Exoscientist

I don’t know what you are referring to. The SLS components have publicly available reliability estimates at 99.9%: 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 The Merlin engines can be estimated to have better than 99.9% reliability based on the number of successful firings, over 1,000, in actual operational flights. But SpaceX offers no estimates on the reliability of the Raptor. For an engine to be relied upon to power manned flights and for which SpaceX obtained a contract from a tax payer funded agency worth billions of dollars that should be a necessity. Based on the number of Raptor’s that leaked fuel and caught fire in the prior Starship tests of the landing procedures I estimated a probability of 1 in 3 that a Raptor would fail in flight, including actually exploding. SpaceX claimed the Raptor 2 to be used on the orbital test flight was a more reliable engine. The result? “Only” 1 in 4 of the Raptors on the booster failed in the first orbital test flight, including some exploding. On the second orbital test flight, the Raptors on the booster were able to successfully fire during ascent but there is strong evidence to suggest they were able to do this by throttling down their thrust level to less than 75%. And even then, some did still explode on restart as had been seen repeatedly on the Starship landing tests. Bob Clark

For the rest of the industry that is just standard industry practice: Rocket Factory Augsburg @rfa_space Jun 2 280 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… https://x.com/rfa_space/status/1664683388928655374 Bob Clark

They might not now be doing full flight duration test firings after the Merlins have been fully qualified for flight, with literally thousands of successful operational firings in flight. But, they did do full thrust, full flight duration testing of the Merlins during its development: Space Launch Report: SpaceX Falcon 9 Data Sheet. SpaceX shipped its first Falcon 9 first stage to McGregor in mid-2007. The stage was erected into the company's mass Big Falcon Test Stand during August. During November, 2007 the first Falcon 9 hot fire test, using only one Merlin 1 engine, was performed. This was followed by a two engine test in January 2008 and a three-engine test in early March, Five engine testing occurred in late May, 2008. The first nine engine test was performed on June 31, 2008, in a test tha produced 385.5 tonnes of total thrust. Two more less-than-full-duration 9-engine tests followed. On November 23, 2008, SpaceX performed the first full-duration nine-engine Falcon 9 test at McGregor. Producing tonnes of total thrust while burning nearly 227 tonnes of propellant, the burn lasted 178 seconds. Two of the nine Mer engines shut down as planned after 160 seconds, a sequence that mimicked the planned flight shutdown method. The late-evening test startled Central Texas residents more than miles away. … Merlin Vacuum Certification On March 7, 2009, SpaceX performed a full mission duration firing of the new Merlin Vacuum engine at McGregor. engine fired for six minutes, consumed 45.36 tonnes of propellant, and demonstrated a vacuum specific impulse of 34 seconds, highest ever for a U.S. hydrocarbon rocket engine. The engine produced 41.96 tonnes of thrust in vacuum conditions. …. November, 2009. The second stage was test fired in a second, smaller McGregor test stand for 40 seconds during November. On January 2010, the Falcon 9 second stage completed a full duration mission firing, its Merlin Vacuum engine producing 41.96 tonn thrust for 329 seconds. The stage was shipped to the Cape, where it arrived on January 29, 2010. There, it joined the fir stage in the new SpaceX SLC 40 horizontal integration hanger. … https://sma.nasa.gov/LaunchVehicle/assets/spacex-falcon-9-data-sheet.pdf Two interesting facts about this article. First, “full duration” is meant to be short for “full mission duration”, where the meaning is perfectly clear. Second, the last example specifically makes a distinction between a short mission firing and a “full duration” firing. Bob Clark

SpaceX still not doing full thrust, full flight duration static fires for either stage. Robert Clark

I’m making a serious charge here. I’m suggesting SpaceX knows the Raptor is unreliable and is obscuring that fact both from the NASA and the American public. For an engine that is supposed to power a craft carrying astronauts and for which billions in American tax dollars have been earmarked it should be essential that it’s reliability be established. Note that the components of the SLS each have publicly available reliability estimates at 99.9% The Raptor has been in development since 2016 and no such reliability estimates are offered for it for an engine intended to power craft carrying astronauts and even civilian passengers. Robert Clark

Engines on the booster and on the Starship both exploded on IFT-2: https://twitter.com/rgregoryclark/status/1729867002226081845 The Raptor has been in development since 2016. That it is still exploding in flight suggests it is not a reliable engine. Robert Clark

I should use his phrasing: What I hear from the ground testing at the McGregor site (6 miles from my front porch) is no more Raptor starts at full power. I suspect they may have learned the hard way not to do that. It’s possible he meant judging by the sound they are not doing full power testing. Robert Clark

The SLS has public reliability estimates for each of its components. For the Merlins engines on the Falcon 9's we can estimate it as better than 99.9% based on the over 100 successful launches and 10 Merlins on each rocket. But for the Raptor engine no such estimate has been publicly provided. Based on the number of engine failures or explosions on actual test flights, for the Starship during landing tests or the SuperHeavy/Starship orbital test launches, we can estimate it as quite low. 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 Bob Clark

Testing to failure might include testing at full power. Bob Clark

I’ve heard from someone who lives near the Starbase site that’s he’s heard scuttlebutt that SpaceX won’t be testing the Raptor at full thrust anymore. Anyone else hear that? Robert Clark

I once jokingly said that if they were asked how much the Ariane 6 SRB’s cost, ArianeSpace and ESA would respond, “We’re not going to tell you that!” Turns out that wasn’t far from the truth: Ariane's New Price Tag Is Bad News for Airbus, Great News for Boeing and Lockheed (and SpaceX). By Rich Smith – Dec 23, 2023 at 7:07AM Recall that Ariane originally targeted a 50% cost reduction between Ariane 5 and Ariane 6. Asked about the price at a press briefing earlier this year, though, Arianespace CEO Stéphane Israël first blamed inflation, complaining that Ariane has to work with a "real economy," then flat-out declined to say how much the rocket will cost, telling reporters to "speak...with our customers," as Ars Technica reported in September. Taking the hint, Ars dug up a June speech from ESA Space Transportation Director Toni-Tolker Nielsen, who confided that Ariane 6 is looking likely to cost about 40% less than Ariane 5 -- not 50%. But now, even 40% looks over-optimistic. https://www.fool.com/investing/2023/12/23/arianes-new-price-tag-is-bad-news-for-airbus/ Bob Clark

I really dislike this phrasing of SpaceX of calling a burn “full duration” to mean it lasted the planned time of the burn, even if it was only 5 seconds. It used to be the term “full duration” meant the length of an actual flight burn, which will be several minutes long. If you want to say the burn lasted the planned length just say it lasted the planned length. It’s hard to believe that both the FAA and NASA would be effected by this “Jedi mind trick”: just call it “full duration” and that means the engines have been fully flight qualified for a full burn time in flight. Bob Clark

The comparison between the RL10A-4 and the J-2X is illustrative. The RL10A-4 has an expansion ratio 84 to 1 and gets an Isp of 451 s. The J-2X has an expansion ratio of 80 to 1 and gets an Isp of 448 s: http://www.astronautix.com/j/j-2x.html . So the J-2X Isp is over 99% of the RL10A-4 value. It is correct in general a gas generator engine loses efficiency compared to an expander cycle engine because the gas generator dumps some of the burned gases overboard. But another consideration is combustion chamber pressure. Higher chamber pressure allows higher Isp. The chamber pressure of the RL10’s is about ~40 bar while for the J-2X, it’s about ~100 bar. So if the same close Isp value holds for the ultra high expansion ratio case then the modeled 482.5 s Isp for an expander cycle engine would only be reduced to ~479.3 s for the J-2X. But you do have to do trades on whether the increased weight is worth the increased payload permitted by the higher Isp. I’ll estimate the increased weight by what happened with the RL10. From the table below adding the longer nozzle extension to the RL10-A4 to get the RL10-B2 added about 110 kg to the engine mass, while increasing the Isp from 451 s to 465.5 s. The J-2X is a 13 times larger engine as measured by thrust, so I’ll estimate the nozzle extension as 13 times heavier so about 1,400 kg. But we might be able to make the added mass somewhat smaller. The current nozzle extension on the J-2X is metallic. We can reduce the mass of the current nozzle extension by a half or more by using carbon fiber. RL10B-2 Active 1998 277 kg (611 lb) 110.1 kN (24,800 lbf) 465.5 s (4.565 km/s) 2.2 m (7 ft 2 in) Extended: 4.15 m (13 ft 7.5 in) 2.15 m (7 ft 1 in) 40:1 5.88:1 280:1 44.12 bar (4,412 kPa) 5-m: 1,125 s 4-m: 700 s Delta Cryogenic Second Stage, Interim Cyrogenic Propulsion Stage [1][42] RL10A-4-1 Retired 2000 167 kg (368 lb) 99.1 kN (22,300 lbf) 451 s (4.42 km/s) 1.78 m (5 ft 10 in) 1.53 m (5 ft 0 in) 61:1 84:1 42 bar (4,200 kPa) 740 s Centaur IIIA [11][43] RL10A-4-2 Active 2002 168 kg (370 lb) 99.1 kN (22,300 lbf) 451 s (4.42 km/s) 1.78 m (5 ft 10 in) 1.17 m (3 ft 10 in) 61:1 84:1 42 bar (4,200 kPa) 740 s Centaur IIIB Centaur SEC Centaur DEC [11][44][45] For a nozzle extension to 465.5 s on the J-2X and adding on also the 1,400 kg increased engine mass this would increase the LEO payload from 119.7 tons to ~122 tons by the Silverbirdastronautics.com estimator. This small size increase would be of doubtful value. If the 482.5 s nozzle extension could be added without much greater mass than the 1,400 kg, when taking into account also the current metallic nozzle could be reduced by carbon fiber, then we could get ~128 tons to LEO. Robert Clark

There have been several stories recently coming out from NASA discussing the number of launches required for the NASA/SpaceX lunar landing. The undercurrent you sense is that NASA does not really like the plan: Starship lunar lander missions to require nearly 20 launches, NASA says Jeff Foust November 17, 2023 https://spacenews.com/starship-lunar-lander-missions-to-require-nearly-20-launches-nasa-says/ And the reception that Destin of “Smarter Every Day” got in his critique of the plan before a room full of NASA engineers suggests he was saying out loud what NASA engineers couldn’t say. I proposed a single launch architecture using an Apollo-sized lunar lander with extra propellant being added to the Orions service module: Possibilities for a single launch architecture of the Artemis missions, Page 2: using the Boeing Exploration Upper Stage. http://exoscientist.blogspot.com/2023/08/possibilities-for-single-launch.html Because of the greater mass to be sent to TLI though, it may require greater payload capability than the SLS 1B’s 105 tons to LEO. So I was investigating methods of giving SLS greater payload capacity. The SLS 2 will get 130 tons to LEO but it won’t be here until the 2030’s and will require further billions to develop the upgraded SRB’s. (If anyone tries to convince you the new SRB’s won’t cost much, please see my sig file.) Robert Clark

The authors of those propellant burn and acceleration graphics deduced the data from the information on the SpaceX launch videos. The flight angle as a graphic is also provided on those videos: Bob Clark

Here’s an estimate of the SLS 1B payload to LEO with the RL-10’s swapped out for a J-2X. For the dry mass of the Boeing EUS I increased the dry mass by 1,200 kg to account for the greater mass of the J-2X compared to the four RL-10’s. I kept the residuals for the engine at 3% though it may be less than this. Then the input data looks like this: And the results like this: About 119.7 tons to LEO. In the next post I’ll investigate increasing the payload by using a longer extension nozzle on the J-2X. Robert Clark

I was using the starting propellant load for the booster at the ~90% range, not 80%, and for the Starship at the ~95% range in my calculations. The authors of those graphs are at the links I provided in my updated blog post. I realized I had not credited them properly so I updated my blog post to give the links where the authors described their calculations. By the way the authors provide links to the actual data not just the graphs so you can calculate with them more accurately. About determining the acceleration provided by the engines, to determine thrust, when you take into account gravity you do need to be able provide separately the horizontal and vertical acceleration. This needs to be calculated from the speed and altitude information provided by the launch video. But since the separate coordinate info is not provided directly, errors are introduced when you derive them numerically. There is an alternative method by using the graphic in the video that shows the angle of flight of the vehicle. From that you can determine the x-and y-velocities, and then the x- and y-accelerations. Robert Clark

Another attempt to get an estimate using more accurate numbers of the payload to LEO of the SLS 1B with the RL-10’s on the Boeing EUS swapped out for the J-2X engine. First, note a short-coming of the SilverbirdAstronautic.com is that it overestimates the payload for rockets with large side boosters, such as the Space Shuttle, Ariane 5, and SLS. This is likely due to the fact these large SRB’s greatly throttle down even well before burnout. So instead of using the vacuum thrust for the SRB’s we’ll input the smaller sea level thrust instead. In regards to the parameters of the SRB, Core stage, and Boeing EUS, we’ll use the data in reports by Boeing engineers: Note a key modification to the input numbers will have to be in the residual fields for each stage. The SilverbirdAstronautics estimator as default takes the residuals as 0.5%. But in the image above with the SLS parameters the residual for the SRB’s amounts to only 0.14% but more importantly the residual for the core is about 1%, and the residual for the ICPS stage is about 3%. Boeing EUS residuals aren’t specified but we’ll take it also as about 3% since it also uses the RL-10 and has similar structural form as the ICPS. Note also the mass of the adapter for the ICPS called the LVSA, “launch vehicle stage adapter”, is about 5 tons. We’ ll take the adapter, more accurately, interstage, for the Boeing EUS as also about 5 tons and add this onto the dry mass of the core. Then the input to the estimator is: With an estimated payload to LEO as: The payload value of 105.7 tons is close to the NASA value of 105 tons. In the next post I’ll do the estimate with the four RL-10’s swapped out for the J-2X. Robert Clark

Two separate, independent methods suggest SpaceX throttled down the booster engines < 75%, while the Starship engines fired at ~90% thrust: Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures? http://exoscientist.blogspot.com/2023/12/did-spacex-throttle-down-booster.html This is important to know because if the engines need to operate at < 75% to be reliable, then I estimate the reusable payload would be lowered from 150 tons to ~100 tons. Then instead of needing perhaps 16 refueling flights for the Artemis landing missions there would need to be perhaps 24. Robert Clark

In this slowed down clip by Zack Golden you can see objects spewing out of the engine section. Judging by the 9 meter diameter of the booster, I would say these objects are 2 to 3 meters long, 6 to 10 feet. That almost certainly means it was an engine explosion: https://twitter.com/CSI_Starbase/status/1734783989871763768 Robert Clark

NEWS 12 December 2023 ‘Head-scratcher’: first look at asteroid dust brought to Earth offers surprises Researchers have begun examining the pristine space rock collected by NASA’s OSIRIS-REx mission. By Alexandra Witze https://www.nature.com/articles/d41586-023-03978-4 Bob Clark

Questions about the Raptor reliability will continue to be raised as long as the Starship continues to explode in flight. The FAA aware of the Raptor’s tendency to leak fuel and catch fire in flight listed rectifying this at the top of list of things it wanted SpaceX to correct. Note even in the last two shown here, SN10 and SN15, there were engine fires on landing. And we already know what happened to SN11 not shown here. 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. Robert Clark Edit: by the way, I understand the point you’re making but perhaps it’s not the best example to give in regards to a rocket engine. For rocket engines the turbo pumps are part of the engine. Indeed they have traditionally been regarded as the most finicky part of a rocket engine. So much so that a saying among propulsion engineers is, “Rockets are turbo pump developments with rockets attached.”

Zack Golden of “CSI Starbase” says Kathy Lueders has confirmed the automatic FTS destroyed both stages, Three points: 1, there were small fires in the engine section prior to AFTS; 2, the flame suppression system has no value if AFTS still activates on fires in engines; 3, Raptor no more reliable now than before on relights: What Happened to Starship SN11? | SpaceX Starship SN11 Test Flight & Explosion Cause Analysis. Bob Clark

ArianeGroup CEO Finally Says Quiet Part Out Loud By Andrew Parsonson - December 8, 2023 https://europeanspaceflight.com/arianegroup-ceo-finally-says-quiet-part-out-loud/ Quite astonishing. The ArianeGroup CEO suggests the previous design of the Ariane 6 using two large SRB’s for its first stage instead of a liquid-fueled core would have been better. He hasn’t gotten the point reusability is essential to be competitive with SpaceX. It’s like Tory Bruno head of ULA questioning whether reusability is worthwhile. Here it is with SpaceX beating ULA into the ground with their price cuts from reusability, with ULA being driven to the brink of bankruptcy, and with ULA opening themselves up for sale to forestall going under, and the CEO doesn’t know why. The New Space starts-ups all recognize the importance of reusability. Old Space has become old and decrepit. Bob Clark

As shown in the image from the twitter post, for that 482.5 ISP, it is a matter of a nozzle extension. Those are already used on the highest performing RL-10 versions that bring their ISP’s to the 465.5 range. You can’t get much higher than this on current rockets that use them such as the Delta IV Heavy and Atlas V because the expansion ratio is limited by the rocket diameter. Why not just use the ultra large extension nozzles on the RL-10’s on the SLS with its larger diameter? The high payload possible is due also to the high thrust of the J-2X engine. You would need 3 times more of the RL-10’s so twelve of them to match the J-2X. Too much added weight in addition to practicality issues. Remember it was NASA’s original plan to use the J-2X engine. It was decided not to use it when simulations showed its high thrust could damage the Orions solar panels once deployed. But I’m envisioning twice as much payload to TLI so the acceleration would be much reduced. Edit: the payload would still be quite high if we added a nozzle extension to the J-2X to bring the ISP to the range of 465.5s already reached by the RL-10 with nozzle extensions: Giving a result: Still over 130 tons to LEO. Robert Clark