Ares Explore is a trade study presenting two concepts for a near term, low cost manned mission to Mars.  Both mission scenarios include In Situ Resource Utilization (ISRU) to extract water, oxygen, and carbon monoxide from the Martian atmosphere for life support during the surface stay, and as propellants for the return trip to Earth.  With ISRU, the total masses launched from Earth in both scenarios are sufficiently low that only three launches of Russian Energia heavy-lift vehicles are required.  Scenario I is an all-ISRU mission, using in situ propellant for both Mars ascent and trans-Earth injection (TEI), while Scenario II is partially ISRU, utilizing in situ propellant for Mars ascent only, with TEI accomplished by means of terrestrial propellants.  Scenario II utilizes smaller ascent and descent vehicles to transfer the crew from a transfer habitat in low Mars orbit to a Martian surface habitat.  Scenario I, in contrast, requires only one habitat for the entire duration, which descends to and ascends from the Martian surface.  Both scenarios utilize solar arrays to furnish power during transit to and from Mars.  On the surface of Mars, power is supplied by a nuclear reactor in Scenario I and solar arrays in Scenario II.  An Environment Control and Life Support System (ECLSS) provides atmosphere revitalization, water recycling, and waste management in the habitats of both scenarios.  A pressurized rover in Scenario I allows the crew to travel to a variety of distant locations to conduct surface and atmospheric studies.  In Scenario II, a shorter range unpressurized rover is used, similar in design to the rovers of the Apollo Lunar Program.  While Scenario I is capable of delivering a larger scientific payload to Mars, Scenario II entails lower risks.  Either mission scenario can provide the initial infrastructure necessary for the start of long-term human exploration of Mars.
 
 

INTRODUCTION

The exploration of Mars has received increased attention in recent years.  Over the past 32 years, missions such as the Mariner and Viking series have returned much useful data.  Currently, NASA’s Mars Exploration Program plans to send a joint orbiter and lander to Mars every two years over the next 10 years.  The first two, Mars Global Surveyor and Mars Pathfinder, are scheduled for launch in late 1996.  All of these robotic missions are important precursors to sending humans to Mars.

Ares Explore is an initial study of two possible scenarios for carrying out a manned Mars mission.  Building upon previous design studies that have shown that the concept of a manned mission is feasible with modest improvements to existing technology [1-8], Ares Explore utilizes only existing or foreseeable technologies and evokes the concept of In Situ Resource Utilization (ISRU).

In Situ Resource Utilization, a concept in which mission resources are obtained at Mars, can produce methane, carbon monoxide, oxygen, and hydrogen from the atmosphere of Mars, all of which can be used as propellants to return a crew to Earth.  This greatly reduces the amount of mass a Mars mission must carry from Earth, hence making a manned mission economically feasible.  The concept of ISRU first appeared in science fiction over fifty years ago, but Ash et al. at Jet Propulsion Laboratory (JPL) in the late 1970’s were the first to do serious research on the subject [4].  ISRU missions and processes were further studied by Frisbee at JPL in the 1980’s [5].  Research has also been carried out by Ramohalli et al. at the University of Arizona and Linne at NASA Lewis Research Center [6,7].  Furthermore, missions based on ISRU have been described by Zubrin [1] and by design teams from the University of Washington under the direction of A.P. Bruckner [2,8].

The two scenarios presented in Ares Explore provide a comparison of ISRU alternatives.  One scenario produces all return propellant by ISRU, while the other utilizes some terrestrial propellant for return.  This also provides a comparison of the option of landing and launching a massive manned transfer vehicle at Mars with the option of providing smaller, less massive descent and ascent vehicles to transfer the crew between low Mars orbit (LMO) and Mars’ surface.  Each of the scenarios is presented in greater detail in the remainder of this paper.  The work represented here is not intended to be definitive, rather it is intended to illustrate two of several possible ISRU-enabled manned Mars mission scenarios and to stimulate further work in this area.
 

MISSION SCENARIOS

Many different mission scenarios were initially considered for Ares Explore.  Down-selection was based on minimizing the number of launches from Earth and decreasing the complexity of the mission.  A comparison of some of the main features of the two scenarios is shown in Figs. 1 and 2.
 

Scenario I

The first scenario, shown in Fig. 1, uses 100% ISRU propellants for the return journey.  In this manner, the only terrestrial propellants required for the return trip from Mars are RCS propellants.  Scenario I has a total of three launches from Earth using the Russian Energia heavy lift vehicle.  Each launch places a Trans-Mars Injection (TMI) stage attached to a specified payload into LEO. Each TMI stage then provides a burn that propels its payload to Mars.  The first launch, scheduled for December of 2013, sends the In Situ Propellant Production (ISPP) Transfer Vehicle (ITV) to Mars.  The ITV transports an ISPP plant with storage tanks, a nuclear reactor power source, three rovers, science equipment, and supplies.  At Mars, the ITV does a series of aerobraking maneuvers, then aerocaptures and uses parachutes and retro-rockets to place it on the Martian surface.  The ITV lands in a horizontal attitude, providing easier access to the rovers, nuclear reactor, equipment, and supplies.  Once on the surface, the main cargo bay door opens and a large rover transports the nuclear reactor to a site away from the ship and mounds regolith around it for radiation containment.

The second launch, the Booster Transfer Vehicle (BTV), also launches in 2013.  At launch, the BTV consists of two empty stages in series.  The two stages are the Trans-Earth Injection (TEI) stage and the Mars Orbit Injection (MOI) booster.  Upon arrival at Mars, the BTV aerocaptures into orbit and descends to the surface using parachutes and retro-rockets to slow its descent, and then lands near the ISPP plant.  The MOI booster is used to launch the TEI stage into Martian orbit at the end of the astronauts’ surface stay.  The initially empty MOI booster and TEI stage of the BTV are fueled with the propellants produced by the ISPP plant prior to the manned vehicle launch from Earth.

The third and final launch, on Feb. 2016, is the launch of the Manned Transfer Vehicle (MTV) which also contains the Earth Descent Vehicle (EDV).  The MTV provides the required life support system, water, and food supplies for a crew of four for the entire mission.  In transit to Mars, a tether deploys and the TMI stage and the MTV are rotated about their center of mass to provide 0.4g of artificial gravity for the crew aboard the MTV.  Upon arrival at Mars, the tether is released and the MTV aerocaptures into LMO where a communications satellite is deployed.  After a status check, the MTV descends to the Martian surface using retro-rockets and parachutes for a landing near the ISPP plant.  The propellant produced by the ISPP plant is then transferred to the tanks of the MTV for the return trip to Earth after a 1.5 years surface stay.

At the end of the surface stay, the fueled BTV launches.  Once in LMO, the MOI booster of the BTV is released and the fully-fueled TEI stage of the BTV waits to dock with the MTV.  The fueled MTV then launches to LMO where its main tanks and engine are jettisoned.  The MTV docks with the TEI stage, which provides the delta V to boost the MTV out of LMO on an Earth-bound trajectory.  Upon arrival at Earth, the manned EDV separates from the MTV and performs an Apollo-type entry and recovery.
 

Scenario II

Scenario II, as shown in Fig. 2, is a mission design that combines scaled-up sample return in situ technology [8] and solar power to maximize the probability of success for a first manned Mars mission.  By designing around scaled-up sample return technology, Scenario II provides a low mass Mars ascent that places a minimum demand on autonomous in situ propellant production.  This minimizes the risks involved in the operation of automated hardware and provides smaller propellant production requirements.

Like Scenario I, Scenario II consists of three launches from Earth using the Russian Energia heavy lift vehicle.  The payload and TMI stage are delivered to LEO, where the TMI stage provides the necessary delta V to place the payload on a transfer orbit to Mars.  The first launch, with the Surface Habitat Vehicle (SHV) as payload, occurs in 2013.  The SHV carries a small Mars Ascent Vehicle (MAV), the ISPP plant, a Water Vapor Adsorption Reactor (WAVAR) [3], solar arrays for power on Mars’ surface, a rover, a habitat with an inflatable addition, science equipment, and food supplies for a 1.5 year surface stay.  Once the SHV arrives at Mars, it is aerocaptured into LMO and descends through the atmosphere, using parachutes and retro-rockets, for a horizontal landing.  Solar arrays are immediately deployed to supply electrical power to the ISPP plant, which begins filling the propellant tanks of the MAV.

Using the same 2013 launch window, the second launch sends a TEI stage to LMO, where it remains until it is used for the return trip to Earth.  The TEI stage carries terrestrial propellants to provide the delta V required from LMO to Earth.

The third and final launch sends the MTV with its crew to Mars.  This launch occurs in the year 2016, only after successful production of MAV propellant has been verified and the TEI stage is safely in LMO.  The MTV provides the required life support system and food supplies for the transit to and from Mars for a crew of four.  It also contains the Mars Descent Vehicle (MDV), a vehicle for transferring the crew from LMO to the surface of Mars.

As in Scenario I, artificial gravity is provided for the crew by tethering the MTV and TEI stage.  The tether is released upon arrival at Mars, allowing the MTV to aerocapture into LMO where it docks with the TEI stage.  After a check of current mission integrity is performed, a go/no-go decision is made for landing.  If the decision is made to land, the MDV separates from the MTV and enters the Martian atmosphere to descend to the surface, landing near the SHV.  If a no-go decision is made, the crew can remain in LMO until an abort window is identified and an adequate Earth-return trajectory can be carried out by means of the TEI stage.  A successful Mars landing places the crew on the surface of Mars for 1.5 Earth years to conduct scientific experiments and carry out planetary exploration.

At the completion of the 1.5 year surface stay, the ISRU-fueled MAV is launched from a platform in a storage bay in the rear of the SHV. It carries the crew to LMO where the MAV docks with the MTV.  The TEI stage then boosts the MTV into an Earth-return transfer orbit.  Nearing Earth in 2018, the crew transfers to the MAV and separates from the MTV.  The MAV makes an Apollo-type atmospheric entry and landing, and the crew and Mars samples are recovered.
 

IN SITU PROPELLANT PRODUCTION

A very important element of Ares Explore is the concept of In Situ Resource Utilization (ISRU), which uses resources available at a destination rather than transporting them from Earth.  For this mission, ISRU reduces the mass required to launch from Earth’s surface.  ISRU is utilized in Ares Explore to produce oxygen and water for the crew, and propellants for all or part of the return journey to Earth, depending on the scenario.  Two years before the manned vehicle leaves Earth, an unmanned spacecraft is sent to Mars carrying the equipment to produce the required propellants for the return trip to Earth in Scenario I, or for ascent to LMO in Scenario II.  In both mission scenarios the propellants are produced over a period of 580 days. After production of the propellant has been confirmed, the astronauts leave Earth for the trip to Mars.

Several different ISRU propellants were considered for this design study. Methane and carbon monoxide, both of which utilize oxygen as an oxidizer, were the two most promising choices for fuel, due to their level of prior development and the simplicity of their production [7,9].  While methane/oxygen burning engines have a higher Isp (360 sec), the difficulty of long-term cryogenic storage of liquid hydrogen, required to produce methane, raises a question of methane’s viability. Although carbon monoxide and oxygen have shortcomings (very high propellant plant temperatures and a relatively low Isp of 300 sec), they were deemed the most viable propellants for this study.
 

Carbon Monoxide Production

The production of CO and O2 from the Martian atmosphere is a simple process.  The two most important components of the process are the zirconia electrolyzer and a Ni/Cu adsorption/desorption bed [9].  A block diagram of the production plant, designed for this mission, is shown in Fig. 3.

The Martian atmosphere consists of 95 % CO2 [10].  An adsorption/desorption pump with a membrane filter at its inlet feeds the Martian atmosphere into the propellant plant.  The inlet filter cleans the atmosphere of any dust particles while the adsorption/desorption pump separates the CO2 from other atmospheric gases, while increasing the pressure of CO2 from 0.007 atm to 1 atm.  This increase in pressure raises the temperature of the CO2 to 650 K.  The CO2 then travels through two regenerative heat exchangers and into the zirconia solid oxide cell electrolyzers [11].

Each zirconia electrolyzer unit consists of three concentric tubes surrounded by insulation [11].  Martian CO2 enters the innermost tube and immediately begins acquiring heat.  The flow moves into the middle layer, which is surrounded by the zirconia cylindrical membrane and resistance heater coils.  The electrolyzer raises the temperature of the flow to 1300 K, causing the oxygen to dissociate from the CO2.  The O2 is then electrostatically separated from the flow.  Approximately 50% of the CO2 is converted to CO and O2 in a single pass, as follows:

Exiting the zirconia electrolyzer unit are two streams, one of oxygen and another of CO and CO2 (see Fig. 3).  Each flow passes through one of the regenerative heat exchangers.  The O2 then passes through a radiator and an expander to allow cryogenic storage at 50 K and 1.5 Bar.  The CO and CO2 pass through an adsorption/desorption bed composed of copper and nickel, separating the CO from the CO2 and releasing the CO2 into the atmosphere.  The CO then travels through a radiator, a pump, and another radiator to increase the pressure and reduce the temperature.  The CO is then expanded to a cryogenic state and stored in at 80 K and 1.5 Bar.

Scenario I ISPP Details

In Scenario I, the ISPP plant must produce enough propellant to return the astronauts to Earth from the surface of Mars.  To inject the MTV into an Earth-bound trajectory, the TEI must contain 23,000 kg of propellant, the MOI requires 126,000 kg, and the MTV tanks must accommodate 103,000 kg of propellant.  The resulting propellant mass is approximately 212,000 kg, which must be produced within 580 days.

The optimum oxidizer to fuel ratio (O/F) required for the rocket engines is 0.55, leading to 161,000 kg fuel (CO), which is the limiting propellant in the production plant due to the O/F ratio of the plant being 0.57.  To produce both propellants in 580 days (the time between when the ISPP plant starts production and the astronauts leave Earth), CO must be produced at a rate of 280 kg/day.  An electrolyzer-adsorption bed combination exists at the University of Arizona, which produces 4.2 kg CO/day with an overall plant mass of 8 kg and a power input of 500 W [11].  This leads to the estimated power and mass values listed in Table 1.

Scenario II ISPP Details

To launch the approximately 3480 kg MAV (including crew and supplies) to LMO in Scenario II, 12,000 kg of propellant is produced.  Using a similar analysis as in Scenario I, the ISPP plant yields the values listed in Table 2.
 

WAVAR (Water Vapor Adsorption Reactor)

Another form of ISRU is the WAVAR [3], which is an adsorption/desorption device designed to extract water from the Martian atmosphere.  The unit designed by Coons et al. is capable of adsorbing 0.5 kg of H2O/day [3].  This unit requires a power input of 99 W and has a mass of 28 kg.  The average person uses 29 kg of water per day.  Of this, 90% can be recycled through the life support system, requiring the production of 12 kg of water per day.  As a safety measure, the water for the surface stay (500 days) and the transfer back to Earth (218 days) is produced before the astronauts leave Earth.  This amounts to 8100 kg of water to be produced in 580 days at a production rate of 14 kg/day.  For redundancy, enough WAVAR units to produce 20 kg/day are taken, although only enough units to produce 14 kg/day will be running at one time.  The resulting mass and power requirements of the WAVAR units are 785 kg and 2.8 kW respectively.
 

LAUNCH VEHICLES

A manned mission to Mars requires the transfer of large masses between Earth and Mars.  The inherent risks of manned space flight and the large financial investment necessary for large scale missions require this mass transfer to be accomplished while minimizing complexity and the possibility of failure.  These requirements lead to the necessity of a reliable Heavy Lift Vehicle (HLV) to accomplish TMI for both Scenario I and Scenario II.

One existing but out of production vehicle is the Russian Energia, and another existing concept (as yet not flown) is the Shuttle-C STS-based (Space Transportation System) vehicle.  They are the prime candidates to meet mission launch needs. Both vehicles are capable of lifting to LEO masses of 100 tons [12].  The Energia is a modular vehicle that, in its baseline configuration, consists of a core with four liquid hydrogen/liquid oxygen engines and four strap-on kerosene fueled stages.  The Energia has made two successful flights to date.  The Shuttle-C, based on proven Space Shuttle technology, is a heavy lift concept that has undergone significant development with the requirement of additional development to bring it to flight maturity.  Its basic configuration consists of an H2/O2 core with two strap-on solid rocket boosters and the requirement of two Space Shuttle Main Engine (SSME) booster engines attached to its payload.

Due primarily to its design maturity and low launch cost, the Energia is assessed as the best vehicle to meet mission launch needs.  The Energia provides a combination of lifting ability, simplicity, flexibility, and economy that make it well suited to meet the needs of a manned mission to Mars.
 

TRANSFER VEHICLES

Each scenario has three transfer vehicles.  To minimize development cost and integration problems with the launch vehicle, all transfer vehicles, except the Trans-Earth Injection (TEI) stage of Scenario II, have the same axially symmetric biconic shape with a circular cross-section (Fig. 4).  The axially symmetric biconic shape gives a high lift to drag ratio (L/D) which increases the size of the entry corridor and provides a cross-range capability for high accuracy landings on Mars.  An analysis of vehicle shape yields a maximum L/D of 1.78 at 10 degrees angle of attack.  Moreover, the L/D remains above 1 for angles of attack ranging from 3 degrees to 32 degrees.  The lift coefficient, CL, reaches a maximum value of 0.9 at 45 degrees angle of attack, and the drag coefficient, CD, is 1.45 at maximum CL.  These parameters provide sufficient capabilities to perform aerocapturing and aerobraking maneuvers without exposing the crew to extreme g-loadings.  The center of pressure remains between 55.5% and 58.0% for a wide range of angles of attack.

Scenario I Vehicles
 

ISPP Transfer Vehicle (ITV)

The first launch from the surface of Earth is the ITV, whose main function is to deliver the ISPP plant and its storage tanks, the nuclear reactor, WAVAR, science equipment, rovers, and various supplies.  A general mass breakdown of the main components of the ITV is shown in Table 3.

The ITV lands horizontally as shown in Fig. 5.  This horizontal landing configuration makes it easier for the rovers to deploy to the Martian surface.  Also, the horizontal landing provides more accessibility to the ISPP system and propellant storage tanks.  Once the ITV lands and is stabilized, the nuclear reactor is removed and setup, and production of propellant begins.  The ITV serves as the base of operations because it houses the ISPP plant and the science equipment.

Booster Transfer Vehicles (BTV)

For Scenario I, the BTV (Fig. 6) serves as the TEI stage and Mars orbit injection (MOI) booster using in situ propellant for both stages.  The BTV lands vertically on the surface and receives propellants from the ISPP plant for both its TEI stage and the MOI booster portions.  The BTV stage is launched into a Mars parking orbit, jettisons the MOI booster portion, and sheds the outer fairing.  The remaining TEI stage docks with the MTV and propels the MTV back to Earth.  The Earth launch mass breakdown of the BTV is shown in Table 4.

All the engines used for each stage are of a common design, with a thrust of 160 kN.  Design particulars of this engine are listed in Table 5.  The MOI booster uses four of these engines and the TEI stage uses one.

Manned Transfer Vehicle (MTV)

Two years after the launch of the ITV and BTV, the manned launch occurs.  The Manned Transfer Vehicle (MTV), shown in Fig. 7, carries a crew of four, their belongings, a life support system, a scientific laboratory, and tanks and propulsion system.  The MTV has a mass of approximately 33,000 kg for Scenario I (Table 6) with an interior habitat volume of about 113 m³.  Scenario I utilizes a single habitat inside the MTV for the entire mission (Fig. 7).  The living area is located above the propellant tanks, 8.3 m above the base of the of the structure.  Floor plans for three levels of the habitat are shown in Fig. 8.  Each level is 2.3 m in height, but the floor and ceiling surface areas vary due to the conical shape of the vehicle.  Level l contains a common living area for the crew.  Level 2 and Level 3 house staterooms, which have similar floor plans, but the diameter of Level 3 is slightly smaller than that of Level 2.  Laboratory equipment to be used on the Martian surface is located on Level 4.  Ladders provide access between the levels of the habitat and also access to the EDV, which houses the communications equipment.

From LEO, the MTV is propelled to Mars by the TMI stage.  The MTV aerobrakes into the Martian atmosphere and lands upright on the surface.  The vehicle is refueled with carbon monoxide and oxygen from the ISPP plant, and at the end of the surface stay launches into a LMO circular orbit, jettisons the tanks and engine assembly, and docks with the TEI stage.  The TEI stage injects the MTV into an Earth-return trajectory.
 

Earth Descent Vehicle (EDV)

The EDV resembles an Apollo command module with an ablative heat shield, and houses the controls and communications systems for the MTV.  When the MTV approaches Earth, the crew enters the EDV and separates from the MTV.  The MTV continues on its hyperbolic trajectory, while the EDV follows a ballistic re-entry flight path into the ocean.  Parachutes are used to slow the EDV prior to splashdown and a flotation ring keeps the capsule afloat until the crew is recovered.
 

Scenario II Vehicles

In Scenario II, only the first of the three transfer vehicles launched from Earth lands on the Martian surface.  This requires smaller transport vehicles to transfer the crew between the MTV and the surface of Mars.  Also, only the Surface Habitat Vehicle (SHV) and the Manned Transfer Vehicle (MTV) are biconic vehicles.  The TEI stage is a large blunt body vehicle, and the Mars Descent Vehicle (MDV) and Mars Ascent Vehicle (MAV) have dimensions and design similar to the EDV of Scenario I.

Surface Habitat Vehicle (SHV)

The SHV is the first vehicle launched from Earth, and contains the Mars Ascent Vehicle (MAV), surface habitat, ISPP plant, WAVAR, and solar arrays for surface power (Fig. 9).  The SHV lands horizontally on Mars.  Once on the surface, the solar panels are deployed and the ISPP plant begins producing the CO and O2 for the MAV.  Water and O2 are produced by the WAVAR and ISPP, respectively.  Due to the possibility of on-site assistance being required, the inflatable habitat addition is deployed after the astronauts arrive.  The inflatable section of the habitat extends perpendicular to the SHV, providing additional work space for the crew.  On the opposite side, an airlock provides access to the Martian surface.  All communication and life support equipment are located in the main vehicle structure with the staterooms in the event that the inflatable section begins to leak and must be sealed off for repairs.  A mass breakdown of the SHV is shown in Table 7.
 

Trans-Earth Injection stage (TEI)

The fully fueled TEI stage (Fig. 10) is launched during the same launch window as the SHV.  The TEI stage aerocaptures into a Mars parking orbit and then jettisons its aeroshell.  The TEI stage waits in orbit to dock with the MTV approximately two years later.  Once docked with the MTV, the TEI stage provides the thrust to propel the MTV back to Earth.  In order to avoid boil-off problems associated with long-term storage of cryogenic propellants, Aerozine-50 (50% hydrazine and 50% UDMH) and nitrogen tetroxide (N2O4) were chosen as the fuel and oxidizer respectively.

The engine chosen for this stage is the Aerojet LR91-AJ-5 [12], which is used in the second stage of the Titan IV.  This engine is a gas generator engine with an Isp of 316 sec and an O/F of 1.8.  It provides an average thrust of 462 kN that generates a g-loading of approximately 3.9 on the MTV at burn out.  A mass breakdown of the TEI stage is shown in Table 8.
 

Manned Transfer Vehicle (MTV)

The MTV, the final vehicle launched from Earth (Fig 11), provides the crew with a habitat during the journey to and from Mars.  The habitat is a 6.6 m diameter spherical enclosure.  The MTV also contains two 3.0 m by 16.7 m solar arrays, four OMS engines, RCS thrusters, and propellant tanks.

Located within the forward cone of the MTV is the MDV, which is connected to the habitat by a 1 m diameter docking tunnel to provide crew transfer between the vehicles.  The MDV provides guidance and control for the MTV in transit to and from Mars, as well as serving as the crew area during launch.  An additional guidance and controls system is located in the MTV for use during the orbit about Mars, after the MDV has descended to the surface.  A mass breakdown for the Scenario II MTV is shown in Table 9.

Once in LMO, the MTV is oriented for communication satellite deployment. Next, the aft section of the MTV docks with the orbiting TEI stage.  After successful docking, the crew transfers to the MDV and descends to the Martian surface.  After 1.5 years in the Surface Habitat Vehicle (SHV), the crew returns to the MTV via the Mars Ascent Vehicle (MAV).  The MAV docks with the MTV for the return to Earth.  This configuration is shown in Fig. 12.

The habitat in the MTV of Scenario II is a sphere inside the vehicle.  The spherical shape minimizes the surface to volume ratio, reducing the mass per unit volume of the structure and providing maximum radiation protection.  Because the habitat structure is a sphere, the walls and furniture have varied shapes.  The four staterooms are located on the lower floor, while the upper floor provides a common living area similar to Level 1 in Fig. 8 (outer diameter of 6.6 m).  The MDV houses the avionics and communication equipment.
 

Mars Descent Vehicle (MDV)

Only Scenario II utilizes a MDV to transport the astronauts from the MTV to the Martian surface.  The capsule is similar to the EDV of Scenario I, with the addition of a docking port through its bottom heat shield and retro-rockets for landing.  Minimal life support is provided on the MDV, thus the crew wear spacesuits during the descent to the Martian surface.  The MDV’s descent into the Martian atmosphere is slowed by parachutes, and retro-rockets are fired to decrease its velocity further.  Finally, airbags are deployed to softly land the capsule on the Martian surface.  When the vehicle lands on the surface, it is depressurized completely to open the side hatch.  A rover is mounted to the side of the MDV and the astronauts must assist in its deployment.  The rover is then used to transport the crew to the SHV.
 

Mars Ascent Vehicle

The Mars Ascent Vehicle (MAV), shown in Fig. 13, is the vehicle used to launch both the crew and the Martian samples into LMO, at approximately a 300 km altitude.  Once in LMO, the MAV drops its spent fuel tanks and carbon monoxide rocket engine.  The capsule then docks with the MTV, which is already docked with the trans-Earth injection (TEI) stage.  This same capsule is used as the vehicle for Earth re-entry.  The MAV engine has a thrust of 70 kN and uses a staged-combustion cycle for high performance.  The stage-combustion cycle results in a more complex and heavier engine than other cycles, but the increased Isp makes up for this penalty.  A mass breakdown of the MAV is shown in Table 10 and the results of the engine design are shown in Table 11.

TRANS-MARS INJECTION (TMI) STAGE FOR ALL VEHICLES

The TMI booster (Fig. 14) is used to inject the transfer vehicles from LEO into a Mars transfer orbit.  In both scenarios, the spent TMI booster is used as a counter mass to the MTV while tethering to generate artificial gravity.  At arrival at Mars, the TMI is released before the MTV aerocaptures into the Martian atmosphere.

The TMI uses the J-2 liquid oxygen/liquid hydrogen (LOX/LH2) engine used on the Saturn [12].  The J-2 has been chosen because its characteristics satisfy all the mission requirements.  The average vacuum thrust is 890 kN, the propellant mass flow rate is 213 kg/s, the O/F ratio is 5.5, and Isp 426 sec.  The TMI burn time is 262 sec.  The amount of propellant carried in the TMI is different for each vehicle.  However, the TMI has been designed to boost the TEI of Scenario II which is the most massive vehicle of the two scenarios.
 

AEROSHELLS & THERMAL PROTECTION SYSTEM

Aeroshells are used to meet the thermal protection requirements during aerocapture for all transfer vehicles.  For biconic spacecraft, the aeroshells (Fig. 15) are mounted on top of the aluminum skin of the spacecraft and consist of an ablating nose cap made of Avcoat-5026-39HC, followed by a reinforced carbon-carbon (RCC) section, and then by fibrous refractory composite insulation tiles (FRCI-12).  The carbon-carbon and the tile materials use radiative heat transfer to dispel the heating loads during atmospheric entry, while the ablator uses the heat of vaporization of the material to carry away the heat flux.  Ablative materials are required at the stagnation region of vehicles during Martian atmospheric entry.  The two reradiating aerobrake materials are existing, well-tested materials used on the Space Shuttle.  RCC is made of a carbon-carbon laminate impregnated with a phenolic resin, has a density of 1,650 kg/m^3, and is highly resistant to fatigue loading.  RCC is used on the aeroshell for heat loads in excess of 68 watts/cm2, which is the maximum loading for the FRCI-12 tiles [13].  RCC is abandoned in favor of the ablating material in regions where the heat fluxes are greater than 100 W/cm².  Avcoat was used on the Apollo reentry capsule.  It has a density of 512.6 kg/m³ [14].  The FRCI-12 tiles are made of mostly silica amorphous fibers, are roughly six inches square, and have a density of 170 kg/m³.  The aeroshells are constructed as units to be detached from the vehicles at various stages during the mission, depending on the function of the vehicle.

Avcoat, needed over the first 5% of the aeroshell, possess an effective heat of ablation of 47 MJ/kg [14].  From the modeling of aerocapture, nearly 20 kg of Avcoat is lost, corresponding to a depth of 0.4 cm.  The total ablating aeroshell has a depth of 2.0 cm and a mass of 77 kg.

The inner surfaces of the aeroshells never climb above 450 K.  From the temperature at the aeroshell outer surface, and the uniform allowable heat input to the vehicle of 3400 W/m² (calculated from an analysis of the transient heat loading to the interior of the ship over the 1000 second aerobrake maneuver), a thickness of the aerobrake was computed.  RCC was chosen to cover the next 5% of the vehicle, with FRCI-12 covering the remaining 90%, because of the maximum allowable heat loading to both materials.  The total mass of the heat shield, including the Avcoat section, was calculated to be 4,100 kg.  Fig. 15 shows a schematic of thermal protection system.

All nonbiconic spacecraft, including the Scenario I EDV and Scenario II TEI stage, MDV and MAV capsule, have blunt body aeroshells reminiscent of an Apollo command module.  The thermal protection system for each consists of a coating of FRCI-12 with an additional layer of Avcoat-5026-39HC on the blunt face of the aeroshell, where the greatest heating occurs during aerobraking maneuvers.
 

CREW HEALTH

Radiation and the lack of gravity are two major problems for humans in the space environment.  In both scenarios, to combat the loss of bone density that comes from weightlessness, artificial gravity is provided in transit to Mars by tethering the MTV with the TMI stage and rotating them at a rate of 2 rpm, which is acceptable for most humans [15].  This rotation rate provides a 0.4 g environment, requiring a tether length of 380 m with the center of rotation 90 m from the MTV.  Radiation is a major concern in transit between Earth and Mars, where the fluxes of Solar Cosmic Rays (SCR) and Galactic Cosmic Rays (GCR) are substantially greater than those on Earth's surface.  In both scenarios, most of the required radiation shielding is provided by a thin aluminum skin, polyethylene layer, and strategic placement of equipment on the perimeter of the MTV.  Additional protection is provided by oral chemical radioprotectants which can reduce the carcinogenic effects of radiation [16].  On the Martian surface, the CO2 atmosphere provides adequate radiation shielding [17].
 

Human Consumables

The mass penalties of bringing all human consumables for the entire mission are severe, therefore regeneration capabilities are incorporated (see Life Support section).  Recycling of oxygen and water results in a mass savings of more than 100,000 kg, based on human daily requirements of 0.84 kg of oxygen and 29.14 kg of water per person [18,19].  Each crew member also requires 0.62 kg of dehydrated food per day. The resulting consumable masses for the entire mission are listed in Table 12.

For Scenario I, oxygen and water for the trip to Mars and food for the entire mission are carried on the MTV.  The MTV in Scenario II carries oxygen and water for both transits and food for the entire mission.  In the event that the descent to the surface is aborted, the crew of Scenario II has sufficient food, oxygen, and water in the MTV to remain in orbit for a maximum of 1.5 years, or until the next opportunity arises to return to Earth.  On the surface, the crew of Scenario II consumes food stored on the SHV.
 

Life Support

Life support is provided by an Environmental Control and Life Support System (ECLSS) which recycles oxygen and water by physical-chemical separation processes.  An ECLSS was chosen over a biological system, in which plants purify water and produce oxygen and food, because the later requires 75 kW of power and 240 m² of plant growing area.  The ECLSS of each habitat has a mass of 2,920 kg, a volume of 20 m³, and a maximum power consumption of 9,020 W.  ECLSS subsystems include atmospheric revitalization, atmospheric control & supply, fire detection and suppression, temperature and humidity control, water recovery and management, and waste management [15] (Table 13), which can operate in both gravity and microgravity environments.  A mass and power inventory for the ECLSS is listed in Table 13.

Extravehicular Activities

Extravehicular Activities (EVA) require an Extravehicular Mobility Unit (EMU) which consists of a space suit assembly and a portable life support system (PLSS).  The space suit assembly has nine protective layers of fabric.  The PLSS provides oxygen from a pressurized tank, CO2 removal by a LiOH canister, odor removal by an activated charcoal filter, ventilation by a fan, cooling by a condensing heat exchanger, and power by batteries.  The mass of the EMU is 100 kg.

EMU pressures (4.3 psi) are maintained lower than that of the MTV and SHV cabin (10 psi) to allow maximum joint mobility and minimum fatigue.  To prevent decompression sickness, astronauts must pre-breathe 100% O2 for two hours prior to an EVA to purge the nitrogen from the body [16].
 

ROVER

The rovers for each scenario are designed for traversing a variety of terrain, collecting uncontaminated samples, and conducting scientific experiments.  Scenario I carries the Pressurized Martian Rover (PMR) which includes a pressurized living area (Fig. 16).  The life support capability of the living area allows for extended surface exploration, thus increasing its range of exploration.  Scenario II uses the Unpressurized Martian Rover (UMR) similar to the Lunar Rover Vehicle (LRV) used in the Apollo missions [20].  Because of its limited life support capability, the UMR has a smaller range of exploration. Both scenarios utilize micro-rovers to do initial surveying and mapping before the astronauts arrive.

Scenario I

The PMR in Scenario I consists of two components, a tractor propelled by CO and O2 and a habitable trailer.  It arrives on the Martian surface in the ISPP Transfer Vehicle (ITV).  The tractor is equipped with a bulldozer attachment and robotic arms for drilling and collecting samples.  It also houses the power plant, generator, fuel tanks, instrumentation and controls.  The power plant is a gas generator turbine that uses liquid carbon monoxide fuel, LOX oxidizer from the in situ propellant plant, and CO2 diluent [21].  The engine drives the generator, which supplies power to all systems on the rover.  These systems include a DC motor in each of the wheels, life support systems, heat rejection systems, computers and instrumentation, and a robotic arm.  The habitable trailer component has life support systems independent of the surface habitat.  The tractor and trailer components combine to form the PMR surface exploration system.  The trailer component also houses a back-up power plant, generator, and fuel tanks.  The maximum radius of exploration is 250 km.  The rover has a maximum speed of 8 km/hr.  The total mass of the PMR is approximately 4200 kg; a detailed mass breakdown is shown in Table 14.  The primary structures of the tractor and trailer components are made of 6061-T6 Aluminum alloy, proven for harsh temperature [20].
 

Scenario II

Because of limited space, Scenario II carriers a smaller unpressurized rover, the UMR, which is similar in design to the Apollo LRV.  It is transported aboard the MDV in a compact folded configuration and used by the crew for transportation to the SHV at Mars arrival.  For extended exploration, the vehicle is outfitted with auxiliary scientific equipment for inquiry and sample collection.  Like the LRV, a rechargeable battery provides vehicle power, but unlike the LRV, the UMR is designed for the 0.38 gravity environment of Mars.  The UMR is also designed to navigate a rougher, boulder strewn surface by incorporating a taller ground clearance height and a sturdier suspension system.  The vehicle range is restricted to a 10 km radius to allow for an emergency walk-back capability.  The maximum speed of the UMR is 5 km/hr.  A mass breakdown of the UMR in its full crew configuration used for transporting the crew from the MDV to the SHV, and its science configuration, capable of transporting two crew members, is listed in Table 15.

Micro-Rovers

Both Scenarios carry two micro-rovers.  Initial exploration of Martian terrain is accomplished with these rovers before crew arrival.  They are also utilized during the crew’s surface stay for teleoperated exploration.  Each micro-rover has a total mass of 14 kg.  The height is 28.0 cm with a ground clearance of 13.0 cm.  Each rover is 63.0 cm long by 48.0 cm wide [22].  Micro-rover technology will be tested on the Mars Pathfinder mission in 1997.
 

SCIENCE PACKAGE

A variety of experimental equipment is used to test the surface and atmospheric conditions on Mars.  Tests range from analysis of sub-surface mineral deposits and surface soil to detecting magnetic fields and atmospheric dust.  Soil samples are taken from an array of geological sites, such as a lava flow field and a suspected water eroded basin.  The rovers assist in reaching areas of special interest which might be away from the Martian landing site.

The science package is broken into four categories.  The first set consists of the experiments for in-transit study.  Plasma, particle, and wave detectors are included in this set [23].  Upon reaching the surface the crew collects soil samples for study.  The rover tools consist of rakes, scoops, a set of drills and other related equipment [23].  These sample are taken to a laboratory at the base for study.  The laboratory will include a variety of spectrometers and analyzers for the detection of various elements and compounds [23].  Placed somewhere near the base is a set of meteorological and dust monitoring equipment [24].  While exploring the surface of the planet the crew also carries out a set of field studies.  Field study equipment includes a magnetometer, seismic equipment, and balloons [23].  This package of experiments allows the astronauts to collect valuable scientific data.
 

POWER

The power requirements for Scenario I and Scenario II are very similar, except during the surface stay on Mars.  Both scenarios use photovoltaic cells and batteries while in transit to and from Mars.  Scenario I utilizes nuclear power on the surface of Mars due to the high power requirement of its ISPP plant (29 kW).  Scenario II uses solar photovoltaic cells and rechargeable batteries on the surface of Mars due to the low power requirement of its ISPP plant (1.8 kW).

The solar arrays used in both scenarios are made of gallium arsinide (GaAs) [25].  The cells have an efficiency of 22%, (beginning of life, BOL), and a specific power of 50 W/kg [25].  The array support structure for all the vehicles during transit is a retractable boom called a STACEBEAM (Stacking Triangular Articulated Compact Beam) [25].

The batteries used in both scenarios are Lithium Ion (LiIon) batteries [26] which have an energy density of 100 W-hr/kg with an average loss per cycle of 0.07%.  Their end of charge voltage is 4.1 V and they can cycle in excess of 800 times to 80% of their initial values [27].

The nuclear reactor used in Scenario I is a Space Power Advanced Core Element-Reactor (SPACE-R) [28].  The SPACE-R power system is based on TOPAZ II technology which uses thermionic power conversion.  The reactor has power outputs of 44 kWe and 611 kWt providing 30 V at 1,470 A DC payload.  The reactor, with its heat rejection radiators, is conical shaped with a height of 6.5 m, a base diameter of 0.75 m, and a top diameter of 2.6 m.  The total reactor system mass is 2,200 kg [28].

The size of the solar arrays is determined by the amount of power needed and the available solar flux.  The 490 W/m² solar flux during transit is modeled for the worst possible case, which occurs at the aphelion of Mars' orbit.  The solar flux on the surface of Mars depends on the latitude and the solar declination, which is an angle dependent on the orbit and seasons of Mars.  The average solar flux at 35° north on the surface of Mars (using its mean distance from the sun) is 160 W/m².

Dust storms also play an important role in the modeling of the solar flux on Mars.  Dust storms on the average decrease the solar flux by 50% [29].  Based on this, the average solar flux used for sizing the solar arrays on Mars is 80 W/m².

In Scenario I, the power requirement for the ITV is 400 W for controls and communication during transit.  To meet this, 50 kg of batteries and 6 m² of solar array are used.  Once on the surface the SPACE-R is used to run the ISPP plant. The SPACE-R is moved by the rover to a safe distance upon landing. The total power system mass is 7,000 kg.

In Scenario II, the SHV power requirements are again 400 W during transit.  The SHV uses the same power system during transit as the ITV.  On the surface, however, the SHV power requirement goes up to 9.6 kW for the life support system.  To supply this power, 600 m² (1,200 kg) of solar arrays are used.  The arrays are mounted on 4 inflatable structures which are rolled up prior to their deployment.  The arrays are inflated with tanks of pressurized gas which causes the arrays to unroll along the ground extending from the sides of the SHV.  A substantial pressure is provided to roll the arrays over rocks in its path.  During the night, 1,350 kg of batteries are needed to supply the power to the SHV.  The batteries are then recharged during the day by the solar arrays.  The total power system mass is 2,650 kg.

The BTV for Scenario I and the TEI stage for Scenario II have the same power requirement of 400 W.  They use the same power system during transit as the ITV.  The BTV in Scenario I, however, connects to the nuclear reactor once on the surface.  The TEI stage in Scenario II continues to use the photovoltaic power system throughout the mission.  The total power system mass for each vehicle is 100 kg.

The MTV’s in Scenarios I and II have the same power requirement of 9.6 kW for their life support system.  To meet this during transit, they both use 114 m² (200 kg) of solar arrays and 1,400 kg of batteries.  In Scenario I, the MTV connects to the nuclear reactor upon landing on the surface.  The Scenario II MTV utilizes the same system for the entire mission.  The total power system mass for both vehicles is 2,200 kg.  The ascent/descent vehicle for Scenario I and the earth reentry capsules for both scenarios require 2400 Whr which is provided by 24 kg of batteries.
 

COMMUNICATIONS

While in transit to Mars, each transfer vehicle is equipped with a small Low Gain Antenna (LGA) and a larger High Gain Antenna (HGA).  The LGA is valuable because it is omnidirectional, but it lacks the range to be effective at great distances.  The LGA is used to establish the connection with NASA’s Deep Space Network(DSN).  The ground lock established with the LGA is then transferred to the HGA.  The HGA remains in use for the entire transit and the LGA becomes a back-up system to the HGA.  An identical model is used for the transit of the MTV back to Earth [25].

To solve the problem of communication black-out, a satellite is inserted into a Mars synchronous orbit above the base in both scenarios.  The satellite operates using the X-band, with a data rate of 60 Mbps [30].  This 150 kg satellite is deployed in LMO upon arrival at Mars by the ITV in Scenario I and the MTV in Scenario II.  It then uses 100 kg of propellant to transfer to a stationary orbit.  The 1.3 kW needed to power the satellite is supplied by solar arrays and back-up batteries [25].  Once in orbit the satellite serves as the relay station for communication between the habitat, the rover, and Earth.
 

ASTRODYNAMICS AND AEROCAPTURE

The determination of optimal transfer orbits was performed using plots of launch opportunities provided by the Jet Propulsion Laboratory [31,32].  The design constraints are delta V requirements, transfer time, and hyperbolic excess velocity at arrival.  Conjunction class Type I minimum energy orbits, i.e. orbits that have a true anomaly at arrival of less than 180°, satisfy these requirements during the mission time frame of interest (2013 - 2020).

The data for the transfer orbits are all based on 20 day launch windows in which the C3 and V infinity values used are the highest in the windows.  The different launch opportunities with launch dates and estimated time of flight are shown in Table 16.  The entry velocities for Mars and Earth were calculated from the data from the JPL plots.  An altitude of 125 km was chosen as the atmospheric entry interface for both Mars and Earth.

Scenario I

For each transfer vehicle, the Earth launch vehicle (Energia) injects the spacecraft into a low Earth orbit where a tangential burn is performed using the TMI stage, sending the transfer vehicle to Mars.  Upon arrival at Mars, the spacecraft performs an aerocapture maneuver into a 300 km altitude circular parking orbit.  The ITV and BTV enter the Martian atmosphere and land using parachutes and retro rockets.  The MTV follows two years later using an identical entry procedure.  The aerocapture maneuver was calculated for the highest of the entry velocities at Mars, 7.8 km/s.  The spacecraft have a flight path angle of 11.6° at the entry altitude.

The aerocapture maneuver has a minimum altitude of 33 km, a maximum g-loading of 4.2 (Fig. 17), maximum heat flux of 160 W/cm² (Fig. 18), and a maximum stagnation point temperature of 2400 K.  The g-loading and heating rates noted above are the highest experienced for any of the vehicles entering the Martian atmosphere.  The initial conditions for re-entry at 125 km altitude are a velocity of 3.6 km/s and a flight path angle of 1.0°.  The aerobraking maneuver is modeled using a computer code for lifting entry.  The maximum g-loading is 1.1, maximum heat flux is less than 4 W/cm², and the maximum stagnation point temperature is 980 K.

On the return trip, the launch from the Martian surface requires a total delta V of 4.27 km/s to inject the spacecraft into a 300 km LMO.  The TEI burn requires a DV of 2.02 km/s.  Upon arrival at Earth the astronauts enter the EDV for an Apollo-style entry.  The entry velocity is 11.7 km/s and the flight path angle is 5.9° at an altitude of 125 km.  The maximum g-loading during entry is 6.8, but the deceleration stays above 4.8g for approximately 180 seconds.  The maximum heat flux is 86 W/cm², and the maximum stagnation point temperature is 2100 K.  The EDV releases parachutes to further decelerate the spacecraft before splashdown in the ocean.
 

Scenario II

Many of the maneuvers employed in Scenario II are identical to those used in Scenario I, consequently, only the maneuvers that differ are described here.  The entry parameters used for the analysis of the TEI stage are the same as those used in Scenario I, except for the flight path angle being 11.8° at 125 km altitude.  This results in a minimum altitude of 32 km above the surface of Mars.  The maximum g-loading is 3.3, the maximum heat flux is 54 W/cm², and the maximum stagnation point temperature is 1850 K.  The TEI stage must perform a burn upon reaching apoapsis after the aerocapture maneuver in order to reach Mars parking orbit.

The MTV aerocaptures into LMO and the astronauts enter the MDV for the entry maneuver to reach the Mars surface.  The maximum g-loading during entry is 3.3, the maximum heat flux is 3.5 W/cm², and the maximum stagnation point temperature is 950 K.
 

CONCLUSIONS

Ares Explore covers two possible scenarios for a manned mission to Mars.  Both scenarios utilize in situ resource utilization (ISRU) technologies to provide an economically feasible approach to manned Mars exploration.  The scenarios include many common elements.  Both use carbon monoxide as an ISRU-produced propellant, both use three Energia launches from Earth, and both use axially symmetric biconic transfer vehicles.  The scenarios also differ in a number of significant ways.  Scenario I uses all ISRU-produced propellant for return from Mars while Scenario II uses both ISRU-produced propellant and terrestrial propellant for the return journey.  This leads to two different approaches to Mars descent/ascent. Scenario I utilizes a single large biconic vehicle, the Manned Transfer Vehicle, for crew descent and ascent to and from low Mars orbit and Mars’ surface.  Scenario II utilizes small vehicles, the Mars Descent Vehicle and Mars Ascent Vehicle, for crew descent and ascent at Mars.  These differences lead to advantages and disadvantages for both scenarios.

Because of its use of all ISRU-produced propellants for the return journey, Scenario I is capable of taking a large pressurized rover for extended surface exploration, and also allows the use of a single habitat for the crew for all phases of the mission, leading to simplicity but also a lack of redundancy.  While providing the valuable benefit of extended surface exploration with a pressurized rover, the architecture of Scenario I requires a nuclear reactor for generating the larger power requirements of its in situ propellant production plant.

Scenario II, while having a reduced surface exploration capability with an unpressurized rover, has a much smaller ISRU propellant requirement, thus lessening the development challenges of moving from small unmanned ISRU precursor missions to larger manned ISRU missions.  Furthermore, Scenario II uses solar arrays, rather than a nuclear reactor, to meet its power needs on Mars.  The use of small vehicles for Mars descent/ascent leaves the Scenario II Manned Transfer Vehicle in low Mars orbits, and together with the Scenario II Surface Habitat Vehicle on the Martian surface, provides two separate and complete habitats at Mars.  This leads to complexity, but also provides redundancy and gives a greater margin of safety for the crew.  If one scenario can be judged as more feasible in the near term than the other, Scenario II, with its lower reliance on ISRU-produced propellant and its greater margin for safety, is the likely candidate.

Both scenarios provide feasible methods of achieving a manned mission to Mars, and one day humans may travel to Mars using techniques similar to those outlined in this paper.  Sending people to Mars is the next practical step following the missions in the Mars Exploration Program.
 

ACKNOWLEDGMENT

The authors wish to acknowledge the contributions of the members of the University of Washington’s 1996 AA 420/421 Space Systems Design class: Matt Arbogast, John Beasley, Matt Buffaloe, Tom Chi, Myron Chornuk, Mike Forney, Jens Gjestvang, Chris Hoffman, Chan Hua, Jason Losey, Rob Osborne, Susana Quintana, Lisa Reid, Rob St. Clair and Quang Tran.
 
 
 
 

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