Crewed Lunar on Falcon 9
On the Prospect of Falcon 9 Doing SLS's Job...
Recently I pointed out that Falcon 9 has flown more lunar missions than SLS, and that this is likely to remain true through the end of the decade, based on contracted and expected launches.
Some rightfully pushed back on this being an apples and oranges comparison, since the Falcon 9 doesn't have nearly the capabilities of SLS, and the spacecraft it is sending towards the Moon are relatively small uncrewed ones.
Now it hadn't been my intention to say they were the same or that the Falcon 9 could do SLS's job. I was just making a tongue-in-cheek jab at "Mega Moon Rocket" launching fewer lunar payloads for the foreseeable future than not-a-Moon-rocket, as well as reminding people that two-stage kerolox rockets like Falcon 9 can in fact competitively launch useful payloads towards the Moon, contrary to popular wisdom until not that long ago.
On the other hand, I do like to remind people that Artemis-like missions can be accomplished without a super heavy launch vehicle like SLS, instead relying on Earth orbit assembly using medium/heavy launch vehicles like Falcon 9. In fact, ULA proposed doing this with their Atlas V and Delta IV Heavy launch vehicles before Falcon 9 first flew.
But this SLS/F9 comparison got me thinking... would it be possible to design a single-launch architecture using Falcon 9 that could safely and comfortably get four astronauts to NRHO (near-rectilinear halo orbit) and back to Earth, as the SLS+Orion stack is designed to do?
This seems absurd on the face of it, considering that per SpaceX, Falcon 9, even with booster expended, maxes out at 8.3 tons to GTO (geostationary transfer orbit) and 4 tons to TMI (trans-Mars injection), suggests around 5.5 tons to TLI (trans-lunar injection), and Orion tips the scales at a whopping 26.5 tons, not including its launch abort system or fairing. Even Crew Dragon, which doesn't have the capability to go in and out of lunar orbit on its own, is around 12.5 tons. (I've written previously about how Crew Dragon could be modified for this kind of mission and how Falcon Heavy could launch it.)
But lighter crewed vehicles, like Soyuz, have been flown and even considered for the Moon, and Falcon 9 leaves a lot of lunar performance on the table by not using an Earth departure stage. So let's see what we can do, shall we?
I'll start by trying to maximize the lunar mass capabilities of Falcon 9 without modifying it below the fairing -- just by adding a sort of Earth departure stage on the existing stack. Then I'll work on outlining a crewed vehicle that can fit within the increased mass budget, sustain a 4-person crew for the duration of the trip to and from NRHO, and safely reenter and splashdown, coming in hot with ~11km/s of lunar return speed.
How much can we launch?
SpaceX rates Falcon 9 at 22.8 tons to LEO, with booster expended. This figure is from SpaceX's website, which hasn't been updated in a few years. Some have proposed that Falcon 9 performance has continued to be improved in recent years, and that 22.8 tons is pessimistic. There's indeed evidence for F9 performance improvements, but this has been seen in launches with booster reuse, and it's not clear how much of the performance improvement is due to more efficient reuse (e.g. higher reentry speeds requiring less propellant for booster recovery), so I'll go with the 22.8t figure for F9E (expendable) while noting there might be some extra margin here.
For F9R (reusable), we've seen 16.7 tons to a 53.2 degree orbit at >300km staging altitude (>300km is typical for Starlink missions, and by my math, after SECO at 235km, the vehicle had enough excess velocity to head to a ~325km apogee, likely followed by a circularization burn). This level of performance would enable 18 tons to a 200km orbit at 28.6 degrees (~200m/s less delta v).
For a crewed launch, I'll reduce the max payload figures by 5% for a crew-safe trajectory, g-limits, and margin. Some might consider this an optimistic margin, but I find it defensible and possibly even conservative, given that for a rocket with relatively high second stage thrust like Falcon 9, g-limits and crew-safe lofting don't need to impact performance very much.
And we don't really need performance margins. Since we do have additional stages in play to handle the TLI burn and following maneuvers, vehicle underperformance doesn't really pose a LOC (loss of crew) risk -- just a LOM (loss of mission) risk, and Falcon 9 is enough of a proven vehicle to not worry too much about that.
This leaves us with ~17.1t for F9R and ~21.7t for F9E to LEO.
Fairings and Abort Systems
One other piece to account for is the fairing. It's hard to find a mass figure for Falcon 9's standard fairing, but estimates fall in the 1 to 2 ton range, with most upwards of 1.5 tons.
As is typical for crewed spacecraft, rather than using a large encapsulating spacecraft, we would be flying with much smaller and lighter custom panels covering certain exposed parts of the vehicle.
So we can save a lot on fairing mass, but fairings get ditched early and only make a difference of a small fraction of a kg to orbit for every kg of fairing mass.
But rather than calculate how much of a performance improvement ditching the fairing would yield for Falcon 9, I'll use that mass budget elsewhere.
While like Dragon and Starliner use integrated launch abort systems that stay with the vehicle throughout the mission, I'd opt for a system that is jettisoned early in flight (around the same time as a fairing would be, as with traditional abort towers), in order to save mass.
Given how light our reentry capsule is going to have to be, the 1-2 tons saved in fairing mass ends up being more than enough for aero panels and an abort system that's jettisoned around the same time the fairing would be.
Note that for Orion and Apollo, the LAS is nearly three quarters of the mass of the capsule. But the push systems on Dragon and Starliner are much lighter, and this isn't just a product of using liquid propellant. In fact, high pressure liquid abort systems can have some structural mass disadvantages, and thanks to their low expansion ratio, the SuperDracos on Dragon don't offer very high isp -- just 235s at sea level -- and that's before cosine losses due to off-axis thrust.
Dragon's launch escape system, based on vehicle mass and propellant used for abort tests, provides roughly 300m/s of delta v, which can be matched with with a solid pusher abort system with the same 235s isp, at less than a third of capsule mass even with a very hefty 50% structural coefficient.
You should be able to get to a much lower structural coefficient than 50% with solid propellants, and you ought to be able to stay under 50% with pressure-fed hypergolics, as well. I mainly prefer solids here because they're likely to be safer, and they're substantially simpler.
I'll be aiming to keep the capsule mass at under 3t, so the escape system will comfortably fit in our fairing mass budget. It's also worth noting that the reduced vehicle diameter will tend to reduce drag losses.
Now, you might ask: how do you jettison a pusher system mid-launch? We'll get to that!
Earth Departure Stage and Beyond
Baselining LEO+3.2km/s for TLI (it can be done with as low as 3.05km/s), a hydrolox Earth departure stage with 450s isp and ~10% structural coefficient (inspired by Centaur III -- could be improved upon that) and 1% residuals (again, lower is possible), an Earth departure stage in very low Earth orbit could send 7.2t of payload to the Moon using F9R, 9.1t using F9E.
NASA baselines 900 m/s for NRHO insertion and exit. Adding another 150 m/s for course corrections and margin, I'll budget 1050m/s for the post TLI maneuvers and start by leaving this to dedicated spacecraft propulsion.
Traditionally, this propulsion system would be pressure-fed hypergolic. With a 15% structural coefficient and 1% propellant residuals at a modest 320s isp, we'd get ~4.8t of payload there and back for F9R, and 6t for F9E.
Switching to a pump-fed methalox system with 10% structural coefficient and 375s isp (less performant than could be in consideration of the small size) bumps that up to 5.2t and 6.6t respectively.
You can also ditch separate spacecraft main propulsion and use the EDS all the way through the trans-Earth injection burn (Centaur V is aiming for many months in space), which gets you virtually identical performance to the methalox 4th stage (higher isp being canceled out by higher terminal structural mass).
But what if instead of carrying all your tankage from LEO, you employ drop tanks, for example in a 7-tank configuration (6 around 1)*, dropping 2 at a time until you get to the last? And what if you use lighter tanks? And what if you stage lower, say around 1km/s before LEO, where you're close enough to orbit to not lose much performance with a relatively low thrust-to-weight ratio, but you can increase the contribution of your higher isp hydrolox stage?
*Yes, there are reasonable ways to make this fit within a 5.2m diameter (fairing diameter), if not a 3.7m diameter (F9 body diameter). There are also ways to achieve similar or better performance with a more conventional two-stage hydrolox setup that would be geometrically easier to deal with.
If you do that and employ a conservative 1t payload adapter to enable suspension balloon tanks w/ a conservative 4% structural coefficient and budget a fixed 500kg for other propulsion elements (the 451s isp RL10 on Centaur III weights under 170kg and has over 2x the thrust needed for this application), still using 1% for propellant residuals, you get 6.2t and 7.9t, respectively for F9R and F9E.
If you go to 460s isp, 3% structural coefficient tanks, 0.5% residuals, a 0.7t payload adapter, and stage 1.5km/s before LEO, even adding a ~50m/s delta v penalty for low twr, you get a whopping 7.9t and 10t.
So there we have it, depending on how ambitious we are with the EDS and whether or not we expend the F9 booster, we can get anywhere between a 4.8t and a 10t mass budget for our spacecraft, and the mass of the launch escape system is already accounted for separately.
But frankly, that would be too easy, and you might accuse me of stretching F9 performance beyond credibility (or just beyond what some might consider F9).
Instead I'll go with the lowest 4.8t figure for F9E + unoptimized hydrolox stage just for TLI + a pressure-fed hypergolic spacecraft propulsion system.
I'll choose this modular approach, but add a twist. Instead of putting the reentry module in the middle of the stack, heat shield down, I'm going to put it at the top, heat shield up (Kistler style).
The biggest disadvantage is that now you have a different load direction during launch than you do during reentry.
Structurally, this is pretty trivial -- the reentry vehicle needs to be build for very high compressive loads coming from the base (shield), both for ballistic reentry contingencies, and for rough landings. 20g is a reasonable design target. Building it for a fraction of that loading in the opposite direction (for launch) won't add much mass, especially when the alternative is for the reentry module to carry thrust to the orbital module plus a launch tower, which would be just as high, if not higher, of a load in that direction.
The issue is seat orientation. Lying back on the way down means "eyeballs out" on the way up. Not good.
But there's a pretty straightforward solution: swivel seats. Bearing, hinges, motors can add mass and complexity and require space. But it really won't come out to that much.
Now, is this all worth it? Maybe, maybe not. You can absolutely design this as a more traditional stack, and it might come out just as light, or lighter, but I wanted to explore the upside down capsule route, and I'm running with it.
I'm going to go with 1.6m x 0.6m x 0.6m seats in a 3+1 configuration (three near the base, one closer to the mission module, with room around the individual seat to get to the other three), with a mostly-reclined with bent knees position. These seats would be able to individually rotate around their long axes so that astronauts will experience g loads in most natural direction both on the way up and the way down.
This allows a 2.5m interior diameter at the base of the pressurized volume and fairly high sidewall angles. Even with a 0.9m seat pitch, we could have a ~25 degree internal wall angle from base to mid-seat, and potentially an even higher external sidewall angle.
It would be feasible to keep the external diameter under 3.2m, and the height around 2.6m, yielding an external surface area of <25m^2, about 1/3 of which is at the base.
If you approximate the inner pressure vessel as a 2.5m diameter sphere with a safety factor of 50 (that's right, 50, so you can excuse the spherical approximation) at 1 bar of pressure, and you use multi-ply carbon fiber with a respectable but not boundary-pushing 1.5GPa tensile strength, you need an average thickness of ~2mm, and at 2t/m^3 of density, you get a pressure vessel structural mass of ~80kg.
If you approximate the primary structure bearing compressive load as a 2.6m carbon fiber tube bearing 3tons at 20g along the entire length with a compressive strength of 500MPa and a safety factor of 20, you get ~120kg for that.
The primary structure would be an integrated, but for convenience and a conservative estimate, I'll approximate it as the sum of the above pressure vessel and compressive structure, and add an extra 50kg for good measure, bringing us to a total mass of 250kg.
Next up is the thermal protection system. Conservatively, let's say we use 8cm of PICA-X on the main shield. This was considered adequate for Dragon 1 reentry from lunar or Mars return speeds (a small fraction ablated for LEO reentry), and we're aiming for a substantially lower ballistic coefficient here. And let's say we use ~2cm of it on the sides, just to be safe (neither version of Dragon used any, though Dragon 2 at least would need slightly beefed up side TPS for lunar return). At 270kg/m^3, we're looking at around 260kg of PICA-X.
Now let's assume we need an insulating layer under the PICA-X that needs to (again, conservatively), handle a 500K average temperature drop (PICA starts decomposing well below 800K, so I think it's safe to assume the cool side of the PICA heat shield will be below 800K). Silica aerogels can operate past the decomposition temperature of PICA, so I think we're safe using that material (and if not, carbon aerogels are quite promising). A relatively high density silica aerogel at 120kg/m^3 w/ 13.5mW/m/K observed thermal conductivity would have adequate compressive strength and low brittleness. Let's say we apply it at a uniform 5cm thickness. This comes out to ~140kg.
And let's say our average thermal conductivity is a lot higher at, 30mW/m/K, due to elevated temperatures and/or other factors, resulting in ~10kW of heating at a 500K temperature drop across the insulating layer. And suppose we approximate reentry heat soak as being 20 minutes at these temps (it's less -- you'll do most of your slowing down in a fraction of that, and then the atmosphere will start cooling the exterior). We're looking at ~12MJ of heat to dissipate (or 13-15MJ including internal heat sources).
This can be dumped with approximately 30kg of propane, which can be stored as a liquid at room temperature under pressure, evaporated to provide cooling at ~400MJ/kg, and which would provide enough delta v to double as the reentry RCS system using cold gas thrusters. We'd get upwards of 5m/s of delta v from such thrusters, which would be enough to slowly adjust reentry roll to manage rate of descent**.
With ~50% structural coefficient for an internal heat sink, canisters, valves, and vents/thrusters, we'd have ~60kg for the terminal cooling system + RCS. Let's make that 200kg to budget for launch, final approach, and recovery, when astronauts are strapped into the reentry module with hatches closed, so you'd have an order of magnitude lower heating rate, but for a longer duration.
Budgeting 50kg apiece for some nice sturdy seats with swivel mechanisms, we're looking at 200kg for the seats. (Given the immense loads and safety factors the 250kg primary structure was estimated for, I hope this doesn't give anyone any qualms.)
Let's say the astros are on average 80kg each, have 60kg of suits, life support systems & consumables (in the short duration reentry module), and a 10kg personal allowance. That's another 600kg.
Let's throw in another 150kg for parachutes. Note that for Apollo, parachutes accounted for less than 5% of vehicle mass, with each of three parachute assemblies weighing around 60kg. And unlike some other designs landing in water, we don't need to worry about flotation devices to keep the vehicle upright after splashdown, since we have swivel seats...
That comes out to 2.3t, leaving 2.5t for the mission module.
One more side note: the mission module, the EDS, and the LAS are all single-use, but the reentry module should lend itself pretty well to reuse. The safety factors on the primary structure are very high, we have no solids or hypergolics with burst valves, and almost the entire structure being wrapped in single-use PICA-X would help with preventing saltwater intrusion.
Mission Module
Let's say we want 25 days of life support for four astronauts (more than enough). At a generous 10kg/astro/day***, that's 1t of consumables.
Let's make this module a 3m x 3m cylinder. This would have an internal volume of ~20m^3 (more than adequate for 4 astronauts), and an external surface area of ~42m^2.
4mm of aluminum arranged in a double-wall (1mm outer bumper + 3mm inner wall) would provide insulation and good strength, while also functioning pretty effectively as a Whipple shield (though there's less need of MMOD shielding outside of LEO). This primary structure would weigh ~460kg. An extra ~40kg would buy ~1kg/m^2 of padding (equivalent to 3 to 9cm of expanded polystyrene) -- which isn't necessary for thermal management, but would be useful for ensuring the walls are comfortable to touch.
And let's add 150kg for two hatches, as with the reentry module. One of these would be a docking port on the side of the mission module, since the mission module sits between the reentry module and propulsion elements.
Now let's think about thermal management.
In space, far from the IR sources of the Earth and the Moon (including in NRHO), passive cooling alone would suffice. A high emissivity coating could radiate ~300W /m^2 from the outer wall at lower than internal temperatures. You just need to move heat from the inside of the craft to the outer wall.
For near Earth operation (which would be brief), we could keep power use minimal and use the propane cooling method inside the reentry module.
Let's add another 200kg for a hypergolic RCS system. Even w/ 50% structural coefficient, 100kg of propellant at 300s isp will provide a 4.8t spacecraft 62m/s of delta v for docking/undocking and course correction maneuvers.
And another 300kg for everything else (toilet, sleeping bags, first aid kits, fans, etc., etc.), bringing us to a total of 2.5t, and the whole vehicle to 4.8t.
Primary propulsion is accounted for separately (see earlier in the launch vehicle discussion), and we don't need a service module for anything else -- thermal management, ECLSS, power, comms, nav, and RCS are accounted for in the mission module.
Yes, this is very much doable, with plenty of room all over our mass budget for potential mass growth, and using an F9 with booster reuse and a not very ambitious Earth Departure stage and hypergolic spacecraft propulsion.
Frankly this is all kind of wild to think about. But not so crazy when you consider Zond and Lunar Gemini, and what was within reach for those programs with 1960s tech and shoehorning early LEO vehicles.
And even more exciting questions tantalize, like what could you do with 10 tons?
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