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

 

  

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

Copyright 2024 Robert Clark

 Elon Musk presented an update on the plans for the Starship post the third Starship test flight:

Elon Musk Starship Presentation: IFT-4 Master Plan, Starship V2 & V3, Raptor V3, Mars, IFT-3 & More.

 About 31 minutes in Elon suggests the current version V1 would be capable of 40 to 50 tons to orbit. This is bad because SpaceX sold NASA on the idea the Starship HLS could serve as an Artemis lander based on 150 tons to orbit reusable and “10ish” refueling flights. If the capability is max 50 tons, then it would take “30ish” refueling flights.

 If they intend to use version V2 then this is bad because it would require further qualification flights for the larger version and more importantly further qualification of the more powerful Raptor 3 engine needed.

 This last is doubly bad because I’d be willing to bet dollars to donuts that they never informed NASA that the current version couldn’t do it and further development would be required for the larger version.

  SpaceX needs a true Chief Engineer. Elon once said that early on when there were still doubts about its viability, they tried to recruit a Chief Engineer for SpaceX but no one good was willing to come. So Elon designated himself Chief Engineer. It is not a role Elon is well suited for. A good Chief Engineer should be scrupulously forthright. He would not refer to the little 5 or 10 second static burns SpaceX does for the SuperHeavy or Starship as "full duration".

 A true Chief Engineer would be aware that "full duration" in the industry is short for "full mission duration". These static fires in the industry are conducted at the full length and the full thrust of an actual flight and are meant to give confidence to potential customers that the engines can perform as expected for the promised capabilities of the launchers.

 However, SpaceX in using the term "full duration" for these little few-second burns, doesn't even tell the public, or its major customer NASA for which they have a billion-dollar contract, if these little burns are even conducted at full thrust.

 This has had majorly negative consequences. The FAA had great concerns in the Raptor reliability after the first test flight. In the "corrective actions" they required of SpaceX prior to a second Starship test flight, at the top of the list was correcting the tendency of the Raptor of leaking fuel and catching on fire while in flight.

 I have argued multiple, independent lines of evidence suggest SpaceX intentionally reduced the throttle of the Raptors on the booster on the second test flight, IFT-2, to improve reliability of the engines:

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

 Running an engine at reduced throttle reduces the pressure levels within the engine, high pressure being a major cause of engine fuel leaks. The Starship upper stage though was run at near full throttle on IFT-2, perhaps because performance would be reduced too much if it also was run at reduced throttle.

 The result was the booster engines worked fine, at least on ascent, while the Starship exploded on ascent on IFT-2. SpaceX has said the Starship RUD was due to an intentional LOX dump they performed to keep that flight as suborbital. However, many knowledgeable observers doubted the LOX dump alone would have caused a RUD. They argue due to the tendency of the Raptor to leak fuel, it's more likely that plus the LOX dump caused the RUD.

  For the third test flight, IFT-3, after reviewing both propellant burn rates and the acceleration profile of the flight, I'm suggesting SpaceX learned their lesson from the second test flight, and this time both stages were run at reduced throttle on this flight. And this time both stages were able to complete the ascent stage of the flight successfully.

 However, this does reduce the payload capability of the launcher. Elon has acknowledged this radically reduced payload capability in his recent update. But it needs to be explained by SpaceX why the payload is so greatly reduced. If it is because the Raptor needs to be run at reduced thrust in order to be reliable then that is an extremely important thing to acknowledge, and to inform NASA on it, because the thrust levels of a rocket go into assessing what its actual capabilities are.

 There is another very important issue about Raptor reliability. Multiple times a Raptor has undergone a RUD doing a relight during prior testing of the Starship planned landing procedure. And on this last Superheavy/Starship test flight as well Raptors underwent a RUD during the booster landing procedure. The boostback back burn appeared to have occurred successfully. But there was venting gas after the bostback back burn suggesting there may have been a fuel leak here as well.

 Note for a successful reuse of the Starship and booster, successful relights have to occur both for boostback burns and landing burns. Then in none of the prior Starship landing tests nor of the Superheavy/Starship flight tests have any flights shown successful Raptor relights without leaking fuel and catching fire, and often undergoing a RUD. 

 SpaceX has called one test of the Starship landing test, SN15, successful because it managed to land without exploding. But it is important to note even in that test a Raptor leaked fuel and caught fire prior to landing. It's just on that test SpaceX managed to extinguish the fire before the ship exploded:

 Note that in the SpaceX plans for a reusable Starship it absolutely can not work if the Raptor can not be made to relight reliably. SpaceX in not publicly providing full mission duration, full thrust testing information on the Raptors have not shown this also for relights of the Raptor.

 That is why it is so important for a launch company to publicly provide details on full mission duration, full thrust level static engine testing.

 SpaceX needs a true Chief Engineer to provide such details in a forthright manner.

    Robert Clark

SpaceX should explore a weight-optimized, expendable Starship upper stage.

 Copyright 2024 Robert Clark

 To me it’s just stunning SpaceX is ignoring that an expendable Starship could be done for 40 ton dry mass, choosing instead the current 120 tons for the reusable version: 

Elon Musk

@elonmusk

Probably no fairing either & just 3 Raptor Vacuum engines. Mass ratio of ~30 (1200 tons full, 40 tons empty) with Isp of 380. Then drop a few dozen modified Starlink satellites from empty engine bays with ~1600 Isp, MR 2. Spread out, see what’s there. Not impossible.

9:14 PM · Mar 29, 2019

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https://twitter.com/elonmusk/status/1111798912141017089

 Keep in mind that every kilo of extra mass in an upper stage subtracts directly from the payload possible. Then that 80 tons difference in the dry mass between the reusable and expendable versions is a huge difference. 

 Now,  note because of size, that, just like with the Falcon 9, the 1st stage is 2/3rd of the cost. So for ~$90 million total for the SuperHeavy/StarShip, the SuperHeavy is $60 million of that. But as the Falcon 9 shows it is much easier to get reusable 1st stage. So assume with reuse of SuperHeavy, its cost, is now, say, $5 million per launch. Now it’s a $35 million total cost for the partially reusable SuperHeavy/StarShip. BUT now because of the radically reduced upper stage dry mass, we have ca. 300 tons payload this version!(Assume SuperHeavy lands down range if you wish to maintain the high payload.) But this is about the same cost per kilo as fully reusable 100 to 150 ton payload fully reusable version at $10 million per flight cost.

 Then the question is how realistic is it the Starship could have 40 ton dry mass as an expendable? I think it is quite realistic. 

 Consider the original Atlas rocket first used to send John Glenn to orbit:

SLV-3 Atlas / Agena B.

Family: Atlas. Country: USA. Status: Hardware. Department of Defence Designation: SLV-3.

Standardized Atlas booster with Agena B upper stage.

Specifications

Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg

inclination trajectory.

Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass:

3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec.

Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m.

Propellants: Lox/Kerosene No Engines: 2. LR-89-5

Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg.

Empty Mass: 2,326 kg. Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length:

20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5

Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867 kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0 sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric

acid/UDMH No Engines: 1. Bell 8081

http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm

 The Atlas had an unusual design however. It dropped its main lift-off engine at altitude and continued on with what was called the “sustainer” engine. This engine due to much of the propellant mass being burned off had much lower thrust, and so much reduced required engine weight. Then looking at the specifications of this stage, note it had nearly a 50 to 1 mass ratio(!)

 The comparison of this sustainer stage to the 3-engine Starship upper stage is appropriate since an upper stage typically doesn’t need to have the thrust of a stage needing to lift off from the ground. Weight growth of the Starship now at 120 tons dry mass required adding 3 additional engines, to now have 6 engines.

 However, a key reason why the Atlas was able to achieve such a high mass ratio was that it used what was called “balloon-tank” design. This was a design that used pressurization to maintain its structure even on the ground. It would actually collapse under its own weight when not pressurized.

 However,  methanolox is at about 80% of the density of kerolox. So a corresponding methanolox version would be at 40 to 1 mass-ratio, better than the 30 to 1 mass ratio Elon suggested. But its not likely SpaceX would want to deal with the operational difficulties of having a stage be continually pressurized even when on the ground, unfueled, especially for a stage intended to have high launch rates.

 So I’ll look at another stage, the S-II hydrolox 2nd stage of the Saturn V rocket. The Saturn V launcher of the Apollo program was remarkable in the lightweight features of its upper stages, the S-II and the S-IVB. This page gives a list of the fueled weights and empty weights of the Saturn V stages:

Ground Ignition Weights

http://history.nasa.gov/SP-4029/Apollo_18-19_Ground_Ignition_Weights.htm

 The later versions of Apollo had improved weight optimization. We'll use the specifications for Apollo 14. The "Ground Ignition Weights" page gives the Apollo 14 S-II dry weight as 78,120 lbs., 35,510 kg, and gross weight as 1,075,887 lbs., 489,040 kg, for a propellant mass of 997,767 lbs., 453,530 kg, resulting in a mass ratio of 13.77 to 1. 

 Now, methanolox is 2.5 times greater density than hydrolox. Then the corresponding mass ratio for methanolox would be at 33 to 1. This comparison is particularly apt because the mass in the same size tanks would be approx. at the 1,200 propellant mass of the Starship.

 So Starship could reach ca. 30 to 1 mass ratio when using the weight optimizing methods used during the Apollo program.

 But if the price per kilo of this partially reusable version would be at about what the current version is what is the advantage? One advantage is as mentioned is you would not have the difficulty of making the upper stage reusable, no problematical heat shield tiles.

 There is another advantage not as concrete, but in my mind just as important if not more so. In my opinion the approach SpaceX is taking with the SuperHeavy/Starship is ill-conceived. It is based on the idea the SuperHeavy/Starship should be the be-all-end-all for ALL of spaceflight.

 But if you look at transport methods throughout history even going back to the horse-drawn era transports always came in different sizes. A comparison to the air traffic is most instructive. It turns our the largest air transports the jumbo-jet size aircraft actually make up a tiny percentage of air traffic. The great bulk of air traffic is carried by smaller aircraft.

 And even looking at SpaceX’s own Falcon Heavy demonstrates this. The per kilo cost is less than that of the Falcon 9. But the number of Falcon Heavy flights is tiny compared to the number of Falcon 9 flights.

 The fixation on the reusable Starship as the be-all-end-all for all spaceflight also leads to the poorly-conceived notion that a Mars or Moon mission must be carried out by multiple refuelings of the reusable Starship. The number of refueling flights for the Artemis lunar missions might be 8 to 16 flights.

But it is a basic principle of orbital mechanics that high delta-v missions such as to the Moon or Mars are more efficiently carried out by using additional stages. Simply by giving the SuperHeavy/Starship an additional 3rd stage, flights to both the Moon and to Mars could be carried out in a single launch.

 An expendable Starship would mean it being regarded as just another stage. And a 3rd stage could be set atop it as needed, such as for high delta-v missions. 

 As another illustration of the fact this approach to the SuperHeavy/Starship is ill-conceived, the payload of the SH/ST to GEO is nearly zero because that Starship dry mass is so high. This is the most lucrative satellite market, but a single SH/ST launch could not service that market. In order to just launch satellites to GEO the SH/ST would have to do multiple refuelings just to launch a satellite to GEO, just like when it had to launch manned interplanetary missions. This is an odd state of affairs for a rocket simply to launch satellites to GEO.

 Or of course it could utilize a 3rd stage. But if you are going to use a third stage then, why not just use it also for the manned interplanetary missions that would allow you to do such missions in a single flight?

 Robert Clark