This is a question I've been pondering for the last few weeks after I back-calculated the details of the Merlin 1D engine. Just as a thought exercise, I wanted to see what upgrades the Merlin 1D would need to reach aircraft gas turbine engine time between overhauls. After backing out some dimensions from the published data about the design, I came up with a series of upgrades. I made the following assumptions to get the dimensions:

  1. Sensor ports, ignition fluid injection ports, and tight bends in piping (such as on the turbine exhaust or propellant exit) generate a stress concentration factor of 3.
  2. In places where the material is not published, aluminum alloy 2021-T6 is assumed on low temperature components, and Rene 88 is assumed on high temperature components. (High temperature is anything above 400 K).
  3. The mixture ratio is stoichiometric for the propellants.
  4. The fuel is assumed to be made up of $C_{12}H_{26}$.
  5. The parts are assumed to be crack free.

From those assumptions, the following: static pressure/static temperature/wall temperature/passage dimensions/wall thickness/average stress/maximum stress/cycle lives were calculated using the published performance specifications (wikipedia and spacenews.com).

Combustion Chamber

  • Pressure: 6.77 MPa
  • Static Temperature: 3006 K
  • Wall Temperature: 566 K
  • Diameter: 0.609 meters
  • Wall Thickness: 0.004 meters
  • Average Stress: 233 MPa
  • Maximum Stress: 700 MPa/1,000,000 cycles or 54,000 hours of engine time


  • Static Pressure: 3.9 MPa
  • Static Temperature: 2613 K
  • Wall Temperature: 529 K
  • Diameter: 0.223 meters
  • Wall Thickness: 0.004 meters
  • Average Stress: 105 MPa
  • Maximum Stress: 316 MPa/Infinite cycles, below endurance limit of 600 MPa


  • Static Pressure: 101 kPa
  • Static Temperature: 1697 K
  • Wall Temperature: 388 K
  • Diameter: 0.6077 meters
  • Wall Thickness: 0.004 meters
  • Average Stress: 7 MPa
  • Maximum Stress: 10 MPa/ Infinite cycles, below endurance limit of 600 MPa

Oxidizer Pump

  • Exit Pressure: 10.6 MPa
  • Casing Diameter: 0.116 meters
  • Wall Thickness: 0.005 meters
  • Average Stress: 123 MPa
  • Maximum Stress: 320 MPa10,000 cycles or 544 hours of engine time

Fuel Pump

  • Exit Pressure: 16.1 MPa
  • Casing Diameter: 0.173 meters
  • Wall Thickness: 0.011 meters
  • Average Stress: 128 MPa
  • Maximum Stress: 383 MPa/8,000 cycles or 435 hours of engine time

Based on these results, the weak point in terms of static endurance is the turbopump, specifically the fuel turbopump casing. To reach 1000 hours between engine overhauls, the average stress on the casing needs to come down. Nesting the fuel pump inside the oxidizer pump would reduce the net stress experienced by the fuel pump by using the pressure developed by the oxidizer pump to compress the fuel pump casing. And the oxidizer pump increases its wall thickness to 0.009 meters to reach 1000 hours between overhauls.

However, this is based on the data that I could find out. In my opinion, the real weak points are the impellers, shafts, and bearings in the turbopump. Since they spin at 20,000 rpm, they can quickly go through fatigue lives if they have an imbalance occur due to cavitation, debris, or external vibration. I'm sure SpaceX and Barbor-Nichols (the folks who designed the turbopumps) are aware of this and maybe have some solutions. Outside of increasing the damping inside the pump by using a ferrofluid/magnet and having magnetic bearings, I don't know how one would increase the overhaul time.

Thus, I ask the community: How would you do it? What would you change in the engine design to reach 1000 hours between engine overhauls?

  • 1
    $\begingroup$ Can I ask why you're aiming for 1000hrs? It doesn't strike me as a very realistic figure. Typical time to MECO is ~140-160s. That means 1000hrs is 22,500 flights. Granted, the three engines using during recovery will get a bit more usage, but you're still an order of magnitude higher than the stated non-refurb flight goals of 10. $\endgroup$
    – Saiboogu
    Jan 15, 2018 at 18:25
  • 1
    $\begingroup$ Nevermind, I see the goal to reach aircraft turbine runtimes. I also see the age of the question - oops. Regardless, I think a more realistic figure to crunch numbers for would be ~36 hrs between refurbishments, to account for the goal of 10 flights plus a little buffer for landing and static fires prior to flights. $\endgroup$
    – Saiboogu
    Jan 15, 2018 at 18:50
  • $\begingroup$ Recently Musk stated de-sooting the turbopump was an issue. I wonder if it would be difficult to build a new turbopump that runs on a less sooty third (pump) fuel. Maybe they could run this new turbopump with alcohol, or dimethyl ether. Or perhaps, water or methanol injection together with rp1 fuel in the gas-generator might mitigate the sootiness issue. $\endgroup$
    – Tobe
    Dec 30, 2020 at 16:05
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    $\begingroup$ I must disagree with point #3--rockets burn fuel rich to ensure they never burn oxygen-rich. An engine burning oxygen-rich generally ends up burning engine-rich. $\endgroup$ Jan 3, 2021 at 21:03

1 Answer 1


"Musk stated de-sooting the turbopump was an issue. I wonder if it would be difficult to build a new turbopump that runs on a less sooty third (pump) fuel."

IMHO this gets to the heart of the Mean Time To Failure issue - Kerosene combustion has an inherent soot problem that is going to be detrimental to substantial increases in MTTF.

The real solution is to build a new engine that from the ground up has maintainability and reuse as design requirements. Hence, Raptor.
Meth-LOX does not have the soot problems Ker-LOX does; and Raptor has other design improvements over Merlin to increase MTTF.


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