sevenperforce

SSTO is the only thing that has a ghost's chance of offering complete RLV capability in less time than it would take to simply build another ELV. An SSTO that runs on a single propellant combination without any sort of augmentation is going to have a negative payload fraction, because engines with high enough specific impulse to get you into orbit will be too heavy to get you off the ground. That's why people propose rocket-based combined-cycle engines. Unfortunately, those are all even heavier for their rated thrust than a high-specific-impulse rocket, but do not offer enough in-atmo specific impulse boost to make up for their added dry weight. SABRE might; we don't know yet for sure. We can get a reusable SSTO with a meaningful payload fraction. We merely need one of the following: A tripropellant engine which can run on a variable-ratio mixture of three different propellants (either one oxidizer and two fuels, or two oxidizers and one fuel) to produce high thrust at launch but high impulse for the orbital insertion burn. A super-lightweight afterburner to inject high-thrust propellant downstream of the high-impulse engine. An ultra-high-bypass ejector ramjet to achieve stupidly high specific impulse in-atmo without significant weight increases. Magic. Each of the above have their pros and cons. Option 1 is the most ideal, but the development challenges are huge and it may not be possible to achieve high efficiency. Option 2 allows you to build a more efficient engine more easily, but adds dry weight and doesn't burn the high-thrust propellant as efficiently. Option 3 is difficult to achieve in the supersonic regime without movable inlets, which makes dry weight really high. Option 4 is our best choice, but my application to Hogwarts was rejected.

Mostly profit. Thinking about the SES-9 mission and the idea of having a propellant depot in space with a tug to pull satellites into their desired orbit. Having an active propellant depot in orbit would be fantastic in terms of generally expanding our access to space, but unless it would be profitable, no one would ever do it. So...under what circumstances would an immediate-term (e.g., within the next five years) propellant depot and space tug system be profitable? Let's assume that the project would be undertaken by an independent private company rather than NASA. Let's assume it's not SpaceX or another current launch provider. Your primary customers would be comsat companies who choose to pay you some stated fee to tow their satellites into a desired orbit, though you could also potentially give the ISS boosts. You need at least one space tug -- probably several, to maximize your profit and to lower launch costs. Each tug needs solar panels, ion engines, a grappling or towing mechanism, and the ability to refuel in space (either by mating to disposable propellant capsules or by transferring propellant to an onboard tank. You'll probably want to buy two Falcon Heavy launches to start -- one for as many space tugs as you can fit, and one for your initial propellant depot. You'd then re-launch a Falcon Heavy each time you needed to top-up your propellant depot. Would anyone actually buy your services? How many comsat companies would say, "Hmm, instead of giving my satellite an engine and fuel tanks to push it into the desired orbit, I'll just pay these folks, and so I can make my satellite a little bigger and give it more capacity"? And would you be able to have low enough prices to actually save them money?

Splashdown requires recovery by ship, which is costly and requires a ton of infrastructure and support. The Soyuz capsule parachutes to land...but it actually requires its own set of solid-fueled retrorockets mounted on the parachute cables to slow down at the very end. A parachute which actually slows down a 7-tonne capsule to a gentle landing speed would be absolutely massive and extremely heavy. If the retros on the Soyuz fail, you live, but you break most of your teeth and possibly your spine. The Dragon V2 uses the landing rockets as an emergency launch abort system as well, saving weight at launch. Most capsules have had an external abort tower that is discarded every mission it isn't used. The landing legs soften touchdown and will probably collapse/crush in the event of a parachute landing on land, destroying themselves in the process but saving the occupants.

A crasher stage is just a fancy term for staging during descent. It can really only be used if there is no atmosphere; you don't want to do stage separation during atmospheric re-entry. You also need a final descent engine with hover capability (and if you plan to leave, it needs to be an engine you don't mind restarting). The Dragon 1 doesn't have a vehicle T/W ratio great enough for hover using RCS, so it's a no go. But the Dragon V2 has manned and unmanned configurations. They can absolutely strip off the parachute and heatshield and use it to land a rover or supplies or anything else. In fact, it would make a lot of sense to do exactly that, as a dry run for a manned landing. They could outfit a "disposable" Dragon V2 with additional payload to exactly match the launch mass for a manned mission, allowing them to test staging, lunar hover, and landing. The internal cargo hold is fine. An external one would be nice, but experiments wouldn't really be the point of the mission. As I discussed above, they could do a dry run first and include a rover that way if they really wanted to. You keep insisting that there is no concern for the moon, but I don't think that is true. The article I linked above has Elon specifically talking about using the Dragon V2 for landing payloads on the moon. The integrity of the heat shield would absolutely be a critical mission design element. I'm sure that they would run a whole suite of tests and experiments and simulations. If necessary, they could use a lightweight disposable blast shield that dropped off during ascent. Added solar panels wouldn't be in the path of the engine exhaust. They could retrofit the outer/upper skin with one and sun-track using RCS on the return, or they could simply use a fuel cell.

I was frustrated with the lack of rovers in the Demo, so I built one from scratch. It only has one axle. Does not play well with others. Once I got the hang of throttling each engine individually, though, I could turn and navigate fairly well. It did have a tendency to dance...and go boom. But if at first you detonate, try try again. He's still out there, trundling into the wilderness....

It has Draco thrusters, but those are RCS, not propulsive engines. In my first post, I explained how installing removable nozzle extensions and an auxiliary fuel tank in the cabin could give the Dragon V2 about 3,010 m/s of dV. In this post, I explained how Falcon Heavy could put the Dragon a few km above the surface and drop velocity down under 400 m/s; with the 10% payload increase from Full Thrust, this would bring velocity down to 70 m/s. The Falcon upper stage would break away and crash while the Dragon would separate from its trunk and make a propulsive landing, burning an estimated 200 m/s of dV (allowing a couple of minutes of hover). This leaves the ship with 2,810 m/s. According to this table, you only need 2,740 m/s for a direct ascent return from the moon if you can aerobrake on the way back. I wasn't talking about the trunk; I was talking about the internal cargo hold. I could be wrong about that, though. The solar panels are an issue I mentioned above; it would either need to be fitted with additional solar panels or use additional batteries, since it would discard the trunk (with its solar panels) prior to landing. If you're skeptical about whether SpaceX would actually be able to make these modifications, or be interested in interplanetary missions with Dragon V2 and Falcon Heavy, try this on for size: So I'm not proposing anything they haven't considered, at least in piecemeal. I'm just putting it all together to show that a direct ascent moon shot is entirely within the system's reach.

With nuclear power the weight of extra propellant isn't going to be much more than the weight of a ram-air augmentation system.

Dragonlab does not have engines.

So...Project Pluto, but with a fuel tank? Could work.

Better performance was vs VTHL SSTO pure rocketplanes.

Call it "buy-in" rather than investment, then. Demonstrating that it's possible to land the same manned spacecraft on multiple worlds is a huge, huge step. 0.o Yeah, silly apples to the thing least like apples you can imagine comparisons are so useful. Seriously, the effort required to design a ludicrously specified vehicle is completely and totally irrelevant to the question at hand - the difficulties involved in converting a capsule as specified. Just because x is simpler than y does *not* means that x is in and of itself simple. My point is that if it's simpler to modify an existing capsule than it is to design a purpose-built craft, without significant performance reductions (let alone prohibitive ones), then we have demonstrated a really critical step towards reusability and flexibility in interplanetary expansion. And we should do it. I can't imagine that lunar dust would be more dangerous to heat shields than plummeting through the atmosphere at fifteen thousand miles per hour. I mean, we would definitely want to do testing to make sure, but it seems like a really minor problem. The crewed model of the Dragon V2 already has a cargo hold, IIRC. There is no drop tank. The ship would use the same landing legs for touchdown that it will use on Earth. The only added aerodynamic modeling would be for the attachable vacuum nozzle extensions. Everything else is standard, with the exception that crew capacity has been dropped in exchange for mounting those auxiliary tanks in the cabin. I'm not saying it wouldn't take extensive research and effort. And yeah, it would have a hefty price tag...probably a quarter billion. But those aren't reasons not to do it.

Compared to designing a direct ascent lander capable of Earth re-entry without staging? Yeah, I'd call designing, building, and installing a couple of tanks (each less than 2 m across) is minimal. Yeah, the low-isp upper stage is why a direct ascent on hypergols results in a lower launch mass than LOR. SpaceX has no plans for colonizing the moon, but they certainly have contracts for lunar missions. A demonstration mission, landing a two-man crew on the moon and returning them using existing hardware would go very far toward convincing investors that a Mars shot is worth investing in. SpaceX would not only become the first private company to land men on another world, but Dragon V2 would become the first manned spacecraft in history to land on more than one world. While, I might add, still being reusable. Yeah, I was expecting that. I read that the quoted 53-tonne-to-LEO figure was for a 200x200 km 26-degree orbit from Boca Chica, though I don't recall how solid a source it was. But that's okay, because we do have a solid source, if we do the math ourselves. That 53-tonne figure was the original stated specification for Falcon Heavy when it was first announced in early 2011. It was accompanied by a quote of 12 tonnes to GTO. But by 2013, the quoted GTO payload had come up to 21.2 tonnes...and that is before the Full Thrust modifications. That quote for GTO is useful for two reasons: first, any inclination changes necessary for TLI would be comparable to those required for GTO, so we can ignore that. Second, we know about what perigee SpaceX uses for GTO when maximizing payload, because they used it with SES-9 just a week and a half ago -- it's roughly 200 km. So a fully-expendable Falcon Heavy can deliver 21.2 tonnes to GTO on a 200x35,786 km orbit. That orbit has a perigee speed of 10.25 km/s, while the 200x200 km parking orbit has a constant orbital speed of 7.784 km/s, corresponding to a dV of 2.47 km/s. The upper stage has a dry mass of 3.9 tonnes, so to execute a 2.5 km/s burn with a 21.2 tonne payload would require 51.7 tonnes of fuel. Looks like that originally-quoted 53-tonne figure was a little low. Remember, this is all before the Full Thrust upgrades. Assuming 13.7 tonnes for our Dragon V2 , an upgraded expendable Falcon Heavy should therefore be able to loft it to a 200x200 orbit with 59.2 tonnes of fuel left in the upper stage. That's a dV of 5.02 km/s. Enough? It should be. Extending the apogee out to EML-1 at 346,016 km requires a perigee speed of 10.96 km/s, eating up 3.18 km/s. With the right transfer trajectory, this will result in lunar capture with a free fall to the lunar surface and an impact velocity of 2.24 km/s, slightly under lunar escape velocity. The upper stage can reignite its engine and burn retrograde to engine-out a few dozen km above the lunar surface, reducing the vehicle's speed to 396 m/s. The Dragon V2 can ignite its engines and break away, leaving the Falcon upper stage to crash-land while it makes a propulsive landing. Cutting it close? Sure. But with the Full Thrust upgrades this should be absolutely doable.

Whoops, sorry. Fixed it.

I provided the citation above; it's this MIT study.

They have postponed development of a crossfeed system for the time, so who knows whether they will ever build it. For most FH launches, they will burn all three engine clusters at 100% to start, then downthrottle the core stage as far as it will go through maxQ until booster separation.

Geez, semantic arguments? Really? It's obvious what the OP meant, so anyone talking about ice is just joking. No need to challenge or dissect it. Besides, while "water" is typically used to refer to a liquid state in lay conversation, any discussion in the context of science, engineering, or space is going to permit the use of "water" to refer to all three phases. "Ice" is not restricted to water ice in scientific circles.

The dV each way between LLO and the lunar surface is 1.87 km/s, meaning that a lander in LOR needs over 3.7 km/s. Definitely above what we could squeeze into the Dragon V2. However, the dV from LLO to an Earth aerobraking trajectory is less than 1 km/s. So if you can get down to the lunar surface using a crasher stage, your ascent vehicle needs only 2.7 km/s to get all the way back to Earth (as long as it has a heat shield for braking). The Dragon is billed as having a two-week minimum crewed endurance so I can't imagine a lunar landing would be a problem. And like I said above, direct ascent to Earth only requires 2.7 km/s if you can zero out lunar orbit velocity by using your transfer module as a crasher stage. No drop tank required. The only potential issue is power for the Dragon on the moon and on the return trip; they would have to use an extended battery because they would shed the trunk with its solar panels before landing. And if SpaceX is billing the Dragon V2 for Mars missions, demonstrating a manned lunar landing with a reusable landing and return vehicle would be exactly up their alley. Modifications to the inside are minimal. The two-man crew allows the existing life support systems to suffice. No changes to maneuvering systems would be anticipated.

All right, so I had been noticing the low impulse on the SuperDraco engines on the Dragon V2 and surmised in another thread that adding attachable vacuum exhaust extenders could bring the impulse high enough for the Dragon V2 o serve as the second stage all on its own. That's probably not going to work. But I thought of something that would. The Dragon V2 will carry 1,388 kg of onboard propellant. However, the crewed version can accommodate 7 with its 10 cubic meter volume. If the crew were reduced to two and their living size was increased by 50%, this would free up 5.7 cubic meters of space for an auxiliary internal tank. The propellant combination used for the SuperDraco engines is quite dense, at 1200 kg per comic meter. So that is an extra 6.84 tonnes of fuel. Add 200 kg for the tanka to the vehicle dry mass and assume your two-man crew and their consumables mass 1,103 kg (a third of the nominal payload capacity for the Dragon), and you end up with a non-fuel mass of 5.5 tonnes and a loaded mass of 13.7 tonnes. With the 336 second specific impulse, that's a dV of 3.01 km/s. What's that good for? Well, Falcon Heavy was originally quoted at 53 tonnes to LEO, but with the Full Thrust modifications this will probably come up to about 10% higher, or 58.3 tonnes to LEO. LEO to the lunar surface is a dV of 5.93 km/s; the Falcon Heavy should be able to drop off around 15 tonnes to the lunar surface as a crasher stage. So a FH can lift a retrofitted Dragon V2 to the moon and then allow the Dragon to return to LEO (or simply Earth aeropcapture) under its own power. Thoughts?

Black Horse wasn't my idea. Neither was using H2O2 for the oxidizer. Specifically, JP-1 and H2O2 with aerial propellant transfer (oxidizer only) and airbreathing up to around Mach 5 outperforms LH2/LOX SSTO.

It wouldn't switch sources. The drop tanks would feed the internal tanks directly.

Hydrogen peroxide can't be decomposed to produce hydrogen; it produces oxygen. And yeah, the SABRE engine runs on LOX. I'm talking about the Black Horse project. The engine would have to be redesigned to use H2O2.

Yes, the pressure feeding does make this a slightly more challenging system. Might be worth it to use lighter-weight tanks with a small turbopump. I hate to throw the turbopump away, of course, but at least it could be exceedingly simple because it could run off MMH as a monopropellant. A major advantage is the extremely high propellant density. The tank size would be physically far far smaller than the Falcon second stage.

There is no SECO required. The drop tank would be designed to feed into the internal tank with flow-through, so that the engines are pulling from the internal tank the whole time. Launch abort just means severing the feed lines (which would redundantly self-seal) and dropping the tank, with the engines remaining on full thrust the whole time. The nozzle extensions would have to break away as well for a low-altitude abort. ISS is an issue, I know. But having a 98% reusable system is good regardless.

I meant that it isn't a redesign of the carrier rocket. It isn't crossfeed because there are no other engines involved. It's just adding external fuel couplings that feed into the internal fuel tank. Sure, it takes work, but it is more a retrofit then a full redesign. Yeah, the second stage would definitely be used for any additional payload. This would strictly be for crew ferry only. Heating is a problem. Not sure how to manage that. Thrust isn't a big deal; the T/W ratio will be fine. But the tilt angle does cut into effective specific impulse. However, you can just stage at a slightly higher speed to take care of that.

The cost of living is really going to be prohibitive unless there is some compelling need to be there. Automation really does eliminate almost all reasons for working in space. So the only reasons I can think of would be the ones I mentioned above which depend on people being outside of any government's jurisdiction.