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An SSTO as "God and Robert Heinlein intended".

Posted by: RGClark - Tue Jan 04, 2011 8:37 am
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An SSTO as "God and Robert Heinlein intended". 
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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Feb 26, 2011 6:27 am
Quote:
Another option for a manned launcher. In this report Boeing proposes heavy lift launchers using existing components:
Heavy Lift Launch Vehicles with Existing Propulsion Systems.
Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
Boeing Phantom Works,Huntsville, AL 35824
Doug Blue
Boeing Space Exploration,Huntington Beach, CA 92605
http://www.launchcomplexmodels.com/Direct/documents/AIAA-2010-2370-650.pdf
One of the proposals is of a manned launcher with the Orion capsule using a shuttle ET propellant tank and four RS-68 engines. This does not use an upper stage but is not a single-stage-to-orbit vehicle because the final push to orbit is made by the onboard thrusters on the Orion spacecraft.
However, it is interesting in this report comparison is made to the [url="http://en.wikipedia.org/wiki/S-IVB"]S-IVB[/url] upper stage on the Apollo rocket. I was reminded of a suggestion of Gary Hudson that the S-IVB would be single-stage-to-orbit with significant payload if it used the high efficiency SSME rather than the J-2 engine:
A Single-Stage-to-Orbit Thought Experiment.
Gary C Hudson
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml
In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg. This would be just enough to carry the crewed version of the Dragon spacecraft without cargo.


It is notable that the upper stage of the Ares I is based on this S-IVB stage. Then this upper stage as well should be able to act as an SSTO with an SSME engine. This is important because the Ares I upper stage was originally planned to use the SSME, so this means much of the technical and financial analysis of using the SSME for the upper stage of the Ares I has already been done.
However, because of the cost of the SSME engine and technical risk in making it airstartable, the decision was made to use the J-2X engine instead. But for the SSTO purpose you don't have the problem of making it airstartable, and as I discussed the reusability maintenance costs can be reduced by an order of magnitude for the SSME.
This report contains some of the specifications on the Ares I upper stage:

NASA’s Ares I Upper Stage.
http://www.nasa.gov/pdf/231430main_UpperStage_FS_final.pdf

The propellant mass is listed as 138 mT, the dry mass of the stage as 17.5 mT, and the interstage mass, as 4.1 mT. See the second attached image below taken from page 2 of the report. The interstage supports the weight of the upper stage on top of the lower stage so won't be needed for the SSTO version. So we can take the dry mass now as 13.4 mT.
We need to add onto this now the extra weight of using the SSME over the J-2X engine. The report lists the J-2X mass as 2.5 mT. The SSME mass is 3.1 mT, .6 mT heavier. This brings the dry weight to 14 mT.
A puzzlingly high value of 2.5 mT however is given for the avionics. You wouldn't think it would need to be this high if it consisted of just electronics and computer systems with modern miniaturization. Most of the avionics is included in the "instrument unit". As you can see from the attached image below, the instrument unit is regarded as a separate element of the upper stage and is contained within the forward skirt of the stage. The forward skirt serves to support the weight of the Orion CEV, so needs to have significant strength and mass to support the 20,000+ kg weight of the Orion spacecraft. So I'm wondering if that 2.5 mT mass is including the mass of this forward skirt.
The forward skirt mass can certainly be reduced if using a Dragon spacecraft at only one quarter the mass of the Orion. So that part of the dry mass will be reduced, though it's uncertain if the avionics mass itself can be reduced. In any case using 14 mT dry mass of the Ares I upper stage, the 138 mT propellant mass, the 425 s trajectory averaged Isp of the SSME given by Hudson, and a 9,200 m/s required delta-V to orbit, we can calculate the payload to orbit can be 3 mT:

425*9.8ln(1 + 138/(14 + 3)) = 9,205 m/s.

This payload mass would not be enough for the Dragon spacecraft but might be enough for an innovative new spacecraft proposal from the University of Maryland:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration Vehicle.
http://www.nianet.org/rascal/forum2006/presentations/1010_umd_paper.pdf

This uses a cylindrical shape for the capsule so would have more space for the crew/passengers. It also uses a new design for a thermal protection system called a "parashield" that would save weight over the traditional ablative design. The mass of the capsule in this study is given as 3,268 kg, so would only have to be reduced by a small proportion to fit within the payload mass constraints.
However, it might be possible to increase the payload capability of the SSME-powered Ares I upper stage to be able to carry even the Dragon spacecraft. First, more propellant can be carried in the same size tanks by densifying the propellant by subcooling:

Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-Scale Flight-Weight Tank for the X-33 Reusable Launch Vehicle.
http://www.grc.nasa.gov/WWW/RT/RT2001/5000/5870tomsik.html

As much as 10% more propellant can be carried by subcooled densification. This corresponds to 10% greater mass that can lofted to orbit. So from a 17 mT total of launch vehicle + payload, up to 18.7 mT. This extra mass can go to extra payload so to 4.7 mT payload.
Secondly, recent research has shown that from 10% to 20% weight savings can be made off the structural weight on launch vehicles:

NASA Recalculates To Save Weight On Launchers.
Jan 5, 2011
By Frank Morring, Jr.
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2011/01/03/AW_01_03_2011_p53-277413.xml&headline=NASA%20Recalculates%20To%20Save%20Weight%20On%20Launchers&channel=space

The structural mass sans engine is 11 mT. If 10% weight can be saved off this then that can be transfered to extra payload, bringing the payload capacity to 5.8 mT. This would then be within the payload capacity to carry the Dragon spacecraft.


Bob Clark


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Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Feb 26, 2011 11:45 am
Quote:
Another option for a manned launcher. In this report Boeing proposes heavy lift launchers using existing components:
Heavy Lift Launch Vehicles with Existing Propulsion Systems.
Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
Boeing Phantom Works,Huntsville, AL 35824
Doug Blue
Boeing Space Exploration,Huntington Beach, CA 92605
http://www.launchcomplexmodels.com/Direct/documents/AIAA-2010-2370-650.pdf
One of the proposals is of a manned launcher with the Orion capsule using a shuttle ET propellant tank and four RS-68 engines. This does not use an upper stage but is not a single-stage-to-orbit vehicle because the final push to orbit is made by the onboard thrusters on the Orion spacecraft.
However, it is interesting in this report comparison is made to the [url="http://en.wikipedia.org/wiki/S-IVB"]S-IVB[/url] upper stage on the Apollo rocket. I was reminded of a suggestion of Gary Hudson that the S-IVB would be single-stage-to-orbit with significant payload if it used the high efficiency SSME rather than the J-2 engine:
A Single-Stage-to-Orbit Thought Experiment.
Gary C Hudson
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml
In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg. This would be just enough to carry the crewed version of the Dragon spacecraft without cargo.


The point of the matter is that if you use highly weight optimized structures and high efficiency engines at the same time then what you wind up with will be a SSTO capable stage. The Ariane 5 core stage is another weight optimized structure using common bulkhead design for its propellant tanks. The Ariane 5 core stage will also become SSTO if using high efficiency SSME's instead of the Vulcain engines.
The specifications of the Ariane 5 are given here:

Ariane 5 Data Sheet.
http://www.spacelaunchreport.com/ariane5.html

The Ariane 5 generic "G" version could be lofted by a single SSME. It's gross mass is listed as 170 mT, and the propellant mass as 158 mT, for a dry mass of 12 mT. The Vulcain engine is listed on this page as weighing 1,700 kg:

Vulcain - Specifications.
http://www.spaceandtech.com/spacedata/engines/vulcain_specs.shtml

Switching to a heavier SSME engine would add 1.4 mT to the vehicle dry mass, so to 13.4 mT for the dry mass. Using a 425s average Isp again for the SSME, this would allow a 6,000 kg payload:

425*9.8ln(1 + 158/(13.4+6)) = 9,218 m/s.

We wish to use this for a man-rated vehicle though. The Ariane 5 was originally intended to be man-rated using the Hermes spaceplane to carry crew. However, it's not certain the degree this was followed-through when the Hermes was canceled.
As with the Ares I upper stage, there are means to increase the payload capacity. Subcooled densification allows 10% greater propellant to be carried, so then 10% greater mass can be lofted to orbit. This brings the total lofted weight from 19.4 mT to 21.3 mT. This extra weight can go to extra payload, so from 6 mT to about 8 mT in payload.
The Ariane 5 uses an aluminum alloy, but not the aluminum-lithium alloy being used now for the lightest weight designs. Switching to aluminum-lithium allows approx. 10% weight saving over the previous aluminum alloy. The structural mass sans the SSME engine is 10.3 mT, so about 1 mT would be saved that could go to extra payload.
I also mentioned before the new research that suggests 10% to 20% can be saved in structural mass because of overly conservative design now used. This would be another 1 mT that could be saved off the dry weight. These weight savings could go to extra payload, bringing the payload capacity to 10 mT.
ESA appears to be amenable to adapting the Ariane 5 core stage for other uses, considering its agreement with ATK to use it for an upper stage. So NASA or a private company should be able to make an agreement with the ESA to use it for this purpose, based on getting sufficient financing. In this regard, to get a prototype done at low cost I suggest using the RD-0120 russian analogue of the SSME's. These are in mothballs and probably can be obtained at greatly reduced price. As a point of comparison the NK-33 was mothballed by the russians and Aerojet was able to buy 36 of them for only $1.1 million each(!) Aerojets version of the NK-33 is now on track to be used by Orbital Sciences on their Taurus II launcher.
Then the Ariane 5 core version of this SSTO has the advantage over the Ares I upper stage and S-IVB versions in being already built and in current use. It also has now the capability when powered by an SSME or RD-0120 to launch a SpaceX Dragon sized spacecraft to orbit without having to use special fuel densifying or lightweighting methods.
NASA has said they want to support commercial space. Support for this launcher would allow for a small, relatively low cost launcher that would permit independent private companies to launch their own manned, or cargo flights to space.



Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Mar 04, 2011 3:21 am
JamesG wrote:
Yes, true. SSTO will require that "quantum leap" in propulsion and materials science. Which is really the only way we would be realistically talking about it. Otherwise we are discussing science fiction, or at most absurdly small payloads.

SSTO is possible now, a stretched Falcon 1 could do it. But the payload is tiny compared to the payload of a normal 2 stage Falcon 1. There just isn't any really good reason to do it, unless recovery, refurb, and reflight of the stage is very fast and cheap. No one has (non-destructively) reentered something the size of a full stage before.

I personally like a disposable upper stage and a VTVL RLV first stage. It optimizes to have the upper stage do more than its share of the delta-v so that the reentry of the first stage isn't as difficult. I'm sure Jon has written about this.

The other option is "stage and a half", where you just drop some of your engines. If you're using turbopumped engines, they represent a significant part of the vehicle's cost, and being able to recover and reuse 2/3 (or whatever) of your engines could reduce total cost. That you end up orbiting some of the engines means that they only need to be built to survive five or so starts, and are then 'retired' to orbit.


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Mar 04, 2011 3:36 am
Ben wrote:
No one has (non-destructively) reentered something the size of a full stage before.


The orbiters? :wink:

Quote:
The other option is "stage and a half", where you just drop some of your engines. If you're using turbopumped engines, they represent a significant part of the vehicle's cost, and being able to recover and reuse 2/3 (or whatever) of your engines could reduce total cost. That you end up orbiting some of the engines means that they only need to be built to survive five or so starts, and are then 'retired' to orbit.


That is probably a good compromise for the present. Although I like using solids that way you can flatten out your "stage" into a lifting body shape without needing the headache of arranging/shaping pressure tanks to oblige. Being able the horizontally land is a lot easier on the hardware than dropping it on the deck.


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Mar 04, 2011 3:58 am
The advantages of a SSTO are best utilized as a reusable vehicle.
Then you would have to subtract from this estimated payload mass the
mass needed for reentry and landing systems.
However, the Ariane core stage SSTO version could still be useful as an expendable vehicle.
Then you could have up to a 9,000 kg payload without
the reentry and landing systems. This is close to the 10,000 kg payload capacity of the Falcon 9.
I saw this article that had an estimate for the price of an expendable version of the SSME's:

PWR Offers Shuttle Engine Alternative.
Jul 15, 2009
By Joseph C. Anselmo
"The company also would manufacture additional engines using the
existing SSME design while beginning work on a modified design that
incorporates advances in the construction of nozzles and combustion
chambers. That would be ready to go into production within 3-4 years.
Maser estimates the modified SSME would cost two-thirds to four-fifths
of the original model - depending on the number ordered - and would be
'a little more expensive' than the company's RS-68 engine 'but in that
ballpark.'"
http://www.aviationweek.com/aw/generic/ ... lternative

Using a price of $40 million for the current SSME's this would
correspond to a price of from $26.7 to $32 million for the expendable
versions. Considering the fact the engines make up the bulk of the
cost of an expendable launcher, this expendable SSTO launcher very
well could be comparable in cost to the Falcon 9 at $50 million.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Mar 12, 2011 4:49 pm
In doing some background web searches, I found that the upper stage of the Direct team's Jupiter-246 vehicle also would become SSTO when switched to a SSME engine. I guess I should not have been surprised by this. The thesis I have been arguing repeatedly via email with individuals in NASA and the industry and on space forums such as this one is that if you use BOTH the most weight optimized designs AND the highest efficiency engines available, then what you will wind up with will be SSTO capable whether you intend it to or not.
By highest efficiency engines I don't mean just an engine optimized to have a high vacuum Isp only. I mean an engine of highest efficiency over the entire flight range to orbit. For hydrogen engines that is the SSME, and the Russian analogue RD-0120. However, the point of the matter is that the same is true of kerosene-fueled vehicles, when using both highly weight optimized structures and highest efficiency engines, such as the NK-33 or RD-180.
That a SSME-powered Jupiter-246 upper stage would be SSTO capable is important since the Direct team is more amenable to thinking outside the box. So they would be more amenable to the idea you could have a SSTO vehicle. And in fact at least the expendable version for this SSTO would be no more difficult than their proposal for the upper stage on the Jupiter-246.
Attached below is a diagram showing the specifications of the Direct team's Jupiter-246.

It uses 6 RL-10B-2 engines, according to the specifications here these weigh about 300 kg each:

RL10B-2
Propulsion System
http://www.pratt-whitney.com/StaticFile ... L10B-2.pdf

So exchanging these for a SSME will add about 1,300 kg to the upper stage weight. The dry mass will be increased then to 13,150 kg, and the gross mass to 204,000 kg. However, by the Space Shuttle main engine thrust specifications, even at 109% thrust this comes to only 417,300 lbs, or 189,700 kgf. So we'll reduce the propellant load to be lifted by the SSME.
We'll take the liftoff thrust/weight ratio to be 1.2. This will bring the gross mass down to 170,000 kg. Then the propellant mass has to be reduced by 34,000 kg. This brings the propellant mass down to 156,850 kg. Note this results in a mass ratio close to 13, well sufficient for SSTO with a hydrogen-fueled engine.
This mass ratio for a hydrogen-fueled stage of 13 is high, but the original number for the Jupiter-246 upper stage is even higher at above 17. These high values for the Direct teams launcher led to some doubts about their calculations, but an analysis by Dr. Steven Pietrobon showed it was in keeping with historical trends for upper stages:

Analysis of Propellant Tank Masses.
http://www.nasa.gov/pdf/382034main_018% ... Masses.pdf

Then using Hudson's 425s average Isp for the SSME and the 9,200 m/s required delta-V value for orbit, this stage as an SSTO could loft 6,200 kg to orbit:

425*9.8ln(1 + 156,850/(13,150 + 6,200)) = 9,200 m/s.

Again, we might be able to loft 10% greater total mass to orbit with propellant densification by subcooling and also shave 10% off the structural mass of the stage with the recent weight saving research. This will bring the payload mass up to about 9,000 kg.
In this calculation I kept the same size tanks and only used them partially filled. This might be useful if for instance the Jupiter-246 upper stage was built to the original Direct teams specifications and you wanted to use the same size stage, though switched to a SSME engine, for the SSTO application to save on costs.
However, you could save additional weight off the stage if you used smaller propellant tanks for the SSTO application. I estimate about 900 kg could be saved with the smaller tanks that could go to additional payload.



Bob Clark


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Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Mar 12, 2011 4:55 pm
There might still be some resistance to using the upper stage as a
SSTO. However, the point of the matter is even if you use these upper
stages as part of a multistage system you are still better off using
both highly weight optimized structures and engines of highest
surface-to-orbit-efficiency (not just vacuum optimized engines) at
the same time
.
We'll use in this case parallel staging of the same sized stages, a
bimese version, but using cross-feed fueling. This is a fueling method
that has both stages firing, as with parallel staging, but all the
propellant is coming from only a single stage at a time. Then when
that stage exhausts its propellant, it is jettisoned, and the
remaining stage proceeds on with its own full tank of propellant still
remaining.
Let's see how much payload we can carry in this case. Again assume the
425s trajectory averaged Isp of Hudson, and the 9,200 m/s required
delta-V for orbit.
Estimate the possible payload as 29,000 kg. For the first segment of
the flight the achieved delta-V would be: 425*9.8ln(1+156,850/
(2*13,150 + 156,850 +29,000)) = 2,305 m/s.
For the second segment, use the 455s vacuum Isp of the SSME's:
455*9.8ln(1 + 156,850/(13,150 + 29,000)) =6,921. And the total delta-V
is 9,226 m/s, sufficient for orbit with a 29,000 kg payload.
Note that a 29,000 kg payload is sufficient to even carry a Orion
capsule, at least in an expendable version of the staged vehicle
without reentry and landing systems.
Then you have different options for the vehicle. As a single stage it
could carry a small capsule such as the SpaceX Dragon, or the Boeing
CST-100. But using twinned copies of it, it would be able to loft the
heavier Orion spacecraft.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


Last edited by RGClark on Tue May 10, 2011 2:23 pm, edited 1 time in total.



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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Mar 12, 2011 4:58 pm
Another highly weight optimized stage was the S-II second stage on the Saturn V. According to this Wikipedia page it was even better optimized than the S-IVB stage :

Saturn V.
S-II second stage.
http://en.wikipedia.org/wiki/Saturn_V#S-II_second_stage

The 5 J-2 engines used had a mass of 1,580 kg each, for a total mass of 7,900 kg. You'll need 3 of the SSME's operating at 109% thrust to lift the mass. So the 7,900 kg mass of the engines is replaced with 9,300 kg. And the 36,000 kg S-II dry mass is raised to 37,400 kg and the gross mass is raised to 481,400 kg.
Now using Gary Hudson's 425s trajectory averaged Isp for the SSME engines, and the 9,200 m/s required delta-V to orbit. We get a 17,000 kg payload:

425*9.8ln((481400 + 17000)/(37400 + 17000)) = 9,225 m/s

However, again we can get 10% greater total mass to orbit by propellant densification. This brings the payload to 22,440 kg. Also perhaps 10% off the structural mass can be saved by using aluminum-lithium alloy. And an additional 10% mass can be saved by the new weight saving methods. These weight savings can go to extra payload to bring the payload mass up to 28,000 kg. Note this is sufficient now to carry the Orion spacecraft as a SSTO.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Mon Mar 14, 2011 9:36 pm
Okay, I think you've made your point that an SSTO is possible with today's technology. Use lightweight aluminium-titanium alloy, weight saving methods and high-efficiency engines, we get it. What I don't see however is why one wouldn't add staging to that list and increase the payload even more. Recovering a first stage hasn't been done that much (really only the Shuttle SRB's I think, and the occasional Ariane booster), but it has been shown to be possible, and staging itself doesn't appear to be especially dangerous. So why not TSTO?

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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Tue May 10, 2011 2:45 pm
Lourens wrote:
Okay, I think you've made your point that an SSTO is possible with today's technology. Use lightweight aluminium-titanium alloy, weight saving methods and high-efficiency engines, we get it. What I don't see however is why one wouldn't add staging to that list and increase the payload even more. Recovering a first stage hasn't been done that much (really only the Shuttle SRB's I think, and the occasional Ariane booster), but it has been shown to be possible, and staging itself doesn't appear to be especially dangerous. So why not TSTO?


I discussed in a post above that with staging you could lift the heavier Orion spacecraft rather than just the Dragon:

viewtopic.php?p=44980#p44980

However, note the type of staging I'm suggesting using is parallel staging with cross-feed fueling. SpaceX has stated by using cross-feed fueling they can increase their payload to orbit on their Falcon Heavy by 50%:

FALCON HEAVY OVERVIEW.
http://www.spacex.com/falcon_heavy.php

But quite key is to realize that with parallel staging the engines are being used from the ground all the way to near vacuum conditions. So you want your engines to be optimized for the entire flight range from sea level to vacuum, not just optimized for the vacuum case as with the upper stages used in serial staging.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Wed Jun 08, 2011 8:19 pm
Nice article here arguing in favor of NASA promoting small manned commercial vehicles:

Human spaceflight for less: the case for smaller launch vehicles,
revisited.
by Grant Bonin
Monday, June 6, 2011
http://www.thespacereview.com/article/1861/1

In the comments section, I commented that the capability to produce
such small, low cost, manned vehicles exists now. I estimated the cost
for a such a reusable vehicle in the range of a few tens of millions
of dollars unit cost, comparable to a medium sized business jet.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Jun 11, 2011 11:11 pm
To R Clark, June 8, others:

SSTO, though just physically possible just makes no sense (note that it has not been done). The high mass ratio and that most of the aceleration is done in vacuum requires a stage that is not suited for ascent from sea level then ascend vertically through the atmosphere.

Great compromise to the engines, or resort to strange, yet unused designs like aerospikes would have to be resorted to.

Two stages to orbit makes sense. For maximum performance, for small launchers at least, would have the 1st stage act as a TAV (transatmospheric vehicle), and use atmosphere optimized engines, and enclose the upper stage(s).

Then the upper stage(s) need not be aerodynamic, and with a lofted trajectory could have a greater portion of the 2nd stage thrust directed nearly horizontally. If it were to start with a 20 deg pitch, 94% of the thrust is the horizontal co0mponent, and 34% vertical component helps support the stage during the earlier part of the burn, where total acceleration might be only 1 g or so.

The first stage can be engineered to return to launch site fairly easily, as the staging could take place at a fairly low velocity at an altitude high enough for a vacuum adapted 2nd stage engine to work well.

This is the design for the 100-200 kg GLOW Microlaunchers, and it works well for larger ones also.


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Thu Jun 16, 2011 7:31 am
Just saw this on Hobbyspace.com:

Boeing proposes SSTO system for AF RBS program.
"The new issue of Aviation Week has a brief blurb about a Boeing
proposal for the Air Force's Reusable Booster System (RBS) program:
Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11
(subscription required).
Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they
are proposing a 3-4 year technology readiness assessment that would
lead up to a demonstration of a X-37B type of system but would be
smaller. Wind tunnel tests have been completed. Davis says the system
would be a single stage capable of reaching low Earth orbit and, with
a booster, higher orbits. The system would return to Earth as a
glider.
Davis says "that advances in lightweight composites warrant another
look" at single-stage-to-orbit launchers."
http://www.hobbyspace.com/nucleus/index ... emid=30110

I don't have a subscription to AV Week. If anyone does perhaps they
could look this up.
I'm curious about the statement it would be "smaller" than the X-37B.
I did some preliminary calculations that if you switched to kerosene
fuel and a high efficiency engine such as the NK-33, and filled every
scrap of internal volume with propellant, then a vehicle twice the
size of the X-37B could be SSTO. I would be surprised they are able to
get it to work with a smaller vehicle than the X-37B.
Perhaps they mean it would be smaller than the booster, Atlas V, and
X-37B system, as the Atlas V weighs upwards of 300,000 kg.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Thu Jun 16, 2011 1:26 pm
ckpooley wrote:
To R Clark, June 8, others:

SSTO, though just physically possible just makes no sense (note that it has not been done). The high mass ratio and that most of the aceleration is done in vacuum requires a stage that is not suited for ascent from sea level then ascend vertically through the atmosphere.
Great compromise to the engines, or resort to strange, yet unused designs like aerospikes would have to be resorted to.
Two stages to orbit makes sense. For maximum performance, for small launchers at least, would have the 1st stage act as a TAV (transatmospheric vehicle), and use atmosphere optimized engines, and enclose the upper stage(s).
Then the upper stage(s) need not be aerodynamic, and with a lofted trajectory could have a greater portion of the 2nd stage thrust directed nearly horizontally. If it were to start with a 20 deg pitch, 94% of the thrust is the horizontal co0mponent, and 34% vertical component helps support the stage during the earlier part of the burn, where total acceleration might be only 1 g or so.
The first stage can be engineered to return to launch site fairly easily, as the staging could take place at a fairly low velocity at an altitude high enough for a vacuum adapted 2nd stage engine to work well.
This is the design for the 100-200 kg GLOW Microlaunchers, and it works well for larger ones also.


Using SSTO's has been criticized on the grounds that you can loft more payload to orbit with staging. However here is a very key point:

Even if you want to increase your payload to orbit by using staging you are still better off using a staging method where each individual stage is SSTO-capable.

I'm referring to the staging method that uses parallel staging in concert with cross-feed fueling. These two techniques together are known to increase your payload to orbit over the usual method of staging with just an upper stage placed above a first stage.
And even if you are already using parallel staging, by using cross-feed fueling you can still increase your payload. For instance SpaceX found it could increase its payload with the Falcon Heavy using cross-feed fueling by 50% over that of the original Falcon 9 Heavy.
So to get greater payload with the X-33, or VentureStar, or X-37B, use two copies of them attached together firing in parallel at the start, called "bimese" staging, so that one of them peels off and returns to the launch site after staging altitude is reached, and the other continues on to orbit with the payload.
I believe it's a general principle you can loft more payload this way. Note with this method you have to use engines optimized for use from sea level to vacuum, even for what is the "upper stage".
For this reason you really should use altitude compensation methods on such a system. But I still think you can increase your payload to orbit using this method even with standard bell nozzles, without altitude compensation.
Try it for your two-system. Note you will be using the same engines on your upper stage now as on the lower stage, not vacuum optimized engines, because both stages will be copies of each other.
Make sure also you take into account you are using cross-feed fueling. This means that at lift-off and up to the staging point both stages are firing, with all the propellant coming from the tanks in only one of the stages.
This ensures that after the staging point when one of the stages is jettisoned that the remaining stage continues on with a full propellant load.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Thu Jun 16, 2011 3:18 pm
Rob's remarks re parallel staging being best can only apply to large rockets, where all stages deal with aerodynamics. For smaller rockets it would be better for the upper stage(s) be inside the first stage, and that to be optimized for taking the upper stages to vacuum. The upper stages being free of aerodynamic restraints can then be lighter, more efficient.

The first stage then could be design for durability, re-useability. It can be structurally heavier, as it would not be called upon to go very fast and could have a low mass ratio. MR for the Microlauncher stage one is to be about 2.5.

And, again, an engine optimized for vacuum cannot work well at sea level. The Atlas center engine was an attempt, and it was a compromise.

So, 2 stage for LEO; 3 for escape makes sense.


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