Some proposals for low cost heavy lift launchers.

[exoscientist]

I showed in this forum post:

viewtopic.php?f=15&t=21008&start=60#p450736

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:

viewtopic.php?f=15&t=21008&start=60#p451104

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 launch 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 (SpaceX’s 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

 

[exoscientist]

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 using dense propellants:

http://en.wikipedia.org/wiki/Single-sta ... t#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 easy.
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

 

[exoscientist]

The original Atlas from the 1960's was close to being SSTO capable:

http://en.wikipedia.org/wiki/Single-sta ... t#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

 

[rcsplinters]

Are you interested in single stage to orbit or heavy lift? In terms of cost effective heavy lift, the SD - HLV numbers are about the most plausible in terms of cost that I've seen. http://www.nasaspaceflight.com/2010/06/sd-hlv-assessment-highlights-post-shuttle-solution/. SSTO is interesting, certainly, but is 100 tons to orbit really practical for that option in the near term, say 10 - 15 years? Sounds like it would be best focused on humans to LEO then pursue the sort of heavy lift needed to support operations beyond LEO.

 

[orionrider]

You mean something like that:
The boosters are fully reusable, so I think you can count it as a kind of SSTO.
http://www.k26.com/buran/Info/Energia_H ... oster.html

Full launch profile: http://www.k26.com/buran/assets/images/enert1.jpg

 

[ZiraldoAerospace]

If the Falcon 1 is SSTO, what about the Falcon 1e? Also, does anyone know if LockMart still has the mostly built X-33 laying around? It would at least be interesting to pull the aerospike engines off and other parts and build a new, kerosene SSTO vehicle, using many of the original parts from the X-33. Just throwing some ideas out there.

 

[neutrino78x]

The X-33 was awesome. I would like to see one that takes off from a horizontal position, though, so it doesn't need a special facility to take off.

It seems like they could fix the issue with the hydrogen storage if they had more funding....I prefer pure hydrogen rockets for environmental reasons.

--Brian

 

[neutrino78x]

(by takes off from a horizontal position, I'm actually thinking, ideally, of a harrier jet type of liftoff.)

 

[ZiraldoAerospace]

The X-33 was awesome. I would like to see one that takes off from a horizontal position, though, so it doesn't need a special facility to take off.

It seems like they could fix the issue with the hydrogen storage if they had more funding....I prefer pure hydrogen rockets for environmental reasons.

--Brian

My question is, why bother spending that funding? It seems to me that hydrogen rockets are inferior, as demonstrated above and on several other forum topics. Also, people whine about how little the payload will be on an SSTO vehicle, but if it took off horizontally, it would be reduced even further, or the vehicle won't make it to orbit at all.

 

[exoscientist]

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 ... ?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

 

[exoscientist]

Nice video here on the high performance Russian engines:

The_Engines_That_Came_In_From_The_Cold.
http://video.google.com/videoplay?docid ... 443&hl=en#


Bob Clark

 

[exoscientist]

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

 

[kelvinzero]

You mean something like that:
The boosters are fully reusable, so I think you can count it as a kind of SSTO.
http://www.k26.com/buran/Info/Energia_H ... oster.html

Some have also coined the expression 1.5STO for 'one-and-a-half-stage-to-orbit', e.g., the Atlas
http://en.wikipedia.org/wiki/Two-stage-to-orbit

I have always liked the reusable booster idea. By using just one or two with small payloads and a whole heap with super heavy lift, you could effectively get a high flight rate (number of boosters launched per year) which is the key to reusables being cost effective. Because the boosters do not carry crew, a failure to recover one would not halt the entire industry. Because they do not make it all the way to orbit, their engines can be more robust and they are not continually exposed to that tough reentry, and fly-back or boost back becomes more practical.

It would also be great if the stages could share fuel, so if one engine fails it is not just dead weight.
This would also be a useful feature of second stages. I see a lot of commonality between the principles of docking a second stage with a fuel depot for a BEO mission, joining the tanks of multiple second stages to make a larger fuel depot, and joining multiple second stages to make a large Earth Departure Stage.

 

[vulture4]

If you wish to converse with me, define your terms. -Voltaire

Reusable Launch Vehicles (RLVs), Single Stage to Orbit launch vehicles (SSTO), and heavy lift launch vehicles (HLVs) are all concepts for dealing with the most serious problem in spaceflight, the extremely high cost of launch. But they are distinctly different strategies.

An HLV attempts to reduce costs by carrying a bigger load with each launch. The Shuttle, with a cargo capacity of about 32 tons to equatorial LEO, and the Delta IV Heavy, with about 25 tons, both are at the low end of the heavy lift spectrum. The Saturn V represented the high end of the HLV spectrum with a LEO payload capacity of well over 100 tons. Virtually all HLVs are large expendable multistage rockets.

An SSTO is a vehicle that can get into orbit with only one stage. An SSTO can be completely expendable,. With all the assertions made regarding the Titan and Falcon, I am unaware that any single stage rocket has ever actually gone into orbit. While the majority of the structure of the Atlas could make orbit, the booster engine section which was dropped was actually a substantial part of the dry mass.

An RLV is a launch vehicle that is partially or fully reusable. The Shuttle is the only RLV in service, and it is only partly reusable since the ET is expended and the boosters must be fully disassembled and rebuilt after each launch. However an RLV can be suborbital, and both the X-15, which flew 199 flights, and the SpaceShip 1 and 2 are successful examples of manned suborbital space launch vehicles. In most cases the upper stage of an RLV also serves as a spacecraft, though it does not remain permanently in orbit.

 

[EarthlingX]

Shuttle : 24t to LEO, not 32t.
http://en.wikipedia.org/wiki/Space_Shuttle

32t perhaps if launched from the equator.

Heavier launcher makes sense, when you either have something heavy to launch, or when you use multiple payloads on the same launcher, to get a bigger share of the launch market with the same launch frequency.

It is a bit of a problem with USA launchers, since you can hardly export them to compete on the global launch market, which is on the rise, just not in the America.

Usually countries limit imports to protect their industry, in this case they are stopping exports. I have a serious problem understanding that, since almost anyone can buy Protons or Soyuzes, and that makes security reasons kinda weak.
Limiting American launcher exports is beneficial to Proton and Ariane sales, if that was the idea.

It also makes military the most important customer, and that doesn't help launch costs.

Nothing new on the ITAR front ..

 

[Astro_Robert]

I don't believe ITAR is the sole driver of the limited use of American launchers in the foreign launch market. I believe cost is the primary driver. EELVs Atlas and Delta were designed to US gov specificationss, and only have launch pads on US soil, requiring use and coordination of US ranges. The cost structure here is a lot higher for those things than say Russia or China. If some company was willing to pay the cost of a US rocket rather than the mostly lower cost for a Russian, Chinese or French one; I do believe it would be possible for them to buy a launch on an EELV. Assuming Falcon continues to be successful, I also believe that they could compete on the world market without violating ITAR (Internation Traffic in ARms regulations), assuming their cost structure is competitive with the Russians and Chinese.

In fact, Boeing briefly had a JV with the Russians to launch their Zenits from a boat, known as sealaunch. Lockheed also had a deal with the Russians for Proton launches in Russia.

ITAR does influence our ability to share our rocket know how with other countries (there was a bit of a scandal maybe ~15-20 years ago about some US companies helping China's launch industry after some LongMarch launch failures) as rocket and ICBM technologies are closely related.

 

[Astro_Robert]

I do not believe that X-33 is available any more, just like Saturn V is also not available. Also, the engines were never installed, but were tesed at Stennis fairly thoroughly.

Both Lockheed and Boeing through their United Launch Aliance, submitted options to NASA to upgrade their respective Atlas and Delta launch vehicles to meet the requirements of crew/cargo launchers for NASA. For some reason NASA seems to be stuck on using the prohibitively expensive Shuttle technology at an excessively low launch rate instead.

Right now, ATK can only make a certain relativley low number of solids per year, so even if the solid was desireable/acceptable to everyone, the launch rate would remain low for the foreseeable future. Thiokol/ATK used to have 2 plants that made the solids, but 1 of them exploded like 10 or so years ago. Being down to 1 plant, it is only possible to support a modest launch rate using solids, whatever their safety/desireability may be.

 

[exoscientist]

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

Found this after searching on Astronautix.com:

RD-0120.
"Engine Model: RD-0120-CH. Manufacturer Name: RD-0120-CH. Designer: Kosberg. Propellants: Lox/LCH4. Thrust(vac): 1,576.000 kN (354,298 lbf). Isp: 363 sec. Mass Engine: 2,370 kg (5,220 lb). Chambers: 1. Chamber Pressure: 172.50 bar. Oxidizer to Fuel Ratio: 3.40. Thrust to Weight Ratio: 67.80. Country: Russia. Status: Design concept 1990's.
Proposed variant of the RD-0120 engine using liquid methane instead of hydrogen as propellant."
http://www.astronautix.com/engines/rd0120.htm

The RD-0120 was the hydrogen fueled engine used on the Russian Energia heavy lift booster, which lifted the Russian Buran space shuttle for instance. I can't tell from this description though if it was actually tested with liquid methane or if these were only theoretical studies.
After searching on the NASA Technical Report server I found some theoretical studies that suggest that the SSME could be converted to hydrocarbon-fueled at relatively low cost (compared to developing a new engine.)

Booster engines derived from the Space Shuttle Main Engine.
Sobin, A. J.; Poynor, S. P.; Cross, E
"By using a majority of the current SSME engine components for the LOX/RE-1 booster engine, engine development time and cost can be significantly reduced compared to the development of a new engine."
Propulsion Conference, 13th, July 11-13, 1977, Orlando, FL
http://ntrs.nasa.gov/search.jsp?N=0&Ntk ... 9770059130 [abstract only]

Tripropellant engine study.
Wheeler, D. B.; Kirby, F. M.
NASA-CR-150808; RI/RD78-215
"SUMMARY.
"The results of these studies have shown that the conversion of an SSME engine to
a high chamber pressure, dual-mode fuel engine will require major modifications
to the hardware and/or the addition of a significant number of new engine cowponents.
However, the study has shown numerous possibilities for the use of SSME
hardware derivatives in a single-mode LOX/hydrocaxbon engines. It was also
shown that a reduced chamber pressure version of a staged combustion SSME is
operationally feasible using the existing fuel-rich preburners and main chamber
injectors. Certain turbomachinery modifications or additions are required for
a total low chamber pressure ( 2300 psia) engine system. This study also has
shown that the engine system concepts applicable to the dual-mode systems are
somewhat narrowed since the operational constraints of two systems must be
considered."
http://hdl.handle.net/2060/19780024238 [full text, 145 pages]

Another possibility might be to adapt the hydrogen-fueled aerospike engines intended for the VentureStar to be hydrocarbon-fueled. This theoretical study from 1977 was on the possibility that an aerospike engine of the linear configuration later adopted for the VentureStar could be dual-fueled, i.e., running on both hydrocarbon and hydrogen:

Linear aerospike engine study.
Diem, H. G.; Kirby, F. M.
NASA-CR-135231; RI/RD77-170
http://hdl.handle.net/2060/19780003139 [full text, 246 pages]

This would have the advantage that it would already have altitude compensation. If the dual-fuel modes are workable this would also increase performance.
This study was primarily on dual-fuel operation but did also study hydrogen only operation. It might be useful to compare the predicted hydrogen only operation with the performance actually found with the aerospike engines created for the X-33 sub-scale demonstrator. If the measured performance does correspond to the predicted values that would give confidence that the dual-fuel version would also be close to the predicted values.


Bob Clark

 

[James_Bull]

Nasa could always just hitch a ride on China's future "moon rocket"!
http://www.bbc.co.uk/news/science-environment-10762634

 

[EarthlingX]

Nasa could always just hitch a ride on China's future "moon rocket"!
http://www.bbc.co.uk/news/science-environment-10762634

That is Anatoly Zak .. :shock:

http://www.bbc.co.uk : China considers big rocket power

26 July 2010 Last updated at 13:02 GMT

By Anatoly Zak Science reporter

Engines on China's Long March-5 will develop 120 tonnes of thrust

Chinese engineers are considering a new super-powerful engine for the next generation of space rockets, say officials.

According to Li Tongyu, general manager of the marketing department at the China Academy of Launch Vehicle Technology (CALT), engineers are currently studying a rocket engine capable of generating thrust of 600 tonnes.

If China succeeds in the development of such power, it would increase the nation's capabilities in space by orders of magnitude.

Wiki : Long March 5

Long March 5 (LM-5, CZ-5, or Changzheng 5) is a Chinese next-generation heavy lift launch system that is currently under development by China Academy of Launch Vehicle Technology (CALT). Currently, six CZ-5 vehicle configurations[1] are planned for different missions, with a maximum payload capacity of 25,000 kg to LEO and 14,000 kg to GTO. The Long March 5 will have the second largest "carrying capacity factor" of any rocket after Boeing's Delta IV Heavy.[2] The CZ-5 rocket is due to be first launched in 2014 from Wenchang Satellite Launch Center on Hainan island.

 

[exoscientist]

Here are some possibilities for lower cost super heavy lift launchers, in the 100,000+ kg payload range. As described in this article the proposals for the heavy lift launchers using kerosene-fueled lower stages are focusing on using diameters for the tanks of that of the large size Delta IV, at 5.1 meters wide or the even larger shuttle ET, at 8.4 meters wide:

All-Liquid: A Super Heavy Lift Alternative?
by Ed Kyle, Updated 11/29/2009
http://www.spacelaunchreport.com/liquidhllv.html

The reason for this is that it is cheaper to create new tanks of the same diameter as already produced ones by using the same tooling as those previous ones. This is true even if switching from hydrogen to kerosene in the new tanks.
However, I will argue that you can get super heavy lift launchers without using the expensive upper stages of the other proposals by using the very high mass ratios proven possible by SpaceX with the Falcon 9 lower stage, at above 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 (SpaceX’s 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

We will use tanks of the same size as these other proposals but will use parallel, "bimese", staging with cross-feed fueling. This method uses two copies of 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 it's expended its fuel. See the attached images below for how parallel staging with cross-feed fueling works.
Do the calculation first for the large 8.4 meter wide tank version. At the bottom of Kyle's "All-Liquid: A Super Heavy Lift Alternative?" article is given the estimated mass values for the gross mass and propellant mass of the 8.4 meter wide core first stage. The gross mass of this single stage is given as 1,423 metric tons and the propellant mass as 1,323 metric tons, so the empty mass of the stage would be approx. 100 metric tons (a proportionally small amount is also taken up by the residual propellant at the end of the flight.) Then the mass ratio is 14 to 1. However, the much smaller Falcon 9 first stage has already demonstrated a mass ratio of over 20 to 1.
A key fact about scaling is that you can increase your payload to orbit more than the proportional amount indicated by scaling the rocket up. Said another way, by scaling your rocket larger your mass ratio in fact gets better. The reason is the volume and mass of your propellant increases by cube of the increase and key weight components such as the engines and tanks do also, but some components such as fairings, avionics, wiring, etc. increase at a much smaller rate. That savings in dry weight translates to a better mass ratio, and so a payload even better than the proportional increase in mass.
This is the reason for example that proponents of the "big dumb booster" concept say you reduce your costs to orbit just by making very large rockets. It's also the reason that for all three of the reusable launch vehicle (RLV's) proposals that had been made to NASA in the 90's, for each them their half-scale demonstrators could only be suborbital.
Then we would get an even better mass ratio for this "super Evolved Atlas" core than the 20 to 1 of the Falcon 9 first stage, if we used the weight saving methods of the Falcon 9 first stage, which used aluminum-lithium tanks with common bulkhead design. It would also work to get a comparable high mass ratio if instead the balloon tanks of the earlier Atlas versions prior to the Atlas V were used.
So I'll use the mass ratio 20 to 1 to get a dry mass of 71.15 mT, call it 70,000 kg, though we should be able to do better than this. We'll calculate the case where we use the standard performance parameters of the RD-180 first, 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 338 s for the standard vacuum Isp of the RD-180. For the required delta-V I'll use 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 115 mT. Then the delta-V for the first leg is 329*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 115)) = 1,960 m/s. For the second leg the delta-V is 338*9.8ln(1 + 1,323/(70 + 115)) = 6,950 m/s. So the total delta-V is 8,910 m/s, sufficient for LEO with the 115 mT payload, by the 8,900 m/s value I'm taking here as required for a dense propellant vehicle.
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 payload as 145,000 kg. For the first leg, the delta-V is 338.3*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 145)) = 1,990 m/s. For the second leg the delta-V is 360*9.8ln(1 + 1,323/(70 + 145)) = 6,940 m/s, for a total delta-V of 8,930 m/s, sufficient for orbit with the 145,000 kg payload.
Now we'll estimate the payload using the higher energy fuel 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 I'll take the vacuum Isp as 380 s. Estimate the payload as 175,000 kg. Then the delta-V over the first leg is 352*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 175)) = 2,040 s. For the second leg the delta-V will be 380*9.8ln(1 + 1,323/(70 + 175)) = 6,910 s, for a total delta-V of 8,950 m/s, sufficient for orbit with the 175,000 kg payload.



Bob Clark

 

[exoscientist]

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

 

[Valcan]

quote]
Why do i have a sudden feeling i know where the plans for the saturn V and its engines are. :roll:

 

[docm]

New from the SpaceX folks at the 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference this week....

 

[exoscientist]

Thanks for that. Curiously member Dwightlooi presented very similar looking proposals for what SpaceX's evolution might be for heavy lift:

Roadmap for SpaceX Heavy Lift Vehicles.
by dwightlooi » Fri Jan 02, 2009 3:06 am

viewtopic.php?f=15&t=11793

Beyond the Falcon 9
by dwightlooi » Fri Jul 24, 2009 8:59 pm
Perhaps SpaceX can fill in for NASA's bungling super heavy lift efforts. Perhaps what we need is not the ARES V, or even DIRECT or JUPITER. Perhaps what we need is a 2-stage, all hydrocarbon launch vehicle designed for economy, reliability, ruggedness and simplicity rather than performance. Perhaps SpaceX's heavy lift road map can look like this...

viewtopic.php?t=19025

Affordable Heavy Lift
by dwightlooi » Sun Feb 21, 2010 4:36 pm
FICTION

viewtopic.php?t=22912


Perhaps DwightLooi had some inside knowledge?


Bob Clark