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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
I hope the E-cat guy is for real. Historically, our best, most dramatic advances took place when somebody upset the applecart of conventional wisdom. It's way past time this happened to the "fusion business".
GW Johnson
Actually, any practical system capable of a flight to an NEO is capable of going to Mars. You just add landers. It's not so much the delta-vee and flight times, it;'s actually crew survival that drives what you do. If you fly weightless, you have to fly fast: one year max is demo'd on the various space stations over the last few decades, and we do not know that Mar's 0.38 gee is enough to be therapeutic. Otherwise, go 9 months one-way, several months there, and about 9 months home, and provide just about 1 gee by spinning the ship end-over-end. There's no way anyone will stay sane cooped up in any capsule; we'll need a Skylab-like module to live in. The bigger, the better. Think orbital rendezvous and assembly here, and orbital rendezvous (like the Apollo missions) at Mars. You'd better start thinking nuclear thermal rockets, maybe even gas core. You might also start thinking about the real point of going there: is it flag-and-footprints like Apollo, or is it real exploration to find out what exactly is there, and where exactly is it?
GW Johnson
Hi gang:
This is GW Johnson the old aero engineer, and ramjet expert from long ago. I surely am glad to see the forums up and running again.
In recent news: I have picked up a consulting client for a possible ramjet launch effort. And, that client and I both think I may be just about the last living US all-around expert in ram propulsion (I seem to have outlived the rest).
I’m particularly glad to see LEO access under active discussion, especially with Josh and Hop talking about reusable vehicles, and perhaps ramjet assist. The last stuff I had is a posting over at http://exrocketman.blogspot.com, where I looked at horizontal takeoff and landing with a winged first stage using separate rocket and ramjet power. That article is dated August 22, 2010, so it’s way down the list (chronological, latest on top). There’s a navigation tool by date and title on the left, under my photo. It looked to me like a staging condition of near Mach 6 at around 60,000 feet altitude might well work out, including booster flyback. And, it looks like ramjet might really pay off in this scenario.
Josh is exactly right: the frontal thrust density of ramjet is too low to support vertical acceleration of a heavily-loaded vertical launch vehicle on ramjet alone. But parallel-burn of otherwise-separate rocket and ramjet might offload some of the thrust requirement temporarily onto the higher-Isp ramjet, thus swapping a smidge of rocket propellant for a smidge of extra payload. That kind of vehicle is moving only around M2 at 60,000 feet, so it’s a “low-speed” pitot inlet design, not the “high-speed” spike inlet that makes sense for the ramjet launch airplane. Whether this idea is actually technically and economically attractive in vertical launch, remains to be seen. I just dunno, yet.
You can think of “high-speed” systems being some sort of spike or ramp inlet, a dump combustor, and a very mild-expansion convergent-divergent nozzle. Min Mach number is around 1.5, and max is around M5 to 6, depending more on vehicle drag than the ramjet design itself. I looked at nose inlets, but side inlets also work. Peak performance is around M2 to 3, and frontal thrust density falls too low to provide effective acceleration above about 60,000 feet. It doesn’t matter a lot whether you analyze RP-1 or JP-5 kerosene, or even RJ-5/Shelldyne-H synthetic, they all come out similar in proportions and performance. I think liquid methane would look very similar to kerosene, too.
“Low speed” systems are a simple pitot (normal-shock) inlet, most likely a nose inlet, a dump combustor, and a convergent-only nozzle. Min Mach is a tad fuzzy, there being thrust greater than drag (in low-drag nacelles) down under Mach 0.5, although Isp is over kero-lox rocket levels only above around Mach 0.7-ish. Max Mach is around 2 to 3, depending more on vehicle drag than the ramjet, with max performance around M1.5 to 2. I think thrust densities fall too low to be useful for acceleration above around 60,000 feet, although this remains to be seen for sure. Again, specific fuel choice is not all that important.
Whatever I do come up with, I’ll let y’all know. This stuff is fun. I haven’t done any of it in almost 2 decades, now. I think I’ll turn my old how-to notebooks into a published book. Otherwise, the art will die with me. (It’s still mostly art, until I can get it all written down). That ain’t easy.
GW