sevenperforce

Is that dV the dV beyond TLI or the dV from LEO? or from C3=0? The PLF and LES are just negligible compared to the really large payloads and rockets we're talking about. They are added to the dry mass of the first stage. It is assumed that they are dropped at or around the first stage burnout.

Someone else (might have been you) playing around on the spreadsheet found that by stretching the first stage tanks to 40 meters and and the second to 25, using nothing but Raptors, you can launch 91.5 tonnes to TLI... ...using 1960s construction techniques and mass fractions.

............and today is Wednesday, isn't it. This shutdown is addling my brain.

Where is the livestream?

Report on plans for lunar surface activity: https://forum.nasaspaceflight.com/index.php?action=dlattach;topic=48676.0;attach=1630122;sess=50751

While you're waiting... I've gotten the Build a Better Saturn spreadsheet closer to where I want it after cleaning up some of the kinks in the math. Please take a look, play around with it, and let me know if you see any inconsistencies or errors I need to know about. https://docs.google.com/spreadsheets/d/1UNuPfjZE-YVIrmLINiUMcOzJk-voZ2TqM6R7MgyaM-I/edit?usp=sharing

Asteroid mining may be a pipe dream that will never be worth anything. Or it could be ridiculously lucrative. The fact is that we don't currently know very much about the early solar system and planet formation. We know enough to know that it might be very lucrative, but we also know that it might not be. But there are new synergies that become possible when launch costs drop. There is plenty of budget money out there for planetary geology, geomorphology, and cosmocheistry. Just like publicly-funded study of terrestrial geology was used by private companies to find oil and coal, it's not difficult to imagine a future where public research into the conditions of the early solar system produce commercially-useful data. The reason this hasn't happened to date is that everything right now has to go through NASA and launch costs are extreme. If launch costs drop low enough, then there could be universities with the budgets to send up an asteroid sample return or even asteroid capture mission. If the data is good (and at this point we have no idea) then commercial entities would start sending up their own prospecting missions. On another note, I've gotten the Build a Better Saturn spreadsheet closer to where I want it after cleaning up some of the kinks in the math. Please take a look, play around with it, and let me know if you see any inconsistencies or errors I need to know about. https://docs.google.com/spreadsheets/d/1UNuPfjZE-YVIrmLINiUMcOzJk-voZ2TqM6R7MgyaM-I/edit?usp=sharing

Flying without any solid boosters, the Atlas V can reach a 28.7-degree LEO with 8210 kg of payload using its large, 5.4-meter PLF and with 9800 kg using its smaller, 5-meter PLF. My model gives it 9957 kg to LEO with the 5.4-meter PLF and 10,030 kg to LEO with the 4-meter fairing, but warns that second-stage T/W is dangerously low, which would require a lofted trajectory and higher gravity drag on both stages. My model adjusts for T/W ratio, but only in terms of how burn time impacts average gravity drag; if a trajectory is significantly different than the Saturn V's then the total gravity drag will be underquoted. So getting relatively close to the actual Atlas V performance using a model designed around the Saturn V ascent profile is encouraging. If I can, I'll add the option to tack on SRBs and see if I can get it to accurately match Atlas V 551 as well as SLS. I went in and added a dV calculator to the spreadsheet so it displays wanting or excess dV from TLI at the very top, hopefully enabling you to compare to various earth escape missions if you know the dV past LEO. If you check the reference sheet you'll see that the free-return TLI used by the Apollo missions required 3,039 m/s past LEO. Required dV from LEO to Mars intercept is 3.6-4.3 km/s, so based on my model the Saturn V could have sent 31.5-40 tonnes direct to Mars. I know that Von Braun's analysis of the Saturn V gave it 45 tonnes to C3 so that's probably in the ballpark.

If you go kerolox-methalox-hydrolox with a 10.1-m diameter all the way up, you can hit 118 tonnes to TLI using RD-191s, Raptor Vacs, and BE-3Us. I still haven't found a good use for the RL-10.

I fixed the issue I was having with stage volumes and dry masses, so now if you plug in everything for the Saturn V exactly as-is, you get the right answer, and the same for other real-life rockets. Can you take another look and let me know if you find any problems? https://docs.google.com/spreadsheets/d/1UNuPfjZE-YVIrmLINiUMcOzJk-voZ2TqM6R7MgyaM-I/edit?usp=sharing With the new data, here's what I'm getting: Replacing the old Saturn V with the exact same engines but newly-built, lighter stages with common bulkheads should increase TLI payload from 48,600 kg to 56,000 kg. Adding new engines increases TLI payload to 84,000 kg. One winning combo appears to be 3 BE-3Us on the third stage, 15 BE-3Us on the second stage, and a whopping 59 Merlin 1Ds on the first stage. You can bump it up to 85 tonnes by swapping out those 59 roaring Merlin 1Ds for 25 RD-191s...what they lack in liftoff thrust and prop densification, they make up for in specific impulse. Going methalox-only (Raptor SL and Raptor Vac) easily breaks 100 tonnes to TLI, with the same tank dimensions as the Saturn V. And that's without changing tank diameter.

It doesn't really depend on field configuration or field strength, no. The plasma inducement is the result of compression heating. If you had a way to use magnetism to interact with the plasma and slow down, you wouldn't have the compression heating and so you wouldn't have the plasma to work with. You can, however, use magnetism in the thermosphere to slow down more slowly and reduce re-entry speed that way.

My point is that magnetism doesn't do anything to plasma under those conditions.

So will electromagnetism...like a heat shield.

What is it about magnets that you suppose will help?

In some ways. However the bigger your vessel, the more cross-section you have to get hit by a meteor large enough to penetrate your prop tanks, and if you need those, you're hosed. That's another reason why the Starship's use of header tanks for landing is a good idea. Many of the SSTO spaceplane concepts from the 80s and 90s used wet wings that would be filled with props (usually liquid hydrogen) on launch, then vented to vacuum once orbit was reached. They used hypergols for deorbit. Their increased drag cut down on g-loading and re-entry heat, and then they could be used for aerodynamic lift to get a nice clean rolling landing for "free". The sort of parachute which works for re-entry is not the sort of parachute which will let you land safely, Unlike in real life, KSP has a sharp cutoff of atmosphere (and hence drag) at 70 km, so you can't get truly slow and gradual deorbits.

Pressure-fed hot-gas thrusters might give you enough dV to do it anyway. Have the Raptor autogenous press feed into dual-redundant maneuvering prop tanks that vent to the main tanks intermittently. In an abort, the vent lines are severed and sealed, and those tanks are used for abort and landing. But this would be for earth to orbit and back, not for landing elsewhere.

It's not actually a bad idea. If you make surface area enough, you can get meaningful drag at very high altitudes, before you hit typical "re-entry conditions" at all. The "fire zone" for re-entry is the boundary between the stratosphere and the mesosphere -- 49 to 80 km. Any drag higher than that simply decreases your speed in the fire zone and thus decreases both g-loading and peak heating. If you can scrub a significant portion of your speed high in the mesosphere, you drop peaking heating and g-loading considerably. Elon has actually mused about this -- giving the Starship "dragon wings" of deployable stainless steel ribbons. It's just very hard to get right, and almost impossible to get right without sacrificing the drag surfaces...which in turn means a separation event during the worst part of the flight envelope.

Ahhh, I bet I mucked up volume by taking stage lengths as inclusive of engine length and not testing against actual volume. Will rework.

I went in and added the J-2, the F-1, and the RD-191 for comparison. I added most of the stats for the HG-3 but I wasn't able to find a good source for its dry weight; if you see one, let me know. Any other engines you want to see? Using the F-1 and J-2 with the other info from the Saturn V is giving me too little dV on the second stage, by a long shot. I am going to look at my structural fractions and propellant volume to see if I can figure out why.

I originally had an "add six engines for each additional engine diameter" equation, which obviously produces the wrong result. I just updated it to calculate based on hexagonal area, so it will fault if the number of engines exceeds of the number of engine-width hexagons that can be inscribed in the stage cross-section.

I think I went linear after 20 when I should have gone quadratic. Redoing some math.

My logic ran up to 20 and then defaulted to an assumed honeycomb pattern above that, but the math may be wanting. Lemme look up higher-order circle packing.

So would I, though it won't be, and I don't like having no abort modes. A HOTOL, an air-augmented mini-Raptor spaceplane with methagox-thruster abort motors might work.

Something y'all might like... Over in the "Build a Better Saturn V" thread, I just posted a link to a spreadsheet I created that will let you play rocket legos with a three-stage moon rocket. It also holds a wealth of information about mass fractions, engine weights and performance, different kinds of drag, and the like. Here it is. It led me to some interesting discoveries about Raptor vs hydrolox for high-energy launches... Hydrolox for a first-stage engine has a bulk density of up to 363 kg/m^3. On an upper-stage engine you want it to run more fuel-rich so it is more like 344 kg/m^3 (that's the RL-10C-2 with its otherwise-impressive 466 s of isp). Compare to subcooled methalox at 903 kg/m^3 and dramatically higher T/W. What he said. If SS/SH works well enough that crewed orbital destinations are cheap, I think we might have a good business case for a small, rapidly reusable crew SSTO shuttle (or perhaps one that uses drop tanks) with fractions for full-envelope launch abort. That would be delightful. EDIT: Re that spreadsheet... ...iIf anyone has suggestions for other engines to add, let me know. I limited it to large engines that are currently (or will soon be) in use on American launch vehicles, but I can go back and add the F-1 for reference or others for comparison.

As could have been predicted from some of my earlier posts, I spent entirely too much time this weekend setting up a spreadsheet that would let me play rocket legos with a three-stage moon launcher. It's actually quite detailed; it automatically accounts for the various sources of drag based on the chosen parameters and gives you T/W ratios and burn times for each stage to make sure you aren't building something that can't actually fly. The spreadsheet lets you choose a payload mass, an overall vehicle diameter, and whether you want a payload fairing, a launch abort system, and so forth. Then, for each stage, you select the engine you want, choose how many engines you plan on clustering, and the total height of your stage. You also have to choose what I call "structural density", a rough measure of the weight of tanks at a given length and diameter, but I use exemplary figures from the Delta IV, SLS, Atlas V, Soyuz-2, and so forth to constrain within reasonable limits. It adds the weight of each interstage based on the length of the engine you choose (and adjusts for engines with extensible skirts). Default is a constant tank diameter but it can be manually changed. It will alert you if your first stage burns out too low, if your second stage is oversized and reaches orbit on its own, or if your fineness ratios are out of envelope. I created a "sandbox" version of the spreadsheet here if people want to play around with it: https://docs.google.com/spreadsheets/d/1UNuPfjZE-YVIrmLINiUMcOzJk-voZ2TqM6R7MgyaM-I/edit?usp=sharing If you want to change the other parameters you'll need to copy it into your own drive and change the permissions. Playing around with numbers produced some interesting results: A simple Saturn V clone with the same tank sizes (but with modern construction) would throw 53 tonnes to TLI (4.4 tonnes more than Apollo 17) with 7 RD-180s on the first stage, 7 BE-3Us on the second stage, and 3 BE-3Us on the third stage. Swapping out BE-3Us for the RL-10C-1-1 is not a good idea. They have so little thrust that even if you fill up the whole 6.6-meter third stage with them so they have no room for gimbal, their gravity drag losses are too high to hit TLI. If you swap out the RD-180s for Merlin 1Ds, you need only 36 due to their high T/W and the fact that they mean densified propellant in the first stage. If you bump up the third stage to the same 10.1-m diameter as the rest of the Saturn V, you can get some very nice results by using kerolox on the first stage, methalox on the second stage, and hydrolox on the third stage. With no change other than that increased third-stage diameter, you can throw 75 tonnes to TLI with 41 Merlin 1Ds, 6 Raptors, and 5 BE-3Us. If you imagine an expendable 3-stage Starship with the same first and second stage sizes, 8 Vac Raptors on the second stage, and clone the second stage as a third stage with 5 Vac Raptors, you could throw 100 tonnes to TLI. The Superheavy Booster, stripped of its grid fins and landing legs, could put 39 tonnes into LEO as an SSTO.