Some proposals for low cost heavy lift launchers.

I showed in this post:

Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark
Date: Tue, 4 May 2010 10:49:50 -0700 (PDT)
Subject: A kerosene-fueled X-33 as a single stage to orbit
vehicle.
http://groups.google.com/group/sci.s...aaf61151?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
Date: Sun, 14 Mar 2010 18:24:37 -0700 (PDT)
Subject: A kerosene-fueled X-33 as a single stage to orbit
vehicle.
http://groups.google.com/group/sci.s...833c4470?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

On Jul 8, 6:58*am, Robert Clark 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 (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.


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-...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 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

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

http://en.wikipedia.org/wiki/Single-...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

On Jul 10, 11:04*am, Robert Clark wrote:
The original Atlas from the 1960's was close to being SSTO capable:

http://en.wikipedia.org/wiki/Single-...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

On Jul 11, 1:08*pm, Brad Guth wrote:
On Jul 10, 11:04*am, Robert Clark wrote:

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

http://en.wikipedia.org/wiki/Single-...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

On Jul 16, 4:21*am, Robert Clark wrote:
On Jul 11, 1:08*pm, Brad Guth wrote:



On Jul 10, 11:04*am, Robert Clark wrote:

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

http://en.wikipedia.org/wiki/Single-...orbit#Examples

It was able to be highly weight-optimized because it used what is
called pressure-stabilized or "balloon tanks". These were tanks of
thinner wall thickness than normal and were able to maintain their
structure in being pressurized. The wall thickness was so thin that
they could not stand alone when not filled with fuel. To be stored the
tanks had to be filled with an inert gas such as nitrogen, otherwise
they would collapse under their own weight.
The Atlas III also uses balloon tanks and a common bulkhead design,
used effectively by the SpaceX Falcon launchers to minimize weight.
The Falcons probably are able to get the good weight optimization
comparable to that of the Atlas launchers without using balloon tanks
because their tanks are made of aluminum instead of the steel used
with the Atlas tanks. The Atlas launchers might be able to weight-
optimize their tanks even further by using aluminum for their balloon
tanks, but there may be structural reasons that for balloon tanks
steel has been preferred.
The specifications for the Atlas III are given on this Astronautix.com
page for the Atlas V:

Atlas Vhttp://www.astronautix.com/lvs/atlasv.htm

The gross mass is given as 195,628 kg and the empty mass is given as
13,725 kg, resulting in a propellant mass of 181,903 kg. The Atlas III
uses an RD-180 engine:

RD-180http://www.astronautix.com/engines/rd180.htm

The Atlas III is actually somewhat overpowered with the RD-180, as
evidenced by the fact that Atlas V carrying 50% more propellant is
still able to use the RD-180. For an SSTO the weight of the engines is
a major factor that has to be tailored to the size of the vehicle. A
engine of greater power may be unsuitable for the SSTO purpose simply
because the larger than needed engine weight may prevent the required
mass ratio to be SSTO.
So again I'll use NK-33's two this time for the engines:

NK-33.http://www.astronautix.com/engines/nk33.htm

Then the engine weight is reduced from 5,393 kg to 2,444 kg. This
brings the dry mass to 10,776 kg, and the gross mass is now 192,679
kg. So the mass ratio is 17.9.
Using aerospike nozzles or other altitude compensation methods on the
NK-33 we might be able to get the vacuum Isp to increase to 360 s and
the average Isp over the flight to be 335 s. Then this would allow a
payload of 4,000 kg, using the 8,500 m/s delta-V I'm taking as that
required for orbit:

335*9.8ln(1 + 181,903/(10,776 + 4,000)) = 8,498 m/s.

Now let's calculate the payload for two Atlas III's mated bimese
fashion and using cross-feed fueling:
with a payload of 22,000 kg, we get a first stage delta-V of
335*9.8ln(1 + 181,903/(2*10,776 + 181,903 + 22,000)) = 1,942 m/s, and
a second stage delta-V of 360*9.8ln(1 + 181,903/(10,776 + 22,000)) =
6,661 m/s for a total delta-V of 8,573 m/s.

Bob Clark

What have they that's new in HTP + hydrocarbons?

*http://www.astronautix.com/engines/rd502.htm#RD-502
*http://www.astronautix.com/props/index.htm

*http://www.dunnspace.com/alternate_ssto_propellants.htm
*propargyl alcohol + HTP Isp = 350
*cyclopropane + HTP Isp = 351.5

*~ BG

*As Dunn's reprt shows there are some fuel combinations using H2O2 as
the oxidizer that give better performance than kerosene/LOX. This
would be most useful for example for Air Force systems intended to be
maneuverable in space, since H2O2 is easier to store in space rather
than LOX since it is non-cryogenic.

* Bob Clark

h2o2 at 99.5% can be stored nearly indefinitely, especially if it's
kept cool and sealed up. The same can be said of viable
hydrocarbons. The Boeing OASIS gateway/outpost at Selene L1 would be
a good location for storing a few thousand tonnes, and using an
artificial shade should be sufficient for easily avoiding their being
toasted to death.

~ BG

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....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

Nice video here on the high performance Russian engines:

The_Engines_That_Came_In_From_The_Cold.
http://video.google.com/videoplay?do...0537443&hl=en#


Bob Clark

On Jul 16, 11:37*am, Robert Clark wrote:
Nice video here on the high performance Russian engines:

The_Engines_That_Came_In_From_The_Cold.http://video.google.com/videoplay?do...0537443&hl=en#

* * Bob Clark

Anyone know if there has been research on converting the shuttle main
engines to hydrocarbon fueled? I was annoyed that NASA had earlier
canceled a program to develop a heavy-thrust hydrocarbon engine after
the Ares I and V were chosen. We would have a reusable and man-rated
heavy-thrust kerosene engine *now* if it weren't for that.
The SSME's have to operate under severe tolerances using cryogenic
hydrogen since the liquid hydrogen is so cold yet LH2/LOX burns at
such high temperature. I would think using kerosene/LOX for instance
would put less severe conditions on the engine operation.
Note that other liquid hydrogen engines have been successfully run on
other fuels under test conditions:

The RL10 (Bruce Dunn; Gary Hudson; Henry Spencer)
http://yarchive.net/space/rocket/rl10.html

And some dense propellant engines have been tested to run on cryogenic
hydrogen:

LR-87 LH2
http://www.astronautix.com/engines/lr87lh2.htm


Bob Clark

In article 6a5f8191-d605-4b39-aa39-ced18b4b90b4
@y11g2000yqm.googlegroups.com, says...

On Jul 16, 11:37*am, Robert Clark wrote:
Nice video here on the high performance Russian engines:

The_Engines_That_Came_In_From_The_Cold.http://video.google.com/videoplay?do...0537443&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?