Explaining the Infeasability of Second Stage Reuse

In an interview in MIT in October 2014 [1], Musk said:

“I don’t expect SpaceX’s Falcon line to have a reusable upper stage. With a kerosene based system, the specific impulse isn’t really high enough to do that, and a lot of the missions we do for commercial satellite deployment are geostationary missions. So we’re really going very far out – these are high delta-velocity missions, so to try and get something back from that is really difficult.”

Let’s see if we can explain why geostationary transfer orbit (GTO) missions are the limiting factor in developing second stage reusablity, as Musk says. We start with SpaceX’s description of their capabilities, from their website [2] [3]:


The thing about second stage reuse is that you have to propel the whole of the second stage to orbit, so any extra mass you add to the second stage has to come directly out of the payload. Knowing this, you can immediately see that F9 GTO missions will be by far the most affected – adding two metric tons to the second stage would be a 40% payload loss for this kind of mission, while only being an 16% loss for F9 LEO, a 10% loss for FH GTO, and a 4% loss for FH LEO.

Even without knowing, even to within an order of magnitude, how much mass reusability would add, we can see why Musk says what he says. A third of Falcon 9’s flights so far have been to GTO, and in the past few years, about half of the Atlas V’s, and almost all the Ariane 5’s launches have been to that (or higher) orbit. And already some of F9’s missions have pushed the launch system’s GTO capability to its limit: AsiaSAT 6, AsiaSAT 8, TurkmenAlem52E/MonacoSAT were all Falcon 9 GTO missions which went within a few hundred kilograms of their upper limit. SES-9 (next in manifest at time of writing) is 5300 kg to GTO, which is actually larger than SpaceX’s claimed capability.

For these, even a small increase in the mass of the second stage would have meant the Falcon 9 could no longer launch them. Second stage reuse would thus cost SpaceX significant numbers of GTO launches, until the Falcon Heavy can fly regularly enough to take over that market.

But the question remains – how much mass would a reusability system on the second stage really add? Would it be 100kg, one ton, or ten tons? Obviously as outsiders, we will probably never know an exact number, but as with many things it’s still possible to estimate an order of magnitude for it.

To do this we first need to consider how SpaceX would have implemented stage two reusability if they were pursuing it. Obviously, this is a hypothetical, but SpaceX did make plans to reuse the upper stage some years ago. In a flight animation from 2011 [4], SpaceX show the stage with a heat-shield on the (blunt) nose. Since this is the only configuration with any known SpaceX endorsement, it seems sensible to base our estimates on it.

So what would be main contributions to the mass in this situation? You would need:

a heat shield for the nose
less powerful heat protection for the body
landing legs
an RCS control system
grid fins
extra fuel for the landing
extremely good avionics / deeper throttle
probably some other things.

Not all of these are possible to estimate, but some are.

Heat shielding (nose + body):

Assuming they use the same material, PICA, for stage 2 (S2) as they do for Dragon, we can estimate approximately what the heat shield would weigh. According to NASA, describing a concept for the Orion heat shield [5]:

“A 5.5-m, 28-deg sidewall concept with a total mass of approximately 11,400 kg requires an aft TPS [thermal protection system] mass of 630 kg and forward TPS mass of 180 kg. The assumed TPS materials for this analysis were PICA for the aft side…”.

The Falcon 9 has a smaller diameter – accounting for this gives a heatshield mass of 270kg. However, it also has a much larger surface area exposed to secondary heating – 247 square meters for F9S2, compared to 27 square meters on the Orion capsule. Again rescaling the NASA value, this gives 1,650kg of secondary thermal protection for the upper stage. This calculation is imperfect – Orion has humans inside, while S2 has residual fuel. SpaceX may be able to use a lighter material, but Orion doesn’t have an engine bell to think about; these aren’t precise values, merely order-of-magnitude estimates. I’ve left the numbers in a precise form because it’s correct to add first and round later.

Landing Legs, RCS & grid fins:

On the first stage, the mass difference between a reusable and non-resuable configuration is estimated [6] to be 2,500kg (25,600kg – 23,100kg between v1.1 and F9R). Now the dry mass of the second stage is estimated [7] to be 4,000 kg, compared to 22,200 for the v1.2 first stage, 5.55 times smaller. So a very minimum estimate for the mass would be 450kg, assuming that all three can be reduced in exact proportion to the mass, since: landing legs must carry 5.55 times less weight and cold gas thrusters and grid fins must push on something with 5.55 times less inertia. The true value will probably be higher than this, because some things (plumbing, maybe) would scale less well than that.

Extra fuel:

Again, we can naïvely suppose that this scales with the inertial mass of the booster. In practice, it will likely be higher because the second stage would have a more difficult flight profile than the first stage, even after most of the speed is taken up by re-entry heating. The first stage separates weighing 54.0 metric tonnes due to fuel remaining (of its original 400 tonnes) for a RTLS landing [8] and 34.6 metric tonnes for a ASDS landing [8]. Again rescaling to dry mass (to ensure the same delta-v as the first stage), the second stage would now weigh 6.23 metric tonnes for an ASDS profile like the first stage assuming approximately the same efficiency. M1D is more efficient in space, less efficient in the atmosphere, so I’m assuming that these cancel each other out, approximately. That’s why these calculations are approximate. We now have effectively 2,230kg of dead weight added to the rocket as landing fuel.

So our overall estimate is 4,600kg. This estimate is very approximate. But it tells us
some important things:

Making the second stage reusable using only 1,000kg extra seems very implausible. Adding even this amount of mass would have made SpaceX incapable of launching any of the previously mentioned GTO satellites. They might still be able to resupply the ISS, but their commercial manifest would be severely hampered.

Reusability would not be expected to use more than maybe 17,000kg, unless the propulsive landing turns out to be way more complicated than anyone thought. This payload penalty would leave the Falcon 9 worthless, but the Falcon Heavy would still
have plenty of low Earth orbit (LEO) capability, and about the same GTO payload as the Falcon 9 has today.

Second stage reusablity is impractical today because of the number of GTO missions on SpaceX’s manifest (GTO missions are hit hardest by the payload penalty), and because the Falcon Heavy isn’t mature enough take over on these flights, yet.


[1] https://www.youtube.com/watch?v=y13jbl7ASxY&t=14m20s
[2] http://www.spacex.com/falcon9
[3] http://www.spacex.com/falcon-heavy
[4] https://www.youtube.com/watch?v=sSF81yjVbJE
[5] https://www.nasa.gov/pdf/695726main_ComingHome-ebook.pdf page 262
[6] http://spaceflight101.com/spacerockets/falcon-9-v1-1-f9r/
[7] http://spaceflight101.com/spacerockets/falcon-9-ft/
[8] https://www.reddit.com/r/spacex/wiki/faq/reusability Note that you have to reverse-engineer the rocket equation to get this result from a “15% drop in performance” – this is left as an exercise to the reader.


14 thoughts on “Explaining the Infeasability of Second Stage Reuse

  1. Neat high level analysis. Hopefully we’ll see second stage reusability with the FH lineup down the road (cuz awesome). I’m curious how long does it take you to write each post?

    Thanks for writing.


  2. I’m sure we all hope that. Hopefully the FH can escape its position of being permanently six months from launch.

    This one took around 4 hours, which is fairly typical. I would never be able to write at this level if I didn’t read the SpaceX subreddit (reddit.com/r/spacex) on a regular basis, though.

    Liked by 1 person

  3. Hello, just a question… how about a reusable second stage engine module that keeps in orbit an receives the fuel and payload for each mision. So you have:

    1 – Full launch configuration ( 1st stage + 2nd stage engine module + 2nd stage fuel module + payload)
    N – Lighter configuration ( 1st stage + 2nd stage fuel module + payload).

    It’s a spacetug-like configuration…. but limited to a smaller number of reuses. What do you think ?


    1. The biggest problem with this approach is that the orbiting tug wouldn’t be able to easily rendezvous with the launch system after 1st stage separation. Stage separation only takes place at about 2 km/s, while orbital velocity is 8 km/s, so even if you got the tug and launch system in the same location, you wouldn’t be able to dock because one would be moving at 6 km/s relative to the other (6 km/s is about 13000 mph). Even if you could slow the engine module down with a big heat shield (and somehow end up in the right location), you would have a matter of minutes to connect the two systems before gravity brings them both crashing down to Earth.


  4. The big assumption made here – by the above authors and Mr. Musk himself – is that reuse requires a return to the Earth’s surface. This is not so. There are a number of conceivable uses for the second stage in LEO: Habitable volume, interplanetary craft modules, shielding mass, counterweights, generic raw materials, propulsion mass, reuse when refuelled with non-terrestrial fuels etc. The fuel/mass calculations are much more favourable, particularly for ISS resupply. I suspect the main issue is that nobody wants the liability for hanging on to the things.


    1. Funny. I thought about this too. The second stage could rendezvous with a Depot which could refuel it and then return home. Could even pickup supplies from a manned space station.


  5. What if future second stages were only given the resources needed to go to an orbit where you could have a large collector spacecraft of sorts that robotically docks,stores and refuels multiple second stages. For future space flights you can sell access to reusable second stages already in orbit at a lower costs than the current option, or new propulsion option altogether to drive revenue growth?

    Liked by 1 person

  6. Love this blog and as an engineer I appreciate the level of effort you put into this page.

    I’m wondering if you would consider this analytical approach to determining how many times the first stage could be reused and how much maintenance would be needed, then backing out to determine how cheap space flight could get.


  7. It’s wrong to scale 1st stage fuel data because the first stage has only minimal heat shielding and most of that fuel is used outside the atmosphere specifically to prevent overheat during re-entry. If we have a heatshield this is no longer necessary since the heatshield has order of magnitude better TWR than an engine used for deceleration. You only need a final landing burn to cancel terminal velocity. For a blunt long cylinder the drag coeficient is about 0.5-0.8 and this gives a terminal velocity of 150-200 m/s at sea level.

    If we assume a degraded vacuum Merlin running at atmospheric pressure with an Isp of 300, then the landing burn will require only 200-300 kg of fuel. That’s about 3s of burn of the same degraded Merlin, at minimal (39%) throttle, 300kN thrust against a 4300kg empty stage, so a really fast approach, maybe not even feasible.

    Regarding the heatshield mass calculation, it’s wrong to use the Orion data since that is a very different system with completely different constraints. Essentially, there are two ways to do re-entry: a shallow angle where most of the energy is dissipated slowly in the upper layers as thermal load against the ablative heatshield, or a steep angle where peak temperature and g forces are high but the total heat load is minimized, requiring less heatshield mass and higher temperature materials that can soak heat and radiate it back in the atmosphere (or even be ejected).

    Since Orion is a crew capsule, it’s limited at about 5g deceleration and must use significant amounts of ablative material. Falcon 2nd stage on the other hand is a perfect candidate for steep re-entry since it can endure very high g values. On the way up, the Falcon Heavy will push 30-40t at a few g of acceleration in addition to it’s own fuel. So on the way down, it’s structurally able to sustain 20-30g of deceleration along it’s vertical axis without any damage. It can lose 7km/s of delta-v in 20-30s without any structural reinforcement, as long as the heatshield can withstand the (enormous) temperatures without melting and can transmit the ~100t of deceleration to the stage without failing. Overall it’s a very different beast and it might well be possible to survive re-entry with a quarter of the heatshield mass you estimated.

    Liked by 1 person

  8. “Falcon Heavy would still have plenty of low Earth orbit (LEO) capability, and about the same GTO payload as the Falcon 9 has today.”
    This is the same conclusion I came to in November 2015. It’s really not rocket science. Only the falcon heavy will be able to loft the heavier 2nd stage. But is that really a problem with first stage reusability and 24 hr turnaround?? I was so convinced this could work I started a tech blog that ended up a blog blog but have a look if you like —> nextstagex.com


  9. An excellent analysis.

    However, you seem to be assuming that the goal would be to recover every second stage, and that doesn’t seem to be how SpaceX operate. With first stages, they frequently don’t attempt recovery where it’s not considered practical for the mission profile, or an older model first stage is being used. If they develop a reuse capability, I could seem them using it when it suits, but continuing to fly expendable second stages for missions that require it.

    I think an interesting analysis would be to consider how much recovering the second stage would reduce cost of a launch by. F9 can launch roughly 5000kg to GTO on an expendable second stage. Assuming 5000kg is deducted from the payload of a reusable second stage, would it be cheaper to launch 5000kg payloads on a Falcon Heavy? If the first and second stages are recovered (as well as the farings), you could potentially fly your satellite for just the cost of the fuel.

    And that’s before you consider the possibility of flying 3 5000kg satellites to GTO on a single Falcon Heavy launch.


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