NASA Produces Oxygen on Mars

Technicians  lower the Mars Oxygen In-Situ Resource Utilization Experiment (MOXIE) instrument into the belly of the Perseverance rover.

(Credits: NASA/JPL-Caltech)

On April 20th on the alien surface of Mars, NASA's Perseverance rover produced oxygen from the Red Planet's carbon dioxide atmosphere-- emitting carbon monoxide as a waste product. 

 

 Links and References

NASA’s Perseverance Mars Rover Extracts First Oxygen from Red Planet

NASA's Perserverence Rover Generates Oxygen on Mars in Amazing First For Science

NASA develops 'MOXIE', a device that generates oxygen from the atmosphere of mars

Rover Makes Oxygen From Mars Atmosphere

Evaluating Lockheed Martin's Reusable Lunar Lander and Orbital Propellant Depot Concept

Notional  reusable lunar landing spacecraft on the lunar surface (Credit: Lockheed Martin)

by  Marcel F. Williams 

At the 69th International Astronautical Congress held in Bremen, Germany this month,  Lockheed Martin  unveiled a new reusable lunar crew lander concept.

For simplicity,  I'll designate the notional Lockheed Martin spacecraft discussed in this article as the R-LL (Reusable Lunar Lander).   According to Lockheed Martin, the R-LL will have dry weight of 22 tonnes and be capable of storing up to 40 tonnes of LOX/LH2 propellant. The R-LL will have up to 5 km/s of  delta-v capability.

Lockheed Martin argues that the R-LL should be capable of crewed round trip  missions to any area of  the lunar surface from NASA's future Deep Space Gateway (DSG) which is to be located at a Near Rectilinear Halo Orbit (NRHO).    Such round trip missions, they argue,  would also be capable of delivering up to one tone of payload to the lunar surface in addition to a crew of four individual astronauts. 



While Lockheed Martin has been rather vague about the exact dimensions of the R-LL, they have indicated that it will consist of only two cryotanks and will be derived from the Centaur upper stage family and its descendants. They also suggest that the R-LL will have a diameter close to that of  the future Orion spacecraft.

Since Lockheed Martin's Centaur V is currently in development as the future upper stage for the ULA's future 5.4 meter in diameter Vulcan rocket, one might speculate that the diameter of the R-LL cryotanks might be the same as  and  is supposed to have the same 5.4 meter diameter as the Centaur V. Such large diameter liquid hydrogen and liquid oxygen tanks should be capable of easily accommodating the 40 tonnes of propellant required for the R-LL. So deriving the lunar vehicle from the Centaur V cryotanks might be the simplest and cheapest path towards rapidly developing the R-LL.

Lockheed Martin's Notional  Reusable Crewed  Lunar Landing Vehicle

Propellant: 40 tonnes of LOX/LH2

Inert Weight: 22 tonnes

Engines: Four RL-10 derived engines

Maximum delta-v capability: 5.0 km/s

Maximum number of crew: Four

Additional cargo capability: one tonne of additional cargo
The R-LL would use four engines to provide engine out capability. This would enhance crew safety during attempted landings in case of a serious malfunction with one of its engines. So just two counter balancing engines could be used during a landing in case of single malfunction engine.     Lockheed Martin says that engines for the R-LL  would be derived from  Aerojet Rocketdyne's  RL-10 family or from Blue Origins restartable BE-3 engine. Aerojet Rocketdyne's RL-10 derived CECE engines would be  capable of at least 50 restarts with a throttling range from 104 percent to  just eight percent of thrust. 

Departing from the Deep Space Gateway, it would take approximately 12 hours for the R-LL to reach any point on the lunar surface. Another 12 hours would be required for the R-LL to return to the  gateway at NRHO.

NRHO: (Near Rectilinear Halo Orbit):

Travel time to and  from LEO:~5 days from LEO (3.95 km/s)

Station keeping: 5 m/s per year

Travel time to and from LLO:~ 12 hours to LLO (0.730 km/s)

Lockheed Martin says that their notional lunar spacecraft would be capable of accommodating  a crew of four astronauts on the lunar surface for up to two weeks. Such a lengthy stay would require at least four tonnes of additional shielding mass to protect astronauts from the inherently  deleterious heavy nuclei component of cosmic radiation and from a major solar flare. So one would assume that such enhanced radiation shielding would be part of the notional space vehicle's 22 tonnes of inert mass.

Lockheed Martin has also suggest that propellant depots could be co-orbited with the Deep Space Gateway so that the R-LL can be refueled at NRHO.

The simplest propellant depots would probably have to be utilized within a month after deployment to NRHO since approximately 3.81% of its liquid hydrogen and 0.49% of its liquid oxygen would boil off within a months time. For the 40 tonne LOX/LH2 requirement for the R-LL, such propellant depots would probably have to NRHO by the SLS or the BFR.

More sophisticated propellant depots could be equipped with cryocoolers and solar arrays capable of re-liquefying fuel boil-off.  Ullage gases from the boil-off of liquid hydrogen could be used to re-liquefy gaseous oxygen while 12 to 15 kWh of electricity would be needed to liquefy one kilogram of gaseous hydrogen. The 5.7 tonnes of liquid hydrogen required for a lunar mission would lose more than 217 kilograms of LH2 per month (7.2 kilograms per day).  But a 10 kWe solar  array deployed to NRHO capable of producing more than  240 kWh of electricity per day would be capable of re-liquefying 16 to 20 kilograms of LH2 per day.  The solar arrays for the Orion spacecraft will be capable of producing more than 11  kW of electric power. So it should be rather simple to deploy propellant depots already equipped with cryocoolers and and solar panels in order to prevent fuel boil-off. 

Solar powered depots that simply re-liquefied its ullage gases and powered pumps for storing and transferring liquid fueles would only  require the continuous delivery of liquid hydrogen and liquid oxygen.  Future Vulcan Heavy/Centaur rocket could deliver 7.3 tonnes of liquid hydrogen or oxygen to NRHO per launch. Monthly launches could deliver more than 87 tonnes of propellant to depots located at NRHO per year, more than enough for two R-LL missions to the lunar surface per year.

Notional propellant producing water depot (Credit: Lockheed Martin)

The most technologically complex propellant depots could use solar power to  actually  produce liquid hydrogen and liquid oxygen directly from water. This would require the addition of an electrolysis plant plus substantially more solar power.  A 375 KWE solar array proposed by Lockheed Martin could produce 40 tonnes of liquid hydrogen and oxygen propellant at NRHO per month. Such huge 375 KWE solar arrays would weigh  less than four tonnes. And two such arrays could be directly delivered to NRHO with a single SLS launch. But much smaller commercial launch vehicles could deploy 300 KWe arrays to LEO for later transport to NRHO by fueled upper stages deployed to LEO. 300 KWE arrays at NRHO could produce 40 tonnes of propellant in five or six weeks rather than just four weeks for the larger arrays.

Solar powered propellant producing water depots would make it much simpler and safer for commercial rockets to deliver fuel to NRHO since the payload would only be water. Propellant producing water depots at NRHO could eventually be supplied with water from the lunar poles.

Of course, water and propellant being produced on the lunar surface itself would dramatically reduce the amount of propellant required for   R-LL departures from NRHO. Reusable tanker vehicles directly derived from the R-LL could deliver more than 40 tonnes of lunar water  to propellant producing water depots at  NRHO per flight.  Just 12 round trips from the lunar surface could deliver enough water to NRHO to manufacture enough fuel for crewed missions to the orbits of Mars or Venus.

Lockheed Martin envisions that astronauts would be deployed to the NRHO gateway via the Orion and the Space Launch System. And then the would take the R-LL to the lunar surface and back to the NRHO gateway. And then they would take the Orion back to Earth.

However, propellant depots deployed at LEO  would make SLS crew launches of the Orion vehicle obsolete.  Refueling at LEO, the R-LL would have more than enough delta-v capability to transport crews from LEO to the  NRHO gateway. And refueling at NRHO, the R-LL would, of course, be capable of returning crews from NRHO back to LEO.  And even with  22 tonne of inert weight, a  5.4 meter in diameter R-LL could be launched to Leo aboard a Vulcan/Centaur launch vehicle within  a  6.4 meter in diameter payload fairing.

So for trips to the lunar surface, astronauts would simply take a Commercial Crew Launch vehicle (Falcon9/Dragon or Vulcan/Centaur/CST-100) to a commercial space habitat at LEO where a propellant depot refueled R-LL was already docked and ready to be boarded.  The R-LL would leave LEO with enough  propellant to take its crew on a 5 day journey to the NRHO gateway where another already depot fueled R-LL would already be docked.  The second R-LL  would take the crew for a round trip to the lunar surface, 12 hours to reach the surface and 12 hours to return to astronauts to the Deep Space Gateway.  The astronauts would return to the gateway with the first R-LL already fueled for their return to a commercial space station at LEO. The Crew would than take a Dragon or CST-100 Starliner back to the Earth's surface.

Such an architecture would, finally,  allow the SLS to be used--exclusively-- as a super heavy lift cargo transport. Such payloads could include: large and spacious microgravity and artificial gravity habitats derived from SLS propellant tank technology,  large water and propellant depots derived from SLS propellant tank technology, interplanetary spacecraft capable of accommodating at least 400 tonnes of propellant derived from SLS propellant tank technology for crewed missions to the orbits of Mars and Venus, 8 meter in diameter space telescopes exceeding the capability of the James Webb telescope,  and large inflatable microgravity and surface habitats that could make it a lot more spacious and comfortable for future astronauts and tourist to live under artificial gravity conditions in space or on the hypogravity surfaces of the Moon and Mars.


Links and References

Concept for a Crewed Lunar Lander Operating from the Lunar Orbiting Platform Gateway

Lockheed Martin unveils lunar lander concept

Cis-Lunar Gateways and the Advantages of Near Rectilinear Orbits

Lockheed Martin's Reusable Extraterrestrial Landing Vehicle Concept for the Moon and Mars

Utilizing Lunar Water Resources for Human Voyages to Mars

At EML4, an OTV-400 interplanetary booster  undocks with a fuel depot (WPD-OTV-400) before proceeding to dock with the  Odyssey interplanetary spacecraft. 

by Marcel F. Williams

Long interplanetary journeys to  Mars, or to the orbit of Venus, or even to some of  the NEO asteroids could take several months or even years before their human occupants finally return to the relative safety of the Earth's surface. Such interplanetary voyages will require a substantial tonnage of water  for drinking, washing, the preparation of food,  the manufacturing of oxygen for air,  the production of  liquid oxygen and hydrogen for propellant, and for appropriately shielding humans inside of  habitat modules from the dangers of cosmic radiation and major solar events.

NASA delta-v estimates for  achieving Mars Transfer Orbits from LEO during the 2030s range above 3.8 km/s to just below 5 km/s-- just to reach the vicinity of Mars. Additionally,  a LEO departure for Mars would require launching the huge amounts of water and propellant for the voyage out of the Earth's enormous gravity well. This would require a delta-v ranging from 9.3 to 10 km/s.

However, if  crewed interplanetary vehicles are launched from an Earth-Moon Lagrange point, the delta-v requirements to achieve a Mars Transfer Orbit would be substantially lower. The delta-v requirements for departing cis-lunar space from one of the Earth-Moon Lagrange points could be less than 2 km/s-- especially if an Earth flyby or an Earth and Moon flyby (even better) are taken advantage of during the initial trajectory burns.

Providing water and propellant for  crewed interplanetary mission from the lunar surface to an Earth-Moon Lagrange point would also have a substantially lower delta-v requirement.  A  delta-v of less than  2.6 km/s would only be required to supply water and propellant to an interplanetary gateway at an  Earth-Moon Lagrange point. Contrast that with the  enormous  9.3  to 10 km/s delta-v that is need to  supply propellant and water to an interplanetary vehicle located at LEO.  The lunar supply of water and propellant to an interplanetary spacecraft at an  Earth-Moon Lagrange point would also have the additional economic advantage of being able to use single stage reusable vehicles. 

Earth-Moon Lagrange Points, the optimal gateways to interplanetary space within cis-lunar space.

So  launching  crewed interplanetary space craft to Mars from one of the Earth-Moon Lagrange points (L1, L2, L4, or L5) has a substantial delta-v advantage over launching crewed interplanetary spacecraft from LEO--  if such spacecraft are supplied with fuel and water from the surface of the Moon. 

During the 2030s, entering High Mars Orbit during a Conjunction Class Mission (330 to 560 day stays) would require an additional delta-v of 0.9 to 1.9 km/s, while entering High Mars Orbit during  Opposition Class Missions (60 day stay)  would require an additional delta-v trajectory burn ranging from  0.9 to 3.8 km/s. The additional delta-v requirements for reaching High Mars Orbit adds further support for minimizing the initial delta-v requirements when departing from cis-lunar space. So, again, supply fuel and water from the Moon from an Earth-Moon Lagrange point gateway would appear to be the optimal way to begin crewed interplanetary journeys. 

But how much water is there on the lunar surface? And how technologically difficult would it be  to extract large quantities of water from the lunar regolith?

A spectral analysis of the ejecta plume from the impact of  a Centaur upper stage  into the  Cabeus crater at the lunar south pole was conducted in 2009.  The analysis indicated that the lunar regolith in the shadowed crater contained water ice with concentrations ranging from 2.7% to 8.5% by mass. This suggest that  in some of the permanently shadowed craters at the southern pole, the lunar regolith there may contain as much as  27 to 85 kilograms of ice per tonne. 

The dark purple and blue areas represent neutron emissions from the Moon's polar regions that indicate  hydrogen-rich regions on the lunar surface covered by desiccated regolith (Credit: NASA) .  

In the Moon's northern polar region, the Mini-RF on board the Chandrayaan-1 orbiting probe strongly suggest that 40  craters in the northern polar region my contain as much as 600 million tonnes of water-ice. So there is clearly no shortage of  water resources on the lunar surface.

In order to provide enough water for human activities at a lunar outpost, a cis-lunar transportation system, and to supply propellant and mass shielding for five Conjunction or Opposition Class missions to Mars during the 2030s: 2030, 2033, 2035, 2037, and 2039, at least 500 to 1000 tonnes of water is going to have to be annually manufactured on the lunar surface.

A lunar water and fuel manufacturing depot along side of mobile LOX and LH2 storage tanks and a mobile microwave water extraction robot (Water Bug).

NASA researchers have demonstrated that a simple one kilowatt microwave oven could extract as much as a tonne of water from the regolith at the lunar poles over a one year period. A single mobile robot with a 100 kw powered microwave oven, therefore, should be  able to annually extract 100 tonnes of water from the regolith from the shadowed areas at the lunar poles. Ten such mobile microwave units might be able to extract 1000 tonnes of water per year from lunar polar regolith resources.

Solar extraction of water from lunar regolith brought from permanently shaded lunar craters at the lunar poles. Transparent domes allows sunlight to heat a layer of lunar regolith from the top while solar heated metal tubes below filled with methanol heat the regolith from below. Water vapor is deposited within regolith insulated cold trap canisters.

However, a simpler method may only require sunlight to passively extract water from the lunar regolith. If one tone of regolith from the shadowed areas of the  lunar poles is composed of  approximately 5% water ice then mobile lunar excavators capable of digging up and depositing at least one tonne of regolith into a solar heater could produce at least 50 kg of water per day (18 tonnes of water per year). Just a few electric powered or fuel cell powered excavation robots on the lunar surface could, in theory,  deposit at least one ton of icy regolith to a solar heater per hour (more than 400 tonnes of water per year).

A reusable water tanker shuttle on a microwave sintered launch pad after being loaded with lunar water and lunar fuel for its flight to a water and propellant depot at one of the Earth-Moon Lagrange points.

The Odyssey interplanetary vehicles would be crewed spacecraft capable of transporting 12 astronauts to High Mars Orbit and back to cis-lunar space using solar photovoltaic powered propellant producing water depots supplied with water from the lunar surface. The Orbital Transfer Vehicle (OTV-400) and the IAGH (Interplanetary Artificial Gravity Habitats) would both be derived from the SLS fuel tank technology.

 The OTV-400 would use  a common bulk-head LOX/LH2 fuel tank derived from the SLS fuel tank technology. An IVL(Integrated Vehicle Fluids) type of technology would be used to utilize ullage gases for attitude control. Space Works has proposed a similar type of OTV using IVL technology that would store more than 450 tonnes of LOX/LH2 fuel. However, the OTV-400 would also use photovoltaic powered cryocoolers to  virtually eliminate any hydrogen and oxygen boil-off.
 
The twin habitat modules of the IAGH would rotate to produce a 0.5g simulated gravity for the six humans inside each module. The IAGH would also  be  appropriately water shielded to protect  its most radiation vulnerable occupants (25 year old female astronauts) from excessive exposure to  cosmic radiation and its  heavy nuclei component in a addition to major solar events-- for up to four years-- during solar minimum conditions. 50 cm of water shielding would be required to appropriately shield the light weight twin SLS fuel tank derived living areas: 118 tonnes of water shielding for each habitat module, 236 tonnes of shielding in total. 

A reusable Orbital Transfer Vehicle (OTV-400) for transporting crewed artificial gravity habitats to the orbits of Mars, Venus, or to the NEO asteroids. While some of the ullage gases are utilized for attitude control, most of the ullage gases are reliquified by photovoltaic powered cryocoolers to prevent fuel loss, a technology already developed by NASA.

Conjunction Class Missions would only require one reusable OTV-400  storing close to 400 tonnes of LOX/LH2 propellant. The higher delta-v Opposition Class Missions will require two reusable OTV-400 boosters. Entering orbit around Mars and reentering cis-lunar space will also require the IAGH  to dump the water shielding from its habitat modules in order to substantially reduce the Odyssey's mass just before the final trajectory burns to enter Mars orbit or to enter cis-lunar space.

Coupled with a large water storage tank and a photovoltaic powered electrolysis plant  and cryocoolers, the OTV-400 would function as a water storage and hydrogen and oxygen producing propeelant depot (WPD-OTV-400). Still equipped with its own rocket engines, it could self deploy itself practically anywhere within the inner part of the solar system  while still being able to manufacture LOX and LH2 anywhere where there is a source of water.

The WPD-OTV-400 water and propellant depot would have the ability to transport itself to Mars orbit from the Earth-Moon Lagrange points after producing enough fuel for its on flight and filling up with enough stored water originating from the Moon. However,  once water is being manufactured on Deimos and Phobos, using the same technologies employed on the lunar surface, propellant from the lunar surface will no longer be required to replenish water and propellant supplies in orbit around Mars. 

Reusable Odyssey I interplanetary space craft with a crew of 12 at EML4 in a trajectory burn configuration for a Conjunction Class mission to High Mars Orbit  in the year 2033. 

After its initial trajectory burns on its way to Mars, the  OTV-400 boosters and the ETLV-2 vehicles (Extraterrestrial Landing Vehicles) would separate from the IAGH  and re-dock at its central axis.The Odyssey would, therefore, be reconfigured  to produce artificial gravity for its 12 person crew for their multi-month journey to Mars.   Liquid carbon dioxide rockets housed in each habitat module, would be used to rotate or to stop the rotation of the Odyssey. Cables will extend from the IAGH core more than 100 meters from the central axis of the vehicle. A series of light weight cylindrical metal or ceramic shells woulds also expand outwards creating rotational arms that would act as levers to increase, decrease, or stop the rotation of the Odyssey. 

Reusable Odyssey II interplanetary space craft with a crew of 12 at EML4 in a trajectory burn configuration for a 1000 day Opposition Class mission to High Mars Orbit (60 day stay) to explore the martian moons Deimos and Phobos while also deploying  satellite constellations at Sun-Mars L1 and Sun-Mars L2 to provide global communications for future human missions to the surface of Mars.

After several months of travel through interplanetary space,  the Odyssey would reconfigure itself again to prepare for an Orbital Insertion trajectory burn and the water shielding within IAGH modules would be dumped into space.  But after a day or two in Mars orbit, the Odyssey will once again reconfigure itself to produce artificial gravity for its crew. The  water shielding for the habitats could be fully  restored within a few hours or a few days from a pre-deployed  WPD-OTV-400 already in high Mars orbit. Returning to Earth will also require the Odyssey to be refueled by the orbiting water and propellant depot in Mars orbit.

Once in high Mars orbit, each fully fueled  crewed ETLV-2 vehicle would have enough fuel to travel to one of the Martian moons for a few days of exploration and sample retrieval  and back to the Odyssey. And if they wanted to travel to the martian moons a second time then they could refuel at the orbiting Mars depot (WPD-OTV-5).

A slowly rotating Odyssey II in an interplanetary configuration to provide a simulated gravity of  0.5 g  for six crew members in each of the  IAGH habitat modules. 

Since the Odyssey is intended for reuse (ten times with its RL-10 engines) for future interplanetary missions, a trajectory burn will be required to return to the Earth-Moon Lagrange points after the trajectory burn for Trans-Earth Injection from Mars orbit. Again, this will require the Odyssey to completely dump its water shielding before the burn. Once the crew is back within cis-lunar space, they will only be a few hours away from protective shelters on the lunar surface or just a few days away from the Earth's surface.

After the 12 person crew completes their 60 day mission in High Mars Orbit, the Odyssey II converts from it's artificial gravity configuration to a trajectory burn configuration for its Trans-Earth Injection burn to begin its journey back to cis-lunar space.

 Because the major Odyssey components are reusable, the recurring cost for the interplanetary vehicle should be substantially lower that other interplanetary vehicle concepts that utilized expendable interplanetary boosters. The Odyssey's RL-10 or RL-10-like rocket engines could also be periodically replaced after perhaps ten round trips between cis-lunar space and Mars orbit which could further reduce their recurring cost.

The human safety advantages of using  lunar water resources for an SLS derived reusable artificial gravity producing  interplanetary spacecraft   deployed at one of the Earth-Moon Lagrange points should also be substantial:


1. The significantly reduced delta-v requirements at an Earth-Moon Lagrange point for fueling and mass shielding a crewed interplanetary vehicle should alleviate any pressure to significantly reduce the appropriate mass shielding of habitat modules against the dangers of cosmic radiation and major solar events. 50 cm of water shielding should also eliminate the possibility of space career ending radiation exposure during a single interplanetary mission for the most vulnerable occupants to radiation.

2. The rotating interplanetary artificial gravity habitats (IAGH) could significantly or totally eliminate long periods of exposure to the  deleterious physical effects of a microgravity environment

3. Having twin AGH habitats allows astronauts the enhanced safety of being able to seek refuge in the opposite habitat in case there is a serious life safety malfunction at the other habitat.

4. The more comfortable accommodations of the spacious SLS derived artificial gravity habitats could significantly reduce the psychological stress experienced by the drew  during several months or years of interplanetary travel.

5. The dangers and the complexity of direct, high-energy aerobraking into the atmospheres of Mars or during a return to directly to Earth would be avoided by limiting the Odyssey's flight path to travel only between the Earth-Moon Lagrange points and High Mars Orbit.

6. With at least one partially fueled  ETLV-2 (Extraterrestrial Landing Vehicle) connected to the Odyssey, any major malfunction of the OTV-400 during an orbital capture trajectory could allow astronauts to safely enter Mars orbit or cis-lunar space via the ETLV-2 vehicle or vehicles.  Even though this would mean the loss of the Odyssey spacecraft, the crew could still safely enter Mars orbit or cis-lunar space via the ETLV-2. However, Entering Mars orbit aboard an ETLV-2, without the Odyssey IAGH would require that an appropriately mass shielded space station already be pre-deployed in Mars orbit.



References and Links


Considerations for Designing a Human Mission to the Martian Moons (NASA)

A Study of CPS Stages for Missions beyond LEO (Space Works)

Mission and Implementation of an Affordable Lunar Return (Spudis & Lavoie) 

 Using the resources of the Moon to create a permanent, cislunar space faring system (Spudis & Lavoie)

Evolving to a Depot-Based Space Transportation Architecture (ULA)

 SLS Fuel Tank Derived Artificial Gravity Habitats, Interplanetary Vehicles, & Fuel Depots

 Utilizing the SLS to Build a Cis-Lunar Highway

Cosmic Radiation and the New Frontier

The Market for Space Fuel from Asteroids