Exoscientist

The delta-v to orbit is commonly taken to be 9.1 km/s for equatorial orbits because you get a 400 m/s boost by the Earth’s rotation for free. Also the dry mass for the BFS upper stage is given as 85 tons by wiki. But this is the version with the passenger quarters for 100 Mars colonists. The tanker version would weigh much less without the passenger quarters, perhaps only in the 50 ton range. Bob Clark

You get 400 m/s for free by launching near the equator, such as from Cape Canaveral. Taking this into account, the delta-v to LEO is often taken to be about 9,100 m/s or 30,000 ft/s: From Modern Engineering for Design of Liquid-Propellant Rocket Engines, p. 12.https://books.google.com/books?id=TKdIbLX51NQC&pg=PA12&source=gbs_toc_r&cad=4#v=onepage&q&f=false Because of the exponential nature of the rocket equation that 900 m/s difference between 10 km/s and 9.1 km/s accounts for a significant amount of payload. Bob Clark

The latest NASA budget suggests the Europa Clipper, an orbiter mission to the Jovian-system to study Europa, won’t fly on the SLS, but instead on commercial rockets: https://mobile.twitter.com/SpcPlcyOnline/status/1105131948903747584 However, instead of just an orbiter mission, by using commercial rockets, we can do it as an actual lander mission at a fraction of the cost of the SLS-based orbiter mission. In fact, it could be so low cost so as to be fully privately financed and at a profit. http://exoscientist.blogspot.com/2015/02/low-cost-europa-lander-missions.html This written in 2015. Since then the F9 has been increases in payload nearly 50% and the FH by nearly 25%. So the landers could be made larger or more capable in-space stages could be used to shorten the flight time. I had assumed that the Falcon Heavy couldn't carry the full Europa Clipper orbiter at 6 ton gross mass to Jupiter. And speculation had been the addition of a Star 48 solid-stage would allow the EC mission on a FH but it would require an Earth gravity-assist that would lengthen the flight time to 6 years. However, I was surprised when I ran the numbers that the upgraded version of the FH could do the mission with plenty of margin with the addition of one of the existing cryogenic upper stages. The extra margin would actually allow you to shorten the flight time from the 2.7 years expected with the SLS. Bob Clark

Didn't know the 7075 was known that long. The X-33 engineers did try replacing the carbon composite with aluminum-lithium, so perhaps the advantage of 7075 over aluminum-lithium was not sufficient to justify its expense. Note though there are now alloys significantly stronger than 7075 as well. Bob Clark

We now know that even reusability of a two-stage vehicle (TSTO) costs significantly in payload. For instance, the Falcon 9 loses 30% of it’s payload with first stage only reuse and 40% payload is lost with full reuse. Since the SSTO doesn’t have the expense of an upper stage, there really has to be an accurate reassessment which comes out ahead on a cost per kilo basis. I emphasize cost per kilo because even though the TSTO will carry more payload it will lose more payload on reusability and cost more because of the upper stage. Bob Clark

Correct about the complex shape being the problem. Actually, cylindrical carbon composite tanks were well understood even then. That’s what led Lockheed engineers to think they could solve the case of conformal, i.e., following the shape of the aircraft, tanks. Unfortunately, it turned out with carbon composites the weight turned out worse than metal tanks rather than saving weight. However, recent high strength metal alloys are even more weight saving than carbon composites. This means we can now build the X-33 with even better than the original expected performance and build the VentureStar with even better than the originally expected payload as an SSTO: DARPA’s Spaceplane:an X-33 version, Page 2. https://exoscientist.blogspot.com/2018/06/darpas-spaceplane-x-33-version-page-2.html Bob Clark

ESA may have no choice in the matter. SpaceX is progressingly rapidly to reusability and to reduced costs. The current version of the Ariane 6 does not allow reusability. By its scheduled time of full operation in 2023, it may already be obsolete. Bob Clark

ESA can get a low cost, reusable version of the Ariane 6 just by adding a second Vulcain to the Ariane 5. Moveover, without needing the solid side boosters, this can be used to finally give Europe an independent manned spaceflight capability. Multi-Vulcain Ariane 6. https://exoscientist.blogspot.com/2018/02/multi-vulcain-ariane-6.html Bob Clark

The delta-v to orbit is dependent on your thrust/weight ratio. What would your delta-v be if the T/W was, say, 1.5? Bob Clark

I was surprised in the study by Hein, et.al. that they didn’t consider the possibility of using ion drive. Ion drive has already been used by NASA for space probes to reach high velocity. Oumuamua was moving at a speed of about 25 km/s. The Dawn mission to Ceres could manage 10 km/s via ion drive. And the Atlas V and in-space stages used for the launch of New Horizons were able to give the spacecraft a speed of about 12 km/s. So using this architecture of chemical propulsion to leave Earth at an initial high speed then using ion drive analogous to that used on Dawn, we could already get 22 km/s. Tweaking both the chemical stages and ion drives should be able to gives us the 25 km/s needed to catch Oumuamua. Bob Clark

I agree we can't be sure of the actual dry mass of the BFR tanker, but it is certainly less than the 85 tons of the BFR spaceship since it won't contain the passenger quarters and supplies for 100 colonists on a six month journey to Mars. Bob Clark

According to Elon the reusable BFR tanker would have a payload of, perhaps, 10 to 15 tons. That's an estimate because all he says is as a reusable the BFR tanker payload is more than an order of magnitude less than that of the full two-stage reusable BFR's payload which is 150 tons. Bob Clark

It comes from consideration of the dry mass of the original ITS spaceship, as in this image: Elon in the video says the dry mass of the half-size BFR spaceship is 85 metric tons. If the dry mass were simply half that of the ITS version it would have been 75 tons. So the actual dry mass is larger than just by proportional scaling by a factor of 85/75 = 1.133. In that image above the ITS tanker is given as having a dry mass of 90 tons. So if the BFR tanker were proportionally half-size, that would put its mass at 45 tons. But the additional scaling factor of 1.133 would put it at 51 tons, which I rounded off to 50 tons. Bob Clark

You have to keep in mind there are two versions of the upper stage, the spaceship and the tanker. It is unlikely Elon is referring to the payload of the spaceship version when he is making that payload comparison to the full two-stage BFR, because clearly he is taking the tanker version as the upper stage in the two-stage payload estimate. For instance see this image from the presentation: From the context of his discussions in the video he is also talking about the reusable payload because that is the only operational mode they are considering for the BFR. I like the idea you are creating a small scale version of the stage with your SFR. However, I think you taking too large an estimate of the BFR tanker dry mass to estimate the dry mass of your SFR. Bob Clark

A quarter-scale version of the BFR upper stage can serve as an SSTO able to carry a Dragon 2 to orbit: A Small Raptor Spaceship. https://exoscientist.blogspot.com/2017/10/a-small-raptor-spaceship.html Bob Clark

Actually what he said in the video presentation is that the BFR upper stage can be SSTO but the full two-stage BFR can carry more than an order of magnitude more payload. Since the payload for the reusable two-stage BFR is 150 tons, the payload for the BFR upper stage as a reusable SSTO might be, say, 10 to 15 tons. Part of the calculation in my blog post was to suggest using winged landing for the SSTO you might lose less of the payload in the reusable case. I estimate less than 10% loss with winged landing, as opposed to 70% to 80% loss with the vertical, propulsive landing approach. If this is true you might want to make also the lower stage do a winged landing. For instance if the 10% loss using winged landing also holds for the two-stage BFR, then instead of losing 40% payload from 250 tons to 150 tons for reusability, it would be only 10% loss from 250 tons to 225 tons. So it is important to do such trades between the different landing modes to see which would result in the smallest loss in payload for reusability. Bob Clark

The SpaceX BFR tanker can serve as a reusable SSTO by switching to a winged, horizontal landing mode: SpaceX BFR tanker as an SSTO. https://exoscientist.blogspot.com/2017/10/spacex-bfr-tanker-as-ssto.html Bob Clark

Perhaps they see the advantage of such a BFR point-to-point transport. http://www.astronautix.com/i/ithacus.html Bob Clark

Perhaps we will see such lander with this new program: NASA preparing call for proposals for commercial lunar landers by Jeff Foust — September 7, 2017 http://spacenews.com/nasa-preparing-call-for-proposals-for-commercial-lunar-landers/ Bob Clark

Very good point. In Elon's description of the Interplanetary Transport System (ITS) from last year both tanker and spaceship version of the upper stage were intended to be reusable multiple times, so likely the heat shield mass was already included in the quoted vehicle mass values: Making Humans a Multi-Planetary Species. Musk Elon. New Space. June 2017, 5(2): 46-61. https://doi.org/10.1089/space.2017.29009.emu Likewise, the descriptions of the tanker and spaceship upper stages in this years BFR version were also already described as being reusable, so the heat shield mass was also included there. In any case, you get a high mass for the payload of the BFR tanker as an expendable SSTO. Estimating the BFR tanker dry mass as approx. half the dry mass of the ITS version, as Elon confirmed with the spaceship case, the tanker dry mass would be in the range of 45 to 50 metric tons, and the payload as expendable SSTO would be in the range of 55 to 50 tons. But this puts it as an expendable SSTO in the payload range of the Falcon Heavy while being in the same size range of the expendable Falcon Heavy. So this SSTO would get the same payload fraction as a 2 and 1/2 stage vehicle. Moreover, judging from the fact the ITS tanker upper stage was to cost $130 million production cost, the half size BFR tanker might only be $65 million, so it would be half the cost of the Falcon Heavy. But the Falcon Heavy as an expendable launcher already would be a significant cut in the cost to orbit. So the BFR tanker as an expendable SSTO could be a great reduction in the cost to space, compared to current values. But Elon wants to go beyond expendables and has implied the reusable version of the BFR upper stage would only get perhaps in the range of 10 to 15 metric tons payload (by saying it's an order of magnitude less than the full BFR 150 ton reusable payload.) That loss in payload seems high, 40 tons, nearly the size of the entire vehicle dry mass, presumably because of the size of the propellant that needs to be kept on reserve for landing on return. I'd like to see a trade study of the payload of instead going with wings for a horizontal landing. Wings typically take up only 10% of an aircraft dry mass. Then with carbon composites, that would be cut to less than 5% of the landed (dry) mass. Keep in mind the loss in payload with vertical, propulsive landing is nearly 100% of the vehicle dry mass. Also, going with short, stubby wings as with the X-37B, you can make the wing weight even less: The areal size of the wings in that case would also be less than that of bottom area of the BFR tanker, perhaps only 1/3rd to 1/2 the areal size. So the increase in heat shield mass would only be at most 1/2 that of the approx. 8,100 kg mass of the current heat shield, so perhaps an extra 4,000 kg. But actually the addition of wings gives a gentler glide slope so probably the heat shield thickness could be reduced. The result might even be the total heat shield mass would be reduced by adding wings. Bob Clark

I'm getting surprisingly high values for the thermal protection of the BFR upper stage, either spaceship and tanker versions. I'm using the fact as indicated in the wiki page on the BFR that it will use the Pica-X thermal protection material. Several references give its density as about 0.25 gm/cc = 250 kg/m3, and the thickness as on the Dragon 2 as 7.5 cm, 0.075 m, about 3 inches. The BFR upper stage has a length of 48 meters and a width of 9 meters. The top part of the stage is conical, so the bottom surface is not rectangular but for simplicity I'll approximate the bottom area to be covered by thermal protection as a rectangle. So the area that needs to be covered is approx. 48*9 = 432 m2. Then the volume of the thermal protection material is 432 * 0.075 = 32.4 m3. At a density of 250 kg/m3, that amounts to a mass for the thermal protection of 32.4 * 250 = 8,100 kg, which is a surprisingly high addition to the dry mass of 85 tons for the spaceship upper stage or to the 50 tons for the tanker upper stage. One possibility, is the thickness of the PICA-X for the Dragon 2 is coming from the fact it is doing a ballistic reentry, thus generating high heat. However, the BFR upper stage will be doing a more gentle gliding reentry. So perhaps the thermal protection will only need to be half as thick. Bob Clark

I like the idea of making an additional even smaller upper stage for this new system. However, when doing your scaling you should consider that the tanker version of the BFR upper stage likely has a dry mass only in the range 45 to 50 metric tons. This is based on how much smaller the tanker version also was than the spaceship version for the original ITS. Then when considering the possible payload in using your new upper stage, the dry mass would be smaller than your estimates, so the payload would be higher. Bob Clark

Here's the description of the original ITS upper stage, both spaceship and tanker versions: And here's the description of the BFR spaceship, half size to the ITS version: You see the BFR spaceship is about half the listed value for the ITS spaceship. Actually during the video Musk says the design mass was 75 tons, but the 85 tons was allowing for weight growth. So it is plausible the BFR tanker is half the mass of the ITS version or a little more, ca. 45+ tons. Bob Clark

That is true for the mass of the propellant tanks that it scales with the size of vehicle. Other components do also such as the mass of the engines. But some do not such as tank insulation which scales more closely to surface area. See this report that gives vehicle component scaling relationships: Mass Estimating Relations. • Review of iterative design approach • Mass Estimating Relations (MERs) • Sample vehicle design analysis http://spacecraft.ssl.umd.edu/academics/483F09/483F09L13.mass_est/483F09L13.MER.pdf By the way, it is interesting that the author, head of the department of aerospace engineering at the University of Maryland, concludes that SSTO's are possible using hydrolox propellant. Bob Clark

The 70 metric ton dry mass estimate seems high to me. The original ITS tanker, about twice as big, had a 90 metric ton dry mass. So I would expect the BFS dry mass to be closer to 45 metric tons. Bob Clark