Exoscientist

“Fool me once, shame on you. Full me twice, shame on me.” Bob Clark

This static test fire was again at only 50% thrust level. We were misled into thinking a 50% test was sufficient to judge the full thrust burn the last time. The only way to be sure is to do a full thrust static burn to judge the water deluge system then. Bob Clark

Does SpaceX want to hide that some engines had to be changed after the last static test? Robert Clark

No. The discussion was about using the SLS for cargo. I think it is too expensive for that purpose. Use the Falcon Heavy for that purpose, or other low cost commercial launchers. No crew module would be included here. I was talking about using two hydrolox stages to get the max cargo on the Falcon Heavy but here’s another way using a single stage: Suppose we use a 45 ton prop load “Centaur-like” stage carried to LEO by the Falcon Heavy for cargo only transport to the lunar surface one-way. As it’s Centaur-like, take the ISP as 465.5s(the max Centaur RL-10 Isp with extended nozzle was 465.5s); and give it a ca. 10 to 1 mass ratio, so a dry mass of 4.5 tons with the 45 ton prop load. Then with 12 tons payload you could get: 465.5*9.81LN(1 + 45/(4.5 + 12)) = 6,000 m/s, which is sufficient for the stage to do both TLI and land on the lunar surface one-way with the 12 tons of payload. But the need for low boiloff for the 3-day flight to the Moon would reduce this payload somewhat. Alternatively, you could let the FH do the TLI burn, and say it could get ca. 20 tons to TLI. Use a 10-ton “Centaur-like” stage, at 10 ton prop load, 1 ton dry mass, and 465.5s ISP. Then it could get 9 tons payload one-way to the lunar surface: 465.5*9.81Ln(1 + 10/(1 + 9 )) = 3,160 m/s, sufficient for the lunar landing once already put on the trans-lunar trajectory to the Moon by the Falcon Heavy. Bob Clark

The current # of mirrors on ELT is 800. Only adding 30 more brings the diameter from 39.3 to 40 meters. You know there’s a considerable psychological effect of that first digit, reason why retailers like pricing their products at $39.95 rather than $40. Having the telescope size at 40 meters puts its size in good stead in relation to the 100 meter, cancelled, OWL telescope. Large ground, segmented mirror telescope costs scale by collecting area, i.e., by square of aperture diameter. Increasing the size by a factor of (40/39.3)^2 would add an additional 3% to the $1.5 billion dollar cost, or $50 million. I’m sure all astronomy enthusiasts world-wide would be willing to add that extra $50 million to bring its size up to 40 meters. Bob Clark

I’ve seen numbers for Falcon Heavy to TLI in the range of ~20 tons NASA chief explains why agency won’t buy a bunch of Falcon Heavy rockets. “It’s going to be large-volume, monolithic pieces that are going to require an SLS.” ERIC BERGER - 3/26/2018, 3:23 PM SpaceX has not publicly stated the TLI capacity of the Falcon Heavy rocket, but for the fully expendable version of the booster it is probably somewhere in the range of 18 and 22 tons. This is a value roughly between the vehicle's published capacity for geostationary orbit, 26.7 tons, and Mars, 16.8 tons. https://arstechnica.com/science/2018/03/nasa-chief-explains-why-agency-wont-buy-a-bunch-of-falcon-heavy-rockets/ But to maximize payload don’t use the kerolox FH upper stage to do the TLI burn. Use the 63.8 ton FH capacity to LEO to carry hydrolox stages for the TLI burn and for the lunar lander stage. For example a 30 ton Centaur-like stage with a 10 ton Centaur-like stage could get about 15 tons one-way to the lunar surface. This assuming low boiloff tech for the lander stage. By “Centaur-like” I mean getting high vacuum Isp and high, for hydrolox, mass ratio also. Robert Clark

Too bad about Luna-25 but one of the rovers sent to the lunar South Pole has to succeed. If the precious metals suggested to exist there by multiple lines of evidence is confirmed then we may finally have the “killer app” that not just for spaceflight to LEO brings us low-cost by large number of flights, but even interplanetary flight as well. See this article on the detections by the LCROSS orbital mission: Prospecting for native metals in lunar polar craters. January 2014 Warren Platts, Dale Boucher, George Randall Gladstone ABSTRACT One of the more astonishing results of the LCROSS mission were spectra indicating large concentrations ofnative precious metals. We hypothesize that the reported metal concentrations represent electrostatic placerdeposits: we theorize that electrostatic dust transport preferentially favors transport of submicron-sized nativemetal particles that get trapped in permanently shadowed regions (PSRs) within much smaller subareas wheresolar wind wake effects are minimal. We review the LRO LAMP and SSC UV/VIS data and note that severalspectral emission lines in the UV are consistent with the presence of platinum, as well as silver and gold. Wealso conduct a numerical simulation that shows that levitation of submicron-sized gold particles is favoredcompared with dielectric dust particles. We then develop an ore genesis model that predicts a soil massabundance of 0.11% for Au within the ore body trap that is in rough agreement with the estimate of 0.52% forAu based on the LRO LAMP column density observations. We apply the same methodology to Hg, and predicta soil mass abundance of 0.53% Hg, compared with an estimated 0.39% Hg based on LRO LAMP columndensities.Greenfield ore grades are determined initially by remote sensing techniques and ore body genesis modeling;secondly by exploratory drilling and sampling; and finally by close-in ore body delineation (detailed samplingand analysis) to provide a 3D picture of the ore body of interest. Now that we have in hand a large body ofvarious remote sensing data sets, and a predictive ore genesis model, we propose to undertake the second step—exploratory drilling. Since the occurrence of electrostatic placer deposits tends to coincide with deposits ofvolatiles, the upcoming Resource Prospector Mission (RPM) will be in an ideal position to detect native preciousmetals as well as volatiles. However, the Lunar Advanced Volatile Analysis (LAVA) instrument can onlycharacterize volatiles below 70 AMU, whereas Ag, Pt, Au, and Hg atoms range in mass from 108 to 200 AMU.Therefore, we propose that an “X-Ray Spectrometer System” (XSS) be added to the RPM rover as a secondaryscientific payload. The XSS instrument will primarily consist of an X-ray fluorescence detector (XRF) thatoffers the right combination of low mass, low power requirements, high speed, and high accuracy (ppm levelfor heavy precious metals). Finally, since water derived from PSRs will eventually be intended for humanconsumption, the likely high concentration of Hg in PSRs is a potentially grave health hazard, and representsa huge knowledge gap in our understanding of how to work and live on the lunar surface that is left unaddressedby the RPM in its present configuration (PDF) Prospecting for native metals in lunar polar craters. Available from: https://www.researchgate.net/publication/286162525_Prospecting_for_native_metals_in_lunar_polar_craters [accessed Aug 21 2023]. Bob Clark

I’ve been thinking about the possibilities of what the upcoming lunar rovers to the Moon’s south pole might find. Too bad about Luna-25, but at least one of the rovers has to succeed. Multiple lines of evidence suggest there may be precious metals there. In that case there would be an economic motive for going to the Moon. In such a scenario more commercial approaches to lunar access would be tried. Robert Clark

SLS at $2 billion per flight shouldn’t be used for cargo. It should only be used for carrying astronauts. For cargo and habitats use the Falcon Heavy. For in-space hydrolox stages on the Falcon Heavy, I estimate 15 tons one-way to the lunar surface if you have low-boiloff tech. If only the stage for TLI is hydrolox, so you don’t need low boiloff, and with a storable propellant lander stage I estimate 10 tons one-way. For Falcon 9 launches it would be 1/3rd of those numbers. Since the crew lander would be only to transfer astronauts to a lunar habitat already stationed and provisioned there the lander could be LEM sized at ca. 15 tons. By the way the total cost for SLS and Orion of $4 billion per launch is unsustainable. For an sustainable presence on the Moon we need cheaper lunar access. The Starship could do it if it gets operational, but I don’t like the multiple refuelings for a single mission. Robert Zubrin noted if you gave the SH/ST a smaller 3rd stage, then it could do single-launch missions to the Moon and to Mars. Edit: I looked it up and the Falcon 9 with cargo Dragon can carry 3,300 kg to the ISS. So with all-hydrolox low boiloff in-space stages, it could get, at a payload of 5,000 kg to the lunar surface, more payload than the Falcon 9/Dragon gets to the ISS. Or by using a hydrolox stage only for TLI, no low boiloff required, with a storable prop lunar lander stage, it could about the same as the F9/Dragon to the ISS at ca. 3,000 kg. The reason why the F9 going all the way to the Moon can get at or more payload than its payload to the ISS is firstly it wouldn’t use the cargo Dragon at 4 tons for the lunar lander and secondly hydrolox is more efficient for upper stages than kerolox. Bob Clark

That was for missions that first went through the Gateway. We’re trying to avoid that. The proposal of using existing ESA space assets to create the lunar lander built by the ESA, would save $3 billion from NASA’s Artemis budget by eliminating the SpaceX lander. And removing the Gateway would delete another $4 billion. That’s $7 billion saved by NASA for the Artemis missions. Bob Clark

I’ll take a look at your calculations and let you know what I think. You appear to be someone who likes delta-v calculations. In that case, take a look at the two cases considered here, the two Vulcain case and the three Vulcain case: Bob Clark

Could be. Any one know of images of the water deluge steel plate after the test without the static burn? Was it covered by a tarp as well then? Bob Clark

The steel deluge plate is covered by a blue tarp there. Bob Clark

Saw this on Reddit: SpaceX tearing up some sections of the concrete after the latest static fire test. Also have they revealed what’s under the tarp covering the steel plate after the deluge: Robert Clark

Sorry, but this is incorrect. Shockwaves or any kind of wave can be reflected any direction. This is a key point. Take a look at this video by Everyday Astronaut: The reverberating shockwaves visible at the 1:57 point in the @Erdayastronaut video suggest the SpaceX approach to water deluge can damage the engines. SpaceX should stop dismissing the lessons of Apollo and learn from them. Use a flame trench. Robert Clark

When impinging on another object liquid, gas, or solid they can travel in any direction, like a ball headed backwards after hitting a wall. Explosion shockwaves for example can bounce off of atmosphere layers and bounce back to Earth known as atmospheric focusing: Atmospheric focusing is a type of wave interaction causing shock waves to affect areas at a greater distance than otherwise expected. Variations in the atmosphere create distortions in the wavefront by refracting a segment, allowing it to converge at certain points and constructively interfere. In the case of destructive shock waves, this may result in areas of damage far beyond the theoretical extent of its blast effect. Examples of this are seen during supersonic booms, large extraterrestrial impacts from objects like meteors, and nuclear explosions. https://en.wikipedia.org/wiki/Atmospheric_focusing Robert Clark

This @Erdayastronaut clip shows quite alot of water reaching the level of the engines even though the water is angled outwards at the base. Then depending on the pressure of the water at this height, the exhaust flow impinging on it can cause reverberating pressure waves back on the engines. In the 5+ second Booster 7 static fire in February, only two engines failed. In the test flight in April with a longer time on the pad to liftoff with a veritable concrete tornado throwing up chunck’s of concrete, only 3 engines failed in the initial liftoff. In this latest test, 4 engines failed after only 2.7 seconds. Cause: the water deluge. Robert Clark

But you would also need some fuel left over to bring the Orion and service module back to Earth. So calculate how much prop needed to get Orion/SM/lander plus return prop to LLO. I estimated this return prop as ~7 tons. Bob Clark

Nice video by CSI Starbase. It corrects a mistaken impression of mine that there was a large hole beneath the area where the water was shooting up from. Actually there are only small holes like an upside down shower head. But starting at about the 6:50 point the video begins discussing that the water pressure is intended to be above the force of the exhaust coming to a halt impinging on the metal plate, where the water jets emanate. SpaceX wants this so that the exhaust gases don’t force their way down into the piping beneath the plate. But this raises the possibility this excess pressure could reach the engines also. Robert Clark

Something just occurred to me. The comparison has been made of the SpaceX SuperHeavy/Starship approach to the Soviet multiple failed N-1 rocket in that they both wanted to test by actually flying the full rocket until it works. The comparison was criticized on the grounds the N-1 engines were not tested individually. Instead, the engineers selected an engine at random from a batch to see if that worked. If it worked the entire batch was chosen. The engines could not be tested individually because the testing was destructive. That engine could not be used if it were tested. The SpaceX Raptor engines on the other hand are tested individually. But here’s the major failing of the Raptor: even if the engine is tested successfully there is still a quite high chance the engine will still fail when used on a flight. That is a major flaw in a rocket engine. No rocket engine would be considered successfully developed with that flaw. Because of the numerous failures of the Raptor both on the test stand and in short test hops of the Starship landing methods prior to the April test flight, I estimated the chance of engine failures of the SuperHeavy/Starship test flight was 1 out of 3. SpaceX claimed prior to the test flight their Raptor 2 was more reliable. The result? Only 1 in 4 of the engines failed. That is still a stunningly high percentage. The upshot of this in a very real sense the Super/Staship is just like the failed Soviet N-1 in flying with engines with poor reliability. Robert Clark

I haven’t heard of a Lockheed design for the Orion service module. Do you have a reference for that? Robert Clark

By that argument they would have realized doing the test flight without a flame diverter was a bad choice as well. And would have realized the FTS system was inadequate to cause immediate destruction of the vehicle. And would have realized using a flipping motion to do stage separation was a bad approach. You can not assume a approach SpaceX decides to take is a good one just because SpaceX chooses it. Again, rather than SpaceX dismissing the lessons of Apollo, they should learn from them. Bob Clark

Actually, its done by rocket engineers all the time who know what they’re doing: For instance it was done for the ESA’s ATV cargo supply vehicle to the ISS in turning it into the Service Module for the Orion. And it was done to the Delta IV Heavy’s upper stage in turning it into the Interim Cryogenic Propulsion Stage(ICPS) for the SLS. Note also that prior to the SLS use, the ATV was last used in 2015. Bob Clark

Has there been any consideration of the 4 engines failing after only 2.7 seconds in the static fire being due to water deluge coming *upwards* from the center, thus with an upward force against the engines? Usually the water deluge comes from the outside horizontally. Water deluge by SpaceX: Water deluge by NASA: Bob Clark