R
RGClark
Guest
sublight said:
http://www.youtube.com/watch?v=W15u-JSV4NY
I have no doubt that Armadillo will at some point get it to work. But while performing their test launches they can afford to lose a few thousand dollar vehicle, you can't afford to do that when the vehicle will cost upwards of a million dollars. Better would be to get someone who has done it successfully for decades going back to the Apollo missions, especially when you consider Draper Labs has been willing to participate in X-Prize type ventures at little to no up front payment as proven by their participation in the Google Lunar X-Prize competition.
Bob Clark
R
RGClark
Guest
I saw this discussed on a space oriented forum:
WSJ: Europe Ends Independent Pursuit of Manned Space Travel.
"LE BOURGET, France—Europe appears to have abandoned all hope of
independently pursuing human space exploration, even as the region's
politicians and aerospace industry leaders complain about shrinking
U.S. commitment to various space ventures.
"After years of sitting on the fence regarding a separate, pan-
European manned space program, comments by senior government and
industry officials at the Paris Air Show here underscore that budget
pressures and other shifting priorities have effectively killed that
longtime dream."
http://www.orbiter-forum.com/showthread.php?t=23006
In this post I discussed getting a SSTO by replacing the Vulcain
engine on the Ariane 5 core with a SSME:
Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Wed, 23 Feb 2011 10:14:42 -0800 (PST)
Subject: Re: Some proposals for low cost heavy lift launchers.
http://groups.google.com/group/sci.physics/msg/e1736e7586cc269f?hl=en
However, in point of fact Europe can produce a manned launch vehicle
from currently *existing*, European components. This will consist of
the Ariane 5 and three Vulcain engines. The calculations below use the
Ariane 5 generic "G" version. You might need to add another Vulcain
for the larger evolution "E" version of the Ariane 5 core.
In a following post I'll also show that the Hermes spaceplane also
can become a SSTO by filling the entire fuselage aft of the cockpit
with hydrocarbon propellant.
The impetus for trying the calculation for a Ariane 5 core based SSTO
using Vulcains instead of the SSME was from a report by SpaceX that
you could get the same performance from a planned heavy lift first
stage using a lower performance Merlin 2 compared to the high
performance RS-84 engine. The reason was the lower Isp of the Merlin
was made up for by its lower weight.
THIS IS A VERY IMPORTANT FACT BECAUSE WHAT IT MEANS IS THAT YOU DON'T
NEED THE HIGH PERFORMANCE ENGINES TO GET THE SSTO. YOU CAN USE ENGINES
OF LOWER CHAMBER PRESSURE AND SIMPLER COMBUSTION CYCLES, SUCH AS THE
VULCAIN WITH A CA. 100 BAR COMBUSTION PRESSURE AND A GAS GENERATOR
CYCLE. THIS MEANS THE ENGINES ARE CHEAPER, EASIER TO MAKE REUSABLE,
REQUIRE LESS ROUTINE MAINTENANCE, AND CAN LAST FOR MANY RESTARTS.
In the discussion of the Ariane/Vulcain SSTO below, I note you can
get a prototype, test vehicle quite quickly since the components are
already existing. To improve the payload though you would want to use
altitude compensation on the Vulcains. In a following post I'll
discuss some methods of altitude compensation.
In regards to achieving this at low cost, I think the most important
accomplishment of SpaceX might turn out to be that they showed in
stark terms that privately financed spacecraft, both launchers and
crew capsules, can be accomplished at 1/10th the developmental cost of
government financed ones. Imagine a manned, reusable orbital launcher,
for example, instead of costing, say, $3 billion, only costing $300
million to develop.
Here's how you can get an all European manned SSTO using the Ariane 5
core stage but with Vulcain engines this time. Note that this is one
that can be produced from currently existing components, aside from
the capsule, so at least an unmanned prototype vehicle can be
manufactured and tested in the short term and at lowered development
cost.
We'll use three Vulcain 2's instead of the 1 normally used with the
Ariane 5 core stage. There are varying specifications given on the
Vulcain 2 depending on the source. I'll use the Astronautix site:
Vulcain 2.
http://www.astronautix.com/engines/vulcain2.htm
From the sea level thrust given there, using three Vulcain 2's will
give us one engine out capability. The weight is given as 1,800 kg. So
adding on two will take the dry mass from 12 mT to 15.6 mT.
To calculate the delta-V achieved I'll use the idea again to just use
the vacuum Isp, but adding the loss due to back pressure onto the
delta-V required for orbit, as I discussed previously. However, here
for hydrogen fuel which has higher gravity loss, I'll use a higher
required delta-V of 9,400 m/s when you add on the back pressure loss.
With the vacuum Isp given for the Vulcain 2 of 434 s, we get a payload
of 3.8 mT:
434*9.8ln(1+158/(15.6+3.8)) = 9,412 m/s.
Note this is just using the standard nozzle Isp for the Vulcain, no
altitude compensation. So this could be tested, like, tomorrow.
However, for a SSTO you definitely want to use altitude compensation.
Using engine performance programs such as ProPEP we can calculate that
using long nozzles, you can get a vacuum Isp of 470 s for this engine.
As a point of comparison of how high an Isp you can get even with a
low chamber pressure engine as long as you have a long nozzle, or
equivalent, note that the RL10-B2 with a ca. 250 to 1 area ratio, and
only a ca. 40 bar chamber pressure, gets a 465 s vacuum Isp. So we'll
assume we can get a comparable Isp by using altitude compensation.
This allows us to get payload of 8 mT:
470*9.8ln(1+158/(15.6+8) = 9,400 m/s.
This allows us to add a Dragon-sized capsule and also the reentry and
landing systems to make it reusable.
Bob Clark
WSJ: Europe Ends Independent Pursuit of Manned Space Travel.
"LE BOURGET, France—Europe appears to have abandoned all hope of
independently pursuing human space exploration, even as the region's
politicians and aerospace industry leaders complain about shrinking
U.S. commitment to various space ventures.
"After years of sitting on the fence regarding a separate, pan-
European manned space program, comments by senior government and
industry officials at the Paris Air Show here underscore that budget
pressures and other shifting priorities have effectively killed that
longtime dream."
http://www.orbiter-forum.com/showthread.php?t=23006
In this post I discussed getting a SSTO by replacing the Vulcain
engine on the Ariane 5 core with a SSME:
Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Wed, 23 Feb 2011 10:14:42 -0800 (PST)
Subject: Re: Some proposals for low cost heavy lift launchers.
http://groups.google.com/group/sci.physics/msg/e1736e7586cc269f?hl=en
However, in point of fact Europe can produce a manned launch vehicle
from currently *existing*, European components. This will consist of
the Ariane 5 and three Vulcain engines. The calculations below use the
Ariane 5 generic "G" version. You might need to add another Vulcain
for the larger evolution "E" version of the Ariane 5 core.
In a following post I'll also show that the Hermes spaceplane also
can become a SSTO by filling the entire fuselage aft of the cockpit
with hydrocarbon propellant.
The impetus for trying the calculation for a Ariane 5 core based SSTO
using Vulcains instead of the SSME was from a report by SpaceX that
you could get the same performance from a planned heavy lift first
stage using a lower performance Merlin 2 compared to the high
performance RS-84 engine. The reason was the lower Isp of the Merlin
was made up for by its lower weight.
THIS IS A VERY IMPORTANT FACT BECAUSE WHAT IT MEANS IS THAT YOU DON'T
NEED THE HIGH PERFORMANCE ENGINES TO GET THE SSTO. YOU CAN USE ENGINES
OF LOWER CHAMBER PRESSURE AND SIMPLER COMBUSTION CYCLES, SUCH AS THE
VULCAIN WITH A CA. 100 BAR COMBUSTION PRESSURE AND A GAS GENERATOR
CYCLE. THIS MEANS THE ENGINES ARE CHEAPER, EASIER TO MAKE REUSABLE,
REQUIRE LESS ROUTINE MAINTENANCE, AND CAN LAST FOR MANY RESTARTS.
In the discussion of the Ariane/Vulcain SSTO below, I note you can
get a prototype, test vehicle quite quickly since the components are
already existing. To improve the payload though you would want to use
altitude compensation on the Vulcains. In a following post I'll
discuss some methods of altitude compensation.
In regards to achieving this at low cost, I think the most important
accomplishment of SpaceX might turn out to be that they showed in
stark terms that privately financed spacecraft, both launchers and
crew capsules, can be accomplished at 1/10th the developmental cost of
government financed ones. Imagine a manned, reusable orbital launcher,
for example, instead of costing, say, $3 billion, only costing $300
million to develop.
Here's how you can get an all European manned SSTO using the Ariane 5
core stage but with Vulcain engines this time. Note that this is one
that can be produced from currently existing components, aside from
the capsule, so at least an unmanned prototype vehicle can be
manufactured and tested in the short term and at lowered development
cost.
We'll use three Vulcain 2's instead of the 1 normally used with the
Ariane 5 core stage. There are varying specifications given on the
Vulcain 2 depending on the source. I'll use the Astronautix site:
Vulcain 2.
http://www.astronautix.com/engines/vulcain2.htm
From the sea level thrust given there, using three Vulcain 2's will
give us one engine out capability. The weight is given as 1,800 kg. So
adding on two will take the dry mass from 12 mT to 15.6 mT.
To calculate the delta-V achieved I'll use the idea again to just use
the vacuum Isp, but adding the loss due to back pressure onto the
delta-V required for orbit, as I discussed previously. However, here
for hydrogen fuel which has higher gravity loss, I'll use a higher
required delta-V of 9,400 m/s when you add on the back pressure loss.
With the vacuum Isp given for the Vulcain 2 of 434 s, we get a payload
of 3.8 mT:
434*9.8ln(1+158/(15.6+3.8)) = 9,412 m/s.
Note this is just using the standard nozzle Isp for the Vulcain, no
altitude compensation. So this could be tested, like, tomorrow.
However, for a SSTO you definitely want to use altitude compensation.
Using engine performance programs such as ProPEP we can calculate that
using long nozzles, you can get a vacuum Isp of 470 s for this engine.
As a point of comparison of how high an Isp you can get even with a
low chamber pressure engine as long as you have a long nozzle, or
equivalent, note that the RL10-B2 with a ca. 250 to 1 area ratio, and
only a ca. 40 bar chamber pressure, gets a 465 s vacuum Isp. So we'll
assume we can get a comparable Isp by using altitude compensation.
This allows us to get payload of 8 mT:
470*9.8ln(1+158/(15.6+8) = 9,400 m/s.
This allows us to add a Dragon-sized capsule and also the reentry and
landing systems to make it reusable.
Bob Clark
- Joined
- 13 August 2007
- Messages
- 7,173
- Reaction score
- 6,574
it nice to see how expensive high tech is all-ways used for SSTO
but wat about low cost solution like Pressure fed Engine ?
I know on fist look it's contradictory because lower performance to SSME or Vulcan
but from cost it's far cheaper
in 1965 to 1969 the USAF look in to minimum cost Design (MCD) as replacement for Titan IIIC
goal bring 45 K pound (21 Tons) in Low orbit with rock-bottom cost
the Rockwell proposal was a two stage rocket used N2O4/UDMH fuels with TRW pressure Fed engine
they evaluate that Unit build and launch cost at total US$42 million (in year 2011 value) at 12 launch/year
drop second stage, reduce payload from 45K to 10k change fuel to Lox/RP-1
a SSTO with US$42 million that is cheaper as a Falcon9
but wat about low cost solution like Pressure fed Engine ?
I know on fist look it's contradictory because lower performance to SSME or Vulcan
but from cost it's far cheaper
in 1965 to 1969 the USAF look in to minimum cost Design (MCD) as replacement for Titan IIIC
goal bring 45 K pound (21 Tons) in Low orbit with rock-bottom cost
the Rockwell proposal was a two stage rocket used N2O4/UDMH fuels with TRW pressure Fed engine
they evaluate that Unit build and launch cost at total US$42 million (in year 2011 value) at 12 launch/year
drop second stage, reduce payload from 45K to 10k change fuel to Lox/RP-1
a SSTO with US$42 million that is cheaper as a Falcon9
R
RGClark
Guest
Michel Van said:
The DIRECT team has proposed a two-stage heavy lift launcher with pressure fed stages only:
DIRECT Project 2 - Leviathan Heavy Lift Launch Vehicle.
http://www.launchcomplexmodels.com/DirectP2/leviathan140.html
And this proposal was for many small, cheap, pressure-fed SSTO's to transport cargo to orbit:
Aquarius Launch Vehicle.
http://en.wikipedia.org/wiki/Aquarius_Launch_Vehicle
Bob Clark
R
RGClark
Guest
In post #66, I argued that small, low cost SSTO's are doable
now using lightweight design and high efficiency engines. However,
I was surprised to find after doing the calculation you don't even need
the high efficiency engines to get the SSTO. The low efficiency SpaceX
Merlin engines would be sufficient for example, IF you have altitude compensation.
The impetus for trying the calculation was from a report by SpaceX
that you could get the same performance from a planned heavy lift
first stage using a lower performance Merlin 2 engine compared to the
high performance RS-84 engine. The reason was the lower Isp of the
Merlin was made up for by its lower weight:
SpaceX Propulsion.
http://images.spaceref.com/news/2010/SpaceX_Propulsion.pdf
Now note that the biggest single contributor to the vacuum Isp of an
engine is not the chamber pressure, but the nozzle length. For
example, the Merlin Vacuum raises its vacuum Isp to 342 s from the 304
s Isp of the Merlin 1C by having a longer nozzle, even though the
chamber pressure remains the same, ca. 100 bar.
So I'll redo the calculation for the SSTO using the SpaceX Falcon 1
first stage but using Merlin engines this time. We'll assume that
using altitude compensation we are able to get an engine with the same
vacuum Isp as the Merlin Vacuum but able to launch from ground.
We'll use the soon to be introduced Merlin 1D:
SpaceX Plans To Be Top World Rocket Maker.
Aug 11, 2011
By Guy Norris
San Diego
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2011/08/08/AW_08_08_2011_p27-354586.xml&headline=SpaceX%20Plans%20To%20Be%20Top %20World%20Rocket%20Maker&channel=defense
Using the 160 to 1 thrust/weight ratio and 155,000 lbs. vacuum thrust
given, it has a mass of 970 lbs., 440 kg. However, this would make it
overpowered for the Falcon 1 first stage only. So we'll use two copies
of this stage powered by a single Merlin 1D.
The original Falcon 1 first stage with the Merlin 1C engine has a dry mass
of 1,360 kg. I estimated the mass of the Merlin 1C in the prior post to
be 650 kg. So without the engine, the stage weighs 710 kg. So two of
them will be 1,420 kg without engines, and adding on the Merlin 1D
engine gives this a mass of 1,860 kg.
The propellant mass of the two copies of the first stage is 43,080
kg. Then to calculate the payload that can be carried I'll again just
use the vacuum Isp and take the required delta-V as 9,150 m/s. We
conclude a payload of 1,140 kg can be lofted:
342*9.8ln(1 + 43,080/(1,860 + 1,140)) = 9,160.
Now we'll estimate how much the payload can be if we use a higher
energy density fuel such as methylacetylene and use lightweight
composites for the stage. I'll get a rough idea how high the Isp can
be for this case by assuming it is increased proportionally to the
same degree as for the high efficiency engine case. That is, using
methylacetylene in the high efficiency case resulted in increasing the
vacuum Isp to 384 s from the 360 s vacuum Isp for the kerosene.
Assuming the vacuum Isp will be increased to the same proportion here
gives us a vacuum Isp of 365 s for methylacetylene and the Merlin 1D
engine.
For the reduced stage weight using composites, assume again it will
be reduced by 40% aside from the engines. Then the stage weight with
the Merlin 1D engine will be .6*1,420 + 440 kg = 1,290 kg. Then will
be able to loft a payload of 2,320 kg:
365*9.8ln(1 + 43,080/(1,290 + 2,320)) = 9,160 m/s.
Also, quite likely SpaceX could make a half-size version of the
Merlin 1D engine. So you could use a single copy of the Falcon 1 first
stage. Then the payload would be approximately cut in half, 570 kg for
the kerosene/standard stage version and 1,160 kg for the
methylacetylene/composite stage version.
Note that low chamber pressure, low performance engines can also be
used to power the SSTO's is extremely important. Such engines have
less complicated combustion cycles and have to withstand much less
strenuous operating regimes. This makes them cheaper, simpler, easier
to maintain, and easier to make reusable. So the most costly component
of any rocket, the engines, become markedly cheaper for the proposed
SSTO.
What is key though is to come up with ways to get the needed altitude
compensation without adding on too much to the engine weight. In a
following post I'll discuss some methods this might be accomplished.
Bob Clark
now using lightweight design and high efficiency engines. However,
I was surprised to find after doing the calculation you don't even need
the high efficiency engines to get the SSTO. The low efficiency SpaceX
Merlin engines would be sufficient for example, IF you have altitude compensation.
The impetus for trying the calculation was from a report by SpaceX
that you could get the same performance from a planned heavy lift
first stage using a lower performance Merlin 2 engine compared to the
high performance RS-84 engine. The reason was the lower Isp of the
Merlin was made up for by its lower weight:
SpaceX Propulsion.
http://images.spaceref.com/news/2010/SpaceX_Propulsion.pdf
Now note that the biggest single contributor to the vacuum Isp of an
engine is not the chamber pressure, but the nozzle length. For
example, the Merlin Vacuum raises its vacuum Isp to 342 s from the 304
s Isp of the Merlin 1C by having a longer nozzle, even though the
chamber pressure remains the same, ca. 100 bar.
So I'll redo the calculation for the SSTO using the SpaceX Falcon 1
first stage but using Merlin engines this time. We'll assume that
using altitude compensation we are able to get an engine with the same
vacuum Isp as the Merlin Vacuum but able to launch from ground.
We'll use the soon to be introduced Merlin 1D:
SpaceX Plans To Be Top World Rocket Maker.
Aug 11, 2011
By Guy Norris
San Diego
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2011/08/08/AW_08_08_2011_p27-354586.xml&headline=SpaceX%20Plans%20To%20Be%20Top %20World%20Rocket%20Maker&channel=defense
Using the 160 to 1 thrust/weight ratio and 155,000 lbs. vacuum thrust
given, it has a mass of 970 lbs., 440 kg. However, this would make it
overpowered for the Falcon 1 first stage only. So we'll use two copies
of this stage powered by a single Merlin 1D.
The original Falcon 1 first stage with the Merlin 1C engine has a dry mass
of 1,360 kg. I estimated the mass of the Merlin 1C in the prior post to
be 650 kg. So without the engine, the stage weighs 710 kg. So two of
them will be 1,420 kg without engines, and adding on the Merlin 1D
engine gives this a mass of 1,860 kg.
The propellant mass of the two copies of the first stage is 43,080
kg. Then to calculate the payload that can be carried I'll again just
use the vacuum Isp and take the required delta-V as 9,150 m/s. We
conclude a payload of 1,140 kg can be lofted:
342*9.8ln(1 + 43,080/(1,860 + 1,140)) = 9,160.
Now we'll estimate how much the payload can be if we use a higher
energy density fuel such as methylacetylene and use lightweight
composites for the stage. I'll get a rough idea how high the Isp can
be for this case by assuming it is increased proportionally to the
same degree as for the high efficiency engine case. That is, using
methylacetylene in the high efficiency case resulted in increasing the
vacuum Isp to 384 s from the 360 s vacuum Isp for the kerosene.
Assuming the vacuum Isp will be increased to the same proportion here
gives us a vacuum Isp of 365 s for methylacetylene and the Merlin 1D
engine.
For the reduced stage weight using composites, assume again it will
be reduced by 40% aside from the engines. Then the stage weight with
the Merlin 1D engine will be .6*1,420 + 440 kg = 1,290 kg. Then will
be able to loft a payload of 2,320 kg:
365*9.8ln(1 + 43,080/(1,290 + 2,320)) = 9,160 m/s.
Also, quite likely SpaceX could make a half-size version of the
Merlin 1D engine. So you could use a single copy of the Falcon 1 first
stage. Then the payload would be approximately cut in half, 570 kg for
the kerosene/standard stage version and 1,160 kg for the
methylacetylene/composite stage version.
Note that low chamber pressure, low performance engines can also be
used to power the SSTO's is extremely important. Such engines have
less complicated combustion cycles and have to withstand much less
strenuous operating regimes. This makes them cheaper, simpler, easier
to maintain, and easier to make reusable. So the most costly component
of any rocket, the engines, become markedly cheaper for the proposed
SSTO.
What is key though is to come up with ways to get the needed altitude
compensation without adding on too much to the engine weight. In a
following post I'll discuss some methods this might be accomplished.
Bob Clark
- Joined
- 9 October 2009
- Messages
- 19,999
- Reaction score
- 10,544
- Joined
- 13 August 2007
- Messages
- 7,173
- Reaction score
- 6,574
There many SSTO proposal with exotic Air breathing Rocket engine
Those need billions in R&D like HOTOL Hardware
but GNOM is unique, it use simple Ramjet from Mach 1.75 to Mach 5.5
Those need billions in R&D like HOTOL Hardware
but GNOM is unique, it use simple Ramjet from Mach 1.75 to Mach 5.5
R
RGClark
Guest
The increasing problem of space junk is getting greater discussion recently:
Space junk at 'tipping point', now getting worse on its own.
More collisions generate more debris, so more collisions.
By Gavin Clarke
Posted in Space, 2nd September 2011 11:18 GMT
http://www.theregister.co.uk/2011/09/02/space_junk_danger/
One company is planning on reusable in-space vehicles to refuel and service satellites. The Air Force favors this since this may also be used to tow inactive satellites out the way of operable satellites, thus reducing the problem of space debris:
Article:
World's First Space Gas Station for Satellites to Launch in 2015.
by Clara Moskowitz, SPACE.com Senior Writer
Date: 15 March 2011 Time: 06:03 PM ET
Then the Air Force recognizes the usefulness of reusable vehicles, when they are in-space. However, the importance of reusable SSTO's is that they could also return these satellites to Earth for repair or salvage.
Remember the old science fiction series Salvage 1:
http://en.wikipedia.org/wiki/Salvage_1
The theme of the show was a small "home made" spaceship was used to return space junk to Earth. I used to think the show was quite implausible because the spaceship went all the way to the Moon and everybody "knows" it takes huge Saturn V sized rockets to do that.
However, as I discussed in post #66 small, low cost SSTO's are indeed possible. And it is a known fact that if you have refueling in LEO then an SSTO can go all the way to the Moon, land, take off, and return to Earth on that one single refueling. So in fact the idea of salvaging spacecraft or satellites from the Moon and/or from GEO is feasible with SSTO's and on-orbit propellant depots.
Then this provides another financial benefit for SSTO's for private developers and for the Air Force. Imagine being able to retrieve satellites, the largest of which can cost upwards of a billion dollars, for reuse or possibly for sale. This does though raise the question of what would be the salvage laws for space.
Bob Clark
Space junk at 'tipping point', now getting worse on its own.
More collisions generate more debris, so more collisions.
By Gavin Clarke
Posted in Space, 2nd September 2011 11:18 GMT
http://www.theregister.co.uk/2011/09/02/space_junk_danger/
One company is planning on reusable in-space vehicles to refuel and service satellites. The Air Force favors this since this may also be used to tow inactive satellites out the way of operable satellites, thus reducing the problem of space debris:
Article:
World's First Space Gas Station for Satellites to Launch in 2015.
by Clara Moskowitz, SPACE.com Senior Writer
Date: 15 March 2011 Time: 06:03 PM ET
http://www.space.com/11135-satellite-refueling-mission-space-debris.html
Then the Air Force recognizes the usefulness of reusable vehicles, when they are in-space. However, the importance of reusable SSTO's is that they could also return these satellites to Earth for repair or salvage.
Remember the old science fiction series Salvage 1:
http://en.wikipedia.org/wiki/Salvage_1
The theme of the show was a small "home made" spaceship was used to return space junk to Earth. I used to think the show was quite implausible because the spaceship went all the way to the Moon and everybody "knows" it takes huge Saturn V sized rockets to do that.
However, as I discussed in post #66 small, low cost SSTO's are indeed possible. And it is a known fact that if you have refueling in LEO then an SSTO can go all the way to the Moon, land, take off, and return to Earth on that one single refueling. So in fact the idea of salvaging spacecraft or satellites from the Moon and/or from GEO is feasible with SSTO's and on-orbit propellant depots.
Then this provides another financial benefit for SSTO's for private developers and for the Air Force. Imagine being able to retrieve satellites, the largest of which can cost upwards of a billion dollars, for reuse or possibly for sale. This does though raise the question of what would be the salvage laws for space.
Bob Clark
R
RGClark
Guest
Just saw this:
The SpaceX
Falcon Heavy Booster: Why Is It Important?
by John K. Strickland, Jr.
September, 2011
"What amazes people is that SpaceX has broken the long-sought 1,000
dollars a pound to orbit price barrier with a rocket which is still
expendable. 'How can he (SpaceX CEO Elon Musk) possibly do this?' they
ask. The Chinese have said flatly that there is no way they can
compete with such a low price. It is important to remember that this
was not done in a single step. The Falcon 9 already has a large price
advantage over other boosters, even though it does not have the
payload capacity of some of the largest ones. The 'Heavy' will even
this score and then some. At last count, SpaceX had a launch manifest
of over 40 payloads, far exceeding any current government contracts,
with more being added every month. These are divided between the
Falcon 9 and the Falcon Heavy."
http://www.nss.org/articles/falconheavy.html
I think the most important accomplishment of SpaceX might turn out
to be that they showed in stark terms that privately financed spacecraft,
both launchers and crew capsules, can be accomplished at 1/10th the
developmental cost of government financed ones. Imagine a manned,
reusable orbital launcher, for example, instead of costing, say, $3 billion,
only costing $300 million to develop.
As I argue, the key variable that made this reduction possible is that
the launcher was privately financed. That is, it was the launch
company's own money that was financing its development. In that case
it makes sense the company would be more fiscally responsible in
developing it.
Then the first step in reducing the price to orbit is making the
vehicles be privately financed. But if the launch companies are going
to spend their own money, they have to be convinced they can make a
profit on them. This will come if there is a significant market.
My view is that there would be a significant market for small,
privately owned, SSTO's. When you consider that with orbital refueling
such craft can also make lunar missions, the market becomes even more
apparent.
An additional finance stream of such vehicles that would make them
marketable I argue could be salvage of satellites in LEO or GEO.
Bob Clark
The SpaceX
Falcon Heavy Booster: Why Is It Important?
by John K. Strickland, Jr.
September, 2011
"What amazes people is that SpaceX has broken the long-sought 1,000
dollars a pound to orbit price barrier with a rocket which is still
expendable. 'How can he (SpaceX CEO Elon Musk) possibly do this?' they
ask. The Chinese have said flatly that there is no way they can
compete with such a low price. It is important to remember that this
was not done in a single step. The Falcon 9 already has a large price
advantage over other boosters, even though it does not have the
payload capacity of some of the largest ones. The 'Heavy' will even
this score and then some. At last count, SpaceX had a launch manifest
of over 40 payloads, far exceeding any current government contracts,
with more being added every month. These are divided between the
Falcon 9 and the Falcon Heavy."
http://www.nss.org/articles/falconheavy.html
I think the most important accomplishment of SpaceX might turn out
to be that they showed in stark terms that privately financed spacecraft,
both launchers and crew capsules, can be accomplished at 1/10th the
developmental cost of government financed ones. Imagine a manned,
reusable orbital launcher, for example, instead of costing, say, $3 billion,
only costing $300 million to develop.
As I argue, the key variable that made this reduction possible is that
the launcher was privately financed. That is, it was the launch
company's own money that was financing its development. In that case
it makes sense the company would be more fiscally responsible in
developing it.
Then the first step in reducing the price to orbit is making the
vehicles be privately financed. But if the launch companies are going
to spend their own money, they have to be convinced they can make a
profit on them. This will come if there is a significant market.
My view is that there would be a significant market for small,
privately owned, SSTO's. When you consider that with orbital refueling
such craft can also make lunar missions, the market becomes even more
apparent.
An additional finance stream of such vehicles that would make them
marketable I argue could be salvage of satellites in LEO or GEO.
Bob Clark
R
RGClark
Guest
At the National Press Club, Elon Musk discussed SpaceX's plans for a fully reusable launcher:
http://www.spacex.com/npc-luncheon-elon-musk.php
A couple of suggestions for this reusable version of the Falcon 9.
First, model it on the DC-X. In the SpaceX video of the proposed
reusable launcher the first and second stages have the same straight
sides of the expendable versions. But having sloping sides helps to
protect the sides of the vehicle during reentry as well as increasing
aerodynamic stability during reentry.
Note that as long as the cross-section remains circular for a conical
shaped stage you should still get the high tankage ratio that obtains
for cylindrical tanks:
Space Access Update #91 2/7/00.
The Last Five Years: NASA Gets Handed The Ball, And Drops It.
"...part of L-M X-33's weight growth was the "multi-
lobed" propellant tanks growing considerably heavier than promised.
Neither Rockwell nor McDonnell-Douglas bid these; both used proven
circular-section tanks. X-33's graphite-epoxy "multi-lobed" liquid
hydrogen tanks have ended up over twice as heavy relative to the
weight of propellant carried as the Shuttle's 70's vintage aluminum
circular-section tanks - yet an X-33 tank still split open in test
last fall. Going over to aluminum will make the problem worse; X-
33's aluminum multi-lobed liquid oxygen tank is nearly four times as
heavy relative to the weight of propellant carried as Shuttle's
aluminum circular-section equivalent."
http://www.space-access.org/updates/sau91.html
The McDonnell-Douglas version mentioned there was the scaled up DC-X.
There are a couple of ways this DC-X styled Falcon 9 could be
implemented. As this is to be a multi-stage launcher, you could have
each stage have the same sloping sides as the DC-C. Then each stage
would have the shape of a truncated cone, a frustum, and when stacked
one on top another the vehicle would have the shape of a single cone.
However, I prefer another method. It is known that you can increase
your payload using parallel staging with cross-feed fueling. Indeed
SpaceX intends to increase the payload of its Falcon Heavy launcher
using this method. Then another method for this reusable Falcon 9
would have each stage in the shape of a full cone, but the second
stage instead of being placed on top of the first stage would be
placed along side of it in parallel fashion.
In addition to increasing the payload this would have an another key
advantage. The high mass ratio of the Falcon 9 first stage, above 20
to 1, means that if it had high efficiency engines such as the NK-33
or RD-180 instead of the rather low efficiency Merlin 1C it would have
SSTO capability. However, because of the high investment of SpaceX in
the Merlin engines they no doubt are committed to its use.
But a key fact is that IF you have altitude compensation then even a
low efficiency, i.e., low chamber pressure, engine can achieve high
vacuum Isp while still providing good performance at sea level.
Methods of altitude compensation such as the aerospike have been
studied since the 60's. Then SpaceX could provide their DC-X styled
Falcon 9 stages with altitude compensation to give their stages SSTO
capability while still using the Merlin engines.
Then these SSTO stages could serve as low cost launchers for smaller
payloads, including being used for private, manned orbital vehicles.
The second model for the reusable Falcon 9 stages would be on the
ESA's proposed Intermediate eXperimental Vehicle (IXV):
Article:
Europe Aims to Launch Robotic Mini-Shuttle By 2020.
Rob Coppinger, SPACE.com ContributorDate: 13 June 2011 Time: 02:58 PM ET
http://www.space.com/11948-robot-space-plane-europe-ixv-launching-2020.html
This does not use the powered landing of the DC-X but rather uses a
glided landing via its lifting body shape. SpaceX does not like the
use of wings for landing because of the extra weight. But this design
would not have wings. It would have larger thermal protection weight
because the horizontal underside would have to be covered, whereas in
the DC-X mode only the base has to be covered. However, it would make
up for this in not requiring fuel for the powered landing.
In this case because the stages would have to maintain the aerodynamic
shape, they could not be stacked as for serial staging. Parallel
staging would have to be used. Once again this means the separate
stages could be used as SSTO's.
Bob Clark
http://www.spacex.com/npc-luncheon-elon-musk.php
A couple of suggestions for this reusable version of the Falcon 9.
First, model it on the DC-X. In the SpaceX video of the proposed
reusable launcher the first and second stages have the same straight
sides of the expendable versions. But having sloping sides helps to
protect the sides of the vehicle during reentry as well as increasing
aerodynamic stability during reentry.
Note that as long as the cross-section remains circular for a conical
shaped stage you should still get the high tankage ratio that obtains
for cylindrical tanks:
Space Access Update #91 2/7/00.
The Last Five Years: NASA Gets Handed The Ball, And Drops It.
"...part of L-M X-33's weight growth was the "multi-
lobed" propellant tanks growing considerably heavier than promised.
Neither Rockwell nor McDonnell-Douglas bid these; both used proven
circular-section tanks. X-33's graphite-epoxy "multi-lobed" liquid
hydrogen tanks have ended up over twice as heavy relative to the
weight of propellant carried as the Shuttle's 70's vintage aluminum
circular-section tanks - yet an X-33 tank still split open in test
last fall. Going over to aluminum will make the problem worse; X-
33's aluminum multi-lobed liquid oxygen tank is nearly four times as
heavy relative to the weight of propellant carried as Shuttle's
aluminum circular-section equivalent."
http://www.space-access.org/updates/sau91.html
The McDonnell-Douglas version mentioned there was the scaled up DC-X.
There are a couple of ways this DC-X styled Falcon 9 could be
implemented. As this is to be a multi-stage launcher, you could have
each stage have the same sloping sides as the DC-C. Then each stage
would have the shape of a truncated cone, a frustum, and when stacked
one on top another the vehicle would have the shape of a single cone.
However, I prefer another method. It is known that you can increase
your payload using parallel staging with cross-feed fueling. Indeed
SpaceX intends to increase the payload of its Falcon Heavy launcher
using this method. Then another method for this reusable Falcon 9
would have each stage in the shape of a full cone, but the second
stage instead of being placed on top of the first stage would be
placed along side of it in parallel fashion.
In addition to increasing the payload this would have an another key
advantage. The high mass ratio of the Falcon 9 first stage, above 20
to 1, means that if it had high efficiency engines such as the NK-33
or RD-180 instead of the rather low efficiency Merlin 1C it would have
SSTO capability. However, because of the high investment of SpaceX in
the Merlin engines they no doubt are committed to its use.
But a key fact is that IF you have altitude compensation then even a
low efficiency, i.e., low chamber pressure, engine can achieve high
vacuum Isp while still providing good performance at sea level.
Methods of altitude compensation such as the aerospike have been
studied since the 60's. Then SpaceX could provide their DC-X styled
Falcon 9 stages with altitude compensation to give their stages SSTO
capability while still using the Merlin engines.
Then these SSTO stages could serve as low cost launchers for smaller
payloads, including being used for private, manned orbital vehicles.
The second model for the reusable Falcon 9 stages would be on the
ESA's proposed Intermediate eXperimental Vehicle (IXV):
Article:
Europe Aims to Launch Robotic Mini-Shuttle By 2020.
Rob Coppinger, SPACE.com ContributorDate: 13 June 2011 Time: 02:58 PM ET
http://www.space.com/11948-robot-space-plane-europe-ixv-launching-2020.html
This does not use the powered landing of the DC-X but rather uses a
glided landing via its lifting body shape. SpaceX does not like the
use of wings for landing because of the extra weight. But this design
would not have wings. It would have larger thermal protection weight
because the horizontal underside would have to be covered, whereas in
the DC-X mode only the base has to be covered. However, it would make
up for this in not requiring fuel for the powered landing.
In this case because the stages would have to maintain the aerodynamic
shape, they could not be stacked as for serial staging. Parallel
staging would have to be used. Once again this means the separate
stages could be used as SSTO's.
Bob Clark
R
RGClark
Guest
Other possible methods to make the Falcon 9 reusable might be to use the "parashield" idea of the Dr. David Akin or the inflatable heat shield NASA is investigating. These might make the reusable Falcon 9 easier and quicker to implement since the usual cylindrical shaped stages could be used:
Phoenix: A Low-Cost Commercial Approach to the Crew Exploration Vehicle.
http://www.nianet.org/rascal/forum2006/presentations/1010_umd_paper.pdf
"Figure 5.9-1: Phoenix ParaShield in stowed and deployed configurations."
Another advantage of the parashield is that it can also serve as a parachute once the vehicle has passed through reentry.
And for NASA's inflatable heat shield:
NASA Launches New Technology: An Inflatable Heat Shield.
UPDATE: 08.17.09
http://www.nasa.gov/topics/aeronautics/features/irve.html
See the video on this page describing the inflatable heat shield.
Bob Clark
Phoenix: A Low-Cost Commercial Approach to the Crew Exploration Vehicle.
http://www.nianet.org/rascal/forum2006/presentations/1010_umd_paper.pdf
"Figure 5.9-1: Phoenix ParaShield in stowed and deployed configurations."
Another advantage of the parashield is that it can also serve as a parachute once the vehicle has passed through reentry.
And for NASA's inflatable heat shield:
NASA Launches New Technology: An Inflatable Heat Shield.
UPDATE: 08.17.09
http://www.nasa.gov/topics/aeronautics/features/irve.html
See the video on this page describing the inflatable heat shield.
Bob Clark
archipeppe
ACCESS: Top Secret
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Umbrella heat-shield capsule, studied and tested recently (with the SCIROCCO hypersonic facility at CIRA, Capua near Naples) by ASI (Italian Space Agency):
http://www.projectirene.com/
http://www.projectirene.com/
R
RGClark
Guest
archipeppe said:
Thanks for that. These could probably work just as well to make the Ariane 5 stages reusable.
Bob Clark
archipeppe
ACCESS: Top Secret
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- Reaction score
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RGClark said:
Frankly speaking I don't know.
The aim of IRENE project is to reduce heatshield volume and mass at launch for small unmanned spacecraft not to allow huge Ariane 5 stage to be eventually recovered on ground...
R
RGClark
Guest
archipeppe said:
In dry mass the core stage is about the same dry mass as the Falcon 9 first stage. The IRENE system also appears similar to NASA's inflatable IRVE heat shield system.
So IF this can work for the Falcon 9 it will probably also work for the Ariane 5.
Bob Clark
R
RGClark
Guest
Byeman said:
The reference was to the post of Oct. 6 that discussed using the inflatable heat shield of NASA for the Falcon 9 stages. As described on the page www.projectirene.com, the IRENE heat shield is of similar design to the NASA one. So if it works for the Falcon 9 stages, likely it would also work for the Ariane 5 stages.
Bob Clark
R
RGClark
Guest
RGClark said:
BTW, I think the payload lost in making the vehicle reusable is being overstated. Elon himself during the speech spoke ruefully of cutting into the 2%-3% payload fraction of launch vehicles. But actually a small percentage of the vehicle's dry weight, which is the important parameter not the gross weight, would need to go the reentry/landing systems.
The reason is this is for a multi-stage launcher, and a key fact is for the larger first stage any extra kilo added to the first stage dry weight subtracts only ca. 1/10th of a kilo from the payload.
And also for multi-stage launchers, the upper stage dry weight is usually rather small, in fact frequently smaller than the payload.
We can estimate the added weight for the Falcon 9. This page estimates the weights for this launcher:
Space Launch Report: SpaceX Falcon Data Sheet.
http://www.spacelaunchreport.com/falcon9.html#components
The dry weight for the first stage is given as 19 mT, and 3 mT for the upper stage. These weights might even be overestimated. Some references for instance give the dry weight for the first stage as in the range of 15 mT.
Now estimate the mass of reentry/landing systems. First, Robert Zubrin gives an estimate of about 15% of the landed weight for reentry thermal protection:
Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_launch_system#Reentry_heat_shields
Secondly, an estimate of 10% is often cited for the wings for glided landing or for the fuel for powered landing:
Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_launch_system#Horizontal_takeoff
Finally the estimated weight for the landing gear is about 3%:
Landing gear weight.
http://yarchive.net/space/launchers/landing_gear_weight.html
This totals to 28%. However, it is important to keep in mind that with modern materials this can probably be reduced to half this.
So 14% of 19 mT on the first stage is 2,660 kg. But remember for a first stage this will only subtract about 1/10th this from the payload. So 270 kg lost.
For the second stage 14% of 3 mT is 420 kg. So the total is in the range of 700 kg lost from the Falcon 9 payload capacity to LEO of 10,000 kg.
But by doing this you are making the vehicle reusable and cutting costs by a factor of 100.
Bob Clark
R
RGClark
Guest
RGClark said:
Another possible lifting-body shape for reusable Falcon 9 stages might be of the Japanese HYFLEX hypersonic test vehicle:
Hypersonic Flight Experiment "HYFLEX".
http://www.jaxa.jp/projects/rockets/hyflex/index_e.html
HYFLEX.
http://www.rocket.jaxa.jp/fstrc/0c02.html
This was successfully tested all the way back in 1996 at a Mach 15 reentry speed.
It's roughly cylindrical shape would mean you would lose a relatively small degree on the mass efficiency of cylindrically shaped tanks. However, rather than redesigning the tanks you might want to just use a composite aeroshell on the usual Falcon 9 stages. A conical aeroshell for example was used on the DC-X.
This would make the reusable Falcon 9 more quickly and easily to be implemented. The mass of the aeroshell though would contribute to the mass lost from payload.
As with the above cases, used with altitude compensating nozzles on the Merlins or with existing high efficiency engines with just their standard nozzles, these HYFLEX-shaped stages could also be SSTO's.
Bob Clark
R
RGClark
Guest
RGClark said:
Space vehicle launches could be routine if they could take off horizontally from airliner runways as a single stage, like aircraft. It was thought that wings would just be dead weight on ascent but in fact following a lifting trajectory can cut in the range of 40% off the propellant requirements from a SSTO if it is at high lift/drag ratio. So wings can "carry their own weight", so to speak even on ascent.
Here's a heuristic argument that an SSTO making a lifting trajectory at high L/D ratio can save on propellant requirements. I'll regard the straight-line path as my X-axis and the perpendicular to this as the Y-axis. Note this means my axes look like they are at an angle to the usual horizontal and vertical axes, but it makes the calculation easier. Call the thrust T, and the mass, M. Then the force component along the straight-line path, our X-axis, is Fx = T - gMsin(θ) - D and the force
component along the Y-axis is Fy = L - gMcos(θ).
We'll set L = gMcos(θ), since the vehicle is traveling along the straight-line, our X-axis, so the force component in the Y-direction is zero. Then the force along the straight-line is Fx = T -gMsin(θ) - gMcos(θ)/(L/D). As with the calculation for the usual rocket equation, divide this by M to get the acceleration along this line, and integrate to get the velocity. The result is V(t) = Ve*ln(M0/Mf) - g*tsin(θ) - g*tcos(θ)/(L/D), with M0 the initial mass, and Mf, the mass at time t, a la the rocket equation. If you make the angle θ (theta) be shallow, the g*tsin(θ) term will be smaller than the usual gravity drag loss of g*t and the (L/D) divisor will make the cosine term smaller as well.
Now note that the equation includes *both* the gravity and air drag. Secondly, note that though using aerodynamic lift generates additional, large, induced drag, this is covered by the fact that the L/D ratio includes this induced drag, since it involves the *total* drag.
Some preliminary calculations I did suggest you could save in the range of 40% off your propellant requirements by reducing the gravity loss in this fashion if indeed your L/D ratio is 7+, at the speed range up to the high supersonic to low hypersonic.
A reduction in the propellant requirements this high means you could carry significant payload to orbit even with standard wing weight, estimated as 10% of the vehicle weight, which would be the gross takeoff weight for a horizontally launched vehicle.
However, we can cut this weight even further. First, even for standard wings you could cut this wing weight by half with modern materials. But a key fact is that you don't even need wings for a horizontal liftoff. Lifting bodies can perform horizontal takeoffs without additional wings. Lifting bodies have long been investigated for use as reentry vehicles, but it seems to have been overlooked to use them also for horizontal takeoff.
The problem though would be to get mass efficient propellant tanks for the vehicle. The propellant tanks are frequently the heaviest component of the dry weight of the vehicle, even more than the engines.
But the lifting body would have non-cylindrical shape. What killed the X-33/VentureStar program was the inability to get light-weight tanks for the X-33's conformally shaped tanks. Even using composite materials the non-circular cross sections would have resulted in tanks twice as heavy as aluminum tanks of usual cylindrical shape [1].
It turns out the determining factor for the heavy weight of rocket propellant tanks is they have to be pressurized. This is because of the requirements for proper operation of the turbopumps on rocket engines [2]. Pressurized tanks have to have a certain thickness to safely hold the contents.
However, quite key is the fact that this is a requirement of turbopumps but not other types of pumps [3]. XCOR because it's using wing tanks for its Lynx suborbital vehicle plans to use reciprocating (piston) pumps [4]. These require reduced pressurization in the tanks, if any. And the tanks in the wings on aircraft commonly are also not pressurized. XCOR had to develop these since they intend to carry the propellant in wing tanks which being non-circular would have been very heavy if they had to be pressurized.
Another possibility would be to use inflatable wings. These can also save weight over standard wings [5].
Bob Clark
1.)Space Access Update #91 2/7/00.
The Last Five Years: NASA Gets Handed The Ball, And Drops It.
"...part of L-M X-33's weight growth was the "multi-
lobed" propellant tanks growing considerably heavier than promised.
Neither Rockwell nor McDonnell-Douglas bid these; both used proven
circular-section tanks. X-33's graphite-epoxy "multi-lobed" liquid
hydrogen tanks have ended up over twice as heavy relative to the
weight of propellant carried as the Shuttle's 70's vintage aluminum
circular-section tanks - yet an X-33 tank still split open in test
last fall. Going over to aluminum will make the problem worse; X-
33's aluminum multi-lobed liquid oxygen tank is nearly four times as
heavy relative to the weight of propellant carried as Shuttle's
aluminum circular-section equivalent."
http://www.space-access.org/updates/sau91.html
2.)Fuel tank scaling laws (Henry Spencer).
http://yarchive.net/space/launchers/fuel_tank_scaling_laws.html
3.)NPSH.
http://en.wikipedia.org/wiki/NPSH
4.)XCOR Aerospace and United Launch Alliance.
Announce Successful Hydrogen Piston Pump Tests.
http://www.xcor.com/press-releases/2010/10-06-08_ULA_and_XCOR_announce_successful_hydrogen_pumping_tests.html
5.)An inflatable wing using the principle of Tensairity.
http://www.empa.ch/plugin/template/empa/*/107170
RGClark said:
A commonly held belief that is essentially a religious belief rather than a scientific one.
R
RGClark
Guest
RGClark said:
Armadillo Aerospace has now succeeded in launching their rocket to high altitude:
Armadillo Launches STIG A Rocket to 137,500 Feet From Spaceport America
Posted by Doug Messier on December 6, 2011, at 12:31 pm in News.
http://www.parabolicarc.com/2011/12/06/armadillo-launches-stiga-a-rocket-to-124000-feet-from-spaceport-america/
http://www.youtube.com/watch?v=yOT2am5VWlk
Congrats to Armadillo and John Carmack. They apparently have solved the problem of instability at high velocity. No doubt getting to the full altitude for space at 100 km will come in short order as well. B)
Bob Clark
R
RGClark
Guest
Altitude compensation on the Falcon 9 first stage and applications.
In post #86 in this thread is an argument for how two copies of the Falcon 1 first stage combined together using a single Merlin engine with an altitude compensating nozzle could be SSTO. I had forgotten though that the first stage of the Falcon 1e, which SpaceX wants to move to for small launches anyway, has about twice the propellant load of the Falcon 1's first stage so it could form the SSTO.
Aerospike nozzles.
So how to get the altitude compensation? The aerospike[1],[2] is the most extensively studied method of altitude compensation so to get to this quickly it would be nice if this could be used. However, the aerospike (or plug nozzle for the shortened version) requires a toroidal combustion chamber. This would require significant modification of the engine.
However, one concept for getting an aerospike engine uses multiple small chambers arranged around a central spike. This was the idea for the X-33/VentureStar[3]. It was also used earlier on the Beta SSTO concept of Koelle[4],[5]. This could be conveniently used on the Falcon 9 first stage because of its multiple engines. The idea would be to greatly shorten the nozzles on the Merlins and arrange them around a central spike.
We need an estimate of the mass of the Falcon 9 first stage. This environmental assess- ment report on the SpaceX Grasshopper VTVL test vehicle[6] gives on p. 7 (page 17 according to the numbering as a PDF file) the Falcon 9 first stage kerosene load as 24,900 gallons and 38,900 gallons of LOX. Using a density of .820 gm/cc for kerosene
and 1.14 gm/cc for LOX, this amounts to about 245,000 kg propellant for the first stage.
Falcon 9 with aerospike nozzle becomes SSTO.
SpaceX has said the Falcon 9 first stage has a better than 20 to 1 mass ratio. This would give it a first stage dry mass of 13,000 kg. But this is using the Merlin 1C engine. From the thrust level and thrust/weight ratio[7] for this engine we can estimate its mass as about 650 kg. The Merlin 1D is supposed to be lighter, estimated as 440 kg. Using these brings the dry mass down to 11,000 kg.
The question is what would be the weight with the truncated nozzles and aerospike? A complaint against the aerospike used on the X-33 is that it had a worse T/W ratio than other LH2/LOX engines. However it had been planned for the full VentureStar version to use lightweight ceramics for the aerospike, expected to double the T/W ratio to about 80 to 1, better than the SSME's. With the advances in ceramics necessitated by the research and test flights with the hypersonic vehicles such high temperature lightweight ceramics should be further along now than they were with the VentureStar. For instance, the method of transpiration cooling using ceramics should make rocket engine combustion chambers and nozzles lighter and more reusable[8]. So I'll assume the total engine weight remains the same with the aerospike.
As before I'll use the Merlin Vacuum Isp in calculating the delta-V and take the required delta-V to orbit as 9,150 m/s for kerosene engines. Then we can get 6,000 kg payload:
342*9.8ln(1 + 245/(11 + 6)) = 9,167 m/s.
An SSTO is best utilized as a reusable though. Estimates of the added weight of reentry/landing systems are in the range of 28% [9]. However, with modern materials this probably can be cut to half that. Then the payload will be reduced to about 4.5 metric tons.
"An increase of 10% in Isp corresponds to an increase in 100% in payload."
This example illustrates well the importance of altitude compensation. Using it we are able to increase our engine Isp by 10% or more, to the extent we can achieve an SSTO with significant payload. Note that because of the rocket equation just being able to increase the Isp by 10% is no trivial feat. A rule of thumb among propulsion engineers is that "an increase of 10% in Isp corresponds to an increase in 100% in payload"[10]. I'll illustrate this with a single stage vehicle. A common estimate is that a kerosene-fueled SSTO needs a mass ratio of 20 to carry significant payload. Let's say a Falcon 9 sized rocket had instead high efficiency engines such as the NK-33[11] with an Isp of 331 s. You would need three of these. These have better thrust/weight than the Merlin 1C. So the dry weight is reduced to about 11 mT. Then it could carry a payload of 4.5 mT:
331*9.8ln(1 + 245/(11 + 4.5)) = 9,150 m/s.
With vacuum optimized nozzles high efficiency engines such as the NK-33 can get an Isp at 360+ s. Ten percent higher Isp than 331 s is at 364 s. Using this as the Isp allows a payload of 9.4 mT: 364*9.8ln(1 + 245/(11 + 9.4)) = 9,150 m/s.
Falcon Heavy with aerospike matches Phase 1 SLS a hundred times cheaper.
Interestingly an improvement in payload using altitude compensation also applies to staged vehicles. A preliminary calculation showed that with a two-stage vehicle you can increase your payload in the range of 25% to 30%. The improvement can be even better if the vehicle uses parallel staging, perhaps up to 50%. This is because with parallel staging the second stage still has to use the same nozzles as a lower stage because they still fire from the ground.
I'll illustrate this with the Falcon Heavy. SpaceX has said its side boosters will achieve a 30 to 1 mass ratio[12]. Reportedly the lower stages also will be stretched to hold more propellant. A posting on the NASASpaceFlight forum suggests a stretched version of the Falcon 9 will have 480,000 kg gross mass[13]. I'll estimate the stretched Falcon 9 based boosters as having a 435 mT propellant load and 15 mt dry mass. The central core stage has to be stronger since it holds the upper stage as well as the heavy payload. I'll take the dry mass for this lower core stage as 20 mT with the same 435 mT propellant load.
For the upper stage, I'll take the propellant load as 30 mT and the dry mass + fairing as 6 mT. For the vacuum Isp of the Merlin 1D I'll take the announced 310 s, and for the Isp of the upper stage, that of the Merlin Vacuum, 342 s. Then we can lift a payload of 51 mT:
310*9.8ln(1 + 2*435/(2*15 + 435 + 20 + 36 + 51)) + 310*9.8ln(1 + 435/(20 + 36 + 51)) + 342*9.8ln(1 + 30/(6 + 51)) = 9,155 m/s.
Now suppose we use altitude compensation to be able to get 342 s vacuum Isp for all the engines. Then we could lift 70 mT:
342*9.8ln(1 + 2*435/(2*15 + 435 + 20 + 36 + 70)) + 342*9.8ln(1 + 435/(20 + 36 + 70)) + 342*9.8ln(1 + 30/(6 + 70)) = 9,150 m/s
This means we could get a 70 mT launcher from essentially the same vehicle as the Falcon Heavy by using altitude compensation on the engines. SpaceX has said they intend to sell the Falcon Heavy for the range of $80 to $125 million per launch. The modifications to the aerospike nozzle should be relatively low cost compared to designing a whole new engine so the price should still be in this range.
Compare this to the estimates of the costs of the SLS program:
Space Launch System.
"Program costs.
During the joint Senate-NASA presentation in September 2011, it was
stated that the SLS program has a projected development cost of $18
billion through 2017, with $10B for the SLS rocket, $6B for the Orion
Multi-Purpose Crew Vehicle and $2B for upgrades to the launch pad and
other facilities at Kennedy Space Center. An unofficial NASA
document estimates the cost of the program through 2025 will total at
least $41B for four 70 metric ton launches (1 unmanned in 2017, 3
manned starting in 2021). The 130 metric ton version should not be
ready earlier than 2030."
http://en.wikipedia.org/wiki/Space_Launch_System#Program_costs
If that estimate for the total costs of the SLS is correct then that's $10 billion per launch for the interim 70 mT payload vehicle. That's two orders of magnitude higher than the Falcon Heavy with altitude compensation.
Reusable first stage Falcon 9 with aerospike also serves as a reusable booster.
This is for the expendable version of the Falcon stages. However, another benefit of the reusable version is that it could serve as the first stage of the reusable booster program (RBS) of the Air Force[14].
Reusable Falcon 9 with aerospike also as next-generation shuttle.
Another interesting possibility is suggested by the recent report of investigations of bringing back the shuttle as a commercial satellite launcher[15],[16]. My view is that the shuttle orbiter is too heavy for that role, ca. 80 mT in dry mass. This cuts greatly into the payload capacity. According to the reports the investigations also considered building their own shuttle but using the advancements made since the shuttle was designed. Then the reusable Falcon SSTO could be used to launch small payloads or even crew capsules. It could be modeled on the aerodynamic design of the shuttle since the aerodynamics for that are so well studied. Since kerolox has an overall density about 1,000 kg/m^3, the propellant could fit within the 300 m^3 shuttle-sized payload bay[17]with 55 m^3 left over. For a cargo only version we could also use the 75 m^3 sized volume of the crew cabin, for a total of 130 m^3.
The stretched version of the Falcon 9 first stage to be used for the side boosters of the Falcon Heavy are expected by SpaceX to have 30 to 1 mass ratio. The improved mass ratio over the current Falcon 9 first stage is probably coming from the fact that making your rocket larger in general improves your mass ratio, the fact that the new Merlin 1D is lighter, and also the fact the boosters do not have to support the weight of the upper stage and payload of the full vehicle. Using a 435 mT propellant load and a 15 mT dry mass, this could launch 15.3 mT: 342*9.8ln(1 + 435/(15 + 15.3)) = 9,155 m/s. If you take the reentry/landing systems mass with modern materials as 14% of the dry mass this takes up 2.1 mT from the payload, so to 13.2 mT. To scale up the space shuttle design for a 435 mT propellant tank compared to a 245 mT tank, the linear dimensions would only have to be scaled up by 20%.
Bob Clark
1.)Aerospike Engine.
http://www.aerospaceweb.org/design/aerospike/main.shtml
2.)Nozzle Design.
by R.A. O'Leary and J. E. Beck, Spring 1992
http://www.rocketdynetech.com/articles/nozzledesign.htm
3.)Aerospike engine.
http://en.wikipedia.org/wiki/Aerospike_engine
4.)Beta, A Single Stage Reusable Ballistic Space Shuttle Concept.
Based on a study contract of the German Federal Ministery for Education and Science, Bonn (RFT 1017).
May 1970
Space Division, Messerschmitt-Bolkow-Blohm ( MBB), Munich, Germany.
http://www.spacefuture.com/archive/beta_a_single_stage_reusable_ballistic_space_shuttle_concept.shtml
5.)A Cost Engineered Launch Vehicle for Space Tourism.
D E Koelle
TCS-TransCostSystems, Ottobrunn, Germany.
IAA-98-IAA.1.5.07
http://www.spacefuture.com/archive/a_cost_engineered_launch_vehicle_for_space_tourism.shtml
6.)Draft Environmental Assessment for Issuing an Experimental Permit to SpaceX
for Operation of the Grasshopper Vehicle at the McGregor Test Site,Texas.
September, 2011
http://www.faa.gov/about/office_org/headquarters_offices/ast/media/20110922%20SpaceX%20Grasshopper%20Draft%20EA.Final.pdf
7.)Merlin(rocket engine).
http://en.wikipedia.org/wiki/Merlin_%28rocket_engine%29
8.)Ceramic Materials for Reusable Liquid Fueled Rocket Engine Combustion Devices.
http://ammtiac.alionscience.com/pdf/AMPQ8_1ART06.pdf
9.)Newsgroups: sci.space.policy, sci.space.history, sci.astro, sci.physics
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Tue, 4 Oct 2011 10:04:42 -0700 (PDT)
Subject: Re: Elon Musk's SpaceX to build 'Grasshopper' hover-rocket
http://groups.google.com/group/sci.physics/msg/f3b4c27da1f13027?hl=en
10.)Discovery of New Molecule Could Lead to More Efficient Rocket Fuel.
ScienceDaily (Dec. 22, 2010)
http://www.sciencedaily.com/releases/2010/12/101222071831.htm
11.)NK-33.
http://www.astronautix.com/engines/nk33.htm
12.)SPACEX ANNOUNCES LAUNCH DATE FOR THE WORLD'S MOST POWERFUL ROCKET.
http://www.spacex.com/press.php?page=20110405
13.)Re: Falcon Heavy Master Update Thread.
http://forum.nasaspaceflight.com/index.php?topic=24711.msg723410#msg723410
14.)Doing a 180 - AFRL's Rocket-back Pathfinder.
Posted by Graham Warwick at 4/7/2010 7:52 AM CDT
http://www.aviationweek.com/aw/blogs/defense/index.jsp?plckController=Blog&plckScript=blogScript&plckElementId=blogDest&plckBlogPage=BlogViewPost&plckPostId=Blog:27ec4a53-dcc8-42d0-bd3a-01329aef79a7Post:1553afc7-cc1e-4b5e-9a6c-ced39704d348
15.)Next Gen Shuttle-Capable vehicle interest as secret effort to save orbiters ends.
December 19th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/12/next-gen-shuttle-vehicle-secret-effort-save-orbiters-ends/
16.)Atlantis Journal – Epilogue.
by MLD on Dec.19, 2011, under Commercial Space, Space Exploration, Space Policy, Space Shuttle Program
http://www.marylynnedittmar.com/?p=1303
17.)Space Shuttle orbiter.
Shuttle Orbiter Specifications.
http://en.wikipedia.org/wiki/Space_Shuttle_orbiter#Shuttle_Orbiter_Specifications
In post #86 in this thread is an argument for how two copies of the Falcon 1 first stage combined together using a single Merlin engine with an altitude compensating nozzle could be SSTO. I had forgotten though that the first stage of the Falcon 1e, which SpaceX wants to move to for small launches anyway, has about twice the propellant load of the Falcon 1's first stage so it could form the SSTO.
Aerospike nozzles.
So how to get the altitude compensation? The aerospike[1],[2] is the most extensively studied method of altitude compensation so to get to this quickly it would be nice if this could be used. However, the aerospike (or plug nozzle for the shortened version) requires a toroidal combustion chamber. This would require significant modification of the engine.
However, one concept for getting an aerospike engine uses multiple small chambers arranged around a central spike. This was the idea for the X-33/VentureStar[3]. It was also used earlier on the Beta SSTO concept of Koelle[4],[5]. This could be conveniently used on the Falcon 9 first stage because of its multiple engines. The idea would be to greatly shorten the nozzles on the Merlins and arrange them around a central spike.
We need an estimate of the mass of the Falcon 9 first stage. This environmental assess- ment report on the SpaceX Grasshopper VTVL test vehicle[6] gives on p. 7 (page 17 according to the numbering as a PDF file) the Falcon 9 first stage kerosene load as 24,900 gallons and 38,900 gallons of LOX. Using a density of .820 gm/cc for kerosene
and 1.14 gm/cc for LOX, this amounts to about 245,000 kg propellant for the first stage.
Falcon 9 with aerospike nozzle becomes SSTO.
SpaceX has said the Falcon 9 first stage has a better than 20 to 1 mass ratio. This would give it a first stage dry mass of 13,000 kg. But this is using the Merlin 1C engine. From the thrust level and thrust/weight ratio[7] for this engine we can estimate its mass as about 650 kg. The Merlin 1D is supposed to be lighter, estimated as 440 kg. Using these brings the dry mass down to 11,000 kg.
The question is what would be the weight with the truncated nozzles and aerospike? A complaint against the aerospike used on the X-33 is that it had a worse T/W ratio than other LH2/LOX engines. However it had been planned for the full VentureStar version to use lightweight ceramics for the aerospike, expected to double the T/W ratio to about 80 to 1, better than the SSME's. With the advances in ceramics necessitated by the research and test flights with the hypersonic vehicles such high temperature lightweight ceramics should be further along now than they were with the VentureStar. For instance, the method of transpiration cooling using ceramics should make rocket engine combustion chambers and nozzles lighter and more reusable[8]. So I'll assume the total engine weight remains the same with the aerospike.
As before I'll use the Merlin Vacuum Isp in calculating the delta-V and take the required delta-V to orbit as 9,150 m/s for kerosene engines. Then we can get 6,000 kg payload:
342*9.8ln(1 + 245/(11 + 6)) = 9,167 m/s.
An SSTO is best utilized as a reusable though. Estimates of the added weight of reentry/landing systems are in the range of 28% [9]. However, with modern materials this probably can be cut to half that. Then the payload will be reduced to about 4.5 metric tons.
"An increase of 10% in Isp corresponds to an increase in 100% in payload."
This example illustrates well the importance of altitude compensation. Using it we are able to increase our engine Isp by 10% or more, to the extent we can achieve an SSTO with significant payload. Note that because of the rocket equation just being able to increase the Isp by 10% is no trivial feat. A rule of thumb among propulsion engineers is that "an increase of 10% in Isp corresponds to an increase in 100% in payload"[10]. I'll illustrate this with a single stage vehicle. A common estimate is that a kerosene-fueled SSTO needs a mass ratio of 20 to carry significant payload. Let's say a Falcon 9 sized rocket had instead high efficiency engines such as the NK-33[11] with an Isp of 331 s. You would need three of these. These have better thrust/weight than the Merlin 1C. So the dry weight is reduced to about 11 mT. Then it could carry a payload of 4.5 mT:
331*9.8ln(1 + 245/(11 + 4.5)) = 9,150 m/s.
With vacuum optimized nozzles high efficiency engines such as the NK-33 can get an Isp at 360+ s. Ten percent higher Isp than 331 s is at 364 s. Using this as the Isp allows a payload of 9.4 mT: 364*9.8ln(1 + 245/(11 + 9.4)) = 9,150 m/s.
Falcon Heavy with aerospike matches Phase 1 SLS a hundred times cheaper.
Interestingly an improvement in payload using altitude compensation also applies to staged vehicles. A preliminary calculation showed that with a two-stage vehicle you can increase your payload in the range of 25% to 30%. The improvement can be even better if the vehicle uses parallel staging, perhaps up to 50%. This is because with parallel staging the second stage still has to use the same nozzles as a lower stage because they still fire from the ground.
I'll illustrate this with the Falcon Heavy. SpaceX has said its side boosters will achieve a 30 to 1 mass ratio[12]. Reportedly the lower stages also will be stretched to hold more propellant. A posting on the NASASpaceFlight forum suggests a stretched version of the Falcon 9 will have 480,000 kg gross mass[13]. I'll estimate the stretched Falcon 9 based boosters as having a 435 mT propellant load and 15 mt dry mass. The central core stage has to be stronger since it holds the upper stage as well as the heavy payload. I'll take the dry mass for this lower core stage as 20 mT with the same 435 mT propellant load.
For the upper stage, I'll take the propellant load as 30 mT and the dry mass + fairing as 6 mT. For the vacuum Isp of the Merlin 1D I'll take the announced 310 s, and for the Isp of the upper stage, that of the Merlin Vacuum, 342 s. Then we can lift a payload of 51 mT:
310*9.8ln(1 + 2*435/(2*15 + 435 + 20 + 36 + 51)) + 310*9.8ln(1 + 435/(20 + 36 + 51)) + 342*9.8ln(1 + 30/(6 + 51)) = 9,155 m/s.
Now suppose we use altitude compensation to be able to get 342 s vacuum Isp for all the engines. Then we could lift 70 mT:
342*9.8ln(1 + 2*435/(2*15 + 435 + 20 + 36 + 70)) + 342*9.8ln(1 + 435/(20 + 36 + 70)) + 342*9.8ln(1 + 30/(6 + 70)) = 9,150 m/s
This means we could get a 70 mT launcher from essentially the same vehicle as the Falcon Heavy by using altitude compensation on the engines. SpaceX has said they intend to sell the Falcon Heavy for the range of $80 to $125 million per launch. The modifications to the aerospike nozzle should be relatively low cost compared to designing a whole new engine so the price should still be in this range.
Compare this to the estimates of the costs of the SLS program:
Space Launch System.
"Program costs.
During the joint Senate-NASA presentation in September 2011, it was
stated that the SLS program has a projected development cost of $18
billion through 2017, with $10B for the SLS rocket, $6B for the Orion
Multi-Purpose Crew Vehicle and $2B for upgrades to the launch pad and
other facilities at Kennedy Space Center. An unofficial NASA
document estimates the cost of the program through 2025 will total at
least $41B for four 70 metric ton launches (1 unmanned in 2017, 3
manned starting in 2021). The 130 metric ton version should not be
ready earlier than 2030."
http://en.wikipedia.org/wiki/Space_Launch_System#Program_costs
If that estimate for the total costs of the SLS is correct then that's $10 billion per launch for the interim 70 mT payload vehicle. That's two orders of magnitude higher than the Falcon Heavy with altitude compensation.
Reusable first stage Falcon 9 with aerospike also serves as a reusable booster.
This is for the expendable version of the Falcon stages. However, another benefit of the reusable version is that it could serve as the first stage of the reusable booster program (RBS) of the Air Force[14].
Reusable Falcon 9 with aerospike also as next-generation shuttle.
Another interesting possibility is suggested by the recent report of investigations of bringing back the shuttle as a commercial satellite launcher[15],[16]. My view is that the shuttle orbiter is too heavy for that role, ca. 80 mT in dry mass. This cuts greatly into the payload capacity. According to the reports the investigations also considered building their own shuttle but using the advancements made since the shuttle was designed. Then the reusable Falcon SSTO could be used to launch small payloads or even crew capsules. It could be modeled on the aerodynamic design of the shuttle since the aerodynamics for that are so well studied. Since kerolox has an overall density about 1,000 kg/m^3, the propellant could fit within the 300 m^3 shuttle-sized payload bay[17]with 55 m^3 left over. For a cargo only version we could also use the 75 m^3 sized volume of the crew cabin, for a total of 130 m^3.
The stretched version of the Falcon 9 first stage to be used for the side boosters of the Falcon Heavy are expected by SpaceX to have 30 to 1 mass ratio. The improved mass ratio over the current Falcon 9 first stage is probably coming from the fact that making your rocket larger in general improves your mass ratio, the fact that the new Merlin 1D is lighter, and also the fact the boosters do not have to support the weight of the upper stage and payload of the full vehicle. Using a 435 mT propellant load and a 15 mT dry mass, this could launch 15.3 mT: 342*9.8ln(1 + 435/(15 + 15.3)) = 9,155 m/s. If you take the reentry/landing systems mass with modern materials as 14% of the dry mass this takes up 2.1 mT from the payload, so to 13.2 mT. To scale up the space shuttle design for a 435 mT propellant tank compared to a 245 mT tank, the linear dimensions would only have to be scaled up by 20%.
Bob Clark
1.)Aerospike Engine.
http://www.aerospaceweb.org/design/aerospike/main.shtml
2.)Nozzle Design.
by R.A. O'Leary and J. E. Beck, Spring 1992
http://www.rocketdynetech.com/articles/nozzledesign.htm
3.)Aerospike engine.
http://en.wikipedia.org/wiki/Aerospike_engine
4.)Beta, A Single Stage Reusable Ballistic Space Shuttle Concept.
Based on a study contract of the German Federal Ministery for Education and Science, Bonn (RFT 1017).
May 1970
Space Division, Messerschmitt-Bolkow-Blohm ( MBB), Munich, Germany.
http://www.spacefuture.com/archive/beta_a_single_stage_reusable_ballistic_space_shuttle_concept.shtml
5.)A Cost Engineered Launch Vehicle for Space Tourism.
D E Koelle
TCS-TransCostSystems, Ottobrunn, Germany.
IAA-98-IAA.1.5.07
http://www.spacefuture.com/archive/a_cost_engineered_launch_vehicle_for_space_tourism.shtml
6.)Draft Environmental Assessment for Issuing an Experimental Permit to SpaceX
for Operation of the Grasshopper Vehicle at the McGregor Test Site,Texas.
September, 2011
http://www.faa.gov/about/office_org/headquarters_offices/ast/media/20110922%20SpaceX%20Grasshopper%20Draft%20EA.Final.pdf
7.)Merlin(rocket engine).
http://en.wikipedia.org/wiki/Merlin_%28rocket_engine%29
8.)Ceramic Materials for Reusable Liquid Fueled Rocket Engine Combustion Devices.
http://ammtiac.alionscience.com/pdf/AMPQ8_1ART06.pdf
9.)Newsgroups: sci.space.policy, sci.space.history, sci.astro, sci.physics
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Tue, 4 Oct 2011 10:04:42 -0700 (PDT)
Subject: Re: Elon Musk's SpaceX to build 'Grasshopper' hover-rocket
http://groups.google.com/group/sci.physics/msg/f3b4c27da1f13027?hl=en
10.)Discovery of New Molecule Could Lead to More Efficient Rocket Fuel.
ScienceDaily (Dec. 22, 2010)
http://www.sciencedaily.com/releases/2010/12/101222071831.htm
11.)NK-33.
http://www.astronautix.com/engines/nk33.htm
12.)SPACEX ANNOUNCES LAUNCH DATE FOR THE WORLD'S MOST POWERFUL ROCKET.
http://www.spacex.com/press.php?page=20110405
13.)Re: Falcon Heavy Master Update Thread.
http://forum.nasaspaceflight.com/index.php?topic=24711.msg723410#msg723410
14.)Doing a 180 - AFRL's Rocket-back Pathfinder.
Posted by Graham Warwick at 4/7/2010 7:52 AM CDT
http://www.aviationweek.com/aw/blogs/defense/index.jsp?plckController=Blog&plckScript=blogScript&plckElementId=blogDest&plckBlogPage=BlogViewPost&plckPostId=Blog:27ec4a53-dcc8-42d0-bd3a-01329aef79a7Post:1553afc7-cc1e-4b5e-9a6c-ced39704d348
15.)Next Gen Shuttle-Capable vehicle interest as secret effort to save orbiters ends.
December 19th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/12/next-gen-shuttle-vehicle-secret-effort-save-orbiters-ends/
16.)Atlantis Journal – Epilogue.
by MLD on Dec.19, 2011, under Commercial Space, Space Exploration, Space Policy, Space Shuttle Program
http://www.marylynnedittmar.com/?p=1303
17.)Space Shuttle orbiter.
Shuttle Orbiter Specifications.
http://en.wikipedia.org/wiki/Space_Shuttle_orbiter#Shuttle_Orbiter_Specifications
This thread, while making me nostalgic for my own alt.* days, reminds me of the physicist who said he could predict the winner of any horse race to multiple decimal points - provided it was a spherical horse moving through a vacuum.
R
RGClark
Guest
Quote from: RGClark on January 03, 2012, 05:47:21 pm
Elon Musk has said he wants to cut the costs to space to the $100 to $200 per kg range by reusability. This is about a two order of magnitude reduction in cost. To put this in perspective, this is like a trans-atlantic flight that costs $1,000 suddenly being cut to cost $10 to $20.
Musk has said this transformation of the Falcon 9 to full reusability will be very hard. I don't believe it will be. But first, keep in mind how important that reduction in cost will be if it succeeds. If it succeeds then SpaceX will monopolize the launch business if the other launch companies do not field their own reusable vehicles. So there is a tremendous financial incentive for SpaceX to invest in reusability. Now, most in the industry believe reusability is very difficult for orbital vehicles and not even worth the expense. So if Musk reinforces that idea then he has a better chance at being able to field one without the other launch providers having one. And since they will not have even started to develop one, it will take them some time to catch up. The effect is that Musk will have a monopoly on all launches for at least a few years.
I don't know if that is Musk's intent in saying reusability is very hard. Actually I'm inclined to believe he is just saying what most in the industry believe including his own engineers. But a key reason why reusability is not very hard is because the cost in mass in reentry and landing systems is surprisingly low. In regards to the technical difficulty, there is none. We know how to do it as the shuttle orbiter and the X-37B and Dragon spacecraft has shown. I include the Dragon in the list of reusables because its heat shield showed minimal degradation on return. Musk has said the same heat shield could make hundreds of flights, at least to LEO.
I made an estimate before of about 28% of the landed mass has to go to reentry/landing systems. This was based on estimates of 15% for thermal protection, 10% for wings or for propellant for vertical landing, and 3% for landing gear. However, I said likely with modern materials this could be cut to half that. In fact, it might even be lower than 10%.
1.)Weight of thermal protection.
Robert Zubrin has given an estimate of 15% of the landed weight for the weight of thermal protection systems(TPS):
Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_launch_system#Reentry_heat_shields
However, I gather this was in relation to the older capsules, Mercury, Gemini, Apollo, etc. Indeed the weight of the ablative heat shield on the Apollo capsule was about 15%:
Apollo Command/Service Module.
2.7 Specifications
http://en.wikipedia.org/wiki/Apollo_Command/Service_Module#Specifications
However, the space shuttle with its mostly silica tiles was able to reduce the TPS weight to about 8% of the maximum landing weight of 104,000 kg:
Space Shuttle thermal protection system.
3.3 Weight considerations.
http://en.wikipedia.org/wiki/Space_shuttle_thermal_protection_system#Weight_considerations
Also, for the X-37B the TUFROC leading edge material instead of the shuttles RCC and the TUFI AETB material instead of the shuttles silica tiles are either of equal or lower weight than the shuttles TPS materials while being tougher and requiring less maintenance:
X-37B Orbital Test Vehicle.
http://www.boeing.com/defense-space/ic/sis/x37b_otv/x37b_otv.html
For ablative TPS, the PICA-X material used on the Dragon capsule weights about half the weight of the AVCOAT material used on the Apollo heat shield:
Re: Dragon v/s Orion.
http://forum.nasaspaceflight.com/index.php?topic=23522.msg754168#msg754168
while being able to still survive lunar and even Martian reentry speeds.
SpaceX has found that at least for LEO reentry speed judging from the minimal degradation on the Falcon 9/Dragon test flight, the PICA-X heat shield could be reused hundreds of times.
Also, for vertical powered landings a la the DC-X, you might not even need an extra heat shield for base first landings. One proposal for a VTVL SSTO uses low thrust during the descent as well as a high temperature-resistant aerospike nozzle to serve as the reentry thermal protection. You would need to retain more mass in propellant or some inert gas for this purpose though.
Another idea for a vertical landing vehicle would be to reenter head first. This was the preferred method of the Air Force since it provided increased cross-range. In that case you would have the blunt heat shield at the top of each stage. I thought this method would be unstable with the heavy engines now at the top during reentry, but since this was considered for the orbital version of the DC-X presumably this was solved.
2.)Weight of the wings and the landing gear.
For horizontal landing, a common estimate is that the weight of wings is 10% of the landed weight. This comes from aircraft examples though where the wings have to carry the weight of the fuel which can be as much as the dry weight of the aircraft itself or more.
An example where the propellant will not be carried in the wings and lightweight composites will be used is the Skylon. According to their released specifications the wing weight will be less than 2% of the take-off weight, which is the appropriate weight to compare to for a horizontal take-off vehicle:
The SKYLON Spaceplane.
by Richard Varvill and Alan Bond
Journal of the British Interplanetary Society, Vol. 57. pp. 22–32, 2004
p. 32.
http://www.reactionengines.co.uk/downloads/JBIS_v57_22-32.pdf
On that same page the landing gear weight is the only 1.5% of the take-off weight.
Then for a vertical take-off vehicle these low weight proportions should apply to the dry, landing weight.
Bob Clark
Elon Musk has said he wants to cut the costs to space to the $100 to $200 per kg range by reusability. This is about a two order of magnitude reduction in cost. To put this in perspective, this is like a trans-atlantic flight that costs $1,000 suddenly being cut to cost $10 to $20.
Musk has said this transformation of the Falcon 9 to full reusability will be very hard. I don't believe it will be. But first, keep in mind how important that reduction in cost will be if it succeeds. If it succeeds then SpaceX will monopolize the launch business if the other launch companies do not field their own reusable vehicles. So there is a tremendous financial incentive for SpaceX to invest in reusability. Now, most in the industry believe reusability is very difficult for orbital vehicles and not even worth the expense. So if Musk reinforces that idea then he has a better chance at being able to field one without the other launch providers having one. And since they will not have even started to develop one, it will take them some time to catch up. The effect is that Musk will have a monopoly on all launches for at least a few years.
I don't know if that is Musk's intent in saying reusability is very hard. Actually I'm inclined to believe he is just saying what most in the industry believe including his own engineers. But a key reason why reusability is not very hard is because the cost in mass in reentry and landing systems is surprisingly low. In regards to the technical difficulty, there is none. We know how to do it as the shuttle orbiter and the X-37B and Dragon spacecraft has shown. I include the Dragon in the list of reusables because its heat shield showed minimal degradation on return. Musk has said the same heat shield could make hundreds of flights, at least to LEO.
I made an estimate before of about 28% of the landed mass has to go to reentry/landing systems. This was based on estimates of 15% for thermal protection, 10% for wings or for propellant for vertical landing, and 3% for landing gear. However, I said likely with modern materials this could be cut to half that. In fact, it might even be lower than 10%.
1.)Weight of thermal protection.
Robert Zubrin has given an estimate of 15% of the landed weight for the weight of thermal protection systems(TPS):
Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_launch_system#Reentry_heat_shields
However, I gather this was in relation to the older capsules, Mercury, Gemini, Apollo, etc. Indeed the weight of the ablative heat shield on the Apollo capsule was about 15%:
Apollo Command/Service Module.
2.7 Specifications
http://en.wikipedia.org/wiki/Apollo_Command/Service_Module#Specifications
However, the space shuttle with its mostly silica tiles was able to reduce the TPS weight to about 8% of the maximum landing weight of 104,000 kg:
Space Shuttle thermal protection system.
3.3 Weight considerations.
http://en.wikipedia.org/wiki/Space_shuttle_thermal_protection_system#Weight_considerations
Also, for the X-37B the TUFROC leading edge material instead of the shuttles RCC and the TUFI AETB material instead of the shuttles silica tiles are either of equal or lower weight than the shuttles TPS materials while being tougher and requiring less maintenance:
X-37B Orbital Test Vehicle.
http://www.boeing.com/defense-space/ic/sis/x37b_otv/x37b_otv.html
For ablative TPS, the PICA-X material used on the Dragon capsule weights about half the weight of the AVCOAT material used on the Apollo heat shield:
Re: Dragon v/s Orion.
http://forum.nasaspaceflight.com/index.php?topic=23522.msg754168#msg754168
while being able to still survive lunar and even Martian reentry speeds.
SpaceX has found that at least for LEO reentry speed judging from the minimal degradation on the Falcon 9/Dragon test flight, the PICA-X heat shield could be reused hundreds of times.
Also, for vertical powered landings a la the DC-X, you might not even need an extra heat shield for base first landings. One proposal for a VTVL SSTO uses low thrust during the descent as well as a high temperature-resistant aerospike nozzle to serve as the reentry thermal protection. You would need to retain more mass in propellant or some inert gas for this purpose though.
Another idea for a vertical landing vehicle would be to reenter head first. This was the preferred method of the Air Force since it provided increased cross-range. In that case you would have the blunt heat shield at the top of each stage. I thought this method would be unstable with the heavy engines now at the top during reentry, but since this was considered for the orbital version of the DC-X presumably this was solved.
2.)Weight of the wings and the landing gear.
For horizontal landing, a common estimate is that the weight of wings is 10% of the landed weight. This comes from aircraft examples though where the wings have to carry the weight of the fuel which can be as much as the dry weight of the aircraft itself or more.
An example where the propellant will not be carried in the wings and lightweight composites will be used is the Skylon. According to their released specifications the wing weight will be less than 2% of the take-off weight, which is the appropriate weight to compare to for a horizontal take-off vehicle:
The SKYLON Spaceplane.
by Richard Varvill and Alan Bond
Journal of the British Interplanetary Society, Vol. 57. pp. 22–32, 2004
p. 32.
http://www.reactionengines.co.uk/downloads/JBIS_v57_22-32.pdf
On that same page the landing gear weight is the only 1.5% of the take-off weight.
Then for a vertical take-off vehicle these low weight proportions should apply to the dry, landing weight.
Bob Clark
Byeman said:
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Don't be so harsh on him, at last i read his post.
R
RGClark
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Orionblamblam said:
Any readers of this forum want routine access to space. Elon and SpaceX believe this is doable and are firmly committed to working toward that goal. The only difference I have with what they are saying is that it's in fact easy, at least compared to the complexity of the design of what is already flying.
That is, compared to the complexity of the design of high performance engines and lightweight structures, adding on reentry and landing systems is trivial both in weight and complexity.
Bob Clark
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Cool! I want one of those violation fingers slapped on me too!
R
RGClark
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RGClark said:
Just saw this discussed on Nasaspaceflight.com :
Elon Musk on SpaceX’s Reusable Rocket Plans.
February 7, 2012 6:00 PM
http://www.popularmechanics.com/science/space/rockets/elon-musk-on-spacexs-reusable-rocket-plans-6653023
Then for the Falcon 9, the payload would be reduced from 10 mT to 6
mT. If the reduction in payload really is this high, then maybe it
would be better to recover the first stage at sea. The loss in payload
is coming from the reduction in the speed of staging as well as the
need to retain a portion of the fuel for the return to base.
Recovering at sea would not have these disadvantages because you could
let the first stage make its usual trajectory at returning to the sea
but use just small amount of propellant for the final slowdown before
the sea impact.
It is ironic that the hardest part is recovering the first stage
instead of the orbital upper stage. Also interesting is that the
payload becomes reduced to about what you can get with an SSTO Falcon
9 first stage using Merlin 1D's with altitude compensation. This would
also be cheaper in not having the upper stage and you would not have
the problem of returning to the launch base for a lower stage.
In this article Musk does mention that returning back to the launch
point allows the turnaround time at least for the first stage to be
just hours. But will we really need that short a turnaround time at
this stage of the game? A turnaround time of a few days would seem to
be sufficient.
Perhaps the idea that retrieval at sea would be so expensive comes
from the experience of the shuttle with the SRB's. But these were
quite large and heavy at ca. 90 mT dry compared to that of the Falcon
9 first stage at less than 15 mT. Also, it is well known the labor
costs for the shuttle were greatly inflated compared to a privately
funded program.
The only additional requirement is that you would need a cover that
could be extended to cover the engine section and would be watertight.
Bob Clark
The amount of fuel needed for recovery doesn't change dramatically if you land on water versus on land.
The first stage needs to slow down from Mach 6 to 0 for land, and from Mach 6 to 30 km/h for water. Those last 30 km/h aren't going to make much of a difference.
A water landing also introduces new complications.
1. A shroud to cover the engines would be at least 3.6 m diameter x 2 m high - and that's with a flat plate covering the engines, no streamlining at all. That's easily going to be several hundred kg to a ton of additional ballast.
2. everything needs to be corrosion-proofed
3. it makes the turnaround time longer
The first stage needs to slow down from Mach 6 to 0 for land, and from Mach 6 to 30 km/h for water. Those last 30 km/h aren't going to make much of a difference.
A water landing also introduces new complications.
1. A shroud to cover the engines would be at least 3.6 m diameter x 2 m high - and that's with a flat plate covering the engines, no streamlining at all. That's easily going to be several hundred kg to a ton of additional ballast.
2. everything needs to be corrosion-proofed
3. it makes the turnaround time longer
R
RGClark
Guest
I've been arguing that SSTO's are actually easy because how to achieve them is perfectly obvious: use the most weight optimized stages and most Isp efficient engines at the same time, i.e., optimize both components of the rocket equation. But I've recently found it's even easier than that! It turns out you don't even need the engines to be of particularly high efficiency.
SpaceX is moving rapidly towards testing its Grasshopper scaled-down version of a reusable VTVL Falcon 9 first stage:
Reusable rocket prototype almost ready for first liftoff.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: July 9, 2012
http://www.spaceflightnow.com/news/n1207/10grasshopper/
SpaceX will be duplicating in this what the DC-X accomplished in the early 90's. The DC-X was a scaled down, low altitude test vehicle for a full-scale SSTO VTVL vehicle. So could the full-sized Falcon 9 first stage act as a VTVL SSTO?
SpaceX deserve kudos for achieving a highly weight optimized Falcon 9 first stage at a 20 to 1 mass ratio. However, the Merlin 1C engine has an Isp no better than the engines we had in the early sixties at 304 s, and the Merlin 1D is only slightly better on the Isp scale at 310 s. This is well below the highest efficiency kerosene engines (Russian) we have now whose Isp's are in the 330's. So I thought that closed the door on the Falcon 9 first stage being SSTO.
However, I was surprised when I did the calculation that because of the Merlin 1D's lower weight the Falcon 9 first stage could indeed be SSTO. I'll use the Falcon 9 specifications estimated by GW Johnson, a former rocket engineer, now math professor:
WEDNESDAY, DECEMBER 14, 2011
Reusability in Launch Rockets.
http://exrocketman.blogspot.com/2011/12/reusability-in-launch-rockets.html
The first stage propellant load is given as 553,000 lbs, 250,000 kg, and the dry weight as 30,000 lbs, 13,600 kg. The Merlin 1C mass hasn't been released, but I'll estimate it as 650 kg, from its reported thrust and thrust/weight ratio. The Merlin 1D mass has been estimated to be 450 kg. Then on replacing the 1C with the 1D we save 9*200 = 1,800kg from the dry weight to bring it to 11,800 kg.
The required delta v to orbit is frequently estimated as 30,000 feet per second for kerosene-fueled vehicles, 9,144 m/s. When calculating the delta v your rocket can achieve, you can just use your engines vacuum Isp since the loss of Isp at sea level is taken into account in the 30,000 fps number. Then this version of the Falcon 9 first stage could lift 1,200 kg to orbit:
310*9.81ln(1 + 250/(11.8 + 1.2)) = 9,145 m/s.
Then the Falcon 9 first stage could serve as a proof of principle SSTO on the switch to the Merlin 1D engine.
Bob Clark
SpaceX is moving rapidly towards testing its Grasshopper scaled-down version of a reusable VTVL Falcon 9 first stage:
Reusable rocket prototype almost ready for first liftoff.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: July 9, 2012
http://www.spaceflightnow.com/news/n1207/10grasshopper/
SpaceX will be duplicating in this what the DC-X accomplished in the early 90's. The DC-X was a scaled down, low altitude test vehicle for a full-scale SSTO VTVL vehicle. So could the full-sized Falcon 9 first stage act as a VTVL SSTO?
SpaceX deserve kudos for achieving a highly weight optimized Falcon 9 first stage at a 20 to 1 mass ratio. However, the Merlin 1C engine has an Isp no better than the engines we had in the early sixties at 304 s, and the Merlin 1D is only slightly better on the Isp scale at 310 s. This is well below the highest efficiency kerosene engines (Russian) we have now whose Isp's are in the 330's. So I thought that closed the door on the Falcon 9 first stage being SSTO.
However, I was surprised when I did the calculation that because of the Merlin 1D's lower weight the Falcon 9 first stage could indeed be SSTO. I'll use the Falcon 9 specifications estimated by GW Johnson, a former rocket engineer, now math professor:
WEDNESDAY, DECEMBER 14, 2011
Reusability in Launch Rockets.
http://exrocketman.blogspot.com/2011/12/reusability-in-launch-rockets.html
The first stage propellant load is given as 553,000 lbs, 250,000 kg, and the dry weight as 30,000 lbs, 13,600 kg. The Merlin 1C mass hasn't been released, but I'll estimate it as 650 kg, from its reported thrust and thrust/weight ratio. The Merlin 1D mass has been estimated to be 450 kg. Then on replacing the 1C with the 1D we save 9*200 = 1,800kg from the dry weight to bring it to 11,800 kg.
The required delta v to orbit is frequently estimated as 30,000 feet per second for kerosene-fueled vehicles, 9,144 m/s. When calculating the delta v your rocket can achieve, you can just use your engines vacuum Isp since the loss of Isp at sea level is taken into account in the 30,000 fps number. Then this version of the Falcon 9 first stage could lift 1,200 kg to orbit:
310*9.81ln(1 + 250/(11.8 + 1.2)) = 9,145 m/s.
Then the Falcon 9 first stage could serve as a proof of principle SSTO on the switch to the Merlin 1D engine.
Bob Clark
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yes, a first stage Falcon 9 with 9xMerlin 1D could bring 1200 kg into equatorial orbit or 800 kg to higher inclined lower orbit
if you wanna trow the stage away after launch
if this first stage Falcon 9 as a reusable SSTO, thing see little bit diverse:
it's large volume/low mass fraction, give a not bad aerodynamic drag to terminal velocity of ~40meter/seconds
Only two problems are here:
To stabilizes the stage on it's way down, that engines show downwards.
And the mass of landing fuel for one Merlin 1D (∆V of 200 meter/seconds) and a landing gear
That mass is a lost on Payload !
if you wanna trow the stage away after launch
if this first stage Falcon 9 as a reusable SSTO, thing see little bit diverse:
it's large volume/low mass fraction, give a not bad aerodynamic drag to terminal velocity of ~40meter/seconds
Only two problems are here:
To stabilizes the stage on it's way down, that engines show downwards.
And the mass of landing fuel for one Merlin 1D (∆V of 200 meter/seconds) and a landing gear
That mass is a lost on Payload !
RGClark said:
http://www.astronautix.com/engines/nk31.htm
Specific impulse: 353 s.
but this is only an upper-stage rocket engine. Maybe it is possible to combine this engine with an ED-nozzle so it could be used from launch to orbit.
Gridlock said:
This is hilarious, and so relevant (sigh...)
Michel - try browsing information on google about "kerosene as fuel for a SSTO", and you'll change your mind. You will REALLY want to be hard with RG Clark after that. Thrust me.
http://www.google.fr/#q=%22The+Coming+SSTO%22&hl=fr&prmd=imvns&ei=3P8IULLhHOGK0AXy-dDCCg&start=10&sa=N&bav=on.2,or.r_gc.r_pw.r_qf.,cf.osb&fp=91c96fcb7ddaab89&biw=1024&bih=621
http://www.google.fr/#hl=fr&sclient=psy-ab&q=%22kerosene+fuel%22%22SSTO%22&oq=%22kerosene+fuel%22%22SSTO%22&gs_l=hp.3...5219.11026.0.11211.21.21.0.0.0.0.143.1301.19j2.21.0...0.0...1c.ew4psCEjCuU&pbx=1&bav=on.2,or.r_gc.r_pw.r_qf.,cf.osb&fp=91c96fcb7ddaab89&biw=1024&bih=621
http://www.google.fr/#q=%22kerosene+fueled%22%22SSTO%22&hl=fr&prmd=imvns&ei=OgAJUIS0OKnF0QXSg8zJCg&start=20&sa=N&bav=on.2,or.r_gc.r_pw.r_qf.,cf.osb&fp=91c96fcb7ddaab89&biw=1024&bih=621
Sure enough, dear RG Clark, you know how to use google to your advantage.
Congrats for that - you could be some outstanding lobbyist for, I don't know - a toothpaste, or a politician, see what I mean ?
The little issue - your ramblings are not exactly *interesting*.
I don't want one - I tried to be constructive and polite (or so I felt !)
Have a nice day everyone !
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