Exoscientist

The proof is in the pudding, as they say. Two key things to look out for is how well the single reused engine performs on this flight, and how many engines from this flight will be reused on the following one. Robert Clark

IF it is scrapped would that include the engines? That would not speak well toward Raptor reusability. Bob Clark

But the one being now flown is essentially the same as flown last time. It would be a great proof of of reusability if the same engines were used. Bob Clark

There appears to be only two possible reasons for this: either the Raptor is not as reliable for reusability as thought or it was damaged during the landing burn. Bob Clark

That was for a crewed mission to another star system, tens of trillions of kilometers away. But the primary focus of this new research is for missions still in the Solar System, 1/10,000th the distance, with cubesat-like probes. Here’s the published report by Rene Heller et.al.: Low-cost precursor of an interstellar mission. René Heller1,2, Guillem Anglada-Escudé3,4, Michael Hippke5,6, and Pierre Kervella7 ABSTRACT The solar photon pressure provides a viable source of thrust for spacecraft in the solar system. Theoretically it could also enable inter- stellar missions, but an extremely small mass per cross section area is required to overcome the solar gravity. We identify aerographite, a synthetic carbon-based foam with a density of 0.18 kg m−3 (15 000 times more lightweight than aluminum) as a versatile material for highly efficient propulsion with sunlight. A hollow aerographite sphere with a shell thickness εshl = 1 mm could go interstellar upon submission to solar radiation in interplanetary space. Upon launch at 1 AU from the Sun, an aerographite shell with εshl = 0.5 mm arrives at the orbit of Mars in 60 d and at Pluto’s orbit in 4.3 yr. Release of an aerographite hollow sphere, whose shell is 1 μm thick, at 0.04 AU (the closest approach of the Parker Solar Probe) results in an escape speed of nearly 6900 km s−1 and 185 yr of travel to the distance of our nearest star, Proxima Centauri. The infrared signature of a meter-sized aerographite sail could be observed with JWST up to 2 AU from the Sun, beyond the orbit of Mars. An aerographite hollow sphere, whose shell is 100 μm thick, of 1 m (5 m) radius weighs 230 mg (5.7 g) and has a 2.2 g (55 g) mass margin to allow interstellar escape. The payload margin is ten times the mass of the spacecraft, whereas the payload on chemical interstellar rockets is typically a thousandth of the weight of the rocket. Using 1 g (10 g) of this margin (e.g., for miniature communication technology with Earth), it would reach the orbit of Pluto 4.7 yr (2.8 yr) after interplanetary launch at 1 AU. Simplistic communication would enable studies of the interplanetary medium and a search for the suspected Planet Nine, and would serve as a precursor mission to α Centauri. We estimate prototype developments costs of 1 million USD, a price of 1000 USD per sail, and a total of <10 million USD including launch for a piggyback concept with an interplanetary mission. https://www.aanda.org/articles/aa/pdf/2020/09/aa38687-20.pdf From the scaling indicated there, it appears the mass scales by the areal size of the sail, i.e., by the square of the diameter. So a 100 meter wide sail, would have a sail mass of 2.3 kg and a payload mass of 22 kg. Bob Clark

A good question. His idea needs to be tested in the lab. If I’m to make a guess, the side of the smaller sail facing the Sun is non-mirrored so has low light pressure pushing it away from the Sun, but the side facing the larger main sail is mirrored. And more importantly that larger main sail focuses all it’s light onto the smaller sail in concentrated fashion. Bob Clark

In that Robert L. Forward conception of using reflected light from the main sail to slow down a smaller sail at the destination, he also included the possibility of using the same method to actually *return* from the far destination. Imagine getting returned samples from ‘Oumuamua, the Jovian and Saturnian moons, and Pluto! However, there is a sticking point in using this method in the case we’re considering here. Forward was imagining it for laser propelled propulsion. In that case you can focus the reflected light that is coherent and collimated. But for our scenario we’re using solar light which will be non coherent and uncollimated. It may not be possible to get the highly focused light at long distances in this scenario. It might be we can carry a laser to do it but that may be too heavy. There may be other light weight methods to do it. An intriguing possibility: IF it did work, then could it be used to do staging by sending focused light from the main sail *forward* to a smaller sail ahead of it? Then we could increase the speed multiple times by doing multiple staging. Bob Clark

The Starship HLS plan for Artemis may be even more complex than thought. It may require two refueling depots, the known one in low Earth orbit plus another one in an elliptical high Earth orbit: SpaceX seeks a single FCC license for multiple future Starship missions, including commercial/Starlink launches and Artemis. Filing shows some technical details about HLS lander, indicating it may require a 2nd refueling in an elliptical Earth orbit. [FCC filing link](https://licensing.fcc.gov/cgi-bin/ws.exe/prod/ib/forms/reports/swr031b.hts?q_set=V_SITE_ANTENNA_FREQ.file_numberC/File+Number/%3D/SATLOA2024121800288&prepare=&column=V_SITE_ANTENNA_FREQ.file_numberC/File+Number), most of the technical details is in the [Technical Annex](https://licensing.fcc.gov/myibfs/download.do?attachment_key=32702913) &nbsp; 1\. The filing covers launch, reentry and in-space operations in the following orbits: * LEO: circular orbit with altitude between 181km and 381 km, all inclinations. This would be the deployment orbit for Starlink and the orbit for HLS LEO depot. * Elliptical Earth Orbit: perigee is between 181km and 381 km, apogee is between 10,534km and 150,534km, inclination between 28 and 33 degrees. Filing refers to these as MEO/HEO but technically they're transfer orbit to circular MEO/HEO. This would cover GTO and transfer to MEO such as orbit of GPS satellites, although the filing didn't mention these. It did mention that this will be the Final Tanking Orbit (FTO) for crewed lunar mission where HLS lander will receive a 2nd propellant transfer. * Translunar Injection (TLI), Lunar orbits (NRHO, LLO) and lunar descent/ascent/surface: These would be for Artemis missions &nbsp; 2\. Communication bands used by Starship * UHF and IEEE 802.11ac 5.8 GHz band: Used for communication between HLS lander and EVA suits on the Moon. I believe these are required by NASA. Range is up to 2km. * S band: Most communication is in this band, including ship to Earth, ship to ship/depot, ship to Orion/Gateway, etc. HLS lander and depot will also use this band to communicate with NASA's TDRSS satellites in Final Tanking Orbit. * Ku band: This is used for radio communication between Starship and Starlink constellation, however it's only usable below 300km. * Ka band: Used by HLS lander for direct to Earth communication &nbsp; 3\. Technical details about HLS lander * As said above, a 2nd propellant transfer from depot to HLS lander may be required in an Elliptical Earth Orbit. Note that someone apparently with sources [mentioned this a few months ago on twitter](https://x.com/Jenakuns/status/1821914964220899650): "Starship HLS conducts 2 refuelling's; 1 in LEO, then a second one in an elliptical orbit to get the architecture delta v down. That's the reason why launch count doesn't line up with wet mass/payload ratio." * HLS lander will carry 4 dual-band (S/Ka) gimbaled parabolic reflector antennas, one in each quadrant. Exact location of these antennas is not disclosed. * HLS lander will carry 2 lunar landing radar in the 35.5-36 GHz band. It'll be activated 4km above the lunar surface and run for approximately 5 minutes until landing. There were [FCC Special Temporary Authority filings for testing this radar on an airplane as early as October 2021](https://apps.fcc.gov/oetcf/els/reports/STA_Print.cfm?mode=current&application_seq=111132&RequestTimeout=1000), call sign is WT9XBJ. https://www.reddit.com/r/SpaceXLounge/s/hEDBg96kBk Bob Clark

Robert Zubrin has made this point numerous times: https://x.com/robert_zubrin/status/1374861051490000896?s=61 https://x.com/robert_zubrin/status/1725747455247979000?s=61 https://x.com/robert_zubrin/status/1278300197664120833?s=61 https://x.com/robert_zubrin/status/1256718751619145728?s=61 https://x.com/robert_zubrin/status/1256574624202018819?s=61 https://x.com/robert_zubrin/status/1256571091100725249?s=61 https://x.com/robert_zubrin/status/1192796619894185987?s=61 https://x.com/robert_zubrin/status/1192785320011415552?s=61 https://x.com/robert_zubrin/status/1192444670191624192?s=61 https://x.com/robert_zubrin/status/1178265541342941184?s=61 https://x.com/robert_zubrin/status/1127195497964478464?s=61 https://x.com/robert_zubrin/status/1127283625366568961?s=61 And finally: Bob Clark

It might be Robert L Forwards idea of reflecting back the light from the main sail to a smaller sail would allow you to slow down the smaller sail to stop at the destination. A problem though is Dr. Forward intended this for the case of laser propulsion where the reflected light could be focused onto the receiving sail. This might not work for the non coherent light from the Sun. However, actually for the solar gravitational lens it still works as long as you are on a line extending out from the SGL so you may not need to slow down for that case. Bob Clark

The Parker Solar Probe recently survived its closest flyby of the Sun at only 0.04 AU. This gives confidence that the proposal to achieve high speed of a solar sail using a close flyby of the Sun using the ultralight, but high temperature material aerographite can work: Interstellar Sails: A New Analysis of Aerographite by Paul Gilster | Sep 27, 2023 | Sail Concepts | https://www.centauri-dreams.org/2023/09/27/interstellar-sails-a-new-analysis-of-aerographite Such a solar sail could reach a speed of 2%c, 6,000 km/s, using this close flyby. At this speed it could reach the solar gravitational lens(SGL) at 550 AU in only 6 months, and ‘Oumuamua in only 11 days(!) The implications are stunning. Aerographite is an existing material. Then this means we currently have this capability. Telescopes placed at the solar gravitational lens(SGL) would have the ability to amplify the images of an Earth-sized exoplanet by 100 billion times. It could resolve continent-sized features on such a planet. ‘Oumuamua is an interstellar object whose unusual motions led some to speculate it could be of artificial origin. Then we now have the capability to directly observe Earth-sized exoplanets in other star systems and to determine features on an interstellar object that came into our solar system which may have been artificially produced. Bob Clark

The new era of heavy launch. By Gary Oleson The Space Review July 24, 2023 https://www.thespacereview.com/article/4626/1 The author Gary Oleson discusses the implications of SpaceX achieving their goal of cutting the costs to orbit to the $100 per kilo range. His key point was costs to orbit in the $100 per kilo range will be transformative not just for spaceflight but because of what capabilities it will unlock, actually transformative for society as a whole. For instance, arguments against space solar power note how expensive it is transporting large mass to orbit. But at $100/kg launch rates, gigawatt scale space solar plants could be launched for less than a billion dollars. This is notable because gigawatt scale nuclear power plants cost multiple billions of dollars. Space solar power plants would literally be cheaper than nuclear power plants. Oleson makes other key points in his article. For instance: The Starship cost per kilogram is so low that it is likely to enable large-scale expansion of industries in space. For perspective, compare the cost of Starship launches to shipping with FedEx. If most of Starship’s huge capacity was used, costs to orbit that start around $200 per kilogram might trend toward $100 per kilogram and below. A recent price for shipping a 10-kilogram package from Washington, DC, to Sydney, Australia, was $69 per kilogram. The price for a 100-kilogram package was $122 per kilogram. It’s hard to imagine the impact of shipping to LEO for FedEx prices. Sending a package via orbit transpacific flight would not only take less than an hour compared to a full day via aircraft, it would actually be *cheaper*. Note this also applies to passenger flights: anywhere in the world at less than an hour, compared to a full day travel time for the longer transpacific flights, and at lower cost for those longer transpacific flights. Oleson Concludes: What could you do with 150 metric tons in LEO for $10 million? The new heavy launchers will relax mass, volume, and launch cost as constraints for many projects. Everyone who is concerned with future space projects should begin asking what will be possible. Given the time it will take to develop projects large enough to take advantage of the new capabilities, there could be huge first mover advantages. If you don’t seize the opportunity, your competitors or adversaries might. Space launch at FedEx prices will change the world. These are the implications of SpaceX succeeding at this goal. However, a surprising fact is SpaceX already has this *capability* now! They only need to implement it: SpaceX routine orbital passenger flights imminent. http://exoscientist.blogspot.com/2024/11/spacex-routine-orbital-passenger.html Bob Clark

Actually, at least for the reusable Starship V3, using the small 3rd stage/lander is cheaper on that measure as well. The current Artemis plan would make the Starship HLS expendable. But instead with a 3rd stage/lander only the cheaper, smaller lander would be expendable. It might be even the lander can be reusable as well. Zubrin has made this point in regard to Mars missions as well. By sending the entire Starship to Mars it is out of use for two years. While using the smaller lander as the stage that goes to Mars, Starship can be returned to Earth for its many reuses after launch, and only the smaller lander is out of use for two years. There really is no logical reason to do multiple refueling, multiple launches for a single Moon or Mars mission when it can be done in a single launch format. Bob Clark

Thanks for that discussion but a key option was missed: using a 3rd stage/lander. Robert Zubrin has made this argument numerous times that addition of a 3rd stage results in a more efficient architecture for the Moon or Mars. In fact it would result in single launch missions for both the Moon and Mars, no orbital refueling flights required at all. The expendable Starship at 250-ton capability and the reusable Starship V3 at 200-ton capacity have about twice the capacity of the Saturn V so would have about twice the capacity for single flight, round-trip missions to the Moon when using additional stage(s). And Zubrin’s Mars Direct approach could do Mars missions using two Saturn V class launches. So expendable Starship or Starship V3 could also do that in single launch format when using additional stage(s) Bob Clark

Yes. In that SpaceWatch.Global interview Zubrin makes clear the Starship upper stage is what Elon says could be made for $10 million, not the Superheavy. But a key point is you can make smaller launcher using the Starship itself as the booster, with a smaller “mini-Starship”, if you will, as the upper stage. At least I assume that is what Zubrin is arguing. I don’t think he would be arguing in favor of a SSTO. Then, if you run the numbers such a smaller two-stage vehicle could be a 100-ton class launcher as an expendable. Bob Clark