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Comparisons - expend./reusable,heavy/light,theory/practice
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 Posted on: Fri Jul 14, 2006 11:11 amHello, WannabeSpaceCadet,
this I didn't expect anymore because it menas that you are misunderstanding. The share is a criterion but not a ratio I want to maximize so. What I have in mind is to use the share or ratio as a criterion to distiguish families or kinds of vehicles. I am working on it a bit using the table I posted. I will have add some oxidator-propellant-combinations from another table under www.benrd-leitenberger.de that seem not to fit into the row of Isps but are essential regarding the share. But doing so - using the share to distiguish families - doesn't have nothing to do with maximization. I don't have in mind maximizations or the like. The goal at present still is to establish comparability. The first measures to do that were to keep V constant and to make other variables constant too - the next measure will be to apply that share. These are measures to be applied at the currect state - might be that later some constants will be made variables again or that I will shift from the share to something else. I don't know but at present the share will be my criterion. The goal is not a ranking now. Dipl.-Volkswirt(bdvb) Augustin (Political Economist) |
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Moon Mission Member  
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Posted on: Fri Jul 14, 2006 1:11 pm WannabeSpaceCadet wrote: what we all want is much less expensive and much larger volume, access to space. The Payload Effciency (payload mass / total mass) is completely irrelevant to this goal. |
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Space Station Member  
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Posted on: Mon Jul 17, 2006 8:48 am campbelp2002 wrote: WannabeSpaceCadet wrote: what we all want is much less expensive and much larger volume, access to space. The Payload Effciency (payload mass / total mass) is completely irrelevant to this goal. Such a vehicle would have already been eliminated by the $/kg (or $/lb ) criteria. But is a 100 ton rocket that puts 5 tons in orbit for $100 million better than a 400 ton rocket that puts 5 tons in orbit for $20 million? I think not, but I suspect LockMart & Boeing might have another opinion. Incidentally, the fuel cost for 380 tons of LOX / Kero is on the order of only $200 thousand. How much for a light steel tank with a common bulkhead divider, a few valves, a rocket engine with a 100:1 T/W ratio, and a basic guidance package? |
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Space Station Member
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Posted on: Mon Jul 17, 2006 8:52 am Ok, Ekkehard, I now see your goal. A list of all the common designs with their results over different categories. Should be interesting.
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Moon Mission Member
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Posted on: Mon Jul 17, 2006 1:54 pm WannabeSpaceCadet wrote: is a 100 ton rocket that puts 5 tons in orbit for $100 million better than a 400 ton rocket that puts 5 tons in orbit for $20 million? |
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Posted on: Mon Jul 17, 2006 2:31 pm In between I grouped the propellants by their normal state - gas, liquid, solid.
Here are the groups: Code: Oxidator Propellant Isp gas/gas ==> pressurization and cryogenics required Oxygen Diborane 3374 Fluorine Ammonia 3500 Fluorine Diborane 3638 Oxygen Hydrogen 3830 Fluorine Hydrogen 4020 gas/liquid ==> propellant flows relatively freely abnd doesn't reqire cryogenics etc. Oxygen Ethanol 2740 Oxygen Cerosene 2945 Oxygen ADMH 3010 Oxygen Hydrazine 3070 Fluorine Cerosene 3139 FLOX (Fluorine and LOX) ADMH 3373 Fluorine Pentaborane 3530 FLOX (Fluorine and LOX) Cerosene 3550 Fluorine Hydrazine 3560 Fluorine ADMH 3569 gas/solid ==> propellant doesn't require cryogenics etc. but doesn't flow freely also Oxygen Lithium hydride 2569 Oxygen HTPB/Aluminum (2776) Fluorine Lithium hydride 3569 Oxygen Beryllium hydride 3628 Fluorine Beryllium hydride 3863 Fluorine Polyethylene (4050) Fluorine/Oxygen Lithium hydride (4335) Ozone Beryllium hydride (5248) liquid/gas ==> oxydator flows relatively freely but the propellant requires cryogenics etc. Hydrogenperoxide 95 % Hydrogen 3187 Hydrogenperoxide 95 % Diborane 3246 Dinitrogen tetroxide Hydrogen 3353 liquid/liquid ==> both flow relatively freely Nitric acid Cerosene 2630 Dinitrogen tetroxide Cerosene 2710 Nitric acid ADMH 2710 Hydrogenperoxide 95 % ADMH 2720 Hydrogenperoxide 95 % Cerosene 2745 Hydrogenperoxide 95 % Hydrazine 2760 Nitric acid Hydrazine 2775 Dinitrogen tetroxide ADMH 2804 Dinitrogen tetroxide Aerozine 50 2820 Dinitrogen tetroxide Hydrazine 2865 liquid/solid ==> the oxidator flows relatively freely but the propellant don't Dinitrogen tetroxide Lithium hydride 2441 Nitric acid Polyurethane (2490) Nitirc acid Polybutadiene (2540) Nitric acid Lithium hydride 2549 Hydrogenperoxide 95 % Lithium hydride 2559 Dinitrogen tetroxide Polyethylene 2696 Dinitrogen tetroxide Beryllium hydride 3432 Hydrogenperoxide 95 % Beryllium hydride 3667 unknown/gas ==> can't say nothing about oxidator but the propellant requires cryogenics etc. Difluoroxide Diborane 3569 Difluoroxide Hydrogen 4020 unknown/liquid ==> can't say nothing about oxidator but the propellant flows relatively freely Difluoroxide Cerosene 3196 Difluoroxide Hydrazine 3383 Difluoroxide ADMH 3442 unknown/solid ==> can't say nothing about oxidator but the propellant doesn't flow Difluoroxide Beryllium hydride 3746 I left away "Chlordifluoride Cerosene 2530" because the oxidator is liquid up to 21° C but gas above that temperature. These groups are of meaning for one of the constants. May be that there is a difference in how freely the different liquids flow but this shouldn't be the focus on at the moment. Dipl.-Volkswirt (bdvb) Augustin (Political Economist) EDIT: I added the additional solid propellants I forgot yesterday and set the Isps into parenthesises because I am suspecting that they are valid for other burning cahmber pressurs, mix ratios or external pressures than those of al the other oxydator-propellant-combinations. But my focus here is on the propellants and oxydators themselves - not on the Isps that much - at present. |
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Space Station Member
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Posted on: Thu Jul 20, 2006 7:27 am campbelp2002 wrote: You will never convince me that a rocket 4 times bigger can cost 80% less no matter how simple it is. I don't see why not, if the cost of the propellants, raw materials, electronics, valves, wires, nut & bolts etc, is all a small fraction of the total cost, then complexity, & the search for the best possible payload fraction, must be the cost drivers. NASA's Minumum Cost Design (MCD) studies of the 80's & 90's, proved this. But scrapping the Shuttle (and now Ares), proved politically impossible. |
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Moon Mission Member
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Posted on: Thu Jul 20, 2006 1:09 pm I might agree that simplifying a design could cut the cost in half. You might get me to agree to a quarter for a really elegant design. But 4 times bigger AND 5 times cheaper is effectively 20 times cheaper for a given size. No way Jose!
Pegasus was supposed to be cheap, because it is all solid propellant and small. What could be simpler? Just a big model rocket; a steel tube filled with propellant. No pumps, valves, pipes, pressurizing gas, regenerative chamber cooling channels, slosh baffles, nothing. But it costs $30 million per launch, far more expensive than Falcon 1 which has a pump fed liquid fuel engine and can put a slightly larger payload into orbit. |
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Space Station Member
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Posted on: Fri Jul 21, 2006 5:14 am Pegasus is small, but not simple. 3 stages, a wing, a converted airliner, (4 stages really). Solid rocket fuel (and hybrids) cost a lot more than most liquids, and require a lot of care and inspection in casting.
Similarly, Rutan's Spaceship-One was the quickest way to sub-orbital for an aircraft company sponsored by a billionaire. But it was nowhere near the cheapest at $200,000+ per flight. An upscaled version of Armadillo's VDR would have costs of around 5 to 10% of that. T/Space believe they can put a manned capsule, weighing almost 5 tons, in orbit, for $20m a flight, and they are still using a carrier aircraft and 2 stages. A non man-rated, expendable BDR could be cheaper than that. |
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Posted on: Mon Jul 24, 2006 11:42 am The groups of my recent table can partially can be aggregated to form the follwoing supergroups:
gas/gas gas/liquid + liquid/gas gas/solid liquid/liquid liquid/solid unknown/gas unknown/liquid unknown/solid Difluoroxide/Beryllium hydride Leaving away the last four I at this point suspect - methodologically - the first group to require the most heavy equipment because for both the oxidator and the propellant crygenics and pressurization are required. The second requires less equipment because only one of the both requires cryogenics and pressurization. The third group I expect additionally to not require equipment liquids need like pumps and valves. The next two now don't need no cryogenics and pressurizations while the last of them may be the group requiring the least amount of equipment. So these groups look as if they are groups of different kinds of rockets or vehicles because Klr and Klt are different in comparison to each other. If the total weight of a rocket is 100% then the percentage of Klr and Klt at this point can be expected to differ between the groups more than within each group. As a consequence it can be expected that the percentage of payload-weight too differs in that manner. Instead of 100% I could have used an absolute value - 100 mT for example. Then it can be seen directly that the expectation is that the payload-weight differs as described - so this may be a good criterion to distinguish 5 lifter-groups where according to the expectations the fifth group may be best fitted to heavy weights while the first may be best fitted to light weights but shouldn't be used for heavy weights. These expections have to be checked later. Before that s going to be done another comparison is required also ... Dipl.-Volkswirt (bdvb) Augustin (POlitical Economist) |
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Posted on: Mon Jul 31, 2006 11:18 am In between I have ready a table providing a synopsis of the groups distinguished - but it doesn't fit into the post in parallel to keeping the table readable and understandable yet.
But what can be seen is that 1. the gas/gas-group, 75% of the gas/solid-group, four members of the gas/liquid-/liquid/gas-group and 25% of the liquid/solid-group have comparable Isps, 2. next there are eight members of the gas/liquid-/liquid/gas-group followed by 3. the liquid/liquid-group and at last 4. five members of the liquid/solid-group. It seems that gases aren't required really and that non-gas-propellants allow for higher payload-shares of the total weight. But this depends on the distinct Isps merely. The issue might still hold because some Isps of solids are even higher than gas/gas. I am going to find a modification of the table that fits into the requirements here - I will post it either here by EDIT or by a separate post. Dipl.-Volkswirt (bdvb) Augustin (Political Economist) EDIT: Here the table I altered the names of the oxydators and propellants into short names first: Translation Names into Short Names Code: Oxydator Propellant Ox Prop Dinitrogen tetroxide Lithium hydride DT Li hy Nitric acid Polyurethane NA Pol u Nitric acid Polybutadiene NA Pol b Nitric acid Lithium hydride NA Li hy Hydrogenperoxide 95 % Lithium hydride HP Li hy Oxygen Lithium hydride LOX Li hy Nitric acid Cerosene NA Cero Dinitrogen tetroxide Polyethylene DT Pol e Dinitrogen tetroxide Cerosene DT Cero Nitric acid ADMH NA ADMH Hydrogenperoxide 95 % ADMH HP ADMH Oxygen Ethanol LOX Ethl Hydrogenperoxide 95 % Cerosene HP Cero Hydrogenperoxide 95 % Hydrazine HP Hzine Nitric acid Hydrazine NA Hzine Oxygen HTPB LOX HTPB Dinitrogen tetroxide ADMH DT ADMH Dinitrogen tetroxide Aerozine 50 DT Ae 50 Dinitrogen tetroxide Hydrazine DT Hzine Oxygen Cerosene LOX Cero Oxygen ADMH LOX ADMH Oxygen Hydrazine LOX Hzine Fluorine Cerosene F Cero Hydrogenperoxide 95 % Hydrogen HP LH2 Hydrogenperoxide 95 % Diborane HP Dibo Dinitrogen tetroxide Hydrogen DT LH2 FLOX (Fluorine and LOX) ADMH FLOX ADMH Oxygen Diborane LOX Dibo Dinitrogen tetroxide Beryllium hydride DT Be hy Fluorine Ammonia F Ammo Fluorine Pentaborane F Penta FLOX (Fluorine and LOX) Cerosene FLOX Cero Fluorine Hydrazine F Hzine Fluorine ADMH; F ADMH gas/solid-case: g/s: Lithium hydride Li hy Oxygen Beryllium hydride LOX Be hy Fluorine Diborane F Dibo Hydrogenperoxide 95 % Beryllium hydride HP Be hy Oxygen Hydrogen LOX LH2 Fluorine Beryllium hydride F Be hy Fluorine Hydrogen F LH2 Fluorine Polyethylene F Pol e Fluorine/Oxygen Lithium hydride F/LOX Li hy Ozone Beryllium hydride Ozone Be hy The table I used to write this post is Comparison of Isps using Short Names Code: hint Ox Prop. gas/ gas/ gas/ liq/ liq/ gas liq soli liq soli DT Li hy 2441 NA Pol u 2490 NA Pol b 2540 NA Li hy 2549 HP Li hy 2559 LOX Li hy 2569 NA Cero 2630 DT Pol e 2696 DT Cero 2710 NA ADMH 2710 HP ADMH 2720 gas/liquid LOX Ethl 2740 HP Cero 2745 HP Hzine 2760 NA Hzine 2775 LOX HTPB 2776 DT ADMH 2804 DT Ae 50 2820 DT Hzine 2865 gas/liquid LOX Cero 2945 gas/liquid LOX ADMH 3010 gas/liquid LOX Hzine 3070 gas/liquid F Cero 3139 liquid/gas HP LH2 3187 liquid/gas HP Dibo 3246 liquid/gas DT LH2 3353 gas/liquid FLOX ADMH 3373 LOX Dibo 3374 DT Be hy 3432 F Ammo 3500 gas/liquid F Penta 3530 gas/liquid FLOX Cero 3550 gas/liquid F Hzine 3560 gas/liquid F ADMH; 3569 3569 or solid g/s: Li hy LOX Be hy 3628 F Dibo 3638 HP Be hy 3667 LOX LH2 3830 F Be hy 3863 F LH2 4020 F Pol e 4050 F/LOX Li hy 4335 Ozone Be hy 5248 Last edited by Ekkehard Augustin on Thu Aug 10, 2006 7:20 am, edited 2 times in total. |
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Posted on: Tue Aug 08, 2006 10:14 am What has been left out of consideration during looking for and at groups is the density of propellants.
Since the density determines the required tank volume it determines the surface of the tank(s) and thus the weight provided the tanks have in common the material they are made of. For comparison purposes I use a standard material - aluminum. The density of the propellant(s) don't establish new groups but simply mean to turn back from the constants to the variables. Dipl.-Volkswirt (bdvb) Augustin (Political Economist) |
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Posted on: Thu Aug 10, 2006 5:16 am Some propellants are not compatible with aluminium. Fluorine for example. They require non-reactive liners. Cryogenic propellants require insulation of the tank, making it heavier.
Pressurization of the propellant initially allows the tank wall to be thinner, then thicker as pressure increases. The trade off with higher pressure (heavier) tanks, is you don't need a (heavy) turbo pump, but ISP is a bit less. ISP varies, for any specific propellant combination, with the chamber pressure, whether produced by tank pressure (medium) or a turbo pump (high). The surface area of the tanks depends on the the shape of the tank as well as the volume. A sphere has the least surface area for a given volume. Spherical tanks also require thinner walls to contain the same pressure. |
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Posted on: Thu Aug 10, 2006 1:04 pm Hello, WannabeSpaceCadet,
under www.bernd-leitenberger.de there is an article explaining why spherical tanks aren't that advantageous but have more disadvantages than advatages. But although spherical tanks are optimal regrading surface and thus material larger amounts of a particular propellant require a larger sphere too which menas more surface and thus more material. The variation of the Isps by chamber pressure I remember to have mentioned when posting the propellant table the first time some posts ago on the previous page of this thread. The table(s) of propellants and Isp I am using up to now explicitly use a constant chamber pressure. At present I am not going to vary that pressure - variations simply would shift the values only I suppose. I might have a look on the thickness or thinness of the tank walls later when I am ready with the propellants and thus the connection between ve and Mve for the first time. At this point I only want to hint to the circumstance that the weight of the tanks is a constant Klt here. In between there are five groups of Klt here - but Klt will be kept a constant but not turned into a variable Mlt again. This all is so simply because the focus is on the connection between ve and Mve at present - which must be kept yet. Dipl.-Volkswirt (bdvb) Augustin (Political Economist) EDIT: Another hint regarding Fluorine and aluminum. At present I am still focussed on theory. This means that I can neglect such incompatibilities - they simply have to be handled later in practice. There might be solutions like thin protecting layers of non-aluminum materials. Each tank might have such layers and it might be possible to select them so that the comparability of the tank weight isn't removed. And LOX-tanks have insulations for wires inside because of the requirement to set the LOX to motion from time to time. What's of meaning at present simply is to apply tanks in these theoretical considerations that are made of the same material or of materials that weight nearly the same at nearly identical amounts. |
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Posted on: Tue Aug 15, 2006 11:24 am In the following table the Isps are replaced by average densities I got from www.bernd-leitenberger.de.
Code: hint Ox Prop. gas/ gas/ gas/ liq/ liq/ gas liq soli liq soli DT Li hy ? NA Pol u 1.391 NA Pol b 1.384 NA Li hy ? HP Li hy ? LOX Li hy ? NA Cero 1.35 DT Pol e ? DT Cero 1.26 NA ADMH 1.25 HP ADMH 1.24 gas/liquid LOX Ethl 1.1 HP Cero ? HP Hzine 1.26 NA Hzine 1.28 LOX HTPB 1.34 DT ADMH ? DT Ae 50 1.2 DT Hzine 1.22 gas/liquid LOX Cero 1.2 gas/liquid LOX ADMH ? gas/liquid LOX Hzine 1.7 gas/liquid F Cero ? liquid/gas HP LH2 ? liquid/gas HP Dibo ? liquid/gas DT LH2 ? gas/liquid FLOX ADMH ? LOX Dibo ? DT Be hy ? F Ammo 1.18 gas/liquid F Penta ? gas/liquid FLOX Cero ? gas/liquid F Hzine 1.31 gas/liquid F ADMH; ? ? or solid g/s: Li hy LOX Be hy ? F Dibo ? HP Be hy ? LOX LH2 0.28 F Be hy ? F LH2 0.45 F Pol e 1.31 F/LOX Li hy 1.327 Ozone Be hy ? The order of the propellants is left unchanged. It turns out that the high densities are concentrated at the propellants with low or medium Isps as far as I found densities listed. Four higher densities only are to be found in the high-Isp-region of the tables. The highest density is to be found in the gas/liquid-column while the only high-density/high-Isp combinations can be found in the gas/solid-column. From this it seems that high Isps are not necessaryly not-outweighed by larger tank-weights. I consciously say "From this" because I might get different results considering it from another perspective - here I still look at the functional links and connections between ve and Mve. Lighter tanks with less Isps - heavier tanks with higher Isps - four exceptions of that only: In general an increase of the ve will be partially consumed by an increase of the weight of the tank(s) and so on - except for 4 exceptions. The conclusion from this seems to tend to be that the slope of the function f(ve) = Mve is not negative. Please note that I am still considering the theory-aspect. The results are weak regarding the function f(ve) = Mve since a lot of average densities aren't available to me at present and they would have to be given under the conditions I listed in an earlier post. Hello, WannabeSpaceCadet, although I am not yet at the point to look at the tanks one hint here. Larger spherical tanks would have a larger diameter. So the rockets using larger spherical tanks would be wider and thus tend to suffer more resistance by the air in front of it/them. This in turn would increase the required amounts of propellant - which again would increase the tank diameter. The shape of rockets might be changed to solve such problems - but this too is or would be a problem to be considered later. ... Dipl.-Volkswirt (bdvb) Augustin (Political Economist) |
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