Whatever happened to single launch crewed Moon missions?
This post is about pitching a bit of alternative history -- a wistful what-could-have-been. But to do that, I need to give a bit of background.
A Bit of History
And some know about the limitations and idiosyncrasies and heritage of Artemis, the modern-day crewed lunar program of the US and international partners. But not all do, so I'll provide a bit of background.
Artemis is the programmatic descendant of Constellation, an abandoned Bush II era program that was described by its chief proponent, then NASA Administrator Michael Griffin, as "Apollo on Steroids". A bigger capsule and a bigger lander, launched on a bigger rocket would take 4 astronauts, rather than 2, to the lunar surface for longer stays in greater safety and comfort.
It turned out to be too difficult to develop a human-rated rocket powerful enough to launch the chosen big capsule (Orion) and big lander (Altair) to the Moon in one shot using existing first stage engines, so Constellation settled on two launches by two to-be-developed rockets based in part on Shuttle hardware. The smaller Ares I would launch Orion with its service module, while the big Ares V would launch the Earth Departure Stage and the Altair lander. The two would rendezvous in LEO (Low Earth Orbit) and continue on to the Moon.
Apollo involved launching around 45 tons of spacecraft and propellant to the Moon. Constellation would have required over 70.
This all ended up being deemed too expensive and Constellation was cancelled under President Obama, but as a political compromise, some of the components and projects were retained/transformed largely for the sake of aerospace jobs that would otherwise have been lost.
Altair was cancelled, but Orion was retained. Ares I and Ares V were cancelled, but a pared down version of Ares V became the SLS. And Orion and SLS became the foundation of Artemis.
Now the US had hardware in development to get 4 astronauts to lunar orbit but no hardware to land them, until the Human Landing System program was spun up at the eleventh hour to see if private industry could put together something cheap enough for Congress to pay for that could enable Artemis to meet Constellation's goals of being Apollo on Steroids.
That lander will need its own ride to the Moon and represents the majority of the mass requirements of Artemis. This is due to the fact that unlike with Apollo, Constellation did not assign the work of lunar orbit entry to the service module of Orion but to the more efficient descent stage of the Altair lander. Without that lander, Orion can only go in and out of a much higher elliptical lunar orbit, which not only increases mission duration for the lander, but greatly increases the lander's delta v requirements.
So now we're left with a hodgepodge of ill-fitting hardware for the intended mission, with many critical components still years away from becoming operational.
But what if instead of Apollo on Steroids, we had tried to do Apollo on a Diet?
Mass to Shave
Apollo was, in many ways, a minimalist design, especially for its time. The command module, which was the only habitation space for the astronauts on the 3 day journey back, was quite cramped. The whole lunar orbit rendezvous scheme (in contrast to taking the reentry capsule all the way to the surface and back) was a hard won mass savings battle that prevented the development of an even larger rocket than the Saturn V that would have used eight rather than five F-1 engines. Mass was pared down at nearly every turn.
So it may surprise space fans to find out that significant mass savings are possible over Apollo.
Let's start with crew size. Apollo only landed crews of 2 on the Moon, but had a total crew of 3, with one person left behind in the orbiting command module. This was because lunar orbit rendezvous was considered risky, and it was important to have someone in orbit to help maneuver the command and service module and open the hatch. With modern electronics and experience, it's no longer such a scary thought to leave that function to automated mechanisms and/or remote control. We could have the whole crew go down to the surface (which, as far as I know, is what's already intended for Artemis IV and beyond), so we could pare the total crew size down to 2.
Next we have the vehicle design. Before NASA settled on the Apollo CSM, they asked for industry proposals. One of these proposals was a design that split the spacecraft into three modules (mission, reentry, and service) very similar to the Soyuz's spacecraft's orbital/reentry/service module split, as opposed to Apollo's command and service module split. Such a design pulls as much of the space and functionality as possible out of the component that has to reenter Earth's atmosphere and land and would have saved over a ton in dry mass in the case of Apollo, and approximately 2.5 times that when taking into account propellant and tankage. Furthermore, it would have allowed the orbital/mission module to double as a lander cabin, saving an additional >1t in dry mass and ~2.5x that in total.
Then we have modern electronics. The guidance and navigations systems of Apollo used very heavy, bulky, and power-hungry electronics by modern standards. These added substantial dry mass to every component. And in the absence of modern photovoltaics, Apollo used relatively heavy fuel cells and batteries to supply power.
The command module alone had 1.5t of navigation, communication, electrical, and telemetry equipment. And yes, the structure, heat shield, and recovery equipment (which added up to 2.6t) had to be sized in part to accommodate all that equipment.
The service module added another 1.2t in electrical equipment. And of course we again have the ~2.5x multiple for both service and command module equipment to take into account propellant and tankage.
Components on the lander were lighter, but there were almost 250kg of batteries alone on the descent stage and over 100kg of batteries on the ascent stage, and we're looking at much higher propellant and tankage multiples for those stages.
Yet another massive reduction in mass is possible due to the secondary effects of modern electronics. The Apollo lunar module had massive spare delta v capacity to enable it to perform its mission safely in the context of imprecise and unreliable navigation and control. The descent stage had 2500m/s of delta v and the ascent stage had 2250m/s of delta v capability in their primary propulsion systems, while 2000m/s is sufficient each way to get in and out of lunar orbit, accounting for gravity losses, circularization, and a small amount of reserve capacity. Such overabundance can also be seen in the reaction control system (used for attitude control -- typically small puffs of thrust) of the ascent stage, which had a whopping 287kg of propellant for attitude control.
I recently tweeted about a potential architecture that would have a budget of 2.5t for the reentry module, 1.5t for the mission/orbital/lunar cabin module, 6.5t for the lunar ascent & descent stages, and 8.5t for the service/transfer module bringing everything into low lunar orbit and taking all but the discarded lunar ascent and descent stages back out of lunar orbit and sending them on their way to Earth. That adds up to a total TLI (translunar injection) mass of 19t.
Let's see if this is achievable.
I baselined with all hypergolic pressure-fed engines, but with 319 second isp matching a more modern variant of the AJ-10 (the AJ10-118K from 1989) as opposed to the 311-314s isp figures of the three main engines used on the Apollo CSM and LM, one of which was an earlier AJ-10 variant
Assuming identical delta v* capability on the ascent stage and the descent stage and identical structural coefficients**, we can break up the 6.5t for descent and ascent into a 1.96t ascent stage and a 4.54t descent stage.
*The ascent stage has to handle rendezvous and might need to do some RAAN correction, but the descent stage will tend to have slightly higher gravity losses, so likely not far off.
That may not seem like a lot, until we realize that an aluminum ladder of the length used on the Apollo LM will tend to weigh around 10kg, and that with modern avionics and with multiple smaller engines rather than one big one, the landing legs need not be very long, sturdy, or foldable. The total landed mass will be less than 3.5t, and on the Moon will weigh about as much as 570kg would on Earth.
So far so good. That leaves the reentry module and the orbital module.
We had a 1.5t loaded mass for the orbital module. Budgeting 300kg for 2 astronauts and their EVA suits, we have a 1.2t mass for the module itself. This isn't too far off from the 1.3t mass of the Soyuz orbital module, which has 10 days' worth of oxygen supply and CO2 removal capability for 3 crew members, has two docking ports (as opposed to the one docking port and one hatch that we need), communication and navigation equipment, more space than we strictly need, and its own RCS. It's also designed to survive for months in the MMOD (micrometeoroid and orbital debris) and thermal environment of LEO.
We do need independent power generation, but a ring of conformal photovoltaics around the mid-section of the module would be adequate. Depending on where we land and when, we won't need active cooling on the surface, and we don't need it in transit (unlike in LEO) -- this dramatically cuts our power requirements and removes the need for radiators. A few hundred Watts for the comms, nav, control, and lighting would be adequate, as it was for the Apollo LM. This would require just a few square meters of conformal PV and could be had for around 10kg with lightweight paneling, or conservatively for a few dozen kg.
We could, in theory, go a lot lighter than the Soyuz OM.
A 2.2m diameter spherical shell made of double-walled aluminum with 1mm thickness for each wall would mass ~100kg with ribbing and the inner wall would have a burst pressure of over 5 atmospheres (plenty) with a 300MPa material strength and provide excellent thermal insulation in vacuum. We don't have to worry much about MMOD risk because of mission duration and the fact that most of it is spent far from LEO.
20 person-days of consumables without water reclamation would still be under 200kg, and we don't have to bring most of it to the surface.
As for the reentry capsule, the Soyuz reentry capsule is 2.95t, seats 3, and has retrorockets and a beefy structure for ground landings. We might need a more capable thermal protection system, but we know that Dragon's 8cm thick PICA-X main shield is adequate for reentry from TEI and PICA weighs on the order of 20-25kg per square meter. And Apollo's heat shield and recovery hardware (parachutes and flotation collar) added up to around 1/5 of the capsule's mass, so we could indeed go a lot lighter than even 2.5t.
What about radiation? Much has been made of the radiation mitigation work that has gone into Orion to make Artemis safer than Apollo, but background radiation on a short 7-10 day transit and landing mission is not enough to cause a serious health risk. A bigger risk is from the possibility of an intense solar storm. Artemis's primary strategy for dealing with this is building a storm shelter by putting stuff already in the capsule for other purposes (like seats) inside of bags and hiding behind them. This has near zero mass effect.
But this 2-person variant of Apollo Redux is also just within the capability of the Falcon Heavy in fully-expended mode. Notably, the Falcon Heavy has been operational since 2018, can handle much higher cadence than SLS, and is about an order of magnitude cheaper per launch.
With a few tweaks, including maybe using a cryogenic pump-fed descent stage like the Altair, it can be brought within the TLI capability of FH with side booster recovery, enabling potentially dramatically higher launch cadence.
It could also have been easily brought within the LEO launch capability of an Atlas V, and could have been launched to TLI using a hydrolox Earth departure stage launched by a Delta IV Heavy, or alternately, going forward, those two launches could be handled by two Vulcans.
Similarly, with one way cargo landers sized for a single FH2R launch, we could have been landing upwards of 5 tons per launch on the lunar surface, and building a small base.
Alas...
But there's even greater hope for the future of human spaceflight now in the form of Starship. Let us not waste that hope on politically driven system architectures.
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