Orbiter-Forum Some proposals for low cost heavy lift launchers.
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 Math & Physics Mathematical and physical problems of space flight and astronomy.

 01-03-2011, 11:51 PM #2 Urwumpe Certain Super User Correction - the practical limit for high-performance vacuum optimized kerolox rocket engines is fixed at 350 s (~3400 m/s), about 100 m/s lower. The problem is not, that you can drive the chamber pressure higher with the energy available in the fuel and still gain a higher specific impulse. By only looking at what happens in the power head and by assuming simple gas dynamics relations between chamber pressure and specific impulse, you are right, 3500 m/s is really the logical result by applying the theoretical maximum chamber pressure that can be reached without higher tank head (pressure at the outlet of the propellant tanks) on known engine data. The problem is the needed expansion ration for getting the optimum, and the resulting nozzle mass, which degrades rocket stage performance. Unless you have a very heavy upper stage, that has a very long burn time, the larger nozzle will eat the performance. Many such engines will increase the negative effect. Just on the paper: 100 m/s more specific impulse mean about twice the chamber pressure, so the combustion chamber and thrust can be only half as large for the same thrust, but the nozzle has to be a bit more than 2 times longer than the original and still reach almost the same exit diameter (so you get the same thrust). Including all weight reductions by the smaller power head, your engine will weight over 5 times more than a less powerful variant, only for getting 100 m/s more specific impulse. You can of course also reduce the nozzle size and let it be underexpanded, but then, instead of reaching the theoretical gain of 100 m/s, maybe (nozzle aerodynamics are more complex than fighter jet aerodynamics) 20 m/s will eventually leave the engine. That combined with an explosion in production costs for an expendable engines. Things will be even worse for an SSTO, because you enter hells kitchen for the many trade-offs and optimizations and experimental measurements for getting essentially an RD-170 (which is about the maximum you can sequence out of kerosene and LOX, there is not much left to be gained at the throat of the engines) to be at optimal expansion for a wide range of pressures. For lift-off it is great to have a high chamber pressure, because you gain by it in any case. In the vacuum, high chamber pressures are suddenly annoying: As tempting and good as they are, in practical engineering, there can be a too good rocket engine. If you think this is now nitpicking...well, I was just counter-nitpicking at you. You pick a few numbers of a friendly favorable calculation (Dunn is a BDB apostle, his numbers have always been too good to be true.), so your big dumb booster becomes good - as long as nobody looks to closely at the numbers. Is a bit like the past in Europe when you wanted to buy an airline ticket. You can get from Berlin to London for 45 €. Plus fees, taxes, fuel costs and the price of an attractive competent stewardess. Don't just look too closely at the price. Also... putting a high performance engine on something that was planned to become an BDB is some sort of ironic, since this cost factor should have been killed by BDB... now it is embraced as necessary evil for reuse. When the launch ground infrastructure comes into place, the calculations will get along "Well, maybe we can do things a bit smaller now, and save a lot of money here" - the evolution of rocket feasibility studies from 1950 to today, repeated on the paper. As current example here, about how different real rockets are to blue print or Colliers Magazine rockets: Just look at Sea Dragon - great idea, the ocean costs nothing and it needs no VAB, since all assembly and a lot of welding will happen on the water in a small protected lagoon. And then you read how many buoys, barges, ships and tugs it needs for each step of the operations. Ooops. Not to mention that the idea of larger combustion chambers for simplicity had been found to be opposing to real conditions in the 1960s, when the F1 engine was designed - larger chambers mean much more trouble for getting stable thrust, downsizing of thrust chambers by higher chamber pressure was also a relief of this problem, since you had been able to get the same thrust ranges with chamber and injector geometries that had already been investigated. SpaceX is also a poor example of simplicity. Their rocket is maybe marketed as simpler as other designs and in terms of engine technology, this is true, but practically, it is not different to LMM or Boeing products. The rocket needed less testing by having less powerful engines, but they needed more painful testing and paid a lot of teaching money for also reducing the amount of engineers on the ground for launching. The Falcon 9 is only as much cheaper as other designs, as it is also less powerful - and it is doubtful at the current development speed, that the Falcon 9 will get the needed number of flights before an successor is already thrown into the field... It is a first one, but not really a good one (The Falcon 9 is actually in one league as the good old Soyuz rocket - and guess which one has currently the better market position). Last edited by Urwumpe; 01-03-2011 at 11:53 PM.
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 01-04-2011, 12:15 AM #3 T.Neo SA 2010 Soccermaniac Quote: Just look at Sea Dragon - great idea, the ocean costs nothing and it needs no VAB, since all assembly and a lot of welding will happen on the water in a small protected lagoon. And then you read how many buoys, barges, ships and tugs it needs for each step of the operations. Ooops. Indeed. I looked into Sea Dragon (and sea launch in general) a while back, and you really struggle with things that you don't get with a land launch, that outweigh the gains of sea launch/assembly. Putting together such a large vehicle would be very difficult regardless, even moderately sized launchers are no easy feat to assemble.
 01-04-2011, 06:48 AM #4 RGClark Mathematician Quote: Originally Posted by T.Neo  Indeed. I looked into Sea Dragon (and sea launch in general) a while back, and you really struggle with things that you don't get with a land launch, that outweigh the gains of sea launch/assembly. Putting together such a large vehicle would be very difficult regardless, even moderately sized launchers are no easy feat to assemble. You can get really large payloads with the 8.4 meter wide super "Evolved Atlas" stage by using parallel, "trimese", staging with cross-feed fueling. This would use now three copies of the lower stages mated together in parallel with the fueling for all the engines coming sequentially from only a single stage, and with that stage being jettisoned when its fuel is expended. Again we'll calculate first the case where we use the standard performance parameters of the RD-180, i.e., without altitude compensation methods. I'll use the average Isp of 329 s given in the Kyle article for the first leg of the trip, and for the required delta-V, again the 8,900 m/s often given for kerosene fueled vehicles when you take into account the reduction of the gravity drag using dense propellants. Estimate the payload as 200 mT. Then the delta-V for the first leg with all three super Evolved Atlas's attached will be 329*9.8ln(1+1,323/(3*70 + 2*1,323 + 200)) = 1,160 m/s. For the second leg we'll use the vacuum Isp of 338 s, then the delta-V will be 338*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 200)) = 1,940 m/s. And for the final leg 338*9.8ln(1 + 1,323/(70 +200)) = 5,880 m/s. So the total delta-V is 8,980 m/s, sufficient for orbit with the 200,000 kg payload. Now let's estimate it assuming we can use altitude compensation methods. We'll use performance numbers given in this report: Alternate Propellants for SSTO Launchers. Dr. Bruce Dunn Adapted from a Presentation at: Space Access 96 Phoenix Arizona April 25 - 27, 1996 http://www.dunnspace.com/alternate_ssto_propellants.htm In table 2 is given the estimated average Isp for a high performance kerolox engine with altitude compensation as 338.3 s. We'll take the vacuum Isp as that reached by high performance vacuum optimized kerolox engines as 360 s. Estimate the payload now as 250 metric tons. Then the delta-V during the first leg will be 338.3*9.8ln(1+1,323/(3*70 + 2*1,323 + 250)) = 1,180 m/s. For the second leg the delta-V will be 360*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 250)) = 2,020 m/s. For the third leg the delta-V will 360*9.8ln(1 + 1,323/(70 + 250)) = 5,770 m/s. So the total will be 8,970 m/s, sufficient for orbit with the 250,000 kg payload. Now we'll estimate the payload using the higher energy methylacetylene. The average Isp is given as 352 s in Dunn's report. The theoretical vacuum Isp is given as 391 s. High performance engines can get quite close to the theoretical value, at 97% and above. So we'll take the vacuum Isp as 380 s. Estimate the payload now as 300 mT. The first leg delta-V will now be 352*9.8ln(1 + 1,323/(3*70 + 2*1,323 +300)) =1,210 m/s. For the second leg 380*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 300)) = 2,080 m/s. For the third leg 380*9.8ln(1 + 1,323/(70 + 300)) = 5,660 m/s. So the total is 8,950 m/s, sufficient for orbit with the 300,000 kg payload. This trimese version of the vehicle would be huge however. For instance it would weigh more than the Saturn V. One of the big cost factors for the development of some of the super heavy lift launchers is that they are so heavy they would require the construction of new and expensive launch platforms. Undoubtedly, the bimese version would be the one to be built first if this launch system is selected. Bob Clark ---------- Post added at 06:48 AM ---------- Previous post was at 06:35 AM ---------- Quote: Originally Posted by Urwumpe  Correction - the practical limit for high-performance vacuum optimized kerolox rocket engines is fixed at 350 s (~3400 m/s), about 100 m/s lower. The problem is not, that you can drive the chamber pressure higher with the energy available in the fuel and still gain a higher specific impulse. By only looking at what happens in the power head and by assuming simple gas dynamics relations between chamber pressure and specific impulse, you are right, 3500 m/s is really the logical result by applying the theoretical maximum chamber pressure that can be reached without higher tank head (pressure at the outlet of the propellant tanks) on known engine data. The problem is the needed expansion ration for getting the optimum, and the resulting nozzle mass, which degrades rocket stage performance. Unless you have a very heavy upper stage, that has a very long burn time, the larger nozzle will eat the performance. Many such engines will increase the negative effect. ... You may be right in regards to using a very high expansion ratio in order to get those high vacuum Isp's while using a standard nozzle. However, an aerospike nozzle may be able to work. Also, note that the first calculation did not require altitude compensation or the very high vacuum Isp's. It just used the standard performance parameters of the RD-180 to get a 115 metric ton payload. Bob Clark
 01-04-2011, 07:03 AM #5 Urwumpe Certain Super User Quote: Originally Posted by RGClark   You may be right in regards to using a very high expansion ratio in order to get those high vacuum Isp's while using a standard nozzle. However, an aerospike nozzle may be able to work. No change then - the mass is then inside the (truncated) spike.
 05-02-2011, 11:40 PM #8 RGClark Mathematician NASA appears to be leaning to a 70 mt payload shuttle-derived launcher as an interim solution to developing a heavy lift vehicle. This would use two 4-segment SRM's as does the shuttle and an ET. But it would not have a shuttle orbiter, nor would this Phase I vehicle have an upper stage: SLS planning focuses on dual phase approach opening with SD HLV. April 25th, 2011 by Chris Bergin http://www.nasaspaceflight.com/2011/...pening-sd-hlv/ However, built into this plan is that at most 4 flights of this vehicle will be made before it is discontinued in favor of a more expensive, 130 mt payload upgrade. These 4 flights are to regarded as "test flights" according to the Bergin article. They will use 3 SSME's at a time and only 12 of those will be available including those taken from the retired space shuttles, thus allowing only 4 flights. Presumably after that either the production of new SSME's will be started or their expendable versions will be, or NASA will choose instead to use kerosene fueled engines for the core stage. However, a better plan in my view would be to explore methods in which this Phase I vehicle could be reusable. Then this low cost HLV could have many more missions as well as cutting costs in being reusable. This would give you more options as to when and if the more expensive vehicle needed to be developed. Many at NASA are not favorably inclined towards reusable systems because of the experience of the shuttle. However, as I mentioned before in the post #6 above, a key reason for why the shuttle was not economical would not hold in this case: it would not have to carry the 80 mt orbiter that took out most of the vehicles payload capacity. Another reason why a reusable vehicle could be done better now is because of the research that has already been done to address the failings of the shuttle system. For instance, for the X-33/VentureStar program the metallic shingles to be used for thermal protection have confirmed in testing they would require less maintenance than the ceramic tiles of the shuttle. The advanced ceramics used on the Air Force's X-37B were also expected to cut maintenance on thermal protection. It would be useful to find out if they have been successful in that regard. The X-37B may also serve as a good model to use for the reentry system for the ET tank to be used on the Phase I vehicle. Note that the X-37B's short stubby wings are much smaller in proportion to the size of the vehicle than those of the space shuttle. That and the composite materials used for the wings would result in much reduced mass used for the wings for the ET tank. Other lightweight reentry systems would be the German IXV program which does not use wings: IXV Program Aims to Put ESA at Cutting Edge of Re-entry Technology Posted by Doug Messier on September 18, 2010, at 4:08 am in ESA. http://www.parabolicarc.com/2010/09/...ry-technology/ and the inflatable one NASA is investigating: NASA Successfully Tests Vacuum-Packed Inflatable Heat Shield. A vacuum-packed inflatable shroud could enable future spacecraft reentry on both Earth and Mars. By Jeremy HsuPosted 08.17.2009 at 3:00 pm http://www.popsci.com/military-aviat...ld-flight-test Bob Clark ---------- Post added at 11:40 PM ---------- Previous post was at 01:56 PM ---------- The above discusses that maintenance costs for thermal protection should be significantly less than for the space shuttle. But another significant recurring cost for the shuttle program was for maintenance on the engines. Now, the SSME's have to be overhauled after every flight, costing ten's of millions of dollars. However, Henry Spencer a highly regarded expert on the history of space flight has said Rocketdyne studies show that with a lot of work to upgrade it, maintenance could be reduced to \$750K per flight per engine: Engine reusability (Henry Spencer) http://yarchive.net/space/rocket/eng...usability.html Spencer here said this would not be satisfactory for really large reductions in space costs. But this would be a reduction in SSME maintenance costs by 1 to 2 orders of magnitude, a major reduction in the costs for using the engine. A key question though is how much would be the cost to make the necessary upgrades to the engine. I also did not estimate the extra mass of the reentry/landing systems. Here's a diagram showing the specifications for the DIRECT team's version of this Phase I ca. 70 mt launcher: DIRECTv3 Jupiter-130 - LEO Cargo Launch Vehicle Configuration. http://www.launchcomplexmodels.com/D...deg_090606.jpg The dry mass of the core stage is given as 63.7 mt. This mass can be reduced by going to a common bulkhead design for the propellant tanks that for instance SpaceX was able to use to reduce dry mass for its Falcon vehicles. As it is now, the intertank on the shuttle ET actually weighs more than the oxygen tank: External Tank. http://science.ksc.nasa.gov/shuttle/...ewsref/et.html Going to a common bulkhead design would eliminate this mass, reducing the dry mass by about 5 mt. Also recent research has shown that dry mass of rocket vehicles in general can be reduced by 10% to 20%. This would take off about another 5 mt to 10 mt. I gave an estimate before in this thread of about 28% of the dry mass for reentry/landing systems. However, as I said probably with modern materials we can cut this in half. Then with all these reductions together the extra mass for reentry/landing systems might only be in the range of 7,000 kg. So we would still maintain 90% of the payload mass while gaining reusability and a longer useful life for this low cost heavy lift launcher. Bob Clark Last edited by RGClark; 05-02-2011 at 11:52 PM.
 05-24-2011, 03:41 PM #9 RGClark Mathematician Quote: Originally Posted by RGClark  NASA appears to be leaning to a 70 mt payload shuttle-derived launcher as an interim solution to developing a heavy lift vehicle. This would use two 4-segment SRM's as does the shuttle and an ET. But it would not have a shuttle orbiter, nor would this Phase I vehicle have an upper stage: SLS planning focuses on dual phase approach opening with SD HLV. April 25th, 2011 by Chris Bergin http://www.nasaspaceflight.com/2011/...pening-sd-hlv/ However, built into this plan is that at most 4 flights of this vehicle will be made before it is discontinued in favor of a more expensive, 130 mt payload upgrade. These 4 flights are to regarded as "test flights" according to the Bergin article. They will use 3 SSME's at a time and only 12 of those will be available including those taken from the retired space shuttles, thus allowing only 4 flights. Presumably after that either the production of new SSME's will be started or their expendable versions will be, or NASA will choose instead to use kerosene fueled engines for the core stage. However, a better plan in my view would be to explore methods in which this Phase I vehicle could be reusable. Then this low cost HLV could have many more missions as well as cutting costs in being reusable. This would give you more options as to when and if the more expensive vehicle needed to be developed... This is for the interim, Phase I, 70 mt launcher. This is to use two SRB's and an external tank as with the shuttle system, but no orbiter and no upper stage. However, the possibilities become especially interesting when we look at the case of making the Phase II, 100+ mt payload launcher reusable. This vehicle will have an additional upper stage. This has a significant advantage for the lightness of the reentry/landing systems in that only the upper stage at a small dry mass would have to have the full reentry systems of an orbiting vehicle. The upper stage of the DIRECT teams's Jupiter-246 for instance weighs less than 12,000 kg. Also, for this case the ET would reach a much reduced velocity and would not reach orbit so its reentry systems would be much simpler and lighter.[1] To get the full benefits of reusability we'll switch out the RL-10's or J-2X engines used on the upper stage for SSME(s). This does have a problem though in that the SSME would have to be made air startable. The benefits for reusability are so significant that costs estimates for this upgrade should be made. However, a different potential solution would also reap additional benefits. If instead of placing this stage atop the ET tank, we put it in parallel with it, then the stage could also be started on the ground. This has a benefit because now we could use cross-feed fueling between the ET and upper stage tanks. Cross-feed fueling with parallel staging is known to be able to increase your payload. For instance by using it for their Falcon Heavy vehicle SpaceX was able to increase its payload by 50%. Note also that we wouldn't have the development cost for a new 4 SSME engine core stage, as is currently planned for the Phase II vehicle. We would use the same 3-engine core stage as used for the Phase I vehicle. The extra thrust for the Phase II vehicle would come from the upper stage now firing in parallel from the start. However, another potentially game changing effect of doing this is that if you look at the mass ratio of this reconfigured upper stage with SSME(s) you see it has SSTO capability. This is because it has the weight optimization of an upper stage and now using an engine optimized to be most efficient during the entire flight to orbit it can reach orbit in a single stage with significant payload.[2] In fact not just the Jupiter-246 upper stage would have this capability, but in fact the Ariane 5's upper stage, the Apollo's S-II and S-IVB, and the planned Ares I upper stage would as well if switched out to use SSME(s). This is important because we will have already existing stages as well as the engines to make at least an initial version of this upper stage. This means we could have a significant cost reduction on an initial version of the upper stage. Another very key fact is because this upper stage can be used as a separate launcher and even manned launcher, thus with its own market, we could initiate it's development and production in parallel to the low cost Phase I vehicle. So we would get in fact not only a 70+ mt vehicle, but we would get the 100+ mt launcher and a manned launcher in just the same short time frame of the Phase I launcher and at a smaller cost than now planned for the Phase II, 100+ mt launcher. Indeed because there would be such a significant market for this manned SSTO vehicle, NASA might not have to pay for its development at all. Bob Clark 1.)http://www.orbiter-forum.com/showthr...68&postcount=6 2.)http://www.orbiter-forum.com/showthr...4&postcount=34

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