sevenperforce

Just occurred to me that the Draco develops 300 s in vacuum. Clustering Dracos wouldn't be enough for a lander, but would they be enough to perform an LOI burn at the gateway? Because that would both save mass and make everything ridiculously simple.

Boiloff wouldn't be an issue for a solar-electrolysis prop depot; you'd already need a massive array of radiators to liquefy everything in the first place.

There's an idea. What's the value per gallon of water in orbit? And hey, that's something that would be enabling even without LOP-G. Go grab about a billion in venture capital, buy a reusable electric-pump-cycle hydrolox engine with accompanying thrusters, and build a solar-electrolysis propellant depot to be lofted into LEO by Falcon 9, then buy Falcon Heavy launches to send as much water as you can to the station, as frequently as it will launch. Use your hydrolox engine and thrusters to build a tug with a metallic heat shield for aerobraking, then offer direct-to-GEO services from basically any orbit. Masten has done a lot with small methalox thrusters that would be usable for a lander engine under the right circumstances.

The usable volume in a Falcon 9 fairing over 216 cubic meters. If you poured water into the fairing's usable volume, you'd only be 1/3 of the way full before you'd hit Falcon Heavy expendable's 63.8 tonne limit. For reusable configurations you'd use even less of the volume before you'd be mass limited. Now, if you're trying to loft hydrolox in bulk, then yes, you're going to run into problems filling up Falcon Heavy expendable's mass budget. A fairing-sized hydrolox tank would mass just about exactly 62 tonnes. But, again, we aren't flying Falcon Heavy expendable, and we aren't lifting pure hydrolox. Common bulk densities for reference: Water: 1 g/mL Hydrolox: 0.32 g/mL Methalox: 0.83 g/mL Kerolox: 1.03 g/mL Hypergolics: 1.2 g/mL

Looking at the Delta IV user guide, ULA offers about a dozen different Delta IV PAFs with suggested maximum load limits ranging between 5.0 for the flimsiest ones and 9.1 tonnes for the heaviest ones. The Atlas V has eight different PAFs with maximum loads ranging from 6.3 tonnes to 9 tonnes. The New Glenn user guide released two months ago and cited explicitly that they could do at least 20 tonnes in a dual-manifested launch, which makes that the lower bound of structural limitations on the upper stage (though I do not know if their PAF can handle 20 tonnes in a single-payload mission). The 45-tonne reference orbit was, naturally, inclusive of residuals. In terms of pure upper-stage structural limits, we know that Delta IV Heavy took the 21-tonne Orion EFT to LEO. So its stage structural limits are more than 230% of its biggest PAF.

Refueling isn't a magic fix for issues like boiloff and ullage. Starship gets around it by having such a massive square-cube advantage that boiloff and generous RCS use are inconsequential, but something like Centaur can't manage nearly so easily. IIRC, Centaur has to actively vent hydrogen boil-off throughout its entire coast just to maintain ullage and prevent the tanks from popping. The Falcon 9 user's guide suggests the current payload adapter is limited to 11 tonnes for center of mass reasons, but says that higher-mass payloads can be accommodated as a special service. I believe the problem is not so much with structural loads on the upper stage (and certainly not for the lower stage) as it is the bending moment on the payload adapter during pitchover under power. If the payload's center of mass is too high above the payload adapter then it will not be able to stay secure. The largest payloads launched in a fairing to date have been the Iridium NEXT clusters at over nine tonnes. But all you need to remedy that is to design a custom PAF, which isn't so bad. Structural limitations on the vehicle itself are probably much higher.

How far along are the XEUS plans for handling boiloff? That's always a challenge, but particularly with hydrolox. And a dual-thrust-axis lander makes the addition of a crew capsule rather challenging, particularly if the crew capsule is sized for independent abort. That's why I suggested putting a refuel module up on a separate LEO launch in order to rendezvous with the lander. Its propellants would be used both to refuel the ascent stage and to act as a drop tank for the descent stage, such that the descent stage can fill the role of tug. Plus, it will be light, meaning that there will be plenty of residuals in the launch vehicle upper stage to push both it and the lander onto TLI. Does anyone know the upper limit of discrete payloads for LEO by most commercial vehicles? Because if we know that number then we can start sizing accordingly.

Logistically speaking... There is value in having hypergolics on the ascent stage, so let's assume an ascent stage powered by either a single SuperDraco or clusters of smaller hypergols. It will need to dock to the lander stage with capacity for propellant transfer. The smaller the better, for pretty much every reason. This also means the lander will need to carry hypergols for refueling the ascent vehicle, though not necessarily for itself. For your mass ratio, you want to shed spent mass whenever possible, and for complexity/reliability you want to drop stages as infrequently as possible. The ideal scenario is to have your ascent module refuel tanks bolted to the same jettisonable module as your lander's lunar orbit insertion tanks (since it's braking itself into cislunar space) so you only have one separation event. For dV reasons, let's assume cryos for the lander. Scenario: Crew and unfueled ascent module are already at LOP-G. Commercial launch 1 places lander module in LEO with just enough propellant for lunar deorbit and landing. Commercial launch 2 takes refueling module into LEO but does not separate. Lander executes rendezvous and docking with refueling module, then Commercial launch 2 restarts its engine for TLI burn. Lander completes TLI burn (if necessary), coasts, and then brakes into cislunar orbit at LOP-G, drawing from cryo tanks in the refueling module. Crew enter ascent module and dock to lander. Ascent module is refueled from refueling module. Refueling module is jettisoned either at the LOP-G or once in LLO. Descent stage lands and the surface mission is conducted. Ascent module returns to LOP-G to repeat the cycle. This expends only one BLEO engine assembly and achieves a full lunar landing in only two commercial launches rather than three. Because the lander is sent to LEO, it can be dramatically larger than otherwise possible. With Atlas V 551 it could go up to 19 tonnes; with Delta IV Heavy, 28 tonnes; with Falcon Heavy reusable it could go up to 30 tonnes or more. New Glenn, Vulcan, or Delta IV Heavy would be ideal for launching the refueling module because the residual hydrolox will give better TLI performance.

Merlin Vacuum without nozzle extension: Merlin Vacuum Nozzle Extension by itself: Combined: It's the same deal with the Kestrel. For the Kestrel, here you can see both the ablatively-cooled chamber and throat attached to the radiatively-cooled nozzle:

It's a Kestrel without the nozzle extension attached.

They need to either dust off the Kestrel and convert to methane-GOX or do that with the SuperDraco.

I don't think I've ever seen a that picture of the Kestrel before.

Spent entirely too long digging into the nine companies tapped by NASA for CLPS to expand the roundup of currently-available engines. Here's what I came up with. BE-3U (Blue Origin), 530 kN. Hydrolox. Broadsword (Masten Space Systems), 160 kN. Methalox. Reaver 1 (Firefly Aerospace), 182.2 kN. Kerolox. RL-10C (Aerojet Rocketdyne), 110 kN. Hydrolox. SuperDraco (SpaceX), 90 kN. Hypergolic. Lightning 1 (Firefly Aerospace), 70.1 kN. Kerolox. AJ-10 (Aerojet Rocketdyne): 27 kN. Hypergolic. HD5 (Intuitive Machines), 24 kN. Methalox. Rutherford (Rocket Lab), 24 kN. Kerolox. Katana (Masten Space Systems), 18 kN. IPA+LOX. PECO (Moon Express), 5.64 kN. RP1+HTP. Machete (Masten Space Systems), 4.4 kN. IPA+HTP. FME (Team Indus for ORBITBeyond), 0.44 kN. Hypergolic. Deep Space Engine (Frontier Aerospace for Astrobotic), 0.44 kN. Hypergolic. Draco (SpaceX), 0.4 kN. Hypergolic. A slightly broader spread, but still not a lot to choose from. It is a shame that SpaceX's gas-gas methane-oxygen thruster has been put on the back burner. Imagine a SuperDraco with a regenerative cooling loop that pre-vaporizes both LOX and CH4, feeds a small amount back to keep the tanks pressed, and then feeds the rest into the chamber with a simple spark igniter. Gas-gas engines are simple, reliable, and while it wouldn't have the best TWR in history, it would have superb efficiency and relatively decent propellant bulk density. NASA's plan currently requires that all vehicle components be limited to what can be sent to TLI by commercial launch providers AND fit inside a five-meter fairing. It's really rather restrictive. Thinking back to Constellation...I would love to see a single-stage transfer, refuel, and descent/surface module which is assembled in LEO via two commercial launches prior to TLI. As it is, the NASA approach requires no less than three commercial launches, plus crew via SLS, for every moon landing, and that's without any supplies being landed on the moon. At least then you would only have a single disposable powerplant braking into cislunar space for each mission instead of literally three.

A human-sized lunar lander needs to start with an appropriately-sized engine...so, what's out there? I thought I would do a quick roundup of engines that could conceivably be used or adapted for use in cislunar space.... BE-3U: 530 kN RL-10: 110 kN SuperDraco: 68-90 kN Rutherford Vac: 24 kN* AJ-10: 27 kN Not a big list, and not much to work with. Rutherford is included because although kerolox is problematic for ignition reasons, the cycle is fantastic for cislunar ops. How many did I miss?

We know the Falcon numbers to a high degree of accuracy. We'd need to estimate for NG based on published imagery and propellant bulk density. Bezos is tight-lipped. To the OP: there is an optimization function for any proposed mission profile or configuration set, but there are a lot of variables. A vehicle intended primarily for launching BLEO will ten to be underpowered for large LEO payloads, and vice versa. Plus, if you use different propellants on the first and second stages, you get an entirely new set of variables. Recovery also complicates.