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

 Copyright 2023 Robert Clark

 I had earlier argued that SpaceX should withdraw the Starship as a lunar lander. The primary basis for this was for safety of the surrounding population in case of an explosion on launch, SpaceX should withdraw its application for the Starship as an Artemis lunar lander.

 However, an additional reason why the Starship should not be used for a lunar lander is for safety of the crew. In the blog post, Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?, I noted two separate methods of calculation suggest the SuperHeavy booster was throttled down to <75%. I also suggested the Starship upper stage was fired at ~90%. Given this difference in thrust power levels, I suggested the booster completed its portion of the ascent because it was throttled down and the upper stage did not because it was at close to full thrust. 

 Even though the booster engines successfully fired during the ascent, the booster exploded during the attempted return. One explanation offered was the engines were damaged by fuel slosh during flip of the booster. However, it should be noted the Starship during tests of the landing procedure, that at least one Raptor always leaked fuel and caught fire.

 Note even in the last two shown here, SN10 and SN15, there were engine fires on landing. For SN10 the engine fire led to the vehicle exploding a few minutes after landing. For SN15 the fire was extinguished before it caused an explosion. SN15 was called  a “successful” landing test because it did not explode. But that a Raptor still caught fire during this test gives further evidence the Raptor is still not a reliable engine. 

 And SN11 experienced a catastrophic explosion after a fuel leak and engine fire: 

 Since relighting the Raptors in flight always resulted in an engine fire, that is the most likely explanation for the IFT-2 booster explosion as well.

SpaceX Misleadingly Characterizes Raptor's Qualification for Flight.

 SpaceX has been using the term "full duration" for their Raptor static fire tests when they might only last 5 seconds. In the rest of the industry other than SpaceX, a full duration static test means firing for the full duration of an actual launch. 

Rocket Factory Augsburg

@rfa_space

280 seconds of glorious hot fire! We are incredibly proud to be the 1st private company in #Europe () to hot fire a staged-combustion upper stage for its full duration. This qualifies our upper stage and Helix engine for flight Enjoy the video and read more in our press release https://bit.ly/3WJY2G4

https://x.com/rfa_space/status/1664683388928655374?s=20

And for the four SSME's on the SLS core stage:

 SpaceX calling their 5 second long test fires "full duration" misleadingly gives the impression that is sufficient to qualify the engines for full mission flight time.

No estimates for Raptor engine reliability publicly provided.

 For engines for a craft intended to carry astronauts and for which billions of dollars of public funds are earmarked there should be provided some indication about the safety and reliability of such engines. For instance this report provides estimates of the reliability of the different components of the SLS:

SLS-RPT-077
VERSION: 1
National Aeronautics and Space Administration
RELEASE DATE: MARCH 8, 2013
SPACE LAUNCH SYSTEM PROGRAM (SLSP)
RELIABILITY ALLOCATION REPORT

https://foia.msfc.nasa.gov/sites/foia.msfc.nasa.gov/files/FOIA%20Docs/42/SLS-RPT-077_SLSP-Reliability-Allocation-Report.pdf

 But no such estimates for the Raptor have been provided. That so many engines have consistently failed in actual flights suggest they have quite low reliability.

 In the scenario of the Merlin engines used for crewed flight, over 80 missions of the Falcon 9 were successfully flown before the first crewed flight. That means over 800 successful firings of the Merlins during that time. And added on after that the many launches since then, over one thousand successful firings of the Merlins have been made.

  Robert Clark

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

 Copyright 2023 Robert Clark

Given the Raptors repeated history of leaking fuel and catching fire I was surprised the booster was able to complete its portion of the ascent with no engine failures.

Hypothesis: the booster flew without engine failures because it throttled down to < 75%. The Starship had engine failures because it ran at ~90%, like the booster did on the first test flight with its multiple engine failures.

Throttle Down Calculated by Propellant vs. Time Graph.

Two separate observers, u/jobo555 and @space_josiah found fairly constant propellant flow rate, and therefore throttle, before where the booster begins to prepare for stage separation. Rocket thrust is given by (thrust) = (exhaust speed)*(propellant flow rate). So can get degree of throttle by propellant flow rate.

The graphs give the percentage of propellant remaining vs time. From this we can calculate the percentage change rate as the slope. For the booster it’s about 0.5%/s, 0.005/s as a decimal. Then given the total propellant load of 3,400 tons, in absolute term that propellant flow rate is 17 tons per second.

But the full thrust propellant flow rate for each Raptor v2 can be calculated as:

props flow rate = thrust/exhaust speed = 230,000*9.81/(327*9.81) = 700 kg/s. Then for all 33 engines on the booster that’s 33*700 kg/s = 23,100 kg/s, 23.1 tons/s. Then the throttle down for the booster amounted to: 17/23.1 = .736, less than 75%.

For the Starship, from the first image below, in its second graph we see from 4 minutes to 8 minutes, 240 seconds, the propellant level dropped from ~80% to ~5%, for a percentage rate drop of 75/240, 0.313%/s. Then the absolute flow rate for a 1,200 ton prop load is 3.756 tons per second. But for the 6 engines the flow rate at full thrust would be 6*700 = 4,200 kg/s, 4.2 tons/s. Then the throttle is .894, ~90%.

Note that throttling down to 75% also correspondingly drops the combustion chamber pressure from 300 bar to about 225 bar, allowing the Raptor to operate without leaks.

But this reduced thrust would also mean the SuperHeavy/Starship could carry less payload. I estimate a drop in payload to ca. 100 tons reusable. In such a scenario, the 16 refueling launches needed for a Starship HLS would be increased to 24 launches.

Throttle Down Calculated by Acceleration Graph.

A completely separate argument allows us to conclude the thrust was throttled down to less than 75%. This observer @meithan42 looked at the velocity and altitude data and derived the acceleration data.

On the acceleration graph I marked where the horizontal acceleration visually appears about 10 m/s². The vertical acceleration there visually appears as about 6 m/s². Visually this occurs at about the 90 second point.

Note that gravity subtracts ~10 m/s² from the vertical acceleration, the actual vertical acceleration produced by the engines thrust is about 16 m/s². Then the actual acceleration generated by the engines thrust is SQRT(10² + 16²) = 18.87 m/s² .

But now lets calculate the actual acceleration that should be produced by the engines assuming they were running at full throttle at the 90 second point. The thrust is, (thrust) = (exhaust speed) * (flow rate). Since we are near vacuum the Isp will be 363 s and the exhaust speed 363*9.81 = 3,560 m/s. Then the thrust at full throttle with a total prop flow rate of 23,000 kg/s, should be thrust = 3,560*23,000 = 81,880,000 N.

We'll take the total mass of the rocket as 4,850,000 considering the tanks are filled slightly less than 100%. If the engines are at full throttle then the mass after 90 seconds is 4,850,000 - 90*23,000 = 2,780,000, and the actual acceleration generated would be 81,880,000/2,780,000 = 29.45 m/s². This is well beyond amount observed.

In contrast, if we take the throttled down propellant flow rate as 17,000 kg/s, then we calculate the actual acceleration as:

363*9.81*17,000/(4,850,000 - 90*17,000) = 18.23 m/s² ,a value much closer to what is actually observed.

Robert Clark

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