Concept of a multi core MCT launcher
The shuttle
With a multi core configuration, it becomes possible to design a vessel with an aerodynamic surface (therefore a lift) significant because it would be no longer necessary to lodge it at the top of the ship-launcher assembly. We already underlined at the beginning of § 1.1, the importance of this fact since the increase of the surface exposed to the aerodynamic flow allows reducing the ballistic coefficient, which is essential to improve braking and to reduce the heat stress during EDL phase. We will therefore place the shuttle parallel to the cores, enough towards the rear so as to ensure a centering allowing piloting in good stability conditions during the different regimes of atmospheric flight.
Thus rid of this constraint, we can increase the width of the shuttle body in the overall design and provide for further increasing the wing surface exposed to braking, while allowing, this time, a glided landing on Earth (without propulsion).
Of course, this leads to a significant penalty in terms of structural masses; shifting from a fuselage section of 12×12 m² to an elliptical fuselage section of 20×10 m², and lengthening the vessel by 5 m, we estimate that the structure masses would rise from 45 tons (in the previous design) to 60 tons (structures – tanks excluded, and thermal protection).
But we largely come out the winners in terms of ballistic coefficient which becomes twice smaller and close to that of the X33. It becomes conceivable to provide the MCT, as was the case for the X33, with a metallic thermal protection, a priori truly reusable without need of any heavy maintenance. Beware though: the MCT (interplanetary) being due to hit the Earth atmosphere, back from Mars, at a speed at least √2 times larger than the X33 (orbital), the energy to cope with during the re-entry process, will, a priori, be at least twice greater. Consequently, the return to Earth must be done in several steps, in order to break the heat load into several parts that are easier to hold off the ship and to dissipate by radiating the thermal protection. Namely, the return would start with a first phase of aerocapture limited to positioning the ship on a highly elliptical orbit; an aerobraking phase would then allow lowering the apogee until reaching a low orbit; finally, a classical EDL would start from that orbit. Still, at that moment, the maximum flows will still remain to be born; the most affected area would be the nose, the radius of which must be maximized; fortunately, this matches the wish and the need to increase the scale of the vehicle to reduce its ballistic coefficient. We chose 3.50m.
Warning
The thermal load and the heat flow matter remains a critical point that we could not study thoroughly. It remains to be checked whether the levels reached during the terrestrial return are manageable with a reusable (non-ablative) Thermal Protection (“TP”). If this were not the case, the concept could be jeopardized at the level of the return mode, and could require either a refurbishment of the TP between each flight or a partial propulsion braking before entering the atmosphere.
It is to be noted that this problem becomes even more acute in case of an immediate return along an « opposition » trajectory, chosen for the sake of a better return on investment. Indeed, in this case the terrestrial return speed could no longer be close to 11 km / s, but would rather reach 14 km / s. Even with a ballistic coefficient half that of the Space Shuttle, we may wonder whether it could be manageable?
Anyway we’ll look at this scenario in order to assess the consequences in terms of sizing of the vehicle and of its propulsion (thrust level, propellant mass for the outward and return legs) as well as the consequences in terms of performance requested from the launcher (or in terms of payload, the launcher remaining the same). We’ll therefore first describe the basic version (standard return after an 18 months’ stay on the surface of Mars) and, afterwards, the « immediate » return version.
Classic return version
In terms of overall design, here are the most notable changes compared to the single core version:
- larger dimensions of the ship in order to increase its braking surface during the EDL phase; we therefore increase the length from 39 to 44 m and give a quasi-elliptical section of 20 m width and 10 m height to the fuselage; a planform of 44 m wingspan further increases the cross section for braking, while allowing a glided landing onto Earth;
- resulting structural mass increase, compelling to reduce the payload mass down to 85 tons;
- installation of 6 engines, of same technology as for the launcher’s Raptor (380 s Isp) but with 70 tons of unit thrust, equipped with a deployable divergent extension (which solves the problem of isolating the jet from the surrounding structures: fuselage, engine hatch); this thrust level is necessary to take off from Mars, fully fueled, with an initial vertical acceleration of 2 m/s2; the 6 engines solution should enhance safety in case of failure of one of them during the critical phase, without possible escape, of the final EDL braking (being noted that the shuttle is much lighter then, than when it takes off from Mars, ie about 200 tons);
- splitting of the tanks, which become spherical; this choice facilitates the drawing of propellant in a transverse propulsion mode that we keep in this version; we adopt a crossed setting, methane – oxygen, which facilitates the feeding of the engines by reducing the length of the supply pipes.
The mass budget, prepared on the same basis as for the single core case is summarized in the following table:
| DATA | |
| Length (m) (without base hood) | 44 |
| Width (m) | 44 |
| Height (m) | 10 |
| Payload mass, outward leg (T) 1 | 85 |
| Payload mass, return leg (T) | 5 |
| Structural coefficient, tanks 2 | 0.040 |
| Prop. mass coefficient Margin | 1.030 |
| Complement Prop. Mass 3rd stage LEO (T) | 400 |
| MANEUVERS (km/s) | |
| ΔV outward leg (Hohmann, Mars at perihelion) | 3.46 |
| ΔV return leg (Hohmann, Mars at aphelion) | 6.52 |
| ΔV maneuvers Mars Orbit/td> | 0.15 |
| ΔV EDL braking & hovering | 0.60 |
| ΔV maneuvers, return to Earth 3 | 0.30 |
| MASS ITEMS (T) | |
| Equipments | 12.0 |
| Engines : 6 of 71 T thrust | 4.0 |
| Structures & Thermal Protection | 60.0 |
| Tanks | 34.6 |
| Dry Mass | 110.6 |
| STATE of TOTAL MASS (T) | |
| On LEO parking | 660.8 |
| During Earth-Mars transfer | 253.7 |
| At EDL Entry | 236.6 |
| At taking off from Mars | 734.1 |
| PROPELLANT MASS (T) | |
| Transfer injection E-M | 407.0 |
| Maneuvers on Mars orbit | 17.1 |
| EDL | 41.0 |
| Return trip | 613.5 |
| Maneuvers, Return Earth 3 | 10.0 |
| Tanks Capacity (fixed by the outward trip) | 865.1 |
| (1) We had to reduce the mass of the payload in order not to exceed the number of 60 engines. |
| (2) Coefficient slightly worse than in the case of a single core, for lack of a common intertank bulkhead. |
| (3) In this case, the ship glides into its final landing, avoiding propulsion braking. |
Features of the shuttle with a multi core launcher version
and a return trajectory in a « conjunction » mode.
Laying out scheme of a shuttle with a multi core launcher.
The required performance in LEO is therefore 661 tons (with a propelling contribution of the shuttle). Note that, as in the single core case, this result is achieved in the most favorable transfer conditions: a Hohmann transfer and Mars being at perihelion at the time of arrival. Yet, this high launcher’s performance is obtained after having reduced by 15 tons the targeted mass payload. This result depends also directly upon the actual mass of structures and TP, for which a 10% inaccuracy makes a 6 tons difference…
When, upon arrival, Mars is near its aphelion, the payload shrinks from 85 tons down to 70 tons.
The mass upon EDL entry is obviously larger than that of the smallest dimension single core ship (237 instead of 222 tons). But, as the aerodynamic surface is, at first approximation, about 2.1 times larger (44 x 44 / 2 versus 39 x 12), the ballistic coefficient dwindles to an estimated 306 kg/m², versus 585 for the Space Shuttle; we get close to the X33 cases (240), a result that allows to consider the implementation of a truly reusable TP without refurbishing. The conditions of re-entry into the Earth’s atmosphere, outlined above, are themselves eased by the surface increase. This gain in ballistic coefficient is the main argument in favor of a multi core launcher formula.
The tank capacity must be 865 tons; it is determined by the needs of the outward leg (including 400 tons allocated for the final setting up of the ship into LEO). This quantity is lodged in 4 spherical tanks inside the elliptical fuselage of a section 20 m x 10 m.
Embarquement pour Mars – Préface Thomas Pesquet
Webinar #2 APM sur STARSHIP disponible en replay
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A very interesting analysis. But I sincerely doubt SpaceX would go the route of sandwiching their passengers in amongst all that rocket fuel. So far everything that they have done has been safety first, re-usability second, cost third, and everything else following on from there.
A single core seems highly improbable, but multiple cores joined like Falcon Heavy (maybe even in a 5, or 9 core config) seems more likely. It maintains their philosophy of doing everything like a production line.
But even multi-core doesn’t preclude the placing of a large craft on the top.
And have you considered that landing of payload my not require the whole thing to return to earth? Payload capsules could be left behind (even maybe acting as shelters) which would reduce the return mass considerably.
I think you need to seriously consider the cost of the relative parts to see which bits they will return and which will get left behind. It’s not all or nothing here.
Anyway, fascinating analysis. Thanks.
hi,
*passengers safety: in any case, as there is no possibility to install a launch escape system, it could prove necessary to ferry the passengers with « taxis » after the mammoth launcher has placed the MCT in LEO.
*placing the craft on the top: as I wrote in the report this poses the dilemma of choosing between a low ballistic coefficient (to improve EDL and thermal conditions) and a lighter vessel (with no wings at all) ; I choose the first option, as I thought it was a condition for a really reusable Thermal Protection. I may be too pessimistic.
*not returning everything on Earth: many people think even not returning the shuttle on Earth and refueling it on LEO… In this case, we certainly come to a less gigantic launcher, but on the cost of multiple launchs and augmented operations complexity.
Thank you for your contribution.
Richard
You start with the assumption that the Mars landing craft must have a minimum habitable volume of 10m^3 per person. Dragon v2 has a habitable volume of 10m^3 for 7 people. Why would the Mars landing craft be any bigger than necessary to accomodate 100 people crammed together in their seats? This landing craft could be a capsule shape. I think the MCT architecture will go something like this… 100 person capsule is launched into LEO where it is docked with (3) BA-330’s and a Mars TMI stage. With a 6 month voyage to Mars, everyone will have about 10 m^3 of space. Once at Mars, the capsule separates and performs EDL. The BA-330’s head back to Earth on a free-return trajectory. A capsule on Mars could launch and rendevous with the BA-330’s if anyone wants to go home.
That makes a lot of sense. I think a permanent spacecraft (as in, a spacecraft that never reenters the atmosphere), is very likely. The use of BA330, if bigelow is successful, is almost a given. Isn’t 10m^3 a little cramped for a month long voyage ? I’m trying to remember how big a dragon is, and I wouldn’t like to live inside one for month. But space is a hard place, so it might do.
Also consider that this is 10m^3 of 3D space in microgravity. Every wall and ceiling is another potential floor. Take a look above you at the ceiling (assuming you’re indoors) and imagine what you could do with all that unused real estate if gravity weren’t in your way. With that in mind 10 m^3 might even be too much room.
10 m3 is not for EDL but for the long transfer journey. It’s very minimum ; 15 m3 preferable (and possible in this design).
Landing 100 pax / 100 T total P/L within a capsule is not feasible ; the b allistic coeff would be very high, with inefficient deceleration and very high thermal conditions. You need to present a large surface to the atmosphere. Furthermore, you need another different vehicle to come back !… The idea of this study was to have only one piece of HW (the shuttle) and to not throw away anything (fully reusable). That said, this principle leads to an over-sized launcher, and raises the question of the reusabillity period (52 months except if it can afford an « immediate » (opposition type) return.
Your proposal of a BA330 on a free return trajectory is seductive. But it looks like the cycler concept, which Elon Musk has explained why it will not be his choice.
You forgot about in orbit refueling,
as per Musks comment:
« I mean, if you do a densified liquid methalox rocket with on-orbit refueling, so like you load the spacecraft into orbit and then you send a whole bunch of refueling missions to fill up the tanks and you have the Mars colonial fleet – essentially – that gets built up during the time between Earth-Mars synchronizations, which occur every 26 months, then the fleet all departs at the optimal transfer point. »
You are righ ; I get a lot of commentaryabout that. But you have to understand that I wanted to look to this extreme solution of only one launcher. As you saw, the result is decouragingly gigantic. So yes, refueling could be a solution, but it would also be more difficult to get low travel price if you have a « bunch » of launches! Also, transefering cryogens in microgravity is not simple ; I would rather consider plugging interchangeable tanks.
Thanks!
Interchangeable tanks have several advantages; they avoid the complications of zero-g pumping, they could be collected in orbit while the expensive parts (the engines) of the tanker returns for more, and if removable at martian gravity, they could be trucked between the landing pad and the fuel factory, allowing these to be located safely apart from each other.
Richard,
I’ve seen a few attempts at modeling this before, but yours is particularly impressive and the first to show the effects of the MCT’s dry mass. This is particularly useful for me, as we have no lack of people at Nasaspaceflight.com convinced that the most logical thing to do is to fuse the second stage and the MCT lander together. The argument they make is based on economics and not mass optimization, and I believe more than a few would change their minds if they could see how heavy the MCT would need to be to make it back from Mars. You should mention to your translator, Pierre, that the term « ton » in North America means 2,000 pounds, not 1,000 kilograms. To avoid any confusion I suggest using the term « tonne », which is the correct term in all English-speaking countries for the unit you mean.
It was mentioned that your reason for this approach was the fact that “I wanted to look to this extreme solution of only one launcher”. While I appreciate the simplicity of single reusable launch to Mars, there are a number of issues besides the fact that Spacex mentioned in-orbit refueling.
–Explosion Risk: A 2,750 tonne Soviet N-1 moon launcher exploded seconds after lift-off, resulting in a 7 kiloton blast that was the largest non-nuclear explosion in human history. You propose a 13,700 tonne rocket, which would mean you would have to sell the public on allowing a rocket able to explode with a force over 2X that of the Fat Man atomic bomb to launch anywhere near them. A smaller rocket would attract less public opposition and be far less dangerous and costly in event of failure.
–Engine Count: The only LV in history to have 30 engines failed in all 4 of its launches, though obviously Spacex are working on a 27-engine rocket. You propose that the risk of a 61-engine first stage is acceptable. I think the risk of engine failure leading to a catastrophic failure is too high, even with the more benign nature of the full-flow staged combustion cycle. A more realistic alternative is a 27-Raptor first stage, which would still enable a 5200 tonne RLV capable of pulling off your Mars mission in 3 launches.
–Economics: Your rocket’s first stage is 20 meters in diameter while the second stage is 12 meters. This requires two sets of tooling, which would add enormous cost. The ideal solution would be two stages with the same diameter, which would let you build them with the same set of tooling. This would make your RLV much less expensive.
–Lack of LAS: If you are going to launch 80-100 people at a time, chances are the US government will insist on there being some way to save the crew if the rocket were to explode at launch. You should probably add LAS engines to reflect this reality.
One possible reason why Spacex’s figures are so hard to attain is because you have used an entirely conventional approach. You probably did not hear that Spacex are considering solar electric propulsion in addition to the chemical engines to up payload and trim transit time. Some calculations on our site showed that the MCT could plausibly generate more than twice the power of the ISS, which would make this a viable option. Hope that helps and thanks again for the superb work!
Why assume the lander and rocket would transport 100 at a time? I imagine a large space station that you could attach many landers to and load it up over many launches to fill this colossal space station and use it as a back and forth transport ship
I heard Musk say that they will probably use a vehicle in constant transit within the Earth-Mars system (transporter) while using tanker launches to refuel the cycler. Can`t remember the source.
I haven`t heard anything regarding EDL but their current interest in hypersonic reentry profiles would indicate that whatever tech they use to land, will not be centered around inflatable decelerators nor cranes (due to high mass of the vehicles)
I guess we`ll have to hold our breath until September 2016 when the architecture is unveiled.
Just happened across your site and I really enjoyed reading your analysis and proposals.
My own opinions are :
1. Booster necessary to lift heavy hardware to Earth orbit.
2. Earth – Mars (maybe Moon use too) Ferry, the MCT
3. Mars SSTO for transfer of material and people between orbit/ground. CH4+O2 engines
4. Nuclear plant on Mars for in situ electrical and CH4 generation.
5. Some kind of reusable entry skin/bladder/container with engines that once the payload is extracted it is folded up and shipped back to Earth for refilling. Very low volume/mass.
Also the first few rockets would carry a constellation of combined Weather, Communication and GPS satellites to place in orbit around Mars.
With the payloads being talked about the MCT could probably deliver all this in once shot.
Investigate a Phoboa and Diemos tankage / rendezvous location as well.