Brachistochrone to Mars with Falcon 9 boosters

[sevenperforce]

If SpaceX gets its booster return down to a science, and Falcon Heavy performs as expected, then an interesting possibility emerges.

The Falcon 9 v1.1 FT Stage 1 booster is capable of SSTO on its own, though without payload or capacity for return. If a Falcon Heavy was launched without any second stage, however, you'd end up with a nearly-full first stage in orbit and two empty strap-on boosters returned safely to the ground, ready to refuel and relaunch. A single Falcon 9 launch costs $61 million, with fuel accounting for roughly $200,000 of that. Thus, Falcon Heavy would allow SpaceX to put a nearly-full Falcon 9 first stage into LEO for marginally more than the cost of a single Falcon 9 launch.

With a $1 billion investment, that would be no less than fifteen nearly-full Falcon 9 first stages in LEO.

Strap them together and you've got a launch stack capable of a Brachistochrone transfer to Mars for a manned mission in a minute fraction of the Hohmann transfer time. A short transfer time means your consumables budget can be much smaller, enabling an even-faster transfer. Can't think of a cheaper way of doing it.

[Robotengineer]

SpaceX is already developing the MCT and BFR, so anything using F9's to get to Mars is off the table. Falcon 9 Lox boil off would be an issue, though I may be wrong about that. We don't know how much the cost savings are going to be for the reusable F9's, but it won't get down to just the fuel cost. 

[p1t1o]

Your thread title reminded me of an Asimov book, "Thiotimoline to the Stars" and was kinda expecting something more weird (I had no idea what a Brachistochrone was).

Your financial math seems very optimistic, but even multiplied by 10, is still a plausible figure.

I don't think finances are the only obstacle to these missions though.

[AngelLestat]

but these booster are not designed to be joint in orbit. A much better approach would be using falcon heavy in reusable mode, then rise all the pieces to make a big tug up there (lego way).

[sevenperforce]

Elon Musk has said SpaceX's likely target date for a manned Mars mission is 2025. I just like speculating about how you could manage it with current vehicles.

You could manage the landing by bringing two descent vehicles along: a full Falcon 9 Stage 1 booster with an empty Dragon V2 already mounted on top plus a separate Dragon V2. After reaching low Martian orbit, the booster would break off first and execute an autonomous powered entry and descent to a tail-first landing. Once it could be confirmed that the booster had landed successfully and had enough fuel for a powered ascent, the astronauts would make an aerobraking-to-powered descent in the separate Dragon V2. They'd hang out, collect soil samples, plant flags, and then transfer to the booster, leaving the descent vehicle behind.

Quote

but these booster are not designed to be joint in orbit. A much better approach would be using falcon heavy in reusable mode, then rise all the pieces to make a big tug up there (lego way).

I'm not quite sure how this differs from what I'm proposing...can you elaborate?

Edited January 29, 2016 by sevenperforce

[AngelLestat]

58 minutes ago, sevenperforce said:

I'm not quite sure how this differs from what I'm proposing...can you elaborate?

How do you:  "Strap them together and you've got a launch stack"?  If they was not designed for that?

If you design each part of the special tug (no falcon9 parts) to be able of merge one with the other in a controlled way with a small engine (why you need 9 merlin engines for each booster with a total of 450 merlin engines which increase the chance of failure?) then you have a tug that can be build depending any deltav and cargo you need.

PD: by the way.. I dont see much point in waste a lot more in deltav to save some  consumables which increasing the risk a lot due extra deltav.

Edited January 29, 2016 by AngelLestat

[Gaarst]

3 hours ago, sevenperforce said:

The Falcon 9 v1.1 FT Stage 1 booster is capable of SSTO on its own, though without payload or capacity for return. If a Falcon Heavy was launched without any second stage, however, you'd end up with a nearly-full first stage in orbit and two empty strap-on boosters returned safely to the ground, ready to refuel and relaunch.

I don't follow you there: putting a 400t heavy first stage in orbit kinda counts for a payload. Putting an empty 20t stage to orbit is doable, as long as you use the fuel in it.
A Falcon 9 cannot put 200t in orbit, so 2 Falcon 9s strapped together cannot put a 400t stage sitting in between them in orbit.

Also, you know that reentry heat thing up right ?
Recovering a stage from a slow suborbital flight is doable.
Recovering a stage from an orbital flight, which subsequently reentered at 7.8 km/s is a bit less doable, unless you shield your stage with a lot of ablative protection and make it a lot stronger to withstand the aerodynamic forces that would really like to rip it off; then you might have a chance to recover it, but only because it would be too heavy to lift itself from the pad.

[sevenperforce]

These are Falcon Heavy components, so they're already designed to be strapped together. And sure, that's a lot of engines, but you needed those engines to lift the fuel into space, so they're kind of already there. Plus, the strap-on boosters are themselves only good for a limited number of launches, so this would be a way of "recycling" boosters that won't last beyond a couple more launches.

I suppose you could lift full second-stage engines into space even more cheaply, though. I don't know how the cost works out...whether the cost per m/s of dV is lower if you're launching nearly-full First Stages with 2/3 booster reuse or completely full Second Stages with 3/3 booster reuse. In the former case you do get vacuum-specific engines, which means a higher Isp.

Using a modular approach where each component of the transfer vehicle stack has its own fuel supply and engine offers a lot more redundancy than trying to launch fuel and engines separately.

Finally, increasing dV saves more than consumables; it saves development time. This is something we could do now, with current tech. No need to design a multi-month hab.

6 minutes ago, Gaarst said:

I don't follow you there: putting a 400t heavy first stage in orbit kinda counts for a payload. Putting an empty 20t stage to orbit is doable, as long as you use the fuel in it.
A Falcon 9 cannot put 200t in orbit, so 2 Falcon 9s strapped together cannot put a 400t stage sitting in between them in orbit.

No, but with propellant crossfeed, the core booster from a Falcon Heavy can have enough remaining fuel at staging to reach orbit in a nearly-full state.

[Gaarst]

4 minutes ago, sevenperforce said:

No, but with propellant crossfeed, the core booster from a Falcon Heavy can have enough remaining fuel at staging to reach orbit in a nearly-full state.

No.

Crossfeed is not magic. A payload fraction of 33% is.

I think you got a bit confused by the fact that the Falcon 9 first stage could SSTO. That means it can put its own mass in orbit, and maybe an astronaut sitting on top but that's it, and using all of its fuel. Even when not considering TWR needs, two F9 stages cannot carry a third one into orbit, unless the third uses all its fuel: carrying unspent fuel decreases dV because you don't use it, and decreases dV because it makes your payload heavier.

A good thing to realise your mistake is to think about other rockets: the F9 first stage is not the first first stage to have SSTO capabilities. Actually it is quite common that first stages have over 9 km/s of dV when on their own, but they have to carry a whole rocket, so they don't SSTO. If strapping 3 core stages was enough to put entire stages in orbit, we would probably have reached Neptune with Delta-IV Heavys by now.

[sgt_flyer]

you won't be able to perform a brachistochrone transfer with kerolox to mars - the ISP is simply not high enough by severals orders of magnitude.

someone has tried to create a table (a bit difficult to read, and the value includes the takeoff from earth) but you're still looking at 370000 m/s of delta-v for a 0,01g brachistochrone transfer... (other sites speaks of around 375km/s of deltav for the same transfer, so his values seems coherent). there's also other kinds of transfers, but without at least an engine of the class of the orion drive (if not more efficient) you can't do brachistochrone.

http://www.projectrho.com/public_html/rocket/appmissiontable.php

 

Edited January 29, 2016 by sgt_flyer

[AngelLestat]

Quote

These are Falcon Heavy components, so they're already designed to be strapped together

no, they are not.  You can put together stages inside a spacex assembly line with lot of machinery and workers, no in orbit by themselves.  

Gaarst also points something very important, not sure how I miss it..  you can not place a first stage in orbit with full fuel.
Like I said, is possible to design a tug with many parts if they are designed with that purpose, but it has also its drawbacks vs a specialized launcher as the MCT.

[sevenperforce]

1 hour ago, Gaarst said:

Crossfeed is not magic. A payload fraction of 33% is.

I think you got a bit confused by the fact that the Falcon 9 first stage could SSTO. That means it can put its own mass in orbit, and maybe an astronaut sitting on top but that's it, and using all of its fuel. Even when not considering TWR needs, two F9 stages cannot carry a third one into orbit, unless the third uses all its fuel: carrying unspent fuel decreases dV because you don't use it, and decreases dV because it makes your payload heavier.

Hmm, let's see.

The Falcon 9v1.1 first stage has a dry mass of roughly 25.6 tonnes in its reusable configuration and carries 395.7 tonnes of fuel. I don't know exactly how much fuel needs to be reserved for landing, but Elon said that landing on a barge allows staging at 2.5 km/s. Assuming first-stage gravity losses are on the order of 0.5 km/s, and the typical launch mass is 541.3 tonnes with specific impulse of 282 s, the rocket equation says that total mass at staging will be 183 tonnes, for a fuel consumption of 358 tonnes. This suggests 37.7 tonnes of fuel must be reserved for landing.

So if three of these are launched together without any upper stage and full crossfeed (not exactly right, but good enough for an estimate), the launch mass is (395.7t + 25.6t)*3 or 1263.9 tonnes. Staging takes place after 716 tonnes of fuel have been consumed, corresponding to a dV of 2.3 km/s. Since there are three Falcon 9 cores firing simultaneously, gravity drag will be lower -- about 0.2 km/s -- so the true dV is 2.1 km/s, leaving 5.7 km/s of dV to get up to a parking orbit. At this point, gravity drag and aerodynamic drag will be negligible, and the higher vacuum Isp of 311 s can be used, so you'll need to burn...

...356.4 tonnes of propellant to get to LEO. Leaving you with less than 40 tonnes of propellant. That's WAY less than I expected.

Well, that's disappointing. Looks like you were right.

What if the upper stage is launched into LEO without payload using a 2/3-reusable Falcon Heavy? The upper stage has an inert mass of 3.9 tonnes and carries 92.7 tonnes of fuel, for a total mass of 96.6 tonnes. Launch mass will be 1361 tonnes; the same math as before gives a booster separation at 1.7 km/s (allowing 0.3 km/s of gravity/aero drag losses). At this point, the vehicle has a mass of 518 tonnes and 395.7 tonnes of fuel in the core stage, with a vacuum Isp of 311 s, for a core-stage dV of 4.4 km/s. At core MECO, the craft will be traveling at 6.1 km/s, so the upper stage will need to burn 36.7 tonnes of propellant for final orbital insertion, leaving it with 56 tonnes of propellant.

That's an improvement, I suppose, but not much of one.

 

1 hour ago, sgt_flyer said:

you won't be able to perform a brachistochrone transfer with kerolox to mars - the ISP is simply not high enough by severals orders of magnitude.

someone has tried to create a table (a bit difficult to read, and the value includes the takeoff from earth) but you're still looking at 370000 m/s of delta-v for a 0,01g brachistochrone transfer... (other sites speaks of around 375km/s of deltav for the same transfer, so his values seems coherent). there's also other kinds of transfers, but without at least an engine of the class of the orion drive (if not more efficient) you can't do brachistochrone.

It would have to be a partial brachistochrone -- burn for a while, then coast, then turn around and burn in the opposite direction. Still potentially an order of magnitude faster than a Hohmann transfer.

[Nuke]

boosters generally dont work well for transfers. they are great for getting to orbit but once there you want a more optimized engine that gives you good isp and uses fuel that wont boil off on you during the voyage. before ion propulsion became a thing most probes just used hydrazine or something like it. apollo service module, and lm ascent and decent stages all used aerozine-50 and n204 and had pretty decent isp around 300+ (and being hypergolic simplified a lot of things). if i wanted to do a brachistochrone trajectory i would want something like that so you can be all tank and a small simple and light engine. not sure what the feasibility of brachistochrone is with ion propulsion.

now you really dont need to go full brachistochrone either, since both the target and return destination have atmospheres, use aerobreaking to save fuel instead of a deceleration burn. this places limits on how fast you can come in but saves a lot of fuel mass. in ksp i would just resort to rather violent aerocapture (ignoring the fact that you need to get uncomfortably low for this to work), jeb seems to think its a good idea.

Edited January 29, 2016 by Nuke

[sgt_flyer]

even without going to full brachistochrone, the dV you'll be able to get from a chemical transfer stage will have a negligible effect on the transit time.

according to the link i've posted above, the fastest non hyperbolic (ie, not on a sun escape trajectory) transfer possible to mars, with a deceleration prior to arriving to mars (and then the same to get back to earth), still makes for a 2 month round trip, and costs around 40km/s of delta-v... (as the values he wrote in the tables are calculated with the d/v required during the launch from the ground inside it).

using on orbit staging, one of the realistic proposals was for a 245 days round trip (with up to 15km/s dV needed for the round trip - and that was using a LH2/lox engine to get higher ISPs, + a 4 stage design to have a reasonable payload fraction). so to get down from an 8 months fast round trip to a 2 month faster round trip, you need 25km/s more dV - unrealistic for chemical engines, even with staging...

http://www.lpi.usra.edu/meetings/marsconcepts2012/pdf/4181.pdf

 

Edited January 30, 2016 by sgt_flyer

[sevenperforce]

I just realized that all those calculations I just performed -- determining that Falcon Heavy's upper stage would be left with 56 tonnes of fuel if it launched without a payload -- could have been avoided if I had merely looked up Falcon Heavy's launch specs and noticed that it is capable of putting a 53-tonne payload in LEO. Oops.

In any case, this means that if SpaceX were to launch Falcon 9 LEO payloads (10-13 tonnes) using the Falcon Heavy instead, it would cost them around $30 million extra per launch, and leave them with 40+ tonnes of fuel in LEO.

The brachistrochrone is far more advantageous on the return trip than it is on an outgoing trip. Dragon V2's integrated ballast sled allows for precision attitude and lift control on re-entry, which (if I understand correctly) gives it unprecedented aerobraking capabilities; it would be able to execute multiple passes through the atmosphere to bleed off a high transfer speed. Thus, the cost of a return brachistrochrone is only 50-60% the cost of an outgoing brachistrochrone. Plus, while the outgoing trip requires a Mars Ascent vehicle, the return trip does not, allowing the same amount of delta-V to be achieved with far less fuel. The Dragon V2 has a manned persistence of something like two weeks, IIRC....

As mentioned above, I think the cheapest manned Martian landing would involve two separate descents: first the unmanned ascent vehicle with a propulsive landing, followed by the manned Dragon V2 on an aerobraking trajectory once the ascent vehicle had made a confirmed landing. With the Apollo missions, it didn't make sense to do two separate landings because there was no way to aerobrake and the ascent delta-V is modest. But splitting them up for a Martian landing would be a significant delta-V savings.

The ascent vehicle would carry the Earth-return Dragon V2, but it would only need enough dv to get up to Low Mars Orbit. At liftoff, the manned Dragon V2 would mass 8,898 kg, plus a standard Falcon-9 upperstage dry mass of 3.9 tonnes, meaning about 30 tonnes of fuel is needed to reach LMO. By this time, propulsive Falcon-9 stage 1 landings will likely have been mastered, so an upper stage retrofit with landings legs and grid fins should be able to make the descent with the remaining 719 m/s of dV. The full-thrust Falcon 9 upper stage is capable of 934 kN at max thrust, allowing nearly 4.8 martian-gees of acceleration: definitely enough.

So the overall mission would look like this:

1. Accumulate upper stages in LEO using Falcon Heavy for Falcon 9 missions with about 40 tonnes of remaining fuel ($30 million each).
2. Launch self-contained high-thrust ion engine with attached power supply and fuel reserves with Falcon Heavy ($90 million for launch + cost of ion engine assembly and fuel).
3. Launch Mars Ascent Vehicle in the form of a Falcon 9 upper stage retrofit with landing legs and attached Dragon V2 (for return transfer and Earth descent) with Falcon Heavy ($90 million for launch + cost of Dragon V2).
4. Launch inflatable transfer hab, Transfer Vehicle Frame, and manned Dragon V2 (for Martian descent) with Falcon Heavy ($90 million for launch + cost of hab, frame, and Dragon V2).
5. Inflate hab, Earth Orbit Rendezvous, assemble transfer vehicle via EVA. (Abort mode: rendezvous with ISS or use Dragon to re-enter.)
6. Use chemical rockets for Hohmann transfer to Mars Capture Orbit; loop around Deimos to enter elliptical orbit around Mars.
7. Mars Ascent Vehicle breaks off along with single Falcon 9 upperstage; upper stage ferries it to a minimum-dV aerobraking trajectory and burns up; Mars Ascent Vehicle descends to a propulsive landing. (Abort mode: skip to step 11 using Dragon V2 used for launch.)
8. Once MAV landing is confirmed, crew enters the Dragon V2 used for launch, uses small burn to cross Deimos/Phobos and enter aerobraking trajectory, use repeated aerobraking passes to descend to a propulsive landing less than 1 km from MAV.
9. Jettison hab, ion engine ferries remaining transfer stages to low martian orbit.
10. After manned EVA on Mars, crew moves to MAV for ascent; rendezvous with ion engine and remaining transfer stages in LMO.
11. Transfer stages fire to place craft on Earth return trajectory; ion engine kicks in to start pushing on a low-thrust one-way brachistochrone to Earth.
12. On Earth approach, jettison ion engine and use repeated aerobraking passes in Dragon V2 to land.

Doable? Maybe.

Edited February 1, 2016 by sevenperforce

[Nibb31]

On 1/29/2016 at 4:40 PM, sevenperforce said:

The Falcon 9 v1.1 FT Stage 1 booster is capable of SSTO on its own, though without payload or capacity for return. If a Falcon Heavy was launched without any second stage, however, you'd end up with a nearly-full first stage in orbit and two empty strap-on boosters returned safely to the ground, ready to refuel and relaunch.

No. You'd end up with a dead core stage in orbit and two boosters burning up in the atmosphere.

The F9 Core stage is not a spacecraft. It's not designed to go to orbit or to reenter from (near) orbital speed. It doesn't have power, avionics, insulation, or attitude control to do anything useful in space.

Edited February 1, 2016 by Nibb31

[sevenperforce]

31 minutes ago, Nibb31 said:

No. You'd end up with a dead core stage in orbit and two boosters burning up in the atmosphere.

The F9 Core stage is not a spacecraft. It's not designed to go to orbit or to reenter from (near) orbital speed. It doesn't have power, avionics, insulation, or attitude control to do anything useful in space.

Yeah, ran the numbers above and figured that out.

Adjusted mission profile is here.

[PB666]

On 1/29/2016 at 10:04 AM, Robotengineer said:

On 1/29/2016 at 9:40 AM, sevenperforce said:

On 1/30/2016 at 7:52 PM, sgt_flyer said:

 

 

 

On 1/30/2016 at 7:52 PM, sgt_flyer said:

 

 

40 kv, lol.

The 29,000 base dV he gives is not the most efficient. he is basically establishing a minimal escape (he doesn't care about time) orbit after which he establishes a mars transfer orbit. Its better to get all your transfer velocity at minimal earth perigee. In fact the horizons space craft did its escape burn before achieve a stable earth orbit. This is because most craft kill engines to mount the apogee before circularizing, but if the apogee trajectory is in line with the desired escape orbit you do not have to waste as much gravity versus time (hoovering, very minute at this point in the flight) losses simply burn an escape vector along the prograde.

Why are we seriously discussing these high g-force accel and decel from orbit, we don' have the energy, we have nothing even close to the energy. To stack rockets you have to essentially add 4 times the mass in each subsequent stage, so you could not simply stack rockets, you would have to create a pyramid structure.

The argument here is not simply getting to Mars, we've done that. In the 3 year period you have to keep humans alive and return them, you have a mass, the longer the journey the more that mass increases.

Lets get a little dose of reality, the voyagers are moving at something like 20km/s alot of the v came from oberth effects and body hopping across the solar system. They are light weight spacecraft. We are talking here about manned space station like objects so the mass is a to two magnitude higher, and then talking about direct dV a couple of times what these craft had. This is something like 20 to 100 times more N*t than any craft that we have managed to produce.

 

 

 

 

[sgt_flyer]

@PB666be careful the way he worded it, he added the delta-V from earth ground to space to the total delta-V needed for the round trip . (heck easy to check with the moon in his tables, 16kdV for a hohmann transfer to the moon orbit, and then come back ^^) - so you have to substract the earth ascent dV ^^ - and it's the value for a round trip anyway.

 

but yes, even removing the dV from earth ascent from his calculations, making an accelerated trip to mars is going to be too much for chemical engines (or as you said, need a huge pyramid ^^).

even the more reasonable paper scenario i've linked above (http://www.lpi.usra.edu/meetings/marsconcepts2012/pdf/4181.pdf) seems over the top for chemical engines (heck, 19km/s needed with a space station sized spacecraft on chemical engines... you'd need the spacecraft to be a tiny fraction of the total weight for chemicals...)

Edited February 1, 2016 by sgt_flyer

[sevenperforce]

Making an accelerated trip to Mars will be too much for chemical propulsion, but making an accelerated trip back might work a little better.

If you have a 53-tonne budget for an self-contained ion thruster to use as a Mars return accelerator, what's the best choice? It will be a fine balance of thrust-to-weight ratio, power-to-weight ratio, and Isp...

If you can't accelerate at least one leg of the trip, then you have to build a very large and complicated hab, in which case it only makes sense to do this. If you aren't going to build a dedicated, persistent transfer hab, then it makes a lot more sense to jettison the hab at Mars, accelerate on your return trip, and aerobrake in.

[fredinno]

4 hours ago, sevenperforce said:

Use chemical rockets for Hohmann transfer to Mars Capture Orbit; loop around Deimos to enter elliptical orbit around Mars.

Deimos is too small to do a reasonable gravity assist with.

2 hours ago, sevenperforce said:

Making an accelerated trip to Mars will be too much for chemical propulsion, but making an accelerated trip back might work a little better.

If you have a 53-tonne budget for an self-contained ion thruster to use as a Mars return accelerator, what's the best choice? It will be a fine balance of thrust-to-weight ratio, power-to-weight ratio, and Isp...

If you can't accelerate at least one leg of the trip, then you have to build a very large and complicated hab, in which case it only makes sense to do this. If you aren't going to build a dedicated, persistent transfer hab, then it makes a lot more sense to jettison the hab at Mars, accelerate on your return trip, and aerobrake in.

Even an accelerated trip back is too much. You should never do accelerated trips when you first go to Mars, it's ridiculous.