MatterBeam

I see. Do you have a reference for the '50 ton BFR tanker' figure I have seen in many places?

Well you need to the drum walls to be solid! I do not have enough information on how exactly the design will be implemented, or if something has even been tested or seriously conceptualized, so I try to avoid going into too much detail. I don't know if hydrogen at 4000K is 'too hot to handle', or if active cooling makes everything pretty easy. I can't assert that the 5800K layer of liquid rhenium directly under the spot of focused sunlight will be thin or thick relative to the depth of the fluid heat exchanger - this depends on the thermal conductivity rate of the fluid and how stable the currents are. If they are chaotic, there very well may be pockets of >4000K rhenium touching the walls of the drum and etching off sections. If it is without turbulence, then any suck pockets will rapidly mix and equalize their temperature with that of the deeper layers. For now, I'm considering an alternative design that relies on liquid rhenium droplets and/or gaseous lithium.

I wished to create flying wings and re-entry gliders that were not limited to 1m thickness wings. Textures are not an issue for me, nor is FPS at this stage.

The 'order of magnitude' comment is only valid if we consider the 75 ton BFR as expendable. The 85 ton 'likely' dry mass of the BFR means it has only 9.6km/s of deltaV with the vacuum engines. After losses at sea-level, it does indeed have no cargo capacity. So, the 75 ton 'design' dry would allow for a 10 ton payload. However... it must be expendable because even a bare minimum landing deltaV of 400m/s would require 9.86 tons of propellant with the 330s Isp landing engines. The actual deltaV for landing being planned is 3-4 times larger. What do you think of the 'SFR' concept I proposed here and submitted to discussion of the SpaceX subreddit here?

Simple pressure. Hydrogen at 4000K has an extremely high pressure. The holes are small just to make sure that there's a good surface area of contact between the hydrogen and the heat exchanging fluid. The rhenium liquefies at 3459K. It is contained in a vessel at 4000K and is constantly warmed by the 5800K upper layer - there is little chance for any part of it to turn solid again. Even if it did, the fluid is in motion. Solid deposits would be whisked around and cleared from holes by the hydrogen flow. Hydrogen's bouyancy ensures that it escapes the rhenium, because the density difference is extreme. Due to the high temperatures, there is no chance for the hydrogen to dissolve either. The balance between the rhenium reversing through the holes and the hydrogen leaving is an engineering problem, which can be solved usually by just increasing the amount of hydrogen entering the drum. If the rhenium does enter the channels, it will not solidify.

Hello. Is it possible to modify the size limits on the B9 procedural wing parts?

As you leave the inner Solar system, your maximum temperature will remain the same but the thrust drops. If you look at the example spaceships, modern and advanced, they have an average acceleration of 0.32 and 3.1g. Around Mars, this drops to 0.14 and 1.37g. Around Jupiter, it is 0.01 and 0.11g. In other words, solar thermal rockets are able to travel to the outer solar system without too much hassle. If you increase the solar collector area without increasing the engine mass, you'll have too much power to handle around Earth but a good amount around Jupiter. For example, instead of 5 tons of solar collectors, use 15 tons. Around Jupiter, you'll retain 11% of you maximum thrust. The complexity of the solar thermal rocket is no worse than a solar electric rocket, if not less. The period between shutting down the engine and cutting the active cooling is an engineering problem. I can imagine gradual transition between 'On' and 'Off', plus the ability to continue the active cooling by just throwing the hot hydrogen overboard. You'll waste a few dozen kgs of hydrogen... but you were throwing tons per second out of the nozzle a few seconds ago, so its not a big deal! STRs will like raise their orbit to be clear of the planet's shadow. This can be accomplished by raising the altitude or launching into a slanted inclination. Also... planning. The burns can be rather short, so its not a big deal like for solar electric craft that slowly wind out of orbit.

Fused quartz or diamond can be used in situations where the sunlight intensity becomes high and the mass per area is less critical. Active cooling with some of the liquid hydrogen will be needed. Let's say we use fused quartz: Over 95% reflectivity and usable up to 1920K. At that temperature, each square meter of the reflector is losing from a black backside 740kW of heat. If the reflector is only 1.128m in diameter, it is absorbing 0.05*927: 46.35MW from the concentrated sunlight. Sum of heat to be removed? 47.09MW The heat capacity of liquid hydrogen at 1920K is close to 449+1920*15: 29249kJ/kg. So, the final small reflector will need to be fed 1.6kg/s of liquid hydrogen to stay at 1920K. It doesn't necessarily have to be directly cooled by the hydrogen - a heat exchanging loop can be established where helium circulates through and around the reflector and dumps its heat into the hydrogen.

What? I calculated the acceleration of the example ships at 0.32 and 3.1 gravities. Thrust per kg is just another term for thrust-to-weight ratio, that I calculated. The engine is a narrow cylinder. Radiations from the internal surface mostly fall back on the internal surface. Light can only escape from the optical window on top and the nozzle opening at the bottom, which represent a small fraction of the radiating surface area. The 1298 is the number of watts per square meter of collector area, so the thrust calculation you made is actually in Newtons per square meter. The Isp at 1224 is more than enough to travel around the solar system at a rapid rate. Two months to Mars without needed massive propellant tanks, no need for a nuclear reactor or nuclear rockets, no need for a laser - just sunlight. There is also an evolution of the Solar Thermal Rocket possible: the Solar ThermoElectric rocket. Concentrated sunlight can be used as a heat source for a high-temperature Carnot cycle electrical generator. With a gas turbine and a 5800K heat source, we can convert 70% or more of the solar power collected into electricity. This electricity can then be used to power electric rockets. For example, the 5 ton solar collector mentioned in the blog could be feeding the 927MW of heat into a gas turbine and making 648MW of electricity out of that. It is enough to make a 3000 Isp electric engine produce 44kN. The nozzle material would have to be a temperature resistant material such as THC that is actively cooled by liquid hydrogen. High operating temperature means massive amounts of heat from the nozzle can be absorbed by the hydrogen. At 3000K, the hydrogen that starts at 20K absorbs 60MJ/kg, and we have nearly a ton of hydrogen per second to use for active cooling of both the rotating drum and the nozzle. A magnetic nozzle isn't necessary here. Quite right.

I'm not sure what you mean by your first line. The liquid rhenium at high temperatures is a near perfect blackbody. It is radiating at 5800K, so that's 64MW/m^2. The solar collectors are taking in 1298W/m^2, so you'd need a concentration factor of 49433. You make a good point about the parabola. The focal length will be quite long - multiple 'corrective' reflectors and lens will be needed to direct the beam into the small opening of a Solar Thermal Rocket. The Earth-Mars escape capture plan? Are you referring to one of the images I used? They are generally for illustrative purposes, especially when not captioned.

Thank you for taking the time to calculate your own numbers! I mentioned in the blog post that the mass ratio gains from increasing your exhaust velocity is exponential, so every little bit helps. One practical advantage I did not want to elaborate in the blog post is the fact that STRs do not face the nuclear controversy that NTRs do. I also mentioned that STRs can have very good power to weight ratios, especially when compared to anemic solar-electric rockets. This makes them practical for human spaceflight, as you don't want your crew spending weeks spiralling our of low orbit and being blasted full-on by Van Allen Belt radiation. STRs can accelerate quickly enough to gain a meaningful boost from the Oberth effect too! I am not sure what you mean by the 'poles', but maybe you mean the structural support for the mirrors in the images I posted? They could be a problem, but I did state that we will be basing our designs on the lessons from solar sail development. This means lightweight distributed structures that are held in place by tension and centripetal force. At 7g/m, a Mylar sheet only has to endure a force of 0.068 Newtons even when accelerating at a full 1g. The structural requirements of handling 0.068 Newtons per square meter of area are tiny! The 20 minute figure was achieved by dividing the ~3.5km/s departure burn for a Mars mission by the 0.32g average acceleration of the STR being discussed in that example. I am not sure where you got the 150000 ton figure from. The power calculations I used are as follows: 7g/m^2 for a modern Solar Collector dish. 1367W/m^2 of sunlight received. 1298W/m^2 is focused into the engine. Engine efficiency is 80%. Engine power density of 167kW/kg to handle the sunlight input. With these figures, I calculated the thrust as Power * 2/Exhaust Velocity. Divide the thrust by the total or average mass of spacecraft and you get initial and average acceleration. A 5 ton modern Solar collector dish would have an area of 714285m^2. This is a disk 953m in diameter. It collects 976MW of sunlight, and reflects 927MW into the engine. The thrust is therefore 123.4kN. A more advanced STR can have even better performance. The failure scenario you described can happen with every single spaceship, just replace 'sheets melt' with 'engine failure' or other examples.

Hi! I've got something new for y'all! Liquid Rhenium Solar Thermal Rocket The maximum temperature concentrated sunlight can heat a material to is 5800K. How do we approach this limit? We will describe existing and potential designs for solar thermal rockets. Solar thermal rockets The Solar Moth The principle of a solar thermal rocket is simple. You collect sunlight and focus it to heat a propellant headed for a nozzle. A rocket engine's performance is determined by its thrust, exhaust velocity and efficiency. A solar thermal rocket's thrust can be increased by sending more propellant through the nozzle. Its exhaust velocity can be increased by raising the propellant temperature. Doing either required more power, so more sunlight needs to be collected. Efficiency will depend on the design. The main advantages of a solar thermal rocket are its potential for high power density, high efficiency and high exhaust velocity. Collecting and heating with sunlight does not need massive equipment - unlike solar electric spacecraft that need solar panels, extremely lightweight reflective metal films can be used. A heat exchanger above a nozzle is compact and masses much less than the electrical equipment and electromagnetic or electrostatic accelerators a solar electric craft uses. Radiators are not needed either, as the propellant carries away the heat it absorbs with it. Put together, a solar thermal rocket can achieve power densities of 1MW/kg while solar electric craft struggle to rise above 1kW/kg. Sunlight would follow the same path as the laser beam here. As the sunlight is being absorbed by a propellant and expanded through a nozzle, there are only two energy conversion steps: sunlight to heat, then heat to kinetic energy. The first step can be assumed to be 99% efficient. The second step depends on nozzle design, but is generally better than 80%. Exhaust velocity will be determined by the root mean square velocity of the gas the propellant turns into. The equation is: Exhaust velocity: (3 * R * Temperature * 1000 / Molar mass ) ^ 0.5 Temperature is in Kelvins. Molar mass is the average g/mol value of the propellant at the temperature it is heated to. R is the molar gas constant, equal to 8.314 J/mol/K. For the very hot gasses we will be considering, we can assume complete dissociation of all molecules. H2 (2g/mol) will become atomic hydrogen (1g/mol), water (18g/mol) becomes a hydrogen-oxygen vapor (6g/mol) and so on. Low molar masses are preferred, with the best propellant being mono-atomic hydrogen unless other factors are considered. These advantages are all the critical elements that allow for travel throughout the inner solar system without requiring vast quantities of propellant. This means smaller spacecraft and lower travel times. Heat exchangers and exhaust velocity The limiting factor for solar thermal rockets is how hot they can heat the propellant. Directly heating the propellant is a difficult task. The lowest molar mass propellant, hydrogen, has terrible absorption. For all practical purposes, it is transparent to sunlight. Seeding the propellant with dust particles that absorb sunlight and heat the hydrogen indirectly through conduction has a major catch: the dust particles get dragged along by the hydrogen propellant flow and increase the average molar mass. A single millimeter-sized carbon dust particle in a cubic meter of hydrogen increases the molar mass from 1g/mol to Indirect heating involved using a heat exchanger as an intermediary between the sunlight collected and the propellant being heated. So far, designs have required the use of a solid mass of metal that is heated up by concentrated sunlight. The propellant is run over the metal, or through channels in the metal, to absorb the heat. Tungsten is often selected for this task, as it has a high resistance to heat, is strong even near its melting point and has a good thermal conductivity. Testing a Hafnium/Silicon Carbide coating. More modern designs make the most of the latest advances in materials technology to allow for higher operating temperatures. Carbon, notably, stays solid at temperatures as high as 4000K. Tantalum hafnium carbide and a new Hafnium-Nitrogen-Carbon compound melt at temperatures of 4200 and 4400K respectively. However, looking at our exhaust velocity equation, the limits of modern materials technology will only provide a 21% increase over common tungsten. This is the reason why so many propulsion technologies that rely on exchanging heat between a heat source, such as a nuclear fuel or a laser beam, and a propellant using a solid interface are said to be 'materials limited' to an exhaust velocity of 9.6km/s with tungsten, or 10km/s with carbon. THC or HNC would allow for an exhaust velocity of 10.5km/s. This is the deltaV equation, also known as the Tsiolkovsky rocket equation: DeltaV = ln (Wet mass / Dry mass) * Exhaust Velocity Wet mass is how much spaceship masses with a full load of propellant. Dry mass is the mass without any propellant. The wet to dry mass is also referred to as the 'mass ratio' of a rocket. We can rewrite the rocket equation to work out the required mass ratio to achieve a certain deltaV using a rocket engine's exhaust velocity: Mass ratio = e ^ (DeltaV required/Exhaust Velocity) 'e' is the exponent 2.7182... in simpler terms, the mass ratio increases exponentially as the deltaV required increases. Or, put another way, the mass ratio required decreases exponentially as the exhaust velocity rises. It is critical to have a higher exhaust velocity for rapid space travel without requiring massive rockets and towers of propellant. You might also have noticed that 'solid' is a keyword up to this point. Why must the heat exchanger remain solid? Liquid Rhenium There is a method to achieve the true maximal performance of a solar thermal rocket, which is heating up the propellant as far as it can go. This is incidentally the temperature of the surface of the sun (5800K). At this temperature, hydrogen propellant reaches an exhaust velocity of 12km/s. A rare, silver-black metal. Rhenium is a rare metal with a surprising number of qualities, one of which is a very high boiling point. Rhenium melts at 3459K but remains liquid up to 5903K. The trick to achieving higher exhaust velocities is to use a molten heat exchanger, specifically liquid rhenium at a temperature of 5800K. Rhenium is also very stable and does not react with hydrogen even at high temperatures, which is something carbon-based materials struggle to survive. It has already been considered as a heat exchanger, in solid form, by NASA. Here is a design that can use liquid rhenium as a heat exchanger: The diagram is for illustrative purposes only - a functional schematic would be more detailed. Here is an explanation for each component: Solar collector: A very large, very lightweight reflective film based on solar sails that can collect sunlight and focus it through a series of lens onto the heat exchanger fluid's inner surface. Rotating drum: The drum's inner surface contains a liquid heat exchanger. The outer surface is actively cooled. The drum is dotted with tiny channels that allow the propellant to enter the liquid from the bottom and bubble through to the top. It is made of Tantalum-Hafnium Carbide. Fluid surface: The fluid here is liquid rhenium. Its surface is heated to 5800K by concentrated sunlight. The lower layers nearer the drum holding the fluid is cooler. The centripetal forces hold the fluid in place Pressure chamber: The rotating gas mix gets separated here. Dense rhenium vapours fall back down, hot hydrogen escapes. Bubble-through heating: The rotation induces artificial gravity, allowing the hydrogen to heat up and rise through the denser rhenium. As it rises, it reaches hotter layers of the fluid heat exchanger. At the surface, it has reached 5800K. Small bubbles in direct contact with the rhenium allows for optimal thermal conductivity. More detail below. Active cooling loop: liquid hydrogen from the propellant tanks makes a first pass through the drum walls, lowering the temperature below the melting point of THC. It emerges as hot, high pressure gaseous hydrogen. High pressure loop: The heated hydrogen is forced through the channels in the drum. It emerges into the fluid heat exchanger as a series of tiny bubbles. Here is a close up of the drum wall, which contains both active cooling and high pressure channels: The configuration displayed above allows the hydrogen to enter the basin bottom at 4000K, then be heated further to 5800K before being ejected into the pressure chamber. If higher quantities of liquid hydrogen for active cooling are used, the drum and high pressure channel temperatures can be lowered to 3800, 3500, 3000K or lower. This pebble-bed nuclear thermal reactor has most of the components of our solar thermal rocket, except that instead using pebbles of nuclear, fuel, we use a liquid rhenium bed heated by sunlight. If the liquid hydrogen active cooling cannot handle the full heat load, radiators will be needed to cool down the drum below its melting point of 4215K. Thankfully, these radiators will receive coolant at 4000K. Their operating temperature will be incredibly high, allowing for tiny surface areas to reject tens of megawatts of waste heat. Electricity can also be generated by exploiting the temperature difference across the radiators' entrance and exit flows, and at very high efficiency. Operation The design is a Rotating Drum Fluid Heat Exchanger Solar Thermal Rocket (RD-FHE STR). It allows for hydrogen propellant to reach 5800K and achieve the maximum performance of a Solar Thermal Rocket. Liquid rhenium does not boil at 5800K, so it remain liquid and can be held inside the basin by simple centripetal forces. Vapor pressure of rhenium at 5800K (0.was determined to be low enough for our purposes. A surface of rhenium exposed to vacuum at that temperature would lose 0.076g/cm^2/s, or 762g/m^2/s. It is unknown how much centripetal force affects the loss rate of rhenium. The pressure chamber would operate at several dozens of atmospheres of pressure, which is known to increase the boiling point and reduce the evaporation rate of fluids. The same techniques used in Open-Cycle Gas Core nuclear reactors to prevent the loss of uranium gas can be applied to reducing the loss of rhenium vapours. At worst, the rhenium heat exchanger loses 0.76 kg of rhenium for square meter per second of operation. Looking at the designs below, the mass flow rate is measured in tons of hydrogen per second. This is a ratio of 1000:1, to be improved by various rhenium-retaining techniques. It should also be noted that rhenium is a very expensive material. A tungsten-rhenium mixhas very similar thermal properties and is much cheaper. Sunlight at 1AU provides 1367W/m^2. A broad-spectrum reflecting surface such as polished aluminium would capture and concentrate over 95% of this energy, so more than 1298W would be available per square meter. Solar sails materials such as 5um Mylar sheets are preferred, massing only 7g/m^2. More advanced materials technology, such as aluminium film resting on graphene foam, might mass as little as 0.1g/m^2. The 'Solar Moth' used inflatable support structure for its mirrors. Based on data for the Solar Moth concept, we have estimated that a solar thermal propulsion system can attain power densities of 1MW/kg. So, each square meter of collector area will require another 1.29 grams of equipment to convert sunlight into propulsive power. Performance Robot Asteroid Prospector We will calculate the performance of two versions of the RD-FHE STR. The first version uses modern materials and technologies, such as a 7g/m^2 Mylar sheet to collect sunlight and a 167kW/kg engine power density. The second version is more advanced, using 0.1g/m^2 sunlight collectors and a 1MW/kg power density. Modern RD-FHE 5 ton collection area => 714285m^2 927MW of sunlight focused onto the drum. 5.56 ton propulsion system Exhaust velocity: 12km/s Thrust: 123.4kN (80% efficiency) Thrust-to-weight ratio: 1.19 Overall power density: 87kW/kg Advanced RD-FHE 5 ton collection area =>50000000m^2 64.9GW of sunlight received 64.9 ton propulsion system Exhaust velocity: 12km/s Thrust: 10.8MN Thrust-to-weight ratio: 15.75 Overall power density: 928kW/kg The principal argument against solar thermal rockets, that their TWR is too low and their acceleration would take too long to justify the increase in Isp, can be beaten by using very high temperatures and very low mass sunlight collectors. For example, a 50 ton propulsion system based on the modern RD-FHE STR design, would be able to push 100 ton payloads to Mars (6km/s mission deltaV) using only 97 tons of propellant. It would leave Earth orbit at a decent 0.24g of acceleration, averaging 0.32g. The departure burn would take only 20 minutes. Using the advanced version of the RD-FHE solar thermal rocket would allow for a positively impressive acceleration of 3.1g. With 12km/s exhaust velocity, multiple missions that chemical rockets struggled to do with low-energy Hohmann transfers can be avoided. A chemical rocket such as SpaceX's BFR might achieve an Isp of 375s, which corresponds to an exhaust velocity of 3.67km/s. It would need a mass ratio of 5.13 to barely produce enough deltaV for a Mars mission. Earth to Destination. If our solar thermal rocket is granted the same mass ratio, it would have a deltaV of 19.6km/s. This allows for a Mars mission to be completed in under two months (10km/s departure, 9km/s insertion). It is also enough deltaV to reach Jupiter with a single stage. Other benefits include a vast reduction in the propellant-producing infrastructure needed to supply orbital refuelling depots and the ability to land on Mercury. Alternative versions: Blown hydrogen: Instead of bubbling hydrogen from the bottom of the liquid rhenium basin, hydrogen is blown into the pressure chamber from the top. It is heated by simply passing over the fluid heat exchanger. The advantage is that the rotating drum does not have to be riddled by microchannels, allowing it to be stronger and rotate faster, which would reduce rhenium losses, and also accept a higher rate of active cooling by leaving more room for liquid hydrogen channels. Another advantage is that there is less chance of hydrogen bubbles merging and exploding in showers at the surface, dragging along rhenium as they escape. The disadvantages is vastly reduced heat conduction rate between the rhenium and the hydrogen. This would require a long and thin pressure chamber to increase the time the hydrogen stays in contact with the rhenium, potentially making the propulsion system heavier than it needs to be and forcing sunlight to enter the chamber at very acute angles. ISRU propellants: Instead of hydrogen, other gaseous propellants might be used. Nitrogen is a good choice, as it is inert and only reduces the exhaust velocity by a factor 3.7 compared to hydrogen. Powering a hydrogen extraction process on Mars requires huge areas of solar panels. Nitrogen is easily sourced from Earth's atmosphere by gas scoops. Other options, such as water or carbon dioxide, are also viable and available on other planets. The advantage is that non-hydrogen propellants are easy to contain and are much denser than hydrogen, so their propellant tanks can be lightweight and small. They are easily sourced and only need to be scooped up and filtered, unlike hydrogen that has to undergo electrolysis. The disadvantage is that there propellants cannot serve as expandable coolant for the rotating drum. A radiator using a closed gas loop is necessary - helium is a likely candidate. This adds mass. A lower exhaust velocity also removes the principal advantage the RD-FHE STR has over other propulsion systems.

A good thing! It just goes to show how much SpaceX has shaken up the market by moving it back into the competitive industry it should have always been. It means that there are multiple players and the US Government is not the sole source of money - in other words, spending your R&D budget on lobbyists is no longer a winning move. Maybe the BFR won't be allowed to carry civilians for a long time or even land near cities, but rapid transport of 100 tons of cargo is very valuable. Imagine shortening the supply chain on valuable items to 30 minutes! Troops will accept risks civilians won't, so good idea. Reading that. Will comment.

Yes! We're back! This time face-to-face with a giant asteroid, with none of that flags-and-footprints nonsense. Straight up nuclear reactors. Although... vacuum toilets?! That's just one accident away from, eh, a de-pooppressurization event.

A BFR with twice the payload cannot make orbit. Even if it gives up trying to land, it is only adding 11 tons to the cargo capacity. Designing a small rocket is vastly less expensive than designing two new and massive stages!! Just look at the difference in development cost for a private jet versus a big airliner. Its tens of millions versus billions. So, an SFR will be cheaper. It won't replace the BFR until the market is ready for the BFR. It's far from a 'giant' and building it from the ground up is simpler than trying to redesign the F9 boosters to accept cross-feed and side-mounting, which is the main reason behind the Falcon Heavy delays. Abandoning the Falcon 1 was a smart decision. It was aimed at the small satellite market but... that market is utterly tiny. SpaceX at the time, and mostly still is, strapped for cash and has to prioritize where it spends its money.

Actually, the larger you make your SSTO the closer you get to the mass ratio of your propellant tanks. Hoop stress is what determines the thickness of the propellant tank walls. Hoop stresses rise linearly with radius, so for the same pressure, you need the same strength, and so the same thickness. However, the smaller tank suffers less stresses due to its own mass; after all, they are not held up by pressure like balloon tanks. Either way, I think the point can be ignored. The current Falcon 9 booster manages an overall mass ratio of 19.5, engines and landing legs and everything else included. That mass ratio is dominated by the propellant tank mass, so the aluminum-lithium propellant tanks in use have at worst a mass ratio of 20. Carbon fibre tanks are expected to have about 3.2x better mass ratios, so 1:62 or better. The BFS's mass ratio is 14... the propellant tanks very roughly influence 22% of the dry mass. That's 3.2 tons out of the 14.4 tons of the SFS. How much worse does scaling down make the carbon fibre tank mass ratio? A full cube/square law ruling would make the volume 14.39 times lower for a surface area 5.9 times lower, so the mass ratio could be 2.4 times worse.... this will increase the mass of the propellant tanks to 7.7 tons and reduce the revised cargo capacity of the SFR from 12-14 tons down to 7.5 - 9.5 tons. At worst.

For spaceships like these, the vast majority of the volume is just a big propellant tank. You can replace the spaceship with a hollow cylinder of the same volume and still be accurate within 10% of the mass by scaling up and down using area instead of volume. Elon Musk's comment seems right... until you do the calculations and realize that it would only be possible to make an SSTO out of the BFS if you also consume the propellant held in reserve for landing. Also, I have revised the numbers with my own Excel table calculation instead of relying on an app. Instead of putting 9 tons in orbit, the SFS can put 12 tons.

Thank you. I try to be extra-conservative with these estimates for systems that have not been developed yet, but if 75/50: 50% less dry mass can be expected for cargo containers, then we can expect significant increases in payload capacity to orbit. Looking at the presentation again, it seems that the crewed BFR is expected to be 75 tons, 85 with development bloat, and that the tanker version can be made lighter because the crew spaces can be removed entirely, leaving a spaceship that is mostly propellant tank. With SpaceX's development and testing pace, TRL4 can be turned into TRL9 in less than five years. The development of the SFS to fix on top of the Falcon 9 booster is a tiny fraction of the cost of developing the entire BFR, both upper stage and booster, plus building new launch platforms and other infrastructure for a 9m diameter rocket. If SpaceX goes ahead with the BFR, it will very likely end up with a rocket they just can't launch more than once or twice a year without leaving the cargo hold entirely empty. That is significant commercial risk. A lot of Falcon Heavy's problems came from the requirement of cross-feeding the boosters, something that has not been done for rockets not designed from the ground-up to handle cross-feed. The BFS would have four vacuum-rated Raptors and two (now three) landing Raptors. The SFS will have one vacuum-rated Raptor and two Merlins/four SuperDracos for landing. If the Raptor has anything like the reliability of existing Merlins, then having a single one should not be a problem... second stages on the Flacon 9 have never failed to ignite and re-start. The SFS would have to keep a different set of propellants in its header tanks for landing if it uses the SuperDracos. It would be a stop-gap measure until a small ~200kN engine running on liquid methane and liquid oxygen is developed... such as converting a Merlin to lower pressures and a different propellant mix. Considering that the Methalox mix is held inside the main tanks for roughly 50 minutes or so, I do not think the insulation requirements will be extreme. Electric heating can be employed. Heat will leak out of the header tanks and into the main tanks if that solution is used... but it will help pressurize the main tanks as they are rapidly drained during launch. The decision to make is between the insulation, heated tanks or methalox Merlins, whichever is best. Just slapping on insulation seems like the cheapest solution to develop.

The BFR in orbit would need 1100/150: 7.3 launches to be filled up too, so we're only talking two to three times the launch cadence but with a much smaller vehicle. The problem as I mentioned with the BFR is that it is unclear whether it is possible to launch the BFRs mostly empty and still make a profit. I find this possibility quite unlikely because if it were the case, then a smaller rocket that is better filled up would make a killing. The mass ratio and engine Isp of the BFR and SFR are the same, so everywhere the BFR can go, the SFR can do the same. Pumping up the launch cadence by refuelling BFRs in orbit still has the problem that there's no money anywhere that will pay for the 8 launches to put a fully fuelled BFR with 150 tons of cargo in orbit. The only envisageable mission would be a large-scale lunar or martian exploration project... but that's mostly up to the governments whether they have the budget. And of course, there's the SLS. I agree with your final point though. If Elon Musk pushes for something that doesn't look or sound economically viable, backed up with his personal cash reserve or over-excited investors, then it will get done. We're can't rely on irrational forces though. Consider the SFR as 'sensible' and the BFR as 'visionary'.

The BFR is already in the works and critical components, like ultra-high-mass-ratio propellant tanks and ultra-high-TWR rockets have already been built and tested. I kind of trust that Elon Musk didn't just throw away figures with the usual disregard given for 'paper' projects. Scaling down using the same technologies should achieve similar results.

I'm a bit of a pessimist in this regard. I think the 'answer' is time. Eventually we'll start start suffocating for new resources and openings for our increasingly educated population. Eventually technology will advance so that high energy propulsion becomes available, even if it was not the primary direction of research. Eventually... we'll reach a position where moving off-world is more a question of 'why not' and we'll think of it as no different than being asked to take a job in another country.

Well, SpaceX shrunk from the ITS to the BFR by a factor 1.8 and maintained the same mass ratio.

I must remind you that my calculations ignored the fact that seven more Raptor engines will add up to 6.7 tons to the dry mass and reduce the payload capacity from razor thin to negative.

I ran the numbers, the BFR cannot function well as a recoverable SSTO. Its deltaV with no cargo is ln(1185/85)*375*9.81 = 9692m/s. However, it has to launch on lower Isp rocket engines, which will cut this deltaV capacity down below the 9400m/s just to reach orbit... and then it it still needs to reserve some propellant to land with. A 400m/s landing deltaV requires 11 tons of propellant using 330s Isp engines. So, we must set the dry mass at 85+11: 96 tons. The BFS is no longer an SSTO is we use the 96 ton figure as dry mass. The picture changes slightly if we use the strict 75 ton mass with no development bloat. The landing reserve drops to 9.86tons and the mass in orbit can be as low as 84.8 tons. This is enough to make the BFS an SSTO again... but with zero payload capacity. This is also ignoring the 7*969 kg engines it needs to ADD to its existing engines to even liftoff, which will add up to 6.7 tons into its dry mass... The increase in Isp from using Raptor engines instead of the current Merlin engines makes the mass budgets more lenient on an upper stage. However, Raptor engines are too valuable to throw away as expendable upper stage engines, so they are well suited to the fully reusable SFS stage.

The full post for you to enjoy and discuss, from here: http://toughsf.blogspot.com/2017/10/spacex-sfr-small-falcon-rocket.html Performance estimates revised: 24 tons expendable, 12-14 tons recoverable to LEO. The Small Falcon Rocket is a scaled down alternative to SpaceX's Big Falcon Spaceship that fits on top of existing Falcon 9 boosters. We will discuss the advantages and disadvantages of such a design. SpaceX's Big Rockets The BFR, or Big Falcon Rocket, is comprised of the Big Falcon Spaceship and the Big Falcon Rocket booster. It is a scaled down and simplified design based on the ITS, or Interplanetary Transport System. The BFR is a BIG rocket. The ITS was revealed in June 2016, although work on the design has begun in 2013 under the name 'Mars Colonial Transporter'. The ITS promised to deliver 300 tons of cargo to Low Earth Orbit, or up to 550 tons if reusability was ignored. It would have massed 10500 tons on the launchpad. The vehicle had a diameter of 12 meters and a height of 122 meters, making it one of the largest rockets ever plausibly considered. And the ITS was positively massive. The upper stage, called the Interplanetary Spaceship, was supposed to hold 1950 tons of propellant with a dry mass of 150 tons. Without a payload, the mass ratio was 14. The BFR replaced the ITS in September 2017. It is a smaller, more sensible design that SpaceX believes it can actually deliver in the next few years. The diameter is reduced to 9 meters and it will mass 4400 tons on the launchpad. Payload capacity is reduced to 150 tons. The upper stage BFS should have a dry mass of 75 tons, but Elon Musk states that this might rise to 85 tons due to development bloat and overruns. It holds 1100 tons of propellant, giving it a mass ratio of 13.9. It is important to note that despite being up to 78% smaller than the previous ITS design, the BFS stage maintains the same mass ratio. Why? Because we are now going to scale down the BFS again. Why go smaller? How big the BFR's booster would be compared to the Falcon 9 booster. Going big is the best way to reduce the cost per kilogram for sending payloads into orbit. SpaceX jumped from the Falcon 1 to the Falcon 9 because the larger rocket can deliver payloads much more cheaply into space. When first considering options on how to make travel to Mars affordable to the general population, SpaceX immediately came up with a gargantuan tower of rocket fuel over three and a half times larger than the Saturn V! A big rocket is also easier to develop. It is more forgiving of development bloat that increases mass over time as the designs are perfected. It has larger safety margins and room for many backups, such as multiple engines. However, bigger is not always better. The total development costs will be higher, as large components need large factories. It is much more difficult to test the components too, and a full testing regime of the completed rocket will require launching and even destroying a full-scale model many times. Remember the failed Falcon 9 booster landing attempts, and imagine them replaced with a vehicle eight times bigger. There is also the fact that the second sure-fire way to reducing launch costs is to have rapid turnover. This involves loading up rockets, sending payloads into space, recovering the rocket and refurbishing it for another launch in a very small time frame, measured in days or even hours. Rapid turnover and minimal refurbishment would allow the space launch industry to more closely resemble existing airline business models. The main benefit of this approach is that a small number of launch vehicles can handle a large volume of missions, critically reducing the initial cost of the vehicles and reducing the amortization rate. Even if SpaceX manages to develop rockets that liftoff and land several times without needing to go to a workshop, they'd still need to solve the issue that there just aren't enough payloads on the market that need to be lifted into space to fill the BFR, let alone the ITS. For example, even the BFR's 150 ton payload capacity can cover all of last year's payloads in about two or three launches. Three launches is far from sufficient. Elon Musk is betting that the space industry will be able to fill the BFR's cargo bays with new satellites and LEO payloads once the lowered cost per kg is offered to them... but there will be a long delay between the launch costs being reduced and the industry contracts appearing en masse. Cost per kg in orbit is only part of the picture. Waiting for more contracts to appear and bundling them together to use the most of a BFR's cargo capacity is not a good solution. It will force SpaceX to delay launches until the mass delivered to orbit reaches a profitable amount - launching BFRs nearly empty with the usual 2 to 5 ton satellite is surely wasteful and a loss for the company. The SFR The SFR, or Small Falcon Rocket, is a possible solution to the development costs, under-utilization and low expected launch rate of the BFR, or Big Falcon Rocket. The SFR is a scaled down Big Falcon Spaceship sitting on top of an existing Falcon 9 booster. It will carry a smaller payload to orbit, but will have a capacity SpaceX is sure to fill up. Existing Falcon 9 boosters can be mated to a fully reusable upper stage, drastically cutting down on development costs. We will now look at the details of the SFR's two stages. The upper stage is the only new part. It is a BFS scaled down to 3.7 meters diameter, using the same Raptor engines rated at 1900kN of thrust at 375 seconds of Isp. We will call it the SFS, or Small Falcon Spaceship. The Raptor engine. The SFS will be (9/3.7)^2: 5.9 times smaller than the BFS. The dry mass is expected to be only 85/5.9: 14.4 tons. It will be 19.7 meters long. Based on the mass ratios calculated above, the SFS will be able to hold 187.2 tons of propellant. An SFS with no cargo and full propellant tanks will therefore mass 201.6 tons and have a deltaV of ln(14)*375*9.81: 9708m/s. The Vacuum-optimized Raptor engine is quite large, with a nozzle opening 2.4 meters wide. It is unlikely that more than one such engine can be fitted under the SFS. It will provide enough thrust for an initial Thrust-to-Weight ratio of 0.96, which must be compared to the current second-stage initial TWRs of 0.8-0.9. For retro-propulsive landing, we will not be able to fit, or even need, the sea-level version of the Raptors. Instead, we will use two of the existing Merlin-1D engines with 420kN of sea-level thrust, but possibly with a lower pressure rating as the thrust generated makes them too powerful for landing. The alternative is the SuperDraco engines with 67kN of thrust and 235s sea-level Isp. Rocket engines in the Raptor + 2x Merlin configuration would represent 13.2% of the overall dry mass, or 8.1% if the Raptor + 4x SuperDraco configuration is used instead. The Raptor engines are assumed to have a TWR of over 200, so their mass should be lower than 969kg. There are no numbers on the SuperDraco's mass, but it should be at most 50kg. These ratios seem not too outrageous when compared to the 7% engine-mass-to-dry-mass ratio in the BFR's original design. Merlin-1D engines. The SFS's mass is based on the 85 ton figure for the BFR's dry mass, but this is a cautious estimate with room given for development bloat and mass budget overruns. The BFR's design on paper gives a dry mass of 75 tons instead. Using the on-paper mass, the SFS could have a dry mass as little as 12.7 tons. The SFR's booster is the Falcon 9 Block 4. The booster will mass 22.2 tons when empty, and can hold 410.9 tons of propellant. This gives it a mass ratio of 19.5. The nine Merlin 1D engines have a sea-level Isp of 282s and an vacuum Isp of 311s. Because the booster stage does not spend a long time at sea level and performs most of the burn at high altitudes with negligible air pressure, we will use 300s as a low-ball estimate of the average Isp. The true average might be a few seconds higher. Taken all together, the SFR will mass 634.7 tons on the launchpad without any payload in the SFS's cargo bays. It stands 89.7 meters tall. We will now calculate how much cargo it can lift into Low Earth Orbit in expendable or reusable mode, and where else it can go. Performance To achieve a Low Earth Orbit, we will set the deltaV requirement as 9400m/s. In reality, it could be achieved with as little as 9200m/s, but we want decent safety margins. Expendable mode is the easy part. It assumes every bit of propellant is consumed and the SFR's stages left dry. Using a multi-stage deltaV calculator and setting the Falcon 9 Block 4's Isp to 300s and the SFS's Isp to 375s, we work out that the booster provides 1899m/s of deltaV and the SFS provides 7488m/s for a total of 9388m/s with a payload of 13.7 tons. Recoverable mode is harder to calculate. The propellants cannot be completely used up: some must be kept in reserve to perform a retro-propulsive landing burn. BFR landing. A landing burn by the SFS requires that about 300m/s of deltaV be held in reserve. This represents 1.65 tons of propellant with Merlin-1Ds or 2 tons of propellant with the SuperDracos. The Falcon 9 booster needs to retain 15% of its propellant reserve to make an ocean landing. This gives it a deltaV of 3910m/s, which is largely enough to cancel most of its forwards velocity and make a very soft landing. However, holding back 61.6 tons of propellant means it boosts the SFS by much less. In recoverable mode, the SFR's cargo capacity drops to 9 tons. If the SFS follows the paper designs more closely and achieves a dry mass of 12.7 tons, it will have cargo capacities of 16.7 tons in expendable mode and 12 tons in recoverable mode. The SFS could achieve a deltaV of 2500m/s after launching on top of a recoverable Falcon 9 booster and without any payload. This is not enough to reach the Moon, so the range of missions the SFR can take payloads on is limited to Low Earth Orbit. Smaller rockets might solve the problem of having to crane down cargo from the top of a tower. However, if it is refuelled in orbit, then the entire Solar System is available. It can deliver 50 tons to Low Lunar Orbit (5km/s mission deltaV). It can send 35 tons to the Mars Low Orbit (5.7km/s mission deltaV) or 21 tons to Mars's surface (6.7km/s mission deltaV). Refueling the SFS will take between 16 and 20 tanker launches. With 14.4 tons of dry mass and a propellant capacity of 187.2 tons, the SFS has a maximal deltaV of 9.7km/s, enough theoretically to put itself far above Jupiter or even Saturn. Conclusions The SFS is a limited vehicle. It is restricted to Low Earth Orbits and can deliver payloads of 9 tons, up to 12 tons, at most. It is far from the multi-purpose machines the BFR or ITS promised to be. However, it is enough to dominate the medium lift launch market, as it is fully recoverable. The re-use of existing Falcon 9 boosters and the smaller number of Raptor engines (one per rocket) will drastically slash the development costs compared to something like the BFR. The smaller payloads are easy to fill, meaning every launch is profitable. Multiple launches promises rapid turnover and a maximization of the return on investment on the craft. With re-fueling, the SFS in orbit can complete missions that require it to send decent payloads to the Moon and Mars. With minor improvements and operating in fleets of multiple vehicles, it can even match the payload capacity of the BFR to various destinations. What do you think?