Some proposals for low cost heavy lift launchers.
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From: Robert Clark on 16 Jul 2010 10:46 Post #1 in this thread showed you could get a low cost heavy lift launcher in the 50,000+ kg class by using a bimese, cross-feed fueled configuration of Falcon 9 first stages, that replaced the Merlin engines with currently available high performance engines, and using known high energy density hydrocarbon fuels. Here I'll show by using this idea with a three stage system, a trimese if you will, you can raise that payload to the 75,000 kg range. Senator Bill Nelson, chairman of the Senate subcommittee on NASA, has said he favors a heavy lift solution to begin development next year that is at least in the 75,000 kg range: Senator Nelson Previews 2010 NASA Reauthorization Bill STATUS REPORT Date Released: Wednesday, July 14, 2010 http://www.spaceref.com/news/viewsr.rss.spacewire.html?pid=34492 Again as in post #1, I'll take the dry weight of the Falcon 9 first stage with the 9 Merlin engines replaced with 3 NK-33's as 12,726 kg and the propellant load as 285,000 kg. You could also do this with a single RD-180 as the engine. You would not get any weight savings in this case in the dry mass, but the Isp would be slightly better than when using NK-33's. Now we will be using three mated together Falcon 9 first stages. Note this looks similar to the Falcon 9 Heavy. But by using higher performance engines, cross-feed fueling, altitude-compensation methods, and high energy density hydrocarbon fuel we will be able to increase the payload to LEO 2.5 to 3 times and without using the upper stage of the Falcon 9 Heavy. As before I will take the average Isp you can get using altitude-compensation methods such as aerospike nozzles with kerolox from table 2 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 It gives the average Isp as 338.3 s. For the vacuum Isp, I'll take the 360 s Isp reached by other Russian high performance engines that were optimized for vacuum performance. Note that such vacuum optimized engines normally get quite poor performance at sea level, so altitude- compensation methods will be a necessity to maintain high performance both at sea level and at high altitude. Then the way the cross-feed fueling will work is that at launch all the engines from all three Falcon 9's will be firing but the propellant for all of them will be coming from only a single Falcon 9 tank. Then when the propellant from that tank is expended, that Falcon 9 will be jettisoned. This will leave two mated Falcon 9's both with their full propellant loads. Now all the engines will again be firing but again all the propellant will be coming from a single Falcon 9 tank. When this tanks propellant is expended this Falcon 9 will also be jettisoned. Finally for the final leg of the trip, the remaining Falcon 9 will still have its full propellant load which will be used to propel the payload to orbit. Let's calculate the delta-V we can achieve. Estimate the payload that can be lofted to orbit as 65,000 kg. For the first leg of the trip with all three Falcon 9's connected, the ending mass of the vehicle for this first first leg will be 3*12,726 + 2*285,000 + 65,000 kg. So the delta-V will be 338.3*9.8ln(1 + 285,000/(3*12,726 + 2*285,000 + 65,000)) = 1,170 m/s. For the second leg using two Falcon 9's, the ending mass will be 2*12,726 + 285,000 + 75,000 kg. This will be at high altitude so we'll use the vacuum Isp of 360 s. Then the delta-V produced by the second leg will be 360*9.8ln(1 + 285,000/(2*12,726 + 285,000 + 65,000)) = 1,992 m/s. For the final leg using a single Falcon 9, the ending mass will be 12,726 + 65,000, so the delta-V here will be 360*9.8ln(1 + 285,000/(12,726 + 65,000)) = 5,435 m/s. Then the total delta-V will be 8,597 m/s, sufficient for orbit using the 8,500 m/s value I'm taking as the delta-V for LEO. Note the 65,000 kg payload is twice that of the Falcon 9 Heavy and without the Falcon 9 upper stage. Now let's calculate the payload using a higher energy hydrocarbon fuel. Again in Dunn's report in table 2 for the fuel methylacetylene, the average Isp is given as 352 s. Dunn also gives what would be the maximum theoretical vacuum Isp in this table as 391.1 s for methylacetylene. High performance engines can get close to this theoretical value, at 97% and above. So I'll take the vacuum Isp of our high performance engine using methylacetylene as the fuel as 380 s. To maximize our fuel load we'll also use the chilled version of our propellant. The overall density will then be slightly above that of kerolox, so we'll take the propellant load as 290,000 kg. Let's calculate the delta-V using the estimate of 80,000 kg as our payload. Then the first leg delta-V is 352*9.8ln(1 + 290,000/(3*12,726 + 2*290,000 + 80,000)) = 1,198 m/s. The second leg delta-V is 380*9.8ln(1 + 290,000/(2*12,726 + 290,000 + 80,000)) = 2,048 m/s. And the third leg delta-V is 380*9.8ln(1 + 290,000/(12,726 + 80,000)) = 5,279 m/s. Then the total delta-V is 8,525 m/s, sufficient for orbit with a 80,000 kg payload. Bob Clark
From: Robert Clark on 16 Jul 2010 11:37 Nice video here on the high performance Russian engines: The_Engines_That_Came_In_From_The_Cold. http://video.google.com/videoplay?docid=-6986776989850537443&hl=en# Bob Clark
From: Robert Clark on 16 Jul 2010 13:45 On Jul 16, 11:37 am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > Nice video here on the high performance Russian engines: > > The_Engines_That_Came_In_From_The_Cold.http://video.google.com/videoplay?docid=-6986776989850537443&hl=en# > > Bob Clark Anyone know if there has been research on converting the shuttle main engines to hydrocarbon fueled? I was annoyed that NASA had earlier canceled a program to develop a heavy-thrust hydrocarbon engine after the Ares I and V were chosen. We would have a reusable and man-rated heavy-thrust kerosene engine *now* if it weren't for that. The SSME's have to operate under severe tolerances using cryogenic hydrogen since the liquid hydrogen is so cold yet LH2/LOX burns at such high temperature. I would think using kerosene/LOX for instance would put less severe conditions on the engine operation. Note that other liquid hydrogen engines have been successfully run on other fuels under test conditions: The RL10 (Bruce Dunn; Gary Hudson; Henry Spencer) http://yarchive.net/space/rocket/rl10.html And some dense propellant engines have been tested to run on cryogenic hydrogen: LR-87 LH2 http://www.astronautix.com/engines/lr87lh2.htm Bob Clark
From: Jeff Findley on 16 Jul 2010 14:24 In article <6a5f8191-d605-4b39-aa39-ced18b4b90b4 @y11g2000yqm.googlegroups.com>, rgregoryclark(a)yahoo.com says... > > On Jul 16, 11:37�am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > > Nice video here on the high performance Russian engines: > > > > The_Engines_That_Came_In_From_The_Cold.http://video.google.com/videoplay?docid=-6986776989850537443&hl=en# > > > > � � Bob Clark > > Anyone know if there has been research on converting the shuttle main > engines to hydrocarbon fueled? Doubtful. It would be too much of a redesign. The lower performance of such engines would make the current design completely unworkable. There were a couple of proposals for LOX/kerosene boosters to replace the SRB's, but those proposals went nowhere. > I was annoyed that NASA had earlier > canceled a program to develop a heavy-thrust hydrocarbon engine after > the Ares I and V were chosen. We would have a reusable and man-rated > heavy-thrust kerosene engine *now* if it weren't for that. Also doubtful. Such an engine development program would take many years and quite a bit of money (money is something NASA is always short of). > The SSME's have to operate under severe tolerances using cryogenic > hydrogen since the liquid hydrogen is so cold yet LH2/LOX burns at > such high temperature. I would think using kerosene/LOX for instance > would put less severe conditions on the engine operation. You'd still have LOX, so you're not getting rid of all of the trouble, but LOX isn't nearly as cryogenic as LH2, so you may have a point. > Note that other liquid hydrogen engines have been successfully run on > other fuels under test conditions: > > The RL10 (Bruce Dunn; Gary Hudson; Henry Spencer) > http://yarchive.net/space/rocket/rl10.html The RL-10's expander cycle is very tolerant of, well, just about anything. It's not the most efficient engine around, but it's no slouch either. > And some dense propellant engines have been tested to run on cryogenic > hydrogen: > > LR-87 LH2 > http://www.astronautix.com/engines/lr87lh2.htm True, but these things aren't quite as easy as it seems. From above: The entire development took place from 1958-1960, and was of the same magnitude as the parallel modification of the LR-87 engine to burn storable propellants for the Titan 2. Three years during this period was quite a bit of time. Rocket engine technology was advancing at a furious pace at the time. Such a program today might take longer. Jeff -- The only decision you'll have to make is Who goes in after the snake in the morning?
From: Brad Guth on 16 Jul 2010 16:11
On Jul 16, 4:21 am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > On Jul 11, 1:08 pm, Brad Guth <bradg...(a)gmail.com> wrote: > > > > > On Jul 10, 11:04 am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > > > > The original Atlas from the 1960's was close to being SSTO capable: > > > >http://en.wikipedia.org/wiki/Single-stage-to-orbit#Examples > > > > It was able to be highly weight-optimized because it used what is > > > called pressure-stabilized or "balloon tanks". These were tanks of > > > thinner wall thickness than normal and were able to maintain their > > > structure in being pressurized. The wall thickness was so thin that > > > they could not stand alone when not filled with fuel. To be stored the > > > tanks had to be filled with an inert gas such as nitrogen, otherwise > > > they would collapse under their own weight. > > > The Atlas III also uses balloon tanks and a common bulkhead design, > > > used effectively by the SpaceX Falcon launchers to minimize weight. > > > The Falcons probably are able to get the good weight optimization > > > comparable to that of the Atlas launchers without using balloon tanks > > > because their tanks are made of aluminum instead of the steel used > > > with the Atlas tanks. The Atlas launchers might be able to weight- > > > optimize their tanks even further by using aluminum for their balloon > > > tanks, but there may be structural reasons that for balloon tanks > > > steel has been preferred. > > > The specifications for the Atlas III are given on this Astronautix.com > > > page for the Atlas V: > > > > Atlas Vhttp://www.astronautix.com/lvs/atlasv.htm > > > > The gross mass is given as 195,628 kg and the empty mass is given as > > > 13,725 kg, resulting in a propellant mass of 181,903 kg. The Atlas III > > > uses an RD-180 engine: > > > > RD-180http://www.astronautix.com/engines/rd180.htm > > > > The Atlas III is actually somewhat overpowered with the RD-180, as > > > evidenced by the fact that Atlas V carrying 50% more propellant is > > > still able to use the RD-180. For an SSTO the weight of the engines is > > > a major factor that has to be tailored to the size of the vehicle. A > > > engine of greater power may be unsuitable for the SSTO purpose simply > > > because the larger than needed engine weight may prevent the required > > > mass ratio to be SSTO. > > > So again I'll use NK-33's two this time for the engines: > > > > NK-33.http://www.astronautix.com/engines/nk33.htm > > > > Then the engine weight is reduced from 5,393 kg to 2,444 kg. This > > > brings the dry mass to 10,776 kg, and the gross mass is now 192,679 > > > kg. So the mass ratio is 17.9. > > > Using aerospike nozzles or other altitude compensation methods on the > > > NK-33 we might be able to get the vacuum Isp to increase to 360 s and > > > the average Isp over the flight to be 335 s. Then this would allow a > > > payload of 4,000 kg, using the 8,500 m/s delta-V I'm taking as that > > > required for orbit: > > > > 335*9.8ln(1 + 181,903/(10,776 + 4,000)) = 8,498 m/s. > > > > Now let's calculate the payload for two Atlas III's mated bimese > > > fashion and using cross-feed fueling: > > > with a payload of 22,000 kg, we get a first stage delta-V of > > > 335*9.8ln(1 + 181,903/(2*10,776 + 181,903 + 22,000)) = 1,942 m/s, and > > > a second stage delta-V of 360*9.8ln(1 + 181,903/(10,776 + 22,000)) = > > > 6,661 m/s for a total delta-V of 8,573 m/s. > > > > Bob Clark > > > What have they that's new in HTP + hydrocarbons? > > > http://www.astronautix.com/engines/rd502.htm#RD-502 > > http://www.astronautix.com/props/index.htm > > > http://www.dunnspace.com/alternate_ssto_propellants.htm > > propargyl alcohol + HTP Isp = 350 > > cyclopropane + HTP Isp = 351.5 > > > ~ BG > > As Dunn's reprt shows there are some fuel combinations using H2O2 as > the oxidizer that give better performance than kerosene/LOX. This > would be most useful for example for Air Force systems intended to be > maneuverable in space, since H2O2 is easier to store in space rather > than LOX since it is non-cryogenic. > > Bob Clark h2o2 at 99.5% can be stored nearly indefinitely, especially if it's kept cool and sealed up. The same can be said of viable hydrocarbons. The Boeing OASIS gateway/outpost at Selene L1 would be a good location for storing a few thousand tonnes, and using an artificial shade should be sufficient for easily avoiding their being toasted to death. ~ BG |