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JoshNH4H, you and Hop and Louis, don't desert us. This discussion is getting very interesting. Air drop launch has entered the fray, perhaps commercially.
I will keep my posts short here. Any long and technical stuff, I will post over at my "exrocketman" site. Y'all know where that is.
GW
More ice on Mars. True, but you have to get there first.
Without a space race motivated by something other than rationality (like last time to the moon), you have to bootstrap your way there, slowly. Too many with political clout dig in their heels about space travel. Gotta do this gradual to sneak it by them.
Right now, the moon is far easier to reach with the puny chemical rockets we have. The propellants for an initial trip to Mars (perhaps along with other stuff intended for use here at home) might be made robotically on the moon, instead of launched up from Earth. It's energetically very easy to ship stuff from the moon. Men might have to get involved in that shipment for safety's sake, which is a good excuse to fly beyond Earth orbit once again.
GW
Wait, I'm just trying to say that we don't know yet what Mars might have to offer economically. Whatever it is, it would have to be a very high value-added physical commodity, to justify the shipping costs. But, it might be an intellectual property, capable of being transmitted electronically. Or something else we simply haven't thought of yet. It will become clear, just give it time once folks are there.
I kind of doubt plain rocks would ever be that valuable. Lots of gold or diamonds might be, at least for a while before the market gets flooded. A supply of high-grade uranium or thorium might be worth it, if enriched and/or bred on Mars to high-grade fission fuels before transport to Earth. (Of course, that last would assume we get over our irrational fears about nuclear power, and proceed with rational solutions to the very real problems of waste disposal and plant vulnerabilities to natural disasters.) Not very likely for a while yet.
I quite agree that what I called "prospecting" would naturally occur, once manned bases get put on Mars. And having robots there working with the men at short distances, is exactly what needs to be done. I rather think we ought to do some serious exploring, based from orbit, at many landing sites, in a single first mission. Then the best 2 or 3 sites get the initial surface bases on the next mission, after we've had time to digest all the data from the first mission.
That's the most practical way to identify what actually might support a future colony. If you don't do that, the colonies never prosper: Spain's mistake 500 years ago with an extractive-mining-only model. Most of those colonies today are 3rd-world countries still.
GW
Once somebody has actually been to Mars the first time and brought some rocks back, I kind of doubt that Mars minerals will remain as valuable commodities. It's a perception thing.
The real value to be derived is as yet unknown, because the exploration is not done. You have to find out what's there and where it is (unevenly distributed, just like here), before you can spend successful time learning to live off the land and figuring out what might actually be useful for trade ("prospecting").
But it is there. Somewhere. You just have to trust that this will be true, because it always has been before, here on Earth.
And you have to get all that exploration and "prospecting" done before it is probable that any colony you plant will be long-term successful. That's history. Just because we're talking about another planet makes no difference to that history lesson. We don't want any failures like Roanoke, or very-marginal survivals like Jamestown.
GW
Hi Rune:
Glad to see you on the new forums. Hi Adaptation - I see you saw the same thing Rune and I saw.
I, too, saw the news articles about a Rutan carrier plane for a Spacex rocket. I believe the rocket is a derivative of the Falcon-5 (5-engine) design they did not originally build, instead going straight to the Falcon-9 (9-engine) design.
The wing is for dropped-rocket pull-up to the steep path angle for a non-lifting gravity turn trajectory. That’s something the carrier plane cannot do at high altitudes in the thin air, especially since it’s a subsonic airplane. 50,000 feet is just about ceiling for most practical designs (the U-2 being an exception) – there is little speed margin between stall speed and maximum speed at such thin-air conditions.
Throwaway designs would drop the wing as unnecessary right after pull-up. A reusable stage might fly to a landing on that wing, at the cost of smaller upper stage(s) and payload. The wing costs weight allowance.
This carrier plane idea is a reprise of the older Pegasus system that Orbital Sciences flew but marketed unsuccessfully about a quarter century ago. Pegasus was a throwaway winged two-stage rocket dropped from a DC-10 airliner. Pretty much the same upper stages now sit (without the wing) on top of an ex-ICBM first stage. Orbital now calls that system Taurus, which is a conventional surface-launched vertical ballistic rocket system.
If anyone can make subsonic air drop work economically, it would be Rutan’s bunch and Spacex. Both understand the need for a small logistical “tail” as the real key to inexpensive access to LEO. Rutan himself just retired, or so I heard. The altitude helps reduce the size of the rocket a little, but is the least effective of the three variables: speed, path angle, and altitude, in that order of importance. The pull-up wing helps a lot more.
Speed is the toughest to achieve: high supersonic or low hypersonic would be really advantageous in reducing the size of the rocket. But, it’s almost impossible to achieve with turbine, and none of the combined cycle development efforts have ever gone anywhere. Ramjet (high-speed designs, not the pitot inlet kind) might work if used separately-but-in-parallel with rockets. I think, based on design analysis numbers I have run, that M5-to-6 is achievable for drop, at altitudes near 50-60,000 feet, complete with carrier plane pull-up to the high path angle (using ramjet and rocket simultaneously). I know an outfit that wants to try this. I may get to help them start trying it, next year late.
GW
It only takes a few centimeters of dirt at 0.38 gee to provide the overburden pressure necessary to keep ice from subliming away in the near vacuum that is Mars's atmosphere. Water-as-ice is easily stored outdoors without a container, if you just bury it in a "shallow grave".
Methane clathrate is only stable at 2 C and 300-meter Earth ocean pressures, which is some 30+ atmospheres pressure. To store it without a container on Mars would require very deep burial indeed. A pressurized container on the surface would be the better deal, if methane clathrate is really what you want to store. Actually, plain liquid methane would be easier to deal with.
I don't think chemical energy storage has been adequately explored, for application here or on Mars. There's more in this world than just batteries and water electrolysis into H2 and O2. What, I dunno.
GW
Scanning through the final page of this conversation, I noticed two things: (1) the huge difference between exploration and colonization efforts is becoming recognized, and (2) there is a need to explore further at Mars, because we don't really know what resources are really available there, not yet.
The paper I gave at the recent August convention of the Mars Society in Dallas deals with those issues, plus the "prospecting" phase that fits between exploration and colonization. Exploration can be done with the tinkertoys we have right now, although it would be easier to do effectively if we were to resurrect the old solid core nuclear thermal rocket technology that did everything but fly 4 decades ago. Exploration seems to me best based from orbit, from which you can do science while you watch over the team on the surface.
"Prospecting" seems to me to be best done with a few surface bases where we learn how to live off the land and to produce some sort of commodity (as yet unknown) that would make a trading colony viable. The same existing spaceflight tinkertoys could be used for this as well as exploration.
Exploration answers two deceptively-simple questions: (1) what all is there? and (2) where exactly is it? And I do mean those questions exactly as worded, that is not slang or dialect. Answering these requires (as one of many parts) the drilling of samples deep under the surface: kilometers, not centimeters. I'd recommend making a bunch of widely-separated landings all in one trip, effecting what amounts to a planetary survey, based from orbit. This concept is way far more than an Apollo-style flag-and-footprints mission. Yet it can be done with chemical or nuclear thermal ships built in LEO massing a few hundred tons, not some ridiculous "Battlestar Galactica". Depending upon who leads it and who does the work, such a mission could be done for 10's of $B (billions), not $T's (trillions). But it cannot be done Apollo style, not for that price. NASA's underestimate for an Apollo-like mission is $450B, last I heard.
Colonization comes later, and actually requires really big ships to be affordable. We don't have anything like that yet. In the absence of any better candidate technologies, I'd suggest the old nuclear pulse propulsion idea, perhaps updated a bit. Half a dozen vessels like that could enable colonies all over the solar system, spread over a century or so. The place to build and test stuff like that safely is the moon.
Just some out-of-the-usual-path ideas for your discussions.
GW
Here's an odd idea. Water seems to be ubiquitously available on a variety of celestial objects, although some purification may be needed. Why not store it and ship it as ice, which requires only the mildest pressure to prevent sublimation into space, and has considerable structural strength in and of itself. If you ship it robotically to your destination months ahead of time, then you can robotically use small-power-level solar or nuclear power to electrolyze it into hydrogen and oxygen.
The fact that electrolysis is inefficient is no problem if the power is essentially free, as with solar or nuclear. The fact that electrolysis production rates are very low is no problem if you have months to get the job done before men ever arrive.
For example, mine ice from the south pole of the weak-gravity moon, and send it with an electrolysis plant to Mars orbit robotically. Do it with a min-energy Hohmann transfer. Have a supply of LH2 and LOX waiting on you when you arrive. Use that supply to support landings and the return trip. Once Phobos has been explored, you may (or may not) have a supply of ice in situ in Mars orbit. But either way, you have the first manned Mars mission covered.
Electrolysis creates H2 and O2 at 2:1 molar, which is 8:1 oxygen:fuel by mass. Even LH2-LOX engines do not use it stoichiometric, they run rich on H2, so there is always excess oxygen to breathe available. Breathing oxygen is really abundant if your engines are nuclear thermal, which typically use only the hydrogen.
GW
I cleaned up my reusability study and illustrated it, and posted it over at http://exrocketman.blogspot.com a few minutes ago. It was based on parachute-slowed ocean impact recovery.
I, too, saw some stuff on Spacex's website about landing the first stage of Falcon-9 on its tail. It was a bit unclear to me exactly how they propose to do this, but I had the impression of parachute-slowed fall to a last-second rocket-braked landing on landing legs. The landing legs and the extra propellant would have the same effect as increased inert weights for ocean impact. Both scenarios have some sort of chute system.
Falcon-Heavy is supposed to fly for the first time out of their new pad at Vandenburg AFB next year, last I heard. It will use the Merlin 1-D, which then retrofits onto Falcon-9 and Falcon-1 later. It's the 1-D that got them to 53 metric tons to LEO, instead of 34 tons.
GW
I re-ran the design trade study I did on the two-stage Falcon-9 with constant payload instead of constant launch weight. I used better estimating models for re-scaling the interstage ring and the payload shroud, which did not affect a payload-constant calculation. What I got differed from the original reusability study for increasingly-heavy stage 1 hardware differed only in the second decimal place.
Stg1 inert fraction payload/launch
4.082% 3.136% (baseline Falcon-9)
8.163% 2.516% (vs 2.512 before)
12.245% 1.898% (vs 1.888 before)
With launch costs proportional to launch weight (all else being equal), a drop in payload fraction is a big increase in cost per unit mass of payload delivered. On the other hand, first stage reusability is very difficult to obtain because of (1) the challenges of reentry at M11.8, and (2) the challenges of ocean chute impact, thermal shock, and saltwater corrosion.
The shuttle SRM’s were sometimes reusable, sometimes not, at inert fractions around 10% for a solid motor pressure vessel case. I rather doubt that Falcon-9 first stage reusability is obtainable much below inert fraction 10%, which would be pretty close to a payload fraction of 2% vs the “stock” 3.1% value. That’s about a factor 1.5 hike in cost per unit mass to LEO, all else being equal. Somehow I don’t see routine reusability of a first stage offsetting that, since hardware and propellant costs pale into insignificance beside logistical costs.
So, I looked at the 3-stage option, based on throwaway second and third stages, and baselining a throwaway first stage for direct comparison to Falcon-9. I would then double and triple the first stage inert fraction to see what effect that had on payload fraction overall. I used payload shroud weight = 15.2% of the payload weight inside, same as Falcon-9. I used interstage ring weight = 0.815% of the weight supported upon it, same as Falcon-9, but both places in a 3-stage vehicle.
I started with mass ratios of 5 and Isp = 304 sec in the third and second stages, similar to the Falcon-9 second stage. I used a mass ratio of 4 and Isp = 289.5 sec in the first stage, similar to the Falcon-9 first stage. For the first stage I used the same 1.10 knockdown for gravity and drag as for Falcon-9 first stage. For the third stage I used the same knockdown of 1.05 as the Falcon-9 second stage. For my second stage, I used an intermediate knockdown of 1.07, being drag-free but more vertical. These gave me an estimated delivered delta-vee of 41,400 ft/sec vs the 26,900 required, so I knocked down all three stages’s delta-vees by 1.506, and recomputed required mass ratios. They are 2.45935 first stage, and 2.84251 in both the second and third stages. Stage 1 dropoff is 7620 ft/sec vs 11,820 for Falcon-9, stage 2 dropoff is at 17,170 ft/sec, and stage 3 burnout is 26,900 ft/sec, same as stage 2 burnout with Facon-9.
Stage propellant fraction is (MR-1)/MR, for 59.339% in the first stag and 64.820% in stages 2 and 3. The remainders must be split between inert fraction and stage “payload” fraction. I figured things as constant payload, top-down. It was easier to cope with the interstage ring weights that way. I used 4.2% inert (one engine) in the third stage, similar to Falcon-9’s second stage, for a stage payload fraction of 30.980%. I used 5% inert in the first stage, similar to the multi-engine Falcon-9 first stage, for a payload allowance of 35.661%. For the second stage, I used an intermediate inert fraction of 4.6%, reflecting multi-engine, but only a few. That stage payload is 30.580%.
The variation was to double and triple first stage inerts (5,10, and 15%), for stage payload fractions of 35.661, 30.661, and 25.661%. I assumed the shroud and first-second interstage ring drop off with the first stage, and the second-third interstage ring drops off with the second stage. Third and second stage weight statements are therefore identical for all three configurations:
Payload 23050 lb
Stg3 dry 3125 lb
Stg3 prop. 48,228 lb
Stg3 ign 74,403 lb
2-3 ring 606 lb
Stg2 payload 75,009 lb
Stg2 dry 11,283 lb
Stg2 prop. 158,996 lb
Stg2 ign 245,288 lb
1-2 ring 1999 lb
Shroud 3500 lb
Stg1 payload 245,288 lb
Stg1 dry at 5% 35,163 at 10% 81,794 at 15% 146,596 lb
Stg1 prop 417,303 485,354 579,925 lb
Stg1 ign = WL 703,253 817,935 977,308 lb
Payload/WL 3.278% 2.188% 2.359%
If you graph the two trends, it is clear the 3-stage option is more tolerant of higher inert weights in the first stage. Combine this with a lower first stage fall-back speed, and reusability seems more certain at 10% inerts, and with a higher payload fraction (nearly 3% 3-stage vs a bit over 2% 2-stage). My conclusion is that 3 stages is a better option than 2, if the first stage is to be reused. The drop from 2-stage non-reusable payload fraction is actually quite small (3.1% to about 2.8%). This is because the all-throwaway 3-stage vehicle actually has a better payload fraction than the 2-stage (3.3% vs 3.1%).
This does raise the question of whether 4 stages might allow first stage reusability at even better payload fraction, or the same payload fraction with both first and second-stage reusability. I leave that for others to investigate.
The main lesson here is you have to do something different to get a different result. Reusability will require a greater inert weight fraction to cover recovery gear, and confer the strength to survive better. It just ain’t gonna happen in the 4-8% inert range. This study sort-of points toward 10% inerts, at least. The more inerts you have to cover, the more stages you need to use, to be tolerant of lowered mass ratio in each stage.
But at least we know the job really can be done, and one well-proven way to do it.
GW
I posted two things, a breakdown of Falcon-9 based on their website, and a design study looking at increasing the metal flown in the first stage vs lost payload fraction. The idea is that extra metal is stronger and more likely to survive ocean parachute impact for reusability.
You can do such things either by a real trajectory code, or by "jigger factors" and the simple rocket equation. I don't know of anything in-between that is reliable. I used a trajectory code at LTV Aerospace 40+ years ago doing work on the Scout launcher. I used the rocket equation and "jigger factors" to set up those studies. Really, one uses both tools.
I've been using a nominal 26,000 feet/sec (7.9 km/sec) for LEO orbital velocity. Sorry about the imperial units, just divide by 39.37/12 and by 1000. There's 2.205 pounds in a kg, BTW. The mass ratios and ideal velocity increments I got for Falcon-9 are good to under a percent, limited by the approximation of multiplying Isp by the standard gravity constant to estimate exhaust velocity. Not technically correct, but pretty close.
My jigger factors are experiential guesses. That's the 1.10 and 1.05 knockdown factors for actual; delivered first and second stage velocity increments. These are the penalties for air and gravity drag, lumped. Doing it this way got me quite close into the ballpark: 8.1 km/sec vs 7.9 required. Not bad for a guess. That was my first post on Falcon-9. Close enough to run a design study realistic enough to trust.
The second study looked at the effects on payload fraction overall of adding more metal to the first stage to make it stronger and thus more survivable and reusable. By the time I tripled inert weight, I'd pretty much cut payload by 3. Since launch costs are more proportional to launch weight than anything to do with payload, this means the reusability has to save factor 3 over non-reusable costs, just to keep the same cost per payload mass, at triple the first stage inert weight. Whether this can be achieved, I dunno, but I doubt it.
At triple the inert weight, the inert fraction starts to look like the solid SRM's of the shuttle, which were reusable, sort-of. Not always. At that level, you become pressure-vessel capable, and could do pressure-feed. History says that is uneconomic for throwaway vehicles, except for some tactical missiles. But it might be economic if you are driven to that much metal by other considerations, such as survivability / reusability.
Reusability is one tough nut to crack. Reentry above M10-12 is really tough, and ocean parachute impact, thermal shock, and corrosion are all far tougher than anyone wants to believe until they've actually tried it.
Spacex has the lowest payload costs in the industry without reusability so far. That's because they have a vehicle that is reliable with a village supporting it, not a major city like all the other contractors or NASA. I think that's the real way to go. If reusability can be achieved too, well then that's gravy. But it's not the driving factor.
I'm still looking at 3 stages vs 2, nothing to conclude yet. But I bet it's logistics, not reusability, that can be achieved with rockets like this. Airplanes, well, that's quite different. Pods assisting rockets that stage off well below M5, that's different, too. But I think we're already seeing the most effective answer with plain chemical rockets.
GW
And, no, I'm not asking anyone to eat a hat. Funny as that might be.
GW
I could do the two-stage launcher with tougher first stage for reusability pretty easy. Remember, these plot as a smooth curve, but thrust levels, engine size and count, common engine considerations, and vehicle acceleration levels will constrain you to specific points, not anywhere along the curve. This is based on Spacex Falcon-9, again.
What I calculated was same launch weight, same 1st stage mass ratio (for same delta-vee), same interstage ring weight (same size 1st stage), reduce the 2nd stage ignition weight plus shroud weight by the added 1st stage inert weight, scale the shroud and 2nd stage weight statements by the factor of current 2nd stage ignition weight-plus-shroud weight to its original value (knowing some of this really doesn’t quite scale that way). Thus 2nd stage mass ratio and delta-vee are the same as original. Then recompute delivered payload weight to launch weight.
Configuration original “double” “triple”
Del ivered payload, lb 23050 18464 13879
Stage 2 dry weight, lb 6250 5007 3763
Stage 2 propellant, lb 118000 94525 71050
Stage 2 ignition, lb 147300 117996 88693
Stage 2 ign + shroud, lb 150800 120800 90800
Scale factor 1.0000 .801061008 .6021220159
Shroud weight, lb 3500 2804 2107
Interstage weight, lb 1200 1200 1200
Dropped at staging, lb 4700 4004 3307
Stage 1 “payload” (2-ig+dropped), lb 152000 122000 92000
Stage 1 dry weight, lb 30000 60000 90000
Stage 1 propellant, lb 553000 553000 553000
Stage 1 ign = launch, lb 735000 735000 735000
Stg 1 str inert fraction 4.082% 8.163% 12.245%
Overall payload fraction 3.136% 2.515% 1.888%
For comparison, I think the Shuttle boosters were somewhere near inert fraction 10%, and a lot of these segments were too dinged-up hitting the sea to reuse. So, an 8% inert fraction (configuration “double”) may well be far too fragile for effective reuse. The “triple” configuration at 12% inert may be more realistic. Or it may still be too fragile.
At any rate, once the tankage gets about as tough as a solid propellant motor, one might as well do pressure feed and eliminate the turbopump machinery. That could save some inerts, and perhaps increase reliability at very low logistical “tail”. It certainly needs to be considered as a viable option.
GW
I'd guess that the thermal injury from cold is a bigger risk that the sharp abrasive characteristics. Up to about 10 and down to about 3, even pH is no real risk. We eat stuff that far from neutral all the time.
That being said, I think there is a lot of merit to a mechanical counterpressure suit made of separate pieces, whose gloves can be doffed in vacuum for several minutes at a time in safety. There is no substitute for handling a sample with your bare fingers, right there on site. It's a big piece of how we were made to sense things, after all.
GW
Take a look at reusable rockets under interplanetary transportation. I posted some reverse-engineering results for the Falcon-9 that apply to this discussion as well as that one.
GW
Spacex’s Falcon-9 is a 2-stage rocket with kerosene-oxygen engines in both stages. It features an interstage ring and a payload shroud (on the satellite version) that I assume both get jettisoned at staging. The same engines are used in both stages, except that the one in the second stage has a longer bell than the nine in the first stage. I looked up most of the basic data from Spacex’s website, and reverse-engineered the rest. Here it is:
Delivered LEO payload 23,050 lb 15.648% of stage 2 ignition
Second stage dry weight 6,250 lb 4.243% of stage 2 ignition (structural inert)
Second stage propellants 118,000 lb 80.109% of stage 2 ignition
Stage 2 ignition weight 147,300 lb 100% of stage 2 ignition
Payload shroud 3,500 lb
Interstage ring 1,200 lb
Dropped at staging 4,700 lb
Stage 2 ign + dropped-at-staging 152,000 lb 20.680% of stage 1 ignition (“payload”)
First stage dry weight 30,000 lb 4.082% of stage 1 ignition (structural inert)
First stage propellants 553,000 lb 75.238% of stage 1 ignition
Stage 1 ignition weight 735,000 lb 100% of stage 1 ignition
Nine first stage kerolox engines (135 sec min burn time) average Isp is average of sea level 275 sec and vacuum 304 sec: 289.5 sec, for estimated Vex = 9314.4 ft/sec.
One second stage kerolox engine (260 sec min burn time) average Isp is vacuum Isp: 304 sec. Estimated Vex = 9780.9 ft/sec.
Use gravity & drag loss factors of 1.1 first stage (near-vertical with air drag), 1.05 second stage (more horizontal and essentially drag-free).
Stage mass ratio ideal delta-V actual delta-V
1 4.03846 13,002 ft/sec 11,820 ft/sec
2 5.02730 15,795 ft/sec 15,043 ft/sec
Overall ---- ---------- 26,863 ft/sec = 8.187 km/sec
The orbital velocity requirement for LEO is commonly said to be 8.1 km/sec. My reverse-engineering results for the Falcon-9 come remarkably close to that value at its rated satellite payload to LEO. That payload is 3.136% of launch weight. And staging does indeed occur way outside the sensible atmosphere: somewhere in the vicinity of 100-150 miles up.
They are having no success yet, re-using any of the first stage tankage or engines, after recovery at sea. There is no attempt to recover and reuse the second stage, the payload shroud, or the interstage ring. Given the 4.082% structural inert mass fraction for that stage, I am unsurprised by that outcome. I doubt they ever will reuse much first stage hardware, as long as it remains that fragile. And as long as it remains that lightweight, it will remain that fragile.
I will turn this same analysis around and see what payload reduction results in a two-stage design at higher first stage inerts, and if I can fly the same payload fraction 3-stage at higher inerts. Haven't done it yet. But when I have, y'all will have a sort of "baseline" on which to hang all these arguments.
GW
Odd I have to log in twice3 to stay logged in. Hmmm.
Cold fusion, or whatever you want to call it, may or may not be "real", in the sense that it may or may not work, regardless of what current theoretical thinking is. Being an engineer, I'm a lot less preoccupied about physics theories, and a whole lot more concerned with results. If this guy's E-cat machines really do generate power without a scam or cheat somewhere, then I say use them, and let's go get some new theories.
I guess you could say I'm skeptically open-minded. The skepticism is proportional to the lifetime of the theory being "violated". Nuclear reaction stuff only dates to the 1930-1940's. A lot of it could still be wrong or incomplete.
On the other hand, conservation of energy / first law of thermodynamics goes back more than 3 centuries without a single violation (for proper system definitions). When people come to me with perpetual motion machines based on weights or magnets, I am very, very skeptical.
I don't see E-cat as violating the first law yet. It's either a scam, or the energy release is coming from something outside prior well-known physics. That second possibility is not something we can rule out. Yet. Being outside an existing theory is not a disqualifier: see Ptolemaic vs Copernican astronomy (both made usable predictions), see also Newtonian mechanics vs relativistic mechanics (both make the same predictions at low speeds and Newton is easier to use).
The original Pons and Fleischman cold fusion incident was not an intended cheat. It turned out to be irreproducible by others. Pons and Fleischman sincerely believed they were on to something.
Why was it irreprodicible? Well, either they were wrong, or else what was happening in their experiment had more parameters than they knew to control for, which means they could not communicate what they didn't know to others. Even today, that second possibility cannot be ruled out, because there are no theories for this sort of thing.
So, until I see what happens with E-cat machines sold for use (do they generate power or not?), I think the jury is out. Polywell fusion is another one on which the jury is still out. The US Navy recognizes it to be a "long shot", but worthy of trials to find out. I'm not holding my breath either, but I hope the polywell guys succeed.
Similarly, I do hope the E-cat guy is successful. We need some breakthroughs. But it's only hope, and a willingness to investigate.
You have to investigate the new ideas that violate preconceptions, which does not mean you bet the farm on them being right. You just have to go find out whether they're right. History says the best advances come from that process. It's like evolution: fits and starts punctuating long periods of stagnation. Science and technology doesn't change very much at all by incremental improvements. Never has.
GW
Well, the "suit" back in 1969 was 6 or 7 layers of porous panty-hose material over torso and limbs. It does not have to be one piece. There was a breathing bag for tidal volume, a restraint vest over that, and a helmet tied to the suit. The gloves and booties were 2 or 3 layers of that same kind of stuff. This stuff got put on the web in Paul Webb's site about the "elastic spacesuit". My dad thinks he knew Webb in the mid 1950's when he (my dad) was designing the F-8 Crusader. Webb was an aerospace medicine type who worked in high altitude crew escape for decades. Somebody at MIT is working mechanical compression suits, but to a 0.33 atmosphere specification from NASA. Today's materials cannot quite meet that.
Other stuff I read said the edema / swelling that hits unprotected body parts takes around 15-20 minutes to occur. That means you could doff the compression gloves and work barehanded in vacuum for maybe 10 minutes at a time, before you have to re-don the compression gloves. The only real worry would be thermal injuries from touching very hot or cold objects. Harder to do with a gas balloon suit, but not impossible. The necessary wrist sealing was on the gas balloon pressure suit Joseph Kittinger wore for his balloon jump in 1960. His right glove failed to pressurize, but he flew the mission anyway. That was several hours exposure to vacuum. It swelled up and was painful and useless, but returned to normal within hours of landing.
GW
For crude estimates, you can factor up the orbital velocity by about 10-15%, as the effective delta-vee required to overcome gravity and drag with a clean, fast-ascent vertical launch vehicle. Good enough for 2, 3, even 4-stage vehicles to zeroth order. To do any better than that requires real trajectory code calculations.
GW
I once built several 2-D trajectory codes out of the concept below, when I was young. One of them was for vertical ballistic launch, the rest were quite different problems. It matched real ballistic (nonlifting) flight performance very well. I might even still have it, but none of these modern PC's and Windows versions now supports the advanced BASIC language that I wrote it in. It was meant for a DOS machine, not Windows. The last Windows variant that still could execute it in its DOS emulator was Windows 98. (I still have a working antique like that. Every time I power it up, I wonder if it will die this time.)
Here's the concept:
Along the flight path forward acts thrust, which you can set as a user-defined set of inputs. Also, drag in the reverse direction, which varies with speed squared and density, and as an empirical coefficient as a function of Mach (different for each stage). The flight path makes an angle to the horizontal, but starts out nearly vertical, usually only about half a degree off vertical. Weight acts vertically downward, but that is related to thrust by the massflow rate, and by the stage configuration. As you climb, gravity really isn't constant, either. (What I wrote was flat-Earth 2-D, but the same stuff can be programmed into spherical-Earth coordinates.)
You can use a simple forward-stepping estimate, if you choose the time step small enough. Long ago we had to be more sophisticated (such as Runge-Kutta), but with the advent of 486 PC's that was simply no longer necessary. Just calculate thrust, drag, and weight at the current time, and ignore their change over the short interval. Keep it really short, though, like under 0.1 second.
When you talk about "gravity loss", you are really talking about the effects of that non-constant vertical weight vector, integrated over time.
GW
Oh yes, X-15 was suborbital. But it did fly in space, it was rocket-powered, and each of the 3 airframes flew lots of times with very minimal ground support. The 40% inert fraction plus its record proves it was tougher than an old boot. (Actually, that service record is because it was tougher than an old boot.) So, that's what you want out of your BDB. 40% inert is no magic number, but I am certain the "right" number (for tougher than an old boot) is one whopping lot more than 8-10%.
There is a size effect for structures: weight (loads) scale as dimension cubed, while strengths scale only as dimension squared. Thus strength / weight ratio scales inversely with size, all other things the same. Gigantic vehicles end up in the same situation as water balloons supported by nails. An easy way to get around this is to build your gigantic stage as a cluster of much smaller tankage, and recover each of them separately. Not every tank needs its own engine, either.
The current optimal throwaway designs to LEO are two-stage rockets, with each stage around 10% inerts. The first stage delivers around 10,000 ft/sec delta-vee, more thrust-limited than Isp-limited, so kero-lox has been hard to beat for decades. The second stage operates not so vertically so it delivers the other 16-17,000 ft/sec, and is more Isp-limited than thrust-limited. Second stages are also physically smaller, so you can match or reduce diameter relative to the first stage, even with a super-low density propellant. That's why LH2-lox works so well.
Built tougher with higher inert fractions may require reverting back to three stages, in order that payload doesn't shrink to zero. That doesn't really make it much more complex (the 3-stage Minuteman was pretty simple, after all, and so was the 4-stage Scout), but by reducing the delta-vee from each stage, one can tolerate existing Isp's and thrust levels, at substantially higher inerts. Higher inerts mean tougher, which can mean smaller logistical tail (not guaranteed, it's more of a cultural change). That's how you make it cheaper, if you can really pull it off.
It's the last stage that worries me the most. I think it needs to burn three times (or even more), to achieve orbit along with the payload. That's not such a dumb, utterly-simple vehicle. That way, the payload can be completely inert dead-head cargo, and, you can ensure the last stage comes down exactly where you want it (it's the lower stage or stages that come down far downrange). I just don't yet have a clear picture of how best to survive reentry with a shape like that, and still be recoverable and reusable. Reentry is really tough. Mach 10-ish hypersonics are not all that bad. Mach 25? Really tough.
GW
Joshnh4h asked about his rocket trajectory. I haven't digested all of this. Old guy, slow. Also about limited trajectory distance over which ramjet adds energy: true enough for vertical launch, that's what worries me about the attractiveness of it. For HTO staged aircraft, the range spent in the air is quite long: that's why it looks so good.
Hop said something about a turbojet/ramjet aircraft. That looks pretty good, and could very well be integrated into the same engine. It would be one step past the J-58's that powered the old SR-71. Those had a 25% of airflow bypass after stage 3 compression, straight to the afterburner duct, meaning 75% of the air still went through the turbine core. Limited to about M3.8 max because of blading temperature problems at both ends of the spool.
On the other hand, if you rigged a 100% bypass capability that you could modulate 0-100, from the inlet straight to the afterburner duct, you could shut down the core turbine (to protect it, and just run the afterburner as a ramjet. Should be capable to M5 or 6.
Depending upon the arrangement, a ramjet pod or nacelle need not be complex or expensive, and can be tough as an old boot. With some sort of recovery built onto it, items like that should be more recoverable than the rocket stages we have so much trouble reusing. I'm thinking a stowed swing wing down the side, skids and a nose wheel, and R/C controlled glideback to a runway.
A ramjet airplane for HTO would be even easier. No integrated engines, just parallel-burn with separate rockets and ramjet. It's just stick-and-rudder flying. The ramjet is your fuselage. Your fuels and propellants go in the wing, the rockets go in the wing strakes, and the pilot goes in the inlet centerbody spike. The payload goes on the back.
GW
The drilling problem must be solved somehow. They did little hand cores a few centimeters deep standing on the moon. It has to be possible. It may be an art that needs to be learned that is peculiar to Mars, I dunno.
But deep drilling is the only reliable way to learn what is really underfoot, same as here on Earth. Without it you cannot truly explore: answer the two deceptively-simple questions (1) what all is there? (2) where exactly is it? To do less than answer those is but a negligible step beyond flag-and-footprints. What's the point of that?
We never really explored the moon, because we never really even tried to answer those two questions until recently.
GW Johnson
Mechanical counterpressure really is the way to go, since mobility is crucial for real exploration/investigation capability. The thing holding it back is materials not capable of 0.33 atmosphere mechanical squeeze. Unnecessary requirement, all that is needed is 0.2 to 0.25 atmosphere, and that was first done successfully back about 1969.
GW Johnson
Whatever basic rocket idea you use, re-usable or not, the real cost is the logistical tail behind it. NASA is famous for being expensive at $1B per shuttle launch (well, until recently, anyway). Why? It took the population of a major American city to support every launch, when you count all the contractors, subcontractors, and vendors. I bet $1B/launch didn't even cover the real costs - that's a lot of people to hire!
Now think "stick-and-rudder" into the black. Maybe a ground crew of under a dozen. You can throw the entire vehicle away and still be cheaper, if you can operate like that. But you cannot put all the bells and whistles on it. Not so extreme, but the same basic idea, is exactly why Spacex is so much cheaper than the majors. Nothing too hard to understand about that.
OK, now add real reusability, which means this thing (or "things" if multi-stage) have to take the abuses of spaceflight for years or even decades. You're going to have to build it tougher than an old boot to take that kind of abuse that long, and still keep the support crew small. There's simply no way around that dilemma. It also means you have to supply enough structure to take that kind of punishment: these 8-10% inert weight fractions I see bandied about are not even in the ballpark for a cheap system.
The most reusable, inexpensive rocket vehicle in all of history was the X-15. Its inert weight fraction (exclusive of its B-52 launcher) was 40%, which is not all that far from the typical supersonic bomber's 50-odd%, and not all that far from the B-52 itself. 3 X-15 airplanes flew 199 times over 2 decades, with only 1 complete airframe rebuild, and that was after destruction in an explosion in ground test. Nice record, for a manned rocket of 1955 design vintage.
GW Johnson