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From: Pat Flannery on 17 Jul 2010 01:12
On 7/16/2010 10:24 AM, Jeff Findley wrote:


> There were a couple of proposals for LOX/kerosene boosters to replace
> the SRB's, but those proposals went nowhere.

Originally, the idea was to use liquid boosters for safety reasons, as
unlike the SRBs they could be shut down and jettisoned if one of them
went out-of-spec, and the orbiter could then hopefully return to the
launch site, either alone or after burning some of the fuel in the ET.
They also looked into a thrust termination system on the SRBs similar to
the one that was going to be used on the manned Titan III's, but the ET
apparently couldn't tolerate the blast effect of the blow-off venting
portals on the SRB nosecone being activated so close to it.
I still like the idea of sticking the Shuttle and ET atop a Saturn V
first stage: http://www.astronautix.com/graphics/s/shusat1c.gif
If something did go wrong with one of the F-1 engines, at least it would
be way behind the orbiter.
Combine that concept with the recoverable S-IC stage proposal, and you
could have had a system as recoverable as the present one that offered
superior safety during ascent due to having the ability to be shut down
in an emergency.

Pat
From: Robert Clark on 24 Jul 2010 20:31
On Jul 16, 1:45 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote:
>
> Anyone know if there has been research on converting the shuttle main
> engines to hydrocarbon fueled? I was annoyed that NASA had earlier
> canceled a program to develop a heavy-thrust hydrocarbon engine after
> the Ares I and V were chosen. We would have a reusable and man-rated
> heavy-thrust kerosene engine *now* if it weren't for that.
> The SSME's have to operate under severe tolerances using cryogenic
> hydrogen since the liquid hydrogen is so cold yet LH2/LOX burns at
> such high temperature. I would think using kerosene/LOX for instance
> would put less severe conditions on the engine operation.
> Note that other liquid hydrogen engines have been successfully run on
> other fuels under test conditions:
>
> The RL10 (Bruce Dunn; Gary Hudson; Henry Spencer)http://yarchive.net/space/rocket/rl10.html
>
> And some dense propellant engines have been tested to run on cryogenic
> hydrogen:
>
> LR-87 LH2http://www.astronautix.com/engines/lr87lh2.htm
>


Found this after searching on Astronautix.com:

RD-0120.
"Engine Model: RD-0120-CH. Manufacturer Name: RD-0120-CH. Designer:
Kosberg. Propellants: Lox/LCH4. Thrust(vac): 1,576.000 kN (354,298
lbf). Isp: 363 sec. Mass Engine: 2,370 kg (5,220 lb). Chambers: 1.
Chamber Pressure: 172.50 bar. Oxidizer to Fuel Ratio: 3.40. Thrust to
Weight Ratio: 67.80. Country: Russia. Status: Design concept 1990's.
Proposed variant of the RD-0120 engine using liquid methane instead of
hydrogen as propellant."
http://www.astronautix.com/engines/rd0120.htm

The RD-0120 was the hydrogen fueled engine used on the Russian Energia
heavy lift booster, which lifted the Russian Buran space shuttle for
instance. I can't tell from this description though if it was actually
tested with liquid methane or if these were only theoretical studies.
After searching on the NASA Technical Report server I found some
theoretical studies that suggest that the SSME could be converted to
hydrocarbon-fueled at relatively low cost (compared to developing a
new engine.)

Booster engines derived from the Space Shuttle Main Engine.
Sobin, A. J.; Poynor, S. P.; Cross, E
"By using a majority of the current SSME engine components for the LOX/
RE-1 booster engine, engine development time and cost can be
significantly reduced compared to the development of a new engine."
Propulsion Conference, 13th, July 11-13, 1977, Orlando, FL
http://ntrs.nasa.gov/search.jsp?N=0&Ntk=all&Ntx=mode%20matchall&Ntt=19770059130
[abstract only]

Tripropellant engine study.
Wheeler, D. B.; Kirby, F. M.
NASA-CR-150808; RI/RD78-215
"SUMMARY.
"The results of these studies have shown that the conversion of an
SSME engine to
a high chamber pressure, dual-mode fuel engine will require major
modifications
to the hardware and/or the addition of a significant number of new
engine cowponents.
However, the study has shown numerous possibilities for the use of
SSME
hardware derivatives in a single-mode LOX/hydrocaxbon engines. It was
also
shown that a reduced chamber pressure version of a staged combustion
SSME is
operationally feasible using the existing fuel-rich preburners and
main chamber
injectors. Certain turbomachinery modifications or additions are
required for
a total low chamber pressure ( 2300 psia) engine system. This study
also has
shown that the engine system concepts applicable to the dual-mode
systems are
somewhat narrowed since the operational constraints of two systems
must be
considered."
http://hdl.handle.net/2060/19780024238 [full text, 145 pages]

Another possibility might be to adapt the hydrogen-fueled aerospike
engines intended for the VentureStar to hydrocarbon-fueled. This
theoretical study from 1977 was on the possibility that an aerospike
engine of the linear configuration later adopted for the VentureStar
could be dual-fueled, i.e., running on both hydrocarbon and hydrogen:

Linear aerospike engine study.
Diem, H. G.; Kirby, F. M.
NASA-CR-135231; RI/RD77-170
http://hdl.handle.net/2060/19780003139 [full text, 246 pages]

This would have the advantage that it would already have altitude
compensation. If the dual-fuel modes are workable this would also
increase performance.
This study was primarily on dual-fuel operation but did also study
hydrogen only operation. It might be useful to compare the predicted
hydrogen only operation with the performance actually found with the
aerospike engines created for the X-33 sub-scale demonstrator. If the
measured performance does correspond to the predicted values that
would give confidence that the dual-fuel version would also be close
to the predicted values.


Bob Clark
From: Robert Clark on 2 Aug 2010 20:34
Here are some possibilities for lower cost super heavy lift launchers,
in the 100,000+ kg payload range. As described in this article the
proposals for the heavy lift launchers using kerosene-fueled lower
stages are focusing on using diameters for the tanks of that of the
large size Delta IV, at 5.1 meters wide or the even larger shuttle ET,
at 8.4 meters wide:

All-Liquid: A Super Heavy Lift Alternative?
by Ed Kyle, Updated 11/29/2009
http://www.spacelaunchreport.com/liquidhllv.html

The reason for this is that it is cheaper to create new tanks of the
same diameter as already produced ones by using the same tooling as
those previous ones. This is true even if switching from hydrogen to
kerosene in the new tanks.
However, I will argue that you can get super heavy lift launchers
without using the expensive upper stages of the other proposals by
using the very high mass ratios proven possible by SpaceX with the
Falcon 9 lower stage, at above 20 to 1:

SPACEX ACHIEVES ORBITAL BULLSEYE WITH INAUGURAL FLIGHT OF FALCON 9
ROCKET.
Cape Canaveral, Florida – June 7, 2010
"The Merlin engine is one of only two orbit class rocket engines
developed in the United States in the last decade (SpaceX’s Kestrel is
the other), and is the highest efficiency American hydrocarbon engine
ever built.
"The Falcon 9 first stage, with a fully fueled to dry weight ratio of
over 20, has the world's best structural efficiency, despite being
designed to higher human rated factors of safety."
http://www.spacex.com/press.php?page=20100607

We will use tanks of the same size as these other proposals but will
use parallel, "bimese", staging with cross-feed fueling. This method
uses two copies of lower stages mated together in parallel with the
fueling for all the engines coming sequentially from only a single
stage, and with that stage being jettisoned when it's expended its
fuel. See the linked image below for how parallel staging with cross-
feed fueling works.
Do the calculation first for the large 8.4 meter wide tank version. At
the bottom of Kyle's "All-Liquid: A Super Heavy Lift Alternative?"
article is given the estimated mass values for the gross mass and
propellant mass of the 8.4 meter wide core first stage. The gross mass
of this single stage is given as 1,423 metric tons and the propellant
mass as 1,323 metric tons, so the empty mass of the stage would be
approx. 100 metric tons (a proportionally small amount is also taken
up by the residual propellant at the end of the flight.) Then the mass
ratio is 14 to 1. However, the much smaller Falcon 9 first stage has
already demonstrated a mass ratio of over 20 to 1.
A key fact about scaling is that you can increase your payload to
orbit more than the proportional amount indicated by scaling the
rocket up. Said another way, by scaling your rocket larger your mass
ratio in fact gets better. The reason is the volume and mass of your
propellant increases by cube of the increase and key weight components
such as the engines and tanks do also, but some components such as
fairings, avionics, wiring, etc. increase at a much smaller rate. That
savings in dry weight translates to a better mass ratio, and so a
payload even better than the proportional increase in mass.
This is the reason for example that proponents of the "big dumb
booster" concept say you reduce your costs to orbit just by making
very large rockets. It's also the reason that for all three of the
reusable launch vehicle (RLV's) proposals that had been made to NASA
in the 90's, for each them their half-scale demonstrators could only
be suborbital.
Then we would get an even better mass ratio for this "super Evolved
Atlas" core than the 20 to 1 of the Falcon 9 first stage, if we used
the weight saving methods of the Falcon 9 first stage, which used
aluminum-lithium tanks with common bulkhead design. It would also work
to get a comparable high mass ratio if instead the balloon tanks of
the earlier Atlas versions prior to the Atlas V were used.
So I'll use the mass ratio 20 to 1 to get a dry mass of 71.15 mT, call
it 70,000 kg, though we should be able to do better than this. We'll
calculate the case where we use the standard performance parameters of
the RD-180 first, i.e., without altitude compensation methods. I'll
use the average Isp of 329 s given in the Kyle article for the first
leg of the trip, and 338 s for the standard vacuum Isp of the RD-180.
For the required delta-V I'll use the 8,900 m/s often given for
kerosene fueled vehicles when you take into account the reduction of
the gravity drag using dense propellants. Estimate the payload as 115
mT. Then the delta-V for the first leg is 329*9.8ln(1 + 1,323/(2*70 +
1*1,323 + 115)) = 1,960 m/s. For the second leg the delta-V is
338*9.8ln(1 + 1,323/(70 + 115)) = 6,950 m/s. So the total delta-V is
8,910 m/s, sufficient for LEO with the 115 mT payload, by the 8,900 m/
s value I'm taking here as required for a dense propellant vehicle.
Now let's estimate it assuming we can use altitude compensation
methods. We'll use performance numbers given in this report:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm

In table 2 is given the estimated average Isp for a high performance
kerolox engine with altitude compensation as 338.3 s. We'll take the
vacuum Isp as that reached by high performance vacuum optimized
kerolox engines as 360 s. Estimate payload as 145,000 kg. For the
first leg, the delta-V is 338.3*9.8ln(1 + 1,323/(2*70 + 1*1,323 +
145)) = 1,990 m/s. For the second leg the delta-V is 360*9.8ln(1 +
1,323/(70 + 145)) = 6,940 m/s, for a total delta-V of 8,930 m/s,
sufficient for orbit with the 145,000 kg payload.
Now we'll estimate the payload using the higher energy fuel
methylacetylene. The average Isp is given as 352 s in Dunn's report.
The theoretical vacuum Isp is given as 391 s. High performance engines
can get quite close to the theoretical value, at 97% and above. So
I'll take the vacuum Isp as 380 s. Estimate the payload as 175,000 kg.
Then the delta-V over the first leg is 352*9.8ln(1 + 1,323/(2*70 +
1*1,323 + 175)) = 2,040 s. For the second leg the delta-V will be
380*9.8ln(1 + 1,323/(70 + 175)) = 6,910 s, for a total delta-V of
8,950 m/s, sufficient for orbit with the 175,000 kg payload.

bimese Falcon 9 launcher
http://i27.tinypic.com/2yxn2oz.jpg

Bob Clark


From: Robert Clark on 2 Aug 2010 20:44
You can get really large payloads with the 8.4 meter wide super
"Evolved Atlas" stage by using parallel, "trimese", staging with cross-
feed fueling. This would use now three copies of the lower stages
mated together in parallel with the fueling for all the engines coming
sequentially from only a single stage, and with that stage being
jettisoned when its fuel is expended.
Again we'll calculate first the case where we use the standard
performance parameters of the RD-180, i.e., without altitude
compensation methods. I'll use the average Isp of 329 s given in the
Kyle article for the first leg of the trip, and for the required delta-
V, again the 8,900 m/s often given for kerosene fueled vehicles when
you take into account the reduction of the gravity drag using dense
propellants. Estimate the payload as 200 mT. Then the delta-V for the
first leg with all three super Evolved Atlas's attached will be
329*9.8ln(1+1,323/(3*70 + 2*1,323 + 200)) = 1,160 m/s. For the second
leg we'll use the vacuum Isp of 338 s, then the delta-V will be
338*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 200)) = 1,940 m/s. And for the
final leg 338*9.8ln(1 + 1,323/(70 +200)) = 5,880 m/s. So the total
delta-V is 8,980 m/s, sufficient for orbit with the 200,000 kg
payload.
Now let's estimate it assuming we can use altitude compensation
methods. We'll use performance numbers given in this report:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm

In table 2 is given the estimated average Isp for a high performance
kerolox engine with altitude compensation as 338.3 s. We'll take the
vacuum Isp as that reached by high performance vacuum optimized
kerolox engines as 360 s. Estimate the payload now as 250 metric tons.
Then the delta-V during the first leg will be 338.3*9.8ln(1+1,323/
(3*70 + 2*1,323 + 250)) = 1,180 m/s. For the second leg the delta-V
will be 360*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 250)) = 2,020 m/s. For
the third leg the delta-V will 360*9.8ln(1 + 1,323/(70 + 250)) = 5,770
m/s. So the total will be 8,970 m/s, sufficient for orbit with the
250,000 kg payload.
Now we'll estimate the payload using the higher energy
methylacetylene. The average Isp is given as 352 s in Dunn's report.
The theoretical vacuum Isp is given as 391 s. High performance engines
can get quite close to the theoretical value, at 97% and above. So
we'll take the vacuum Isp as 380 s. Estimate the payload now as 300
mT. The first leg delta-V will now be 352*9.8ln(1 + 1,323/(3*70 +
2*1,323 +300)) =1,210 m/s. For the second leg 380*9.8ln(1 + 1,323/
(2*70 + 1*1,323 + 300)) = 2,080 m/s. For the third leg 380*9.8ln(1 +
1,323/(70 + 300)) = 5,660 m/s. So the total is 8,950 m/s, sufficient
for orbit with the 300,000 kg payload.

This trimese version of the vehicle would be huge however. For
instance it would weigh more than the Saturn V. One of the big cost
factors for the development of some of the super heavy lift launchers
is that they are so heavy they would require the construction of new
and expensive launch platforms. Undoubtedly, the bimese version would
be the one to be built first if this launch system is selected.



Bob Clark
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