Exoscientist

Funnily enough Zubrin’s proposal of staging off the Starship accomplishes a Mars mission at only ~200 tons IMLEO at a single launch and it is better than the SpaceX plan:

Good point. Apollo did it at 3 million kilo launch mass. SpaceX 50 years later, with advances in materials and propulsion, proposes to do it at 50 million kilo launch mass and perhaps as much as 100 million kilo launch mass if Zubrin is right. Another way of saying this is the measure of large space projects known as “initial mass to LEO”, IMLEO, both stage mass and payload mass delivered to LEO. This was a common way of comparing Mars architectures. It can also be used for comparing any large space project. For Apollo it was 130,000 kilos. For the SpaceX plan it would be 2,500,000 kilos and perhaps as much as 5,000,000 kilos according to Zubrin. This for a single lunar mission. Robert Clark

We are agreed on that. But the thing to remember is the only reason why the Gateway was proposed was that without an extended upper stage, the SLS didn’t have enough power to send the Orion to low lunar orbit. But NASA couldn’t afford the extended upper stage proposed by Boeing. But with an upper stage already built and flying reliably by the ESA for decades such a stage could be made available relatively cheaply, i.e., compared to building such a stage entirely from scratch. Robert Clark

The intent is to make manned spaceflight to the Moon routine, or at least as routine as flights to the ISS are now. That’s not going to happen when it takes 8 to 16 launches of a Saturn V class launcher for a single mission to the Moon. By the way, Robert Zubrin estimated it would actually take ~20 launches of the Starship under the current proposal that has the Starship meetup with the astronauts in the Orion capsule at the Gateway to take them to the Moon and then back again to the Gateway, and then from there using the Orion for the trip back to Earth. This is because of the large amount of fuel needed for going back and forth to the Gateway: Op-ed | Toward a coherent Artemis plan. by Robert Zubrin — May 18, 2020 https://spacenews.com/op-ed-toward-a-coherent-artemis-plan/https://spacenews.com/op-ed-toward-a-coherent-artemis-plan/ It is extremely important to keep in mind the only reason why NASA proposes using the Gateway in the first place is because the SLS without an extended upper stage does not have enough power to take the Orion to low lunar orbit about the Moon and back again. With an extended upper stage, that eliminates the need for the, much derided, Gateway with its great added cost and time delay, and eliminates the need for the ~20 flights for the Starship. The SLS could mount the mission to the lunar surface by itself with no need for the Starship. Note the ESA is actually more in favor of a lunar base than is NASA. By the supplying an Ariane 5/6 to serve as the extended upper stage for the SLS that would go a long way to insure that it happens. Bob Clark

Alien hunters should look for city lights from 'urbanized planets,' study suggests. By Leonard David published about 21 hours ago Lights from alien cities are an intriguing potential technosignature. https://www.leonarddavid.com/search-for-extraterrestrial-technology-city-lights-from-urbanized-planets/ Bob Clark

Even Apollo-era TPS was at ~15% of landed mass. Landing gear at about 3%, so < 20% of landed mass for both. Even subtracting off from the payload a mass of 20% of landed mass could still give you a reusable SSTO vehicle at over 100 ton payload. Robert Clark

Elon also says in that video the tank mass is at 80 tons. Then if SpaceX can reduce the tank wall thickness from 4 mm to 3 mm as they expect that’s cutting 25% off the tank mass, or 20 tons. Then the expendable SSTO Superheavy payload could be 175 tons to 135 tons, depending on if the SuperHeavy dry mass, prior to the tank mass saving, were 160 tons to 200 tons. The grid fins also can be reduced in weight. Elon was unhappy with their mass in the video at 3 tons each for a total mass of 12 tons. Their mass can be reduced to a fraction of this weight using ceramics, subtracting an additional 10 tons from the dry mass, increasing the expendable payload to somewhere in the range of 185 tons to 155 tons. Robert Clark

Elon seems to be acknowledging the larger number of engines on this vehicle increased the likelihood of this occurring: 12:29 AM · Jul 12, 2022·Twitter for iPhone 296 Retweets 52 Quote Tweets 4,583 Likes

Operational simplicity for rapid deployment if needed. Also the same vehicle can be used to come back or rescue if not needing first stage left back at the launch site. Robert Clark

I’m not actually sanguine about the 33 engines on the Superheavy being in common use for passenger flights, either for a TSTO or a SSTO. Better would be to develop the 9-engined Starship either as a first stage booster, with a mini-Starship as upper stage or as an SSTO. Robert Clark

Thanks for the link to the video. In addition to the dry weight range it also gives the propellant load as 3,600 tons, higher than the 3,400 tons I used. Also the Raptor 2 will have greater thrust to 230 tons at sea level. I’ll update the Silverbird estimator with these numbers. Since the estimator uses the vacuum thrust I’ll estimate it as proportionally larger by the 358s vacuum Isp compared to the 330s sea level Isp. That brings the vacuum thrust to 230*(358/330) = 249.5 tons, so to 8,230 tons for 33 engines. The input to the Silverbird estimator, still using the 140 ton dry mass, is: And the result for expendable payload would be: If the dry mass is 160 tons, then the expendable payload according to the estimator would be 155 tons. And if the dry mass is 200 tons then the expendable payload would be 115 tons. Using the 10% of dry mass estimate for the payload loss due to reusability, it would still be at 95 tons reusable payload even for the 200 ton dry mass estimate for the SuperHeavy. By the way, I wonder if the nozzle exit area can be increased due to the higher chamber pressure. Could the sea level Raptor have the vacuum Isp increased to 372s? Small Isp increases are known to allow large increases in payload. The Silverbird estimator then gives a payload estimate of 208 tons. Robert Clark

Do you have a reference for the 200 ton dry mass of the SuperHeavy? The earlier ITS incarnation had a 25 to 1 mass ratio for the first stage: For the ITS first stage at 6,975 ton gross mass and 275 ton dry mass, that’s a mass ratio of 25.4 to 1. For the SuperHeavy at a propellant load of 3,400 tons a 25 to 1 mass ratio gives a dry mass of 140 tons. Robert Clark

A fair point. Doing a google search, on “PICA-X”, “density”, the first google result that pops up gives it as .27 gm/cc, 270 kg/m3. But we need to know the thickness used. This video shows an edge view at about the 1:20 point: It appears to be about 1 inch thick, 2.5 cm. Then the volume of the PICA-X to cover only the windward half of the cylindrical vertical side of the Superheavy is: (1/2)*Pi*9*70*.025 = 24.7 m3. Then the mass is 270*24.7 = 6,700 kg, 4.4% of the dry mass of the SuperHeavy stage. Robert Clark

Rocket engineers regard reducing dry mass and increasing propellant load like gold. Yet by requiring that large amount of propellant to remain onboard the first stage unused during ascent is doing the opposite of both. Losing 50% of payload is a huge loss for payload. Actually, not. There are lots of ways of having extra deadweight on ascent for a stage. For instance suppose you gave a stage wings for return. Then as far as payload to orbit is concerned that is deadweight. You could still use all the propellant though. Robert Clark

The fully reusable Superheavy/Starship is variously estimated to get from 100 to 150 tons to LEO. The expendable SuperHeavy gets 160 tons to orbit per the Silverbird Astronautics payload estimater. But only ~10% of dry mass used for reusability systems, or ~15 tons for the 150 dry mass SuperHeavy. So reusable payload at ~145 tons for the SSTO SuperHeavy. This is in the range already for the reusable SH/SS. BUT not needing the upper stage its actually better than the reusable TSTO on a per gross mass basis. Robert Clark

Yes. Look up “magnetars”. On Earth strong magnetic fields can be created for a short period explosively Robert Clark

I’ve acknowledged the expendable TSTO gets more payload to orbit than the expendable SSTO. The point I’m making is the reusable SSTO meets or exceeds the reusable TSTO payload. The reason is the fully reusable TSTO loses 50% of its payload due to the large amount of propellant that has to remain unused on ascent to orbit in order to cancel out the first stage forward motion and to boost it back to the launch site. Note this unused propellant is doubly disadvantage as far as orbital payload is concerned. First, it adds deadweight on ascent to orbit, and secondly it reduces the propellant load that can be used for that ascent. Elon has stated this reduction in payload on full reusability numerous times: Elon Musk @elonmusk Replying to @PPathole and @SpaceX Optimized, fully-reusable Starship is ~150t to same reference orbit as Saturn V. In expendable mode, Starship payload would be 250t to 300t. 4:41 PM · Mar 9, 2022·Twitter for iPhone 747 Retweets 56 Quote Tweets 12.5K Likes In contrast, the SSTO loses less than 10% of the dry mass on adding reusability systems, allowing it to meet or exceed the payload to orbit of the reusable TSTO. Robert Clark

The key point is the large amount of payload lost for a TSTO due to unused propellant in the first stage during the ascent to orbit, since it has to be used for return to launch site. Elon has acknowledged this cuts full reusability payload by 50%. When you consider the lost payload for a SSTO due to reusability can be less than 10%, the reusable SSTO beats the reusable TSTO. The estimates of the dry mass of the Starship always is for the passenger version with the passenger quarters for ~50 colonists to Mars. That would not be the mass for the tanker version with just a big empty space where the passenger quarters would have been. SpaceX acknowledged this for the prior version the Interplanetary Transport System: Note the first stage already has a mass ratio of 25 to 1, sufficient for SSTO with significant payload. And the upper stage tanker version has mass ratio of 30 to 1, and would also be at 25 to 1 mass ratio when given further engines for ground launch. Robert Clark

I forgot that SpaceX does want to add 3 more engines to the Starship to bring it to 9 engines: We’ll replace the vacuum Raptors with all sea level engines. Again using the equation, (thrust) = (exhaust velocity)*(flow rate), a sea level Isp of 330s and propellant flow rate of 650 kg/s gives a sea level thrust of 330*9.81*650 = 2,104,245 N, so 9 would be 18,938,205 N, or 1,900 tons of thrust. Then depending on how much the propellant is increased this might be enough thrust to lift off from ground. But I wanted to check the results using the current propellant load of 1,200 tons using the Silverbird Astronautics calculator. For 9 sea level engines with a vacuum Isp of 358 s, the vacuum thrust would be 358*9.81*650*9 = 20,545,000 N. Using again the 50 ton dry mass estimate and 1,200 ton propellant load for the 9-engined Starship with no passenger quarters the Silverbird calculator gives: Again you have to select the “No” option for “Restartable Upper Stage”, and enter launch inclination of 28.5 degrees to match the Cape Canaveral latitude, so as not to reduce the calculated payload. Then the results are: Quite close to the estimate we got using the rocket equation of 50 tons. This is the expendable payload, but again the reusability systems should subtract less than 10% of the dry mass from this payload. Robert Clark

As I said, in actuality the reusable SSTO gets at or above the payload of the reusable TSTO because of the severe payload loss from the reusable TSTO having to keep a large amount of propellant on reserve in the first stage to cancel out the forward motion then boost back to the landing site. Robert Clark

The ~70 m stage length contains the top and bottom domes and the engine length as well as a forward skirt at the top above the tank. Estimating the top and bottom tank domes as hemispheres, the two together would be a single sphere of 9 m diameter. So subtract that off the 70 m to get the vertical side length of the tanks. The Raptor engines at the bottom are 3.1 meters long so subtract that off as well. Commonly, the engine nozzles of the upper stage extend into the forward skirt of the first stage. However, it appears the Starship upper stage engines are recessed into the upper stage, so the nozzles do not extend down below the stage. Still the forward skirt has some length. I’ll estimate it as 2 m. Then the actual length of the vertical part of the first stage tanks might be 70 - 9 -3 -2 = 56 m. The volume of the vertical tank walls, i.e., not counting the domes is Pi*(diameter)*(thickness)*(length) = 3.14*9*.004*56 = 6.3 m3 . The density of stainless steel is in the range of 7,850 kg/m3. The vertical tank walls would mass 6.3*7850 = 49,500 kg. To this we would also have to add the mass of the two domes but these usually are a fraction of the vertical tank wall mass. If the wall thickness can be shaved down to 0.003 m then we can cut perhaps 12,000 kg off the tank mass. Robert Clark

By the way, an irritation of mine is that SpaceX is so completely focused on the idea the SuperHeavy/Starship has to be the be all, end all for all launchers. SpaceX is being insightful in recognizing reusability has been the name of the game for any transportation system. But an aspect they are not recognizing is that transports always come in various sizes, going all the way back to the horse-drawn era. In point of fact Starship itself can form an independent launcher without the SuperHeavy and it would have been advantageous to develop a Starhopper-sized stage as well. If they had, then we already would have had a launcher capable of single launch lunar and Mars missions, using the Starhopper-sized stage used as the 3rd stage for the SuperHeavy/Starship. However, the military wanting an orbital troop transport may encourage SpaceX to develop such smaller stages for independent orbital flight. Robert Clark

The military is considering using the Starship for troop transport: TECH & SCIENCE Pentagon Mulls Using Elon Musk's Rockets to Deploy Troops From Space BY ED BROWNE ON 6/24/22 AT 12:33 PM EDT https://www.newsweek.com/pentagon-military-elon-musk-starship-deploy-troops-space-1718969 Clearly for such a use it would be better to have this capability as a single stage. Considering that the military, like NASA, overpays for everything SpaceX could probably get a billion dollar deal for developing this capability. Having a SSTO capability would be a great selling point for this purpose. Robert Clark

Thanks for the response with calculations. The general principle that scaling a rocket up improves mass ratio is coming from the fact that components such as tanks and engines scale approximately linearly with size but things like insulation, wiring, avionics grow at a much lower rate. This was the reason why it was proposed to get lowered launch costs, create a big dumb booster, and to create an SSTO, go large. For using the Silverbird Astronautics calculator, you have to be aware of some quirks of the program. First use the vacuum Isp and vacuum thrust levels in the engine fields since the program takes into account the diminution at sea level. Also, select “No” for the “Restartable upper stage” since otherwise that reduces payload, possibly due to keeping propellant on reserve. Important also is to match the launch angle to the latitude of the launch site. So at Cape Canaveral, set it to 28.5 degrees. For the Raptor engine , several different Isp and thrust levels have been given. I’ll use the vacuum Isp of 358 s for the sea level Raptor. For estimating the vacuum thrust, I’ll use the 650 kg/s flow rate of the Raptor in the Wiki page, then at a 358 s vacuum Isp, using (thrust) = (exhaust speed)x(flow rate), we get a vacuum thrust of 358*9.81*650 = 2,280,000 N. So for 33 engines, 75,300,000 N, ~75,000 kN. Inputting this data into the Silverbird calculator looks like this: And the results is: So a 160 ton payload to LEO as an expendable. I’ve found the Silverbird calculator to be approximately 10% accurate plus or minus. For the fairing, I’ve found inputting it reduces the payload ~10%. But there are various ways of reducing the fairing weight to reduce this lost payload even further. For the PICA-X thermal protection it has reusability for dozens of uses. Even if the dry mass of the SuperHeavy is 200 tons, that still leaves an LEO payload of ~100 tons. Robert Clark

Running some numbers for the SuperHeavy+Starship launcher, I was surprised to get that an expendable SuperHeavy alone could be SSTO with quite high payload. Wikipedia gives the propellant mass of the SuperHeavy as 3,400 tons, but does not give the dry mass. We can do an estimate of that based on information Elon provided in a tweet: Elon Musk @elonmusk Replying to @Erdayastronaut and @DiscoverMag 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·Twitter for iPhone 90 Retweets 32 Quote Tweets 1,498 Likes https://twitter.com/elonmusk/status/1111798912141017089?s=20&t=L-xcKvWnRTmbSDa_YI0OyA This is for a stripped down Starship, no reusability systems, no passenger quarters, and reduced number of engines. But this could not lift-off from ground because of the reduced thrust with only 3 engines plus being vacuum optimized these could not operate at sea level. So up the number of engines to 9 using sea level Raptors. According to wiki the Raptors have a mass of 1,500 kg. So adding 6 more brings the dry mass to 49 tons, call it 50 tons, for a mass ratio of 25 to 1. By the way, there have been many estimates of the capabilities of the Starship for a use other than that with the many passengers, say 50 to 100 , to LEO or as colonists to Mars, for instance, such as the tanker use or only as the lander vehicle transporting a capsule for astronauts for lunar missions. But surprisingly they all use the ca. 100 ton dry mass of the passenger Starship. But without this large passenger compartment it should be a much smaller dry mass used in the calculations. For instance, the Dragon 2 crew capsule dry mass without the trunk is in the range of 7 to 8 tons for up to 7 astronauts. So imagine a scaled up passenger compartment for 50 passengers or more. That passenger compartment itself could well mass over 60 tons. So the dry mass estimate of a stripped down, expendable, reduced engine Starship of 40 tons offered by Elon does make sense. Based on this, an expendable Starship with sufficient engines for ground launch could be SSTO: the ISP of the Raptors for both sea level and vacuum-optimized versions have been given various numbers. I’ll use 358 s as the vacuum ISP of the sea level Raptor. For calculating payload using the rocket equation, the vacuum Isp is commonly used even for the ground stage, since the diminution in Isp at sea level can be regarded as a loss just like air drag and gravity loss for which you compensate by adding additional amount to required delta-v to orbit just like the other losses. Then 3580ln(1 +1200/(50 + 50)) = 9,180 m/s sufficient for LEO. But as of now, SpaceX has no plans of making the Starship a ground-launched vehicle. So we’ll look instead at the SuperHeavy. For an expendable version with no reusability systems, we’ll estimate the dry mass using a mass ratio of 25 to 1, same as for a ground-launched expendable Starship. Actually, likely the Superheavy mass ratio will be even better than this since it is known scaling a rocket up improves the mass ratio. So this gives a dry mass of 136 tons. Then the expendable SuperHeavy could get 150 tons to LEO as an expendable SSTO: 3580ln(1 + 3,400/(136 + 150)) = 9,150 m/s, sufficient for LEO. But what about a reusable version? Reusability systems added to a stage should add less than 10% to the dry mass: ________________________________________________________________________________________________________________ From: [email protected] (Henry Spencer) Newsgroups: sci.space.tech Subject: Re: The cost (in weight) for Reusable SSTO Date: Sun, 28 Mar 1999 22:37:10 GMT In article <[email protected]>, Larry Gales <[email protected]> wrote: >An SSTO with a useful payload using Kero/LOX is easy to do -- provided that >it is *expendable*. All of the difficulty lies in making it reusable... There are people who are sufficiently anti-SSTO that they will dispute the feasibility of even expendable SSTOs (apparently not having read the specs for the Titan II first stage carefully). > (1) De-orbit fuel: I understand that it takes about 100 m/s to de-orbit. That's roughly right. Of course, in favorable circumstances you could play tricks like using a tether to simultaneously boost a payload higher and de-orbit your vehicle. (As NASA's Ivan Bekey pointed out, this is one case where the extra dry mass of a reusable vehicle is an *advantage*, because the heavier the vehicle, the greater the boost given to the payload.) > (2) TPS (heat shield): the figures I hear for this are around 15% of the >orbital mass Could be... but one should be very suspicious of this sort of parametric estimate. It's often possible to beat such numbers, often by quite a large margin, by being clever and exploiting favorable conditions. Any single number for TPS in particular has a *lot* of assumptions in it. > (4) Landing gear: about 3% Gary Hudson pointed out a couple of years ago that, while 3% is common wisdom, the B-58 landing gear was 1.5%... and that was a very tall and mechanically complex gear designed in the 1950s. See comment above about cleverness. I would be very suspicious of any parametric number for landing gear which doesn't at least distinguish between vertical and horizontal landing. > (5) Additional structure to meet loads from differnet directions (e.g., >vertical > takeoff, semi-horizontal re-enttry, horizontal landing). This is >purely > guesswork on my part, but I assume about 8% Of course, here the assumptions are up front: you're assuming a flight profile that many of us would say is simply inferior -- overly complex, difficult to test incrementally, and hard on the structure. >I would appreciate it if anyone could supply more accurate figures. More accurate figures either have to be for a specific vehicle design, or are so hedged about with assumptions that they are nearly meaningless. -- The good old days | Henry Spencer [email protected] weren't. | (aka [email protected]) https://yarchive.net/space/launchers/landing_gear_weight.html The 15% mentioned for thermal protecton(TPS) is for Apollo-era heat shields. But the PICA-X developed by SpaceX is 50% lighter so call it 7.5% for TPS. And for the landing gear ca. 3%, but with carbon composites say half of that at 1.5%. But this would put the reusable payload at ca. 136 tons which is in the range of 100 to 150 tons of the full two stage reusable vehicle! How is that possible? A reusable multistage vehicle has a severe disadvantage. The fuel that needs to be kept on reserve for the first stage to slow down and boost back to the launch site subtracts greatly from the payload possible. But for a reusable SSTO it can remain in orbit until the Earth rotates below until the landing site is once again below the vehicle. Robert Clark