Some proposals for low cost heavy lift launchers.
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From: Robert Clark on 8 Jul 2010 06:58 I showed in this post: Newsgroups: sci.space.policy, sci.astro, sci.physics, sci.space.history From: Robert Clark <rgregorycl...(a)yahoo.com> Date: Tue, 4 May 2010 10:49:50 -0700 (PDT) Subject: Re: A kerosene-fueled X-33 as a single stage to orbit vehicle. http://groups.google.com/group/sci.space.policy/msg/eea2c9e8aaf61151?hl=en that two reconfigured X-33's mated bimese fashion and using a cross- feed fueling system could reduce the costs to orbit by *two orders* of magnitude. This shows there really is no logical objection to developing an SSTO. Because even if it is argued multistaged systems can carry more payload, you can carry *even* more payload by making those stages be separately SSTO capable. *Multiple times* more. I want to emphasize again the only reason why I used the Lockheed version of the X-33 was because it was already largely built. The other two proposed versions of a suborbital X-33 demonstrator by Rockwell and McDonnell-Douglas would also become fully orbital when switched from hydrogen to kerosene-fueled at comparable costs. These would be easier to make because you wouldn't have the problem that led to the X-33's downfall of lightweighting the tanks. Then the only thing keeping us from $100/lbs. launch costs is the acceptance that SSTO is indeed possible. That is why it is so imperative that the Falcon 1 first stage derived SSTO I discussed before be done because it would be so easy and CHEAP to achieve: Newsgroups: sci.space.policy, sci.astro, sci.physics, sci.space.history From: Robert Clark <rgregorycl...(a)yahoo.com> Date: Sun, 14 Mar 2010 18:24:37 -0700 (PDT) Subject: Re: A kerosene-fueled X-33 as a single stage to orbit vehicle. http://groups.google.com/group/sci.space.policy/msg/b2dfd3ce833c4470?hl=en Then finally the light bulb would come on. However, the bimese X-33 would involve some technical risk in that it would require the building of a second hydrocarbon-fueled X-33 and the low payload cost, due to the high payload capacity, would only obtain if the untested tank lightweighting methods really did bring the tankage ratio of the conformal tanks to be more in line with that of cylindrical tanks. Therefore I'll show here that an (expendable) heavy lift system can be produced with a payload capacity in the range of 40,000 kg to 60,000 kg at a minimal cost compared to the other heavy lift systems being proposed, and while using already existing components and at minimal technical risk. Previously I had argued that both the Falcon 1 and Falcon 9 first stages had a 20 to 1 mass ratio, and that this was important because this was the mass ratio often cited for a kerosene-fueled rocket to have SSTO capability. But that was based on the data on the SpaceLaunchReport.com site. The numbers on this site though are estimates and can be inaccurate. For instance from numbers actually released by SpaceX, the Falcon 1 first stage mass ratio is actually about 16.8 to 1. However, I was surprised to see in this recent news release from SpaceX that the Falcon 9 first stage mass ratio is actually better than 20 to 1(!): SPACEX ACHIEVES ORBITAL BULLSEYE WITH INAUGURAL FLIGHT OF FALCON 9 ROCKET. Cape Canaveral, Florida June 7, 2010 "The Merlin engine is one of only two orbit class rocket engines developed in the United States in the last decade (SpaceXs Kestrel is the other), and is the highest efficiency American hydrocarbon engine ever built. The Falcon 9 first stage, with a fully fueled to dry weight ratio of over 20, has the world's best structural efficiency, despite being designed to higher human rated factors of safety." http://www.spacex.com/press.php?page=20100607 Undoubtedly it is able to achieve this high mass ratio because it also uses common bulkhead design for the propellant tanks as does Falcon 1. Note that the original Atlas and the Saturn V upper stages nearly had SSTO mass ratios because they used common bulkheads. From this news release, we can also estimate the dry mass of the first stage: UPDATES: JULY 2009 - DECEMBER 2009. DRAGON/FALCON 9 UPDATE. Wednesday, September 23rd, 2009 "Weighing in at over 7,700 kg (17,000 lbs), the thrust assembly and nine Merlin engines represents over half the dry mass of the Falcon 9 first stage." http://www.spacex.com/updates_archive.php?page=2009_2 So I'll estimate the dry mass of the first stage as 15,000 kg, and the first stage total mass as 300,000 kg, and so the propellant mass as 285,000 kg. I'll again use three NK-33's as the engines, replacing the nine Merlin's. Using 660 kg as an estimate of the Merlin 1C mass, and 1,222 as the NK-33 mass, the dry mass becomes 15,000 - 9*660 + 3*1,222 = 12,726 kg. Again let's calculate what payload we can get using two of these Falcon 9's mated bimese fashion using cross-feed propellant transfer. This time I'll use a little more conservative average Isp of 335 s for the first portion of the trip where they are still mated together, but still assume some altitude compensation method is being used such as an aerospike. Then I'll still take the vacuum Isp as 360 s. Let's estimate the payload as 40,000 kg. Then we get a delta-V of: 335*9.8ln(1+285,000/(2*12,726+285,000+40,000)) = 1,954 m/s, for the first mated-together portion of the flight, and then: 360*9.8ln(1+285,000/(12,726+40,000)) = 6,552 m/s, for the upper stage portion, giving a total of about 8,500 m/s. Note again that by using more energetic hydrocarbon fuels, perhaps also densified by subcooling, you can get perhaps 50% higher payload to orbit than the 40,000 kg, so to perhaps 60,000 kg. This certainly qualifies as heavy lift if not super heavy lift. And could satisfy the requirements of a lunar mission at least for the launch system by using two launches. Bob Clark
From: Robert Clark on 10 Jul 2010 13:40 On Jul 8, 6:58 am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > ... I was surprised to see in this recent news release from > SpaceX that the Falcon 9 first stage mass ratio is actually better > than 20 to 1(!): > > SPACEX ACHIEVES ORBITAL BULLSEYE WITH INAUGURAL FLIGHT OF FALCON 9 > ROCKET. > Cape Canaveral, Florida June 7, 2010 > "The Merlin engine is one of only two orbit class rocket engines > developed in > the United States in the last decade (SpaceXs Kestrel is the other), > and is > the highest efficiency American hydrocarbon engine ever built. The > Falcon 9 > first stage, with a fully fueled to dry weight ratio of over 20, has > the > world's best structural efficiency, despite being designed to higher > human > rated factors of safety."http://www.spacex.com/press.php?page=20100607 > > Undoubtedly it is able to achieve this high mass ratio because it also > uses common bulkhead design for the propellant tanks as does Falcon 1. > Note that the original Atlas and the Saturn V upper stages nearly had > SSTO mass ratios because they used common bulkheads. > From this news release, we can also estimate the dry mass of the first > stage: > > UPDATES: JULY 2009 - DECEMBER 2009. > DRAGON/FALCON 9 UPDATE. > Wednesday, September 23rd, 2009 > "Weighing in at over 7,700 kg (17,000 lbs), the thrust assembly and > nine > Merlin engines represents over half the dry mass of the Falcon 9 first > stage."http://www.spacex.com/updates_archive.php?page=2009_2 > > So I'll estimate the dry mass of the first stage as 15,000 kg, and the > first stage total mass as 300,000 kg, and so the propellant mass as > 285,000 kg. > I'll again use three NK-33's as the engines, replacing the nine > Merlin's. Using 660 kg as an estimate of the Merlin 1C mass, and 1,222 > as the NK-33 mass, the dry mass becomes 15,000 - 9*660 + 3*1,222 = > 12,726 kg. > Again let's calculate what payload we can get using two of these > Falcon 9's mated bimese fashion using cross-feed propellant transfer. > This time I'll use a little more conservative average Isp of 335 s for > the first portion of the trip where they are still mated together, but > still assume some altitude compensation method is being used such as > an aerospike. Then I'll still take the vacuum Isp as 360 s. > Let's estimate the payload as 40,000 kg. Then we get a delta-V of: > > 335*9.8ln(1+285,000/(2*12,726+285,000+40,000)) = 1,954 m/s, for the > first > mated-together portion of the flight, and then: > 360*9.8ln(1+285,000/(12,726+40,000)) = 6,552 m/s, for the upper stage > portion, giving a total of about 8,500 m/s. > Several studies made during the 90's showed that it was actually easier to make a SSTO using dense fuels rather than hydrogen, such as this one: 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 The two key reasons for this is that though hydrogen's higher Isp means it needs only about half the mass ratio of, for example, kerosene it requires twice as much engine weight for the thrust produced and *3 times* as much tank weight for the propellant weight. These two advantages of the dense fuel over hydrogen swamp the hydrogen Isp advantage with the result that a similarly sized dense- fueled SSTO can carry *multiple* times more payload that a hydrogen- fueled one. This is what the math shows. And the actually produced Titan II rocket gives real world evidence for this as well. The Titan II stems from the earliest days of orbital rockets in the early 1960's yet its first stage had SSTO capability even then [i]using dense propellants[/ i]: http://en.wikipedia.org/wiki/Single-stage-to-orbit#Examples And now the Falcon 9 first stage having SSTO capability with a 20 to 1 mass ratio confirms this as well, while using standard structural techniques known for decades in the industry. Note that neither for the Titan II first stage or the Falcon 9 first stage was the intent to create an SSTO. The intent was to optimize the combination of the vehicle's weight and engine performance, the SSTO capability just happened accidentally. Why? Because getting SSTO-capability with dense propellant vehicles is [i]easy[/i]. Let's calculate the payload we can carry for the Falcon 9 first stage used as an SSTO. Since we're doing an SSTO where we need to maximize performance I'll assume altitude compensation methods are used such as an aerospike nozzle. In Dunn's paper "Alternate Propellants for SSTO Launchers." He gives an estimate of the average Isp over the flight with altitude compensation for kerosene (RP-1) as 338.3 s. Using the 8,500 m/s delta-V value I've been using to reach orbit, this would allow a payload of 11,000 kg : 338.3*9.8ln(1 + 285,000/(12,726 + 11,000)) = 8,507 m/s. But kerosene is not the most energetic hydrocarbon fuel. Another one described in Dunn's report is given as having an average Isp of 352 s, methylacetylene. With supercooling its overall density with LOX oxidizer is slightly above that of kerolox, so I'll take the propellant amount as 290,000 kg, then this would allow a payload of 14,200 kg: 352*9.8ln(1 + 290,000/(12,726 + 14,200)) = 8,505 m/s. Bob Clark
From: Robert Clark on 10 Jul 2010 14:04 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 V http://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-180 http://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
From: Brad Guth on 11 Jul 2010 13:08 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
From: Robert Clark on 16 Jul 2010 07:21
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 |