Armchair Rocket Scientist

Fortunately, even going to polar orbits, countries don't let a rocket launch overfly populated areas. Here's a ground track of a polar orbit launch from Guiana Space Center: NASA never launches to polar orbits from KSC, because in either direction you'd be overflying populated areas. Instead, they launch from Vandenburg, like this: Russia can't really avoid launching over land, but their trajectories seem to mostly go over Kazakhstan, Mongolia, and the Asian part of Russia, all of which are sparsely populated.

Well, boron is fairly rare, and tungsten is extremely rare, so I don't think either of them are good candidates. Nitrogen seems unlikely without carbon being involved as well, since complex nitrogen-containing molecules with no carbon seem to mostly be high explosives.

Well, ships still require maintenance, fuel, etc. That being said, a landing platform wouldn't have a lot of the fancy on-deck equipment an operating oil platform has. The most important equipment would be some type of crane to secure the rocket after landing and possibly load it onto a transport ship; there are already semisubmersible Crane Vessels, but they're designed to lift thousands of tons, and probably that useful something delicate like a rocket. There would also probably need to be systems to handle leftover fuel; LOX could just be vented, as could nitrogen or helium from cold gas systems, and methane would be best off burned (it isn't particularly toxic, but it's a much stronger greenhouse gas than its combustion products). Kerosene and other liquid hydrocarbons would need to be captured and stored, as would any hypergolic propellants. Pretty much. I believe SpaceX has said a downrange propulsive landing (which they're only doing for tests) means a 15% payload hit, vs. 30% for boostback. Essentially, you're sacrificing a fixed investment equivalent to building several new 1st stage cores, and maybe 1 core worth of maintenance per year, for a 20% payload increase. With the Falcon Heavy core though, boostback would be pretty much impossible, and there's not much the right distance downrange from Texas. For that matter, polar orbits don't offer many convenient islands. Launching north from Puerto Rico it might be possible to reach the US East Coast with some fancy manuevering, but it would take a lot of dV. The Gulf of Mexico is that deep, at least in the parts where it's convenient to land a rocket. Most of the oil wells are in relatively shallow water on the continental shelf and slope, but from SpaceX's Texas site, the best landing sites would be 2-3 km deep. Besides which, IIRC Elon Musk has said that the FH core would overshoot Florida from Texas, which means the landing area would be over 5 km deep. The same would be true for any launches from Florida, and any polar launches from California or Puerto Rico.

Half the dV on the first stage is about what the Falcon 9 first stage gets, so reentry should be fine with a fairly small braking burn. However, even with a metholox stage with a mass ratio of almost 25, the boosters would have to be 80% of the rocket's liftoff mass (by comparison, if the boosters are the same size as the core, they'd be less than 67% of the liftoff mass). Also, since the core stage would be burning at the same time as the boosters, they'd actually be going a bit faster at burnout. First off, you'd have to run literally thousands of little pipes around the entire outside of the stage. This plumbing is going to be heavy and very expensive to manufacture. Running oxygen or methane through the pipes is a very bad idea: even at room temperature you can't liquify them with any amount of pressure. That means you'd have to boil off the fuel to cool the vehicle, essentially using it like an ablative heat shield. Since oxygen and methane can't absorb that much heat by boiling, this fuel would be put to much better use with a braking burn, with the added advantage of not requiring the heavy plumbing. For that matter, an actual ablative heat shield - most likely using some kind of spray-on coating would be more effective. Active cooling systems have been proposed for spacecraft, but generally for things like SSTO spaceplanes on the way up (which would have significant aerodynamic heating with the engines running, meaning you could immediately burn the boiled fuel), or orbital reentry (where the speeds are so high that braking burns aren't practical). Propulsive splashdown partially solves the structural issues, although the stage could still topple over after its engines were underwater and smack its upper section into the surface (I've lost some stages like this in KSP), and liquid engines still don't like saltwater. One idea I really like for lower-stage reusability is the BDB (Big Dumb Boat) approach. Basically, you land the stage on a floating platform downrange of the launch site. Semi-submersible vessels are very popular for deep water oil drilling because they're very stable even with wave action, which also happens to be a desirable trait when you're trying to land a tall rocket stage on one. Used oil platforms are about a hundred million dollars, and new ones a few times more, so $200M for a landing platform is reasonable. The platform would be towed into position approximately where the first stage would land by tugboats. At launch time, the tugboats and all personnel would station themselves several miles away, and the stage would guide itself to a propulsive landing. Once the stage is safely landed, the crew would return and perform any necessary post-flight checks. At this point, the stage could either be partially refueled and fitted with an aerodynamic nose cap, then fly itself back to the launch site (giving turnaround times of a few hours) or be loaded onto a boat and sailed back to the launch site (still allowing turnaround of less than 24 hours). Now, SpaceX wants "single digit hours" turnaround with their lower stages. However, this wouldn't allow much more than a preliminary "are there gaping holes in the airframe or missing engines) inspection before slapping on a second stage and payload, refueling, and launching again. Besides, a launch provider doing 200 flights a year (a lot) with a fleet of 5 reusable vehicles would only need to launch an individual vehicle every 9 days. Also, on a fully reusable vehicle the second stage wouldn't even be back yet. In theory, a second landing platform could be used, with one for low-inclination orbits and one for high-inclination orbits, meaning less time and fuel spent towing multi-kiloton platforms around. Now, this approach would be great for the core stage of a rocket with big boosters (such as the Falcon Heavy core), which would be going too fast to return to the launch site. However, the boosters would probably be going slowly enough to just boost back to the launch site.

Lemme quote XKCD: Now, the pressure differential between sea level and Mexico City is only about 0.25 bars, as opposed to 90 bars. The equation used breaks down for mach numbers above 0.3, but suffice to say that after the ice melted the airflow would probably be well above the speed of sound.

Reusability isn't associated with strapon boosters: every operational rocket I know of besides the Space Shuttle just lets spent SRBs crash into the ocean (this includes the SLS). So, basically, you discovered staging? The SLS/shuttle already does adopt "huge boosters smaller core," as do most rockets that use SRBs. For example, the Shuttle SRBs have more than twice the thrust of the entire main engine cluster... each. The Ariane 5's SRBs each have nearly 5 times the thrust of the main engine. The boosters may look small, but on the Ariane 5 each booster is heavier than the core, second stage, fairing, and payload. The only reason the boosters look small is that the cores of all the rockets mentioned use hydrogen fuel, which is much less dense than the solid propellants used. There are several major problems with this approach. 1: The rocket you proposed would have more than twice the cross-sectional area of the SLS, which is not good for aerodynamics. It's the same reason we don't see "pancake" rockets in real life. It would be more efficient to put all 8 of the booster engines on a giant central core, then replace the central pair with some sort of vacuum-optimized upper stage... which is basically what SLS does. 2: Hydrogen-fueled engines have poor thrust-to-weight ratios, and both the engines and the stages in general have poor thrust-to-cost ratios. The latter is especially true of the SSME. Only one rocket (the Delta IV) launches off of hydrogen alone; most rockets with hydrogen-fueled core stages have powerful SRBs so that lighter and cheaper main engines can be used. 3: Reusability in real life is a lot more complicated than just slapping a parachute on. First of all, the space shuttle SRBs are built like tanks compared to any liquid-fueled booster. At the descent speeds achievable with a parachutes, an LFB's fuel tanks would crumple like a soda can, to say nothing of delicate parts like engine nozzles. For that matter, LFBs would be more susceptible to being torn apart by aerodynamic forces in an uncontrolled descent. Second, the SRBs don't have too many moving parts, so dunking them in seawater isn't that bad, but liquid-fueled engines would require a LOT of refurbishment after a splashdown. Third, the more dV the boosters contribute, the faster they'll be going at separation. Too fast, and they'll burn up or be torn apart on reentry. So, let's say we want to redesign the SLS with giant reusable liquid-fueled boosters. First of all, we don't want hydrogen providing most of the takeoff thrust, so we'll switch to kerosene. While we're at it, let's make the core stage kerosene too so we don't have to deal with different refueling equipment. The boosters can't handle a splashdown, so they'll have to perform a powered landing either back at the launch site or downrange on a floating platform. This means many engines on the booster so that it can provide very low thrust when it needs to. They also can't contribute too much dV or we'd waste a lot of fuel slowing down enough to survive reentry. About the size of the core stage is a good maximum. But if our boosters are about the size of the core stage, we might as well make them identical and use a lot of the same hardware. Since they're using the same fuel, we can boost the payload capacity by adding fuel crossfeed. Ya know, this rocket is starting to look kinda familiar...

This. For example, approximating the trajectory as a 400,000 km long straight line trip, with an acceleration of 1 G, the trip would take about 3.5 hours... but require 125 km/s of dV.

Yep. An ocean is a large body of liquid water. The groundwater you find in Earth's crust is sometimes referred to as forming underground rivers or oceans, which makes some people think of giant caverns full of water. However, except in actual cave systems, which are relatively rare, the water actually exists in microscopic cracks or pores in the rock, or in buried sediment which hasn't experienced enough heat and pressure to turn it to rock. In fact, some of the best aquifers are basically wet sand or gravel. The water in Earth's mantle is not like that. It's actually locked up in the crystal structure of the mineral ringwoodite as hydroxide ions, and is well under 1% of its mass. For comparison, gypsum is about 20% water by weight, and it's not "wet" in the conventional sense. Aside from the lack of liquid water, that part of the mantle is several times deeper than the zone where the temperature and pressure are right to form diamonds. No organic molecule could possibly survive under those conditions.

This isn't accurate. Let's check the math. With an ISP of 400 (about midway between kerosene and hydrogen) and assuming a stage's mass is 10% the mass of its fuel, I'll consider various configurations for a rocket with a total of 10,000 m/s of dV and estimate the payload fraction. 0/10000 m/s (SSTO case): Payload fraction -1.4% (not possible). 1000/9000 m/s: Stage 1 payload fraction 75%, stage 2 payload fraction 1.1%. Total payload fraction: 0.8%. 2000/8000 m/s: Stage 1 payload fraction 56%, stage 2 payload fraction 4.3%. Total payload fraction: 2.5%. 3000/7000 m/s: Stage 1 payload fraction 41%, stage 2 payload fraction 8.5%. Total payload fraction: 3.5%. 4000/6000 m/s: Stage 1 payload fraction 30%, stage 2 payload fraction 14%. Total payload fraction: 4.1%. 5000/5000 m/s: Stage 1 payload fraction 21%, stage 2 payload fraction 21%. Total payload fraction 4.3%. As you can see, payload fraction is maximized when the stages contribute equal dV. However, as you mentioned, this is affected by upper stages having higher ISP, which tips the balance in favor of upper stages. This is outright wrong. The lowest dV for a liquid-fueled stage that I've heard of is 2750 m/s on the Proton-M, and it has 3-4 stages. The Saturn V's lower stage gets about 3800 m/s (using the vacuum ISP stats) and as I said, Delta IV and Atlas V both get over 6 km/s from their lower stages. The Ariane 5 probably gets more like 8.5, but its huge SRBs do a lot of the work compared to the GEM-60s on a Delta or Atlas. The reason for this, incidentally, is that many rockets are optimized for launches to GTO, which requires an extra 2.5 km/s. In reality, an Atlas or Delta's dV is split about evenly, with each stage providing 6000 m/s. On the other hand, the Falcon 9's stages each provide about half the dV for an LEO launch.

A major factor is how fast the first stage is moving when the second stage ignites. For example, according to my calculations based on this, the Proton-M's first stage only provides 2750 m/s of dV, while the Atlas V and Delta IV CCBs can provide 6000 m/s of dV (in the lightest configuration). According to , the Ariane 5 is moving 6.91 km/s at first-stage cutoff.Now, a higher dV first stage is a higher fraction of the vehicle's launch mass, so the ratio between vehicle mass at first stage ignition and vehicle mass at second stage ignition is higher. For example, the Proton second stage + everything above it is 37% of the vehicle's launch mass. For the Delta IV or Atlas V, it's more like 10%. If the second stage needs to accelerate less mass at a certain minimum TWR, it needs less thrust. In addition, because the second stage is contributes less of the dV required to reach a stable orbit, it can safely use a lower TWR and still have time to reach orbital velocity without falling back into the atmosphere. The second stage of the Proton-M has a TWR of about 0.9, while the DCSS and Centaur get more like 0.3 depending on the payload. Falcon 9's second stage probably has a bit more thrust than it needs to because (a) SpaceX wants to use the same engine for both stages to reduce cost ( F9 will probably fly a steeper flight profile than most rockets to reduce return fuel requirements for the first stage.

The problem with that is some of the major physiological effects of long-term spaceflight are muscle atrophy (including the heart) and loss of bone density. These occur precisely because there isn't weight on the muscular, cardiovascular, and skeletal systems. The real danger isn't microgravity itself, it's going back to normal gravity. A human body in good condition is typically able to run and jump around, so all those systems are quite a bit stronger than is necessary to perform basic functions like pumping blood to the brain while standing up. A body in poor condition has a narrower safety margin.

I don't know a whole lot about the process of getting permission, but I'd say the US is fairly "rocket-friendly" compared to other countries. However, I don't know how much of that is the FAA and equivalent agencies and how much is fire protection, explosives, etc. It's hard to say why this is, but I'd guess that two possible reasons are (a) America is pretty economically libertarian compared to other developed nations (i.e. not receptive to government regulations on what you can purchase) and has something of a love affair with firearms. It's possible that this has also extended to other hobbies involving flames and "explosions." ( America isn't as empty as Canada or Australia, but compared to most European countries it's sparsely populated. In particular, the Great Plains region and the western/southwestern deserts have a bunch of wheat/corn fields, ranches, and dry lakebeds in the middle of nowhere, which means that American high power launches can get waivers to really high altitudes.

The systems described in the article both make heavy use of multiple planets in the same orbit. The exact situation you mentioned, with two planets on opposite sides of a star, would be a planet at L3. However, this situation is unstable; if the two planets come even one nanometer out of alignment, they will be pulled towards each other, ultimately resulting in either a horseshoe orbit or a very elongated tadpole orbit.

For a suborbital mission, the flight would last less than half an hour. In the US, I think they basically give you clearance for airspace over a given area up to a given altitude during a given time interval, and for larger projects you are expected to show that the vehicle will not leave the cleared area. For a launch from international waters, I'd guess that whatever country you're from will still make sure you notify air traffic control or otherwise make sure that no airplanes will fly over any area that might have a rocket launching from it.

Really, as I mentioned, I don't think you'd actually get anywhere near 60 habitable planets in one system. There's definitely room for three distinct "orbital lanes" in the HZ, but I'm not sure about 6 unless you consider a habitable zone out to 3 AU. That's for Earth-sized planets. With the right orbital resonance, you might be able to fit 2 gas giants. However, I'd expect such system's to mostly be like Saturn's: most of the mass in one, or maybe two habitable moons. In addition, a single gas giant could definitely have trojans. I'm not sure if they'd be stable for a pair. One option not mentioned is the possibility of habitable planets in orbital resonances with gas giants (other than 1:1, aka a trojan or horseshoe orbit). For example, the 2:3 and 3:2 resonances (as seen with Pluto and the Hilda family of asteroids respectively) seem to be stable. They might both result in eccentric orbits, but not enough to make a planet uninhabitable. However planets in such resonances could destabilize the orbits of trojans, and vice versa. In a best-case scenario, including: 4-5 terrestrial planets, one or two in horseshoe orbits or double planet relationships with others Two gas giants with two habitable moons each and maybe a resonant planet One gas giant with two habitable moons, two trojans, and a resonant planet I'd say 4-6 is the maximum number of EARTHLIKE habitable planets a single star could support. If you also includes the possiblity of ammonia-based or methane-based life, the habitable zone gets much bigger. However, ammonia and methane require a narrower range of conditions to remain liquid on a planetary surface than water does, so I wouldn't expect to actually see more than one such planet in a single system. Icy moons with subsurface oceans are likely to be very common wherever there are cold gas giants, but life that develops there won't be building spaceships, and life from other worlds will have a very hard time contacting it.

The problem with that is pretty much every chemical reaction in a living organism takes place dissolved in water. I can't think of any way a dragon would prevent the sodium from reacting the instant it was produced. An animal could almost certainly produce hydrogen, methane, propane, or whatever. However, gases would be hard to store in large quantities due to their being gas. A liquid, sprayed into the dragon's breath and aerosolized, might be more effective. As mentioned, ignition would be a problem. The pistol shrimp can produce very high temperatures by cavitation, which might work as a spark, but I have no idea if that's possible in air. On the other hand, hydrazine is literally synthesized from ammonia or urea, which animals already produce as waste and ordinarily have to use up energy (and water) to get rid of. For that matter, dinitrogen tetroxide shouldn't be too hard to make. Due to their extreme toxicity, storing them might be a challenge, but they could be used to ignite the main flame.

Actually, one species emerging and colonizing the others is probably what would happen. Let's take Drake's equation, and apply it to a single star system. We can ignore the first couple variables, because we know we have 60 planets in the habitable zone. Now: not all planets in the HZ will actually be habitable. Some will be like Venus or Mars, for various reasons. Still, I'll be generous and assume that all 60 planets are in fact habitable. Not all habitable planets necessarily develop life, but currently most scientists think that pretty much all of them do. Here's where things get tricky. Life and complex life are not the same thing. Earth has supported life for 3.5 billion years. It has only supported complex, land-dwelling life (aquatic organisms, however smart they may be, are unlikely to discover fire. No fire means no metal smelting, which means no electronics for communication, and definitely no rocket engines) for 400 million years. Now, that life seem to have been steadily increasing in intelligence, but let's say that Earth has had animals with the potential to evolve into something that could build a spaceship for about 50 million years. Now it gets worse. Humans have been smart enough to make stone tools for about 2 million years. We've only had agriculture for ten thousand. In another 1000 years, we'll probably have colonies all over our solar system, if not beyond. With 60 life-supporting planets in a system, 50 are either "slimeworlds" or have no land. Maybe 5 of the remaining 10 will develop intelligent (as in smart enough to make and use tools) life. Those that do will, like Earth, develop many such species. Eventually one will develop advanced technology like agriculture and cities, and in the blink of an eye that species will be in orbit. If that one species is environmentally conscious, they'll mostly study and explore the other 59 worlds, and in the off chance one of them has any life that's advanced to the stone age, they might share technology. If not, they'll colonize, terraform, and exploit the others. If a system with multiple habitable planets is colonized by multiple species, then unless they have radically different biochemistry I'd put my money on them coming from the same planet.

I have a lot of things to say about this article. Regarding system 1: none of the planets are actually in a Trojan configuration. For an object to be a trojan, it must be definition be at or orbiting the L4 or L5 point of another object. L4 and L5 are only stable if (a) M1 is much larger than M2 - this means you can't have a binary star with a planet orbiting at one of the star's trojan points unless the stars differ greatly in mass - and ( M2 is much larger than M3 - this means that two planets cannot be mutual trojans, which is the scenario in system 1. In general, the rule of thumb is that M1 must be at least 10X heavier than M2 and M2 must be at least 10X heavier than M3. There is, however, a way for two objects of similar mass to share an orbit. It's called a horseshoe orbit. Basically, M2 and M3 are both orbiting M1. M2 is in a slightly lower orbit, so it eventually catches up to M3. Gravitational attraction between the two causes M2 to speed up and M3 to slow down. This causes M2 to be pulled into a higher orbit, slowing down as a result, while M3 is pulled into a lower orbit and speeds up. Now M3 is moving faster, so it eventually catches up to M2, and the cycle repeats itself. If M2 is much, much larger than M3 (as is the case for the several asteroids in horseshoe orbits with Earth), M2's orbit doesn't really change. Note that a full horseshoe orbit cycle can take hundreds of orbits. For example, http://en.wikipedia.org/wiki/2002_AA29's cycle with Earth takes a total of 190 years. In any case, both the trojan points and horseshoe orbits may be disrupted by external perturbations. For example, the moon's trojan points are unstable due to the sun. The same would likely be true for planets very close together. Unless there are long-term simulations on these kinds of arrangement, I would guess that the horseshoe orbits in system 1 may be unstable, and the trojan points in system 2 definitely are. Actually, the situation is even worse for system 2. First of all, to retain a life-supporting atmosphere a moon probably has to be at least 10% the mass of Earth. Those gas-giants have 5 moons each, so at least 0.5 Earths of material in the moon system. Now, Jupiter and Saturn are both about 5000 times more massive than their moon systems, rings, etc. This means that each of those "Jupiters" would actually be at least 8 jupiter masses, and likely some would actually be small brown dwarfs, at 13 jupiter masses or greater. The larger an object is compared to its parent, the more effectively it destabilizes the orbits of everything nearby. For example, Saturn's moons get along quite nicely despite being so tightly packed, and Ceres is too small to clear out the rest of the asteroid belt. On the other hand, objects like Jupiter and Earth's moon throw anything that gets too close out of the system. Trying to pack four 8+ jupiter mass planets into the habitable zone of a red dwarf star would result in celestial Highlander. You'll end up with fewer planets than you started with, and at most one will remain in an orbit capable of supporting life. There's another problem with system 2. Any planet in the HZ of a red dwarf star is going to be tidally locked to its parent. Extremely large moons might overpower the tidal forces from the star, but the habitable moons are way too small to tidelock the gas giants. Now this means that moons will all have orbital periods shorter than their parents' rotational periods. In this situation, the moons will eventually spiral in and collide with the parents. This same effect is likely why Venus and Mercury have no moons (neither of them are technically tidally locked, but they spin too slowly for a moon in a stable orbit. Basically, a red dwarf's habitable zone might be around for hundreds of billions of years, but any moons there won't be. Pretty much all the the concepts used in this "ultimate system" are feasible. Two habitable planets in a horseshoe orbit with each other? Possible. A gas giant with habitable trojans? Possible. A "double planet"? Possible. A binary star with habitable worlds around both stars? Possible. Gas giants with habitable moons? Possible. Habitable planets around red dwarfs? Possible. Several planets in a star's habitable zone? Possible. All of those at once? Not possible.

Those stats sound about right: they're very similar to my Hatchet, which is the same diameter but a bit longer. A word of caution: unless you have very good eyesight, you will not be able to see a rocket that size at that altitude. To avoid losing the rocket, make sure that you keep your descent relatively fast (minimizes wind drift and time spent at out-of-sight altitude). For a rocket that size, you can probably land safely with a streamer instead of a parachute. The streamer should also be either highly reflective or a bright color such as fluorescent orange, for maximum visibility both in the sky and on the ground. For the vast majority of rockets, the deployed parachute or streamer will have a much larger cross-sectional area than any other component. In addition, for larger (and more expensive) rockets you can take other measures to making losing the rocket less likely. For example, personal keychain alarms are compact enough to fit in many low-power or mid-power rockets, but loud enough to be heard from over 100 feet away. While rockets often land farther away than that, if a rocket lands in tall grass being able to locate it by sound is very useful if you know the general area. Larger rockets may disperse colored powder or smoke for better visibility in the air, and be equipped with radio transmitters or even GPS modules.

All right. First of all, I strongly recommend downloading OpenRocket. It's an open source simulation program for model rockets, and it allows you to verify that your rocket will actually be stable, etc. Unstable rockets are VERY dangerous. As for TWR, you generally need to be moving at least 13 m/s at the end of the launch rod/rail for the rocket to be stable. With a 'C' engine, you will most likely use a launch rod about 1 m long. This means that your rocket must accelerate at 85 m/s^2. Add 10 m/s for the force of gravity, plus a safety factor, and your rocket needs a minimum TWR at launch of 10. If your motor produced constant thrust, your rocket would have to weigh less than 61 grams. Fortunately, real rocket motors do not produce constant thrust. According to this: http://www.thrustcurve.org/motorsearch.jsp?id=21 the C6 has a "spike" of much higher thrust at the start of its burn, with a maximum thrust of 14.1 N. With a real thrust curve, simple math won't allow you to predict how heavy the motor can safely lift. However, my simulations show that the maximum safe liftoff mass for a rocket with a C6 motor is about 100 grams. With a low-drag rocket, an ejection delay of 5 seconds is appropriate for this mass. However, higher-drag designs may require you to use C6-3 motors instead. Again, get OpenRocket and simulate your design before you build anything.

Well, the required mass ratios have already been mentioned, but: en.wikipedia.org/wiki/Single-stage-to-orbit Basically, rocket-only SSTOs are known to be possible, and a few existing lower stages, including the Titan II and possibly the Saturn V and Falcon 9 could be turned into SSTOs. The problem is, you get crappy payload fractions and no hope of reusability, so unless you went with a total Big Dumb Booster approach, you couldn't compete with the launch cost of multistage rockets. Also, interestingly, according to the Wikipedia article hydrogen performs about the same as dense fuels in an SSTO as far as payload fraction is concerned. However, as Iskierka mentioned, a hydrogen-fueled vehicle would be very low density with its fuel drained, which would make reentry easier. However, the use of an inflatable heat shield could make that not a problem. I think the biggest problem with a reusable rocket-only SSTO, besides the bit where it might not actually be possible, is that it would be just as hard to build as either a multi-stage reusable rocket or an air-breathing SSTO, but wouldn't be able to compete with either of them on payload fraction. The one other option sometimes considered is a nuclear rocket-only SSTO. This would be heavily dependent on how much thrust we can actually get out of a nuclear engine, but in theory such a vehicle could manage some decent payload fractions. The problem is, the extreme safety standards a nuclear-powered vehicle would need to meet to be allowed to fly would make refurbishing the space shuttle seem like a walk in the park.

I've already claimed 'Lucifer' for Kepler-70b The latin adjective 'lucifer' means "light-bringing." And that planet isn't just so hot it glows, it's hotter than the surface of the sun, as hot as an A9 or F0 class star. It's so hot it emits enough ultraviolet radiation to give you "sunburns" (assuming you somehow blocked all the radiation from its parent star and positioned yourself so that you received the same total radiation intensity from the planet that Earth gets from the sun). Compare that with typical glowing-hot planets, which emit mostly infrared with a bit of red light. Also, Kepler-70b has actually been through hell and back; it started out as a gas giant that was engulfed when its star turned into a red giant, but the star prematurely blew off its outer layers before the planet could be entirely vaporized, becoming a B subdwarf star and leaving the planet's rocky core still in orbit. The same is also true for Kepler-70c.

It seems like spinning the spacecraft up and down would require much more RCS fuel and more powerful thrusters than rolling a cone. Also, as you mentioned, the G-forces would be continuously changing in direction. The human body can withstand the highest G-forces lying on its back, followed by downward along the length of the spine. "Upward" g-loads in particular are dangerous, because high-pressure blood is forced into the eyes and brain and may cause blood vessels to rupture. Besides that, being in a spinning capsule with little or no outside visibility would probably give even trained astronauts motion sickness.

Also, spheres have no lift, so the spacecraft must reenter on a ballistic trajectory. This means the crew are subjected to higher G-forces than for a conical or headlight-shaped capsule, which can generate lift by reentering at an angle of attack and making its trajectory slightly shallower. I think a Vostok (or a Soyuz performing a ballistic reentry by accident) experiences about 9 Gs, compared to 6 for Apollo (from the moon) or 3.5 for Dragon (from LEO).

The 0.5-0.7 seems to only apply to liquid-fueled upper stages. Antares appears to start with a TWR of 1.3 on the upper stage. Scout's final stage has an absolutely ridiculous TWR. Like, at least 5.