Speculation thread: intermediate-lift Raptor-derived RLV

[DOCinCT]

How does the lander get vertical again, for takeoff? Once fueled the landing engines will not be enough to get it off the ground will they?

Obviously lwith just an appropriate application of levers.  
(Image from Colliers Magazine April 30, 1954 as reproduced in Horizons Vol 39 Issue 2 Sept/Oct 2013, used with permission)

[sevenperforce]

This makes sense, except that you're comparing to the ITS tanker. As I pointed out already, the tanker isn't designed to fly to Mars. It most likely doesn't have TPS for an interplanetary entry, which will probably mass about twice as much as TPS for Earth LEO entry. It probably doesn't have significant solar power, operating mostly off batteries for the few hours it's in LEO. And it probably doesn't have active or passive cooling to keep the landing fuel from boiling off during a 6 month transit, which is how long a mass-optimized transfer to Mars takes. I'd figure dry mass of at least 5% of wet, before payload.

A little OT -
I figured that orbital refuel will be done first to one of the tankers on orbit, and than from the refueled tanker to a crewed spaceship. That means the tanker will stay longer on orbit and will act as a fuel depot for the spaceship.
The spaceship will have it's fuel ready on orbit when it is launched thus increasing crew safty.
For that to happen, at least one of three tankers, if not all, need to have built in cooling and power to act as a depot.

Yes, that's what I've assumed as well. It doesn't make sense for the crewed spaceship to just sit around for 5+ tanker launches; much more sense if the tanker goes up and acts as a fuel depot. If they can hit their mass fractions, it's not gonna be too tough to get 380 tonnes to orbit each launch, or maybe even a little bit more.

In any case I'll keep my mass fractions for conservatism.

Before running more (incorrect) numbers for different trajectories, you should try to match your numbers with the ACTUAL numbers released by spaceX.

Errrr...what? What SpaceX numbers am I missing?

You need 1000 m/s for a capsule, but not for a biconic lifting body. Even 800 m/s is an overestimate, I think. The Mid-L/D MAV concept allocates 650 m/s for the landing burn, and my upper vehicle would be much fluffier (higher-drag, lower-mass) than the Mid-L/D. Speaking of which, the re-entry profile is quite similar:



I designed this from the ground up as an Earth-based reusable upper stage, but the adjustments to make it work equally well for Mars or the moon are really, really minor. The few tweaks merely make it a more capable spacecraft for Earth missions.

How does the lander get vertical again, for takeoff? Once fueled the landing engines will not be enough to get it off the ground will they?

Ah, see, that's the beautiful thing. Providing enough thrust for aggressive, efficient SL landing on Earth means that the landing thrusters are able to push a total of 736 kN in the lower-pressure Martian atmosphere. That's enough to lift 198 tonnes on Mars, meaning it could theoretically take up to 45 tonnes of payload to Martian orbit.

[sevenperforce]

Notionally, consider the following plan for a Mars Sample Return mission:

Two-tonne rover capable of acquiring samples is launched to LEO in the cargo-variant upper stage. A tanker-variant upper stage is also launched to LEO. Both are fully refueled in orbit and exit together on TMI. After the TMI burn, the mission spacecraft has 35 tonnes of propellant remaining; the tanker has 55 tonnes of propellant remaining.

Immediately after the TMI burn, the two upper stages rendezvous and the tanker transfers 53 tonnes of propellant to the mission spacecraft, then adjusts its trajectory to perform a Martian free-return. It will have enough residuals for high-energy Earth EDL after its loop around Mars. The mission spacecraft is now 62% fueled, with 88 tonnes of propellant.

The mission spacecraft performs a high-energy entry and landing on Mars, reaching the surface on its auxiliary thrusters with 69 tonnes of propellant. The rover exits and picks up a series of samples, then returns to the mission spacecraft.

The mission spacecraft lifts off on its thrusters, fires its main engines, and rockets toward the solset on a direct ascent to Earth Injection. It performs a high-energy entry and lands on Earth with 5 tonnes of propellant to spare.

Fully-reusable Mars Sample Return with no ISRU required, at the cost of only two reusable Falcon-family launches plus refueling runs.

EDIT:

Adapting the same mission plan as before, but for a crewed lunar mission:

Fully-fueled manned vehicle (dry mass 20.5 tonnes, payload 4 tonnes including crew) and tanker head for TLI out of LEO together as before, with the tanker transferring its propellant reserves to the manned vehicle immediately after the TLI burn and coming back on a free-return trajectory to land. Manned vehicle reaches cislunar space with 112.8 tonnes of propellant, executes orbital entry, deorbit, and landing to reach New Tranquility Base with 48.8 tonnes of propellant remaining.

After the mission (which can last quite a while, given that total delivered payload is the same as the entire gross mass of the Apollo Lunar Ascent Module), the manned vehicle lifts off on its thrusters, ignites its main engines, and heads on a direct ascent to Earth. EDL is completed with a whopping 10.2 tonnes of propellant to spare.

[rakaydos]

This makes sense, except that you're comparing to the ITS tanker. As I pointed out already, the tanker isn't designed to fly to Mars. It most likely doesn't have TPS for an interplanetary entry, which will probably mass about twice as much as TPS for Earth LEO entry. It probably doesn't have significant solar power, operating mostly off batteries for the few hours it's in LEO. And it probably doesn't have active or passive cooling to keep the landing fuel from boiling off during a 6 month transit, which is how long a mass-optimized transfer to Mars takes. I'd figure dry mass of at least 5% of wet, before payload.

A little OT -
I figured that orbital refuel will be done first to one of the tankers on orbit, and than from the refueled tanker to a crewed spaceship. That means the tanker will stay longer on orbit and will act as a fuel depot for the spaceship.
The spaceship will have it's fuel ready on orbit when it is launched thus increasing crew safty.
For that to happen, at least one of three tankers, if not all, need to have built in cooling and power to act as a depot.

That's what EVERYONE figures... but that's not what SpaceX is actually saying.

It's always been "send the interplanetary craft up, then refuel it, and if that takes more than a week, send the crew after."

this "take out EVERYTHING that isnt absolutely essential to short-stay LEO refueling to increase fuel capacity" idea explains the seemingly-obvous disconnect between the "right" thing and what spaceX is saying. And since SpaceX has run the REAL numbers, I'd say they probably have the right of it.

[rakaydos]

Notionally, consider the following plan for a Mars Sample Return mission:

Two-tonne rover capable of acquiring samples is launched to LEO in the cargo-variant upper stage. A tanker-variant upper stage is also launched to LEO. Both are fully refueled in orbit and exit together on TMI. After the TMI burn, the mission spacecraft has 35 tonnes of propellant remaining; the tanker has 55 tonnes of propellant remaining.

Immediately after the TMI burn, the two upper stages rendezvous and the tanker transfers 53 tonnes of propellant to the mission spacecraft, then adjusts its trajectory to perform a Martian free-return. It will have enough residuals for high-energy Earth EDL after its loop around Mars. The mission spacecraft is now 62% fueled, with 88 tonnes of propellant.

The mission spacecraft performs a high-energy entry and landing on Mars, reaching the surface on its auxiliary thrusters with 69 tonnes of propellant. The rover exits and picks up a series of samples, then returns to the mission spacecraft.

The mission spacecraft lifts off on its thrusters, fires its main engines, and rockets toward the solset on a direct ascent to Earth Injection. It performs a high-energy entry and lands on Earth with 5 tonnes of propellant to spare.

Fully-reusable Mars Sample Return with no ISRU required, at the cost of only two reusable Falcon-family launches plus refueling runs.

EDIT:

Adapting the same mission plan as before, but for a crewed lunar mission:

Fully-fueled manned vehicle (dry mass 20.5 tonnes, payload 4 tonnes including crew) and tanker head for TLI out of LEO together as before, with the tanker transferring its propellant reserves to the manned vehicle immediately after the TLI burn and coming back on a free-return trajectory to land. Manned vehicle reaches cislunar space with 112.8 tonnes of propellant, executes orbital entry, deorbit, and landing to reach New Tranquility Base with 48.8 tonnes of propellant remaining.

After the mission (which can last quite a while, given that total delivered payload is the same as the entire gross mass of the Apollo Lunar Ascent Module), the manned vehicle lifts off on its thrusters, ignites its main engines, and heads on a direct ascent to Earth. EDL is completed with a whopping 10.2 tonnes of propellant to spare.

We may be overestimating the capability of the tanker. That dry mass is amazing, but it's got to come from somewhere... and SpaceX has gone out of their way never to claim  the tanker has the endurance to be a fuel depot. Instead it's always "send the spaceship, refuel with tankers, and if that takes too long, send (another spaceship with) crew after. Batteries, insulation, probably more- the tanker's going to be shaved to the bone to squeeze in more fuel.

Cutting the heat shield back to being able to barely handle LEO is another way to save on dry mass... and one synergestic with other dry mass reductions, as a lighter tanker (with the same aerodynamic footprint) has less reentry stress. Does a hollow carbon fiber tube even need heat shielding?

But while this improves the tanker's ability to do it's primary task, it means any cargo or fuel depot has to use the SPACESHIP's dry mass numbers.

[sevenperforce]

We may be overestimating the capability of the tanker. That dry mass is amazing, but it's got to come from somewhere... and SpaceX has gone out of their way never to claim  the tanker has the endurance to be a fuel depot. Instead it's always "send the spaceship, refuel with tankers, and if that takes too long, send (another spaceship with) crew after. Batteries, insulation, probably more- the tanker's going to be shaved to the bone to squeeze in more fuel.

Cutting the heat shield back to being able to barely handle LEO is another way to save on dry mass... and one synergestic with other dry mass reductions, as a lighter tanker (with the same aerodynamic footprint) has less reentry stress. Does a hollow carbon fiber tube even need heat shielding?

But while this improves the tanker's ability to do it's primary task, it means any cargo or fuel depot has to use the SPACESHIP's dry mass numbers.

Yeah, envy pointed this out a page back or so, and I redid my dry mass numbers to adjust. My dry mass, for the cargo and tanker and crew variants, are all based on the ITS Spaceship's dry mass numbers.

By that, I mean that I used the dry mass of the spaceship (less projected engine mass) divided by the propellant capacity of the tanker, and used that to find my structural/tankage mass fraction. The two ships have the same external dimensions.

[spacenut]

With a horizontal landing, what about dust and debris hitting the heat shield material landing on Mars?  Will it be strong enough to take some "pings"?  Can it be repaired while waiting for refueling?  Can longer landing legs work? 

[sevenperforce]

With a horizontal landing, what about dust and debris hitting the heat shield material landing on Mars?  Will it be strong enough to take some "pings"?  Can it be repaired while waiting for refueling?  Can longer landing legs work?

The landing thrusters are set in the wing extensions, rather high off the ground...maybe 4 meters up? With the low air pressure on Mars, the plume is going to be really diffuse by the time it hits the ground, so debris impingement should be low or nonexistent. Same with landings on the moon.

For Earth return, landing would be on a pad, so no worries there.

This is a problem for the ITS; I don't really know how they are going to handle it. The engines are going to be directly in the debris impingement zone even if the heat shield is not.

[dror]

With a horizontal landing, what about dust and debris hitting the heat shield material landing on Mars?  Will it be strong enough to take some "pings"?  Can it be repaired while waiting for refueling?  Can longer landing legs work?

The landing thrusters are set in the wing extensions, rather high off the ground...maybe 4 meters up? With the low air pressure on Mars, the plume is going to be really diffuse by the time it hits the ground, so debris impingement should be low or nonexistent. Same with landings on the moon.

For Earth return, landing would be on a pad, so no worries there.

You may be able to further improve by gimbaling the thrusters out a bit at the last second of landing so that debris will be diverted away

[sevenperforce]

The landing thrusters are set in the wing extensions, rather high off the ground...maybe 4 meters up? With the low air pressure on Mars, the plume is going to be really diffuse by the time it hits the ground, so debris impingement should be low or nonexistent. Same with landings on the moon.

For Earth return, landing would be on a pad, so no worries there.

You may be able to further improve by gimbaling the thrusters out a bit at the last second of landing so that debris will be diverted away

Oh, good idea. I was planning on having the thrusters fixed rather than gimballed, to reduce dry mass, since their placement allows total pitch and roll authority with differential thrust alone. But gimballing could also add yaw authority in addition to protecting the heat shield on landing.

Another possibility would be having the landing panels open up in the opposite direction, protecting the underside of the ship. Not quite as stable, though.

[sevenperforce]

Another option is a second stage for FH as a half scaled ITS. 6m diameter, 24m long, 1 raptor and 3 flavours:
1- Cargo version with integrated fairing (crocodile style) to deploy satellites
2- Mini crewed ITS launched as FH: minibus of 20 PAX to LEO, 10 PAX to moon vicinity and 5 to asteroids or even Phobos. Crewed part is 6m diameter, 9m long with around 180m3. Could copy-paste beefed up ECLSS systems from Dragon and become a very polyvalent spaceship.
3 - Tanker version

That would be the spaceship that NASA needs for BLEO and would match beautifully with the SLS. But I guess reminds too much the shuttle. Would delay too much the bigger brother, but a very good exploration spaceship.

Landing is a problem here. You'd need auxiliary thrusters for touchdown. And protecting that Raptor Vac would be tough.

[Jimmy Murdok]

Landing is a problem herd. You'd need auxiliary thrusters for touchdown. And protecting that Raptor Vac would be tough.

SL Raptor has around 300tons, throttle at advertised 20% and you are at 60 tones for a 2nd stage of FH + spaceship all in one and on a hoverslam. I guess should work.
Regarding the nozzle extension either disposable or retractable.
I don't know, the ITS full system is such a giant leap forward that we all look at logical intermediate steps to make it "chewable". My guess is that from Elon perspective, this might make sense as a good cost/benefit step, allow test and become a good solution itself, but I'm far from being convinced. In few weeks we might have a hint about if mid stage in the architecture. If so, Zubrin will breath

[dror]

Landing is a problem herd. You'd need auxiliary thrusters for touchdown. And protecting that Raptor Vac would be tough.

SL Raptor has around 300tons, throttle at advertised 20% and you are at 60 tones for a 2nd stage of FH + spaceship all in one and on a hoverslam. I guess should work.
Regarding the nozzle extension either disposable or retractable.
I don't know, the ITS full system is such a giant leap forward that we all look at logical intermediate steps to make it "chewable". My guess is that from Elon perspective, this might make sense as a good cost/benefit step, allow test and become a good solution itself, but I'm far from being convinced. In few weeks we might have a hint about if mid stage in the architecture. If so, Zubrin will breath

I fear that from Elon perspective the ITS is a chewable intermadiate step...

[dror]

This is obviously plenty of margin to have a crewed version, which would use the same tank and body as the rest of the orbiter but have a crew cabin in place of the cargo bay. Payload capacity is high enough that the crew cabin could carry at least a dozen crew members plus unpressurized cargo and still have independent LES and re-entry capability (lifeboat).

Can you please elaborate on the crew cabin?
Surely you have a sketch of it 
If you want it to have independent LES and re-entry capability then it is basically a Dragon2

[sevenperforce]

Landing is a problem herd. You'd need auxiliary thrusters for touchdown. And protecting that Raptor Vac would be tough.

SL Raptor has around 300tons, throttle at advertised 20% and you are at 60 tones for a 2nd stage of FH + spaceship all in one and on a hoverslam. I guess should work.
Regarding the nozzle extension either disposable or retractable.
I don't know, the ITS full system is such a giant leap forward that we all look at logical intermediate steps to make it "chewable". My guess is that from Elon perspective, this might make sense as a good cost/benefit step, allow test and become a good solution itself, but I'm far from being convinced. In few weeks we might have a hint about if mid stage in the architecture. If so, Zubrin will breath

I don't see jettisoned or retractable nozzle extensions as likely; either one would hurt the 380+ second specific impulse that the Raptor Vacuum can boast. So I'm pretty sure landing on the main engine is right out. 60 tonnes is a LOT of dry mass compared to the 4-tonne dry mass of the current upper stage. With any reasonable fuel fraction you'd be getting to the upper limits of what Falcon Heavy can lift. In order to maintain a launch TWR of 1.36:1, the lower limit for F9, the upper stage and payload cannot exceed 400 tonnes.

The reason I think a dev-Raptor-powered composite methalox upper stage is promising is that the Air Force helped fund Raptor based on the promise of a methalox upper stage. This would be a perfect test bed for composite staging, upper-stage reuse, and BLEO operation.

This is obviously plenty of margin to have a crewed version, which would use the same tank and body as the rest of the orbiter but have a crew cabin in place of the cargo bay. Payload capacity is high enough that the crew cabin could carry at least a dozen crew members plus unpressurized cargo and still have independent LES and re-entry capability (lifeboat).

Can you please elaborate on the crew cabin?
Surely you have a sketch of it 
If you want it to have independent LES and re-entry capability then it is basically a Dragon2

Yes, the idea was basically to have the same overall systems as a Dragon 2. That's what my dry mass considerations were based on, just for conservatism.

No sketch at present but I've attached outer and cutaway line drawings. Essentially a Dragon 2, but turned on its side. Re-entry is not particularly pretty and probably close to ballistic, but that's for emergencies only, so I'm not too worried.

As I've depicted it, there are two pressurized regions -- a "trunk" with its own docking port on top and a fair bit of volume for pressurized cargo, and the forward crew cabin. In a launch abort, solids between the fore and aft sections activate and push the upper section free by itself; recovery is via chute. As you can see, the heat shield on the overall craft doubles as the heat shield for the crew compartment if the stage is damaged and the crew must perform an emergency lifeboat re-entry. Unlike Dragon 2, crew-capsule-only landing is NOT nominal, so sacrificial elements are okay.

[GWH]

IMO it would be a more efficient use of resources to go with a blunt nosed stubby mini-ITS for launching Dragon on top. Then not going to the trouble of developing a fully integrated crew man rated crew cabin for what is essentially a stepping stone program with a short life.  Or just keep crew on a Merlin based upper stage and rendezvous in orbit.

Still would retain the in flight abort ability, independent reentry - so a suicide burn landing is acceptable.
Mars Ascent Vehicle would just be a dragon in a fairing where the fairing is ditched after landing on Mars.

I'd been meaning to post this strategy on Robotbeat's humans to Mars in 2020, I think a cargo MITS + Dragon 2 could be a workable architecture in the short term.

[sevenperforce]

IMO it would be a more efficient use of resources to go with a blunt nosed stubby mini-ITS for launching Dragon on top. Then not going to the trouble of developing a fully integrated crew man rated crew cabin for what is essentially a stepping stone program with a short life.  Or just keep crew on a Merlin based upper stage and rendezvous in orbit.

Still would retain the in flight abort ability, independent reentry - so a suicide burn landing is acceptable.
Mars Ascent Vehicle would just be a dragon in a fairing where the fairing is ditched after landing on Mars.

I'd been meaning to post this strategy on Robotbeat's humans to Mars in 2020, I think a cargo MITS + Dragon 2 could be a workable architecture in the short term.

While I think a reusable methalox upper stage/MITS is probable, I know this particular configuration is less so. That being said, it represents my own personal wish-fulfillment fantasy for a fully-reusable LEO crew shuttle.

The trouble, as always, is recovery mode. If your stage loses its front end, how does it shield itself from heat on re-entry? And while suicide burns may be possible without crew, landing on the vacuum engine(s) is not, so you still need landing thrusters. Biconic re-entry makes a lot of sense, so biconic landing should as well.

I'm not sure I am picturing your MAV proposal correctly....

[GWH]

The heat shield "stubby nose" would need to be under the Dragon trunk, and it may need an expendable stage adapter to smooth out the airflow. 
For the MAV, take your sketch of the hinged nose cone and place a dragon under the cone, with the trunk mounted in the upper half of the payload section.  Add a tapered skirt if necessary for aerodynamics, which is less of a concern as an MAV.

[GWH]

Landing is a problem here. You'd need auxiliary thrusters for touchdown. And protecting that Raptor Vac would be tough.
[/quote]

If the center core interstage was shortened significantly with the Raptor Vac being shoved up inside the MITS rather than dangling below, or at least wrapped in a 120+degree shield/flaps then re-entry damage wouldn't be a problem

[sevenperforce]

The heat shield "stubby nose" would need to be under the Dragon trunk, and it may need an expendable stage adapter to smooth out the airflow. 

For the MAV, take your sketch of the hinged nose cone and place a dragon under the cone, with the trunk mounted in the upper half of the payload section.  Add a tapered skirt if necessary for aerodynamics, which is less of a concern as an MAV.

Dragon doesn't have nearly enough dV to be an MAV, though.

The crewed version I've outlined above would be able to do a Mars EDL and return to orbit with any ratio better than 26.3% (that's including losses for the thruster burns), so if you send a tanker into Martian orbit ahead of (or along with) your lander, you can manage easily.