Monday, October 22, 2018

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

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