Some proposals for low cost heavy lift launchers.
Robert Clark
Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@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...@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 (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
Robert Clark
> ... I was surprised to see in this recent news release from
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
Robert Clark
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
Brad Guth
> 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:
>
>
> 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:
>
>
> 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:
>
>
> 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
Robert Clark
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
Robert Clark
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
Robert Clark
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
Robert 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
Jeff Findley
@y11g2000yqm.googlegroups.com>, rgrego...@yahoo.com says...
>
> On Jul 16, 11:37 am, Robert Clark <rgregorycl...@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?
Brad Guth
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
Pat Flannery
> There were a couple of proposals for LOX/kerosene boosters to replace
> the SRB's, but those proposals went nowhere.
Originally, the idea was to use liquid boosters for safety reasons, as
unlike the SRBs they could be shut down and jettisoned if one of them
went out-of-spec, and the orbiter could then hopefully return to the
launch site, either alone or after burning some of the fuel in the ET.
They also looked into a thrust termination system on the SRBs similar to
the one that was going to be used on the manned Titan III's, but the ET
apparently couldn't tolerate the blast effect of the blow-off venting
portals on the SRB nosecone being activated so close to it.
I still like the idea of sticking the Shuttle and ET atop a Saturn V
first stage: http://www.astronautix.com/graphics/s/shusat1c.gif
If something did go wrong with one of the F-1 engines, at least it would
be way behind the orbiter.
Combine that concept with the recoverable S-IC stage proposal, and you
could have had a system as recoverable as the present one that offered
superior safety during ascent due to having the ability to be shut down
in an emergency.
Pat
Robert 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:
>
>
> And some dense propellant engines have been tested to run on cryogenic
> hydrogen:
>
> LR-87 LH2http://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=all&Ntx=mode%20matchall&Ntt=19770059130
[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 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
eric gisse
[...]
If you are going to talk to yourself, don't do it in sci.physics.
Robert Clark
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 linked image 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.
bimese Falcon 9 launcher
http://i27.tinypic.com/2yxn2oz.jpg
Bob Clark
Robert Clark
"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
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
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
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