errors in execution of station-keeping maneuver and attitudethruster control effects). A Monte Carlo simulation was setup to randomly vary these eight parameters and propagatethe baseline L2 trajectory for 90 days. A station-keeping maneuver was then targeted at that time to return the trajectoryto that of a periodic orbit for the remainder of the Martian year. The statistics were gathered on the magnitude of thestation-keeping maneuver required to correct the trajectory.

The results of the Monte Carlo simulation with 100 runs yielded an average station-keeping ∆V magnitude of 0.044 m/s(with standard deviation of 0.003). The same simulation was run for a large amplitude L2 orbit in the earth system (forcomparison) and the average ∆V was 0.45 m/s and standard deviation of 0.03. Since the period of the Earth isapproximately half that of Mars, a second Earth-centered L2 Monte Carlo run was made where the station-keepingmaneuver was done after only 45 days of propagation. The average ∆V was 0.43 m/s with the same standard deviation.These results seem to indicate that the Martian Lissajous orbits require an order of magnitude less station-keeping ∆V thanthose in the Earth system. One possibility for this that was considered is Lunar effects; but the examination of a Lissajousabout L2 of the Sun-Earth system with the moon removed from the model yielded no significant difference in results.There are several other possibilities to be explored, including the different distances from the Sun and planet sizes.

Conclusions22,23

This research began by re-examining a 2-vehicle communication relay system orbiting the Sun-Mars libration points,including transfer orbits, injection strategies, and station-keeping, to see how past studies and data compared to that fromcurrent desktop computing techniques using full-force model targeting and propagation (namely, the Satellite Tool Kit(STK) / Astrogator module). Earth-Mars transfers and Lissajous orbit injections for a 2016 mission were analyzed. Itwas found that trajectory trends from the previous studies were still valid when using full force models, however theactual magnitudes of the maneuvers could increase. This work also highlights the fact that the minimum departure C3energy does not always correspond to the minimum LOI maneuver. Also revealed was that using a braking maneuverat a low altitude (200 km) Mars periapsis prior to LOI saves significant spacecraft on-board fuel, for certain approachtrajectories. One can take advantage of the geometry of this close approach to control the Z amplitude and class of theLissajous orbit as well. It was also determined that the loose control technique for station-keeping could be appropriatefor the L1 and L2 communication relay concept. The stability of these orbits are on the same order as the Sun-Earthorbits in terms of deviations from nominal as a function of time.

Three methods were explored to achieve the 180 degree relative phasing of the spacecraft in their respective Lissajousorbits:

1. Adjusting the time of arrival at Mars periapsis using a midcourse correction;2. Adjusting the time of flight from periapsis Mars to LOI by altering the amplitude of the Lissajous orbit; and3. The addition of a phasing loop before the transfer to L1.

The first method proved too costly in terms of ∆V. The second method did not move the LOI epochs close enoughtogether. The third method was successful, and the targeting algorithm was described.

The quality of coverage was investigated using the fully numerically integrated trajectories and the actual motion ofMars’ polar axis. For most latitudes, the maximum gap was found to be about a half an hour, which is slightly longerthan previous papers suggested, but still within the scope of the missions described. The poles behave somewhatdifferently, with longer gaps, but far fewer. Future work could include investigation of the effect on coverage of orbitphasing with the Martian seasons.

An estimate of the station keeping cost for a Mars L2 orbit was calculated using a Monte Carlo technique, varying theinitial orbit state, area, and maneuver execution errors. This was compared with a similar Earth L2 orbit, and the Mars

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orbit requires about an order of magnitude less ∆V for the maneuver. The reasons behind this are not fully understood,and could be pursued in future work.

Overall, the study has provided useful initial data on the trajectory designs to place vehicles in Lissajous orbits aboutthe Sun-Mars L1 and L2 points, while showing that the innovative use of the two-satellite communication system is apractical concept for Mars exploration.

AcknowledgmentsThe authors wish to thank Hank Pernicka and Chauncey Uphoff for their invaluable assistance with portions of thisstudy. We also thank Bob Farquhar and Dave Dunham of the Applied Physics Laboratory at Johns Hopkins Universityfor their correspondence and discussions on the topic. Thanks go as well to Professor Don Danielson and the SpaceSystems Academic Group at the US Naval Postgraduate School. Roger Martinez of AGI provided key support for theMars Society presentation. We especially recognize Dean Rudy Panholzer at the US Naval Postgraduate School for hissupport of all the conference presentation efforts.

References1. J. Carrico et al., “Rapid Design of Gravity Assist Trajectories,” Proceedings of the ESA Symposium on Spacecraft Flight Dynamics held in

Darmstadt, Germany, 30 September - 4 October 1991 (ESA SP-326, December 1991).2. C.D. Brown, “Spacecraft Mission Design,” AIAA Education Series, 1992, pp. 96-98.3. V.G. Szebehely, “Adventures in Celestial Mechanics,” University of Texas Press, Austin Texas 1989.4. D.A. Vallado, “Fundamentals of Astrodynamics and Applications,” McGraw-Hill, New York, 1997.5. D. Halliday, R. Resnick, J. Walker, “Fundamentals of Physics, Extended,” John Wiley & Sons, New York, 1997.6. H. Pernicka, D. Henry, M Chan, “Use of Halo Orbits to Provide a Communication Link Between Earth and Mars,” AIAA Paper 92-4584, 1992.7. V.G. Szebehely, “Theory of Orbits: The Restricted Problem of Three Bodies,” Academic Press, New York and London, 1967.8. Website: www.users.skynet.be, “The Guide to the Universe.”9. K.C. Howell, “Families of Orbits in the Vicinity of the Colinear Libration Points,” AIAA Astrodynamics Specialist Conference, Paper 98-

4465, August 1998.10. Website: www.map.gsfc.nasa.gov, “Microwave Anisotropy Probe.”11. G. Gomez, K.C. Howell, J. Masdemont, C. Simo, “Station-Keeping Strategies for Translunar Libration Point Orbits,” AIAA Spaceflight

Mechanics Meeting, Paper 98-168, 1998.12. P. Keaton, “A Moon Base/Mars Base Transportation Depot,” Los Alamos National Laboratory, LA-10552-MS, UC-34B, September 1985.13. E. Belbruno and J. Carrico, “Calculation of Weak Stability Boundary Ballistic Lunar Transfer Trajectories,” AIAA/AAS Astrodynamics

Specialist Conference, AIAA Paper 2000-4142, 14-17 August 2000.14. D. Dunham and C. Roberts, “Stationkeeping Techniques For Libration-Point Satellites,” AIAA/AAS Astrodynamics Specialist Conference,

AIAA Paper 98-4466, 10-12 August 1998.15. W. Kok-Fai Tai, “Mars Communication Network Design Trade Study,” Master of Science Thesis, San Jose State University, 199816. M. Danehy, “Martian Communications Network Design Trade Study,” Master of Science Thesis, San Jose State University, December, 1997.17. S. Matousec and A. Sergeyevsky, “Feasible Ballistic Earth to Mars Trajectories from 2002 to 2020,” chart for Mars Surveyor Program

Advanced Missions Studies Office, March, 1998.18. R. Farquhar, “The Flight of ISEE-3/ICE: Origins, Mission History, and a Legacy,” AIAA Paper 98-4464, August 1998.19. R. Farquhar, “Stationkeeping in the Vicinity of Collinear Libration Points with an Application to a Lunar Communications Problem,” AAS

Science and Technology Series: Space Flight Mechanics Specialist Symposium, Vol. 11, pp. 519-535 (Presented in Denver, CO, July 1966).20. R. Farquhar, et al, “Trajectories and Orbital Maneuvers for the First Libration-Point Satellite,” Journal of Guidance and Control, Vol. 3, No.

6, Nov-Dec, 1980, pp. 549-554.21. P. Sharer, J. Zsoldos, and D. Folta, “Control of Libration Point Orbits Using Lunar Gravity-Assisted Transfer,” AAS-93-295, Spaceflight

Dynamics 1993, Vol. 84, Part 1: Advances in the Astronautical Sciences, Proceedings of AAS/NASA symposium, April, 1993.22. J. Strizzi, J. Kutrieb, P. Damphousse, and J. Carrico, “Sun-Mars Libration Points and Mars Mission Simulations,” AAS-01-159, February 2001.23. J. Carrico, J. Strizzi, J. Kutrieb, and P. Damphousse, “Trajectory Sensitivities for Sun-Mars Libration Point Missions,” AAS-01-327, July 2001.24. R. Farquhar, “Future Missions for Libration-point Satellites,” Astronautics and Aeronautics, May 1969.25. R. Farquhar, “Lunar Communications with Libration-Point Satellites,” Journal of Spacecraft and Rockets, 1967.26. Website: www.stk.com27. NASA, “Mission Analysis and Design Tool (Swingby): Mathematical Principles,” Rev 1, Sept 1995 (Draft), Sect 4.4.1.Various on-line resources for some figure art.

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Martian Emergency Medical Crisis Management

C. Marsh Cuttino, MD[1999]

IntroductionAccidents and injuries happen. Frequently.

Despite superior training and caution, despite careful preparation, injuries and accidents occur which have significantimpact upon our lives. In the everyday world, when an accident happens, people are taken to the emergency departmentfor treatment, usually by an ambulance with trained emergency responders. This combination of emergency medicaltechnicians, paramedics, emergency medicine physicians and trauma surgeons provide highly skilled medical care tovictims of accidents and injuries.

Distance and time separate people in space from prompt medical care. When the first manned crew goes to Mars theywill be the most isolated people in history. Prompt emergency medical care from outside sources will be impossible toobtain. The crew will have to be self-sufficient and provide for their medical care.

Medical EmergenciesNASA and other sources estimates that in a prolonged mission such as the projected Mars missions there will be almostcertainly an injury, severe enough to require a visit to an emergency department if it had occurred on Earth.1 Therefore thequestion should not be “How do we manage a medical emergency IF it occurs on a Mars mission,” but “How do we managea medical emergency WHEN it occurs on a Mars mission?” There are several different methods and models to choose.

Should the problem be handled from Earth by telemedicine, or should the crew manage the crisis, with later assistancefrom Earth.2 Should a physician be a member of the crew, or could the crew be trained adequately in emergency medicaltreatment and use checklists.

The technique most often cited is the use of telemedicine to provide expert medical care in space.3 This is the methodof choice for the International Space Station and Earth orbit operations.4 This allows specialists on Earth to examineand direct medical care to a remote location, while allowing the participants on site to have a lower level of training.5The technology is rapidly expanding in this field, including remote sensing, haptic force feedback technology, andtelepresent surgical techniques.6 The primary limitation is time delay for transmissions between the subject andoperator. This becomes a major factor on a Mars mission as the round trip transmission time from Mars to the Earthand back is about forty minutes. There is a two month period where transmission between Earth and Mars is not possiblewhen they are on opposite sides of the Sun, but this could be avoided by placing a satellite relay system to bounce signalsaround the Sun.7

For the majority of medical tasks, such as routine health surveillance, data collection, and minor medical care, this timedelay is acceptable. The time delay will only be a factor in an acute medical emergency. Emergency medical problemsmust be handled in a prompt manner. Emergency Medicine and Trauma Surgery recognize that there is a “Golden Hour”in which treatment and patient stabilization must be performed to ensure patient survival. Delaying emergent medicaltreatment past an hour will often prove fatal. This time limit requires a high degree of autonomy by the astronaut crew.They must be self sufficient to the extreme, and must have a high degree of training.

Why Send A Physician?Should a physician be sent on the Mars crew? Why should a physician be sent, taking a spot that would otherwise go toa geologist or other scientist? The answer is yes, a physician should be sent on the Mars Crew. The reason is simplythat the cost of not having a highly trained individual available to treat acute illness and injury is simply too high. The

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C. Marsh Cuttino, MD; Assistant Professor, Medical College of Virginia – Virginia Commonwealth University,Department of Emergency Medicine, Main Hospital, G-503, 401 North 12th Street, P.O. Box 980401, Richmond, VA 23298-0401;

Email: [email protected]; Web site: www.EmergencyMedOnline.com

loss of the function of a crew member would severely affect the crew’s ability to complete their mission functions, andcould affect the survival of the entire crew.

It is possible to invent multiple scenarios where the entire crew is injured or killed. Some opponents to having a crewmember as a physician cite this as a reason to not have a physician on board. “If anything happens” it is said, “the entirecrew will perish, why bother wasting a crew spot on a physician if they will not be able to help?” This is a fallacy. Itis more likely that an injury will occur to an individual. Accidents that happen to the entire crew may occur, but thisdoes not mean that there should not be a physician to treat and minimize injuries on the crew.

The loss of a single crew member would mean the loss of research and exploration time and capabilities. Loss of thepilot could complicate or prevent safe return of the entire crew. A severely injured crew member would require that theother members of the crew spend additional time from their schedules in direct medical care of that crewman, thusremoving a second or even a third crew member from a large portion of their daily activities. This would have a majornegative impact upon mission capabilities. The public relations fallout from a preventable death would be enormous,and could hurt further missions.

Cross TrainingPhysicians often have varied scientific backgrounds prior to their training in medical school. In the United Statesphysicians must have completed an undergraduate degree prior to matriculation to medical school. Many physicianshave backgrounds in rigorous scientific fields such as chemistry, physics, biology and engineering. Degrees in thesefields are required to be competitive for admission to medical school. It would be simpler to take a physician with sucha background and provide them with the necessary training to be a field scientist then it would be to take a field scientistand make them a physician.

It takes a minimum of seven years to become a general physician, and training in surgical specialties may take up totwelve years. By choosing a physician who has an undergraduate degree in science, field training will be shortened toan acceptable level in preparation for a Mars exploration mission. This means that the doctor on board will not be anextraneous “Ship’s Doctor” waiting around for an emergency to happen, but will instead be a productive, fullyfunctioning member of the crew.

Types of EmergenciesThere are three basic types of accidents with injuries:

1. Those that will kill you no matter what you do2. Those that will get better no matter what you do3. Those that will get better only with prompt treatment

In any given situation the type of injury is essentially random. For serious injuries, some will kill the personimmediately, some will not be life threatening, but will require treatment, and some will go either way. The third set iswhere emergency training can have the greatest benefit, but emergency training can also reduce the severity of injuryand shorten recovery time in the second set as well. Therefore in the majority of injuries the presence of a physicianwill make a significant difference in the acute management of the patient, and can make a life or death difference.

The actual types of injuries that could be expected on a manned mission to Mars include a litany too long to be exploredin detail in this paper. They include traumatic injuries such as falls, blunt trauma, fires, electrical injuries,decompression injuries and injuries from equipment failure. Psychological factors and inattention can contribute toerrors and accidents.8 There are unknown environmental injuries, such as the effects of accidental exposure to theMartian atmosphere. This could result in hypothermia, toxic side effects, and pulmonary damage from particulateinhalation, and hypobaric injuries. There are numerous dangerous toxins and explosive items carried onboard thespacecraft.9 There are inherent medical problems involved with long-term space flight and microgravity that havedetrimental consequences such as calcium bone loss, autonomic instability, and muscle atrophy10 which could increase

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the risk of accidents and increase the severity.11 Prolonged microgravity may increase the risk of dangerous cardiacarrhythmias12 and kidney stones.13 There are risks from infectious disease organisms on board the spacecraft14 and thedangers of the crew’s weakening immune systems. There are numerous difficulties inherent in predicting what type ofaccidents will occur, because if you could predict them then you would be able to prevent them.15 Therefore the bestresponse is to be prepared for the worst possible event that can happen. By taking care of the downside, the upside willtake care of itself.

Who Should Go?Because medical care that is not time dependent can be taken care of from Earth, a physician chosen for a Mars missionshould be from a field that deals with emergencies. Two types of physicians primarily deal with emergencies. The firstare Trauma Surgeons, who specialize in the care of patients with blunt and penetrating traumatic injuries. The second,Emergency Medicine Physicians, are a type of physician who work in the emergency department, and are experiencedin handling all types of emergencies, including toxic exposures, medical and cardiac emergencies, trauma, andenvironmental injuries.

The Emergency Medicine specialist may provide more of an advantage on the expedition. The emergency medicinephysician specializes in resuscitation and management of all types of emergencies. The emergency physician works in achaotic and unpredictable environment in the emergency department. Trauma surgeons focus on the surgical managementof blunt and penetrating injuries. This is a more focused approach then the generalist approach of the emergencyphysician. The emergency physician may be better equipped to treat injuries that do not require emergent surgery.

Patients who have injuries severe enough to require emergent surgery have a much higher morbidity and mortality rate.Even with proper treatment and intervention there is a high risk of death. In the final risk / benefit analysis it may seemthat if the patient is sick enough to need a trauma surgeon their survivability is so low it may not be worth having thephysician on board. For the severely injured but not immediately fatal patient, an emergency physician trained in abroader range of emergencies would seem to be a more appropriate choice to provide “more bang for the buck.”

Physician vs. Crew Medical OfficerThe next question often asked is “Why take a physician when you can train a pilot / geologist / chemist / engineer toprovide medical care?”16 The best answer is that specialists do better then non-specialists.

Prompt recognition and treatment of emergency conditions allows for improved treatment. Operational experience inhandling emergencies provides a background that would be difficult to obtain during the heavy training schedulerequired for a Mars crew.

To demonstrate the possible impact of varying levels of training to an emergency situation an experiment was performedusing the human patient simulator.

Emergency Medical Crisis ManagementThe design of the experiment was to compare the performance of three levels of care providers. To simulate theresponse of individuals trained at various levels, and their response to emergency medical crises a series of simulationswas performed.

Three groups were utilized. The first group consisted of Emergency Medicine Attendants. These individuals havecompleted all specialty training and are Board Certified by the national specialty boards. This is the highest level ofmedical training available. The second group consisted of Paramedics and Emergency Medical technicians. These areindividuals who daily handle the initial care of all types of emergency problems. The final group consisted of third yearmedical students. These students have completed two years of classroom training, and are involved in the clinical aspectof their medical training, which is an additional two years.

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Three standardized emergencies deemed likely to occur during a manned mission to Mars were created and implementedusing the Eagle patient simulator. This is a complex, sophisticated mannequin that very accurately represents humanresponse to injury. All encounters were videotaped and graded on a standardized system such as is used for AdvancedTrauma Life Support Training and Emergency Medicine oral examinations.

The three scenarios chosen involved a single astronaut and three crew responders. The crew was informed that theycould ask “Mission Control” for any information or suggestions, but all responses would be delayed by 20 minutes. Thecrew received a short 5-minute briefing of the scenario and then was sent to the room containing the simulator.

The first scenario involved a trauma scenario in which the astronaut develops a pneumothorax and internal bleedingleading to shock (see appendix for further information). The second scenario involved a rapid decompression withembolism, hypothermia and neurologic deficit. The third scenario was an exposure to hydrazine resulting in respiratorydistress, stridor and hypoxia.

All settings, key actions, and results were determined prior to crew testing. The results were tallied based on analysisof the videotape. Scoring was based on completion of critical actions and time to complete the scenario. Delay due tocontacting mission control for guidance was noted.

ResultsAs expected the attending physicians performed at a significantly superior level. All critical actions were rapidlyperformed, there were no adverse events, and mission control was not contacted prior to disposition.

The paramedics and EMT’s performed at an intermediate level. All critical actions were performed, but not asefficiently as for the attending physicians. Time to diagnosis and treatment was about 33% longer than that of theattending physicians. Mission control was contacted for guidance once. There were no adverse events.

The medical students were significantly poorer in performance. Their time to diagnosis was up to 75% longer than thatof the attending physicians. Mission control was contacted in each case, and this was responsible for the majority ofthe delay. There was a death due to unrecognized esophageal intubation with the endotracheal tube, and failure tooxygenate in the third scenario.

ConclusionCrews on a mission to Mars will encounter emergency medical crisis situations. In order to handle these situationsefficiently and provide the greatest margin of safety for the crew, physicians trained in Emergency Medicine and TraumaSurgery should be incorporated into the crew. The higher level of training provided by these physicians and the skillsthey possess would be difficult for non-physicians to obtain. These skills can be crucial in determining the differencebetween life and death for an injured crew member. Therefore it is imperative that an attending level physician who isappropriately trained be included in a manned mission to Mars.

AcknowledgmentI would like to thank Mimi Hotchkiss, RN, CRNA and the Medical College of Virginia Human Simulation Center fordonating their time and resources to this project.

Appendix – Martian Emergency Medical Crisis ManagementSimulator Scenarios:

• Anesthesia and monitoring equipment removed from simulation center• placed on floor with body perpendicular to tilt-pan-zoom camera• patient cables placed along side wall• space suit on simulator, work boots, helmet off to the side(if available)• speaker to patient relocated to skull flap

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• IV access available

Props:• storage shelves in area stacked with empty, closed boxes• Thomas Pack (located in back, far wall under control room window) includes:

• Airway equipment(scope / blades / ETT / mask / NC / OA / NA)• IV setup• stethoscope• LR• Oxygen tank placed on side, against wall with app. Valve• Ambu bag• Pulse oximeter machine• ETCO2 detector• Silver rewarming blankets• Drugs:• Albuterol• Aminophylline• Corticosteroids• (5) of each: epinephrine, atropine, lidocaine, morphine• LifePak 12• LifePak 12 simulator attached; remote in control room

Patient Profile:Age: 35 years oldGender: MalePMH: NonePSH: Appendectomy 20 years ago without complications. Cholecystectomy 2 years ago without complications.MEDS: NoneAllergies: None

Operator NotesDecompression Scenario (Hypothermia, Air Embolism +/- Pneumothorax, Neuro deficit):Site: Mars Hab Module

Situation:Four astronauts have been on the Mars Hab Module for the last six months. (This allows time for the body to equilibrateto gravity changes.) An astronaut develops a leak in his space suit and quickly returns to the air lock. During the re-pressurization process he looses consciousness and drops to the floor. When the other astronauts arrive, they find theircolleague in and out of consciousness. Initial VS are BP in the 90’s, HR about 110. Patient quickly decompensatedover the next 5 minutes. VS changes are BP in 70-80’s, HR 130-150 and SaO2 falls to 80-90’s.

Role:First responder astronaut (Captain) [participant placed on wireless microphone and remote headset with microphonelocated on top of head and out of contact]

Objectives:1. Identify signs and symptoms of hypothermia.2. Identify signs and symptoms of air embolism.3. Identify signs and symptoms of pneumothorax.4. Perform appropriate medical interventions for the above conditions:

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• Airway management• IV access• Rhythm recognition• Vital signs assessment• Pharmacological intervention

• rewarming• neuro exam• recognize need for re-pressurization

Toxic Exposure to Hydrazine (Asthma / Pulmonary Edema with stridor, Hypoxia)Site: Mars Hab Module

Situation:Four astronauts have been on the Mars Hab Module for the last six months. (This allows time for the body to equilibrateto gravity changes.) An astronaut is coming in through the airlock and performs inadequate decontamination proceduresafter removing his space suit. He immediately begins to experience SOB. He calls on the intercom system to notify his

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colleagues that he is in distress. When the other astronauts arrive, they find their colleague on the floor experiencingdifficulty in breathing. Initial VS are BP in the 120’s, HR about 110. Patient quickly decompensated over the next 5minutes. VS changes are BP in 100’s, HR 120-130 and SaO2 falls to 70-80’s.

Objectives:5. Identify signs and symptoms of hydrazine exposure.6. Identify signs and symptoms of pulmonary edema.7. Identify signs and symptoms of hypoxia.8. Perform appropriate medical interventions for the above conditions:

• Begin decontamination procedures• Airway management• IV access• Rhythm recognition• Vital signs assessment• Pharmacological intervention

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Multi-Trauma Scenario (Hypovolemia / shock, Pneumothorax, Pelvic Fracture and Concussion)Site: Mars Hab Module

Situation:Four astronauts have been on the Mars Hab Module for the last six months. (This allows time for the body to equilibrateto gravity changes.) An astronaut is in the storage area and attempting to remove boxes from the top of the shelves. Asthe astronaut is removing a large box, he suddenly falls and knocks several heavy boxes down on top of him. Otherastronauts hear the loud crash and come rushing into the storage area. When the other astronauts arrive, they find theircolleague awake but in severe pain. Initial VS are BP in the 90’s, HR about 110. Patient quickly decompensates overthe next 5 minutes. VS changes are BP in 60 – 70’s, HR 130-140 and SaO2 falls to 70’s.

Objectives:9. Identify signs and symptoms of hypovolemia / shock.

10. Identify signs and symptoms of pneumothorax.11. Identify signs and symptoms of pelvic fracture.12. Perform appropriate medical interventions for the above conditions:

• Airway management• IV access• Stabilization of pelvis• Rhythm recognition• Vital signs assessment

• Pharmacological intervention• Pain management• Evaluation of blood loss for possible transfusion

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References1. “Human Exploration of Mars: The Reference Mission of the NASA Mars Exploration Study Team,” NASA SP-6017, Stephen J. Hoffman and

David I. Kaplan, editors, Lyndon B. Johnson Space Center, Houston, Texas, July 1997.2. Holland D. Some virtual reality and telemedicine applications useful for long duration spaceflight from a systems engineering perspective.

Stud Health Technol Inform. 1999; 62: 141-7.3. Dorn CR. Applications of telemedicine in the United States space program. Telemed J. 1998 Spring; 4(1): 19-30.4. Grigoriev AI. Medical monitoring in long-term space missions. Adv Space Biol Med. 1997; 6: 167-91.5. Crump WJ. An application of telemedicine technology for otorhinolaryngology diagnosis. Laryngoscope. 1996 May; 106(5 Pt 1): 595-8.6. Campbell MR. Surgical care in space. Tex Med. 1998 Feb; 94(2): 69-74.7. Davis JR. Medical issues for a mission to Mars. Tex Med. 1998 Feb; 94(2): 47-55.8. Manzey D. Mental performance in extreme environments: results from a performance monitoring study during a 438-day spaceflight.

Ergonomics. 1998 Apr; 41(4): 537-59.9. James JT. Carcinogens in spacecraft air. Radiat Res. 1997 Nov; 148(5 Suppl): S11-6.10. Vernikos J. Human physiology in space. Bioessays. 1996 Dec; 18(12): 1029-37.11. Buckley JC Jr. Preparing for Mars: the physiologic and medical challenges. Eur J Med Res. 9-Sep-1999; 4(9): 353-612. Rossum AC. Evaluation of cardiac rhythm disturbances during extravehicular activity. Am J Cardiol. 1997 Apr 15; 79(8): 1153-513. Whitson PA. Renal stone risk assessment during Space Shuttle flights. J Urol. 1997 Dec; 158(6) 2305-10.14. Viktorov AN. Residential colonization of orbital complex “Mir” environment by penicillium chrysogenum and problem of ecological safety

in long-term space flight (abstract). Aviakosm Ekolog Med. 1998; 32(5): 57-62.15. Billica RD. Perception of the medical risk of spaceflight. Avait Space Environ Med. 1996 May; 67(5): 467-73.16. Gonzalez MA. An integrated logistics support system for training crew medical officers in advanced cardiac life support management. Comput

Methods Programs Biomed. 1999 May; 59(2): 115-29.

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Manned Missions To Mars: Use Of Diagnostic Sonography

Arthur C. Fleischer, M.D.[1999]

List of Abbreviations:AG = Artificial GravityDRM = Design Reference MissionISS = International Space StationEVA = Extravehicular ActivitySMS = Space Motion SicknessTCD = Transcranial Doppler

IntroductionIt has been said that all the “hardware” for a manned mission to Mars is currently available or can be readily modifiedfrom available technology.1 However, the major limitations in accomplishing such a mission primarily involveuncertainties involving human factors.

Thee are major unknowns involving the effect of long-duration (greater than the longest Mir mission of 14 months)exposure to microgravity with intervening episodes of up to 29 forces expected on entry to and exit from Mars as wellas during Earth re-entry. This web site provides an overview of the potential perils of such a mission and proposespossible countermeasures to these. It is divided into discussions of the predictable perils during transit to and from Mars(the Journey) and those that may occur while the space crew is on the surface of Mars (Habitation and Exploration).The potential uses of diagnostic sonography during such a mission are also addressed.

The Mission And Important QuestionsWith currently available technology and equipment, it has been estimated that a Mars mission would require up tobetween 18 and 24 months; 6-8 months in transit to, and 6-8 months transit from Mars (total of 12-16 months), with 6-8 months of habitation and exploration on Mars’ surface depending on predetermined mission parameters. Accordingto the Design Reference Mission (DRM) proposed by John Charles, Ph.D. of NASA’s Johnson Space Center, theplanetary alignment of Mars and Earth would be optimal for Mars departure on January 20, 2014 with Mars’ arrivalJune 30, 2014, Mars departure on June 24, 2016, and Earth arrival June 26, 2016.2 The proposed mission would involve161 days transit out; 573 days Mars’ surface stay and 154 days for return. The actual duration of each of thesecomponents depends on the specific planetary alignment during a specific launch window. It can be predicted that Mars’perihelion will be January 22, 2013 and December 10, 2014.

A few of the important questions that come to mind regarding this mission are:

1. Will the astronauts “weather” the relatively long trip to Mars and be physically and mentally able to complete theirtasks when they arrive?

2. Will they be able to cope with long duration (minimum 30 months) space flight and mission and the ride “home”?3. Will they be able to cope with acute and chronic medical problems that can occur in transit or while on Mars?4. What are the predictable perils (both physical, such as meteor collision, as well as mental) and possible

countermeasures to these potential problems?5. How can diagnostic sonography be used during this and other long duration missions?

Some of the predictable perils of a manned Mars mission have been formulated in the Design Reference Mission (DRM)described by Dr. Charles.2 It is proposed to test some of these when the Mars Habitat is attached to ISS. Several otherimportant points made in his presentation include:

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Arthur C. Fleischer, M.D.; Chief, Diagnostic Sonography; Vanderbilt Center for Space Physiology and Medicine, 21st Avenue S. and GarlandSt., Nashville, Tennessee USA 37232-2675; Phone: (615) 322-3274; Fax: (615) 343-2512; E-mail: [email protected];

1. A Mars mission would require a minimum duration of 30 months. The length alone of such a mission is a majorfactor considering the longest duration experienced by a spacecrew member in Earth’s orbit up to now is as part ofa Mir spacecrew. The proposed duration of stay on the International Space Station (ISS) for the space crew is only3-6 months. There is no data on effects of long (greater than several months) duration spaceflight.

2. The number and severity of episodes of hypergravity and G transitions will be greater than Earth missions with 1.5-2.0g for Earth aerobreaking compared to 3-5g x 2 for Earth and Mars aerobraking. There will be a doubling of Gtransitions for a Mars mission (Earth = 2, Mars = 4).

3. There will be frequent extravehicular activities (EVA’s) on a Mars mission versus infrequent for Earth orbitmissions.

4. One has to consider extraterrestrial toxins on Mars as opposed to only spacecraft and terrestrial toxins for Earthrelated missions.

Dr. Charles’ excellent overview serves as a guide to formulate countermeasures to predictable perils.2 Other issues suchas development and implementation of bioregenerative systems needed for a long duration mission and specialrequirements of a proposed mission deploying an Earth return vehicle and fuel generation plant to Mars’ surface beforea manned mission is launched also need to be addressed. Some of what follows is based on this work.

During the JourneyDepending on the launch timing relative to the position of Earth and Mars, under optimal conditions just the journeyover 250,000,000 miles between Earth and Mars will take anywhere between 12 and 16 months. During this time, thespace crew will traverse space that has some known and unknown hazards such as possible meteorite collisions andcosmic radiation. Mars orbit is 227,940,000 km (1.5 Astronomical unit) from the sun whereas Earth’s is 149,6000,000km (1 Astronomical unit).

Delays in communication between the spacecrew and Earth centers will vary from 3-20 minutes depending on thedistance the space crew are from the Earth. This necessitates handling of some decisions without real-timecommunication from controllers on Earth. It also mandates the ability of the spacecraft to alter course immediately ifthere is a perceived danger of unexpected collision with meteors or space debris.

Effects of Microgravity During TransitThis discussion primarily involves problems that may occur when the space crew is in transit towards and away fromMars. In general, the predictable effects of prolonged weightlessness involve gravity receptors, fluid shifts, and changesin weight bearing structures. These effects are nicely discussed and illustrated in an article by Professor White inScientific American.3 Assessment of these changes that occur in space, interestingly enough, may help us betterunderstand a variety of common disorders, such as those involved in aging, that are encountered in humans on Earth.

It should be pointed out that these predictable physiologic changes are based on our experience with relatively shortduration missions, however. Prior to the actual Mars mission, it is proposed that the Mars Habitat Module should betested while attached to the International Space Station in low Earth orbit. The “livability” of the module may be testedduring 3-6 month mission of ISS before being optimized for the Mars mission.

Changes in gravity receptors have extensive effects including effects on vestibular apparatus in the inner ear. Theneurovestibular appropriates detection of normal orientation of the body in space as well as acceleration / decelerationforces. Some have advocated the use of cushioned shoes to better simulate gravity’s effect while in space. In space,touch and pressure receptors no longer signal the “down” direction. These problems contribute to visual-orientationillusions and feelings of self-inversion, such as feeling that the body or the spacecraft is spontaneously reorienting.

Space motion sickness (SMS) will affect some astronauts, more the rookies than the pros, during the first few days ofthe mission. This is why major tasks are typically not planned in the first few days of a mission. Conceivably, SMSmay also affect space crew that returns from Mars missions after an extended stay on the planet. The symptoms of SMS

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include headache, impaired concentration, loss of appetite, stomach “awareness,” and vomiting. Although typicallyexperienced in the first few days of spaceflight, some cosmonauts have reported SMS toward the end of long missions.3

Another major effect of weightlessness involves fluid shifts within the body. Within minutes of arriving in amicrogravity environment, the astronaut’s neck veins (internal jugular veins) bulge and the face becomes puffy. As theprocess of fluid migration to the chest and head continues, sinus and nasal congestion results – altering taste and smell.These fluid shifts occur because the body fluids no longer have appreciable weight in microgravity. It is important torealize that about 60% of a person’s weight is water which is contained either in the cells of the body (intracellular), inarteries or veins (plasma volume), or in the spaces between blood vessels and cells (interstitial spaces). Average loss ofplasma volume while in space is 9%. When a person stands up, gravity contributes to physiologic changes, termedhydrostatic changes, in these three compartments. In feet, both arterial and venous pressure can increase by 100 mmHg, double the normal arterial pressure and many times the normal venous pressure.

In space, hydrostatic pressure does not play an important role in fluid distribution. Direct measurements of weightvolume have shown that each leg loses approximately 1 liter of fluid (almost 1/10th of the total volume) in the first day.This contributes to the development of “chicken legs” and is associated with a plasma volume diminution ofapproximately 20% accompanied by an increase in the frequency and volume of urination. These changes initiate acascade of renal, hormonal, and electrolyte ones which include an approximate 20% increase in kidney filtration rate,and a relative anemia due to decreased production of red blood cells (erythrocytes) due to the body’s perceivedoverabundance of erythrocytes.

To counteract this effect, daily and vigorous treadmill exercising while using “vacuum pants” or a “skirt” while theastronaut is tethered to the machine can be used to encourage fluid into the legs. In addition, the exposure of theastronauts to artificial gravity by rotation of the space habitat may reduce these effects.

Artificial gravity (AG) may be used to pre-adapt the space crew to Mars gravity (outbound) and re-adapt to Earth gravity(inbound). For a Mars mission, it could provide extended physiologic protection from 1g, ease transition throughout1/3g exposure, and provide 1g outbound and inbound. However, as of now, there is no consensus regarding the AGlevels needed for an exploratory mission. According to the DRM, a research program should be initiated to determineoptimal characteristics for intermittent AG, identification of threshold values needed to optimize human performance,and to determine optimal AG characteristics.2

Another major effect of space flight that will affect space crew in a Mars mission relates to changes in muscle and bonemass. Because the force of gravity is removed in space, the spine straightens and people actually grow 2” or so. Inaddition, lungs, heart, and other organs expand. As one shuttle physician / astronaut (Dr. F. Andrew Gaffney) stated,“You feel your guts floating up. I found myself tightening my abdomen, pushing things back.”3

Because skeletal muscle, which evolved to allow an upright position in gravity, atrophies in microgravity, changes in thephysiologic function are altered in space. The muscle groups used with the space environment, namely for fast contractilefibers differ significantly from those needed on Earth. Exercise is vital to maintain adequate muscle and bone density.

Changes in bone density can also be predicted as a result of the change in osteoclasts vs. osteoblasts (construction vs.destruction of bone elements). The process also changes calcium concentrations and distribution. Joint Russian andAmerican studies have shown that cosmonauts lose bone from the lower lumbar vertebrae, hips, and upper femur at therate of 1% per month.3 This would compute to up to 36% loss for a 3 year mission which is very substantial! Increasedcalcium turnover could also contribute to formation to renal calculi.

This problem with bone loss can be compared to the problem of decreased bone density in post-menopausal women.The major countermeasure to this is for astronauts to perform vigorous onboard exercise which increases blood flow tomuscle and provides a stimulus to maintain bone density.

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Other predictable changes include changes in the immune system, human factors involving poor quality sleep,psychological irritation of close living quarters and radiation effects. Immune system changes may contribute to arelative immunodeficiency.4 Changes in day/night cycles can contribute to adverse effects on circadian rhythms, whichnegatively impact on proper and sufficient sleep patterns. And we all know what poor sleep does to our interaction withour fellow human beings!

The radiation affects, both predictable and unpredictable, will be an important field on a Mars mission. According tothe DRM, a Mars mission should involve 26-30 months exposure to both galactic cosmic radiation and solar particleevents. In addition, there would be significant exposure risk if a nuclear power reactor were used for the spaceship. Anastronaut spending 1-year in Earth orbit would receive a radiation dose ten times that of ground inhabitants.1 It ispredicted that a one-year stay on the Moon would result in a dose seven times higher and a flight to Mars would be evenworse. Sudden outflows of particles from the sun, like that that occurred in August of 1972, can deliver a dose of morethan one thousand times annual ground dose in less than 1 day.1 Fortunately, such events are rare and spacecraft willcontain areas of special shielding in which the astronauts can temporarily shelter from such a threat. The long-termeffects of radiation, namely the risk of potentiating cancer, cannot be definitively assessed since it is impossible tocompletely replicate the radiation dose that an astronaut would receive on a long duration mission.

For short-term missions, returning to Earth affords spontaneous and complete reversal of the previously mentionedeffects of space travel. However, the actual effects of space flight can only be clearly appreciated by the astronaut andground based observers. For example, space crew find it difficult to stand up for more than 10 minutes without feelingfaint immediately after landing. This phenomenon is termed orthostatic intolerance and can be observed in bedriddenand some elderly people on Earth. Other effects of long-term duration flight that have Earth-related counterparts involvewobbliness created by space travel and tendency of the elderly to fall, bone and muscle loss experienced in space andosteoporosis, immunodeficiency, poor sleep patterns with associated decrease in motor coordination. Analysis of thedata derived from the return of Astronaut / Senator Glenn may help clarify some of these issues as they relate to spacetravel as well their counterparts involved in aging processes in Earth-bound folk.

What Can Go Wrong . . . Lessons from MirThe United States’ involvement with Mir missions has shown potential problems of long term, multicultural missions inspace. These have recently been vividly documented by US astronaut Jerry Leninger on a Nova videotape presentationentitled “Terror in Space”.5 Physical dangers such as onboard fires initiated by faulty oxygen canisters and uncontrolledcollision of an unmanned spacecraft with Mir made for life-threatening and highly stressful situations on these missions.

Experience coping with Mir and ISS missions will be vital to better understand and deal with issues involving longduration spaceflight. In addition, the optimal uses and integration of such low Earth orbit space stations for futurelaunch, control, or other activities for support of long duration space flight might be determined.

Probably, the most important crew member on a long duration mission is the “mechanic,” followed by a physician whocould tend to the medical and emotional needs of the crew as discussed by Dr. Zubrin on Destination Mars, 1997.1Multicultural differences need to be taken into account, as well as the ability of the crew to respond and communicatewith one another and act as a team.

During Mars HabitationThe Mars environment poses significant perils, both predictable and unpredictable. Hopefully, the space crew wouldhave conditioned themselves physically and mentally during transit to be able to tolerate Mars gravity and the spacesuitneeded for extravehicular activities (EVA) on Mars’ surface.

While on Mars’ SurfaceScientists tell us that Mars and Earth have similar evolutionary paths, but that Mars, sometimes referred to as Earth’s“little brother,” lost a thick atmosphere and its water million of years ago.9 Many of the environmental differences

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between Earth and Mars can be traced to a lack of an atmosphere to contain vital gases. The thin (less than 1% of Earth’s)atmosphere consists primarily of carbon dioxide. Water on Mars is concentrated at the polar ice caps and deep within thecrust. Mars has approximately one tenth of the water present on Earth. Its gravity is approximately one third of Earth’s.The temperature extremes on the surface of Mars are formidable, from -120ºC at the poles in “winter” to +25ºC at theequator in “summer.”9 The average temperature on the surface of Mars is -50ºC. There is no atmosphere to blockultraviolet rays from reaching Mars surface. These factors must be considered while organizing a manned Mars mission.

As of now, it is proposed that the habitat and fuel generation facility be operational prior to launching the astronauts.2A Mars “day” is nearly that of Earth’s (24 hrs 37 minutes) but its orbit around the sun takes about 2 Earth years. TheMars habitat must contain a radiation shelter for contingent use during a solar flare. The space suits used should beshielded as well. The space suit used on a Mars mission needs to be maximally flexible as well as protective in orderto provide safety with flexibility during excursions on the surface of Mars.

Other predictable perils to humans while on Mars include violent wind and sand storms (with winds up to 80 mph) andunpredictable episodes of massive cosmic radiation.1 Thus the space crew can not venture too far (less than 1 mile)from the habitat or have some means of surface transport that can get them back to home base as soon as possible ifeither of these occur.

Table 2 lists some of the important parameters pertinent to planning of a Mars mission. While there are some similaritiesbetween Earth and Mars, i.e., the length of its “day” and surface temperature extremes, many other parameters are quitedissimilar and need to be considered in planning a manned mission to Mars.

Potential Medical ProblemsBecause of the long duration (a minimum of approximately three years) and hazards of the mission, and delay incommunication with ground control experts, it is recommended that one of the space crew be a physician familiar withconditions which may affect fellow space crew. The physician on the space crew should have the responsibility for earlydiagnosis and treatment of a variety of medical disorders. One of the most reasonable and practical tools that thephysician will use is diagnostic sonography (Figure 1). What follows is a brief discussion of the use of diagnosticsonography during transit to and from Mars as well as attending to those disorders that may be encountered while onMars’ surface.

The Potential Use of Diagnostic SonographyCertain issues need to be addressed because their occurrence during such a mission is virtually certain. These are:

1. Changes in bone density and matrix and muscle in microgravity.2. Cardiovascular and intracerebral blood flow changes in microgravity.3. Possible adverse affects of cosmic radiation and other physical dangers, such as meteorite collisions.4. Other unpredictable acute disorders such as rotator cuff, Achilles tendon tear, and carpal tunnel syndrome related to

trauma or “occupational” hazards.5. Other physiologic and/or psychological effects

Accordingly, it is highly recommended that those planning a Mars mission to have diagnostic sonography available bothduring transit and within the Mars habitat. Each of these potential applications will be addressed. A more in-depthdiscussion can be found elsewhere.6

Bone Density Changes:Prolonged exposure to microgravity is associated with profound change in bone density and matrix. Bone sonometeryuses a small machine about the size of a small microwave oven and can be used to monitor changes in bone density(Figure 2).6

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Maintenance of bone density and muscle turgor will require that the astronauts exercise a minimum of two hours whilein transit. As with any form of vigorous exercise, the possibility for acute injury exists. Diagnostic sonography providesa means to diagnose a variety of these, including rotator cuff tear, carpal tunnel syndrome, Achilles tendon rupture, renal/ liver / pancreatic trauma, testicular or adnexal torsions, to name a few.

Ultrasound can also be used to monitor an astronaut’s bone density by determining the speed of sound (SOS), boneattenuation (B. A. in dB/cm) and a mathematically derived value of “bone stiffness” (T-score), which is related totrabecular structure. It is possible that bone sonometery will be used weekly to monitor and detect early significant bonedensity loss. It will also be a means to monitor changes in bone density with various therapeutic regimens.

3-D color Doppler sonography may also be used to depict changes in blood flow to muscle before and after exercise(Figures 3A, 3B, 3C). This information may be useful in monitoring muscle health and early detection of atrophy.

Changes in Cardiovascular and Intracerebral blood flow:With long exposures to microgravity, there are major changes in the cardiovascular system including “cephalization” ofblood volume, changes in stroke volume and decreased venous return due to microgravity. How will this best betreated? The use of vacuum body suits to encourage flow into the legs may help. Production of “artificial gravity” byrotation of the space vehicle may also help.

Some of these changes can be assessed using echocardiography. They should be correlated with changes inhemodynamics that have been described through the use of continuous central venous pressure monitoring.7 This bringsto mind a few more questions:

1. Who will perform the echocardiograms?2. How much time / experience is needed in order to perform and interpret the scans?3. Is it possible to transmit real-time images to an expert at a ground station?4. What are the important cardiovascular changes and how best to treat them?5. How does this affect intracerebral blood flow?

These questions need to be addressed on ground-based studies as well as ones done on ISS. There is extensive literatureon the cardiovascular effects of space travel but does it apply to long duration missions? How is intracerebral bloodflow affected by cardiovascular changes? Transcranial Doppler US may help. This technique is difficult even in themost experienced hands on Earth let alone in space! (Figure 4)

Radiation effects:Excessive radiation can result in a variety of disorders including death of any rapidly growing cell. This may lead to avariety of symptoms arising from radiation sickness, ranging from acute diarrhea to chronic potentiation of cancers.Abdominal sonography may be used to distinguish other causes of abdominal pain (i.e., pancreatitis, cholecystitis) fromthose symptoms of acute radiation sickness.

Radiation hazards may arise for unpredictable sources and there needs to be an “early warning system” to detect them.This would allow the space crew to seek shelter and protection as soon as possible.

Acute trauma:The rigors of space travel and work (exercise, EVAs) may be associated with acute trauma, be it musculoskeletal or otherorgan system disorders (Figure 5). Diagnostic sonography can be used to diagnose and assess the extent and responseto treatment of the disorders such as pancreatitis secondary to blunt abdominal trauma.5

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Other:Another peril to the space crew is possible collision with meteorites and/or space junk. An ever-vigilant warning systemis needed for this problem. Can the astronauts deal with sudden terror? The “cool” demeanor of astronaut / cosmonautJerry Lenninger allowed him to survive a life threatening series of mishaps on Mir.5 If a rescue mission was required,could the equipment stand the rigors of the trip? The recent attempt to airlift medical supplies and equipment to theAntarctic exemplified this!

SummaryFull and proper consideration of human element issues is the most complex and vital part of a manned Mars Mission.Manned Mars missions will pose significant physiologic and psychological challenges to the spacecrew. Some of thepotential perils can be simulated with ground based experiments while others will only become apparent during the mission.

This work provides an overview of the predictable perils proposed countermeasures involved in manned Mars missions.Table 1 lists the predictable perils and possible countermeasures to these. I believe it is wise to “expect the unexpected.”I also believe having a trained physician on board will be vital to the success of such missions.

Table 2 compares some of the Earth and Martian parameters that are important for factoring into a manned Marsmission. Although much of the Martian surface can be considered a “rusty desert” there are high volcanoes and deepcanyons to be reckoned with.

A unified and inspired effort is needed to realize the goal of safe journey and habitation of Mars in the near (next 15years) future. Diagnostic sonography provides an important part of ensuring the health and safety of Mars spacetravelers. Table 3 summarizes some of the sonographic techniques that may be useful in a manned Mars Mission. Assonographic techniques improve, additional technology will and should be added to that already in space or awaitingdelivery into space.

AcknowledgmentsThe author would like to thank ATL / Phillips, Inc. for their financial support of this project. Specifically I would liketo acknowledge Mr. Bob Dockendorf, Vice President of Sales, for his support of this and other research projects. Bradenand Lynn Fleischer are thanked for their review and help creating this web page. Lorene Walter and Janet Staley arerecognized for their editorial assistance. The helpful comments of David Martin, RDMS of Wiley Labs, John Charles,Ph.D. of NASA Johnson Space Center and Rhea Seddon, M.D., F. Andrew Gaffney, M.D., and David Robertson, M.D.,Ph.D., of Vanderbilt University Center for Space Physiology and Medicine are also appreciated.

ReferencesDestination Mars. Discovery Series Videotape, 1997. Charles, John Human Health and Performance Aspects of the Mars Design ReferenceMission.

1. Charles, John Design Reference Mission (DRM) Human Space Life Sciences Program Office (HSLSPO) NASA – Johnson Space Center,Houston TX, 1997.

2. White, R. Weightlessness and the Human Body Scientific American. Sept. 1998 p51-63.3. Jensen, W Notes from SpSt 410, University of N. Dakota, Olegard School of Aerospace Studies.4. Terror in Space. Nova, 1997 (WGBH PBS Boston).5. Fleischer AC: The use of diagnostic sonography on long duration spaceflight. Paper for SpSt 410. University of North Dakota Olegard School

of Aerospace Science. May 3, 1999.6. Buckey, J. Gaffney, F, et al. Central venous pressure in space: letter to editor NEJM 328(25):1853-1854, 1993.7. Raeburn, P Uncovering the secrets of the Red Planet, N.Y., Nat. Geographic, 1998.8. Couper H, Henbest N. The Space Atlas - A Pictoral Atlas of our Universe. London: D. Kindersly Pub 1992 p. 28-29.9. Zubrin, R. The Case for Mars - The Plan to Settle the Red Planet and Why We Must. N.Y.: Simon and Schuster 1996 p.3.

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Tables

Table 1. Predictable Perils and Proposed Countermeasures of Manned Mars Mission

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Table 2. Earth vs. Mars Comparisons

* from Ref. 9

Table 3. Potential Uses of Diagnostic Sonography on Manned Mars Mission

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Figures

Figure 3B

Figure 3C

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Figure 1. Photo of medical instrumentation panel including anultrasound scanner (ATL 5000) within the wall of the HumanResearch Module of the International space station. The re-configured scanner is incorporated into a panel containing avariety of medical instruments. (Courtesy of Lockheed -Martin, Inc.)

Figure 2. Bone sonometer. Thisdevice assesses bone speed ofsound and attenuation in thecalcaneus without the use of awaterbath. A T-score whichcorrelates to bone density iscalculated and plotted againstnormative values. (Courtesy ofHologic, Inc.)

Figure 3A. 3D color Doppler sonographyshowing blood flow picture taken during anultrasound scan of the author’s formidablebiceps muscle, before 20 curls. Whencompared to image taken prior to exercise (Fig.3B) there is more flow within the biceps muscleafter exercise (Fig. 3C). There is increased flowinto the muscle, as evidenced by more vesselswithin the biceps after exercise. The flow withinbrachial artery was used as a reference.

Figure 4. Transcranial Doppler US (TCD)

Figure 4A. Diagram showing typical transducer positions(transtemporal or transoccipital) and intracerebral vesselsthat can be interrogated (courtesy of ATL / Phillips, Inc.)

Figure 4B. Typical TCD image for TCD of right middlecerebral artery. Time averaged peak velocity (T.A.P.) is67.7cm/s, which is normal.

Figure 5A. Superficial organ or “small parts” transducer usedfor musculoskeletal sonography. Picture of high frequencylinear array transducer as used for imaging superficialstructures. The transducer is small and can be used like a pen.Here it is being used to evaluate the rotator cuff. (Drawing byPaul Gross, M.S.)

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Figure 5 B,C. Sonogram showing normal (B) and torn(C) rotator cuff. Notice the full thickness break in the outlineof the supraspinatus ligament.

Figure 5D Sonogram showing torn Achilles tendon in long axis. (Courtesy ATL / Phillips)

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Preliminary Design Considerations For The M.A.R.S. Wastewater Treatment System:Physico-Chemical Or Living Machine?

David M. Blersch; Erik Biermann; Patrick Kangas, Ph.D.[1999]

AbstractThe Mars Society has been preparing designs for a terrestrial-based analog human habitat for Mars. The Mars ArcticResearch Station (M.A.R.S.), currently in the design phase, is slated for construction beginning in July of 2000 inHaughton Crater on Devon Island in northern Canada. An important component of the M.A.R.S. will be the wastetreatment system. Housing four to six people, the M.A.R.S. will produce both gray water and domestic sewage thatrequires treatment. The external environment of the M.A.R.S. imposes unique design constraints, affecting size, cost,reliability, and other critical features of the waste treatment system. The focus of this paper is on the treatment of thewastewater, but the ultimate design for the M.A.R.S. – and for Mars missions in general – will need to be multipurposewith considerations of water recycling, nutrient recycling, and food production. To begin the design process, it isimportant to review alternative treatment systems that are available. The literature on alternatives is reviewed, includingconventional physico-chemical treatment, microbial reactors, algal turf scrubbers, and treatment wetlands. An analysisis made of the relative qualities of these systems to assist in the choice of designs for the M.A.R.S. Some systems, suchas some conventional physico-chemical systems, may serve only the purpose of waste treatment, while other systems(especially biologically based systems) have regenerative potential. Two of these systems – algal turf scrubbers andtreatment wetlands – are selected for more in-depth engineering analysis, with preliminary calculations for scaling to asix-man crew. Ultimately, the selected design will probably be a modular system with treatment unit processes hookedup in series for progressive stepwise treatment. Finally, consideration is given to adapting Arctic species to treatmentsystems to enhance performance in a colder, low-light environment.

IntroductionAny manned Mars mission will require a certain degree of self-sufficiency in food production and waste recycling. TheMars Society’s Mars Arctic Research Station (M.A.R.S.) planned for Haughton Crater in Devon Island will be anopportunity to test some strategies and technologies for achieving the goal of closure and self-sufficiency. Designing aclosed life-support system for a mission to Mars presents a variety of challenges, such as low temperatures, low wateravailability, high ultraviolet light exposure, and low availability of nutrients. Many of these same challenges will befaced in the planning of the Devon Island experiment. To overcome these challenges, resource-recovery waste treatmentprocesses must be incorporated into the mission that are multipurpose and multifunctional. The purpose of this paperis to begin the analysis of a wastewater treatment system for the Devon Island station within the context of multipleobjectives and optimal choice of alternatives.

There is a large literature base on the design of life support systems for manned space travel that is relevant to the designof the Devon Island station and ultimately to a manned Mars station. The kinds of technologies that have beenconsidered can be classified along a gradient of resource re-utilization, ranging from non-renewable to renewable /regenerative. At the non-renewable end, the mission strategy is to carry from the mission’s start all the resources (water,oxygen, food, etc.) required for the entire mission as storage, and wastes are discharged as they are produced or storedfor later disposal. At the other end of this gradient are regenerative technologies, which emphasize the production ofconsumables from the recycling of waste materials. The literature includes designs for life support systems developedfor spacecraft, space stations, and extra-terrestrial settlements.

One of the continuing areas of discussion and investigation is the degree to which life support systems are mechanicallybased versus biologically based. Physico-chemical life support alternatives have been well developed through the pastfew decades (Eckart, 1996). While they have proven successful for the short-term or proximate Earth-orbital mannedmissions of the past, these technologies exhibit few characteristics required for regenerative life support design.

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David M. Blersch; University of Maryland, Department of Biological Resources Engineering; 1426 Animal Science / Ag. Engineering Bldg.,College Park, Maryland 20742-5711 USA; Phone (301) 405-1183; Fax (301) 314-9023; email: [email protected], [email protected]

Research continues on developing viable biologically based regenerative “closed” life support systems in which some orall wastes generated by a human crew are recycled into usable consumables. Examples of these technologies include theNASA Advanced Life Support (ALS) development program, the Soviet / Russian experiments in BIOS-3, the Biosphere-2 effort in Arizona, and a variety of interesting but lesser known experiments since the 1950’s (Taub, 1974; Eckart, 1996).

Wastewater Treatment Technology AlternativesThe treatment of wastewater produced by humans during space missions is a fundamental part of life support systems.Wastewater can be deleterious to humans because of the existence of pathogens or the concentration of toxic substances.However, wastewater also contains valuable nutrients that might be recycled into a food-production system. Wastehandling systems must therefore effectively isolate the waste stream and distribute it to treatment alternatives forresource recovery. The simplest approach to waste handling has been to dispose of it to the external environment or tostore it for later disposal (Eckart, 1996). Conversion by various forms of combustion or incineration is a promisingapproach for short-term missions, but often entails high-energy expenditures and few paths of resource recovery. Analternative class of waste handling systems is biologically based using controlled communities of biological species thatconsume wastes and that form the basis of food chains that can lead back to humans.

Table 1 provides a preliminary review of some of the technologies that have been examined specifically for wastewaterhandling and treatment for space missions. A number of criteria are considered and rated qualitatively for each of thetechnologies. These criteria form an important subset of a larger list of design features and tradeoffs that must beconsidered for the treatment approach (Wieland, 1994), including:

Table 1 summarizes some of the criteria evaluated qualitatively to choose between alternatives. The first four rows inTable 1 are examples of physico-chemical treatment technologies that have been or are currently being studied in spaceprograms (Wieland, 1994; Eckart, 1995). A detailed explanation of the physico-chemical processes will not bepresented here; refer to Eckart (1995) for a discussion of these technologies. The last two rows represent examples of“living machines” – a series of biologically based treatment processes utilizing controlled ecological design. Thecriteria selected for analysis in Table 1 include some of the traditional tradeoff considerations, such as relative mass andpower consumption, along with considerations related to regenerative capacity and multifunctional potentials.

All of the systems considered in Table 1 have positive and negative aspects. Biological systems are usually more easilyintegrated into regenerative life support systems with food production. They can efficiently recapture valuable nutrientswhile providing a number of secondary services such as O2 production and CO2 absorption, and water purificationthrough plant transpiration. These systems, however, almost invariably require high mass investment to maintainadequate growth conditions for the ecosystems utilized. High-energy inputs might also be required in the form of lightor heat, but these can be easily augmented by direct solar energy. Some physico-chemical systems offer someregenerative characteristics comparable to biological systems, often recycling wastes into usable atmosphericconstituents. While these systems are generally more compact and of less mass than biologically-based systems, mostrequire a high concentrated energy input for the process to function. Nutrient recapture by these systems is generallyminimal, and they are thus more difficult to integrate into a food-production mission scenario. All these aspects mustbe considered further and quantified in the context of specific mission design to determine which is appropriate for theDevon Island station.

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Preliminary Design Considerations For The M.A.R.S. Wastewater Treatment System: Physico-Chemical Or Living Machine?

• design risk• design simplicity• cost• serviceability• contingency operating modes• noise• mass• power consumption• similarity to “home” environment

• reliability• safety• ease of operation• waste recycling capability• training requirements• expendables consumption• pathogens and odor control• crew time requirements

Table 1. Preliminary review of wastewater treatment alternatives for the Mars Arctic Research Station (M.A.R.S.)

Sample Designs For Ecologically Engineered SystemTo further the analysis of the appropriate life support technology for mission design, the two biologically based systemsin Table 1 are scaled for a potential M.A.R.S. mission scenario. An important preliminary step in the design process isthe sizing of the unit processes that make up the overall treatment system. This was performed here using standardequations from the literature (Metcalf and Eddy, 1991; Reed, et al. 1995; Van Haandel and Lettinga, 1994). Twoscenarios for waste stream generation come from opposing mission-diet scenarios presented by Hall & Brewer (1985),representing complete food-storage on the one extreme and complete food production on the other. The food storagescenario consists entirely of preprocessed, prepackaged foods designed for their reduced mass and for their maximumcaloric and nutritional value. The food production scenario assumes a vegetarian diet from greenhouse production; itswaste stream is larger than that of the food storage scenario, as it includes (1) a greater proportion of biomass from thehigh percent of inedible and indigestible materials in plant structure, and (2) a greater wastewater volume from irrigationand nutrient delivery subsystems. Based upon these two waste stream scenarios, the sequences of biological unitprocesses of the waste treatment systems are presented in Figure 1, sized for a six-person crew. In the figure, the boxes

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represent the treatment unit processes, and the arrows represent the flow either of total organic carbon (reported aschemical oxygen demand or COD) or of total nitrogen (TN) through the processes. Each system begins with a holdingtank, intended to provide flow capacitance to ensure an even flow rate to subsequent components. The tank is followedfirst by an upflow anaerobic sludge blanket reactor (UASB) and then by an aerobic trickling filter (ATF). Thesecomponents are primarily responsible for the breakdown and digestion of the complex organic compounds in the waste,the reduction of the COD to acceptable levels, and the nitrification of the nitrogen compounds. The UASB provides ahigh-rate digestion process that is efficient at removing the high concentrations of COD with low sludge production (vanHaandel & Lettinga, 1994). The ATF efficiently completes the breakdown of the organic compounds and nitrificationof ammonia (Metcalf and Eddy, 1991).

Figure 1a. Wastewater treatment unit processes for a six-person food production mission scenario

Figure 1b. Wastewater treatment unit processes for a six-person food storage mission scenario

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At the end of each of the treatment systems, two photosynthetic controlled-ecology units, designed for the primaryfunction of reducing total nitrogen levels in the wastewater, are contrasted. The algal turf scrubber uses a complex,stable community of attached filamentous algae as the basis for the nitrogen uptake (Adey, et al. 1993). In conventionalterrestrial wastewater treatment systems, algal turf scrubbers have exhibited a high rate of nutrient uptake in waters ofvarious nutrient concentrations (Craggs, et al. 1996). Optimizing the light and heat available to the algae communitymaximizes primary productivity of the algae, and thus the nutrient uptake rate. Because they are complex communities,algal turf scrubbers do not show the sensitivities and instabilities often found with monocultures of algae (Adey, et al.,1993). Optimal operation of the algae system requires periodic harvesting of the algae biomass to remove metabolitesand grazer organisms. This stimulates increased productivity through regeneration of the algal thallus (Craggs, et al.1996). For the calculations here, a primary productivity rate of 60.9 g•m-2•day-1 of algae biomass is assumed, with a5% by mass nitrogen content (Adey, et al., 1993).

In contrast to the algal turf scrubbers, treatment wetlands consist of communities of higher emergent plants, usuallycattails, reeds, and bulrushes in standard terrestrial applications. Wetland cells may be designed where the waterundergoing treatment is maintained at a level below the top of a basin of gravel in which the plants grow. It is thoughtthat the biological reactions responsible for waste treatment are performed primarily by attached growth biofilmorganisms (Reed, et al., 1995). Thus, because of the high surface area for biofilm growth provided by the gravelsubstrate and the plant roots, subsurface flow wetlands exhibit relatively high treatment efficiencies per unit of squarearea of wetland. The gravel bed also provides a measure of thermal insulation, a positive consideration for designingwetlands for colder climates (Reed, et al. 1995). Design of a wetland system for nitrogen removal is performed usingstandard equations from the literature relating the flow rate, influent and desired effluent concentrations, and a reactioncoefficient dependent on temperature, here assumed to be 20°C (Reed, et al. 1995).

One of the fundamental differences between these two systems is the ultimate fate of nitrogen, important inconsiderations of nutrient recycling in a closed life support system. In the algal turf scrubber, most of the nitrogen iscaptured in the algae biomass. Thus, the biomass accumulated through the regular and periodic harvesting regime canbe used directly for fertilization of crops or for feeding of animals. In the treatment wetlands, however, nitrogen is lostto the atmosphere as N2 gas through biological denitrification (Reed, et al., 1995). While this nitrogen might remaincaptured by a tightly closed life support system, its recycling as a nutrient for crop growth would require additional massand energy investment in a process or processes for nitrogen fixation.

For the analysis presented here, the contrasting controlled-ecology unit processes are scaled to reduce the total nitrogenloads to a level of 5 mg•L-1. Previous research has shown that most pathogens can be removed by these systems sizedfor nutrient and COD removal, although post-processing (disinfection and filtration) would be desirable for potablewater generation (Metcalf and Eddy, 1991). Figure 1 summarizes the results of all calculations, showing the requiredsizes for waste treatment for a six-person crew. Because of the greater waste volume in the food production scenario(Figure 1a), these treatment unit processes are necessarily larger than for the food storage scenario (Figure 1b). Thepreliminary unit processes were calculated to be relatively small and compact. Based upon the respective volumetricflow rates for the waste stream, the holding tank was sized for each scenario to contain two days worth of wastewatervolume. The required anaerobic reactor, sized for 80 per cent removal efficiency of COD, was determined to be 0.093m3 for the food production scenario and 0.035 m3 for the food storage scenario. The aerobic trickling filter, scaled for85 per cent removal efficiency of COD, was sized at 0.082 m3 for the food production scenario and 0.031 m3 for thefood storage scenario.

A large size difference is seen between the two contrasting photosynthetic unit processes intended for nutrient recapture.For each scenario, the algal turf scrubber was considerably larger than the contrasting treatment wetland – between 8and 14 times larger on an area basis, and between 1½ and 3 times larger on a volume basis. These differences are basedon the fundamentally different ecosystem processes. Because its efficiency relies on maximization of primaryproductivity, an algal turf scrubber is designed as a long, shallow raceway. Surface area is maximized to providemaximum light availability to algae cells, and agitation of the thin film of water in the raceway maximizes nutrient

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mixing and delivery to the cells. Treatment wetlands, however, rely upon the extensive surface area of the gravelsubstrate and of the plant root zone biomass to support the complex denitrifying bacterial community. Hence, they donot require as much total square greenhouse area or growth basin volume for efficient treatment. It may be hypothesizedthat these differences in size extrapolate directly to different energy requirements. The algal turf scrubbers naturallydemand a greater light intensity than do treatment wetlands for maximum treatment efficiency. Thus, for the algalsystems, the available solar light may have to be augmented by artificially produced light. Also, the higher surface areaof the algal turf scrubbers results in a greater rate of heat loss, demanding a more carefully controlled and larger, warmergrowth climate than the wetland cells.

ConclusionsChoosing between wastewater treatment alternatives will not be a simple decision for a Mars mission. A variety ofalternatives are available, and there are many criteria that need to be balanced and weighed in choosing between them,as evidenced in Table 1. Will wastewater treatment be approached as an isolated system in the station design, servingas a sink for the many waste compounds produced? Or does the treatment system need to be multifunctional andaccomplish resource recovery to integrate with other systems such as food production or atmospheric management? Ifa Mars mission is to include provisions for food production, then issues such as nutrient recycling for food productionbecome a necessary consideration. If, however, food will be stored or supplied by regular shipment from Earth, nutrientrecapture may not be so much of a concern. Thus the details of the particular mission scenario need to be defined totailor the design to the specific requirements. It will be best to address these issues at the Devon Island station.

The ecologically engineered alternatives examined in more detail here require significant areas of controlledtemperature and light regimes for optimal performance. This will require significant investment in mass and energy fora greenhouse or growth-chamber. At first, this requirement might seem to negate the viability of these options.However, the benefit of these options over some physical-chemical systems is the ready production of O2 andconsumption of CO2, and the capture of critical nutrients in usable form for food production. Creativity in design mightoffset some of the large area requirements of these systems. For example, algal scrubber raceways might be stacked oneon top of another so long as the proper environmental requirements (light, temperature, and water turbulence) wereprovided. Additionally, these large area requirements might be advantageous in the context of food production: agreenhouse designed to house a large algal scrubber raceway would necessarily have excellent conditions for cropgrowth. Indeed, even larger areas of complex ecosystems have been advocated for remote life support. For example,H. T. Odum (1963) predicted that as much as 2.5 acres (10,000 m2) of ecosystem per person would be required toprovide adequate functions of waste treatment, food production, and atmospheric regeneration. Large ecosystem areasare thought to buffer the closed system environment and provide reliability for life support through the inherent andredundant self-organizational and regenerative capabilities of complex biological systems.

The analysis here is certainly preliminary, and a number of refinements can be proposed. The qualitative assessmentsin Table 1 can be quantified, particularly for the biological systems presented. A detailed design of a system wouldinclude cost analysis, performance, mass analysis and power consumption, and materials and peripheral control andsupport equipment requirements. Also, Table 1 presents only a sampling of the waste treatment technologies, bothphysico-chemical and biological, that exist. Other options might be evaluated, either alone or in different combinationswith other unit processes. Some of the alternatives are better understood than others; for example, the algal turfscrubbers, while exhibiting some attractive regenerative characteristics, are relatively early in their development stages.Additional research and modeling of promising technologies, both biological and physico-chemical, are required beforesound engineering choices can be made. In the analysis presented here, the dynamics of critical atmosphericconstituents – particularly oxygen, carbon dioxide, and nitrogen – have not been addressed for all the biological unitprocesses. This will be an important consideration that needs to be examined for determining the final design in a closedsystem, as each process will potentially act as a source or a sink. Finally, the ecologically engineered designs presentedhere assume temperate climate conditions, based upon the data available in the literature. Research into utilizingbiological systems of algae and plant communities pre-adapted to colder or low light environments may yield smaller,more optimized designs ideal for the M.A.R.S. and eventually for Mars.

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References1. Adey, W., Luckett, C., Jensen, K. 1993. Phosphorus removal from natural waters using controlled algal production. Restoration Ecology

March, 1993, 29-39.2. Craggs, R.J., Adey, W.H., Jessup, B.K., Oswald, W.J. 1996. A controlled stream mesocosm for tertiary treatment of sewage. Ecological

Engineering 6: 149-169.3. Eckart, P. 1996. Spaceflight Life Support and Biospherics. Torrance, CA: Microcosm Press.4. Hall, J.B., Brewer, D.A. 1986. Supercritical water oxidation: concept analysis for evolutionary space station application. In: Aerospace

Environmental Systems: Proceedings of the Sixteenth ICES Conference P-177. Warrendale, PA: Society of Automotive Engineers.5. Metcalf and Eddy, Inc. 1991. Wastewater Engineering: Treatment Disposal Reuse. New York: McGraw-Hill.6. Odum, H.T. 1963. Limits of remote ecosystems containing man. Am. Biol. Teach. 25: 429-43.7. Reed, S.C., Crites, R.W., Middlebrooks, E.J. 1995. Natural Systems for Waste Management and Treatment. New York: McGraw-Hill.8. Taub, F. B. 1974. Closed ecological systems. Annu. Rev. Ecol. Syst. 5:139-160.9. Van Haandel, A.C., Lettinga, G. 1994. Anaerobic Sewage Treatment. New York: John Wiley & Sons.10. Wieland, P.O. 1994. Designing for human presence in space: An introduction of environmental control and life support systems. NASA RP-

1324.

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Design and Resource Requirements for Successful Wind Energy Production on Mars

George James; Gregory Chamitoff; Donald Barker[1999]

AbstractManned Mars missions (and some unmanned precursor missions) are expected to be long duration expeditions that relyon the utilization of in-situ resources to the maximum extent possible. Traditionally, nuclear energy has been suggestedas the power supply of choice for such missions. However, in the event that nuclear power is unavailable, solar energyis the only alternative power source that has received significant study to date. Unfortunately, the periodic and longduration dust storms on Mars drive the need for extensive solar arrays to provide energy needs. This work considersthe possibility that wind energy may provide a secondary power source in an all-solar mission.

Previous studies have shown that a viable Martian wind energy system must be ultra-lightweight, deployable, androbust. In this study, the possibility of meeting these criteria with buoyant wind energy conversion structures is studied.Conceptual designs have been assessed in terms of mass, volume, and power production. Based on current estimates ofsolar energy production efficiency during dust storm conditions, the design constraints on a complimentary wind energysystem are determined. Using system mass per kW-hour as a figure of merit, the feasible wind speeds for three Martianbuoyant wind energy conversion systems (a Savonius, a Darrieus, and a spherical) were determined and compared toprevious rigid designs. The large sizes required for buoyant systems resulted in large turbines producing 10 to 30 kWin 25 m/s winds. The efficient inclusion of the buoyant chambers into the structure was found to be critical, as the massof the balloon is significant. As a result the spherical system was seen to provide the best design. Although it isaerodynamically inefficient and untested, it was found to be feasible in a 29 m/s wind based on an estimated mass of429 kg and 19 kW estimated power.

Martian Wind and Power ProductionMars Energy Needs and ApplicationsA critical aspect of Mars mission planning is the development of efficient and cost effective sources of electrical power.Nuclear power represents the current baseline for long duration habitable missions. However, the technical and politicalissues of emplacing and maintaining a nuclear reactor on the surface might be detrimental to a near term Martianoutpost. Alternatively, solar and wind-generating systems are secondary and tertiary options that utilize the readilyavailable natural resources. Solar power is abundant, inexpensive, and can be used with few safety concerns ortechnological risks. However, solar power is a relatively variable source of energy since Mars is farther from Sun, hasan atmosphere which is prone to seasonal dust storms, exhibits continuous surface dust accumulation, has a substantiallyeccentric orbit, and has a twelve hour night (Haberle et. al., 1993; Landis, 1997; James et. al., 1998). Power output, forexample, from solar photovoltaic cells would be expected to degrade by up to 85% during dust storm conditions. Inaddition, findings from the Pathfinder mission have shown that dust accumulation can obscure as much as 30% of agiven surface over a 100-day period (Landis, 1997). Hence, a manned all-solar Mars mission requires extremely largearrays that have built in routines for maintenance and cleaning.

Wind energy represents an alternative power source that can offset the size of the solar arrays by reducing dust storminterference (i.e., due to the increased scattering of incident light and dust deposition), and nighttime losses (James et.al., 1999). There are two primary lines of query that are needed in order to determine the feasibility of wind powerproduction on Mars. First, in-situ measurements, global monitoring / prospecting, and analytical model developmentare needed to effectively characterize the Martian wind as a potentially extractable energy resource. Second, noveldesigns (and design tools) specifically suited to Martian environmental conditions are required to establish wind energyproduction systems.

Wind energy generation has several niches in Mars mission planning and implementation:

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George James; ETM, Inc / Gregory Chamitoff; NASA Johnson Space Center / Donald Barker; Barrios Technology

1. a tertiary power supply in a primarily nuclear mission to emphasize safety, reliability and mission success;2. a secondary power supply in an all-solar mission to reduce the effects due to dust-storm power reductions and day

/ night cycles;3. a primary power supply in an early Martian settlement with an emphasis on rudimentary in-situ construction capabilities;4. a mobile power supply option to enhance and/or enable long-distance rover operations;5. a cooperative power supply to expand the potential and duration of non-nuclear unmanned precursor missions; and6. a stand-alone power source for the development of portable in-situ fuel and life support resource production

facilities (i.e., methane, oxygen and water).

The work detailed herein is focused on determining resource needs and developing design tools and criteria chiefly toexplore application number two above. However, the results are applicable to each of the scenarios listed.

The current estimates of energy needs for an all-solar mission call for an energy budget of 17 kW continuous energyduring clear day conditions and 9 kW continuous energy during the night (Hemmet et. al., 1999). This includes 1 kWcontinuous energy during the day for rover operations. During dust storm conditions, the daytime utilization needs dropto 16 kW continuous, as rover operations will be curtailed. Hence, the baseline energy needs for an initial outpost areassumed to be 16 + 9 = 25 kW continuous during daylight hours (assuming no energy storage losses). Hence, if aMartian day is assumed to be 12 hrs per day, the daily energy needs are 25 x 12 = 300 kW-hr. However, due to lossesduring dust storms (solar radiation reaching the array may drop to 15% of clear condition values), an all-solar missionmust utilize a solar array eight times larger than needed for the baseline requirements during clear conditions. Hence,the daily solar power produced during clear conditions is 8 x 300 = 2400 kW-hr (James et. al., 1999).

Therefore, for 600 Martian days, total baseline energy requirement is 600 x 300 / 1000 = 180 MW-hr. Additionally,daily rover operation requirements during a clear 12 hour day equals 12 kW-hr or 5 MW-hr for 450 clear days of themission. Also, (Zubrin et. al. 1991) propose that return propellant can be produced on Mars at a cost of 370 MW-hrover the 600-day mission. Hence total energy needs for the entire 600 day mission is 180 + 5 + 370 = 555 MW-hr.Assuming a worst case scenario in which a planet encircling dust storm occurs for 150 Martian days (i.e., the entirelength of the storm season), the total energy production over 600 Martian days with such a production system would be:(450 x 2400 + 150 x 300) / 1000 = 1,125 MW-hr. Therefore, the excess energy production capability is 1,125 – 555 =570 MW-hr. The utility of wind energy production systems in an all-solar mission would be to allow the reduction ofmass (and therefore cost) of the solar arrays needed to meet dust storm conditions.

Environmental Conditions Favoring Wind Energy ProductionWith an atmospheric density 1/75 of the Earth, Mars would at first appear to be an unlikely candidate for wind energy.However, Mars has several advantages for successful wind energy extraction: less gravity (less massive components),large temperature and pressure swings (producing high winds), and tremendous surface relief and low atmosphericthermal inertia (produces consistent wind patterns). Also, the extraction potential for wind power is a function ofvelocity cubed and only proportional to density as shown in equation (1):

where, P is power produced;C is an efficiency coefficient that ranges from .2 to .6;ρ is the density of the Martian atmosphere;A is the swept area of the turbine; andv is the wind speed.

To date, the most direct observations of wind speed on Mars are limited to the Viking landers and from Pathfinder (asof the production of this paper). Wind speeds at these locations were observed to average about 5 m/s, with a peak of25 to 30 m/s recorded at the Viking Lander 1 site (measured 1.6 meters above the surface). As observed from orbit, a

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P = ½ CρAv3 (1)

local dust storm over Chryse Planitia accompanied these peak readings (Greeley, 1982). The Viking landing sites,however, were selected on the basis of mission safety, which precluded the exploration of the more complex or steepterrain that is more likely to harbor the high winds of interest. Other contributors to wind production, variation andlocalization include boundary layer turbulence, surface roughness, local topography and the traveling planetary wavesthat are responsible for short-period fluctuations in the daily averaged surface pressures and temperatures (Keiffer et.al., 1992). Computer models of the Martian atmosphere based on remote sensing data and extrapolations from the windblown sand streaks on Mars have predicted significantly larger values for surface wind speeds (Haslach, 1989). It hasbeen estimated that a well-chosen site could harbor sustained speeds approaching 14 m s-1. Possible sites include thehorseshoe vortices around raised rim craters (as seen by dark streaks), and natural wind channels due to the topographyof hills and valleys (such sites have been used successfully on Earth). Also, regions such as Hellas basin (the lowestregion on Mars) have up to a 44% denser atmosphere (and hence a 44% increase in power). These regions would befavorable sites if high local wind speeds can be identified. Long low angle slopes (as seen on the shield volcanoes orslopes of large basins or plateaus) may produce winds of 25 to 33 m s-1 at approximately 25 meters above the surface.It should be noted that the wind patterns at the Viking 1 landing site were believed to be dominated by this type of slopewind pattern (Zurek, 1992). Recent measurements made at the nearby Pathfinder landing site further support thisconjecture (Schofield, 1997). Additionally, these winds are expected to operate at pre-dawn and during dust storms(times when solar energy is reduced or ineffective).

Our emphasis here is to develop the cursory design and analytical techniques used to support wind energy productionfor a manned mission in light of differential production levels between clear and dusty conditions. The effects of dust(solid, dark, µm diameter particles) and seasonal dust storms directly relate to the work at hand in that they affect boththe performances of solar collecting devices as well as alter atmospheric structure and circulation patterns. Currenttheories (positive feedback models) show that increased levels of suspended dust may cyclically amplify the diabaticdrive for wind production (Keiffer et. al., 1992). Dust storms of various size, distribution and place of origin have beenrecorded in observations since the late 1800’s. Planet-Encircling dust storms have the propensity to directly affectsurface solar energy production for up to several months at a time, and current data show that the storm season may beas long as 1/3 of a Martian year. Providing a solar alternative energy source that work within the boundaries of theaforementioned environment conditions and takes advantage of in-situ resources should prove to be more effective,efficient and cost effective.

Wind Turbine DesignWind turbine designs can be categorized into two different groups – turbines that depend on aerodynamic lift and turbinesthat utilize aerodynamic drag. For the same swept area the power produced by lift type turbines far exceeds the powergenerated by drag type turbines. Drag type turbines have lower rotation rates than the corresponding wind speed withrelatively high shaft torque. Lift type turbines have high rotation rates (linear speed of blades is generally faster than thewind speed) and low shaft torque (Walker and Jenkins, 1997). The lift devices are generally more efficient and thereforesmaller for a given power output. However, the design of the airfoil components for the specific atmospheric and windspeed conditions are important. Also, the drag devices tend to produce power at lower wind speeds.

Wind turbines can be further classified into horizontal axis and vertical axis machines. The horizontal axis or propellertype turbines are more abundant and this technology is highly developed. A Horizontal Axis Wind Turbine (HAWT)typically has blades that can pitch to extract energy over a broad range of wind speeds. However, the blades arecantilevered from the hub, which is at a significant height. Likewise the generator and critical rotating components areat height. HAWT’s must be slewed into the wind unless the resource is relatively unidirectional. Our previous workconsidered an 18-meter diameter Horizontal Axis Wind Turbine (HAWT) that would produce 2.5 kW in a wind speedof 13 m/s. Alternatively, a 30-meter diameter turbine in a 25 m/s wind would produce 28 kW. This design effortsuggested that wind turbines with sizes approaching large utility scale terrestrial wind turbines would be required.However, the chord lengths would be three times the values for similar turbines on Earth. Likewise, the thickness tochord ratio could be expected to be 1.5 times that of terrestrial turbines. Also, the power output (and imposed torquevalues) would be 1/10 the values seen on terrestrial turbines of a similar size. This work utilized a modified code

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developed for terrestrial turbine design. Therefore, the airfoils and blade twist parameters were all optimized forterrestrial not Martian conditions (Ferrell, et. al., 1998).

A Vertical Axis Wind Turbine (VAWT) can accept winds from any direction. Also the primary generator and bearingcomponents are located at ground level. A VAWT with straight lift type blades is called a giromill. Although a straightblade is more susceptible to rotational fatigue stresses, such a devise can change blade pitch to capture low speed winds.Haslach (1989) introduced a concept that called for a 17.25-meter tall giromill turbine situated atop a 21.5 meter landingvehicle with an estimated weight of 175 kg. Using an efficiency value of C = 0.47, as is common on terrestrial turbines,such a 200 m2 turbine could produce 2 kW in a 14 m/s wind and 12 kW in a 25 m/s wind. The authors’ previous workproduced a Darrieus style VAWT with troposkien-shaped blades that was 30 meters tall and produced 14 kW in a 25m/s wind. The turbine was designed using a modified terrestrial wind turbine analysis code. In addition, this workdeveloped a strategy to assess feasibility and subsequently showed that a feasible Martian wind turbine would need tobe of an ultra-lightweight design. Assuming lightweight blades, the mass of such a system was estimated at 944 kg(Hemmet et. al., 1999, James et. al., 2000). The work presented herein is a further extension of the work that exploresa VAWT design specific for Mars utilizing a buoyant configuration.

Buoyant Wind Turbine DesignsDesign Considerations for This StudyA set of Martian-specific design tools is under development to allow trade-off studies, prototype development, resourceprospecting, and feasibility analysis. A VAWT design is the current baseline due to the advantages of keeping the heavycomponents at ground level as well as the simplicity of the rotating sub-systems. A buoyant design is being studied inthis phase of the work. The initial assumption was that buoyant structures would allow large deployable capture areasto be emplaced using pressurized minimum weight structures. Also such gas-filled structures may provide significantstiffness and rigidity with little weight penalty, thus the possibility of using in-situ atmospheric materials. Thermalstructural stresses might be controlled with pressure variations of the inflation medium. The most significant issue withthis design will be the structure needed to transfer the torque to the generator on the ground. However, a buoyantstructure will place the torque transfer structure in tension as opposed to compression. A hollow cylinder is currentlybaselined. This structure may be inflated or foam-filled as needed to maintain stability under torque. The next sectionwill review the lift capabilities of balloons in the Martian atmosphere.

Lift Capability of Hydrogen-Filled Balloons on MarsThe estimation of buoyancy forces due to hydrogen filled structures on Mars is provided by the following equation:

where BF is the buoyancy force;

ρ is the ambient atmospheric density;ρi is the internal gas density;

gM is the acceleration due to Martian gravity; and

V is the volume of the balloon.

The mass of the balloon skin is calculated using a small thickness approximation:

where MB is the mass of the balloon;

RB is the radius of the balloon;

tB is the thickness of the balloon material; and

ρB is the density of the balloon material.

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(2)

(3)

The mass that can be lifted by the balloon is then given by the following:

where ML is the mass that can be lifted by the balloon.

Using the appropriate values for Mars, a hydrogen filled interior, a balloon density of 1250 kg/m3 and a thickness of0.00005 meters, a series of calculations were performed to estimate the lift potential (Figure 1).

It can be seen from Figure 1 that there is a radius below which the balloon cannot even lift itself. For the parameters chosenabove this occurs at a radius of around 12 meters and a corresponding mass of 60 kg. However, reducing the thickness,reducing the material density, or heating the inflation gas can improve this situation. Above this threshold, the mass of theballoon is roughly equivalent to the lift capability over the range of radii shown above. Hence, a buoyant system musteither be able to tolerate the addition of such a mass or integrate the buoyant chamber into the primary structure.

Figure 1. Mass and Lifted Mass for a Hydrogen-Filled Balloon on Mars.

Design of a Buoyant Drag-Type Savonius VAWTA Savonius turbine is a vertical axis machine that uses alternating blades to capture the wind. Drag forces then rotate theturbine. The primary body of the turbine designed herein is assumed to be a ring-stiffened inflatable cylinder with a

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(4)

spherical chamber on the top. The rings also support vertical members that form attachments for the blades. Theconstruction of the blades and the straight vertical members consist of fabric stretched over semicircular frames. Forcesare transferred from the blades to the turbine body via cables attached to the ring stiffeners and the skin of the turbinebody. The torque is transferred from the lower ring to a smaller diameter torque tube / tether via a woven and inflatedfrustum. The hollow or foam filled torque transfer mechanism is assumed to transfer torque but not bending forces. It isalso intended to operate under tension due to buoyancy of the main body. A generator / gearbox is situated on the ground.

An optimization code was written which maximized an objective function based on the power to mass ratio and thebuoyancy. The code was able to modify the radius and length of the cylindrical body, the diameter of the blades, thelength and thickness of the torque transfer cable, and the height of an attachment frustum. The radius of the buoyantchamber was constrained to be equal to the radius of the cylinder plus the diameter of a blade. Skin thickness andsupport radii were generally allowed to vary to assure stresses remained allowable. However, upper and lower boundswere provided which typically constrained these parameters. The number of blades was variable.

The best results were obtained with two blades. These were half cylinders 7.2 m in diameter attached to the sphericalchamber and running the entire length of the 25-meter cylindrical body. This body was 15 meters in radius. The frustumwas 1 meter high which gave it the aspect ratio of a plate. All of these components were assumed to have a skinthickness of 0.0001 meters. An aramid material of density 970 kg/m3 was assumed. The cable was 25 meters long withan outer radius of 1.533 cm and a thickness of 7.633 mm. These torque transfer structure dimensions were chosen basedon the ability to handle the applied torque only. No other stability criteria were used in the estimation of the cabledimensions at this time. The spherical chamber was assumed to have a density of 1250 kg/m3, a thickness of 0.00005meters, and a radius of 22.2 meters. The total system mass was 997 kg including 48 kg of hydrogen inflation gas. Thesystem had a low efficiency (in terms of power per kg) but produced 29 kW in a 25 m/s wind. Figure 2 shows thedimensions of the structure from the front and Figure 3 shows the dimensions from the bottom.

This example showed that the buoyant turbine structures would, by necessity, be a large system, and would produce powerin the 10 to 30 kW range. It should be noted that the optimization process was free to produce a turbine of power outputranging from 0.5 to 30 kW in a 25 m/s wind. However, the buoyancy requirement drove the system to the large size.

Design of a Buoyant Lift-Type Darrieus VAWTThe next design exercise was to modify the previously mentioned Darrieus-type VAWT design to include buoyancy(Hemmet et. al., 1999, James et. al., 2000). The geometric configuration of the previous design was retained. Thisconfiguration included a 30.5 meter tall VAWT with two troposkien shaped blades of 19-meter diameter. Figure 4 showsthe configuration of the turbine including the 0.3-meter diameter central tower. Three blade sections were used withblade sections of 3.2, 2.8, and 1.8-meter chords. The system was assumed to rotate at 75 rpm. The analytical modelsuggested that the device would produce 14.1 kW in a 25 m/s wind with a maximum efficiency coefficient of 0.59. Thiscoefficient was the theoretical maximum, and was likely biased by the lack of appropriate Reynolds number informationfor the airfoils used in the code. The original design used an aluminum tower with guy wire supports.

This study replaces the guy wires with a buoyant support chamber (a balloon) attached to the top of the turbine. Thecurrent study assumes that the blades and torque tube are pressure-stabilized although the tools to analyze the stiffeningeffects have not been produced nor included in the design process. The blade sections were thinned from the previousdesign from 0.0005 to 0.0001 meters. Using an aramid fiber, these blades were still found to support the rotationalstresses. The structural area of the blade section was approximated by the following formula:

where As is the structural area of the blade cross-section;

Cs is the chord of the blade section; and

ts is the thickness of the blade section.

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(5)

Figure 2. Dimensions of the Buoyant Savonius VAWT – Front View

Figure 3. Dimensions of the Buoyant Savonius VAWT – Bottom View.

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This formula was provided in (Sullivan, 1979) for a similarly shaped blade. The length of the each blade was originallyestimated at 48 meters (the circumference of a 30.5 meter circle). This was an overly conservative approach and wascorrected in this work by assuming that each blade was 40 meters (slightly more than the average circumference of a30.5 and a 19 meter circle). The total structural mass of the blades was estimated at 50 kg using the largest chord of 3.2meters. It was found that the performance / mass ratio was not greatly affected by changing the blade thickness from0.0005 to 0.0001 meters (blade mass reduced from 250 kg to 50 kg).

The torque tube from the previous design was chosen to carry the torque and carry the appropriate buckling loads.Therefore, the dimensions were retained and the material was assumed to be an aramid material of 970 kg/m3. Anadditional 15 meters of tube was added to eliminate potential ground interference. The resulting component weighed205 kg and produced a lifted mass of 255 kg. Equations (2-4) were used to estimate the turbine size needed to lift thismass (plus 50 kg of additional mass as a safety margin). The resulting 0.00005 meter thick chamber was 367 kg.Assuming an additional 150 kg of ground support material and 36 kg of hydrogen gas, the total mass was 808 kg. Thiscompares to the 944 kg of the original design. Figure 5 shows the resulting system with the 21.6-meter radius buoyancychamber and 15-meter torque tube extension.

Figure 4. Shape and Dimensions of the Baseline Vertical Axis Wind Turbine.

Design of a Spherical Drag-Type VAWTBased on the information gained during the last two design exercises, the most efficient use of a buoyant chamber is tointegrate it directly into the body of the turbine. A third design exercise was undertaken which assumed the blades weredirectly attached to the spherical shell. These blades were assumed to be flaps that are shaped like a geometric lunearound a spherical body. They were each of 15-degree angular extent that covered one quarter of the surface area. Oneside of each flap was tightly attached to a cable running a great circle. The other side was loosely attached to a series

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of netted cables that were eventually tied into the torque-bearing tether. The flaps were assumed to create pockets thatcaught the wind on the downwind side and closed on the upwind side. The same tether design as was used in theSavonius design was assumed: 25 meters long, 1.53 cm radius, .77 cm thickness, and 13.43 kg. Assuming the samematerial as the buoyant chambers in the other design problems, the required sphere was 17.76 meters in radius and247.72 kg. This size was selected to provide 50 kg of extra lift mass. The total mass was 428 kg including 150 kg ofground equipment and 17.75 kg of inflation gas. Since this is a drag-type turbine it can be expected to have a lowefficiency. Also the spherical shape is not often used for large-scale wind turbines. Therefore, a turbine efficiency of0.15 was assumed. Hence, the turbine produced 19 kW in a 25 m/s wind. Figure 6 shows an artist’s rendering of sucha turbine in use on Mars.

Figure 5. Dimensions of Buoyant Darrieus VAWT.

Assessment of Wind Resource RequirementsThe feasible wind speed selection for the designs discussed above was evaluated using the energy to mass ratio of solarcells as derived during surface dust storm conditions. This process was discussed in previous work (Hemmet et. al.,1999, James et. al., 2000). The mass of a solar-based energy production system for an all-solar mission is estimated toinclude approximately 6 metric tons of solar photovoltaic cells. The energy output for these arrays per unit mass duringdust storms dropped from 400 kW-hr to 50 kW-hr per metric ton. The following relationship was used to definefeasibility of the initial design:

where MW is the mass of the wind turbine; andEW is the energy produced in one day.

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MW (metric ton) ≤ (1/50) EW (kW-hr) = .02 EW (6)

The energy produced by the wind turbine was estimated using equation (1) as the integrated value of power producedover a Martian day. However, a typical approximation is to define this energy based on the maximum wind speed theturbine sees for at least one hour in a given Martian day. A design is feasible if the above inequality holds for a givenwind speed. Equation (6) is solved for the wind speed in terms of the other quantities

Figure 6. Concept of a Buoyant Spherical VAWT in Operation.

Table 1 provides a comparison of the turbine designs listed in this work. The information in this table points out thefact that wind speeds in the mid to lower 20 m/s range will be needed for at least one hour per day during dust stormseason to efficiently (kW/kg) offset solar energy production losses using wind energy.

Table 1. Feasible Wind Speeds for the VAWT Designs Provided

* Assumes turbine is atop a 21.5-meter vehicle.

Ongoing WorkDesign ImprovementsThe designs provided herein represent work in progress and are not final products. In fact these designs were generallyproduced with only mass, power, and buoyancy in mind (structural stress was considered in a limited number ofcomponents). Critical issues such as robustness, fatigue, elastic stability, structural response, dynamic response,pressure stabilization, partial buoyancy, and gas loss must eventually be considered in the design process. Issues suchas airfoil design are needed to produce Mars-specific systems.

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Prototype Development / TestingThe entire concept of an inflatable or partially inflated wind turbine is a new concept and will require significant prototypetesting. Fortunately, there are terrestrial analogues (such as low wind speeds, stratospheric testing, wind tunnel testing)that can be used for some of this work. However, a test object will eventually need to be emplaced on Mars.

Resource Assessment and SupplyThe use of wind power for a Martian outpost will require some significant information on the local, global, and seasonalwind conditions. This means that surface and orbital measurements will be needed. Also, if buoyant structures are tobe used then an on-site source of hydrogen (water) will be most advantageous for maintenance, gas loss, and turbineproduction. This work has not addressed the logistics of the storage and initial supply of hydrogen.

ConclusionsThis project has explored the feasibility of wind power generation on Mars. Three designs using completely buoyantstructures were used. This was an attempt to produce ultra-light, and potentially mobile, systems for Mars. It was foundhowever, that there is a significant mass offset that a buoyant system must overcome. Hence, buoyant systems will tendto be large systems in the 10 to 30 kW range. Rigid or semi-buoyant systems will be required for smaller power stations.The Savonius and Darrieus examples required such large systems that they may not be useful. These designs requireddaily winds of 28 m/s or better for a least one hour to be feasible. The spherical system was the most efficient in itsinclusion of the buoyant chamber. Although it suffers from a low efficiency and a lack of operational experience, it wasseen to be feasible with a 26 m/s or better wind. Such on going research into a solar alternative / complementary energysource that takes advantage of in-situ resources should prove to be more practical, efficient and cost effective for theestablishment of a permanent human presence on the planet Mars.

AcknowledgmentsWe would like to thank the following people: L. Michelle Matanic and Mark Fischer, The Texas Space GrantConsortium; Michael Duke, Center for Advanced Space Studies; Jeff George; NASA Johnson Space Center; Paul Veers,Dale Berg, and Thomas Ashwill, Sandia National Laboratories; and Jason Ferrell, Sherie Hensley, Jesse Le Blanc, JohnRoach, Emir Hemmet, Chi Ngyuen, Kier Wylie, and Bharat Singh, University of Houston.

References1. Beer, F., and Johnston, E., Mechanics of Materials, McGraw-Hill, N.Y., NY, 1981.2. Berg, D. and Rumsey, M., “SLICEIT Input Data Field Definitions,” Sandia National Laboratories internal memo, Department 6225, August 22, 1991.3. Berg, D., “Changes to SLICEIT Input Data Field Definitions,” Sandia National Laboratories internal memo, Department 6225, June 4, 1992.4. Eggers, A., Ashley, H., and Dihumarthi, R., “Considerations of Gravity Effects on VAWT Rotor Configurations Which Minimize Flatwise

Moments and Stresses,” 10th ASME Wind Energy Symposium, Palo Alto, CA, pp. 99-110, 1991.5. Ferrell, J., S. Hensley, J. LeBlanc, and J. Roach, “Conceptual Design and Feasibility of a Martian Power Generation System Utilizing Solar

and Wind Energy,” final report for MECE 4334 Applications from Engineering, Mechanical Engineering Department, University of Houston,Houston, TX, December 1998.

6. Haberle, R. M., C. P. McKay, J. B. Pollack, O. E. Gwynne, D. H. Atkinson, J. Appelbaum, G. A.7. Landis, and D. J. Flood, “Atmospheric Effects on the Utility of Solar Power on Mars,” in Resources of Near-Earth Space, edited by J. Lewis,

M. S. Matthews, and M. L. Guerrieri, University of Arizona, Tucson, AZ 1993.8. Haslach, H. W. Jr., “Wind Energy: A Resource for a Human Mission to Mars,” Journal of the British Interplanetary Society, Vol. 42, No. 4,

April 1989, pp. 171-178.9. Hemmat, A., Nguyen, C., Singh, B., and Wylie, K., “Conceptual Design of a Martian Power Generating System Utilizing Solar and Wind

Energy,” final report for MECE 4334 Applications from Engineering, Mechanical Engineering Department, University of Houston, Houston,TX, May 1999.

10. Keiffer, H. H., Jakosky, B. M., Snyder, C. W. and Matthews, M. S., Mars, University of Arizona Press, Tuscon, 1992.11. James, G., Chamitoff, G., and Barker, D., “Surviving on Mars without Nuclear Energy,” proceedings of the Founding Convention of the Mars

Society, Boulder, CO, August 13-16, 1998.12. James, G., G. Chamitoff, and D. Barker, “Feasibility of the Utilization of Wind Energy on Mars,” to be presented at the 1999 AIAA - Houston

Chapter Annual Technical Symposium, Houston, TX, May 28, 1999.13. James, G., B. Singh, E. Hemmet, C. Nguyen, and G. Chamitoff, “Design of a Wind Turbine for Martian Power Generation,” to appear in the

Proceedings of the ASCE Space 2000: 7th International Conference / Exposition on Engineering, Construction, Operations, and Business inSpace, Albuquerque, NM, February 29 – March 2, 2000.

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14. Landis, G. A., “Mars Dust Removal Technology,” NASA / IECEC 97-97340, NASA, August 1997.15. Magalhaes, J. and Gierasch, P., “A Model of Martian Slope Winds: Implications for Eolian Transport,” Journal of Geophysical Research, Vol.

87, No. B12, Nov. 1982, pp. 9975-9984.16. Paraschivoiu, I., “Aerodynamic Loads and Performance of the Darrieus Rotor,” 2nd AIAA Energy Systems Conference, Vol. 6, pp. 406-412,

Colorado Springs, CO, Dec. 1982.17. Spera, D. A., “Introduction to Modern Wind Turbines,” Wind Turbine Technology: Fundamental Concepts of Wind Turbine Engineering,

ASME Press, NY, NY, 1994.18. Sullivan, W. N., “Economic Analysis of Darrieus Vertical Axis Wind Turbine Systems for the Generation of Utility Grid Electrical Power,”

Volume II - The Economic Optimization Model, Sandia Laboratories, Albuquerque, NM Sand78-0962 1979.19. Walker, J.F., and Jenkins, N., Wind Energy Technology, John Wiley and Sons, New York, NY, 1997.20. Zubrin, R. M., D. A. Baker, and G. Owen, “Mars Direct: A Simple, Robust and Cost Effective Architecture for the Space Exploration

Initiative,” 29th Aerospace Sciences Meeting, AIAA 91-0329, January 1991.

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The “Martian Farmer” – Mining Water from the Martian Regolith

Daniel D. Slosberg[2000]

AbstractMars’ atmosphere is saturated with water. On Mars, saturated means a partial pressure of about 0.1 Pa (compared withabout 600 Pa on Earth (Moran & Shapiro, 2000)). That represents only 1 mg of water in every cubic meter of Martianair. In contrast, every kilogram of Martian regolith (soil) contains up to 40 grams of water (Zent & Quinn, 1997). Asabatier reactor big enough to fuel an Earth Return Vehicle such as that used in Mars Direct requires 9.7 grams ofhydrogen per hour (Zubrin et al., 1997), which can be produced from 88.1 g of water, or as much water as is containedin a little over 2 kg of soil. This paper discusses several methods for extracting water from atmosphere enhanced withwater from the regolith. Methods include the WAVAR system previously discussed by Grover & Bruckner in the 1998Mars Society Conference; freezing water out with a thermoelectric conveyor belt; and using a compression /refrigeration unit to first compress the bulk atmosphere and then freeze out the water. In all cases it is seen that theMartian Farmer is an enabling technology.

The Martian FarmerThe Martian Farmer is an autonomous rover designed to collect water from the Martian regolith and store it for humanconsumption, life support, in situ propellant production, or any number of other uses. The platform of the MartianFarmer is a rover roughly 3 meters long by 1 meter wide and less than 1 meter deep. Such a rover would be coveredwith solar panels supplying it with 150 watts of power. On each corner of the rover, a simple greenhouse dome wouldbe deployed. Each dome would have a radius of 1 meter and cover an area of 3.1 square meters. The bottom of eachdome would be a flexible skirt (think of saran wrap cut in vertical strips) that could traverse easily over small rocks.

The sun heats the air and regolith under each dome, releasing water from the regolith into the atmosphere. Between thedomes and under the solar panels would be a system for extracting water from the moist Martian air. Air is drawn fromunder the four domes and through the extraction system. The water is removed and stored on board. Periodically theon board water stores are pumped or taken to a central storage facility.

According to my calculations (available upon request from [email protected], but beyond the scope of this paper), thearea covered by each dome has 50 g of water available for release. The solar energy available, however, is enough torelease 49 g of water from under each dome every hour (see calculations in the WAVAR section). Please note that warm,wet air leaking out from under the dome will frost out around the dome edges as it comes in contact with the colderoutside air. For this reason, it may be advantageous to have the rear domes slightly (1-2 cm) larger than the forwarddomes so that they can sublimate this frost ring.

Water Vapor Adsorption Reactor (WAVAR)One of the most promising techniques for atmospheric water extraction, the WAVAR reactor adsorbs water from theMartian atmosphere into a bed of Zeolite 3A (a sort of sponge) from which it is desorbed (released) and piped into a storagedevice. This concept has been developed under Adam P. Bruckner of the University of Washington.4,5,6 His teams haverun simulations with 100, 200, and 400 W reactors under various Mars conditions. For a more extensive generaldescription of how WAVAR works, see the Proceedings of the Founding Convention of the Mars Society: Part II.4 For amore detailed technical description, see the 33rd AIAA / ASME / SAE / ASEE Joint Propulsion Conference & Exhibit.6

The energy requirements for the WAVAR device are determined by the amount of water available in the airflow. “Theavailable water is cut roughly in half for every 5ºK drop in the frost point temperature. Comparing how much water isavailable at the Viking Lander 1 site (based on average frost point) with the optimistic assumed value, there is 20 timesmore water available at 213ºK than at 193ºK. The available water quickly becomes minuscule if the frost point

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Daniel D. Slosberg; Martian Water Research, 3010 Applelane Avenue, Kalamazoo, MI 49008-2183, USA; E-mail: [email protected]

temperature begins dropping below 200ºK, and correspondingly, the energy requirements for extraction becomeunacceptably high.”6 WAVAR draws 8000 cubic meters of Martian air every hour.6 This much atmosphere containsonly 8 grams of water.

Under nominal Martian conditions, each dome can release 49 grams of water per hour. (Given the amount of water inthe regolith, the Martian Farmer needs to move 4 meters per hour to release this amount of water.) This much waterraises the frost point of the air reaching the WAVAR from 195ºK to 220ºK, raising the water content from 1 mg/m3 to25.5 mg/m3. The greenhouse makes WAVAR a much more feasible water extraction technology. While studies showmore of the available water can be extracted at higher wattage, only the 100 watt WAVAR fits in the Martian Farmer’s150 watt energy budget. By increasing the available water by an order of magnitude, however, even the low energyWAVAR could capture copious water for storage and later use. Operating for 8 hours per sol, the WAVAR can generate1.6 kg of water, more than half the 2.2 kg/sol needed for a manned Mars mission.4

As you can see, WAVAR is well suited to the Martian Farmer: water available to the WAVAR is increased by an orderof magnitude; and the power requirements of the WAVAR are well within the Martian Farmer’s energy budget.

Thermoelectric Conveyor BeltOriginally discussed at the Think Mars conference at MIT in the fall of 1999,1,b I have not seen the concept of thethermoelectric conveyor belt discussed anywhere in the existing literature, so I shall attempt to give a completedescription here. The main idea: keep a refrigerator door open so it collects water in the form of frost. A thermoelectricrefrigerator consists of two dissimilar metals with a voltage applied across them causing one side to heat up while theother side cools down. It has no refrigerant to lose, and is therefore easier to deal with far from maintenance facilities.Pairs of these metal strips make up the slats of the conveyor belt. They are held closely together along a straight sectionwhere ice forms. One by one the slats peel off and turn a corner while the ice, held together by it’s own crystallinestructure (and perhaps released from the slat by a reversal of current and the consequent heating of the surface of theslat), moves off the conveyor into a collection area. When a large enough sheet of ice has come off the belt, it can becut at a specified length and added to a stack of such ice sheets, stored for further processing at a later time.

The major constraint on the thermoelectric conveyor belt is that the water frozen onto the belt must be thick enough togenerate usable sheets of ice during each sol. The speed of the conveyor belt can control how thick a sheet of ice couldtheoretically get, but there must also be enough water available to reach that theoretical capacity of the sheet.Thermoelectric couples with a temperature difference of over 80ºK are available.8 Splitting this difference keeps thecold side of the thermoelectric refrigerator 40ºK below the ambient temperature, allowing us to potentially extract 99.7%of the water from any atmosphere coming in contact with the refrigeration unit.6 If the conveyor belt is 1 meter longby 10 cm wide, then 1 gram of water evenly distributed on its surface will produce a film 10 micrometers thick. If weforce as much air over the conveyor as we do through the WAVAR, from the ambient Martian air we can remove 8 gramsof water per hour, producing a sheet 0.64 mm thick after an 8 hour sol. Not only is this sheet unacceptably thin, but Ihave serious doubts as to the ability of a thermoelectric refrigerator to remove the water from 8,000 cubic meters everyhour. That is a lot of air to process!

How do things change when we add the greenhouse domes? First of all, the air going over the conveyor belt will beheated, so there will be a larger temperature difference. This can only help the device. Secondly, up to 200 grams ofwater will be available each hour at half the flow rate of the WAVAR, and the airflow rate can be further reduced withoutsignificantly reducing the available water. 200 grams of water is enough for an ice sheet 2 mm thick to be producedevery hour, or 1.6 kg of ice each sol if it is in operation 8 hours/sol. The speed of the belt can be adjusted to optimizeice sheet thickness, but typical values are 0.5-2.0 cm.a,8

As can be seen, the thermoelectric conveyor belt is a technology that is enabled by the Martian Farmer’s greenhousedome. Rough calculations show that the Martian Farmer can supply 1/3 the energy the refrigerator would need to extractthis amount of water.

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Compression and RefrigerationW. Mitchell Clapp described a compression, refrigeration apparatus to remove water from the Martian atmosphereduring The Case for Mars II.9 The atmosphere is drawn into an axial compressor, and then forced through an annulusbetween a cylindrical refrigeration unit and the wall of the device. Given a frost point of 213ºK, Mitchell finds that100,000 cubic meters of atmosphere are required to generate 2.79 kg of water per sol, given that it operates only during6 hours of maximum solar input. According to data from the first Viking lander, the frost point on Mars was notobserved to rise above 205ºK naturally during the first 350 sols of the lander’s stay. I recalculated the numbers for afrost point of 195ºK and found merely 100 grams of water could be generated from the same mass flow of atmosphere.

Even with the Martian Farmer’s greenhouse domes, the frost point of so much air can only be raised to 205 from 195ºK,or 210 from 205ºK.6 Using the 210ºK frost point, 4.4 kg of water are available, of which the compression, refrigerationcycle recovers just under 20%. This yields 770 grams of water per sol, more than enough for a small sample returnmission,10 but only one third as much as would be needed for a manned mission. Unfortunately the Martian Farmercannot supply the enormous amounts of power this method requires.

This technology may be enabled by the Martian Farmer, but it does not show as much promise as the other two methods.By reducing the flow rate, and extracting a larger percentage of the available water, more water may be extractable.Some external source of power is required, however.

ConclusionAll three methods of extracting water from the Martian atmosphere benefit by additional water extracted from theregolith. The WAVAR moves from a borderline possibility an enabling technology. The thermoelectric conveyor belt,something that cannot function in ambient Martian conditions, can function if feed atmosphere from under a greenhousedome. The compression and refrigeration device comes much closer to the design assumption if it draws enriched airfrom the domes. The Martian Farmer can supply not only moist air, but also power for the extraction devices. However,the WAVAR is the only device that can be entirely supplied by the rover’s modest energy budget.

I feel further research into all enabling technologies is warranted, and I would be glad to share my calculations andefforts with anyone working on these technologies. Please contact me by email.

Notesa. In the original discussion, the refrigerator on a Volvo was cited as an example of a thermoelectric refrigeration system, and the concept was

therefore referred to as the “Volvo Device.”b. The equation of thickness, including derivation, is available upon request ([email protected]).

References1. Moran, M. J. and Shapiro H. N., Fundamentals of Engineering Thermodynamics, 4th ed., John Wiley & Sons, New York, 2000.2. Zent, A. P. and Quinn, R. C., “Measurement of H2O adsorption under Mars-like conditions: Effects of adsorbent heterogeneity,” Journal of

Geophysical Research, Vol. 102, 1997, pp. 9085-9095.3. Zubrin, R. M., Frankie, B. and Kito, T., “Mars In Situ Resource Utilization Based on the Reverse Water Gas Shift,” 1997.4. Grover, M. R., Hilstad, M. O., Elias, L. M., Carpenter, K. G., Schneider, M. A., Hoffman, C. S., Adan-Plaza, S. and Bruckner, A. P., “Extraction

of Atmospheric Water on Mars in Support of the Mars Reference Mission,” MAR 98-062, Proceedings of the Founding Convention of theMars Society: Part II, ed. R. M. Zubrin and M. Zubrin, Boulder, CO, August 13-16, 1998, pp. 659-679.

5. Williams, J. D., Coons, S. C. and Bruckner, A. P., “Design of a Water Vapor Adsorption Reactor for Martian In Situ Resource Utilization,”From Imagination to Reality: Mars Exploration Studies of the Journal of the British Interplanetary Society (Part I: Precursors and EarlyPiloted Exploration Missions), ed. R. M. Zubrin, Vol. 91, AAS Science and Technology Series, 1997, pp. 59-73.

6. Coons, S. C., Williams, J. D. and Bruckner, A. P., “Feasibility Study of Water Vapor Adsorption on Mars for In Situ Resource Utilization,”AIAA 97-2765, 33rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Seattle, WA, July 6-9, 1997.

7. “Thermoelectric Cooling,” 1977 Fundamentals Handbook, 1977, pp. 1.24-1.32.8. Incropera, F. P. and DeWitt, D. P., Fundamentals of Heat and Mass Transfer. John Wiley & Sons, New York, 1996.9. Clapp, W. M., “Water Supply for a Manned Mars Base,” AAS 84-181, The Case for Mars II, ed. C. P. McKay, Boulder, CO, July 10-14, 1984,

Vol. 62, AAS Science and Technology Series, 1985, pp. 557-566.10. Clark, D. L., “In Situ Propellant Production on Mars,” 1997.11. Ash, R. L., Assorted Notes on the Properties of Water, 1998.

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12. Belahcene, K., “Martian Atmospheric Modeling,” Graduate Seminar of the Atmospheric, Oceanic, and Space Sciences Department, Universityof Michigan, Ann Arbor, MI, June 23, 2000.

13. Halliday, D., Resnick, R. and Krane K., Physics, Vol. 1-2, 4th ed., John Wiley & Sons, New York, 1992.14. Sittig, J., Tour of Kalamazoo Greenhouses, June 30, 2000.

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Modified Martian Lava Tubes Revisited

R. D. “Gus” Frederick[1999]

AbstractOne of the key elements for successful long-term human occupation of Mars is a viable habitation scheme. Countlessideas have been proposed along these lines from converted landers to inflatable domes. The advantages of mostschemes thus far are that they are location independent, to an extent. The lander lands and the habitation is set up. Inother words, bring the habitat to Mars.

But what if ready-made habitats were available – select locations on the planet which, with minor modifications, wouldeasily serve as a semi-permanent base of operations? These locations could well be lava tubes.

Lava tubes are caves formed by flows of highly fluid lava – a “river” of molten rock flowing from an eruption source,either volcano or fissure. Often as the flow progresses, the tops and sides solidify. If the flow source stops, theremaining lava may pour out, leaving a hollow “tube” of rock. Not all lava flows produce tubes. Sometimes the flowsides form large “levees” as the sides harden, and the top remains liquid.

On the Earth, the author has personally visited lava tubes on the flanks of Mount St. Helens, in Washington State,Central Oregon, and the Big Island of Hawaii, as well as tubes formed by fissure eruptions in Iceland. Many of the lavaflows identified on the planet Mars feature the same characteristics as terrestrial flows, including lava tubes and levees.The main difference is a matter of scale: The Martian features dramatically dwarf their Earth-based counterparts.

This paper offers some speculations on the utilization of these landforms for the construction of viable human habitats.With examples from many lava tube-related features here on Earth, I will demonstrate how their much larger Martianversions could provide a quick, easy and inexpensive way to provide long-term human outposts on the Red Planet.

DisclaimerOne point to keep in mind during the reading of this paper is that I am not trained as a geologist, and that this paper ispurely an exercise in speculation.

Volcanic BackgroundsGrowing up in the Pacific Northwest guaranteed a strong exposure to volcanic terrain. My family stressed anappreciation for the great outdoors, so every summer weekend found us bivouacked in any number of remotecampgrounds in the High Cascades of Oregon and Southern Washington. A favorite family get-away was the MountFuji of North America, Mount St. Helens.

I recall my siblings and I standing at the bottom of steep piles of very fresh-looking pumice overlooking the dark bluewaters of Spirit Lake. My dad related how he had read somewhere that a couple of geologists had suggested that MountSt. Helens would erupt in the next twenty years.9 Imagine the glee my pre-teen brothers and I expressed as we repliedthat we hoped it would happen when we were around!

We got our wish in 1980. The very location of that conversation is now occupied by a gaping hole and view of majesticLoowitt’s guts. Spirit Lake, while still visible, is radically changed. A reminder of Mother Nature’s periodic “rentcollections.”

Another favorite Frederick family destination was located on the other side of Mount St. Helens along the south slopes.Here, some two thousand years earlier, massive fluid flows of pahoehoe lava spilled from the flanks of this Cascade

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R. D. “Gus” Frederick; 401 Silver Street, Silverton, Oregon 97381; 503-873-6216; E-Mail: [email protected]

cone. It is within these flows that the longest and best-preserved lava tube systems in the U.S. are located.

Liquid basalt, heated to more than 1200 degrees C, issued from fissures along the sides of Mount St. Helens. Flowinglike a river down hill, it followed the path of least resistance, and like wax from a candle, began to freeze along the sidesof the flow. The farther from the vent, the cooler the flow became. A crust soon developed over the top of the flow,which tended to insulate the liquid within as it continued downhill.

When the source of magma was exhausted, the remaining liquid lava drained out of the flow, leaving a series of hollowrock tubes to mark its path. Where the crust was too thin, the top of the tube collapsed creating skylight sinkholes.These would provide the future entrances for later human visitation.

The youth of these tubes is the reason that they are so well preserved, as many terrestrial tubes succumb to erosion,earthquakes and sedimentation. Ape Cave, the longest and most famous of the St. Helens lava tubes, traverses a totalof 3,904 meters, (12,810 ft.). In many places within the tube, the ceiling is over 7 m tall, and the cave resembles asubway tunnel in places with its smooth, symmetrical meanderings.

Years later while serving in the U.S. Navy, I found myself in Iceland, stationed at the Naval Station in Keflavic. Duringa tour of the countryside, I was introduced to the Icelandic lava tubes. Many of these were much larger than the St.Helens flows. In fact, one, “Surtshellir,” was used in the 17th century as a hideout for a band of marauding bandits.These tubes were the result of floods of basalt that erupted from fissures, or cracks in the Earth’s crust.

Types of TubesOn the Earth, there are five basic types of lava tubes. While a tube’s structure is dependent upon many factors, thereare two key variables: the surrounding terrain and length of flow.5

Interior Tube:This is probably the most basic lava tube. These are like blood vessels in flesh, conveying the fresh lava to the flowfront. These occur mainly in flood basalts and seldom, if ever, become “caves” since they do not “drain out.” The fluidlava simply stops flowing and hardens like concrete poured into a pipe.

Surface Tubes:Small, individual lava streams sometimes crust over, and then drain to form lava tubes on the surface of the surroundingground. These can sometimes be identified as ridges radiating from known eruptive centers. On Earth, these types oftubes tend to be small in size. Their Martian counterparts could well be much larger.

Semitrenches:A Semitrench tube results from lava overflowing a channel, which in turn builds up walls or “levees.” If the flowcontinues long enough, a roof forms over the flow completing the tube.

True Trenches:These kinds of lava tubes form from a continuous flow that “eats into” the surface, eroding it down like hot waterflowing over ice. The walls of these tubes are composed of the surrounding native rock with a veneer of fresh lava asa glaze.

Rift Tubes:These lava tubes offer the most variety in shape and size, due to the fact that they are dependent upon forces other thanjust the flowing lava. Rift Tubes form as lava flows down an existing rift. As a result, the tubes take on attributes ofthe rift itself. Many times, subsequent flows can pile up on top of one another resulting in multi-level lava tubes.

Eruption of Martian Data

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After Mount St. Helens came back to life, and the data from Mars started to accumulate, I was drawn to the vast volcanicareas of the Red Planet. The orbital views showed many of the same igneous landforms found here on Earth, with thedifference being a matter of scale. The Martian features were much larger, and this included the lava flows andpresumably the lava tubes as well.

The tubes are there, many being identified from the “skylights” of their collapsed ceilings. These tubes are considerablybigger than their terrestrial counterparts. Since we have yet to explore the tubes of Mars, we can only assume that theirinternal structure would be similar to their Earth cousins.

The West flank of Olympus-Mons for example exhibits many landforms that look like collapsed tube skylights.

Along the Southeast flank of Arsia Mons, a series of well-defined tubes and channels are also visible, as well as thesides of the Northern shield Alba Patera.

The relative elevation may eventually be a factor in locating the tube systems for future habitation. These range fromthe “low” 1 km of the Northern Alba Patera area upwards to 10 km and greater on the Tharsis shields. An elevation of“0 km” was defined as that elevation where mean atmospheric pressure at the surface is equivalent to the triple-pointpressure of water, or 6.1 mbars.1 By comparison, my barometer here in Silverton, Oregon is at this moment reading29.78 mbars. Silverton is at about 200 feet above sea level.

Logic dictates that many of the common features of terrestrial lava tubes would also be present in larger versions onMars. What we would have would be long tubes of solid rock. It would be a relatively simple matter to build a colonyin one of these tubes. It would be air-tight and would offer superior shielding against the raw environment of theMartian surface – its thin atmosphere and resulting exposure to the elements – solar flares, radiation, cosmic rays andthe like.

In the second volume of his trilogy of the colonization of Mars, “Green Mars,”11 Kim Stanley Robinson located one ofhis rebel colonist groups within a hypothetical modified lava tube situated in the Martian Southern hemisphere,approximately 64S, 290W, in the Northern Dorsa Brevia region. In his scenario, the colonists blocked off sections ofthe huge tube with bulkheads of a pliable, airtight fabric.

A small dome was erected over one of the skylights to admit light, and the tube was partially flooded, creating alandscape of underground forests, fields, lakes and islands within this enclosed world. Expansion of the colony wasachieved simply by moving deeper into the tube system. He sized his tubes using a 2-to-1 ratio created by thegravitational and other uniquely Martian conditions. NASA observations put the ratio at 10-to-1.13

His tube was wider than its Earth kin by a factor of several hundred, and was 40 km long. This scheme closed off 12km of the lava tube, divided into 1 km segments.

Robinson’s tube system was apparently of a single flow. It featured only one main (albeit large) single tube. On Earth,and presumably on Mars too, multiple lava flows over-lap each other,12 sometimes creating lava tubes on top of olderlava tubes.

In Ape Cave, this phenomenon is present. In fact, one such area is like a large “bubble” located above the main tunnel.By clambering up a side channel, and crawling through a narrow half meter-high opening for about 3 meters, one entersinto a large domed chamber about 7 m in diameter and some 5 m high. (The last time I visited this amazing place thirty-some years ago, it had been vandalized by juvenile spray painters. A trade-off for making such exotic places easilyaccessible to the masses.)

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The Cascade volcanoes including St. Helens are composite cones that alternately erupt silica-rich rhyolites as ash andflows of fluid basalt.

Tubes as Habitats and ReservoirsThe Icelandic tubes I explored tended to have little siltation. This no doubt is due to the fact that the tubes were causedby fissure eruptions that tend towards flood basalts. One of the larger and more popular of the Icelandic tubes isSurshellir, a cave located east of Iceland’s capitol of Reykjavik. Between 1752 - 1757, the King of Denmark, who then“owned” Iceland, commissioned a complete survey of this northern outpost of Denmark that was later published. Whilein Iceland, I purchased the then newly released English version.10 In it, the authors describe in much detail, theirpassage through this cavern:

“There is no doubt that this cave has been inhabited, not by giants but by vagabonds, who escaped toavoid punishment for their crimes, which is probable both from its situation and the following anecdote.In two of the ancient histories it is stated that in the tenth century, a body of thieves took refuge hereand found a safe retreat, because, from superstition, no person would approach the cave. When theywent out to commit their depredations, they had on one side a number of farms, and on the other theland of Arnarvatn, which was always covered with sheep and oxen at pasture. Several other tales aretold of different bands of robbers, who have successively resided in this cavern, which have made suchan impression on the minds of the people, that none of them will attempt to enter it.

“The entrance to the cavern of Surtshellir is gloomy, and runs from N. W. to S. E. but preserves itsheight, which is from thirty to thirty-six feet, while its width is from fifty to fifty-four. Its soil or bottomis uneven, sometimes rising, and at others falling; its partitions are the same, only that there is an equaldistance between them. On advancing, it is perceived that the cavern turns to the south, and afterwardsto the S. W. and W. in proportion as it diminishes in width.

“. . . At the end of this declivity, our travelers found a lake of fresh water, the bottom of which wasfrozen. They passed it with the water up to their knees, and at every step they had additional proof thatthe whole of these caves had been formed by the melting or dissolution of stones. The great channelbeing at length blocked up for some time, and the fire not being to able to find a vent, acted upon thesides, and melted the more dissoluble earth and stones; but before the fiery matter could thus find anoutlet, the great canal had forced its way, and had ceased to have any action on the caverns.”

This last paragraph reveals another aspect of the Earth tubes: They tend to collect water. In fact, most of the lava tubesI have explored on Mount St. Helens have small pools of water at their ends. One tube is named “Lake Cave” for it’slarge “lake” that fills the end of it. Another, “Little Red River Cave,” has a little stream running through its length. Ofcourse, one also finds fine sandy floors of volcanic ash in places from the various “lahar” or mud flows that have spilleddown the mountain over the years.

In Central Oregon, this is even more pronounced. Many of the lava tube caves there have such names as “Arnolds IceCave,” “Surveyors Ice Cave,” and “South Ice Cave.” These caves are not “ice caves” but rather common lava tubeswith seasonal, and sometimes permanent ice deposits.7 What occurs in these caves is that during the winter months,cold, sub-freezing air sinks into the cave’s depths. Owing to the surrounding basalt’s superior insulating properties, theair remains below 0ºC. When the spring thaw occurs, meltwater trickling into the depths of the caves encounters thiscold winter air and freezes out – many times as spectacular formations of stalactites and stalagmites of transparent ice.

Arnolds Ice Cave in fact was regularly “mined” for its ice, which was sold in near by Bend, Oregon during the warmsummer months. These mining operations opened up the cave for the first time in many years, and allowed for itsthorough exploration. After the advent of electric refrigeration, the ice miners were out of work, and the cave slowlystarted to refill with ice. Today, it is totally inaccessible, with only the top of the stairway built in the 1960s poking out

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of the top of the ice.

On Mars, we may also find water, frozen at the end of the tubes. These may in fact be huge natural cisterns. The tubesmay well prove to be a good place to look for water. But where to look for the lava tubes? Robinson’s Tube Colonywas located in the ancient Southern flows. But examination of the Viking orbiter images, as well as those from thecurrent Mars Global Surveyor, show lava tubes and features relating to lava tubes in a host of areas around the planet.6

Building in the proximity of a skylight would allow for the “piping” of sunlight into the tube chamber. Robinson’sInflatable bulkheads could be created to block off large sections. An easier approach may be inflating a single, verylarge balloon within the hollow. This would be like blowing up a toy balloon within a cup. It would conform to theshape of the space, and provide a quick and easy habitable area.

Hornito HabitatAnother feature of lava tubes on this planet are “hornitos.” These are breaches in the solid roofs of an active lava flow,which results in lava spattering out of the holes. The lava soon builds up a “spatter cone” around the opening, and canresult in a “chimney” to the inner tube after the flow stops. And unlike a traditional skylight sinkhole, these openingsdo not have large rubble piles under them. Skylight Cave in Central Oregon has three hornitos in its ceiling, which allowfor outside light to illuminate the interior.

A hornito habitat could provide a sheltered habitat complete with a natural light source and minimal excavation.

A Procedure for setting up Hornito Habitat might involve the following steps:

1. Identify likely candidate lava tubes from orbit. Look for a series of skylight pits and spatter cones arranged in alinear pattern along lava flows.

2. Establish ground contact, and do a preliminary evaluation of candidate tubes.3. Prepare the site. This may involve clearing some minor debris from the skylight cave-in.4. Install the deflated balloon. It would be constructed of a tough, insulated material. The portion where the skylight

fits would be clear or translucent to allow for light transmission.5. Setup the balloon for inflation. This could be done either from supplied compressed air, or by the Martian

atmosphere, with a “slow pump” sucking in Martian air over an extended period of time. (Plants could beintroduced at a point to start converting the carbon dioxide into oxygen.)

6. Once inflated, we move in! Establish an airlock; setup inner walls, partitions, etc.7. Communications antenna and solar power units would be setup outside, with the cables running down into the tube.

With a hornito habitat, the domed-over hornito open would provide a light channel for illumination.

Advantages• A large protected habitable space could be setup in a very short period of time; maybe within 24 hours if one used

supplied compressed air.• The surrounding rock would provide an excellent radiation shield.• The lava tubes might contain frozen water deposits.• By deflating and moving on to other lava tubes, the colony could be semi-portable.

Disadvantages• Location-specific. The scheme would rely on the location of large lava tubes. This would exclude the majority of

the surface.• Lighting might be a problem. Locating directly under a skylight would help, but usually these areas (in terrestrial

tubes) contain huge mounds of debris from the cave-ins that created the skylights. A better solution would be tolocate caves with hornito openings.

• No Martian lava tubes have been explored, so we can only guess at this point as to their viability as shelters. I think

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it would be a good guess!

The idea of making use of lava tubes has been explored by many others. The largest body of research I have encounteredcomes from the Oregon L5 Society. This Portland-based space advocacy group is the Oregon chapter of the NationalSpace Society, has been actively involved in this concept for over a decade, only focusing on lunar lava tubes rather thanMartian ones. Between 1987 and 1988, a series of experimental bases were constructed by the group in CentralOregon2.

In conjunction with the Oregon Museum of Science and Industry, (OMSI) and the Oregon Young Astronauts, OregonL5 contacted the City of Bend, Oregon and received permission for the use of Youngs Cave, a small lava tube withinthe city limits of Bend. They setup a temporary base inside the cave using inflatable structural beams.

The work of the Oregon L5 Moon Base team lead to several conclusions:

• Successful educational simulations of lunar lava tube bases can be run in lava tubes of central Oregon• A lunar base simulation can be performed with personnel of minor age• A permanent site would allow better simulations through use of an evolving base infrastructure on site;• A full-time staff would improve organization, scheduling of activities, and data analysis• Using a lava tube saves money, time, labor, and material.

Since then, the group has been involved with various methods for identifying lunar tubes, both from existing data,notably the Clementine data set making use of the JPL pattern recognition software “JARTool.” Another methodproposed my members makes use of a “Radar Flash Bulb” that would send out a short-lived yet powerful burst of groundpenetrating radar to identify subsurface cavities. Most of the published papers of Oregon L5 are available via thegroup’s Web site listed below.

There you have my flight of fancy. As I said at the onset, I do not pretend to fully understand all the intricacies of settingup a viable habitat on another world. But then again, you never know when some crazy half-baked idea might proveuseful some distant day!

Thank you for taking the time to read this paper.

References1. Batson, R.M., Bridges, P.M. and Inge, J.L., 1979. “Atlas of Mars.” NASA SP-438, Appendix C: Contour Mapping by Sherman S.C. Wu, 131.2. Billings, T. L., Dabrowski, J. and Walden, B. 1988. Evolving Lunar Lava Tube Base Simulations with Integral Instructional Capabilities.

Oregon L5 Society3. Carr, M.H. & Greeley, R. 1980. Volcanic Features of Hawaii: A Basis for Comparison with Mars. NASA SP-4034. Francis, P. 1976. Volcanoes. Penguin Books, London5. Harter, R. & J.W. 1979. Geology of Lava Tubes from NSSAC Geology & Biology Field Trip Guidebook. National Speleoogical Society,

Huntsville, Alabama6. Hoges, C. & Moore, H. 1994. Atlas of Volcanic Landforms of Mars. USGS Professional Paper 15347. Larson, C. & J. 1987. Central Oregon Caves. ABC Publishing, Vancouver, Washington8. Macdonald, G.A. & A.T. Abbott, 1970. Volcanoes in the Sea: Geology of Hawaii, University of Hawaii Press.9. Mullineaux, Donald R. and Crandell, Donald R., 1962. “Recent lahars from Mount St. Helens, Washington.” Geological Society of America

Bulletin 73, 855-869.10. Olafsson, E. & Palsson, B. 1760. Description of the Cavern Surtshellirs. Revised English version published by Bokaut gafan Orn Orlygur

Copenhagen 1975.11. Robinson, Kim Stanley, 1994 “Green Mars,” Trade Edition, Bantam Books. Part 6 - Tariqat, 282-28412. Williams, Howel and McBirney, Alexander R., 1979. “Volcanology,” Freeman, Cooper & Co. Chapter 5, 106-10913. Various. 1976. Mars as Viewed by Mariner 9. NASA SP-329.14. Oregon L5 Society Web Site: http://www.teleport.com/~rfrederi/L5

All illustrations and photos by Gus Frederick, except for Surtshel Icecave graphic from reference number 10 above.

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BiographyOregon native Gus Frederick, (DOB 10/4/54) lives in Silverton, Oregon with his 14 year-old daughter Genevieve. He works as an InstructionalTechnologist for the Oregon Public Education Network, and has been a longtime Mars enthusiast and amateur caver.

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The Martian Time Poll: One Martian Year Of Data

Thomas Gangale; Marilyn Dudley-Rowley[2000]

AbstractThe design of a Martian time-keeping system must be as much a social construct as an astronomical one if it is to gainwide acceptance within the Martian community. Not only must such a system accurately mark the passage of theMartian diurnal and annual cycles; it must also incorporate features that satisfy human social needs. What kind of aclock and calendar do Martians want? The Martian Time Web Site began conducting an on-line poll in September 1998.The Martian Time Poll consists of 25 questions on the basic elements of Martian time-keeping. The results of the firstMartian year of data are reported and discussed.

IntroductionAs we humans establish ourselves as a multiplanetary species, spreading throughout the Solar System during this newcentury, we will leave behind the 24-hour day and the 365-day year. These are cycles that are peculiar to Earth, and asa product of billions of years of evolution on this planet, we are designed to operate by them. Humans will have no usefor diurnal periods that are hundreds of hours long. Similarly, years of 12 or 29 times the duration of the terrestrial year(the orbital periods of Jupiter and Saturn, respectively) will be of no practical use in human affairs. We define a standardunit, the second, in as abstract a way as possible for the physical sciences, but time is a social measurement, first andforemost. We awaken, we work, we eat, and we sleep. We gather to transact business and recreate. We are born, wemature, and we die. How will we measure ourselves, our biological and social needs, according to the passage of timeon alien worlds? What social measurements of time will we bring with us from Earth to make our new homes less alien?To what physical cycles of these new worlds will we adapt ourselves and our new societies?

The study of extraterrestrial social measurements of time has been confined almost entirely to Mars, although systemshave recently been proposed for the Galilean satellites of Jupiter (Gangale 1998). There are a number of reasons whyMars dominated the subject. Mars in one of the nearest planets to Earth, and therefore one on which humans are likelyto establish themselves in advance to voyages to other worlds. Furthermore, in the past half-century, while we havecome to know both Venus and Mars as being less hospitable environments than pre-spacefaring civilization had hoped,Mars has clearly emerged as the best prospect for humanity’s second home. Finally, the cycles of Mars are Earthlikeenough that humans living there will find it terribly inconvenient to ignore them. Living and working by Earth’s 24-hour day, humans would find themselves rising 40 minutes earlier each Martian sol. The Gregorian calendar will beuseless for marking the regular passage of the Martian dust storm season and other annual weather phenomena, muchwhich has yet to be discovered. Martian society will require a Martian clock and calendar for its own specific, localizedpurposes, and will refer to Earth’s Universal Time only as its off-world interests require.

History Of Ideas:The first ideas on Martian time-keeping arose 120 years ago as novelists began to speculate on the possibility of aMartian society. The earliest tales envisioned humans encountering indigenous Martian civilizations. Later, as ourincreasing scientific knowledge of Mars reduced the prospect of advanced forms of Martian life, the trend was towardstories about humans establishing their own cultures on Mars. As incidental minutiae in a fictional narrative, the subjectoften received superficial treatment, lacking the detail to be a complete and useful system (Heinlein 1949, Clarke 1951,Piper 1957). Occasionally, such ideas were based on a faulty knowledge of astronomy (Burroughs 1913, Compton 1966,Lovelock and Allaby 1984). Even when complete systems were described that fairly accurately accounted for the orbitalfactors of Mars, they did not take into account all the time-keeping needs of a human society (Greg 1880).

The first complete Martian calendar was developed by an astronomer who was active in the calendar reform movementin the 1930s (Aitken 1936). Another astronomer invented a complete time-keeping system in the 1950s, going so far as

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Thomas Gangale; 430 Pinewood Drive, San Rafael, CA 94903 USA; Email: [email protected] Dudley-Rowley; OPS-Alaska, c/o 1030 Carl Shealy Road, Irmo, SC 29063 USA; Email: [email protected]

to have a functioning Earth-Mars clock-calendar constructed (Levitt 1954). Not only did these systems accurately reflectthe astronomical phenomena of Mars, but they also took into account many of the sociological aspects of time-keeping.

More ideas on Martian time-keeping have been generated as interest in sending humans to Mars has increased. The Casefor Mars series of conferences included two presentations on Martian time (Mackenzie 1989, Gangale 1997). In the 1990s,roughly 20 authors wrote on the subject. The first commercially printed Martian calendar is available for the current Martianyear (Graham and Elliott 1999). A number of real-time Martian clocks are currently posted on the World Wide Web. Linksto several dozen on-line Martian time-keeping topics are available on the Martian Time Web Site at www.martiana.org, alongwith an in-depth discussion of the systems that are known to the primary author of this presentation.

The Social Construction Of Measurement:A clock or a calendar does more than “tell time,” it “measures the measurers,” it tells the story of those who constructedit and where they came from. Measurement and all who do it are part of human culture (Sydenham 1979, p. 29). Theroots of measurement are in the social process itself – even when it strives to be precise, scientific, and abstract. Thestudy of the history of measurement has demonstrated that the procedures that natural and social scientists use inmeasurement were invented to solve problems of everyday life (Duncan 1984, p. 2). For instance, during the fifthmillennium BC, Egyptian priest-astronomers recognized that the solar cycle heralded the rise and fall of the Nile. TheSun eventually became all-consuming object of astronomical observation, entirely displacing the Moon in importance,which was the primary astronomical time-keeping device for most other cultures. The Egyptians were the first todevelop a calendar based solely on the solar cycle, in which the months were uniform divisions of the year that weredivorced from the phases of the Moon. The scientist usually comes into the picture when the measuring instrumentneeds to be improved. An excellent example is the idea of measuring temperature with a thermometer, a vague conceptthat was made less vague through instrumentation (p. 2).

We take our familiarity with the dimensions of Nature for granted. However, the historical study of measurement hasrevealed that not only are the familiar units of mass, distance, and time socially constructed, but the physical dimensionsthemselves are social constructs and have not always been conceived of in the same way throughout history (p. 14).There has been an evolution toward greater abstraction and standardization, but the fact remains that Nature does notdictate the duration of a second or an hour. The hour has not had a fixed duration over human history (p. 15). Theduration of a second, until recent history, was not agreed upon. Until 1967, “time was bound up with the classicalmechanics of Newton; today it is defined in terms of quantum mechanics, and it is not certain that the two are the same(Danloux-Dumesnils 1969, p. 64).” It is the quest for ever greater precision in measures that led to the discovery of theirillusory character (Langevin 1961). In trying to tie the metric system to Nature, its creators discovered that the systemwas not so natural and immutable. The International Meter is 0.2 mm shorter than the Metre des Archives, based nowon a different standard than a fraction of the arc of meridian (Duncan 1984, p. 22). By 1928, the distinguished physicistP. W. Bridgman wondered whether, “from a strict operationist standpoint, physics was justified in treating as one andthe same concept the notion of length pertaining to ultramicroscopic dimensions, the tactual concept suited to everydaylife, and the optical concept, which is required for astronomical measures of length (p. 15).”

The truth of the matter is that there is an “idealization of the measurement process” which our scientific method is sodependent upon (pp. 120-121). Much of the philosophy of science is a neat ex post facto rationalization (p. 120). Ourdefinitions of physical measurements and our conceptualizations of the architecture of the Cosmos are only as solid asour experience of everyday life. As we move outward into the Cosmos, the new challenges we face as a people redefineour experiences. In seeking to accommodate this process, we will enhance our vision of the universe and invent newinstrumentation to measure it. Time measurement, as any other of our measurements, illustrates how social needs andprocesses influence the framework and conventions of physical measurement. Timing, sequence, tempo, and duration arefundamental features of social events (Duncan 1984, p. 30, citing Zerubavel 1982). It is logical to expect that so long asthose features remain tied to the everyday experiences of current terrestrial life, they will not change much. However,when the everyday experiences of humans ranges farther afield, those features will change. They may even begin tochange as humans start to consider new data from possible human ecological niches elsewhere in the solar system.

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So, it was no mere idle exercise in creativity to gauge the attitudes and opinions of those interested in timing, sequence,tempo, and duration on Mars, even though no actual Martian respondents exist. We do stand on the verge of theacquisition of the Red Planet, with new data being deposited periodically into the human collective consciousnessregarding conditions there. We are becoming Martian. What would those “becoming Martian” have to say abouttemporal measurement on Mars?

MethodsThe history of time-keeping on Earth suggests that people resist change, even small and prudent ones. An example ofthis is the reform of the Julian calendar that was promulgated by Pope Gregory XIII in 1582, which deleted one leapday every 400 years. The Gregorian calendar did not completely replace the Julian calendar for civil purposes until the20th century, and even today the Julian calendar continues to be observed as a religious calendar by Orthodox Christians,despite the fact that it is now off by 14 days. History is replete with examples of more radical reforms that failed. Thusthe authors surmised that the proposed design features in Martian time-keeping that most closely mimicked terrestrialtime-keeping conventions would be the ones best received. As Niccolo Machiavelli observed, “It must be rememberedthat there is nothing more difficult to plan, more doubtful of success, nor more dangerous to manage than the creationof a new system . . . The hesitation . . . arises . . . in part from the general skepticism of mankind which does not reallybelieve in an innovation until experience proves its value.”

At the same time, the authors were mindful of the possibility that the demographic that was likely to respond to theMartian Time Poll, i.e., the Internet community in general and Mars enthusiasts in particular, both representingpioneering populations, would be to some degree more sophisticated and more open to innovation than society as awhole. But to what degree? In this respect, the Martian Time Poll was an exploratory study – polling for responses inorder to frame testable hypotheses and to later be able to ask cogent research questions.

The Martian Time Web Site began conducting an on-line poll on 20 September 1998. The Martian Time Poll consistsof 25 questions on the basic elements of Martian time-keeping. The questions break down into two categories. The firstof these deals with the structural details of time-keeping: how many hours should there be to a Martian sol, how manymonths to a Martian year, et cetera? The second set of questions pertains to the nomenclature of Martian time: shouldwe devise new names for the Martian units of time, and how should we name the sols of the week and the months ofthe year? This presentation reports only the results of the questions that address the preferred structure of a Martian clockand calendar. While preferences regarding the shape of a Martian time-keeping system are certainly influenced byculture, it must be noted that most of human society has become familiar with the 24-hour, 60-minute, 60-second clockand the structure of the Gregorian calendar to varying extents. That which we call Universal Time becomes more trulyuniversal every day. On the other hand, the names that we apply to these common elements of time are heavilyinfluenced by history, language, and culture. Since the Martian Time Poll is entirely in English, the results of thequestions addressing the nomenclature of Martian time will, to some degree, be culturally biased.

Another possible bias in the poll was that some respondents might not understand all the questions. A Frequently AskedQuestions page is available on the web site, which they may or may not have seen.

The data reported in this presentation were recorded on 29 July 2000. At that time, the poll had been open for 708 days,21 days longer than a Martian year. The results are presented in tabular form in the Appendix. For each question, therewas a drop-down list of options from which the respondent could choose. As noted in the Appendix, some questionswere added to the poll at later dates, as were response options to individual questions in some cases.

ResultsFor each of the 14 structural questions, the total number of respondents is reported, then the number and percentage ofrespondents to the highest ranked responses. The lowest ranked responses are reported as a group. Finally, there is adiscussion of the data.

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1. How should the Martian sol be divided?Of 109 respondents, 52 (47.71%) chose to “stretch the second” to make up for the 3% longer length of the Martian solto imitate an Earth day of 24 hours, 60 minutes to an hour, 60 seconds to a minute. This choice was far and away themost popular of the possible selections. The other 57 respondents’ choices were distributed over 16 other configurationsand a choice to respond “No opinion.” “No opinion” was the second-most popular choice with 12 respondents (11.01%)opting for it. The range of response for the 16 other configurations was 0 to 8 respondents.

Refer to Figure 3.1 below.

Figure 3.1

This is a dramatic example of taking what works on Earth and stretching it to make it work on Mars. A recurring themein the data is that if the proposition is straightforward enough, people choose the most conservative option every time.The Martian sol is only 2.7% longer than the Earth day, and simply stretching the familiar units of time by this smallamount would be insensible to most people.

But as physicist Robert L. Forward once insisted in a letter to the primary author, “A second is a second is a second!”Yes, there must always be a universally accepted standard technical unit of time, but it is the society that owns time, notthe scientists and engineers. The Martian unit of time that mimics the standard second might one day be called by amore distinctive name, but that is a subject for a future presentation.

It is instructive to note here that in the late 18th century, the idea of dividing the day by powers of ten was proposed aspart of the metric system, and was discarded early on. Ideas on “metric time” for Mars appear destined to a similar fate.In a sense, the Martian clock is already a fait accompli, since the stretched 24-60-60 clock was used routinely at the JetPropulsion Laboratory during the operation of the Viking 1, Viking 2, and Pathfinder landers on the surface of Mars.

2. Should Martian months always begin on the same sol of the week?Of 109 respondents, 50 (45.87%) chose “No” and 36 (33.03%) chose “Yes.” The next popular choice was “No opinion”with 16 respondents making up 14.68% of the response to this question. The two other selections had a range ofresponse of 3 to 4 respondents.

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Figure 3.2

Once again, the data can be interpreted as being a conservative response. On the Gregorian calendar, the day of theweek beginning a calendar month varies not only from one month to the next, but from year to year for any given month.This chaotic system has been in place since the seven-day week was incorporated into the calendar during the reign ofthe Roman emperor Constantine I. Several attempts at reform in the late 19th and early 20th centuries sought toregularize the calendar and make it “perpetual,” i.e., make each calendar year identical. One proposal, championed byGeorge Eastman of Kodak (Eastman 1926), would have had each month begin on Sunday, and while the InternationalFixed calendar never gained acceptance as a civil calendar, some major corporations continued to use it for accountingpurposes toward the end of the 20th century. Another proposal, the World calendar, would have had months begin onvarious days of the week in a regular pattern that repeated every three months (McCarty). The failure of these reformson Earth suggests that the current system is not a great inconvenience to most people. The response to Question 2 ofthe Martian Time Poll shows that the idea of regularizing each month of the calendar year is so unfamiliar as to outweighwhatever advantages it might have. However, in the response to Question 14, we will see that there is support for somesort of a perpetual calendar for Mars.

3. How many major divisions of the Martian year should there be?Of 109 respondents, 62 (56.88%) opted for four major divisions of approximately 167 sols each. This choice was themost popular choice. The next popular choice was “No opinion” with 21 respondents (19.27%). The third-most popularchoice was eight major divisions with approximately 84 sols each with 18 respondents representing 16.51% of theresponse to this question. The other two choices received 3 and 5 responses.

Refer to Figure 3.3 on the next page.

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Figure 3.3

Much of Earth’s population lives in temperate zones, where the passage of the year through spring, summer, autumn,and winter are obvious, and the response to Question 3 reflects the desire to retain these seasonal concepts on Mars.However, it should be noted that in Earth’s tropical climates, the year tends to be viewed in terms of wet and dry seasons.The human experience of the Martian climate may be similarly bifurcated, as tropical populations may be more affectedby the passage of the dust storm season, or of aphelion and perihelion, while the four seasons marked by the equinoxesand solstices may have a more pronounced effect on the temperate zones.

4. Should the Martian calendar have a leap sol or a leap week?Of the 109 respondents, 81 (74.31%) chose the leap sol. “No opinion” was the next popular choice with 16 respondentsrepresenting 14.68% of the response to this question. The leap week was selected by 12 respondents (11.01% of the response).Refer to Figure 3.4 below.

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Figure 3.4

The leap week is a device used on some perpetual calendars to keep the months and weeks in synchronization. In thecase of the seven-sol week, which was overwhelmingly preferred by respondents (see the responses for Question 7),such calendars vary the length of the Martian year between 95 and 96 weeks, i.e., either 665 or 672 sols. On the otherhand, perpetual calendars that use leap sols must either include sols that do not count as part of a week if 95 weeks arecounted, or periodically shorten weeks to only six sols if 96 weeks are counted, since the Martian year containsapproximately 668.6 sols. There is no example of a leap week calendar in practice on Earth, so respondents preferredthe traditional leap year calendar.

5. Should Martian months be of equal duration, or should they span equal arcs in Mars’ orbit?Of the 109 respondents, 74 (67.89%) selected months of equal duration, a nearly 3 to 1 advantage over months spanningequal arcs, an option chosen by 25 respondents (22.94%). The responses to the three other options ranged from 2 to 5,with “No opinion” leading the pack.


Figure 3.5

Nearly all Earth calendars divide the year into approximately equal segments of time, based either on the 29.53-daylunar cycle or on divisions of the solar year. While it is true that on the Gregorian calendar, February is two days shorterthan any other month, most people do not consider this arrangement to be inconveniently lopsided. The unattractivenessof the equal-arc type of calendar for Mars may chiefly lie in the fact that it results in months containing highly variablenumbers of sols because of the ellipticity of the Martian orbit. For example, in the best-known equal-arc calendar(Zubrin 1993), the 12 months vary from 46 to 66 sols. Moreover, there are only two months that have the same numberof sols, requiring a mnemonic rhyme to be much more than a simple ditty. The difficulties that Martian accountantswould face in dealing with such calendars can scarcely be imagined.

6. How many sols should there usually be in a Martian month? (Equal duration months)Of the 109 respondents, 49 (44.95%) favored a calendar of 24 months containing 28 sols each. This was more than twicethe number of selections for the second-most popular choice, a calendar of 12 month containing 56 sols each, which wasfavored by 22 respondents (20.18%). Other options, which ranged from 0 to 7 choices, ranked no higher than “No opinion.”

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Figure 3.6

Nearly two-thirds of the respondents preferred a year that is divisible either by 12 or 24. The advantage of these twonumbers is that they have a lot of other numbers as factors. For a variety of sociological purposes, the 12-month yearcan be divided by 2, 3, 4, and 6. The 24-month year can be further divided by 8, and since eighths of a Martian yearare analogous to quarters of an Earth year, a standard accounting period, it is possible that the eighth would be a usefulperiod of time for reporting Martian finances. Surely Martian investors will want to count their money about as oftenas their Terran counterparts, and as was remarked early in the Space Age, “No bucks, no Buck Rogers.” Anotheradvantage of calendars comprising 24 months of 28 sols each is that such systems accommodate a fundamental humanbiological cycle. The statistical average of the menstrual cycle is about 28 sols. Since the purpose of a calendar is tomark the passage of time in human terms, the more human factors that are designed into a calendar, the better.

7. How many sols should there be in a Martian week?Of the 109 respondents, 75 (68.81%) chose a 7-sol week, nearly a 4 to 1 advantage over the second most popular choice,a 10-sol week, which was selected by 19 respondents (17.43%). The other four options received between 0 to 6 choices.


As mentioned in the discussion of Question 1, the application of powers of ten to units of time was originally envisionedin the metric system. Revolutionary France enacted a calendar (Weisstein 1996) comprising 10-day weeks, however itwas widely ignored and was eventually discarded by Napoleon Bonaparte. The Soviet Union made several short-livedattempts to deviate from the 7-day week. It is clear from the responses to Question 7 that calendars that retain thisancient scheme are far more likely to gain acceptance.

8. On what sol should the Martian calendar year begin?Of the 109 respondents, 51 (46.79%) opted for the vernal (northward) equinox. An equal number of respondents weredistributed nearly evenly in their choices, 10 to 11 each, for the anniversary of the first human landing, 16 sols after thewinter solstice (January 1st), “No opinion,” perihelion, and the winter (southern) solstice. The remaining respondentsselected from among the eight remaining choices with a range of response of 0 to 2.


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Figure 3.7

Figure 3.8

There was a consensus in favor of marking the beginning of the Martian calendar year with some annual astronomicalevent. There were 73 (66.97%) responses in this general category. This is one of three examples of the respondentsdeviating significantly from the Gregorian calendar (Questions 13 and 14 being the other two). However, it should bepointed out that the Roman calendar originally began on the vernal equinox, and that most of the other calendars of Earthwere tied to a specific astronomical event. Also, it has been standard astronomical practice for centuries to referencethe longitudinal position of celestial objects from the point of the vernal equinox.

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9. What event should mark the first Martian calendar year?Of the 109 respondents, 31 (28.44%) selected the first human landing on Mars as the epoch from which to count theMartian calendar years. This was double the number of choices of the second-most popular choice, the Viking 1landing, representing 15 respondents (13.76%). A close third was “No opinion” at 13 responses (11.93%). Combined,the Viking 1 landing and “No opinion” received 28 choices (25.69%), a bit less than the front-runner. The 11 otherchoices received between 0 to 9 responses, but accounted for 50 (45.87%) responses.


Figure 3.9

Of the 14 structural questions in the Martian Time Poll, this one produced by far the lowest ranking first choice. Theweak consensus expressed by the respondents was to begin counting the Martian calendar year with an event that hasyet to occur.

10. Should the first Martian calendar year be numbered 0 or 1?Of the 109 respondents, 52 (47.71%) favored beginning the counting of Martian calendar years with the year 1, while51 (46.79) preferred a year 0. This was a nearly even split, with a slight preference for beginning with the year 1.Understandably, so simple a question with so few choices has resulted in only 6 (5.50%) respondents expressing “Noopinion.”


This is the one question in the Martian Time Poll that failed to produce a consensus. We humans are rather schizoidwhen it comes to counting. There was no year 0 in the Gregorian calendar, nor do the months begin with a zero day.These social measurements date from an ancient time when the concept of zero was largely unknown outside the worldof mathematicians. Yet in the modern world it has become common practice to begin each day with the 24-hour clockreading 00:00:00.

11. What would be an acceptable leap year scheme for a Martian calendar?Of the 67 respondents, 34 (50.75%) preferred leap years to occur in odd-numbered years and years divisible by 10. Thisis a 3 to 1 advantage over the nearest competitor, the idea of having some scheme that produces three leap years every

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five years, which received 11 (16.42%) choices. A close third was the concept of having no predefined scheme at all,but rather inserting leap years as astronomical observations pointed out the need for them. The “By observation” optionreceived 9 (13.43%) responses, barely ahead of “No opinion,” which received 8 (11.94%) responses. The two otherchoices received 2 and 4 responses


Figure 3.10

Figure 3.11

This was a rather complex question, and it is surprising that it produced such a clear result. The preferred algorithmproduces six leap years every 10 years, which is necessary since the Martian year contains approximately 668.6 sols. Apattern that is repeatable every five years would be a bit more accurate; however, the most popular scheme is thesimplest to remember, while other patterns require more of an explanation.

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12. How should the length of calendar year vary?Of the 67 respondents, 51 (76.12%) chose to have calendar years of either 668 or 669 sols. A far distant second, “Noopinion” received 10 (14.93%) choices, while the six other options received between 0 to 3 choices.


Figure 3.12

Again, people chose the simplest and smoothest solution. Other choices either required calendar years of three differentlengths rather than only two, or required the length of the years to vary by more than one sol. In contrast, the calendaryears on Earth are either 365 or 366 days; there are only two types of calendar years and they vary by only a day.Unsurprisingly, the preferred solution for Mars embodies the same principles.

13. When should the leap sols occur?Of the 55 respondents, 32 (58.18%) favored placing the leap sol at the end of the year. Placing the leap sol at mid-yearwas favored by 7 (12.73%) respondents. Choices for the six other options ranged from 2 to 4.


The response to this question marked a significant departure from Earth’s Gregorian calendar, in which the leap dayoccurs at the end of the second month (February). Only two (3.64%) respondents opted for this scheme on a 12-monthMartian calendar, and in the case of a 24-month Martian calendar, only three (5.45%) people chose to have the leap solat the end of the fourth month. Certainly most people are unaware of the fact that the Roman year originally began withMarch and that February was once the last month of the year, where it made sense to put the leap day.

14. Should the calendar be perpetual (each year occurring on the same sol of the week)?Of the 55 respondents, 28 (50.91%) preferred a perpetual calendar, while 19 (34.55%) chose to have a non-perpetualcalendar. The two other choices received 3 to 5 choices.


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Figure 3.13

Figure 3.14

Again, we see here a response that deviated from current practice on Earth. Ideas on regularizing the calendar in termsof reconciling the months and the 7-day week in a repeatable pattern go back at least as far back as 18th centuryMaryland, when a colonist writing under the pen name of Hirossa Ap-Iccim dedicated a perpetual calendar to GeorgeII (Ap-Iccim 1745). As discussed earlier, the cause of calendar reform experienced a golden age on Earth in the late19th and early 20th centuries, then lost momentum. And so, decade after decade and century after century, none of usknows on what day of the week the 12th of next month will fall without referring to a printed calendar. We seem all butresigned to this inconvenience here on Earth, but the response to Question 14 indicates support for a perpetual calendarfor Mars. Some proposals establish a repeatable pattern over a period of a year or two, while others have every monthinvariably begin with the first day of the week.

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Figure 3.13

Discussion And ConclusionRegarding the division of the Martian sol into sub-units of time, the overwhelming favorite was the 24-60-60 clock, in whichthe hour, minute, and second are stretched by 2.7% to accommodate a diurnal cycle that is just a bit longer than Earth’s.

The results also demonstrate clearly that respondents preferred an Earth-like, modern Western calendar for Mars,consisting of:

• Four seasons.• A seven-sol week.• A leap sol versus a leap week.• Common years comprising 668 sols and leap years totaling 669 sols.• Equal duration months of approximately 28 sols each.

The leap year scheme preferred by respondents was the simplest one to express: odd-numbered years plus decennial years.

The three notable exceptions to the Gregorian calendar were the desire to:

• Begin the year on the vernal equinox.• Place the leap sol at the end of the year.• Begin each year on the same sol of the week (a perpetual calendar).

While starting the new year on or near the vernal equinox is not a feature of the most modern Western calendar, it wasa feature of past Western calendars and non-Western ones extant on Earth today. All of these latter calendars wereinformed by the view that the arrival of spring was the beginning of the new year and the end of winter, appealing toboth the hunter-gatherer and the agriculturist. It is interesting that post-industrial respondents want to import this artifactfrom the Old World to the new.

Placing the leap sol at end of year rather than somewhere else inside the year is a logical idea, but not a new one. Thiswas once a feature of the Roman calendar that is the basis of the Gregorian calendar. What we see here is a popular desireto return to a basic concept that got lost during several thousands of years of priests and politicians tinkering with time.

The quest for a perpetual calendar, one in which all common years and all leap years are identical in relation to the daysof the week, is over 250 years old. The most momentous innovation in time-keeping in the Western world occurredmore than 2000 years ago, when Gaius Julius Caesar, wielding absolute power, took Rome from a lunar calendar to asolar one designed by a Hellenistic Egyptian astronomer. The reform instituted by Pope Gregory XIII in 1582 was anextremely modest one that merely eliminated one leap day every 400 years, a reform that was required for the calendarto keep in step with the seasons. Even so, Protestant and Orthodox Europe ignored this prudent reform for centuries.Social inertia may doom further attempts at calendar reform on Earth, but on Mars we will have a new society, possiblymore disposed to judge ideas on their merits rather than on their history. For the most part, the Martian Time Poll hasshown that the proto-Martian community is firmly wedded to the time-keeping traditions that have served humanity wellon Earth; it remains to be seen how Martian society will evolve new social measurements suited to its otherworldlyenvironment.

Two issues of Martian time that remain to be resolved are:

• Defining the epoch as the year 0 or 1.• Calibrating the epoch to a Gregorian date.

There is no consensus whatsoever regarding whether to begin counting the Martian years with 0 or 1. There are anumber of social measurements that start from 0, while many others begin with the number 1. We speak of “the eleventh

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hour” as a metaphor for time running out, but if 11 p.m. begins that last hour of the day, why don’t we then speak of thebeginning of the day as “the zeroth hour?” If the Martian calendar began with the year 0, calculating the interveningyears between a date before the epoch and after it would be a simple matter. If one was born in the year -22 and diedin the year 11, that person lived to be 33 Martian years old, more or less. Starting with a year 0 would also mean thatthe next century would begin with the year 100. This makes intuitive sense, whereas on Earth many people areconfounded by the fact that the new millennium won’t really arrive until 2001. However, if Mars were to have a year0, would we then speak of the years 0 through 99 as “the zeroth century?”

The weak consensus expressed by the respondents was to begin counting the Martian calendar year with an event thathas yet to occur. This presents a problem to those who would like to begin using a Martian calendar as soon as possible,for without an epoch, we cannot possibly know what year it is! On the other hand, to those who believe that the adoptionof a calendar should be a decision left to future Martians, this is not an issue.

The Martian Time Poll will continue to collect data. Old questions may be reconsidered. New questions may be posed.The educational value of the Martian Time Web Site will be enhanced to better ensure that respondents are wellinformed on the issues. In particular, as consensus develops regarding the structure of Martian time, the questions ofnomenclature will become more relevant, and a carefully considered questionnaire will need to be designed to avoidcultural biases as much as possible on these issues.

Who speaks for Mars? Who has the power to decide the shape of Martian time? The decision is not one to be renderedby Caesar, nor is it conceivable that the Roman Catholic Church will play as important a role as it did in the 16th century.Neither the League of Nations nor the United Nations acted decisively on calendar reform in the 20th century. Does thepromulgation of a system of time for another world fall within the purview of NASA? JPL? The IAU?

We believe that it is the people who are “becoming Martian,” those who somehow have a vested interest in Mars andits acquisition, who will develop the social measurements of Mars. A question that the proto-Martian community mustask itself is whether a clock and calendar could be important symbols of the emerging Martian culture on Earth, whetherthe early adoption of such a system of social measurement could be a factor in the coalescing of a cultural identity, whichin turn could serve to hasten the date of the first human landing. We must ask ourselves whether we who areEarthbound, yet whose hearts are bound to Mars, are Martian enough to take a hand in designing tomorrow.

AcknowledgmentsThanks are in order to Bill Woods, Lance Latham, and Alan Hensel for contributing many of the questions for theMartian Time Poll.

References1. Ap-Iccim, Hirossa. 1745. “An essay on the British computation of time, coin, weights, and measures.” The Gentleman’s Magazine, (London,

July). http://personal.ecu.edu/mccartyr/hirossa.html.2. Aitken, Robert G. 1936. “Time Measures on Mars.” Journal of Calendar Reform 6, 65. http://www.martiana.org/mars/other/aitken.htm.3. Burroughs, Edgar Rice. 1913. The Gods of Mars. New York: Random House, Inc.4. Clarke, Arthur C. 1951. The Sands of Mars. London: Sidgwick & Jackson, Ltd.5. Compton, D. G. 1966. Farewell Earth’s Bliss. Hodder & Stoughton, Ltd.6. Danloux-Dumesnils, Maurice. 1969. The Metric System: A Critical Study of Its Principles and Practice. London: Athlone Press7. Duncan, Otis Dudley. 1984. Notes on Social Measurement: Historical and Critical. New York: Russell Sage Foundation.8. Eastman, George. 1926. “The Importance of Calendar Reform to the Business World.” Nation’s Business (May).

http://personal.ecu.edu/mccartyr/eastman.html.9. Gangale, Thomas. 1998. “The Calendars of Jupiter.” http://www.martiana.org/mars/mst/jupifrm.htm.

10. Gangale, Thomas. 1997. “Mare Chronium: A Brief History of Martian Time.” Pp. 381-393 in The Case for Mars IV: The InternationalExploration of Mars, edited by Thomas R. Meyer. San Diego. Univelt, Inc. http://www.martiana.org/mars/chronium/chronfrm.htm.

11. Graham, James M., and Elliott, Kandis. 1999. “The Millennium Mars Calendar.” Pp. 1031-1033 in Proceedings of the Founding Conventionof the Mars Society, edited by R. M. Zubrin and M. Zubrin. San Diego. Univelt, Inc. http://www.martiana.org/mars/other/millenn.htm.

12. Greg, Percy. 1880. Across the Zodiac: The Story of a Wrecked Record.13. Heinlein, Robert A. 1949. Red Planet. New York: Charles Scribner’s Sons.

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14. Langevin, Luce. 1961. “The Introduction of the Metric System.” Impact of Science on Society 11 (August): 77-95.15. Levitt, I. M. 1954. “Mars Clock and Calendar.” Sky & Telescope (May). http://www.martiana.org/mars/other/levitt.htm.16. Lovelock, James; and Allaby, Michael. 1984. The Greening of Mars. New York: Warner Books, Inc.17. Mackenzie, Bruce A. 1989. “Metric Time for Mars.” Pp. 539-543 in The Case for Mars III: Strategies for Exploration, edited by Carol Stoker.

San Diego. Univelt, Inc. http://www.martiana.org/mars/other/mcknzfrm.htm.18. McCarty, Rick. “The World Calendar.” http://personal.ecu.edu/mccartyr/world-calendar.html.19. Piper, H. Beam 1957. “Omnilingual.” Astounding.20. Sydenham, P.H. 1979. Measuring Instruments: Tools of Knowledge and Control. Stevenage, UK: Peter Peregrinus.21. Weisstein, Eric W. 1996. “French Revolutionary Calendar.” http://www.treasure-troves.com/astro/FrenchRevolutionaryCalendar.html.22. Zerubavel, Eviatar. 1882. “The Standardization of Time: A Sociohistorical Perspective.” American Journal of Sociology 88 (July): 1-23.23. Zubrin, Robert. 1993. “A Calendar for Mars.” Ad Astra (November/December). http://www.drfast.net/mars/Zubrin.html.

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A Master Plan for Mars: From Transporting the First Coloniststo Total Terraformation of the Red Planet

Tanya Harrison[2000]

IntroductionThis paper was done as a part of NASA’s Mars Millennium Project, a project to design a colony for one hundred peopleon Mars in the year 2030. The project began in the spring of 1999 and was made for teams of students in kindergartenthrough 12th grade. I became involved with the project when a commercial aired for it during Star Trek: Voyager onenight.

The point of the project was to create a community for the first hundred colonists on Mars, but I have expanded myproject from just the colony to designing a ship with state of the art propulsion systems to cut down on transit time, andterraforming Mars. This is entitled the Emissary Project, as these 100 intrepid voyagers will be our ambassadors to theRed Planet. Emissary is also the name of the ship transporting the crew to Mars. It will be launched in July 2030 toarrive at Mars, avoiding dust storms and catching Mars at its closest point in its orbit to Earth.

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Tanya Harrison; Kentridge High School, 12033 256th Kent, WA 98031

I. Construction and SchematicsEmissary is approximately 400 meters in length with 17 decks. It was built at NASA’s Spacecraft Yards in the Sea ofVapours on the Moon. A Moon base for scientific study, 3He mining outposts, and an engineering base for buildingspacecraft that would be highly expensive to build on Earth were constructed in 2008. Since the Emissary has nuclearengines, being built on the Moon keeps it from breaking the agreement signed after the Cold War about nuclear anythingin the atmosphere that could be a potential danger. The main parts of the ship, including the body, rings, and engineswere built on the Moon, while the more delicate pieces, such as the computer components, were constructed on Earthand shipped up to the Moon. Materials could be mined on the Moon with which to construct the Emissary; during theApollo Moon missions it was found that Moon rocks contained about 10% aluminum, magnesium and titanium in theform of oxides. These could be refined using hydrogen for reduction. The ship is constructed with titanium andmagnesium alloys, aluminum, and carbon plus some other elements to help shield the crew from radiation exposure.

The ship has two main components: the fuselage and the rings. The fuselage contains the crew quarters and all of theship’s control functions. The rings rotate with the ship at 4 rpm, which generates close to Martian gravity (0.38 g). Therings have an approximate diameter of 115 meters and encircle the entire ship except for a small section in the aft thatcontains the fuel tanks and is the location of the exhaust nozzles.

On Mars, the fuselage of the ship would house crew members until the colony structure had been completed. The ringswould be used for material with which to construct the colony, as well as material from the ship, materials sent to Marsbeforehand, and materials mined on Mars (such as magnesium, alloyed to create metal).

II. Propulsion SystemsThe ship has three types of propulsion systems: nuclear fusion propulsion, ion engines, and pulsed plasma thrusters. Thenuclear engines run by a series of deuterium-3He fusion reactions occurring in a nuclear reactor core. Deuterium-3Hefusion was chosen because it produces the safest by-product, H4 and protons. A magnetic resonance field is generatedon the inside walls of the core to prevent it from becoming unstable. If for any reason the core were to become unstable,it would be ejected and the auxiliary systems would kick in, which are the ion engines.

Instead of using a chemical reaction, such as the combustion of hydrogen with oxygen, these engines ionize atoms ofxenon gas and then, with a strong electric field, expel these ions at high speed. Ion engines have a very high efficiencywhen compared to chemical propulsion methods, but the drawback is that the levels of thrust produced by these systemstend to be several orders of magnitude lower than those produced by chemical methods. This means that to achieve thesame overall change in momentum the engine must operate for longer and must therefore be more reliable than itschemical counterparts.

Nuclear Core

Figure 1.3. The Pulsed PlasmaThruster Operating Principle

(Source: NASA)

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A Master Plan for Mars: From Transporting the First Colonists to Total Terraformation of the Red Planet

Pulsed plasma thrusters are used for orbit and precision positioning. The pulsed plasma thruster system is composed ofpower and control electronics, a high-energy capacitor, electrodes, and chlorofluorocarbon-based fuel bars. The use ofsolid, non-toxic propellant bars eliminates fluid propellant systems and the complex ground handling used in traditionalpropulsion methods. The fuel is ablated, ionized, and accelerated electromagnetically from the pulsed plasma thrusterduring a high-voltage capacitor discharge across the face of the fuel bar. The pulsed plasma thruster system offers asimple, lightweight, microthrust propulsion capability with minimal spacecraft interface requirements for spacecraftattitude control, orbit-raising and translation, and precision positioning.

III. Computer Systems

Figure 2.1 (Source: NASA JPL)

Emissary uses a holographic data storage system. A hologram is a photographic record of the spatial interference patternformed by the mixing of two coherent laser beams. One beam carries the spatial information, called the “object” beam;the other is distinguished by its particular direction of travel, called the “reference” beam.

Holographic data storage promises fast access times, because holograms encode a large block of data as a single entityin a single write operation, and reading the hologram retrieves the entire block simultaneously.

The HDS system can hold an incredible amount of data when compared to other methods of data storage. For example,the new hard drives that just came out are 10-gigabyte drives. A holographic cube the size of a sugar cube can hold 10gigabytes per cubic centimeter. A block of optical media the size of a deck of cards would hold a terabyte of data.

Fortunately, there are many advantages to holographic data storage. Reading out of images instead of single hits seriallyprovides a huge improvement in bandwidth. Also, the ability for light to be launched through space and easily deflectedwill eliminate the need for rotation of the medium. HDS is a convenient way to address a storage medium in threedimensions while only scanning beams in two.

HDS also uses a relatively small amount of components to run, minimizing the space they take up. The maincomponents are:

• Spatial light modulator to properly shape the object beam• Optical beam scanner to point the reference beam

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• Detector array to convert the reconstructed data output object data into an electronic bit stream• Electronics to control the entire process and condition the input / output electronic information• Sufficiently powerful laser to overcome the optical losses of the system

Holography records information from a three-dimensional object in such a way that a three-dimensional image maysubsequently be constructed. Holographic memory uses lasers for both reading and writing the data blocks into thephotosensitive material. Recording the interference pattern between a carefully modulated coherent wave front and areference beam on a photosensitive material forms a digital hologram.

Iron doped lithium niobate is currently the medium used for holographic storage. Unfortunately, it has two mainfailings: a destructive readout of the data and a relatively low sensitivity. Currently, it is the only material that has theoptical quality that is critical for a system application. Hopefully by 2030, we will have found a better medium forholographic storage. Superfluid helium (4He) is beginning to look promising for this type of data storage, and its use isstill being tested.

The computer core contains 10,000 holographic data storage disks. Each of the disks is 9 x 6 x 1 centimeters and hasa volume of 54 cm3. Each disk can hold a terabyte of data, which gives the computer a maximum storage capacity of10,000 terabytes. To break that down: the computer has a capacity of 18,518,518,519 bytes per cubic centimeter, or18.518518519 gigabytes. With 54 cm3 in a disk, each disk holds one trillion bytes, or a terabyte. This means that thetotal storage capacity of the computer is 1016 (ten quadrillion) bytes! This is equivalent to the amount of data that canbe held on ~71,429 floppy disks. The core can process approximately one trillion calculations per microsecond. Toovercome the problem of computer error, computer redundancy is used in the ship’s computers. The computers mustreach consensus before they can issue a command.

Emissary has two-dimensional graphical interfaces and voiceprint technology. The verbal controls are multi-lingual andwork with English, Russian, Japanese and French. Either of these methods can be used to access the ship’s library,which contains all the data we have on Mars, including photographs, videos, maps, and writings. Even H. G. Wells’The War of the Worlds is stored in the library. Also stored are all of the crew profiles and mission data. Profiles includebiographical information, educational record (degrees, etc.), department aboard the ship, and closest living relatives incase of emergency.

Security systems on board have two purposes: to protect the computer files and to keep crew members where theybelong. Sections of the ship that are restricted access areas are protected by iris scanners, as are encrypted / restrictedfiles. Also, no one can open your letters from home except for you (and the commanding officer, if need be).

Iris scanning has many advantages over fingerprints and voiceprints. The iris has 266 measurable characteristics, whilethe fingerprint only has about 35. No two people have the same irises; even your right and left irises are different.Fingerprints can be manipulated (and even removed), and voices can be imitated, whereas irises cannot. Iris scanningworks with glasses and contacts, with an identification time of two seconds. Currently, this kind of scanning requires512 bytes per eye for storage. Fingerprints require over one thousand bytes per print. For the entire crew, it would onlytake up 102.4 gigabytes of the 10,000 terabytes in the ship’s computer (if data were stored for both eyes).

The computer also has an advanced encryption standard from IBM called MARS. MARS is a shared-key block cipher,with a block size of 128 bits and a variable key size, ranging from 128 to over 400 bits. It is a resilient system, becauseall of the known cryptanalytical attacks (including linear and differential cryptanalysis) require more data than isavailable (2128), making these attacks would be impossible against MARS. MARS is also more secure than othersystems because it was designed using a mixed structure, where the top and bottom rounds were designed differentlythan the middle rounds. This was done because different parts in a cipher play different roles in assuring security. Topand bottom rounds usually have a different role than the middle rounds in protecting against cryptanalytical attacks.

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The computer has a Facial Action Encoding System on it. This is a system that analyses the facial muscles to deducethe emotion of the user, and reacts with a series of pre-programmed actions. For example, if the computer senses youare nervous or impatient, it speeds up its response time. FACS (pronounced “faces”) deals with what is clearly visiblein the face and ignores “invisible” changes or visible changes too subtle for distinction. FACS does not include skincoloration, facial sweating, tears, rashes, pimples, and permanent facial characteristics in its analysis. Specific actionsare described: the movements of the skin, the temporary changes in the shape and location of the features, and thegathering, pouching and wrinkling of the skin. FACS has action units (which are numbers) that have assigned namesto them and the muscular basis of these actions, as seen here:

Single Action Units

The table indicates where they have collapsed more than one muscle into a single Action Unit from a single muscle.(Source: http://www.cs.wpi.edu/~matt/courses/cs563/talks/face_anim/ekman.html)

As an example, FACS would detect the action units 6 (cheek raiser), 12 (lip corner puller), 14 (dimpler), 20 (lipstretcher), and possibly 44 (squint) and deduce from these that the user is smiling.

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Emissary’s computer has a sort of artificial intelligence: it can answer questions (i.e., “Are the deuterium injectorsperforming within tolerable parameters?”) and process commands (i.e., “Access the internal sensors.”). The computeruses the bottom-up approach. Bottom-up is the top-down approach (a heuristic IF-THEN method, which uses decisiontrees) plus induction and many of the subtle nuances of human thought. It codes known human behaviors and thoughtpatterns into the computer as symbols and instructions. It learns from what it does, devises its own rules, and createsits own data and conclusions. The computer adapts and grows in knowledge based on the network environment in whichit lives.

IV. Flight Plan D

Figure 4.1. Flight Plan of Emissary

The flight plan includes a Venus swing-by, which involves a Hohmann transfer to a specific point in space near Venusso that the trajectory of the ship is redirected toward Mars. The main problem with this is that the ship will come towardMars at a high velocity, which could require a large amount of fuel to slow down the ship at Mars. The way out of thisproblem is to bring along a heat shield and use the upper atmosphere of Mars as friction to slow down to a gradual stop

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as opposed to using up propellant (Emissary has an inflatable heat shield, stored in the front of the ship). Emissary willalso slow down to a calculated speed before encountering Venus so that the velocity of the ship after encountering Venusis the same as it was before the encounter.

The Emissary has a somewhat complicated landing procedure. When the ship reaches Mars, it uses the pulsed plasmathrusters to insert itself into a parking orbit. The clasps that connect the rotating rings to the ship are released, and therings are jettisoned down to the planet near the landing spot at Utopia Planitia. This leaves the crew in microgravity forthe descent. The ship angles itself so as not to burn up in the atmosphere, coming in at a slow speed. Four large landingbraces, two at the fore and two at the aft, are deployed. Since there are obviously no runways on Mars and most of thelandscape is rock-strewn, traditional wheel landing gear would not be appropriate for landing on Mars. As the shiplevels out horizontally, retrorockets on the bottom of the braces are activated to slow the descent even further. The finallanding is a bit rough, so the crew is instructed to hold onto something. After landing, ramps are lowered from theairlocks on the bottom deck (the same place that the landing braces are stored during the trip) so the crew can get outof the ship and set foot on Mars.

V. ComplicationsNow, there are a few problems with this ship. First of all, it has an exhaust velocity of 26,400 km/s, which is 95,040,000km/hr. With speeds this high, gravitational forces would crush the crew long before reaching Mars (don’t forget, shipsusually travel twice as fast as their exhaust velocity). Also, it would take awhile to accelerate, and if one were cruisingalong at 95,040,000 km/hr in a place with no friction, deceleration would take an incredibly long time. So, one couldnot accelerate all the way for a trip as short as to Mars. And if one wanted to go at top speed, they’d need to get a holdof Captain Picard and the Enterprise and ask how those inertial dampers work.

As you go faster, your length begins to shorten. The Emissary is 400 meters long, but travels at ~6.8% the speed oflight. Since the ship would shrink by a factor of gamma, which is:

where v is the velocity of the object and c is the speed of light, the ship’s gamma equals 1.003900073. Take the originallength of the ship and divide by gamma. Emissary would end up being 398.4460314 meters long. Everyone andeverything aboard would also shrink by the same gamma, but it is not always the same amount. The Emissary shrinksby about two centimeters, but when I put in my height:

I only shrink by less than a centimeter. And the trip to Mars shrinks by 304,577.85 kilometers. This means that if theship were going at 26,400 km/s, it would take approximately 50 minutes to reach Mars.

When one goes faster, time slows down as well. Going at 6.8% the speed of light would cause a small time dilation.The equation for time dilation is:

where ∆t is the observer’s time (the time on Earth) and ∆t1 is the time on the moving system (the time aboard the ship).Using this equation, we find that the time dilation if the ship were going at top speed the entire trip would only be 11.47seconds.

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(1.1)

150 cm/γ = 149.4173 (1.2)

(1.3)

The ColonyI. Construction and Schematics

Figure 2.1. Colony Base at Utopia Planitia(Photograph of a model of the base)

The colony structure, built at Utopia Planitia, is circular in shape with a diameter of approximately one kilometer. Theinner section holds all of the colony’s control functions and duty stations; the outer ring holds crew quarters andrecreational facilities, such as the gymnasium. It is composed of aluminum, titanium, and magnesium alloys, and a fewvarious carbon compounds to protect against the sun’s deadly UV rays. The colony will be constructed underground tohelp protect the crew from radiation as well. Mirrors are used to reflect sunlight down to the base’s greenhouse to allowfor photosynthesis, as well as a place for the crew to walk around in “natural” light somewhat safely.

The base has safeguard systems in case of emergency. If a containment breach occurs, alarms go off and the computershows the location of the breach. All crew members put on environmental suits and wait for the breach to be repaired.If a quarantine situation were to occur, everyone would change into environmental suits and evacuate the structure. Thestation is then flooded with ozone to clean it.

II. Life Support System and Aspects Requiring a Technological SolutionThere are many challenges with creating contained biospheres on Mars. The first, which was noted in the Biosphere 2project, is the amount of power used per person. The world average primary power is ~2 kWe/person, and the worldaverage electrical power is ~0.2 kWe/person. In Biosphere 2, a whopping 100 kWe/person was used! A settlementdesigned by the Obayashi Corporation in Japan for 150 Martian colonists estimated a power consumption of ~50kWe/person, including life support and routine use, energy for mobility, expansion and emergencies. Either way, this issignificantly more power than the amount used regularly on Earth. This is why the station has solar and wind power:

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to regenerate enough power for the 100 colonists. Since Mars will not be exactly hospitable to the colonists when theyfirst arrive, there are aspects requiring a technological solution:

Maintenance of habitable enclosureReplacement of leaked gasesThe loss of the artificial atmosphere to the outside on Mars would not be a severe problem since Mars has an atmosphereof its own that can re-supply the base. If the artificial atmosphere in the structure leaks out due to a breach, particlegenerators will generate the atmosphere and the computer will notify the crew as to when the amount of oxygen,nitrogen and trace gases is suitable for them to return to the base (or to remove their EV suits). If noxious gases leakinto the structure, or the nitrogen / oxygen / trace gases levels in the atmosphere become harmful to the crew, the entireatmosphere will be vented out in the section with the problem (after that section has been sealed from the rest of thestructure to prevent total depressurization) and generators will be reconfigured to correct the malfunction. The entirestation’s atmosphere is vented if the problem becomes widespread. The artificial atmosphere system must be carefullymonitored and controlled, since minor imbalances can have significant affects on all of the base’s inhabitants – humans,plants and microbes.

Radiation protectionThe radiation that reaches the Martian surface is approximately 2,000 times that of Earth! The materials the station iscomposed of give radiation protection, as well as the location on the planet (Utopia Planitia is located in what appearsto be an ancient ocean bed, giving it a low elevation, and the colony is underground there).

Temperature regulationTemperature regulation systems must be installed due to the lack of natural weather to keep circulating fresh air.

Hydrological cycleAn artificial hydrological cycle must be created due to the inability to form clouds and rain inside of an enclosed habitat.After transpiration by plants, water is condensed from the air and piped back into the greenhouse sprinklers and thecolony’s “watershed.” Aquifers will be placed near the colony to reach Mars’ ground water (if there is any), andmachines will be placed outside to create water from elements in the Martian atmosphere. Water is recycled / conservedas much as possible; the colony has employed sonic showers, which would be more sanitary and require only electricity(to generate the sound waves) to run.

Management of horticultureThe hydroponics bay will house the agricultural department’s food and plant species. Most will grow under blue or redlights, since experiments have shown that blue and red lights promote larger crops with less of the wasted, inedible partsof the plants. In fact, radishes, potatoes, and wheat grown very well in NASA’s simulated Martian soil. When thecolony (and eventually colonies) becomes larger, biodomes will be built. The biodomes are composed of H2O andsilicon. More plant life can be grown in these biodomes to support more humans on Mars. Recycled water (from humanwaste) would be used as rain inside these domes.

Artificial atmosphere maintenance and air revitalisation (Equations: Fogg, Martyn J. Terraforming: EngineeringPlanetary Environments. 1995 by SAE International, Warrendale, PA, pgs. 44-46, 51)

Oxygen is produced from H2O by electrolysis:

CO2 scrubbers will be in place like on nuclear submarines, to keep the levels of carbon dioxide in the atmosphere at asafe level. The CO2 scrubbers run by precipitating calcium carbonate in two steps:

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2H2O —> 2H2 + O2 (2.1)

CO2 + 2NaOH —> Na2CO3 + H2O (2.2)

After removal from the air, CO2 is reduced by using the Sabatier process, converting the CO2 to methane at 370°C inthe presence of a catalyst:

Combining respiration, electrolysis, and the Sabatier process, the overall reaction is:

Food and water are consumed, producing oxygen and excess waste methane. When combining the Sabatier process witha hydrazine-based nitrogen generation subsystem for maintenance of cabin pressure in the base:

Hydrogen is provided for CO2 reduction, conserving water. The dispensable atoms from this are carbon and hydrogen.The advantages of using the Sabatier process (over the option of the Bosch process, which is used in some systems tominimize the loss of hydrogen) are that the machinery employed is lightweight and the reduction of CO2 occurs at ahighly efficient rate.

Fans circulate the air inside of the base. There is also an emergency oxygen injection system in case of breaches incontainment. The oxygen generators run by using hydrogen, brought to Mars from Earth (and eventually mined fromthe asteroid belt), and mixing it with carbon dioxide from the Martian atmosphere to create oxygen (and methane rocketpropellant, according to Zubrin’s Mars Direct plan).

Emergency life supportThere are backup emergency environmental controls and life support in case of emergency (i.e., if the solar panels thatcollected the energy to run the base were hit by meteorites, which hit Mars about 1000 times a day).

III. Rovers and NIMFsOutside of the station is the rover bay, which holds the colony’s twenty solar-powered land rovers, each with a five personmaximum capacity. They must be checked out to use them and communications is kept open with the base so that ifanything happens, they know where the crew members are and how to find them. Each rover is equipped with a microchipthat keeps track of where the rover is on the planet. Thecoordinates are transmitted back to the command center.

The rovers, unlike those used on the Moon when Apollo11 landed there, are totally enclosed and pressurized,sort of like a high-tech Martian car. If the crew wantedto ride in the rovers without environmental suits on, theycould do so, but they’d need one to leave the rover, sothe rovers are equipped with two environmental suits.The extra suits could also come in handy if there was acontainment breach in a crew member’s suit.

To provide quick, efficient access to other parts of Mars,the base is equipped with two NIMFs, which stands forNuclear rocket using Indigenous Martian Fuel. The

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Na2CO3 + CaO + H2O —> CaCO3 + 2NaOH (2.3)

4H2 + CO2 —> 2H2O + CH4 (2.4)

(CH2O) + H2O —> CH4 + O2 (2.5)

N2H4 —> N2 + H2 (2.6)

Figure 2.2. NIMF Design

idea of NIMFs was invented by Robert Zubrin. NIMFs use raw carbon dioxide from the Martian atmosphere as apropellant. The carbon dioxide is heated by an onboard nuclear thermal rocket, or an NTR, creating a hot rocketexhaust. NTRs don’t convert their heat into electricity, allowing the NIMFs to be both lightweight and small. Since thepropellant is raw carbon dioxide, chemical synthesis gear is also eliminated. NIMFs have total global range becausethey make their own fuel After more colonies has been established on Mars, NIMFs could be used for cargo transport,and for transporting people from city to city.

NIMFs could also be used to transport crew members to various spots on Mars to set up weather monitoring stations tokeep track of the infamous dust storms. Mars Global Surveyor (or the newer version of it in 2030) would also assist bytransmitting data to the base.

IV. The Space Elevator

Figure 2.3. Space Elevator

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Figure 2.4. C-60 Molecule(Source: www.carbon60.co.uk)

Figure 2.5. Buckytube(Source: www.carbon60.co.uk)

Also outside is the space elevator, which goes up to an asteroid with a space station constructed on it where the yearlysupply shuttle docks. The elevator car travels up an areosynchronous elevator rail, which is held in place by anareosynchronous counterbalance. A carbonaceous asteroid a couple of kilometers in diameter, brought to Mars to serveas a tether anchor, could provide all of the carbon 60 required to construct an orbital tether (a buckytube). The tethercould be created by automated machines on the surface of the asteroid and slowly spun down to the surface.

Figure 2.6. Tall Tower Concept(Source:Source: Rawlings, R.P.: NASA Artwork by Pat Rawlings, Science Applications

International Corporation, prepared during the Advanced Infrastructure Workshop on GeostationaryOrbiting Tether “Space Elevator” Concepts. NASA Marshall Space Flight Center, June 8-10, 1999.)

The tether would need to be approximately 34,000 kilometers long to have the elevator car reach the (asteroid) dockingport in areosynchronous orbit. It extends down from the asteroid and is attached to a 30-kilometer-tall tower that isconstructed on the surface above the base. The rover bay is to one side of the tower and the NIMF bay and launch siteis to the other side. At the base of the tower is a high-power laser that propels the elevator car, which is a lightcraft.Lightcrafts work by using a high-power laser to heat up the air behind the craft until it explodes behind it, propelling itup the tether. These cars are lightweight and relatively fast; the car would take about five hours to reach the asteroid.

Terraforming MarsI. Planet ShiftingThe Emissary’s crew’s biggest undertaking will be to move Mars from a position 228,000,000 kilometers away fromthe sun to a position 211,054,200 kilometers away from the sun. This will be done to heat up the surface, melt some ofthe permafrost to give Mars some of its water back, and to assist in the release of volatiles trapped in the Martianregolith. They will use a multi-swingby technique to move the planet.

Perhaps the least daunting of planet shifting techniques, since it involves a process that has already been experimentedwith on a small scale is what Oberg called the “multi-swingby” technique. When two bodies in space pass close enoughfor significant gravitational interaction, an exchange of momentum occurs. An example of this is when the Voyagerprobes encountered Jupiter to accelerate and deflect their trajectory toward Uranus and Neptune. But since the massratio between Jupiter and the probes is so minute (~10-25), the planet experiences no discernible deflection at all.However, close gravitational encounters with larger bodies similar in mass to the planet itself would have a large effecton the planet’s orbit.

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Oberg conceived of the possibility of engineering repeated encounters with a succession of ~100-km-wide objects, whichhe called “cue balls,” each one giving the planet a small and harmless nudge in the desired direction. Although the “cueballs” are now more manageable in size, the scheme becomes extremely complex. You would have to have more than1000 cue balls that encountered the planet every three hours or so for about 100 years, and the mass of all of the cue ballswould be enough to amount to a moon-sized planet on their own – not a readily accessible source (most likely from theasteroid belt). Another difficulty is that as each mass gives the planet a small shove, it receives, in relative terms, apowerful kick in return. The planet would scatter the cue balls far and wide, further complicating the problem.

The multi-swingby idea does have a number of points in its favor, though. For a start, it is a highly efficient way oftransferring both momentum and kinetic energy; the only energy dissipating is a tiny amount converted to heat in thetidal bulges raised during the encounter. A possible way to remedy the scattering problem is to have the cue ballsencounter another much larger object, such as Jupiter, which will then send them back into the inner Solar System.Thus, the vast reservoir of angular momentum and energy embodied in the orbit of Jupiter could be tapped. Setting thewhole process in motion might be done easily if we conceive a hierarchical approach, where small objects deflect largerones, which deflect still larger ones, and so on.

Obviously, the moving of Mars would not be completed soon after the colonists arrive, but they could be the ones to setit in motion.

II. Atmospheric Enhancement and Surface HeatingAfter writing about the colony performing planet-shifting to terraform Mars, another less drastic measure was decidedon to help give Mars back its water and atmosphere, based on the work of Carl Sagan, Chris McKay, and Robert Zubrin.Instead of moving Mars closer to the Sun, a runaway greenhouse effect will be induced on Mars – but not to the pointwhere it gets as bad as Venus (or some may say as bad as Earth).

Table 3.1. Mars AtmosphericComposition

Mars has a thin atmosphere, but not nearly a breathable or protective one. The current Martian atmosphere is dominatedby carbon dioxide and nitrogen, together making up over 98% of the atmosphere. Oxygen is only .07%, and ozone isthe scarcest, with a mere 0.04 to 0.2 ppm.

These are the levels of oxygen currently on Mars. It shows the volume percent of oxygen compared to the atmosphericpressure. As you can see, anyone walking around on Mars without a pressure suit would suffer from nearly immediatehypoxia. There are scarce amounts of oxygen in the air, and hardly any ozone. Four massive ozone generators will beemployed on Mars. Since ozone molecules bond with each other, the generators could assist in the creation of an ozone

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layer to increase the planet’s insolation and reduce the levels of surface UV flux. Solettas will also be placed in orbitto enhance insolation, with multiple solettas at the poles and higher latitudes to melt permafrost. Enhancing Mars’atmosphere will cause permafrost to melt, giving it back some of its water and releasing volatiles from the regolith, asI said before. This will help thicken the atmosphere, making Mars easier to live on, and perhaps eventually will makeit so that you could walk on Mars without a pressure suit. You’d still need your handy pack of compressed oxygen,though, until total terraformation had been completed.

Mars’ mean global temperature is ~-56ºC. Equatorial temperatures range from 20ºC in the afternoon to -90ºC beforedawn. Winter poles are usually below -130ºC. With temperatures like this, water remains ice, and humans can’t survivein temperatures this cold. In a way, a solution to this could kill two birds with one stone. With Mars’ lack of atmosphereand a protective ozone layer, hardly any of the sun’s heat is trapped to warm the surface. To create a thicker atmosphere,we’d need to put more greenhouse gases in the atmosphere, which could be done by particle generators (placed at 45spots around the planet) and other planetary engineering techniques.

In Carl Sagan’s schematic for the runaway greenhouse scenario, various planetary engineering techniques are used towarm regions of Mars that are rich in volatiles. Carbon dioxide in the polar caps and the regolith then begins tovaporize. The thicker atmosphere warms the surface, causing a further release of gases. If positive feedback is strongenough, self-sustaining outgassing may occur as a result of a comparatively trivial forcing.

Properties of Ideal Greenhouse Gas Mixture1. Uniform gray absorption2. Strong radiative forcing at low concentrations3. Non-destructive to ozone4. Long atmospheric life5. Resistance to photodissociation6. Non-toxic in ppm concentrations7. Manufacturable from elements likely to be abundant on Mars

(Source: McKay, C.P., Toon, O.B. and Kasting, J.F. “Making Mars Habitable.” Nature, 352, 489-496, 1991.)

This graph shoes the thermal balance of a model 1-bar Earth-like atmosphere on Mars with 10 mbar of CO2 inequilibrium with water.5 The surface temperature is set at 15ºC and is shown as the upper dotted curve; the temperature

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at which the planet should radiate to space is shown as the lower dotted curve. The solid line is the actual infraredradiation emitted at the top of the atmosphere showing that the surface is far too hot to be in equilibrium.

Now, this shows the effect of adding trace gases to the Earth-like Martian atmosphere from the preceding graph (the solidlines) and to the present-day atmosphere of Mars (dotted lines). Carbon, hydrogen, and fluorine are added to theatmosphere, which are “super-greenhouse” gases. All of these are already on Mars. We would not generate chlorine orbromine, which are destructive to ozone, even though they are major greenhouse gases. Chlorine and bromine break theproperties of the ideal greenhouse gas mixture. Trace gas absorption in the window region gives maximum warmingincrements of ~40ºC and ~30ºC, respectively. A uniform gray absorber, however, a hypothetical mixture of gases that isactive over the entire infrared spectrum, can warm Mars above freezing. This would make living on Mars a bit morehospitable to humans, though temperatures would resemble those of Northern Canada, Alaska, and parts of Eastern Russia.

III. Water on MarsThere is water on Mars. Getting to it or finding it where it isn’t ice is the problem. The most popular idea to put wateron planets that lack it is to crash land comets on the surface. The problem is if you put humans on the planet before youintend to crash-land the comets, you’ll kill the humans when you decide to crash the comets. So, this option’s too late,but there’s another option involving comets. The ship could target a comet and then, using a monopolaric beam ofcoherent energy, shoot the comet so it breaks into pieces small enough to be placed in the cargo bay. These pieces couldbe melted on the surface. Granted, this doesn’t seem like a very easy idea; this would take a while. So, the solution togetting water on Mars is using what is already there. The majority of Mars’ water is frozen in the polar ice caps,permafrost in the higher latitudes, or possibly ground water.

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(Source: Fogg, Martyn J. Terraforming: Engineering Planetary Environments. 1995 by SAE International, Warrendale, PA.)

A simplified cross-section of the Martian regolith from the North Pole to the South Pole shows that ground ice is stablepoleward of about 40º latitude where the temperature is permanently below freezing. Ice at lower latitudes must beisolated from the atmosphere by a diffusive barrier.

IV. Stages of Martian Terraforming ProcessAll terraforming processes have their difficulties and faults, but combining various planetary engineering techniquescould help to compensate for all of the faults of the different techniques. Here are the stages for the terraforming ofMars:

STAGE ONE: ECOPOIESIS1. Increase insolation

a. use solettas to gain a 30% increase in the Martian solar constant (fs = 1.3)*reflect 6370 TW onto Mars

2. A Mini-Runaway Greenhouse (controlled)3. Devolatilization of Carbonates and Nitrates

a. nuclear miningb. focused sunlight from solettas

4. Add Artificial Greenhouse Gases to Atmosphere5. Establishment of a Dynamic Hydrological Cycle

a. hydrological cycle working outside of the enclosed habitat*precipitation, lakes, rivers

STAGE TWO: TOTAL TERRAFORMATION1. A Breathable Partial Pressure of Oxygen2. A Solution to the Problem of Mars’ Lack of Nitrogen

a. possibility of mining nitrogen from the thick atmosphere of Titan, which is mainly nitrogen and methane3. Topping Up and Maintaining Martian Seas

The Martian Clock and CalendarThe Martian day (referred to as a “sol”) is 24 hours and 37 minutes long. Instead of trying to work with that extra 37minutes, I decided to manipulate the length of an hour on Mars (when I did this, I had never heard about changing thelength of the second or minute...that just seems complicated, and this is fairly easy). If you round the 37 minutes up to40, you find that this gives you 1480 minutes/sol. That can give you 37 hours of 40 minutes per hour.

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How does this differ from the actual time? Refer to the chart below:

If you use the leap day system (subtracting one ‘leap day’ every 494 sols to make up for rounding up 3 minutes asol), the length of your year in reality is no longer 669 sols (though it is on the calendar, which I will explain next),but 668.6457489 sols, almost the same length as the actual Martian year (this is factoring in the .4 days you gain inrounding up to 669 from 668.6)!

Now, the calendar consists of 22 months with nine months having 31 days and 13 having 30, giving you 669 days in theyear. I managed to keep the seven-day week with 96 weeks in a year. The months are named after features on Mars(Tharsis, Elysium, Hesperia, etc.), but names of the days of the week were left the same, since I was more concernedwith numbers than names.

I also made a calendar for Mars after its shifting had been completed (if humans decided to go that route). The newcalendar has 615 sols in a year, adding a leap day every third year to even things out. This year has 20 months withfifteen months having 31 days and five having 30 days.

Where did those numbers come from? Put the number of days in an Earth year over its average distance from the sun,and a variable “x” over Mars’ new distance from the sun:

Take your cross products to find the value of x, which comes out to be ~631.334. This number is the number of Earthdays in the new Martian year. Then, place the current number of Earth days in a Martian year over the number of solsin the current Martian year, and the number 631.334 over x.

Using cross products again, you discover that x equals ~615.344. This is the number of sols in the new Martian year.Round this off to 615, and you end up the calendar I explained before. If you round off the decimal to .3, then everythird year, you’d add a leap day to make up for your rounding down. This is only off by approximately .1 days.

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365 / 149,000,000 = x / 211,054,200 (3.1)

687 / 668.6 = 631.334 / x (3.2)

Martian / Terran Physiological DifferencesMartian humans will be physically different than those on Earth, and these changes will not take thousands of years tobecome evident; you could see them with the first generation of children born on Mars. With the lower gravity, theskeleton and muscles will grow and develop differently. Bones would be elongated, resulting in a height approximatelytwo feet taller than the person would have been if they had been born and grew up on Earth. As an example, I am 4’10”tall here on Earth. If I had been born and grew up on Mars, I would be about 6’6”. This would be how tall the shortpeople would be on Mars (I’m about a half a foot shorter than the average female my age)! So, that would mean therange of height on Mars (for fully-grown Martians) would be from about 6’6” (people who are short, like me) to over9’ tall (basketball player height people)! Women on Mars would, on average, be taller than Earth men! Along withbeing taller, the fingers would be a bit longer, as well as the shape of the skull (but not by that much; it would probablybe noticeable, if you stuck a Terran and a Martian together, who was who. Now, those people living on the Moon andthe Belters are a whole other story . . .).

Conclusion“Humans will go forth to Mars. Of this we may be as sure as we can be of anything. In the next twoor three decades perhaps, in the twenty-first century almost certainly, astronauts will make the longjourney . . . For Mars is not only a destination; it can be the beginning of the irreversible expansion ofhumans into the cosmos.”

— John Noble Wilford, Mars Beckons

This is my view of humanity’s future on Mars. And we will go to Mars, because it is in the nature of humanity toexplore. We’ve gone to the limits on Earth, and must now turn our views to the skies. There is an endless amount ofexploration to do out there, and I plan to be one of the ones doing the exploring. It’s the future, and we’ve got to checkit out.

Emissary Mission, Mars 2030. We’re the Martians now.

AcknowledgmentsThe majority of this research was done at home using books obtained from the library, books I owned, and the Internet.I worked on the project alone, but I do owe a few acknowledgments to some very important people. All of the picturesand graphs in this paper were either hand drawn and scanned into the computer or drawn with Paint© by the authorunless otherwise noted. The figures in the terraforming section came from the noted sources, but were drawn on thecomputer by the author in Paint.©

First of all, I would like to thank all of the members of the Mars Society Puget Sound Chapter, who made me a member,helped me find places to present my project, and let me help out on raising public awareness of getting humans to Mars.The National Space Society is also owed a thank you, since they let me present for them at Boeing’s Museum of Flightin April of 2000, the first time I really presented my first research.

I also owe thanks to Don Scott (NASA educator), Michael Okuda (Star Trek technical director), Daniel Slosberg(college student and fellow Mars Society member), and Thomas Gangale (head of Martiana). The picture of theterraformed Mars on the cover page was done by Michael Carroll, so I also owe him thanks for doing this beautifulpainting.

A dedication goes out to the Great Bird of the Galaxy himself, Gene Roddenberry, for creating Star Trek, which sparkedmy interests in astronomy. He will always be remembered by his fans.

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Appendices

Appendix 1 – Physical Parameters of Mars

Appendix 2 – Bulk Parameters of Mars

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Appendix 3 – Martian Seasons

Appendix 4 – Martian Eras

• Early and Middle Noachian: More than 4.4 billion years ago. Primitive ancient crust; early bombardment.• Late Noachian: 4.4 billion to 3.8 billion years ago. Intercrater plains; lava flows; sinuous channels.• Hesperian: 3.8 to 3.6 billion years ago. Lava flows; complex ridged plains.• Early Amazonian: 3.6 to 2.3 billion years ago. Smooth plains such as Acidalia; extensive volcanism.• Middle Amazonian: 2.3 billion years ago to 700 million years ago. Continued volcanism.• Late Amazonian: 700 million years ago to the present time. Major volcanic activity in Tharsis and elsewhere dying

out at a relatively late stage; disappearance of surface water.

Appendix 5 – Comparison of Plant Nutrients in Soils on Earth and Mars

* Mg = 25 kilos/m3 of regolith

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Appendix 6 – Articles of the Terran System

Planetary CodeARTICLES OF THE TERRAN SYSTEM

ISSUE: 473250

Solar System HeadquartersTerran System Alliance

Tanya HarrisonPlanetary Relations Ambassador

Terran System Command

ARTICLES OF THE TERRAN SYSTEM

WE THE MEMBERS OF THE TERRAN SYSTEM ALLIANCE DETERMINED:

To reaffirm faith in the fundamental human rights. In the dignity and worth of the human person, to the equal rights ofmale and female and of planetary social systems large and small, and To establish conditions under which justice andmutual respect for the obligations arising from treaties and other sources of interplanetary law can be maintained, and

To promote social progress and better standards of life in larger freedom,

AND TO THESE ENDS

To practice benevolent tolerance and live together in peace with one another as good neighbors, and

To ensure by the acceptance of principles and the institution of methods that armed force shall not be used except in thecommon defense, and

To employ machinery for the promotion of the economic and social advancement of all humans,

HAVE RESOLVED TO COMBINE OUR EFFORTS TO ACCOMPLISH THESE AIMS.

Accordingly, the respective social systems, through representatives assembled on the planet Earth, who have exhibitedtheir full powers to be in good and due form, have agreed to these articles of the Terran System, and do hereby establishan interplanetary organization to be known as the Terran System Alliance.

Chapter IPurposes And Principles

ARTICLE 1The purposes of the Terran System Alliance are:

1. To maintain interplanetary peace and security within the Terran System, and to that end: to take effective collectivemeasures for the prevention of threats to the peace, the suppression of acts of aggression, and to bring about bypeaceful means, and employing the principles of justice and interplanetary law, adjustment or settlement ofinterplanetary disputes which might lead to a breach of the peace;

2. To develop friendly relations among planets based on respect for the principles of equal rights and self-determination of humans, and to other appropriate measures to strengthen universal peace;

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3. To achieve interplanetary cooperation in solving problems of economic, social, cultural, or humanitarian character;in promoting and encouraging respect for human rights; and for fundamental freedoms for all without distinctionbetween culture, sex, race, or religious belief.

ARTICLE 2The Alliance and its members, in pursuit of purposes stated, shall act in accordance with the following principles:

1. The Alliance is based on the sovereign equality of all its members;2. In order to ensure to all of them the rights and benefits resulting from membership, all members shall fulfill in good

faith the obligations assumed by them in accordance with these articles;3. All members shall settle their interplanetary disputes by peaceful means in such a manner that interplanetary peace,

security, and justice, are not endangered;4. In all interplanetary relations, all members shall refrain from the threat, or use, or force against the territorial

integrity or political independence of any planetary social system, or in any manner inconsistent with the purposesof the Terran Alliance.

Chapter IIMembership

ARTICLE 3All planetary bodies in this Solar System that are colonized are members of the Terran Alliance. If a body that waspreviously uninhabited is colonized, it becomes a member of the Terran Alliance. This includes all nine planets orbitingour Sun and all of their moons.

ARTICLE 4Any member of the Terran Alliance which has persistently violated the purposes and principles contained in thesearticles of the Terran Alliance may be expelled from the Alliance by the Supreme Assembly by recommendation of theAlliance Council. With this action, the member that has been expelled loses all rights and protection given to them bythe Alliance.

Chapter IIIAgencies

ARTICLE 51. There are established as the principal agencies of the Terran Alliance: a Supreme Assembly, an Alliance Council,

an economic and social council, an interplanetary supreme court of justice, a combined peace-keeping force, and asecretariat;

2. Such subsidiary agencies as may be deemed necessary from time to time may be established in accordance withthese Articles of the Terran Alliance.

ARTICLE 6The Terran Alliance shall place no restriction on the eligibility of male and female persons of any member planetarysocial system to participate in any capacity under conditions of equality in its principal and subsidiary agencies.

Chapter IVThe Supreme Assembly

ARTICLE 6CompositionThe Supreme Assembly shall consist of all the members of the Terran Alliance. Each member shall be entitled to havenot more than five (5) representatives in this body;

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FUNCTIONS AND POWERS

ARTICLE 7The Supreme Assembly may discuss any questions on any matters within the scope of these Articles of the TerranAlliance or relating to the powers and functions of any agencies provided for in these articles and, except as providedin Article 9, may make recommendations to the members and the council or both on any such questions or matters;

ARTICLE 81. The Supreme Assembly may discuss any questions relative to the maintenance of interplanetary peace and security

put to it by any member or the Alliance Council, and, except as provided in Article 9, may make recommendationswith regard to any such questions to the members, the Alliance Council, or the pleading planetary social system, orto all of them. Any such question on which action is necessary shall be referred to the Alliance Council by theSupreme Assembly either before or after discussion;

2. The Supreme Assembly may call situations which are likely to endanger the interplanetary peace and security to theattention of the Alliance Council;

3. The powers of the Supreme Assembly as set force in this article shall not limit the scope or Article 7;

ARTICLE 91. Where the Alliance Council is executing the functions assigned to it under these articles with respect to any dispute

or situation, the Supreme Assembly shall make no recommendation with regard to that dispute or situation unlessso requested by the Alliance Council;

2. The Supreme-Secretariat, with the consent of the Alliance Council, shall notify the Supreme Assembly at eachsession of any matters relating to the maintenance of interplanetary peace and security which are under discussionin the Alliance Council, and shall notify the Supreme Assembly, or the members if the Supreme Assembly is not insession, immediately when the Alliance Council completes its deliberations on any such matters;

ARTICLE 101. The Supreme Assembly shall initiate studies and make recommendations for the purpose of:

A) Promoting interplanetary cooperation in political fields and encouraging the progressive development ofinterplanetary law and its codification;

B) Promoting interplanetary cooperation in the economic, social, cultural, educational, and health fields, andassisting in the realization of human rights and fundamental freedoms for all without distinction as to culture,sex, language, or religion;

2. The further responsibilities, functions, and powers of the Supreme Assembly with respect to matters mentioned inParagraph 1(B) above are set forth in later Chapters.

ARTICLE 11Subject to the provisions of Article 10, the Supreme Assembly may recommend measures for the peaceful adjustmentof any situation, regardless of origin, which it deems likely to impair general welfare or friendly relations among planets,including situations resulting from violations of the provisions of these articles setting forth the purposes and principlesof the Terran Alliance;

ARTICLE 121. The Supreme Assembly shall receive and consider regular and special reports from the Alliance Council; which

reports shall include an account of the measures that the Alliance Council has decided upon or taken to maintaininterplanetary peace and security;

2. The Supreme Assembly shall receive and consider reports from the other agencies of the Terran Alliance on agreedupon regular periods or reporting;

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ARTICLE 131. The Supreme Assembly shall consider and approve the budget of the Terran Alliance;2. The expenses of the Terran Alliance shall be borne by the members as appointed by the Supreme Assembly;3. The Supreme Assembly shall consider and approve any financial and budgetary arrangements with specialized

agencies and shall examine the administrative budgets of such specialized agencies with a view to makingrecommendations to the agencies concerned;

4. All budgets of, and expenses of the Terran Alliance shall be made and paid in common interplanetary credit. Thecommon interplanetary credit shall be the official medium of exchange within the Terran Alliance.

VOTING

ARTICLE 141. Each member of the Supreme Assembly shall have one vote;2. Decisions of the Supreme Assembly on important questions shall be made on a two-thirds (2/3) majority vote of the

members present and voting. These questions shall include: recommendations with respect to the maintenance ofinterplanetary peace and security; the election of non-permanent members to the Alliance Council; the suspensionof the rights and privileges of membership; the expulsion of members; and budgetary questions;

3. Decisions on other questions, including the determination of additional categories of questions to be decided by atwo-thirds (2/3) majority, shall be made by a majority vote of the members present and voting;

ARTICLE 15A member of the Terran Alliance which is in arrears in the payment of its financial obligations to the Alliance shall haveno vote in the Supreme Assembly if the amount it is in arrears equals or exceeds the amount of the contributions duefrom it for the preceding two accounting periods. The Supreme Assembly may, nevertheless, permit such a member tovote if it is satisfied that the failure to pay is due to conditions beyond the control of the member.

PROCEDURE

ARTICLE 16The Supreme Assembly shall meet in regular periodic sessions and in such special sessions as occasion may require.Special sessions shall be convoked by the Supreme-Secretariat at the request of the Alliance Council or of a majority ofthe members of the Terran Alliance;

ARTICLE 17The Supreme Assembly shall adopt its own rules of procedure. It shall elect its president for each session;

ARTICLE 18The Supreme Assembly may establish such subsidiary agencies as it deems necessary for the performance of its functions.

Chapter VThe Alliance Council

ARTICLE 19Composition

1. The Alliance Council shall consist of eleven (11) members of the Terran Alliance. The Supreme Assembly shallelect six (6) other members of the Alliance to the non-permanent members of the Alliance Council, due regard beingespecially paid, in the first instance, to the contribution of the members of the Terran Alliance to maintenance ofinterplanetary peace and security and to the other purposes of the Alliance;

2. The non-permanent members of the Alliance Council shall be elected for a term of two (2) session periods. In thefirst election of non-permanent members, however, three (2) shall be elected for a term of one (1) session period.A retiring member shall not be eligible for immediate re-election;

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FUNCTIONS AND POWERS

ARTICLE 201. In order to assure prompt and effective action by the Terran Alliance, its members confer on the Alliance Council

primary responsibility for the maintenance of interplanetary peace and security, and agree that in carrying out itsduties under this responsibility the Alliance Council acts on their behalf;

2. In discharging these duties the Alliance Council shall act in accordance with the purposes and principles of theTerran Alliance. The specific powers granted to the Alliance Council for the discharge of these duties are laid downin Chapters VI, and VII;

3. The Alliance Council shall submit regular and, when necessary, special reports to the Supreme Assembly for itsconsideration;

ARTICLE 21The members of the Terran Alliance agree to accept and carry out the decisions of the Alliance Council in accordancewith these Articles of the Terran Alliance;

ARTICLE 22In order to promote the establishment and maintenance of interplanetary peace and security with the least diversion ofthe Alliance’s people, and economic resources for armaments, the Alliance Council shall be responsible for formulating,with the assistance of Terran System Headquarters staff referred to in Article 46, plans to be submitted to the membersof the Alliance for the establishment of a system for the regulation of armaments;

ARTICLE 23Voting

1. Each member of the Alliance Council shall have one vote;2. Decisions of the Alliance Council on procedural matters shall be made by an affirmative vote of seven (7) members;3. Decisions of the Alliance Council on all other matters shall be made on affirmative vote of (7) members including

the concurring votes of the permanent members, provided that, in decisions under Chapter VI, a party to the disputeshall refrain from voting;

PROCEDURE

ARTICLE 241. The Alliance Council shall be so organized as to be able to function continuously. Each member of the Alliance

Council shall, for this purpose, be represented at all times at the seat of the Alliance;2. The Alliance Council shall hold periodic meetings at which each of its members may, if it so desires, be represented

by a member of its government or by some other specially designed representative;3. The Alliance Council may hold meetings at such places other than the seat of the Alliance as in its judgment will

facilitate its work;

ARTICLE 25The Alliance Council may establish such subsidiary agencies as it deems necessary for the performance of its functions;

ARTICLE 26The Alliance Council shall adopt its own rules of procedure, including the method of selecting is governor;

ARTICLE 27Any member of the Terran Alliance which is not a member of the Alliance Council may participate, without vote, in thediscussion of any question brought before the Alliance Council whenever the latter considers that the interests of themember are specifically affected.

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Chapter VIPacific Settlement Of Disputes

ARTICLE 281. The parties to any dispute, the continuance of which is likely to endanger the maintenance of interplanetary peace

and security, shall, first of all, seek a solution by negotiation, inquiry, mediation, conciliation, arbitration, judicialsettlement, resort to regional agencies or arrangements, or other peaceful means of their own choice;

2. The Alliance Council shall, when it deems necessary, call upon the parties to settle their dispute by such means;

ARTICLE 29The Alliance Council may investigate any dispute, or any situation that might lead to interplanetary friction or give riseto a dispute, in order to determine whether the continuance of the dispute or situation is likely to endanger themaintenance of interplanetary peace and security;

ARTICLE 301. Any member of the Terran Alliance may bring any dispute, or any situation of the nature referring to in Article 29,

to the attention of the Alliance Council or the Supreme Assembly;2. The proceedings of the Supreme Assembly in respect to matters brought to its attention under this article will be

subject to the provisions of Articles 8 and 9;

ARTICLE 311. The Alliance Council may, at any stage of a dispute of the nature referred to in Article 28 or of a situation of like

nature, recommend procedures or appropriate methods of adjustment;2. The Alliance Council shall take into consideration any procedure for the settlement of the dispute that has already

been adopted by the parties;3. In making recommendations under the article of the Alliance Council should also take into consideration that legal

disputes should as a general rule be referred to the Interplanetary Supreme Court of Justice in accordance with theprovisions of the statute of the court;

ARTICLE 321. Should the parties to a dispute as referred to in Article 28 fail to settle it by means indicated in that article, they shall

refer it to the Alliance Council;2. If the Alliance Council deems that the continuance of the dispute is in fact likely to endanger the maintenance of

interplanetary peace and security, it shall decide whether to take action under Article 31 or to recommend such termsas it may consider appropriate;

ARTICLE 33Without prejudice to the provisions of Articles 28 through 32, the Alliance Council may, if all the parties to any disputeso request, make recommendations to the parties with a view to a pacific settlement of the dispute.

Chapter VIIAction With Respect To Threats To The Peace, Breaches Of The Peace, And Acts Of Aggression

ARTICLE 34The Alliance Council shall determine the existence of any threat to the peace, breach of the peace, or act of aggressionand shall make recommendations to maintain or restore interplanetary peace and security;

ARTICLE 35In order to prevent aggravation of the situation, the Alliance Council may call upon the parties concerned to complywith such provisional measures as it deems necessary or desirable. Such provisional measures shall be without prejudice

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to the rights, claims, or position of the parties concerned. The Alliance Council shall take into account any failure tocomply with such provisional measures;

ARTICLE 36All members of the Terran Alliance in obligation to the maintenance of interplanetary peace and security, agree to makeavailable to the Alliance, on call of the Alliance Council, armed forces, assistance, and facilities, including rights ofpassage, necessary for the maintenance of interplanetary peace and security;

ARTICLE 37When the Alliance Council has decided to use force it shall, before calling upon a member not represented on it toprovide armed forces in fulfillment of obligations assumed under Article 36, invite that member to participate in thedecisions of the Alliance Council relating to the employment of contingents of the member’s armed forces;

ARTICLE 38In order to enable the Terran Alliance to take urgent military measures, all members so capable, shall assign contingentsof their armed forces to the Alliance to be employed as a single peacekeeping force of the Terran Alliance. Allcontingents so assigned, and for the duration of their assignment, shall hold full faith and loyalty to the Terran Allianceand the protection of purposes and principles of these Articles;

ARTICLE 39Plans for the application of Alliance armed forces shall be made by the Alliance Council with the assistance of themilitary staff committee of Alliance Headquarters;

ARTICLE 40There shall be established within the Alliance a military staff committee to advise and assist the Alliance Council on allmatters relating to the Terran Alliance’s military requirements for maintaining interplanetary peace and security;

ARTICLE 41The action required to carry out decisions of the Alliance council for the maintenance of interplanetary peace andsecurity shall be taken by the Alliance, using such contingents as appropriate to the specific action.

Chapter VIIIThe Interplanetary Supreme Court Of Justice

ARTICLE 42The Interplanetary Supreme Court of Justice shall be the principle judicial instrument of the Terran Alliance. It shallfunction in accordance with the appended statute, and forms and integral part of these Articles;

ARTICLE 43All members of the Terran Alliance are “ipso facto” parties to the statute of the Interplanetary Supreme Court of Justice;

ARTICLE 441. Each member of the Terran Alliance undertakes to comply with the decision of the Interplanetary Supreme Court

of Justice in any case to which it is a party;2. If any party to a case fails to perform the obligations incumbent upon it under a judgment rendered by the court, the

other party may have recourse to the Alliance Council, which may, if it deems necessary, make recommendationsor decide upon measures to be taken to give effect to the judgment;

ARTICLE 451. The Supreme Assembly or the Alliance Council may request the Interplanetary Supreme Court of Justice to give an

advisory opinion on any legal question;

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2. Other bodies of the Terran Alliance and the specialized agencies, which may at any time be so authorized by theSupreme Assembly, may also request advisory opinions of the court on legal questions arising within the scope oftheir activities.

Chapter IXThe Supreme-Secretariat

ARTICLE 46The secretariat shall be comprised of a supreme-secretariat and such staff as the Alliance may require. The supreme-secretariat shall be appointed by the Supreme Assembly upon the recommendation of the Alliance Council, and shall bethe chief administrative officer of the Terran Alliance;

ARTICLE 47The supreme-secretariat shall act in that capacity in all meetings of the Supreme Assembly, of the Alliance Council, of theEconomic and Social Council, and shall perform such other functions as are entrusted to the secretariat by these bodies.The supreme-secretariat shall make a periodic report to the Supreme Assembly on the work of the Terran Alliance;

ARTICLE 48The supreme-secretariat may bring to the attention of the Alliance Council any matter which in his opinion may threatenthe maintenance of interplanetary peace and security;

ARTICLE 491. The staff shall be appointed by the supreme-secretariat under regulations established by the Supreme Assembly;2. Appropriate staffs shall be permanently assigned to the economic and social council, and, as required, to other

bodies of the Alliance. These staffs shall form a part of the secretariat;3. The paramount consideration in the employment of the staff and in the determination of the conditions of service

shall be the necessity of securing the highest standards of efficiency, competence, and integrity. Due regard shallbe paid to the importance of recruiting the staff on as wide a basis as possible.

Chapter XRatification And Signature

ARTICLE 501. These Articles of the Terran Alliance shall be ratified by the signatory governments in accordance with their

respective statutory processes;2. The ratifications shall be deposited with the government of the United Nations of Planet Earth, which shall notify

all of the signatory governments of each deposit as well as the supreme-secretariat of the organization when he/shehas been appointed;

3. These Articles of the Terran Alliance shall come into full force upon the deposit of the ratifications by the UnitedNations of Planet Earth, the Lunar Alliance, and the Federal Republic of Mars. A protocol of the ratificationsdeposited shall thereupon be drawn up by the government of the United Nations, which shall communicate copiesthereof to all of the signatory governments;

4. The governments signatory to these Articles of the Terran Alliance which ratify it after it has come into force willbecome original members of the Terran Alliance on the date of the deposit of their respective ratifications;

ARTICLE 51These Articles of the Terran Alliance, of which the various language texts are equally authentic, upon the coming intofull force of the Terran Alliance, shall be transferred by the United Nations to the organization for permanent deposit inits archives. Duly certified copies thereof shall be transmitted by the supreme-secretariat to the governments of all thesignatory social systems.

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In faith whereof the representatives of the Governments of the Terran Alliance have signed these Articles of the TerranAlliance.

Done at the Planet Earth, Earth date / Lunar date 25 April 2038Martian date 15 Hecates, M.Y. 8.

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• Dobbins, Thomas and Sheehan, William. “The Colours of Mars: Reality and Illusion.” “Sky and Telescope,” April 1999, pgs. 116-120.• Doody, David F. “Three Ships to Mars.” “Sky and Telescope,” Dec. 1999, pgs. 24-26.• Ferris, Timothy. “Personal: Human Seeks Alien.” “Popular Science,” Jan. 2000, pgs. 51-53.• Fogg, Martyn J. Terraforming: Engineering Planetary Environments. 1995 by SAE International, Warrendale.• Gresh, Lois and Weinberg, Robert. The Computers of Star Trek. 1999 by Basic Books, New York.• Groopman, Jerome. “Medicine on Mars.” “The New Yorker,” Jan. 2000, pgs. ?• Hammond, George and Kuck, Valerie J. Fullerenes: Synthesis, Properties, and Chemistry of Large Carbon Clusters. 1992 by the American

Chemical Society, Washington, D.C.• Hartmann, William K. “Invading Martian Territory.” “Astronomy,” April 1999, pgs. 46-51.• Hoagland, Richard C. The Monuments of Mars: A City on the Edge of Forever (4th ed.). 1987 by North Atlantic Books, Berkeley, California.• Joëls, K.M. The Mars One Crew Manual. 1985 by Ballentine Books, New York.• Kaplan, David. Environment of Mars. 1988 from the NASA Technical Memorandum 100470, Washington, D.C.• Kieffer, Hugh H. (with 114 collaborating authors) Mars. 1992 by University of Arizona Press, Tucson.• Krauss, Lawrence M. The Physics of Star Trek. 1995 by Harper-Perennial, New York.• Malin, Michael C. “Visions of Mars.” “Sky and Telescope,” April 1999, pgs. 42-49.• “Mars.” New Standard Encyclopaedia, Vol. 9, pgs. M-156-M-158.• McDaniel, Stanley V. The McDaniel Report. 1993 by North Atlantic Books, Berkeley, California.• McKay, C.P., Toon, O.B. and Kasting, J.F. “Making Mars Habitable.” Nature, 352, 489-496, 1991.• Michaun, C.M. Handbook of the Physical Properties of the Planet Mars. 1967 by the U.S. Government Printing Office, Washington, D.C.• Moore, Patrick. On Mars. 1998 by Cassell, London.• “More Martian Microbes?” “Sky and Telescope,” June 1999, pgs. 24-25.• NASA/CP – 2000-210429, Compiled by Smitherman, Jr., D.V. Space Elevators: An Advanced Earth-Space Infrastructure for the New

Millennium. 2000 by NASA Marshall Space Flight Centre, Huntsville, AL.• Newton, Sir Isaac. Mathematical Principles of Natural Philosophy (Principia Mathematica). 1934 (Written 1687) by Encyclopaedia

Britannica, Inc., London.• “Nuclear Physics.” New Standard Encyclopaedia, Vol. 10, pgs. N-427-N-428.• Odenwald, Sten. The Astronomy Café. 1998 by W.H. Freeman and Company, New York.• Okuda, Michael and Denise. The Star Trek Encyclopaedia: Updated and Expanded Edition. 1997 by Pocket Books, New York.• Okuda, Michael and Sternbach, Rick. Star Trek: The Next Generation Technical Manual. 1991 by Pocket Books, New York.• Pauls, Michael and Facaros, Dana. The Traveller’s Guide to Mars. 1997 by Cadogan Books, London.• Raeburn, Paul. “Manned Mission to Mars.” “Popular Science,” Feb. 1999, pgs. 40-47.• Raeburn, Paul. Uncovering the Secrets of the Red Planet. 1998 by National Geographic Society, Washington, D.C.• Robinson, Kim Stanley. Green Mars. 1994 by Bantam Spectra, New York.• Robinson, Kim Stanley. Red Mars. 1994 by Bantam Spectra, New York.• Schefter, Jim. “NASA’s Changing Fortunes.” “Popular Science,” April 2000, pgs. 56-61.• Shirley, Donna. Managing Martians. 1998 by Broadway Books, New York.• Slipher, E.C. A Photographic History of Mars: 1905-1961. 1962 by Northland Press, Flagstaff, Arizona.• Stewart, John. Moons of the Solar System: An Illustrated Encyclopaedia. 1991 by McFarland & Company, Inc. Publishers, Jefferson, North

Carolina.• Sweetman, Bill. “Runway to Space.” “Popular Science,” June 1999, pgs. 72-77.• “A Tale of Two Polar Caps.” “Sky and Telescope,” April 1999, pg. 17.• Utter, E.C. Parliamentary Law at a Glance. 1928 by The Reilly & Lee Company, Chicago.• Watters, Thomas R. “Planetary Face-Off.” “Astronomy,” Jan. 1999, pgs. 58-63.• Wells, R.A. Geophysics of Mars. 1979 by Elsevier Scientific Publishing Company, New York.• Wilford, John Noble. Mars Beckons. 1990 by Alfred A. Knopf, Inc, New York.• Zubrin, Robert. The Case for Mars: The Plan to Settle the Red Planet and Why We Must. 1996 by the Free Press, New York.• Zubrin, Robert. Entering Space. 1999 by Tarcher-Puntam, New York.

Documentaries• Life on Mars?• Destination: Mars• Mars: Pioneering a Planet• Inside the Space Shuttle• Into Orbit: The Astronaut• The Science of Star Trek• Understanding Space Travel• Rocket Ships• Solar Empire• Terraforming• Cosmic Safari: Are We Alone?• NASA Press Conference: Possible Evidence of Liquid Water On Mars

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• Space Colonies• Science of the Impossible: Can We Reach the Stars?

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Mobile Nuclear Fission Engines for Mars Surface Vehicles and/or Aircraft

Robert D. Woolley[2000]

AbstractWhen humans go to Mars, we should take vehicles such as “rovers” or helicopters, powered by compact nuclear fissionreactor engines. This would provide portable “muscle power” for digging, e.g., in search of life or water, and for longdistance mobility to explore the entire planet without having to periodically return to a fixed ground base. Reactorpower would provide continuous life support, including heat to survive the frigid Martian nights.

Technical requirements for Martian mobile nuclear fission engines are discussed. The fundamental design issue isminimizing mass while providing adequate power conversion and radiation shielding.

Nuclear fission reactors have supplied more than 20% of Earth’s electricity over the last several years, using eithernatural uranium fuel or uranium slightly enriched in the fissionable isotope U235. However, a compact reactor must useeither Highly Enriched Uranium (HEU) or plutonium fuel in its core. A plutonium core would be smaller than a HEUcore but would suffer from a smaller experience base. Compact nuclear fission reactors using HEU are ubiquitous inthe submarines of several nations’ navies and are even used to power some aircraft carriers.

Substantial design, development and testing activity was pursued in the 1950s and 1960s to adapt mobile nuclear fissionreactors to propel jet aircraft and nuclear thermal rockets. The aircraft and rocket applications are not in use today, inpart due to radiological safety issues that could accompany a crash in a populated region. In contrast to plutonium, HEUfuel carries the safety advantage of having negligible radioactivity before it is inserted into a reactor and fissioned. If aMars mission with an unoperated HEU-fueled reactor as cargo were to crash while being launched from Earth,radiological safety issues would be insignificant.

Open Brayton CycleThe “open Brayton cycle” provides an effective implementation of mobile engines. Atmospheric air is compressed toan elevated pressure, heat is added at constant pressure, and then the compressed, heated air expands through a turbine.Expander work exceeds compressor work and provides net power as shaft rotation torque.1,2

Fixed installations employing gas turbines for electric power production usually add various heat exchangers betweenairflows to more efficiently use heat, via regeneration and/or intercooling. But these additions are not appropriate for amobile engine since they add significant weight. For a mobile engine, the simple Figure 1 diagram is proper.

On Earth, the open Brayton cycle has been used in aircraft and surface vehicles. For such terrestrial implementations,the “heat addition” stage is usually implemented by burning hydrocarbon fuels in the compressed air. Although furthertemperature increase would in principle boost output power and efficiency, air temperatures are not usually designed toexceed about 1300°C due to turbine blade materials limits.

Why We Need Long Range Mobility and Nuclear EnginesNASA’s present stated cost of launching into low Earth orbit is $10,000 per pound mass, i.e., $22,000/kg. Using ahydrogen-oxygen rocket to inject into a transfer orbit to Mars and making a minor allowance for aerobraking equipmentand for deadweight such as fuel tanks and rocket engines, NASA’s present delivery cost to the surface of Mars mustexceed $71 million/metric ton. The delivery charge alone for a 53.8 metric ton Mars Direct mission would thus exceed$3.8 billion. Median US household income in 1995 is stated by the census bureau as being $34,076. Thus the deliverycharge for each Mars Direct mission is equivalent to the total annual income of over one hundred thousand medianincome US households.

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Robert D. Woolley; email: [email protected]

Figure 1. The Open Brayton Cycle

Because the crew of each manned mission to Mars will spend about 18 months on the planet Mars, and because eachmission will be expensive, it would be wasteful for manned exploration to be limited to the immediate vicinity of eachlanding site. For instance, if round-trip long distance mobility on Mars’ surface were severely limited to not exceed 100km (62 miles) from each landing site, then 4554 separate missions to Mars would be required to visit the entire planetonce. A happier situation would provide round-trip long distance mobility to explore any location within a 2,171 km(1,349 mile) distance from each landing site, allowing access to 10% of Mars’ surface area during each mission. Theultimate capability would support round-trip excursions up to 10600 km away, providing mobility to explore any placeon Mars from any single landing site.

The alternative to using a mobile nuclear reactor engine would be to use a conventional engine powered by the chemicalreaction of fuel and oxidizer, such as an internal combustion engine or a combustion gas turbine. But Mars has no globalnetwork of fuel and oxidizer manufacturing plants, so chemically fueled vehicle engines would need to be supported bythe infrastructure of a fuel and oxidizer manufacturing plant built at a ground base. The Martian air, composed primarilyof CO2 gas, cannot support combustion, so in addition to the weight of a fuel tank, a vehicle powered by internalcombustion would also need to carry another tank containing an even greater weight of oxidizer. Since Mars lackspetrochemicals to serve as chemical feedstocks, the fuel and oxidizer would need to be manufactured from availablesimple materials, e.g., CO2 and perhaps H2O. Engine-relevant quantities of oxygen practically require a cryogenic tankalong with the weight of its thermal insulation and the complexity of its boil-off pressure controls. Depending on whichfuel is manufactured, e.g., CH4 (methane), the fuel itself could also practically require cryogenic storage. Long durationtrips would then require further increasing the amount of cryogenic oxidizer and fuel in order to accommodate boil-offduring the excursion. The maximum weight of fuel plus oxidizer plus tanks, which the vehicle is capable of carrying,would determine the vehicle’s maximum range. The maximum distance of points accessible via round-trips from theground base would be about one third of that maximum range if a safety margin is included, but could never exceed onehalf of that maximum range.

On Earth, cars typically can carry enough gasoline to operate for up to 8 hours at cruising speed. The oxygen consumedhas a mass 3.5 times the mass of gasoline burned, so if it were necessary to also carry oxidizer within the same total weight,the time between refueling would need to be reduced below 2 hours and the driving range similarly reduced. Some longdistance passenger aircraft can cruise for up to 12 hours, but that duration would be reduced below 3 hours if it werenecessary to carry oxidizer along with the jet fuel. Martian excursion duration could be extended somewhat by increasingfuel and oxidizer tank sizes at the expense of payload, or by reducing engine size and power (thus jeopardizing the crew).

To safely venture even 100 km round-trip from the landing site, the one-way range should be 300 km. For a surfacevehicle with an average ground speed of 20 km/hr through rough terrain, the 300 km range would require driving for 15hours. This is probably near the limit achievable for engines powered by combustion of chemical fuel and oxidizer.Truly long-range mobility excursions lasting for days or weeks using chemical fuels and with an adequate power levelappear impossible. Nuclear power appears to be necessary for long range mobility on Mars.

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General RequirementsLong range mobility implies that between returns to the ground base, crewmembers may be inside a mobile vehicle forextended excursion periods of time, measured in days or even weeks. Such excursion durations are too long forcontinuously wearing space suits as done in rover excursions during Apollo missions to the Moon. Each Mars vehiclewith long distance mobility must therefore have a pressurized crew compartment with air rebreathing life supportequipment (i.e., CO2 removal and O2 addition), sanitary facilities, food preparation facilities, communicationsequipment, sleeping bunks, and storage for consumables such as oxygen, drinking water, and food. The vehicle’s enginemust be capable of powering the internal equipment in addition to providing mobility. Because the nighttime Martianair temperature dips to –100°C (-148°F) at the equator and gets even colder elsewhere, the vehicle’s engine must alsoprovide adjustable heat as needed to keep the crew compartment warm.

Nuclear fission has the advantage of very high energy content, about one million times the energy content of the samemass of chemical fuels. The fission of one gram of fuel provides an amount of heat equal to a 1 megawatt power levelfor a 24 hour day, so a one megawatt high temperature heat source can be provided continuously throughout an 18 monthMars surface stay duration by the nuclear fission of only 0.55 kg (1.2 pounds mass) of fuel. This level of fission “burn-up” can easily be provided by a single compact fission core without requiring any refueling operation during the 18month mission. If the one megawatt high temperature heat source operated a thermal conversion engine with 25%thermal conversion efficiency, the engine output would be 350 horsepower of mechanical work plus 750 kW of lowertemperature “warm” heat.

It is expected that mobile nuclear reactor engines for Mars will need to develop maximum mechanical power levelsranging from around one hundred horsepower for small surface vehicles to several thousand horsepower for aircraft.Fission (and thermal) power levels to consider for mobile nuclear fission engine design purposes therefore range fromabout 0.3 MW to 10 MW, depending on application details.

To save weight, the energy conversion system should be open cycle, using the Martian air as its working fluid tominimize the requirement for heavy heat exchangers. The open Brayton cycle appears to be the best choice of thermalconversion system. That, in turn, requires that the reactor have as high an operating temperature as feasible in order toavoid low conversion efficiency.

The Nuclear Heated Open Brayton CycleAlthough possible, it is not advisable to directly heat compressed Martian air by routing it through a nuclear reactor.That would require designing the nuclear reactor with sufficient internal cooling spaces to accommodate a gas, whichwould increase the reactor’s critical mass and prevent the reactor from having a compact size. Also, oxygen in the air’scarbon-dioxide, and free nitrogen, would become activated during passage through the reactor, becoming N16 , a gammaray emitter with a 7.1 second half-life, and N17 , a neutron emitter with a 4.2 second half-life.8 A continuously operatingengine routing air through its nuclear reactor would thus trail a short, radioactive exhaust plume.

A good fluid to use for transferring heat between a nuclear reactor and air is the liquid metal, lithium. It has not yetbeen used much as a heat transfer fluid in nuclear reactors because one of its two naturally occurring isotopes, Li6, (7.5% abundance) has a very high cross section for absorbing neutrons and tends to prevent the fission chain reaction fromproceeding. The other isotope, Li7, (92.5 % abundance) has a very low cross section for absorbing neutrons and iscompatible with use in a reactor. Happily, the two isotopes can be separated (at some cost) so that nearly pure Li7 canbe used.4,5

Sodium is another liquid metal that has been used extensively in certain nuclear reactor designs (i.e., breeders).However, sodium becomes temporarily radioactive within a reactor, so the heat exchanger between sodium and airwould need its own gamma ray shielding for personnel protection. In contrast, Li7 does not become activated in thereactor. The liquid / air heat exchanger for a design using Li7 would not require shielding.

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Figure 2. Nuclear Heated Open Brayton Cycle

Lithium, with an atomic number of three, is a member of the chemical group known as the alkali metals. It is used invarious conventional applications, e.g., in high energy density batteries, in fireworks where it imparts a crimson color,etc. Lithium melts at 180.54°C, so a lithium heat-transfer loop must be kept warmer to prevent freeze-up. As a liquid,lithium’s density is about one-half the density of water. Lithium’s specific heat is close to 1.0 cal/g-°C throughout itsliquid range, matching the specific heat of water. But whereas water can be kept liquid up to 300°C only by increasingits pressure above 90 atm, the boiling point temperature of lithium at one atm pressure (101.3 kPa) is 1,347°C.

Although lithium is corrosive to many materials, tests have shown lithium exhibits long-term chemical compatibilitywith tungsten (a.k.a. wolfram) at temperatures up to 1300°C. Tungsten has the highest melting point (3,422°C) of allknown materials. It also is a very hard metal. It is used commercially in light bulb filaments and welding equipment.With present technology, it will be difficult to manufacture complicated shapes out of tungsten.

However, there are benefits to using tungsten beyond the fact that it facilitates lithium’s use. Tungsten’s high density(19,300 kg/m3) makes it a more effective gamma ray shield material than lead (density 11,350 kg/m3), allowing reducedgamma ray shield mass. Tungsten is more effective than lead in scattering high-energy neutrons into lower energyranges where the moderator becomes effective. Tungsten’s thermal conductivity follows closely behind the thermalconductivities of gold, silver, copper, and aluminum, exceeding thermal conductivities of other metals. Tungsten’schemical inactivity renders it a good container for fission products. And its high melting temperature might help if areactor malfunction occurred.

Pattern Process for Nuclear Heated Open Brayton CycleThe NIST-JANAF thermochemical tables3 were used to prepare thermodynamic properties simulating Martian air,assuming its composition is 95.7% CO2, 2.7 % N2, and 1.6 % Ar, approximating Viking data. The calculated enthalpy(H), entropy (S), and specific heat (Cp) of this “MarsMix” versus Temperature are plotted in Figure 3. Note that aMarsMix mole’s mass is 43.5 grams.

Using thermodynamic relationships for perfect but nonideal gases, adiabatic process pressure changes are calculated viathe relation

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Figure 3. Calculated “MarsMix” Thermodynamic Data

Given the choice of lithium and tungsten materials, the high temperature compressed Martian air emerging from the heatexchanger is assumed to be 997°C (1,270°K = T4). (This choice may be conservatively low.) The planetary averageMartian air temperature is about -63°C (210°K = T0), and is chosen as inlet air temperature, although it could varybetween –133°C (140°K) and +7°C (280°K) depending on local conditions. For purposes of these calculations, theMartian air pressure was taken as 680 Pa, which is consistent with Viking data. It can be shown1 that the optimumchoice of open Brayton cycle compression ratio to maximize output power per unit airflow is the one which sets T2 =T5. Using this approach, the optimum compression ratio was determined numerically from the MarsMixthermodynamic data, and found to be 112; the corresponding value for T2 = T5 is 580°K (which also is well abovelithium’s freezing point temperature).

Heat flows and compressor / expander powers are obtained from the enthalpy plot. For the ideal case, heating from state2 to state 4 adds 35.5 kJ/mole, the net work output is 20.9 kJ/mole, and the thermal efficiency is (20.9/35.5) = 59 %.Actual components may have isentropic efficiencies near 80 %, as depicted by the dashed lines of Figures 5 and 6.Heating from 2’ to 4 adds 31.2 kJ/mole, but the net work is 28 -18.8 = 9.2 kJ/mole. Calculated thermal efficiency isthen a more realistic 29.5%.

Assuming 31.2 kJ/mole heat addition from dashed line state 2’ to 4, in Figures 5 and 6, a 1 MW nuclear reactor heatsource would correspond to a Martian airflow of

(106 Watts) / (31.2 kJ/mole) x (43.5 grams/mole) = (32 moles/second) x (43.5 grams/mole) = 1.4 kg/sec airflow

The output power produced by the engine at 29.5% efficiency would be 295 kW, equivalent to 395 horsepower.

Gas flows are not excessive inside the engines. But due to the low density of the Martian air, special designs may berequired for air intake ducts. If the Martian air density is 0.020 kg/m3, the volumetric flows at low pressure are easilycalculated.

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Table 1. Range of Mobile Engine Parameters

Figure 4

Figure 5

Liquid FlowSince this process pattern includes heating the compressed Martian air by 690°C from 580°K to 1270°K, it would bereasonable to also select a temperature difference of 690°C for the liquid lithium loop. Using lithium’s specific heat anddensity shows that a 690° temperature change corresponds to changing lithium’s heat content by 1.44 x 109J/m3.Required volumetric flow rates of the lithium are obtained by dividing the reactor power by this value, i.e.,

Volumetric Flowrate of Lithium = (Power in watts) / ( 1.44 x 109 J/m3)

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Figure 6

The total flow lithium duct cross section area is given by the volumetric flowrate divided by the designed flow velocity.A commonly used flow velocity for many cooling applications is 3 m/s. Arbitrarily selecting it, we have

Lithium Duct Area = Power / (1.44 x 109J/m3)/(3 m/s)

This leads to the following table:

Table 2. Lithium Flow Parameters

Assuming that lithium flow within the nuclear reactor occurs in ducts formed between parallel tungsten plates separated2 millimeters apart, the Reynolds number of the flow is 23077, which is well into the turbulent regime. Moody’sdiagram7 gives a friction factor of 0.025. This allows calculation of the lithium pressure drop needed to pump throughthe reactor. If the flow path length between the parallel plates is 10 cm, the reactor’s pressure drop will be only 1,406Pa = 0.2 psi.

The heat transfer coefficient within the reactor can be simply estimated by assuming slug-flow conditions in the liquidmetal, as suggested by Eckert and Drake.6 The resulting coefficient is h = 3x105 W/m2-°C. If we allow a 100°Cmaximum temperature drop between tungsten and lithium, then the heat flux must be limited to

Qmax = 3x107 W/m2 (= 3 kW/cm2)

For a 10 MW reactor, a rectangular core measuring about 10 cm X 10 cm X 10 cm would provide adequate heat transferarea to limit heat flux to this level. The core volume would be approximately25 % liquid lithium, 25% tungsten, and50% HEU uranium almost pure in the fissionable U235 isotope. Lower power reactors could be more compact.

Neutron Moderator and ReflectorThe Li7 isotope is light enough to serve as a neutron moderator, but its cross section for neutron scattering is so smallthat an excessive quantity of lithium would be needed. Beryllium has a much larger scattering cross section, and could

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be used as a moderator. But hydrogen is a far better neutron moderator than any other element, which is one reason thatwater is used so often in reactor designs. Natural lithium reacts with hydrogen to form the hydride, LiH, with averagemolecular weight 7.95, and which remains a solid with density 0.78 g/cm3 at temperatures below its 688°C (961°K)melting point temperature. Lithium hydride’s hydrogen density can be directly calculated from these values, and isfound to be 88% of the hydrogen density in water. So Li7H can be expected to be a superior moderator at the hightemperatures involved in mobile nuclear reactors for Mars.

To approximately estimate the effect of using lithium hydride, the “migration length” value listed for water in Glasstone& Sesonske’s Nuclear Engineering text8 was divided by 0.88 to yield M = 6.7 cm. The “rules of thumb” recommendedthere state that a neutron reflector with a thickness of two migration lengths behaves essentially the same as an infinitelythick reflector, and that the corresponding reflector savings in the radius of the critical mass is reduced by one migrationlength below a nonreflected reactor.

Neutron Absorbing ShieldThe boron-10 isotope has a very large cross-section for absorbing thermalized neutrons, 3,838 barns=3,838x10-28 m2.9Multiplying by its atomic number density N = 12.81x1028 atoms/m3 gives the macroscopic cross section of 49,164/m.So a 1 cm thick layer of B10 attenuates thermalized neutrons by a factor of exp(-491.64) = 3x10-214, which essentiallyeliminates them. Incompletely thermalized neutrons are attenuated less, but may be reduced dramatically by a 1-cmthick B10 shield.

Gamma Ray ShieldThus, the radius of the reactor and its surrounding moderator / reflector shell and a B10 neutron absorbing shell may beabout 20 cm for a 10 MW reactor, and slightly smaller for lower power level reactors.

For a 10 MW reactor with maximal equilibrium accumulated fission products, the calculated gamma ray dose withoutattenuation at a location 3 m from the reactor’s center would be 63000 Sv/hr. To be in conformance with USA rules forcontinuous occupation by radiation workers, the gamma rays in occupied regions should be attenuated to not exceed5.7x10-6 Sv/hr. If a location 3 m from the 10 MW reactor center is to be occupied, gamma rays should therefore beattenuated by a factor of 9x10-11.

A numerical analysis using the tabulated fission gamma energy spectrum, energy absorption coefficients for differentphoton energies in tungsten, and photon “buildup factors” to estimate the effects of collisions on photon energy,9 leadsto the conclusion that the attenuation factor of 9x10-11 adequate for a 10 MW reactor is provided by a 30 cm thick layerof tungsten. This ignores the substantial gamma ray shielding effects of tungsten and uranium in the reactor core wheremost of the gamma rays are produced. And it ignores the fact that mechanical engine component could be placedbetween the reactor and occupied areas, providing some of the shielding.

If all 3 m distant locations in all directions surrounding the reactor were to be continuously occupied, it would makesense to provide the gamma ray shield as a spherical shell from R=20 cm to R=50 cm. The volume of that shell is easilycalculated to be 0.49 m3, and with tungsten’s 19300 kg/m3 density the tungsten mass would be 9.46 metric tons.

If the tungsten shell thickness were reduced to 25 cm, its mass would be reduced to 6.7 tons. However, the radiationdose at 3 m from the reactor’s center would reach the allowable annual limit in 187 hours. If the tungsten shell thicknesswere reduced to 20 cm, the shell mass would be reduced to 4.5 tons. However, the radiation dose at 3 m from the reactorcenter would reach the allowable annual limit in 4 hours.

The design of long range mobile vehicles should place the nuclear reactor external to the crew compartment, preferablyin the rear. Then only one side of the reactor needs to be well shielded. Thinner reactor shielding could be used on thereactor sides where a suited astronaut might only rarely approach during an EVA, and on the top, where an astronautmight never venture at all.

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By using these strategies, it should be possible to keep the total mass of a mobile engine including a 10 MW nuclearreactor to about 5 tons. Lower power reactors could have a slightly reduced mass.

Aircraft Characteristics and RequirementsThe interest in using aircraft on Mars derives from the same advantages that motivate the use of aircraft on Earth.Aircraft can provide long distance mobility at high speed without excessive energy consumption. As stated in HighSpeed Wing Theory by R.T.Jones and D. Cohen.11

“The chief advantage of air transportation over other forms of travel is the great speed that can beachieved with a relatively moderate cost in terms of energy, or fuel expended per mile of flight. Anairplane of efficient aerodynamic form may have a drag less than one twentieth of its weight. Theenergy required in steady flight is therefore less than one twentieth of the weight times the distanceflown . . .

“It is easily seen that the efficiency of the airplane is the result of the favorable aerodynamic propertiesof the wing. If one contemplates travel by a rocket, which overcomes gravity and achieves its distancesolely by virtue of the kinetic energy imparted at the beginning of the motion, then it is found that theenergy requirement is much greater – of the order of the whole weight time the distance. The energyexpenditure of a wingless rocket or projectile thus corresponds to a lift-drag ratio of about one, a figurethat can be surpassed easily by almost any form of winged body . . .”

On Mars, the aircraft situation differs from the situation on Earth in several respects. The Martian atmosphere is far lessdense than Earth’s atmosphere, so therefore Martian aircraft designs must be different from Earth designs, requiringmuch larger areas of aerofoils or wings to develop similar aerodynamic forces. Martian aerofoils or wings can be muchthinner since their forces per unit area are considerably less, and they must be much thinner in order to avoid excessiveweight. However, their weight is also reduced by the lower Martian gravity.

Traditional airplane designs land or take off while running horizontally on sufficiently smooth surfaces. However, muchof Mars’ surface is rough and strewn with rocks or boulders, and Mars also has no bodies of liquid water with theircharacteristically smooth surfaces. With no existing global network of Martian airport runways, the number of locationssuitable for a traditional design of airplane to land horizontally would be extremely limited or even non-existent.

For application in Mars exploration missions, it would thus be desirable for aircraft designs to be capable of verticallandings and vertical takeoffs at most locations on Mars’ surface. Helicopters are the most well studied machines thatprovide a capability for vertical landing and takeoff on Earth. If helicopter designs were suitably modified to work onMars, they could provide the most useful capability for long-range mobility.

Dry Earth air at 0°C has a measured speed of sound of 331.5 m/s (742 mph) whereas carbon dioxide, the mainconstituent of Mars’ atmosphere, has a speed of sound at the same temperature of 259 m/s (579 mph). The speed ofsound in colder Mars air is slower, perhaps approaching 210 m/s (470 mph) during polar winter night.

To avoid the drag increase associated with breaking the sound barrier, the designed rotor tip speed could be limited to140 m/s (313 mph) while the maximum designed cruising speed limited to 70 m/s (157 mph). Then the maximum rotortip speed through still air would not exceed 70 m/s + 140 m/s = 210 m/s while cruising.

The lift force is given by the formula, L = CL 0.5 ρ V2S , where ρ represents the air density, V represents the airfoil speedwith respect to the air, S represents the planform area of the aerofoil, and CL represents the lift coefficient of that aerofoilgeometry. For most aerofoil shapes, CL varies in rough proportion to the angle of attack, reaching a maximum lift valuetypically near 1.0 at an angle several degrees from the zero lift orientation.10 Using ρ = 0.017 kg/m3 as a typical Martianair density, and using V=110 m/s as an average airspeed of the helicopter’s rotor, this gives LMAX = (200 Pa) S.

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Thus for instance, to hover above Mars’ surface with a mass of 10 metric tons, a helicopter would need a total aerofoilarea of at least S=(10000 kg)(3.7 m/sec2) / (200 Pa) = 185 m2. If two 10-meter radius rotor disks were used, slightlyless than 30% of the area of the two disks would need to be aerofoils. This is a far greater solid fraction than typical ofEarth’s helicopters, but it appears possible by using multiple blades per rotor. If the aerofoil Lift / Drag ratio were aslow as 10, the engine would need to supply (3700 N)(110 m/s)= 0.4 MW (536 hp) of mechanical power just to maintainhovering. However, the engine must also have additional power for takeoff, landing, cruising, and furthermore shouldhave reserve margins for emergency conditions such as high winds. Probably several thousand horsepower is moreappropriate, using a 10 MW reactor driving an open Brayton cycle.

At the proposed 70 m/s cruising speed of this example, it would require 21 hours flying time to travel from Mars’equator to either its North or South pole, or equivalently, 84 hours flying time to completely circumnavigate the entireglobe. Because of the high-speed travel, the extent of recycling in the vehicle’s life support system could probably bereduced to save weight. For long distance flights, the pressurized cabin space should accommodate at least 2crewmembers to alternate as pilots.

ConclusionsThe engineering design and development of nuclear powered engines and vehicles for long range mobility on Marsshould be pursued now. The developed engines, rovers, and aircraft will provide significant value on Mars at adevelopment and delivery cost which is a miniscule fraction of the entire Mars exploration program cost. Without suchequipment, long-range mobility for exploring Mars and for exploiting Martian resources would be extremely limited.The mobility limitations would either cause a large increase in Mars program costs, perhaps leading to early terminationof a Mars program, or could even inhibit the first manned missions from occurring. To not undertake design anddevelopment of nuclear powered vehicles for Mars application would thus be “penny-wise but pound-foolish.”

References1. R. Decher, Energy Conversion, Oxford University Press, 1994, ISBN-0-19-507959-02. W. Reynolds, Thermodynamics, McGraw-Hill, 19683. M. Chase, Ed., NIST-JANAF Thermochemical Tables, 4th ed, 19894. J. Ballif et al, TC-1000 Lithium Literature Review: Lithium’S Properties And Interactions, Hanford Engineering Development Laboratory,

19785. R. OHSE, ed., Handbook Of Thermodynamic And Transport Properties Of Alkali Metals, Blackwell Scientific Pubs, 1985, ISBN-0-632-01447-

46. E. Eckert and R. Drake, Heat and Mass Transfer, McGraw-Hill, 19597. R.Fox and A. MacDonald, Introduction to Fluid Mechanics, John Wiley and Sons, 19858. S. Glasstone and A. Sesonske, Nuclear Reactor Engineering Fourth Edition (Vols 1 and 2), Chapman and Hall, 19949. J. Shultis and R. Faw, Radiation Shielding, Prentice Hall, 199610. L. Milne-Thomson, Theoretical Aerodynamics, Dover, 197311. R. Jones and D. Cohen, High Speed Wing Theory, Princeton University Press,1960

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A Model for Minimizing the Risk of Sudden DeathDuring Exercise in Long-Term Space Flight

George D. Swanson Ph.D[2001]

AbstractCrew members in long-term space flight may be relatively sedentary for months. An exercise program will be necessaryto maintain health and physical capacity. However, vigorous exercise stresses the heart and can trigger sudden death.This risk increases for those relatively sedentary individuals who occasionally attempt vigorous exercise. Alternatively,the risk is also high for those who exercise regularly for longer periods of time. The purpose of this paper is to introducea model that characterizes the risk of sudden death during exercise and to use that model to determine an optimal amountof regular vigorous exercise that minimizes that risk. The Physicians Health Study (Albert et al., N Engl J Med 343:1353-1361, 2000) followed 21,481 physicians for 12 years during which 122 died of sudden death. These physicianswere free of cardiovascular disease at the beginning of the study and will form the reference group for our analysis(Swanson, N Engl J Med 344: 854-855, 2001). Nested case-crossover methods were used to lay out contingency tablesin terms of sudden death and person hours of exercise or sedentary time. Combining these tables yields an odds ratiofor sudden death during exercise compared to rest: OR = [Fx / (1-Fx)] [(1-Px) / Px] where Fx is the fraction of suddendeaths that occur during exercise and is Px is the proportion of time that exercise could trigger sudden death. As in thePhysicians Health Study, Fx is linearly related to Px: Fx = α + β Px. Combining these two equations yields a quadraticequation for the minimum risk: β(β-1) Px2 + 2αβ Px + α(α-1) = 0. The solution characterizes the optimal amount ofexercise. For the Physicians Health Study, this was about 30 min per day six days a week. The implications of thismodel will be explored for a Mars mission.

IntroductionCrew members in long-term space flight with out artificial gravity will experience an environment of micro gravity andrelative inactivity for months. Consequently, crew members will experience cardiovascular de-conditioning, loss ofmuscle strength, bone deterioration and a depressed immune response.2 These effects would be detrimental for crewmembers expected to be fully functional upon arrival at Mars.

Exercise has been proposed as a countermeasure. As shown in Figures 1 and 2, a variety of resistance and aerobicexercises have been explored to counteract these effects.

However, excessive exercise increases the risk of overuse injuries,3 inflammation,4 upper respiratory track infection,10

stroke7 and vigorous exercise stresses the heart and can trigger sudden death.13 This risk increases for relativelysedentary men and women who occasionally attempt vigorous exercise.1 Alternatively, the risk is also high for men andwomen who exercise regularly but for longer periods of time.13 These risks may become important considerations fora mission to Mars.

Therefore, an exercise prescription has as a goal to balance maintenance of health and functional capacity of crewmembers with the risk of exercise. An overall model of risk / benefit ratio would be helpful here. As a first step towardthe development of such a model, we shall introduce an epidemiological model of sudden death. This model predictsan optimal exercise exists that minimizes risk of sudden death during exercise as compared to rest.

The ModelLack of physical activity and or lack of physical fitness are considered risk factors for heart disease.1 Therefore, life-time physical activity has been considered protective for heart disease. However, physical activity also stresses the heartand consequently can trigger a sudden cardiac death.13 From this point of view, habitual physical inactivity carries ahigh risk of heart disease but a minimal exposure to an exercise trigger. In contrast, high habitual physical activity

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George D. Swanson Ph.D; Department of Physical Education and Exercise Science; California State University; Chico, CA 95929-0330;[email protected]

carries a lower risk of heart disease but a more extensive exposure to an exercise trigger. These concepts suggest thatthere may be an optimal level of habitual physical activity – one that balances cardiac risk and exercise trigger exposureso as to minimize the risk of sudden cardiac death.

Figure 1. Resistance exercises for crew members on long-term space flights.Figure 2. Aerobic exercise for crew members on long-term space flights.

Our approach extends the analysis of recently published Physicians Health Study data.1 Case-crossover methods8 wereused to formulate the model and optimization problem.

The case-crossover approach lays out a contingency tables in terms of person hours for each subject who dies duringexercise or rest and yields an odds ratio for sudden death during exercise versus rest. Let Fx, be the fraction of suddendeaths that occur during exercise and Px the proportion of time (say in a week) in an exercise trigger.Then the odds ratio8 is given by,

The data from Table 2 in the Albert, et al.1 paper can be used to estimate a relationship between the relative frequencyof habitual vigorous exercise activity Px and Fx, the fraction of deaths occurring in exercise. Following the approach ofKessler,6 a linear relationship is derived using regression methods. Therefore, let:

Combining equations 1 and 2 and taking the partial derivative with respect to Px and setting the result equal zero yieldsthe optimal value for Px. This results in quadratic equation,

Model ImplicationsEquation 1 and 2 yields an odds ratio as a function of Px, while equation 3 yields the optimal Px, which minimizes therelative risk (odds ratio) See Figure 3.

Figure 3 Odds ratio for sudden death in exercise versus in rest. Note: Px includes exercise time (in hours) plus 0.5 hourpost exercise trigger (per exercise session) as a fraction of the 168 hours in a week . OR =[Fx / (1-Fx)] [(1-Px) / Px] with

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A Model for Minimizing the Risk of Sudden Death During Exercise in Long-Term Space Flight

OR = [ Fx / (1-Fx)] [(I-Px) / Px] (1)

Fx = α + β Px (2)

β (β-1) Px2 + 2αβ Px + α(α-1) = 0 (3)

Fx = α + β. For the Albert et al. data, Fx = .0876 + 6.31 Px, r = 0.99.6 The optimal Px is the solution of the quadraticequation β(β-1) Px2 + 2αβ Px + α(α-1) = 0, with the optimal Px = 0.035 for the Physicians Health Study data.1

Figure 3 plots the odds ratio (relative risk) as a function of exercise amount.Note the minimum occurs for Px = 0.035, which translates into a habitual vigorous

exercise activity (> 6 METS ) of about 30 minutes a day, six days a week.The minimum is reasonably broad so that a range from 4 to 7 days yields a similar risk (OR ~12).

Alternatively, one day a week for a 30 minutes doubles the risk as does 2 hours per day 7 days a week.

DiscussionThe male subjects in the Physicians Health Study were aged 40-84 years. Therefore, the present model may not applydirectly to a Mars crew. As pointed out by Thompson,15 the relative risk increases as age decreases because the absoluterisk at rest decreases. However, for younger men, the present model concepts should generally apply.

The incidence of sudden cardiac death during exercise is in the range of 1 per 15,000 (6.7/100,000) joggers per year.9Analysis of sparse data from the Seattle study13,15 indicates the absolute risk of sudden death as 1 death per 17,000(5.9/100,000) in men who exercise 1 to 19 min/week in vigorous activity; 1 death in 23,000 (4.3/100,000) for 20 to 139min/week; and 1 death per 13,000 (7.7/100,000) for more than 140 min/week. These data suggest that the absolute riskof sudden death during exercise is relatively low but follows a “J’ shape curve so that higher exercise levels may leadto much higher absolute risk.

Alternatively, our modeling allows us to interpolate among the sparse data of the Physicians Health Study yielding a “J”shape curve. The model facilitates a prediction of an optimal exercise level for minimizing the relative risk of suddendeath during exercise.

Although the absolute risk is low, the loss of a crew member to sudden death would be catastrophic. Therefore,minimizing the risk of sudden death would seem to be an important goal. However, other considerations may dictatelonger exercise periods.

In attempt to maintain health and function, exercise periods on Soviet Salyut-6 space station and Mir space station wereas high as 2-4 hours per day.2 This amount of vigorous exercise not only increases the risk of exercise but may also becostly to an operational work day and lead to an excessive caloric expenditure. An average daily cost of 1,450 kcal for

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2.5 hours of exercise would increase the relative risk of death during exercise to 50 (see Figure 3, Px = 0.125) and addan extra 527,800 kcal in one year. This would mean an extra yearly consumption of 605 kg of food, which would haveto be either supplied or grown.

Alternatively, perhaps the crew will remain relatively sedentary during the trip to Mars and “ramp” up their exerciseprogram prior to arrival. Can an exercise prescription be designed to minimize risk while restoring function?2

Additional modeling may help answer this question. However, even the best exercise strategy may not be enough tomaintain or restore health and function. The bottom line – a partial artificial gravity may be necessary during transport.

References1. Albert, C. M., M. A. Mittleman, C. U. Chae, I. M. Lee, C. H. Hennekens, and J. E. Manson. Triggering of sudden death from cardiac causes

by vigorous exertion. N Engl J Med 343: 1355-61., 2000.2. Convertino, V. A. Physiological adaptations to weightlessness: Effects on exerice and work performance. In: Exercise and Sport Reviews, K.

B. Pandolf and J. O. Holloszy (Eds). Baltimore: Williams & Wilkins, P 119-166, 1990..3. Fahey, T. D. Biological markers of overtraining. Biol. Sport 14: 1-19, 1997.4. Geffken, D., M. Cushman, G. Burke, J. Polak, P. Sakkinen, and R. Tracy. Association between physical activity and markers of inflammation

in a healthy elderly population. Am J Epidemiol 153: 242-50., 2001.5. Huikuri, H. V., A. Castellanos and R. J. Myerurg. Sudden death ddue to cardiac arrhythmias. N. Engl. J. Med. 345: 1473-1482, 2001.6. Kessler, K. M. Triggering of sudden death from cardiac causes by vigorous exertion (Correspondence). N. Engl. J. Med. 344: 854-855, 2001.7. Lee, I., and R. Ralph S. Paffenbarger. Physical Activity and Stroke Incidence: The Harvard Alumni Health Study. Stroke 29: 2049-2054, 1998.8. Maclure, M. The case-crossover design: a method for studying transient effects on the risk of acute events. Am. J. Epidemiol. 133: 144-153,

1991.9. Maron, B., C. Araujo, P. Thompson, G. Fletcher, A. Luna, J. Fleg, A. Pelliccia, G. Balady, F. Furlanello, S. VanCamp, R. Elosua, B. Chaitman,

and T. Bazzarre. Recommendations for preparticipation screening and the assessment of cardiovascular disease in masters athletes. Circulation103: 327-343, 2001.

10. Nieman, D. Exercise, upper respiratory tract infecytion, and the immune system. Med. Sci. Sports Exerc. 26: 128-139, 1994.11. Shah, P. K. Plaque disruption and thrombosis: potential role of inflammation and infection. Cardiol Rev 8: 31-9., 2000.12. Siegel, A. J., J. J. Stec, I. Lipinska, E. M. Van Cott, K. B. Lewandrowski, P. M. Ridker, and G. H. Tofler. Effect of marathon running on

inflammatory and hemostatic markers. Am J Cardiol 88: 918-20, A9., 2001.13. Siscovick, D. S., N. S. Weiss, R. H. Fletcher, and T. Lasky. The incidence of primary cardiac arrest during vigorous exercise. N Engl J Med

311: 874-7., 1984.14. Swanson, G. D. Triggering of sudden death from cardiac causes by vigorous exertion (Correspondence). N. Engl. J. Med. 344: 854-855, 2001.15. Thompson, P. Cardiovascular risks of exercise. Phys. Sportsmed. 29: 33-47, 2001.

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2001mars society convention part 3

Documents

human physiology

university research

meters altitude

physiological research

mars related research

blood gases

life support systems

altitude capability