sevenperforce

Latest prediction is at 2:05 EST +/- 5 hours, which gives us a meaningfully narrowed ground track. https://www.tiktok.com/t/ZTRDq4Tf5/ I estimate odds of impact over the USA at 2.8% and odds of land impact generally at 24.8%.

You can always massage the Roche limit equations. If the primary is not very dense and the moon is extremely dense then the Roche limit is lower.

If it took 8-16 launches of a Saturn V class launcher to do a single moon mission, then moon missions would be extremely rare. It does not take 8-16 launches of a Saturn V class launcher to do a single moon mission, because Starship is not a Saturn V class launcher. Zubrin is wrong and his numbers are silly and he doesn't even understand the current proposal. SLS with an extended upper stage does not have enough power to take Orion to LLO and back again, because Orion itself does not have enough dV to go to LLO and back again. The long pole is Orion, not SLS. Evidence suggests that it could not. No matter how extended an SLS upper stage may be, Orion still cannot go to LLO and come back to Earth. What proposed stage are you referring to? Boeing's original LUS proposal evolved into the DUUS proposal, which was renamed the EUS, which NASA *is* buying. ESA does not have a suitable upper stage already built and flying reliably for decades.

The term you’re looking for is “absolute ceiling” and it demarcates the altitude at which the engines at full throttle exactly match the aerodynamic drag at the optimal lift-to-drag ratio, meaning the aircraft physically cannot climb any higher. This highlights a common misconception about aircraft. When a pilot wants to gain altitude, she doesn’t pitch the plane’s nose up; rather, she increases the throttle. Similarly, when she wants to increase speed, she doesn’t increase the throttle; she nudges the nose down. What’s odd about @Spacescifi’s hypo is that he’s describing air being moved across the wing surface by magnetic fields to create aerodynamic lift. But if you can use magnetic fields to move air, you can just lift off the ground vertically. If he’s proposing a vehicle which does not have an T/W ratio greater than 1, then it will assuredly run shy of net acceleration long before aerodynamic hearing becomes a problem.

It should also be noted that there would absolutely be no direct correlation between the total weight of a vehicle and the aerodynamic heating in this scenario. A larger spaceship will have a greater surface area and so it will have more space to generate aerodynamic lift.

Oh yes I could tell. Too soon?

At least OTRAG would have worked. This is nonsense.

Do you have any particular source for thermal power plants causing anoxia in bodies of water? I’d like to read more about that.

And people are literally glossing over this HORRIFYING reality as if it is just any ordinary Tuesday.

Because scifi drives based on the power it would actually take to pull of such feats likely leave craters behind. My idea is that at a planetary spaceport you don't want craters. The ship still has chemical rockets though for VTOL because again... who wants to labd inside a smoking crater? You are correct that in order to get single-stage interplanetary spaceships, you're going to need a very high-energy propulsion system. And you are correct that the kind of high-energy propulsion system you will need to make it to another world and back without staging or refueling is going to tend to produce rather spectacular damage to anything close to the exhaust flow. However, there are a number of solutions to this, most of which are much simpler than the solutions you have suggested. As I've said in this thread and others, if you want the solutions you've suggested simply because you like their aesthetic, that's great, but they are by no means the inevitable outcome. If you want to avoid melting your launchpad, then you can always use an afterburner. Afterburners sound like something that would cause more damage, not less, but afterburners on rockets do the opposite. If you have an extremely high-energy engine and you dump extra propellant into the nozzle downstream of the throat, you increase thrust dramatically but decrease the velocity of your exhaust. Then, as your vehicle climbs, you can maintain the main throttle while throttling down your afterburner injection to increase specific impulse when you no longer need quite as much thrust. The easiest way to throttle down for landing is to simply use clustered engines. Each engine in the cluster can have its own afterburner for launch, after all, and so using fewer engines to land will give you whatever throttle setting you want. The only situation where it would make sense to use completely separate engines for landing is if your super-high-energy main engine cannot be throttled and needs to be huge in order to work at all. Then sure, use separate landing engines. You can use these same landing engines to provide added liftoff thrust, too -- the gravity drag savings will make up for the lower specific impulse. Launch ramps make sense exactly never, unless the ramp itself is providing the impulse. [snip]

@sevenperforce recently said in another thread that having lots of small engines turns out to be more efficient than one large. I'll let him explain better, or even if this is applicable to whether 9 is a special number or not. Well, the issue in the other thread was primarily combustion instability. Once you get into the world of really really big single-chamber engines like the Rocketdyne F-1, you're looking at combustion taking place inside a chamber significantly larger than a bathtub. Modern hydrodynamics modeling is significantly better than it was in the 1950s, but even so it's not perfect, and turbulent flow inside that gigantic chamber is going to mess with the ability to maintain a nice constant steady-state thrust situation. That's why the Russians built so many multi-chamber engines: This is not four engines, but one engine with four chambers. There's a single turbopump feeding propellant to all four chambers, and all four chambers are throttled (down to an impressively deep 56%) at the same rate by throttling down the single turbopump. By keeping the thrust chambers physically smaller, you dramatically reduce combustion instability. But while smaller startup launch companies DO have occasional issues with combustion instability, that isn't the real driver for using multiple first-stage engines. Those engines simply aren't large enough to encounter the endpoint problems that you get when you're working with something on the order of the F-1. AFAIK that's speculation from Tim Dodd; I don't know that it's at all accurate. Also I believe the outer engine bells on the Falcon 9 are already too close to each other to gimbal inward together. You might get a little higher pressure around the base of the center engine due to plume recirculation, reducing parasitic drag, but you're not going to get any sort of proper aerospike effect without an actual thrust surface against which to push. I do think 1:9 is a pretty good ratio, overall. Generally speaking, you need to have roughly equal ΔV on each stage for optimal efficiency. If we take LEO as requiring ~9.5 km/s of total Δv (this will allow us to treat pressure drag as part of the total Δv, allowing us to use vacuum isp for our first stage), then the total propellant on the first stage as a percentage of gross liftoff weight is going to need to be around 79.5% for hypergols, 77.3% for kerolox, 73.6% for methalox, or 68.3% for hydrolox. In modern aerospace, the dry mass of a stage is usually about 10-15% of the propellant it carries, although this varies based on propellant type. If we say it's about 13% just to put a number on it, then you can readily solve for the mass of the upper stage as a percentage of total gross liftoff weight: 10.2% for hypergols, 12.7% for kerolox, 16.8% for methalox, and 22.8% for hydrolox. If you're using common engines, your upper stage engine is going to have roughly 15% higher thrust compared to the first-stage engine thrust at sea level, and your first stage TWR needs to be about 20% higher than your upper stage. So the total thrust of your first stage as compared to your second stage is going to need to be around 13.5x for hypergols, 10.9x for kerolox, 8.2x for methalox, and 6.1x for hydrolox. So for a hydrolox TSTO using common engines on the first and second stage, you could get away with using 5-7 sea level engines on the first stage and one vacuum-expanded engine on the second stage. However, most smallsat and startup launch companies don't wanna mess with hydrolox; it's expensive AF and it's difficult to handle, and the development costs for a hydrolox engine are much higher. Same with hypergols. In fact, I can't think of any hydrolox or hypergolic rockets which use common clustered engines. So smallsat and startup launchers are typically looking at kerolox and methalox, where an engine ratio of 8-11:1 is optimal. Then why don't we see smallsat launchers with 8 methalox engines on the first stage or with 11 kerolox engines on the first stage? Well, circle packing becomes an issue. To reduce drag, you want your rocket to be as thin as possible, but you still need enough space on the back end to fit all the engines. The optimal arrangement of circles-within-a-circle is trivially symmetric at first: However, once you go over nine engines, things start to get messy. Optimal circle-in-circle packing with more than 9 circles is no longer trivially symmetric, which means you either waste valuable space on the back end of your booster or you end up with a thrust structure that is not axisymmetric. Plus, it helps with booster design if you can have a single engine at the center with an even number of evenly-spaced engines around it (so you can shut down opposite engines to maintain balanced thrust if you need to). The only circle-in-circle arrangements where those constraints are met by the optimal circle-in-circle packing are 1, 7, 9, and 19: So yeah, that's why you're most often going to see 9 engines on the first stage of a launch vehicle that uses common engines. Also note that 9 gives the central engine extra gimbal room which is especially nice if you want to recover your first stage propulsively.

This was done intentionally, in hope that de Laval reflectors make the rays parallel. Lord Vader needed to fake a moon landing and had to come up with a parallel light source to do it.

I think the OP asks about interplanetary travel not interstellar travel. To the OP's question: I second @kerbiloid; Project Rho is excellent and you should definitely spend some time bingeing that site, as I have. However, here is the answer to your question. There are basically three different paths we could take to regular interplanetary travel. I will call them Brute Force, Outpost, and Nuclear Option. Brute Force Earth's gravity well is punishing, but it does give us a nice boost from the Oberth effect. So if we're going interplanetary, the most efficient path (from an orbital mechanics perspective) is to use LEO propellant depots, fill up our interplanetary vehicles, and go from there. You can do this if you have rapidly reusable rockets; otherwise it's prohibitively expensive. Your rockets will need to be methalox or hydrolox. Hydrolox has a slightly better ISRU potential but it is not as storable as methalox so that's a tradeoff. An advantage is that you can use your engines both in-space and on the surfaces of other worlds. Outpost With enough investment and ISRU, you can stage interplanetary missions from the moon. There's tons and tons of water on the moon you can use for this and you can construct vehicles entirely in a vacuum with no worry about aerodynamics. Your rockets will need to be hydrolox. Nuclear Option Stage your interplanetary vehicles in LEO, but use near-future tech like Z-pinch fusion or VASMIR to propel your vehicle. Unlike the other options, there's no way to go to the surface and back using an engine designed solely for vacuum operations, so you will need extra dV to enter your destination orbit. However, your specific impulse will be RIDICULOUS so that's not as big of a problem, and you can even do continual-burn trajectories so that you have the lowest transit times between planets. This is the approach that is most likely to make manned missions to the outer planets feasible.

If you just like the look of a giant rocket plume coming out of a single giant nozzle on a spaceship, then say so. Rule of cool is totally valid. But that's not what you've been saying. You've been saying, "I think engineering principles would require a large spaceship to have a single large nozzle, isn't that right?" And then we say no, it isn't right, and you come up with new, even more ill-founded reasons why it really is right. If you've already decided what you want your fictional spaceship to look like and you are having trouble coming up with an in-universe explanation, that's fine. We can think of in-universe explanations easily enough. No one is really going to pay close attention to why the rocket looks the way it does and your explanations don't have to be perfect. Start with your aesthetic and work from there. For example, if you want a really large nozzle as part of your vehicle aesthetic, just lean into the square-cube law. Say that your propulsion system uses a variant of inertial confinement fusion with some sort of critical mass ignition element, so engines need to be gigantic in order to function at all. Boom, there you go. Yes, you are correct. The "regenerative coolant lining" proposed by @Spacescifi sounds more like ablative cooling than regenerative cooling. If you want a very thick nozzle for whatever aesthetic reason, make it so and then handwave it. For instance, you could propose an engine which uses inertial confinement fusion to induce helical magnetic fields which in turn produce eddy currents which are conducted to the nozzle, yeeting the propellant out of the nozzle much faster than an ordinary de Laval nozzle could sustain. Then you could just say that the super thick nozzle is necessary due to needing enough ferromagnetic material in the nozzle to properly conduct the eddy currents and shape the exhaust field. Easy.

Lol, what method did you mean?

Using a ramp doesn't save you any propellant costs. Using a ramp requires more dV than merely launching VTOL. So this makes no sense.

I think one of the reasons these perennially sprouting threads give us all headaches (and yet induce us all to answer) is that they are SO close to being reasonable questions, but they aren't. Like, we inevitably imagine that by explaining it just one more time, then something will "click" and the questions will start to make sense. And they WOULD be fascinating questions IF the correct assumptions were applied so that they made sense. But no matter how many ways or how many times it is explained, nothing ever clicks. The questions never actually start making sense. Essentially, we are being continually nerd sniped. Like, this question: This question LOOKS reasonable. It involves apparent maths. It cries out for a thoughtful response. And yet it is unanswerable because it is based on a fundamental lack of understanding of the basic principles of rocket science. Specifically, it's based on the OP's apparent notion that the "seconds of delta v" carried by a spaceship are somehow directly related to the engine or nozzle design. As @KSK pointed out, delta-v is not measured in seconds, but in meters per second, so it is already unclear whether the OP is trying to talk about burn duration or about m/s of dV. But even more fundamentally, almost all of the "specific SSTO data" are irrelevant because neither the staging design nor the pulsed vs continuous nature nor the total stage dV have any bearing on the pressures experienced by the engine; that only comes down to one thing: 100 ton SSTO capable of 3g pulse acceleration for 20,000 seconds of delta v 300 ton SSTO capable of 2g continous thrust for 10,000 seconds of delta v 1000 ton SSTO capable of 2g pulse acceleration for 1000 seconds of delta v The only information relevant to estimate the pressures sustained is what remains: total mass and acceleration rate, which together provide thrust. To push a 100 tonne vehicle at 3 gees you'll need a thrust of 2.2 MN. To push a 300 tonne vehicle at 2 gees you'll need a thrust of 4.4 MN. To push a 1000 tonne vehicle at 2 gees you'll need a thrust of 14.6 MN. These are big numbers, but not that big. A single Raptor 2 produces 2.3 Newtons at sea level, more than enough to push a 100 tonne vehicle at 3 gees. A single F-1 engine produced 6.7 MN, significantly more than what you'd need to push your 300 tonne vehicle at 2 gees; if you don't mind using dual nozzles the RD-180 will get you to 4.1 MN which is nearly where you want to be. The original Shuttle SRBs produced 14.7 MN of thrust each, plenty to push 1000 tonnes at over 2 gees. So if you want to get an idea of how large or thick the nozzle needs to be for any desired application, just look at the size of the actual nozzles of actual engines that actually produce that much thrust. First of all, it's SSTOs, not SSTO's. There's no possessive. Ugh. But more to the point, this assumption is wrong. In order to get increased performance in the world of orbital rocketry, you need to increase exhaust velocity. That doesn't necessarily translate to higher nozzle pressures, though. Behold the magic of a de Laval nozzle: The green "p" line above is the gas pressure. The beauty of a de Laval nozzle is that when you pass the shock at the nozzle throat, your exhaust flow becomes supersonic and so pressure and temperature both drop precipitously in favor of increased gas velocity. All other things being equal, an engine with higher performance will have higher chamber pressure and higher chamber temperature than an engine with lower performance, but once the gas leaves the chamber and enters the nozzle its pressure and temperature fall exponentially, so you never actually need a "large and thick" nozzle. Unless, of course, you just want a thicc nozzle for aesthetics...in which case, get it. I see..... so to even warrant a large and thick nozzle I would need an SSTO both massive and heavy.... in other words like project Orion only a fusion rocket is propelling it. Thanks... so apparently I can go large... but yeah... the plume is likely going to blow up anything it hits nearby. No part of what @KSK said would suggest, imply, or encourage the conclusion you subsequently drew. I do believe you are wrong in so assuming. As I explained, the pressure in the nozzle drops off exponentially, so there is not a linear (or even near-linear) relationship between specific impulse and nozzle pressure. But the question you should be asking is whether a single large combustion chamber is better at handling high temperatures and pressures than multiple smaller combustion chambers. The answer to that question is a resounding NO. The square-cube law is often your friend in the world of engineering. Bigger is usually better, because the volume in a given space increases faster than the surface area of the enclosing solid. For example, a 10-meter spherical tank will carry more propellant for its weight (at a given tank pressure) than a 5-meter spherical tank. That's geometry for you. But that doesn't always translate. In particular, it doesn't translate to rocket engine combustion chambers. Combustion chambers are receiving a constant flow of propellant and so volume is constant; making it bigger doesn't change the tensile strength requirements of the chamber wall. Chamber wall pressure depends on internal area, which is quadratic; chamber wall weight depends on thickness times circumference, which is also quadratic. Double the pressure and you double the weight. Bigger is not better and bigger is not worse. NOTE: I'm using the term "combustion chamber" here but it does not mean we're talking about conventional chemical combustion. A combustion chamber is merely shorthand for "that place where exhaust gets really energetic" -- it doesn't matter whether the potential energy comes from chemical sources or nuclear sources or whatever other sources you invent. Bigger chambers do have a tendency to induce combustion instability which is why the Russians love their single-turbopump multi-chambered engines. Further proof that a bigger chamber doesn't actually help and may often hurt. You may or may not know what a torchdrive is. If you're talking about an engine which can execute a true Brachistochrone trajectory between planets then it's not going to be dependent on anything remotely similar to chemical rockets. It's like asking, "What would a car look like if it could fly like a plane but without wings?" It's just not a meaningful question. If you just mean a spaceship that can produce high thrust and high specific impulse, enough to fly between planets on a single stage, then you're just going to need a high-density, high-power, and high-specific-energy propellant combination. There are any number of ways to achieve this in a science fiction universe. Pick one. But for any spaceship which is capable of providing single-stage-to-destination performance with a de Laval nozzle, the engine design can be as big or as small as you want it. But a cluster of small engines is usually better, because it allows you better throttling, better differential pitch and yaw control, and lower overall loading on your roll gimbals. Nope, not even slightly. The Superheavy booster carries 3600 tonnes of propellant and it could easily reach orbit in a single stage if launched without Starship. It does not need an F-1 engine nozzle. When and where in an addled universe did you come up with this ramp idea? Just launch from a floating platform and use pumps to create a water deluge. Or just launch from underwater. Ramps are one of the worst ideas I have seen from you, to date, so far.

Update: using only solids and no control inputs (just a KAL-1000 to time the staging events) I've now got a three-stage fire-and-forget launch vehicle.

Sure, here it is. The small solid rocket motors provide some impulse but mostly just spin. The larger solids help it get off the ground and out of the atmosphere quickly. The main engine is a Reliant so there's no gimbal. I put a Terrier on the upper stage but I disabled gimbal, as you can see. The tiny solids up top are just to help with rotation once I'm out of the atmosphere. The launch: No Thud required.

I made one pretty trivially. Timewarp.

So what if I just do a spin-stabilized rocket?

And landed.

It is grounded in stuff we can actually do, with a single exception, which is all of it. Your latest fixation seems to have drifted to "fusion pulse with large and thick bodied nozzle with uber magnetic fields". None of those things make sense in connection with each other. If you have the capacity to control nuclear fusion then just heat your propellant using nuclear fusion and use regenerative cooling on your nozzle; no need for magnetic fields or a "thick bodied nozzle" (whatever that is supposed to mean). I'm not sure why you would imagine that a fusion rocket engine delivering over 8,000 tonnes of thrust would have problems with air ENTERING the chamber. If your wingless vehicle can accelerate into orbit after yeeting off the end of a ramp, then its TWR was already greater than 1 to begin with, so the ramp and wheels were all unnecessary to begin with. No part of this makes sense. If it has no wings then it will fall vertically into the ground.

If it's expendable, why would it need grid fins at all? That's why this whole notion just seems so silly -- it's like you're not thinking about the vehicle as an actual vehicle with actual meaningful parts. You're still going to need a fairing for whatever payload you're payloading. Even if the fairing is no longer than the one on Falcon 9 and only as wide as the rest of the vehicle, it's still going to have a mass of at least 5.7 tonnes. You're going to need a mechanism to open and close that fairing, but let's ignore that. You claim that the weight of the grid fins can be reduced by over 80% "using ceramics" but you give absolutely no explanation or citation for these magical featherweight ceramics; aerospace/industrial ceramic parts are typically closer to half the weight of equivalent metal parts. In the video @tater linked, Elon notes 58 tonnes worth of engines, 80 tonnes tank and structural mass, 20 tonnes for the interstage and grid fins and avionics, and roughly 20 tonnes of residuals after landing. So that would be 178 tonnes right there. Of course we're going up to 33 engines instead of 29 so that brings us up to 186 tonnes. Let's just hypothesize that they can use "ceramics" and other savings to shave 10 tonnes off the interstage and grid fins and avionics. You talked about going to 3mm steel instead of 4mm steel but that's only going to be for the tank walls themselves, not for structural mass. Let's say they can cut 10 tonnes off of tank mass. Add in the weight of the fairing and you're looking at 172 tonnes landed mass. You claim landing gear is 3%, which would be 5.2 tonnes, but I don't know how you're going to manage that; Falcon 9's four landing legs have a total mass of 2.4 tonnes for a vehicle with a dry mass of around 25 tonnes. So 10% would be closer to the right number. Let's be generous and put it at 6% (closer to your 3% number than to F9's 10% number) and peg the landing gear at 10.3 tonnes which takes us up to 182 tonnes. Let's pretend, somehow, that you could put TPS on the butt end and somehow prevent the engines from burning to slag. Then it can enter rear-first. It's absolutely not possible, but let's wave our hands and pretend it is. If Apollo was 15% let's give Superheavy a 10% landed mass margin. So that's 18.2 tonnes of TPS, bringing our landed mass up to 200 tonnes. I'll ignore the mass growth requirements on the landing gear, etc. So what, then, does the landing burn look like? We can assume that like Falcon 9, Superheavy will need to initiate its burn in the transonic regime, at around 310 m/s vertical speed. It needs to limit its hoverslam to around 3 gees to avoid damage; the landing burn will thus take at least 11 seconds (probably longer but we're being generous to your SSTO idea). That 11 seconds will be 108 m/s of gravity drag, so the total burn needs to be 418 m/s. At the sea level specific impulse of 330 seconds, that's 27.8 tonnes of landing propellant, bringing our re-entry mass to 228 tonnes. We will again ignore mass growth, here related to TPS. In space, you'll need roughly 100 m/s of dV to deorbit. We'll ignore ullage and maneuvering RCS and everything else like that. You'll benefit from the higher vacuum specific impulse, so that's nice; it'll cost you a little over 6 tonnes of propellant, bringing our effective dry mass up to 234 tonnes. So let's plug all this into Silverbird and see what it tells us now. This *is* a restartable stage so I checked yes for that box (starting engines uses extra props). I set residuals at 0% because we are already factoring them into dry mass, and I reduced propellant by the amount of propellant needed for deorbit, landing burn, and residuals. So yes. If you can find a way to magically keep the grid fins and butt end of Superheavy from melting to slag on re-entry with 18.2 tonnes of TPS, then you could convert it into a reusable SSTO that delivers ~97 tonnes of payload to LEO. How you plan to do that, or what 97-tonne payload you plan to launch in a 13x9 meter fairing, is anybody's guess.

Probably not better in the sense of absolute resolution. Voyager 2 came within 4950 km of Neptune during its flyby, close enough for its Imagine Science System cameras to achieve 500 meter resolution. We are 19 AU away from Neptune; you would need a primary mirror 35% larger than the dwarf planet Ceres to match that resolution from Earth's location. Voyager 2 only came within 84,000 km of Uranus, so the resolution during that flyby was not as high as the Neptune flyby. Even so it's not much better; you would need a primary mirror just slightly smaller than Ceres to match that resolution from Earth's location. However, JWST can take more pictures of Uranus and Neptune, and it can take them over a longer period of time, and it can take them in a broader spectrum, allowing us to see deeper. We can monitor weather and identify cloud layers and all kinds of cool things.