A route to aircraft-like reusability for rocket engines.

  Copyright 2024 Robert Clark

  A general fact about aircraft jet engines may offer a route to achieve aircraft-like reusability for rockets:

https://www.airliners.net/forum/viewtopic.php?t=739359

The key question: does this fact about jet engines also hold for rocket engines?

 If so, increasing a turbopump rocket engine power just 10% to 15% cuts engine life in half. And conversely, decreasing it by 10% to 15% doubles engine life. And if so, would this still work if we repeated the concept multiple times? If we reduced the thrust by .9^5 = .60, i.e., to 60%, which most turbopump engines can manage, then we could increase the lifetime by a factor of 2^5 = 32 times? Then a Merlin engine with a lifetime of, say, 30 reuses by running it only 60% power could have its lifetime extended to 1,000 reuses? 

 This would be in the range of number of reuses of the type of jet engines used on long haul flights. Rocket engines with that number of reuses probably also would allow “gas and go” operation. That is, no major refurbishment needed in between flights, as with jet engines.

Is reduced temperature the key?

 In examining this question of rocket engine longevity versus jet engine longevity I once hypothesized it had to do with the high temperatures rocket engines operated at, typically ca. 3,000 °C, whereas jet engines might only operate at ca. 1,200 °C to 1,500 °C.

 It might be thought it would be the high pressures of rocket engines but that can’t be the primary reason since automobile diesel engines can operate at hundreds of bars of pressure, above even that of rocket engines for many hours of service:

ToughSF

@ToughSf

The pump pressures in rockets are impressive, but let's not forget that the injection system in modern diesel engines operate at 2500 bar. They are also fast enough to accomplish up to eight separate fuel injections with each cylinder cycle. Bosch CRS3-25. https://youtu.be/T7o2hvoJE-Q

0:20

8:00 AM · Sep 12, 2020

 However, temperature can not be the only factor involved either. The hydrogen turbopump on the Space Shuttle Main Engine only operates at 1,310 °F, 710 °C, and turbopumps for rocket engines are a primary limiting factor in rocket engine longevity. 

 Still, the SSME hydrogen turbopump is still a type of turbomachinery and the longevity extension mentioned above known to hold with turbomachinery in general will still apply to it: derating its power level by a small relatively small percentage will result in a quite large increase in longevity.

 Note an intriguing possibility also exists. The increase in longevity is actually in regards to turbomachinery specific power, that is power per weight. Note for example diesel engines are quite ruggedly built. So rather than derating the engine to, say, 60% power we could ruggedize it, i.e., using thicker, heavier components, increasing weight by ~50% to achieve the same reduction of specific power to 60% while keeping the same power level.

 The reason you might want to do it that way is you may want to maintain the same power, i.e, thrust level. Lowering the thrust level would increase gravity losses, thus reducing payload, especially for a first stage engine, where most gravity loss is incurred. On the other hand at least for first stages increasing the mass of the stage only results in a fraction of that increase in the first stage mass being loss in payload, typically only 5% to 10% of that increase being lost from payload capability.

 On the other hand rebuilding the turbopumps and possibly the combustion chamber using heavier components would be an expensive rebuild. The other approach of just derating its operational power level simply means running the existing engine at the lower power.

Fast ramp up to full power as cause of shortened longevity.

 There is another interesting fact in that passage I cited on jet engine longevity:

Once a turbine engine has been shut down, usually it needs to cool down before restarting, depending on power levels prior to shut down. Cooling down can be done at ground idle power setting. Turbine engines generally don't like to be shut down straight from take-off power. They also require warming up before slamming to take-off power.

 Take a look at this video of the SSME startup sequence:

STS-134 - The final launch of Endeavour - Full Launch in HD. 

                                    

 
 At about the 9:52 point the video shows the thrust level of the SSME goes from 0% thrust at T minus 7 seconds to 100% thrust at T minus 2 seconds, i.e., in only 5 seconds. 

 For jet engines, it is known such rapid changes are detrimental to engine longevity. It is quite likely the same is true for rocket engines. But for turbopump rocket engines you need almost immediate ramp up to the safe 60% to 100% throttle range.

 But pressure-fed engines are usually more forgiven in this regard. They have a much wider range of throttleability levels. In that case you could have more gradual ramp up to full power, full thrust, or whatever you set the operational thrust level now to be.

 Then it would be helpful if the turbopump engine could be operated in a pressure-fed mode, both in start-up and cool-down. This is known to be possible for some gas-generator engines which have an "idle mode". This is when the turbopumps are bypassed and the engine runs just on the pressure on the propellant in the tanks, i.e., a pressure-fed mode.

 It's not known if staged combustion engines such as the SSME, RD-180 or Raptor can be run in idle mode. Given the potential of having such high performance engines have reuse lifetimes measuring in the thousands, I think it is a possibility that should be explored. A couple of possible ways need to be explored as possible ways of accomplishing it: one, the propellant could be allowed to flow through the turbopumps to the engine but the turbos being locked so they don't operate. And two, different piping being installed so that the turbos could be bypassed entirely if desired and the propellant go directly to the engines.

 Another possibility is that there are turbopump engines that can run at highly reduced throttle level, to below 10% throttle for example. One such is the Common Extensible Cryogenic Engine(CECE), variant of the famous RL10 engine. These engines operate on a lower power engine cycle however known as the expander-cycle. Rather than using preburners to run their turbopumps as do gas-generator and staged-combustion engines, they run their turbos on the heat produced from the hydrogen run through the engine cooling channels. It may be possible to get higher power gas-generator and staged-combustion engines to run in a low power mode by adding additional piping to bypass the preburners,  sending the hydrogen from the cooling channel directly to the turbopumps.

Temperature variations as cause of longevity reduction.
 I'm informed by someone in the industry of another cause reducing rocket engine longevity: temperature variations across the combustion chamber causing differential thermal expansion.

 This is primarily because of the high temperature in the combustion chamber at thousands of degrees compared to the low temperatures in the cooling channels at hundreds of degrees. If this is a problem of steady-state operation, i.e., not an issue of temperature ramp up speed, then it would seem to be the only solution would be operating the engine at lower temperature.

 I mentioned that typically rocket engines operate ca. 3,000 °C, whereas jet engines might only operate at ca. 1,200 °C to 1,500 °C. This can not be the only issue with rocket engine longevity as we saw as the relatively low temperatures turbopumps operate at shows. Still, it could be an issue with combustion chambers. 

 There is a way to run rocket engines at the lower temperatures for most of the time of the flight. For rocket engines at near vacuum which is most of the flight, to get high performance, high Isp, what's really important is high expansion ratio. You don't really need high chamber pressure and the associated high temperature then. 

 Then if you had an adaptive nozzle, i.e., altitude compensation, that could extend to high expansion ratio at high altitude, near vacuum, you could get high Isp even using relatively low pressure and low combustion temperature in the near vacuum regime. This though would result in reduced thrust.

 At low altitude soon after takeoff though you do need the high thrust. You might then run the engines at high pressure, high temperature then. The engine lifetime would still be extended longer than before in this being a relatively short portion of the engine firing time. 

 Alternatively, there might be a way to run the engines even at low altitude at low pressure, low temperature. It's related to the concept of altitude compensation. It's a method called the  "thrust augmented nozzle(TAN)". It's quite analogous to afterburners for jet engines.

 In TAN, additional propellant is injected into the nozzle and combusted there. This allows relatively low chamber pressures to operate in the engine while at the same time having a high expansion ratio even at sea level, and providing higher thrust.

Is SpaceX using the reduced throttle effect on longevity to improve Raptor reliability?I have argued that multiple lines of evidence suggest SpaceX throttled down the Raptor on the second and third test flights to improve reliability:

Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?

https://exoscientist.blogspot.com/2023/12/did-spacex-throttle-down-booster.html

SpaceX should withdraw its application for the Starship as an Artemis lunar lander, Page 2: The Raptor is an unreliable engine.

https://exoscientist.blogspot.com/2023/12/spacex-should-withdraw-its-application.html

SpaceX should withdraw its application for the Starship as an Artemis lunar lander, Page 3: Starship has radically reduced capability than promised.

https://exoscientist.blogspot.com/2024/04/spacex-should-withdraw-its-application.html

 Elon Musk gave the current version of the Superheavy/Starship launcher in his recent update to the Starship program as only having a payload capability of 40 to 50 tons to orbit. The problem is going by the specifications SpaceX has given for the dry mass and propellant mass of both stages and for the Raptor thrust and Isp values, the vehicle should easily be able to reach 100+ tons to orbit even as reusable. Then either the dry mass values or Raptor values or more likely both must be significantly worse than the cited values.

If so, then SpaceX is missing a golden opportunity to develop aircraft-like reuse capability for the Raptor and for the Merlin. It would require acknowledging however they are running the Raptor at reduced power to improve reliability, and in effect, longevity.

  Bob Clark