Exoscientist

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

Musk prefers the vertical, propulsive landing approach. If you look at the discussion here: http://yarchive.net/space/launchers/horizontal_vs_vertical_landing.html one side of that debate argues you can save significant mass using the winged, horizontal landing approach over the vertical landing approach. When you take into account carbon composites can save significantly on the wing weight, it may very well be the reusable SSTO can carry significant payload as a point-to-point transport. Bob Clark

This simulation shows the upper stage of the original ITS from last year as an expendable SSTO could get a total 190 metric tons to orbit, including the stage and the payload: https://www.youtube.com/watch?v=Kzyzwr-5XXY Since the stage was estimated to weigh 90 tons, this would mean 100 metric tons payload as an expendable. Then the question is how much mass would be taken up for propellant for return using vertical landing? If the new version of the upper stage, the BFS, is about half size, then we might estimate the expendable SSTO payload as ca. 50 metric tons. But again the question is how much would be taken up for reusability systems? The discussion here is this should be a small percentage of the landed weight: http://yarchive.net/space/launchers/horizontal_vs_vertical_landing.html So it should still be feasible with significant payload as a reusable. Bob Clark

That's puzzling. This screen grab from the IAC 2017 video only shows two sea level engines: Bob Clark

Thanks for that. I like the part where he acknowledges that the new version of the upper stage which he calls BFS will be given hovering capability for landing on Earth. Bob Clark

I hadn't heard of trisomy. But I heard of cases where fraternal twins actually had different fathers. In this case separate eggs were fertilized by different sperm during the same pregnancy. https://en.wikipedia.org/wiki/Superfecundation Bob Clark

The highest level solid rockets available to amateurs are called the O-class. They weigh in the range of 30+ kg: and can have 4,000+ lbs. thrust: http://www.thrustcurve.org/browser.jsp?1category=l3&2class=O You need to be certified by one of the recognized high power rocketry associations to purchase them: http://www.tripoli.org/Certification And you have to get clearance from the FAA to launch them, since they have to clear the airspace above the launch site before you launch. Bob Clark

Yes, liquid fuel rockets are technically more challenging than solid rockets. Also, they have to construct their own engines rather than.taking them "off-the-shelf". Bob Clark

It's the same type of construction of the lower stages and the upper stages. It's a general principle that the first stage makes up the bulk of the size and cost of a launcher that's true across very different kinds of launchers. With staging for orbital rockets, the upper stages are a fraction of the size of the first stage, so you wouldn't multiply the size of the first stage to get the size of the full rocket. Bob Clark

These solid motors also have very high thrust/weight ratios in the range of 25 to 50 and above. Good for reducing gravity drag but bad since this increases air drag. Bob Clark

As an educator I loved that segment in the "Amateur Rocketeers Reach For The Stars - KQED QUEST" video dealing with the high school students building their own rocket. The professional engineers who founded this program instructing high school students in rocketry deserve great kudos for inspiring interest in STEM fields in young people. I'm suggesting that the rockets that can be possible for young people under professional supervision to build extends to orbital rockets. But then Robert Heinlein has said, once you reach orbit you're halfway to anywhere. So it would also be possible for them to build cubesat-based planetary missions. Imagine the enthusiasm for science generated among young people who built their own planetary orbiters and landers. But beyond that there are technical reasons why it is important that such high altitude rockets can be built by amateurs. Elon Musk has said the cost of the first stage is 3/4ths the cost of the Falcon 9. And this is the case in general that the first, booster stage makes up the largest bulk of the cost. So finding the cost that amateurs spent in constructing these high altitude rockets gives a good way of estimating the full cost of the launch system. Also, one of the questions raised about this solid rocket proposal for an orbital rocket is the fact that the solids used have a quite high thrust/weight ratio, in the range of 25 to 1, while orbital rockets typically have a T/W ratio in the range only about 1.2 or so. The high T/W ratio for the solid would mean it would reach high speed while still in the dense, lower part of the atmosphere. On top of that also smaller rockets generate more drag than larger rockets. These two facts mean this rocket would generate significantly more air drag than standard orbital rockets. The solid rocket though would have the advantage that the gravity drag would be reduced. Accurate trajectory simulations need to be done to see how these two competing effects would effect the final delta-v of the rocket. Therefore the fact the amateur built rockets even with the high T/W solids can reach high altitude of 100,000+ feet suggests that the remaining stages, at greatly reduced air density of 1% or less of sea level, could attain the required delta-v for orbit. Bob Clark

It's remarkable what some amateurs groups have done in high powered rocketry. For instance, In this video at about the 7:15 point is discussed a rocket built by high school students under the supervision of professional rocket engineers consisting of two stages and a clustered first stage: Amateur Rocketeers Reach For The Stars - KQED QUEST. Their rocket reached Mach 2.8 and 100,000 feet altitude. And this amateur group built a rocket that crossed the Karman line of space of 100 km altitude: DIY Rocket Fly Into Space - Amateur High Power Solid Model Rocket Launch to Space. Bob Clark

This indicates even with "small" launchers the development costs are so high that even with an easily correctable defect, there is no scheduled retest of the concept. Bob Clark

I hadn't heard of that one before. Do you have a reference? These ain't your "Estes size" solid motors bud: http://www.nar.org/high-power-rocketry-info/ Bob Clark

If they are orbital, they are still million dollar rockets. Both DARPA and the U.S. Army funded programs to develop orbital rockets for payloads in the few 10's of kilos range for the radical low price of $1 million per launch. Both those programs failed. These were using liquid-fueled rockets though, which are more complicated than simple solid rocket motors. I'm suggesting using essentially "off-the-shelf" high power solid motors such a low cost, small payload rocket can be produced. Bob Clark

Two possibilities, one liquid fueled: Orbital rockets are now easy. https://exoscientist.blogspot.com/2015/08/orbital-rockets-are-now-easy.html the other solid-fueled: Orbital rockets are now easy, page 2: solid-rockets for cube-sats. https://exoscientist.blogspot.com/2017/08/orbital-rockets-are-now-easy-page-2.html Bob Clark

The Super Strypi was planned to get 300 kg to LEO for $15 million, or about $50,000 per kilo. The large aerospace companies are just not interested in launchers that don't cost in the millions of dollars range. Rail-launched Super Strypi Rocket Packed with Cubesats Fails in Debut. by Mike Gruss — November 4, 2015 The Super Strypi launcher was expected to be able to place as much as 300 kilograms of payload into low Earth orbit. Based on designs developed by the U.S. Department of Energy’s Sandia National Laboratory in New Mexico as part of nuclear testing programs dating back to the 1960s, the Super Strypi ultimately was expected to cost about $15 million per mission, officials had said. http://spacenews.com/rail-launched-super-strypi-rocket-packed-with-cubesats-fails-after-liftoff/ The small solid motor launchers in contrast would only get about 1kg to 10kg to orbit for a cost only in the tens of thousands of dollars range. You're right about the guidance being a significant issue. But some earlier rockets such as the rocket that sent the first Explorer 1 satellite to orbit had simple systems that used spinning platforms and the gyroscope effect to maintain stability. They didn't have the accuracy in their final orbital position but they did allow the rocket to reach orbit. Producing your own ammonium perchlorate can indeed be quite dangerous if you don't know what you're doing. That's why I suggest in the blog post using the commercially produced solid motors sold to amateurs who do experimental flights in high power rocketry. Bob Clark