sevenperforce

I don’t believe so, certainly not with Skylab and pretty sure on the ISS. Wasn’t much need, since they all had crew tunnels, and seems very high risk. Just looked it up -- there were plenty of spacewalks to assemble ISS, but the first that would have allowed an EVA between vehicles was STS-104 when the Quest airlock was actually installed. Although there were three spacewalks during this mission, two from the Atlantis airlock and one from the Quest airlock, nobody ever left one airlock and entered through another. Maybe it has never happened! Apollo 15 required an EVA to retrieve a recording device from the Apollo service module, but this was after the lunar module had been detached. For Skylab 2 (the first crewed mission to Skylab) the crew had to EVA, climb onto Skylab, and manually disassemble part of the Skylab docking ring in order to get the ring to function for hard capture, so this was at least a transfer between vehicles, even if they didn't enter Skylab via EVA. But yeah, more interesting to see instances where EVA transfer was planned and integral, like in the N1 lunar lander. No, I'm talking about the 1L/Vostok-7 spacecraft (not to be confused with the cancelled Vostok-7 mission, the Vostok-7 I refer to is of the same category of designations as the Vostok-3KA, the official name of the original Vostok). Wow, cool! I love how I'm continually learning new stuff here. Very Kerbal. But that begs the question, why not just launch in the 1L? Not the first time the Soviets launched with no abort ability… They may have had no concerns about launch abort, but the mission profile called for sequential crew-guided assembly of the propulsion modules with the circumlunar crew vehicle attached last: The IL vehicle was added last, so a separate crew vehicle was required.

Wait Vostok-derived lunar flyby?? Probably talking about the early Soyuz-derived lunar flyby using the Soyuz-A-B-V design, where two crewed Soyuz capsules would go up to meet a Soyuz-based tug that would have already been refueled by several subsequent R-7 missions. The two capsules would have had some EVA transfers while assembling the whole circumlunar stack. I'm guessing @SunlitZelkova means EVA transfer for a lunar mission architecture. Weren't there at least some EVA transfers during Skylab and during the construction of the ISS? There was an EVA transfer proposed (in hindsight) as a rescue mission for Columbia, but obviously that wasn't operational.

I think there's likely a top port for fuel transfer, then the side port for crew transfer See, the trick is that they don their spacesuits and swim through the liquid hydrogen. They needed the spacesuits anyway! The best part is no part! Brilliant! If they had a four-engine cluster at the center instead of a five-engine cluster, they could do "sweep" the surface on touchdown by vectoring two engines inward and two engines outward, then reversing. I suppose they could do the same thing with the five-engine cluster since the central engine is going to be doing the work of dusting off the center anyway. Still, it does make a true surface-level egress impossible because you need vertical space for the engines (just less vertical space than a single engine of comparable thrust).

I suppose cranes are easier to manage on the moon thanks to the low gravity. But still, geez. I'm guessing there are some fairly straightforward maths that relate the amount of nozzle exit area under a lander to its carrying capacity....

Any indication of how the cargo offload would actually work? Where does the cargo go, anyway? For a one-way hydrolox moon vehicle, I think it makes a lot of sense to have the fluffy hydrogen tank on top, the LOX tanks on the sides, and 4-6 smaller engines mounted on or around the LOX tanks to straddle a large central cargo bay at ground level. Something like this: But of course NASA is in love with a single central thrust path, which I suppose makes sense for structural reasons as well as heritage.

Good examples, but arguably those are things which would be one-way launches, which obviates the "up and down" element needed for the whole manufacturing economy of scale to operate. One possibility would be higher-capacity microchips, or even quantum computing chips. Moore's Law is reaching its limits and will soon be obsolete (some are saying it already is). Once we have lower-cost access to space, I'm sure that some of these tech companies will be itching to experiment with ways that microgravity can increase the density or function of integrated circuit chips. With crystal growth in microgravity, perhaps we could see more true monolithic 3DIC chips, where transistors are "grown" on a three-dimensional lattice to achieve exponentially greater processing speed with a smaller footprint and reduced power consumption. That could be one of the enabling technologies that pushes manufacturing offworld, since the demand for higher transistor speeds scales in a way that niche/bespoke things like organ manufacturing do not. Another potential avenue could be some sort of yet-unimagined solar powered fuel production system. We have a lot of technologies that depend on liquid-based fossil fuels and internal combustion engines, and there appear to be limits to battery density. What if we could send up a few hundred tonnes of waste plastic with some sort of solar-powered bacteria or bacteria-fungal symbiote that, once exposed to sunlight and microgravity, could "eat" through the plastic and produce hydrocarbons or some other hydrogen-dense fuel at scale, creating its own lattice structure to continually absorb more and more sunlight?

They are really maximizing the utilization of space by having the big LH2 tank up top. If you recall from the Artemis thread, the volumetric utilization for the old Altair lander design was awful, on the order of 19%. Doing this really big lightweight LH2 tank up top takes maximum advantage of the square-cube law so that you can fit a LOT of liquid hydrogen into a relatively small volume.

Oh, certainly. Obviously none of these things can be done, not at this point. I was just musing about what omelets could have been made if we had been willing to break a few eggs. The comanifested volume is one of the big issues. The Altair lander would have been 10 meters wide and 15 meters high. Of course the descent stage would have been hydrolox which drove up volume, but even with a denser propellant choice I don't see how you can fit a lander underneath Orion on SLS. Revisiting this quickly just for general edification. The Altair lander was a two-stage design with a hypergolic ascent stage using an AJ-10, a cryogenic stage using a pressure-fed deep-throttling RL-10 for lunar orbit insertion and lunar surface descent, and a separate airlock and cargo capability on the descent stage: Total mass launched to TLI would have been 45.9 tonnes, plus the 26.5 tonnes of Orion. The pressure-fed version of the RL10 was built and test-fired as the CECE demonstrator and was capable of throttling down to 8% with a max specific impulse of 445 seconds. To brake the entire stack into low lunar orbit, Altair would have needed to burn 13.5 tonnes of hydrolox. Orion would have detached, leaving the weight of the sortie vehicle at 32.4 tonnes; it would need to burn another 11.3 tonnes of hydrolox to get down to the lunar surface. So in total, Altair would have had a propellant capacity of roughly 24.8 tonnes of hydrolox. Schematically, the Altair lander had a central CECE RL10 with a height of 1.53 meters and a diameter of roughly 1 meter (allowing for gimbal); it also shrouded the ascent vehicle's AJ-10 engine. Thus the hydrolox all would have been in the octagonal outer envelope surrounding these two engines: The RL10 has a mixture ratio of 5.88:1, meaning that those 24.8 tonnes of hydrolox include 3.6 tonnes of liquid hydrogen, occupying 50.71 cubic meters, and 21.2 tonnes of LOX, occupying 18.58 cubic meters. The portion of the vehicle containing the tanks can be approximated as a right toroidal cylinder with an external diameter of 10 meters, an internal diameter of 3.33 meters, and a height on the order of 5.2 meters (based on some rough pixel counting). In total that's a "tank carrying volume" of 363 cubic meters, giving a volumetric efficiency of ~19.1% (not unexpected given how fluffy hydrogen is and how suboptimally the tanks need to be arranged here). So if we somehow had a beefed-up SLS capable of launching an arbitrary payload to TLI, what could we fit in the 10-meter-high, 8.4-meter-wide space under Orion on top of the EUS? Let's start by making this a slimmed-down sortie lander, assuming pre-emplaced surface assets. First let's get a better grip on Altair. NASA's assumptions about an Artemis ascent stage assume 9-12 tonnes, including extra props to get to NRHO, and we know Altair was bloated, so let's take Altair's ascent vehicle (with props) as 9.5 tonnes. This means the dry mass of the descent stage was 11.6 tonnes. Subtract the airlock and we save around 1.5 tonnes, dropping the descent stage to 10.1 tonnes dry mass. Remove the 350-kg engine and the expected 500 kg of unpressurized payload to the surface, and we get ~9.2 tonnes for bare structure, landing legs, and tankage weight. So the structural-and-tankage ratio here is a disappointing 2.7:1. These tanks are roughly the same diameter as the tanks on Centaur (which boasts a propellant fraction of 10:1) and held 20% more props, but these would also have been pressure-fed tanks with five times as many hemispheric bulkheads, so I'd ballpark the tankage weight on Altair at 2.5x that of Centaur or 5.2 tonnes. This suggests that the load-bearing structure, RCS propellant, RCS thrusters, radiators, landing legs, and so forth all come in at 4 tonnes. What would an Altair built for SLS look like, then? Well, the minimal-mass ascent vehicle would come in at 9 tonnes...except that we only need to go to LLO, not all the way to NRHO, so we only need 1,870 m/s instead of 2,600 m/s. At 316 seconds of specific impulse, a 9-tonne ascent vehicle would burn 5.12 tonnes of propellant to develop 2,600 m/s of Δv, whereas we will only need ~3.23 tonnes of props, meaning our ascent vehicle need only have a wet mass of 7.11 tonnes. Let's borrow @Exoscientist's idea and use the old discontinued Standard-sized Cygnus as our pressure vessel OML, making it 5.75 meters high including the engine (3.65 meters without). Let's ditch any extensible solar panels and have it run on batteries and fixed panels after detaching from the descent module. Since we are using hypergols, we can fit our 2.2 tonnes of dinitrogen tetroxide in a pair of 1.2-meter external spherical tanks placed as far forward as possible, and fit our 1.3 tonnes of hydrazine (let's take extra for monoprop RCS) in four 0.9-meter spherical tanks clustered underneath: This gives us a good estimate for the absolute maximum amount of volume (purple) we can allocate to a lander and descent stage. The cylindrical region below the engine bell has a volume of approximately 188 cubic meters, the toroidal region surrounding the engine bell has a volume of approximately 95 cubic meters, and the available volume on either side of the ascent vehicle (leaving the front and back open for both thrust balance and egress) is roughly 21 cubic meters, for a total available structural volume of 304 cubic meters. Of course the descent vehicle needs engines. To conserve space (and perhaps utilize a little of the wasted volume around the top of the EUS tank), let's use four BE-7 engines (two would do it, but we need redundancy). With gimbal allowance, each of those takes up a height of 80" and a diameter of 47.4", but since we can let them hang down slightly around the top of the EUS tank we only end up losing about one cubic meter of volume for each engine, bringing us to ~300 cubic meters of structural volume. Setting volumetric efficiency at 19.1% as before, this gives us 57.3 cubic meters of usable tank volume. Bulk density of hydrolox is on the order of 0.365 g/cc, so that's just under 21 tonnes of propellant. Tankage weight will be a little better than Altair since we aren't using pressure-fed engines -- let's say 3.7 tonnes. If the load-bearing structure of Altair was 4 tonnes for 363 cubic meters of structural space, then we can ballpark the load-structure here at 3.3 tonnes. All four engines together probably come to something like 300 kg. So we have a descent stage with 21 tonnes of propellant and a dry mass of 7.3 tonnes, and an ascent stage with a total wet mass of 7.1 tonnes. So let's throw this to TLI on a magically-beefed-up SLS and see what happens. With this approach, Orion doesn't need quite as much propellant since it will be braked into LLO by the lower stage. Orion ordinarily carries 8.6 tonnes of props, but we can drain about 2.5 tonnes out and still have enough for the return voyage (there really won't be a meaningful decrease in dry mass with this approach) to make it lighter. Total injected mass to the moon is 59.4 tonnes. Ballparking the BE-7 at 449 seconds specific impulse, the lower stage burns through 11.1 tonnes of hydrolox to brake itself and Orion into low lunar orbit. Orion detaches, leaving the lander at a wet mass of 24.3 tonnes with just 9.9 tonnes of hydrolox remaining. Fortunately, that gives it 2.3 km/s of Δv, which is sufficient margin for reaching the lunar surface. We can probably shave a total of 2.5 tonnes of hydrolox off and still make it work, reducing our lunar injection mass to 56.4 tonnes. Cramped, but doable...if it was possible to make an SLS Block 2+ with 10-15 tonnes more capability than SLS Block 2, which is already never going to happen.

Oh, certainly. Obviously none of these things can be done, not at this point. I was just musing about what omelets could have been made if we had been willing to break a few eggs. The comanifested volume is one of the big issues. The Altair lander would have been 10 meters wide and 15 meters high. Of course the descent stage would have been hydrolox which drove up volume, but even with a denser propellant choice I don't see how you can fit a lander underneath Orion on SLS.

Any architecture based on SLS is also going to have the issue of limited volume for a co-manifested lander. 8.4x10 meters of volume is definitely a good bit -- probably enough for a flags-and-footprints lander -- but it's not enough to include a dedicated LOI braking stage for a single-launch architecture, and Orion certainly can't do the braking burn. So even if you were able to beef up SLS (SLS Block 2B?) enough to throw an arbitrary amount of payload to TLI, it's still unclear how to make that work. Then again, maybe I'm wrong and there IS a way to fit a good lander and a braking stage into that volume. It just makes much more sense to have all of the assets delivered to cislunar space ahead of sending crew. And yet the fiddling I did yesterday is making me very curious about other possibilities. I wonder what would happen if SLS used 4-segment boosters and put liquid kerosene tanks on top, and then put RD-180s there on the outer jettisonable engine skirt, fed from the kerosene tanks on the boosters and from the core LOX tank.

90 tonnes to LEO is not enough for single-launch manned lunar missions. To be clear, the Falcon 9 SLS booster concept I describe above is absolutely NOT a reusable architecture. SLS liquid booster reuse is a VERY long pole.

Oh, I doubt the recovery elements would be particularly reusable. That's in line with ULA's overall approach to reuse. With SMART, they aren't going to be reusing the tanks or the heat shield or the parachutes; just the engines. So it would make sense for them to use disposable recovery elements to recover the more expensive hardware. I would really love to see an exploration of new recovery modes and an expansion of available data. One near-future concept for reusable re-entry vehicles is the reversible deployment of large-volume drag structures (deployed either by pressurants or electromagnetically) that slow vehicles rapidly at ultra-high altitude before there is appreciable heating. We need more data on inflatable heat shields to see how that would actually work.

Would not fit in the VAB as he said, so why bother doing the math? Actually it would. The limits for the length of an upper stage for the SLS are based on specs for a supposed second mobile launcher for the SLS which has to be provided to the manufacturer of the mobile launcher. But that second mobile launcher has not been constructed yet, and the manufacturer has been greatly criticized for costs overruns and delays. A commercial approach would probably just choose a different manufacturer for the second mobile launcher such as SpaceX. The doors of the vehicle assembly building are quite huge and can accommodate the size of the SLS core and Ariane 5 as a second stage: Vehicle Assembly Building. There are four entries to the bays located inside the building, which are the four largest doors in the world.[14] Each door is 456 feet (139.0 m) high... With the existing mobile launcher design, there is an 18-meter height available for the upper stage. The current mobile launcher base comes 7.6 meters above the ground and has a launch umbilical tower that is 108 meters higher than that, reaching 5.7 meters higher than the max height of Block 1B. The Ariane 5 core is 23.8 meters high. Putting it into SLS as the second stage would lift the height of this new "Block 1F" (F for Frankenrocket) above the height of the current launch umbilical tower, which is already what limits the height relative to the VAB doors. The only option is to build a crawler-transporter that carries SLS lower to the ground. I'm not sure how we are going to manage that. SLS is heavy. Do the calculation with the Ariane 5 core as the upper stage. A commercial space approach wouldn’t use such a tiny upper stage such as the ICPS for a rocket this size when larger, appropriately-sized upper stages are available As I said, you need to start with a one-to-one comparison to see if ditching boosters and using RS-68s improves anything on the lower stage, before proposing a different upper stage. Otherwise you're changing multiple variables at once, which muddies the water. If changing the lower stage makes things worse, then you should drop that plan entirely, before talking about combining a bad lower stage with a new upper stage. Also, the Ariane 5 core is not "available" at all; it's not even being made. Also also, the Ariane 5 core had a Vulcain 2 on it, which is not a vacuum engine and gets really poor vacuum specific impulse compared to a proper upper stage. But if you're calling the ICPS "tiny" then you should also be calling the Ariane 5 core "tiny" in comparison to the EUS for Block 1B. The EUS carries 76% more propellant than the Ariane 5 core. The Ariane 5 core is extremely narrow compared to EUS. A reasonable space approach wouldn't use such a skinny, narrow stage as the A5 core for an upper stage when larger, appropriate-width upper stages are available. As I recall, the crew Dragon for a lunar mission would need only minor modifications such as a more powerful communications system for communication from the Moon. It was already given sufficient heat shield capability in its design for return from escape velocity, which is higher than just return from LEO. Crew Dragon has some re-entry stability issues due to the greater height-to-width ratio of its OML compared to more traditional capsules. It's fine for LEO returns but I don't believe it can handle the higher angle of attack needed for cislunar returns. More importantly, though, it doesn't only need a more powerful communications system; it needs a service module. Its onboard propellant reserves are not enough for a lunar return, and certainly not enough for a lunar orbital insertion (in fact it can't even fire its main propulsive vacuum thrusters while attached to another module). And again, at this point, why are we bothering with SLS at all? I frickin love that thing. Now you've got me wondering what the maths would look like. The SLS core carries 29.9% more propellant than the Shuttle did. The intertank is essentially identical (not unlike the Ariane 5, the thrust of the solid boosters is transmitted in large part through the intertank in both the SLS and STS designs, leaving the hydrogen tank to hang in tension until booster separation). The SLS core masses 85.3 tonnes without engines. Crunching the numbers, you get roughly 34.4 tonnes for the tankage, 16.1 tonnes for the forward skirt section, and 34.7 tonnes for the engine boattail section sans engines. Note that SLS Block 1 also carries the 4.5-tonne LVSA adapter for ICPS. Let's imagine a few alternate possibilities inspired by the S-1D design. With all of these, I'm going to compare solely to SLS Block 1 to get an idea of what is actually possible, with the understanding that adding EUS to the design will create further improvements. To that end, everything north of the LOX tank remains identical: the same forward skirt, the same stage adapter, the same ICPS, and so forth, and total height must remain the same as well. For reference, here's the aft end of the current SLS: We plug in the following values for SLS Block 1 to LEO with Orion: Booster 1st Stage 2nd Stage Dry Mass, kg 102058.3 103872.1 3490 Propellant, kg 623689.5 952544 30710 Thrust, kN 17595 9116 110.1 Isp, seconds 268 452.3 462 We use default prop residuals, a restartable upper stage, and set a "payload fairing" of 6,926 kg (that's Orion's LAS) to jettison at 120 seconds. Launching from Cape Canaveral to a reference orbit of 185x185 km at 28.5 degrees using a two-burn ascent gives 85,671 kg to LEO. Note that this is under the earlier estimate both because of the addition of Orion's LAS and because I was using a number for the dry mass of the first stage that didn't include the mass of the engines or the LVSA. Time to get creative! Let's start small. Suppose that we keep the existing RS-25s but split that 34.7-tonne boattail engine section into two parts, one of which is attached to the core stage and one of which can be jettisoned: When do we jettison the outboard engines? Well, the boosters burn out and separate at T+131 seconds, after the core has burned 269.1 tonnes of hydrolox. At separation, then, SLS Block 1 has a T/W ratio of 1.13, not including payload. If we want this same T/W ratio to be maintained at outboard engine jettison, where our thrust will be cut in half, we need to wait to T+331 seconds, at which point there will be just 272.7 tonnes of hydrolox remaining in the tanks. For the purposes of plugging everything into the launch performance calculator, this creates a simulated three-stage architecture, where the "dry mass" of the "1st Stage" is merely the mass of the discarded outboard engines and engine section and the "propellant" of the "1st Stage" is the total prop burned at jettison, with all four engines as the "thrust" of the "1st Stage" and only two for the "2nd Stage": Booster 1st Stage 2nd Stage 3rd Stage Dry Mass, kg 102058.3 24404 79468.1 3490 Propellant, kg 623689.5 679858 272700 30710 Thrust, kN 17595 9116 4558 110.1 Isp, seconds 268 452.3 452.3 462 Plugging this into the calculator, with the same other parameters listed above, gives 94,523 kg to LEO. Now, before you go and say "wait, that's only what SLS Block 1 can already do!", remember that this is comparative to the 85.7 tonnes that the calculator gave us for SLS Block 1 before, which was notably low (in part because of the Orion LAS jettison). So in reality, we're talking about a ~10% improvement in payload just by dropping these engines at the correct time. Now, let's go one step further and replace those two outboard engines with RS-68As and only put a single RS-25 on the core: The RS-68A is only ever so slightly larger than the RS-25. The major issue with putting them on the aft end of the planned Ares V was that the heat flux from the solid boosters would have melted them, but hopefully with greater distance from the solid motors they will be in better shape here. I'm sure we can afford to slap some extra heat shielding on here somewhere if necessary. Together, a single RS-25 and a pair of RS-68As produce 9396 kN of thrust, just slightly more than four RS-25s. Good start so far. The specific impulse of the trio in combination is now going to be 421.8 seconds, based on thrust-specific Isp. The weight of the engine section is a function of the max thrust it has to transfer to the vehicle, so the total engine section weight is going to go up from 34.7 tonnes san engines to 35.8 tonnes. However, 75.7% of that is going to be in the thrust structure for the RS-68As, meaning more mass can be jettisoned. However, we'll have to burn all three for longer in order to make sure our thrust-to-weight is in good shape when we jettison. Because we are going to be jettisoning later (as in, nearly-to-LEO), gravity drag will be much lower and we can get away with a T/W ratio on the order of ~0.85. In the end, we get this: Booster 1st Stage 2nd Stage 3rd Stage Dry Mass, kg 102058.3 40608.4 62730.2 3490 Propellant, kg 623689.5 782100 170444 30710 Thrust, kN 17595 9396 2279 110.1 Isp, seconds 268 421.8 452.3 462 Once again plugging this into the calculator with the same parameters as before, we get a disappointing 86,717 kg to LEO. This is still technically an improvement over SLS Block 1, but it's not nearly as much of an improvement as our first concept. That's not too surprising; that extra 40 seconds of specific impulse on the RS-25 over the RS-68A makes a big difference. This is a sustainer architecture and so the extra thrust on the core stage doesn't do all that much for us. Finally, let's try a combination of these two designs: two RS-25s in the center and two RS-68As outboard: Now the core is really starting to get beefy. The total core thrust comes up to 11.7 MN so we need to beef up our engine section weight to 44.5 tonnes. Doing all the math as above, we get this: Booster 1st Stage 2nd Stage 3rd Stage Dry Mass, kg 102058.3 40608.4 79468.1 3490 Propellant, kg 623689.5 679858 272700 30710 Thrust, kN 17595 11675.3 4558 110.1 Isp, seconds 268 427.7 452.3 462 Unfortunately, this only gets us 84,567 kg to LEO. The extra thrust isn't being utilized well, and the extra weight is prohibitive. The RS-25 is a hard engine to beat. Just for the fun of it, let's come up with a REALLY wild proposal. Let's get rid of the SRBs entirely and replace them with Atlas V first stage boosters, and let's put a grand total of five engines on the core (four RS-68s and one RS-25). We'll stack a pair of additional Atlas V LOX tanks on top of each one, which we will use to crossfeed the RS-68As: Things are a little bit different now. The boattail section doesn't have to transmit nearly as much booster force and so I'll leave its total mass no greater than for the 2-and-2 design (though obviously it is getting two more RS-68A engines at an unpleasant 6,740 kg each). The empty mass of each booster is 21,054 kg, plus the weight of two LOX tanks which I will ballpark at an additional 30 tonnes per pair. Here the dry mass is added to the boosters but the propellant mass is added to the core, effectively, thanks to crossfeeding. Each pair of extra LOX tanks will hold 415442 kg of LOX (we will move the bulkhead in the core but otherwise everything will be as it was). Our numbers look like this: Booster 1st Stage 2nd Stage 3rd Stage Dry Mass, kg 51054 54088.4 62730.2 3490 Propellant, kg 284089 1510742 170444 30710 Thrust, kN 4152 20072.2 2279 110.1 Isp, seconds 337.8 417.6 452.3 462 To my delight, this gets 86,952 kg to LEO -- not as high as some of the other designs above, but really very impressive considering that we've completely ditched those big solid boosters entirely! We can also do the same design but use expendable Falcon Heavy side boosters (still sticking with the Atlas LOX crossfeed tanks so I don't have to rework the math): The numbers: Booster 1st Stage 2nd Stage 3rd Stage Dry Mass, kg 55600 54088.4 62730.2 3490 Propellant, kg 395700 1510742 170444 30710 Thrust, kN 8227 20072.2 2279 110.1 Isp, seconds 311 417.6 452.3 462 This configuration gets an absolutely whopping 121,720 kg to LEO with just ICPS on top. Replace ICPS with EUS as planned for Block 1B, and it can send 144 tonnes to LEO or 63.7 tonnes to TLI, nearly 40% more than the mythical SLS Block II and well over what Saturn V could do.

Agree that ULA is less likely to just ape SpaceX on the second stage. A Stoke-like aerospike adaptation of ACES isn't out of the question. But of course I want to see new and better ideas. It would be really cool if they used SMART reuse as a testbed for an inflatable ballute heat shield. ULA is much more likely than SpaceX to utilize disposable elements on an otherwise-reusable vehicle (see, e.g., Starliner). Prior to re-entry, the upper stage could vent remaining hydrogen pressurant gas into a series of inflatable phenol-impregnated bladders that would form a ring around the payload adapter, stretch up the sides of the vehicle, and shroud the engine bay. Those inflatable bladders would both increase the re-entry cross-section (which decreases heating) and absorb the bulk of the heat flux, then finally act as airbags to cushion the vehicle on its desert touchdown. It's still not the VTOL spaceplane we all REALLY want, but it would be cool if they could get it working. If they REALLY wanted to think outside of the box, they could build a lifting-body design with aerospike engines located in the wing strakes, perpendicular to the long axis of the vehicle. The booster carries the vehicle up and out of the atmosphere, so after separation the vehicle could simply rotate 90 degrees and thrust to orbit that way. That helps with the descent and landing portions of the re-entry because the vehicle can glide to the landing site and then fire its landing engines to hover its way down.

Damn I'm jealous

Actually, the relevant scenario is the two-stage case of a half-size Ariane core with a ca. 10 ton upper stage, so it is loftable by a single Vulcain. Your post is the one that said 2,500 kg is enough for a 3-person crew capsule, which is not correct. As for this other proposed vehicle... Per my notes above, the dry mass of this core would actually be 8,175 kg, taking the rest of your new numbers as assumptively correct (which I think is a generous assumption). For a crew launch vehicle you're also going to need an LAS, which I will ballpark at 2,030 kg for a notional 2,900 kg command module (the weight of the Cygnus-based capsule described above less the 1,800 kg service module) based on weight ratios of the Orion and other abort systems. I will set it to jettison 10 seconds after booster burnout to give time for separation and second stage engine startup. 1.34 MN divided by 431 seconds of specific impulse gives a propellant flow rate of 316.9 kg/s, so those 66.7 tonnes of propellant will be burned through in 210 seconds. Thus LAS jettison takes place at 220 seconds. Finally, to try and begin to account for the extra 300-400 m/s of gravity drag, I'm going to set the destination orbit to 185x1300 km, which is the equivalent of burning into a 185x185 orbit and then adding 300 m/s of additional dV. All told, this results in an estimated payload of 3,834 kg, still about a tonne shy of what is actually needed. Such a first stage would not in any way be reusable. No matter how cheap it is to add a second Vulcain to the Ariane core, it's still too expensive if the resulting vehicle can't reach orbit.

The Ariane 5 core is twice the max height of the VAB. Well, let's not get ahead of ourselves. To start with, let's do a one-to-one comparison with the current SLS to see if we have improved anything. If it's not improved, then swapping out upper stages is something that could as readily be done on the current SLS. Taking the specifications of SLS Block 1 as follows... Booster 1st Stage 2nd Stage Dry Mass, kg 102058.5 85275 3490 Propellant, kg 623689.5 952543.9 30710 Thrust, kN 17595 9116 110.1 Isp, seconds 268 452.3 462 ...the Silverbird Astronautics calculator gives an estimated 97.2 tonnes to a 185x185 orbit from Cape Canaveral, with the 95% confidence interval running from 79.0 tonnes to 118.6 tonnes. That's just 2% over the actual stated LEO performance of SLS Block 1 according to NASA, so we can be pretty confident that Silverbird is giving us good results in this configuration. So now let's redo it, but get rid of the boosters altogether and replace the four RS-25s with five RS-68s. First stage dry mass goes up by 19 tonnes (each of the RS-68s is almost twice as heavy as each RS-25), first stage vacuum thrust goes up to 17800 kN, and first stage vacuum specific impulse drops to 414 seconds: 1st Stage 2nd Stage Dry Mass, kg 104275 3490 Propellant, kg 952543.9 30710 Thrust, kN 17800 110.1 Isp, seconds 414 462 Running the numbers to the same orbit, we get a truly abysmal 35.7 tonnes, with a 95% confidence interval running from 28.7 tonnes to 44.6 tonnes. I don't like large solid boosters any more than the next person, but you can't discount the heavy lifting they are (literally) doing in the SLS design. Removing the boosters would mean removing 40% of the total propellant in the stack and you can't make up for that much of a loss with a higher-thrust core. If we are speculating wildly about a better SLS design that would still use the same core tankage, one tempting option would be to ape the old Saturn S-1D Mega Atlas design: One RS-25 in the center, four RS-68s on the outer thrust ring. Use Atlas Vs or Falcon 9s as side boosters and mount Delta IV common booster core oxygen tanks on top of them, so that you feed the two of the RS-68s their LOX from the side boosters. All four of the RS-68s get all of their liquid hydrogen from the core. Two of the RS-68s lose LOX and shut down at side booster separation, and then the other two shut down a minute or so later and the entire skirt is dropped once the thrust to weight ratio of the stack is optimal. No crossfeed necessary (although crossfeed would definitely improve performance). The core common bulkhead would need to be shifted slightly but the tankage mass and vehicle height would remain the same.

Your blog post at that link contains assumptions and calculations that are not representative of reality. As previously discussed, the Ariane 5G hydrogen tank is a hyper-thin balloon tank that is "hung" from the JAVE element; it is supported by tension rather than compression when it is under regular Earth gravity. If you remove the JAVE element then you'd need to strengthen the balloon tank by at least half the weight of the JAVE if not more. This completely newly-designed first stage you're contemplating would have a dry mass of at least 8525 kg. With your contemplated ~2500 kg payload and 79 tonnes of hydrolox propellant, gross liftoff weight is just a hair over 90 tonnes. At sea level, the Vulcain 2 developed 318 seconds of specific impulse and 0.99 MN of thrust; in vacuum it developed 431 seconds of specific impulse and 1.34 MN of thrust. With a 90-tonne GLOW, that's a liftoff T/W ratio of 1.1, which is too sluggish for an SSTO design that needs high initial acceleration to avoid gravity drag penalties. Compared to a more typical launch vehicle we are talking about an additional 300-400 m/s of gravity drag. Plugging the wrong numbers into the launch performance calculator will produce fictitious results. If you give it the wrong specific impulse then it's going to overestimate, particularly with an SSTO concept where specific impulse has such a high impact. Additionally, because the Vulcain 2 is highly underexpanded at sea level, the calculator will be using too high of a sea level specific impulse and too high of a sea level thrust, so it will underestimate gravity drag. Using the actual dry mass and actual vacuum specific impulse, the calculator gives just 1539 kg of estimated payload, and that's using too generous of a sea level specific impulse and ignoring that extra gravity drag. Your blog post also speculates that 2,500 kg is enough for a 3-person crew capsule. It is not. Your example, Cygnus, has a mass of 3,300 kg in its smallest configuration, BEFORE a heat shield, aeroshell, life support, parachutes, seats, or actual passengers. Using the two-person Gemini capsule as an example, the heat shield alone needs to be on the order of 7-10% of the total re-entry vehicle weight. Even setting aside the need for a new aeroshell, a Cygnus-based capsule would be on the order of 4,700 kg (339 kg heat shield, 230 kg for chutes, 133 kg integrated RCS, 78 kg RCS propellant, 350 kg crew seats and provisions, 270 kg crew). We haven't even talked about a launch escape system, which adds extra weight for the first few minutes of launch. You then suggest that altitude compensation can raise the vacuum specific impulse of Vulcain 2 to 466 seconds. It cannot. It is thermodynamically impossible for a gas-generator hydrolox engine to achieve 466 seconds of vacuum specific impulse. The reality: as a sustainer architecture, Vulcain 2 is already an altitude-compensated nozzle akin to the RS-25. Adding a more complex altitude-compensating nozzle might help with sea level thrust and thus counteract some of the gravity drag issues, but it would decrease vacuum specific impulse. So putting those fictitious numbers into the calculator will produce fictitious results (and even the fictitious 4,544 kg isn't enough to get a Cygnus-based capsule into space). Subsequent calculations merely compound the error.

Yeah... pretty big difference. Long enough for the panic of "Oh no, they lost one!" to start. I really wonder why. Even the guy on the NASA livestream (I refuse to watch this stuff on Xwitter) said it was noticeably different. There are a few options, I guess: SpaceX could be transitioning to a longer, staggered booster separation process, either to reduce the odds of a post-separation impact or to increase performance SpaceX could be experimenting with a new return boostback trajectory to increase performance One booster could have experienced a randomly greater wind buffeting, engine thrust fluctuation, or other random event which altered its return trajectory Of these, 3 seems least likely (those things can happen but it's hard to imagine them having THAT big an impact). Both of the others seem reasonably possible. A staggered booster separation could increase performance by allowing one booster to burn out sooner than the other, making it more like three separate parallel stages than two.

If the blaster bolts in Star Wars were moving at light speed in a galaxy with a slower speed of light, fast punches would produce relativistic effects, and ground speeders would experience time dilation.

And finally, SECO-1 with nominal parking orbit insertion. Go Psyche!!

T-60 seconds, vehicle on internal power, all systems go. All 27 engines firing -- leaving Terra Firma! Center core throttling down (visible in fire trail). Booster burnout. Booster separation successful. Center booster burnout. Successful stage separation, MVac ignition, and fairing separation. Side boosters started landing burns. Massive sonic booms. Successfully landed the boosters! Quite a significant difference in timing between the return of the two boosters...more than I remember from earlier FH flights.

SpaceX launches the Big Boy to a metal asteroid and I get my bar results, all on the same day. Metal.

Have you ever seen someone fire a gun? Unless someone has a very good suppressor, you can still see exactly where "slug thrower" shots are coming from.

...but, unlike the Falcon 9, incapable of reaching orbit with meaningful payload.