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From: Robert Clark on 8 Aug 2010 11:10
On Aug 8, 9:21 am, Sam Wormley <sworml...(a)gmail.com> wrote:
> On 8/8/10 3:41 AM, Robert Clark wrote:
>
>
> > On Aug 7, 4:59 pm, Sam Wormley<sworml...(a)gmail.com>  wrote:
> >> On 8/7/10 1:41 PM, Robert Clark wrote:
>
> >>>    The point I have been arguing in this thread is that not only is SSTO
> >>> technically doable, it is in fact *easy*  if you use DENSE propellants
> >>> This point is made in this article:
>
> >>     You have you ask yourself the question: Why is it not routinely used?
>
> >   Good question. There are two reasons for this. One, the Isp issue.
> > The Isp is the most important parameter for creating a launch vehicle.
> > A high Isp means you need less fuel for the same mass vehicle. For a
> > SSTO you would think then you should use the fuel that has the highest
> > Isp. Hydrogen-LOX provides the highest Isp among the practical
> > chemical propellants.
> >   However, ironically hydrogen-LOX turns out to be the *worst* liquid
> > fuel combination you can use. The reason is its low density. That low
> > density drives it to have much worse engine T/W and propellant tankage
> > ratios than dense hydrocarbon fuels.
> >   While LH2/LOX does have an a 1.5 times better Isp than kerolox, it
> > has a 2 times worse engine T/W ratio and *3 times* worse tankage
> > ratio. The mass of the tanks and the mass of the engines are the two
> > biggest components of the dry mass of the vehicle.  The combination of
> > the reduction in the dry mass of the vehicle you get from these two
> > factors for dense fuels well swamps the Isp advantage of hydrogen.
> > That all NASA proposals for SSTO's focused on using hydrogen
> > ironically meant you were using the *hardest* fuel combination to
> > use.
> >   Second, I said getting a SSTO using dense propellants is actually
> > easy. But to do it you have to use BOTH the most well weight optimized
> > vehicles AND the most efficient engines. However, the Russians had the
> > most efficient kerosene engines, but rather poorly weight optimized
> > vehicles. And the Americans had the best weight optimized vehicles,
> > but their kerosene engines were of rather poor efficiency. (The
> > Americans did have very good hydrogen-fueled engines such as the
> > shuttle SSME's, but as I said hydrogen is the *worst* fuel to use for
> > a SSTO.)
>
> >     Bob Clark
>
>    You never addressed reasons as to why kerosene is not routinely
>    used.

Well kerosene of course is routinely used for multistaged rockets. It
was used for the first stage of the Saturn V. It has always been used
for the Soyuz launchers. And it is currently used for the first stage
of the Atlas launchers.
I presumed your question was why it is not routinely used for SSTO's.
As I said the presumption has been that a SSTO has to be hydrogen-
fueled. Ironically doing that way makes the problem *much* harder.
Also, you need both the lightweight vehicles the Americans have and
the high performance kerosene engines the Russians have. Using both
would not have been possible before the end of the Cold War in the
early 90's. Now that is possible it is clear this is the approach to
take to achieve near term SSTO's.
I want to reiterate another point I mentioned in a prior post.
Accepting that you can get SSTO-capability using dense propellants is
now *essential*. For by using this fact it becomes easy and cheap to
get super heavy lift capacity by using these SSTO's as the components
of a multistage system, and indeed at a fraction of the cost of the
multi-billion dollar proposals being presented as solutions to the
question of super heavy lift launchers.


Bob Clark
From: Sam Wormley on 8 Aug 2010 14:59

See: http://en.wikipedia.org/wiki/Rocket_propellant#Mixture_ratio
From: Pat Flannery on 8 Aug 2010 21:24
On 8/8/2010 6:47 AM, Dr.Smith wrote:


> Atlas SSTO link:
>
> http://en.wikipedia.org/wiki/Project_SCORE

I wouldn't call it SSTO; it drops off two of its three engines on the
way up.
I once did the math on whether it could reach orbit with the two booster
engines still attached, but it didn't work.

Pat


From: Robert Clark on 10 Aug 2010 23:38
On Aug 7, 2:41 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote:
>  The point I have been arguing in this thread is that not only is SSTO
> technically doable, it is in fact *easy* if you use DENSE propellants
> This point is made in this article:
>
> Single-stage-to-orbit.
> "The early Atlas rocket is an expendable SSTO by some definitions. It
> is a "stage-and-a-half" rocket, jettisoning two of its three engines
> during ascent but retaining its fuel tanks and other structural
> elements. However, by modern standards the engines ran at low pressure
> and thus not particularly high specific impulse and were not
> especially lightweight; using engines operating with a higher specific
> impulse would have eliminated the need to drop engines in the first
> place.
> The first stage of the Titan II had the mass ratio required for single-
> stage-to-orbit capability with a small payload. A rocket stage is not
> a complete launch vehicle, but this demonstrates that an expendable
> SSTO was probably achievable with 1962 technology."http://en.wikipedia.org/wiki/Single-stage-to-orbit#Examples
>
>  The Titan II first stage did have SSTO capability using dense
> propellants. And I'll show here the kerosene-fueled Atlas III does
> have SSTO capability if switched to using the lighter NK-33.
>  The original Atlas from the 1960's was close to being SSTO capable.
> It was able to be highly weight-optimized because it used what is
> called pressure-stabilized or "balloon tanks". These were tanks of
> thinner wall thickness than normal and were able to maintain their
> structure in being pressurized. The wall thickness was so thin that
> they could not stand alone when not filled with fuel. To be stored the
> tanks had to be filled with an inert gas such as nitrogen, otherwise
> they would collapse under their own weight.
> The Atlas III first stage also uses balloon tanks and a common
> bulkhead design, used effectively by the SpaceX Falcon launchers to
> minimize weight. The Falcons probably are able to get the good weight
> optimization comparable to that of the Atlas launchers without using
> balloon tanks because their tanks are made of aluminum instead of the
> steel used with the Atlas tanks. The Atlas launchers might be able to
> weight-optimize their tanks even further by using aluminum for their
> balloon tanks, but there may be structural reasons that for balloon
> tanks steel has been preferred.
> The specifications for the Atlas III first stage are given on this
> Astronautix.com page:
>
> Atlas IIIAhttp://www.astronautix.com/stages/atlsiiia.htm
>
> The gross mass is given as 195,628 kg and the empty mass is given as
> 13,725 kg, resulting in a propellant mass of 181,903 kg. The Atlas III
> uses an RD-180 engine:
>
> RD-180http://www.astronautix.com/engines/rd180.htm
>
> The Atlas III first stage is actually somewhat overpowered with the
> RD-180, as evidenced by the fact that Atlas V first stage carrying 50%
> more propellant is still able to use the RD-180. For an SSTO the
> weight of the engines is a major factor that has to be tailored to the
> size of the vehicle. A engine of greater power may be unsuitable for
> the SSTO purpose simply because the larger than needed engine weight
> may prevent the required mass ratio to be SSTO.
> So again I'll use NK-33's two this time for the engines:
>
> NK-33.http://www.astronautix.com/engines/nk33.htm
>
> Then the engine weight is reduced from 5,393 kg to 2,444 kg. This
> brings the dry mass to 10,776 kg, and the gross mass is now 192,679
> kg. So the mass ratio is 17.9.
>  Using aerospike nozzles or other altitude compensation methods on the
> NK-33 we might be able to get the vacuum Isp to increase to the 360 s
> reached by other vacuum optimized high performance Russian engines.
> For the average Isp over the flight we'll use the value 338.3 s
> estimated for high performance kerolox engines using altitude
> compensation given in table 2 of this report:
>
> Alternate Propellants for SSTO Launchers.
> Dr. Bruce Dunn
> Adapted from a Presentation at:
> Space Access 96
> Phoenix Arizona
> April 25 - 27, 1996
> http://www.dunnspace.com/alternate_ssto_propellants.htm
>
>  For the delta-V to orbit use 8,900 m/s, approx. 300 m/s less than
> that required for hydrogen fueled rockets due to the reduction in
> gravity loss using dense propellants:
>
> Single-stage-to-orbit.
> 4 Dense versus hydrogen fuels.
> http://en.wikipedia.org/wiki/Single-stage-to-orbit#Dense_versus_hydro...
>
>  Then this would allow a payload of 2,500 kg:
>
> 338.3*9.8ln(1 + 181,903/(10,776 + 2500)) = 8,911 m/s.
>
> Now let's calculate the payload for these two Atlas III's mated bimese
> fashion and using cross-feed fueling:
> with a payload of 19,000 kg, we get a first stage delta-V of
> 338.3*9.8ln(1 + 181,903/(2*10,776 + 181,903 + 19,000)) = 1,981 m/s,
> and a second stage delta-V of 360*9.8ln(1 + 181,903/(10,776 + 19,000))
> = 6,919 m/s for a total delta-V of 8,900 m/s.
>
>  But there are other hydrocarbon fuels that would give even better
> performance, for instance, methylacetylene. Table 2 in Dunn's report
> gives its average Isp as 352 s with altitude compensation. The max
> theoretical vacuum Isp is given as 391.1 s. High performance engines
> can get upwards of 97% of the theoretical value. So we'll take the
> vacuum Isp with the methylacetylene fuel as 380 s.
>  For the SSTO version we would get a payload of 4,000 kg:
> 352*9.8ln(1 + 181903/(10,776 + 4,000)) = 8,929 m/s.
>  For the bimese, cross-feed fueled version, estimate the payload as
> 23,000 kg:
>
> 352*9.8ln(1+181,903/(2*10,776 + 1*181,903 + 23,000)) = 2,033 m/s for
> the first stage delta-V and 380*9.8ln(1 + 181,903/(10,776 + 23,000)) =
> 6,904 m/s for the second stage, for a total of 8,937 m/s.
>  The "Atlas IIIA" page on Astronautix.com gives the cost for the Atlas
> III first stage as $50,000,000. Two NK-33's are actually slightly
> cheaper than the one RD-180 used. So keep the price of the NK-33 powered
> version the same and take the price of the bimese version as twice as
> high.  Then the cost to orbit per kilo would be $100,000,000/23,000 =
> $4,350/kg, about half-current launch rates.
>


So we saw the Atlas III switched to a pair of lighter but still high
performance NK-33 engines becomes SSTO with significant payload.
This is the import of the report of Dr. John C. Whitehead:

Single Stage To Orbit Mass Budgets Derived From Propellant Density and
Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference
Lake Buena Vista, FL July 1-3, 1996
http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf

that it is easier to produce a SSTO with dense propellants rather than
using hydrogen fuel.
Indeed, this is a common state of affairs: if you use highly weight
optimized stages, such as with the Atlas's "balloon tanks", AND you
also use high performance kerosene engines, then what you will wind up
with
will be a SSTO.
Further examples are provided by earlier, smaller Atlas versions. Use
for instance the Atlas I first stage, also called the "sustainer"
stage:

Atlas-I, Design
http://www.b14643.de/Spacerockets_2/United_States_3/Atlas-Centaur/Design/AtlasG_I.htm

The first stage gross mass is given as 143,200 kg, and the propellant
mass as 137,530 kg, giving a dry mass of 5,670 kg. It would appear
this Atlas first stage had a mass ratio of over 25 to 1(!) However,
the design of these early Atlas versions was for most of the lift off
thrust to be provided by the drop off booster engines. The engine
that came with the first stage was smaller and did not have sufficient
thrust to lift the first stage on its own.
So we'll replace it with one NK-33 engine, with a mass of 1,222 kg:

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

The mass of the engine that came with the Atlas I first stage was 460
kg:

LR-105-7.
http://www.friends-partners.org/partners/mwade/engines/lr1057.htm

So replacing it with the NK-33 adds 762 kg to the dry mass to bring
it to 6,432 kg kg, and the gross mass becomes 143,962 kg. This is
still within the lift capacity of a single NK-33, and the mass ratio
is still a very good 22.4.
To estimate payload, again use the Isp values given in Table 2 in
Dunn's report. The average Isp for kerolox with altitude compensation
is given as 338.3 s, and for high energy density methylacetelyne its
352. Then for kerolox we could get a payload of 3,600 kg:
338.3*9.8ln(1 + 137,530/( 6,432 + 3,600)) = 8,900 m/s, and for
methylacetelyne its 4800 kg: 352*9.8ln(1 + 137,530/( 6,432 + 4,800)) =
8,910 m/s.

The John Whitehead article showed the mathematics for why SSTO's with
dense propellants are achievable. And real world examples bear this
out if you use both weight optimized vehicles and high performance
kerosene engines at the same time.



Bob Clark
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