sevenperforce

Also, they say that Musk has already been "testing" the Falcon Heavy "spacecraft" since 2011 or something. Neither of those things make sense. It's called the Dragon V2, not the Crew Dragon, because it will replace the current Dragon for unmanned cargo launch. Moreover, the launch abort system has been pad-tested but not flight-tested, something the author seems completely uncertain about. Further testing is not to "make sure passengers are completely safe"; it's to test the damn system in action.

Yeah, the internal payload bay makes little sense. I really need to do some math to have a better idea of how it can be scaled. Most cargo deliverables are on the order of a Dragon V2's mass or greater. So to use the same RLV for crew transfer as for cargo flights, you want a vehicle which can SSTO a manned capsule in its base, stripped-down configuration but can have its thrust and its tankage capacity augmented to lift heavier payloads. That's the major advantage of using an open flowpath: you get to play around with the possibility of multiple fuel injection points.

Well, Falcon Heavy can take a manned Dragon V2 to Mars, no problem. You'll just die enroute...probably from carbon dioxide poisoning, if nothing else kills you first. But even if you did miraculously survive and land on Mars in the Dragon (the final aspect of which, I must point out, is totally workable)? Good luck getting off-planet, let alone back to Terra.

IIRC, the lack of efficacy associated with the V1 had less to do with inherent limitations of rockets and more to do with Britain's excellent counterespionage efforts that effectively kept the Germans blind to where their rockets were falling.

Not sure if that is the case for fuel-rich combustion or only for stoichiometric combustion.

Are we just ignoring the V1?

To answer why you can't simply keep increasing chamber pressure to get higher ISP: eventually, as you approach that upper theoretical limit, your fuel no longer has any more chemical energy available to increase the heat and pressure in the combustion chamber. Full-flow staged combustion has the disadvantage of requiring two separate preburners. However, this means the requirements on each one is lower, and since you already need separate pumps, this isn't much additional weight cost. And yes, the Russians did oxy-rich staged combustion with RP1. That's not problematic because burning oxygen-rich eliminates coking. When there is much more oxygen available then needed, all the fuel fully combusts. In contrast, burning hydrocarbons in a fuel-rich mixture is begging for coking. Staged combustion has a huge advantage because thrust-specific fuel consumption is the single biggest factor in keeping your mass ratio low. If you are throwing away fuel that doesn't contribute to thrust, you're throwing away specific impulse.

I think Skylon hasn't gotten a lot of press because it has made only very incremental developments. While important developments, they are still "small stuff" by press standards. It doesn't get a lot of discussion on the forums because it's a fairly noncontroversial and limited thing. If it works, great; we'll have a cheaper way of getting to orbit. If not, oh well. But it doesn't really expand our access to space. It is limited to LEO; it can't send something on GTO like SpaceX's Falcon 9. And that's if it works at all, which remains to be seen. Using airbreathers to orbit is a really tricky proposition. It's not just about altitude, either. The faster you're moving, the greater the kinetic energy of the incoming air. You've got to get that air moving out the back of your engine faster than it came into your engine. So even if you can design a high-thrust engine and somehow minimize drag, there is a limit to how fast you can be going and still gain anything. Eventually, the airstream will be moving in faster than you can accelerate it back out.

Falcon Heavy can take a Dragon V2 to Mars but not a manned one.

Errr... They think it is possible to land on Jupiter?

It is easy to forget just how primitive the moon landings were by modern standards. The Apollo landers used two separate engines, one for descent and one for ascent. Why? They didn't have the technology for deep throttling, even with hypergolics. Nor could they reliably execute a drop tank design. Today, we would obviously use a single engine/engine cluster with separate tanks for descent and ascent and merely drop the descent tanks along with the landing legs. But they had to use two separate engines. If Apollo was done today, we would probably leave the command module unmanned during the lunar landing and use a single engine for lunar orbit injection, descent, ascent, and LEO transfer.

A dazzling one. Any tripropellant design is an engineering nightmare; an efficient one ever so much more so.

Depends. Could we build a spacecraft that looked like that? Sure, I suppose. Would it have the same capabilities? Probably not.

More efficient than the 1.5-stage design, though also more complicated, is the variable-ratio tripropellant liquid-fueled engine. Design an engine that can burn two different types of fuel (typically LH2 and RP1, though other combinations are possible) along with a single oxidizer. Most importantly, design the engine such that it will be able to vary the fuel and oxidizer ratios infinitely while maintaining maximum efficiency. That way, you can start your launch with a high proportion of RP-1 for maximum thrust, then vary the ratio slowly to more and more LH2 as your thrust requirements drop, increasing specific impulse. Putting the RP1 in a drop tank makes this design even better.

I thought the capsule-as-payload configuration was pretty obviously spelled out with, "The crew-carrying variant would be mated at the top to a Dragon V2 capsule, while the cargo variant would hold an internal payload." Maybe I should have said "configuration" instead of "variant"? Also, if you're familiar with tthe Chrysler SERV, you'd recognize the same arrangement: internal payload bay for cargo; mated at the top to a separate crew vehicle for crew transfer.

Whoa, sheesh. You said that having a separation event was a thing to be avoided. The only alternative I could think of would be to not have a separation event, i.e., make the cabin and rocket a single body. Am I missing something? Docking with some destination would presumably be the target. I mean, you can go up in a capsule and come down again after a couple of orbits, I suppose, but the present primary need for a manned launch vehicle is to ferry crew to the ISS. Well, we'll get good at it, but we're already good at returning capsules. The proposed model is a rocket capable of delivering payloads to orbit via SSTO. You're throwing out big warnings about one of those payloads being a manned capsule (and yes, it would obviously require a service module) and suggesting that separation is a problem. This would presumably mean you have some alternative configuration in mind...so, what would that be?

We must be talking past each other, or something. Would you have the crew cabin be built into the nose of the rocket, so that the rocket goes and docks with the ISS like the Shuttle did? The point is to have a RLV which can launch different payloads based on its configuration. I don't see what would be gained by building the crew cabin into the rocket body. Why is the on-orbit lifetime of the capsule measured in hours? The Dragon V2 is supposed to have a pretty lengthy orbital manned persistence. How is it less safe to be in orbit in a Dragon V2 than in a gigantic capsule-shaped rocket? Separation accomplishes quite a few things. For one, it returns the launch vehicle immediately, so it can be refurbished and reused while the capsule remains in orbit (or docked with the ISS) for as long as required. It also allows the whole system to be modular, so that the same launch vehicle that put a satellite into orbit last month can launch a manned capsule to the ISS this month. Well, at the very minimum, it's easier to get a 7-tonne capsule back to Earth than it is to get a 50-tonne nearly-empty rocket back to Earth. It's something we're good at. A capsule like the Dragon V2 carries its own engines for either propulsive landing or launch abort, and (as I said) can carry backup parachutes. Keeping the entire monster in orbit and then deorbiting it with the crew cabin still in it is much less safe because it can't carry backup parachutes, and because its re-entry profile will be different. A crew capsule is optimized for one thing: returning a crew safely. The launch vehicle, on the other hand, has to be optimized first for launch and only secondarily for re-entry. Yes. But the ducts go all the way through. You're using PICA-X type heat shielding on the base and it extends up into the ducts. Plus, the plasma escapes through the ducts.

That's beautiful. I concur with what the commenters said: the Shuttle is undoubtedly the most awe-inspiring, aesthetically breathtaking, mission-capable thing we've ever put into space, but WOW was it a waste of money. I wonder how recovery would have gone with that ring.

In case of abort? And I quote from the original post, just as I did in my post. It's not an abort mode - it's a normal mode. Wait, I'm completely confused. What is wrong with a separation event? Particularly an on-orbit separation event? Every single spacecraft (not just manned ones) to enter orbit has had at least one separation event. This one would be far safer because it occurs in orbit. And having a separate crew capsule which lands independently allows for catastrophic launch abort while also being safer than the alternative. It allows the rocket to be landed in the way that's best for the rocket and the capsule to be landed in the way that's best for the capsule. The capsule can have backup parachutes which would not be feasible for the rocket. Indeed, and this would have fine crossrange performance. It just wasn't something possible when SERV was proposed. The goal of the central bypass design is to allow the airflow through the duct to self-correct; the rocket would use a launch profile designed to optimize this. Not sure if this is possible; if not, I'll probably design a variable-attack-angle craft. And yeah, I'll need to smooth the exhaust surfaces for re-entry. No problem getting re-entry plasma in the ducts, though; they're going to be handling rocket exhaust as it is. The worst part of the idea. Yeah, the payload fairing needs to be fixed. The ship would be scaled such that the nominal payload to LEO would be something around twice the mass of the Dragon V2; this will allow it to carry both a crew capsule and a small cargo load simultaneously. With crew capsule alone, it could probably take the capsule to the Earth-Moon Lagrange points. You don't like hydrazine, do you? Haha. I wanted a monopropellant that could burn with the exhaust plume in the ducts and thus serve as an afterburner.

Why in the world did I say that?

If you aren't familiar with the Chrysler SERV concept, check it out. Quite possibly the most innovative of the proposed Shuttle concepts, though it was never seriously considered. The military need for crossrange maneuvering during abort-once-around polar orbit launches, coupled with a lack of reliably restartable rocket engines, led to them adding dozens of jet engines to the design...which really doomed it. But with the air augmentation of a ducted rocket design, inner side-mounted deeply throttleable linear engines, and an overall ballistic re-entry profile, this could be promising. The crew-carrying variant would be mated at the top to a Dragon V2 capsule, while the cargo variant would hold an internal payload ducting would allow for weight-free afterburn injection to increase launch thrust for larger payloads with an extended upper fairing: The ducts at the top would allow maximum airflow and thrust augmentation without adding significant drag: The lower base would have a passive heat shield, dissipating heat both around the body and up through the heat-resistant ducting. The linear engines are positioned in such a way as to make maximal use of the atmosphere, from aerospike effect and thrust augmentation all the way up into vacuum operation, without needing to use any moveable flaps or panels: Downstream injectors (not shown) allow for added thrust at takeoff, using a denser fuel (yes, I favor hydrazine) to combust with the existing exhaust stream and the airflow. This would be used when the extended fairing holds a heavier payload and requires additional thrust. Although this requires a greater weight of fuel, thrust augmentation would be significant enough to increase the T/W overall and keep the total dV constant. Due to the possibility of afterburn-style thrust augmentation, this vehicle can have its thrust increased dramatically at the cost of reduced delta-v, making it a fantastic super-heavy lift first stage. After reaching orbit and ejecting payload, re-entry would take place using the base heat shield, landing on a hover at the launch pad with the throttleable main engines. In no case would re-entry be manned, as the capsule would re-enter separately. Since orbit would be achieved on virtually all missions (with the exception of super-heavy lift scenarios as noted above), this wouldn't have the SpaceX downrange problem or need to land on a barge; it could just abort once around and come back to the launch pad every time.

insert_name is right. All objects have a center of mass, and solid objects can spin stably on any axis passing through that center of mass as long as the angular momentum around that axis is maintained. However, with a liquid-filled object, the centrifugal forces caused by spinning produce internal forces on the liquid, causing it to circulate. The circulation transfers angular momentum away from the spin axis, producing a new spin axis, which produces additional force, further misaligning the spin axis. Plus, if there are bubbles in the liquid, this can also alter the center of mass, further messing things up. With the long cylinder, the liquid still circulates, but the shape of the body causes it to circulate in perfect alignment with the spin axis, so there is no misalignment and no shift in angular momentum.

I'd still maintain that once you relax the requirement of direct launch from Earth, a lunar mission with full reusability is not necessarily an open-shut win for LOR. With LOR and a separate reusable lander, you'll need to leave LEO with: Earth orbit boost engines (to get you out of and into LEO quickly enough to avoid the Van Allen belts) Fuel tanks for Earth orbit boost engines Transfer engines Fuel tanks for transfer engines Transfer hab and control center Supplies for transfer hab Descent hab and descent control center Supplies for descent hab Descent/ascent engines Fuel tanks for descent/ascent engines Two separate bodies, one with landing legs and one without With a single craft, you need to leave LEO with: A single hab and control center Supplies for the hab A single engine cluster A single rather large fuel tank A single body with landing legs With LOR, you have the advantage of leaving some stuff in orbit...but when it all boils down, the only things you actually end up saving on are the fuel for the return trip and half the supplies for the transfer hab. Everything else you either bring down with you, or is an extra requirement of LOR. On the one hand, you have to build two separate habs with separate life support systems, two separate crafts, and a maze of fuel tanks of varying sizes. On the other hand, you have more fuel in one tank.

Why not?

Ditch the jet engines in favor of throttleable linear aerospike central-bypass ducted rockets positioned around the payload bay, and you'd have something.