Stephen Horgan,
Apparently our American Usenet folks (aka rusemasters) don't actually
give a tinkers damn about your reliable and efficient Ariane-5 being at
best 82.3:1 for getting whatever's 9.6 tonnes into GSO. So what's so
great about that?
I've relocated somewhat interesting Saturn-V and Apollo mission numbers
that are even more impressive than I'd thought, as for appreciating
what their LXO/RP-1 first stage managed to get off the ground, as
supposedly being much better Isp than having to use modern day SRBs
that are apparently nowhere as thrust/inert mass efficient as per those
extremely old LOX/RP-1 methods.
At 55.43:1 represents perhaps the "smallest orbital launch vehicle"
that should manage to get fairly compact, and supposedly smaller yet
with using SRBs instead of a massive LOX/RP-1 first stage. Meaning that
a 10X microsatellite collective deployment worth of 10 each 10 kg
satellites as a deployment group representing 100 kg intended for
orbiting the moon, at perhaps their starting in at only 25 km off the
deck, shouldn't involve a liftoff mass of at most not much greater than
5 tonnes. I'm thinking that a modern SRB instead of LOX/RP-1 and a good
usage of nifty composites should eventually represent much less
inert/dry mass, whereas getting such things down to as good as 25:1,
thereby a 2.5 tonne rocket and payload at liftoff should manage to get
those ten microsatellites into orbiting and subsequently a few
semi-soft landings upon our moon. Is that good news or what?
http://space.kursknet.ru/cosmos/engl...hines/ap17.sht
Fully loaded spacecraft mass (including crew and the return ascent
stage): 52,740 kg
For some reason this Russian information source has our Saturn-V inert
mass a wee bit on the high side (must have included the spacecraft
portion): MO/inert mass = 1,057.942 tonnes
http://www.nasm.si.edu/collections/i...lo/saturnV.htm
The complete assembly including the Apollo spacecraft and the Saturn
launch vehicle stood 363 feet tall (110.6 meters) and weighed over 6
million pounds (2.7 million kg).
http://www.nasm.si.edu/collections/i...RES/Fig49a.jpg
Saturn-V First stage
LXO: 3,258,280 lbs = 1,477.93 tonnes
RP-1: 417,334 lbs = 189.30 tonnes
? First stage mix at 7.8:1 = thrust: 7,680,982 lbs = 3,484 tonnes
? First stage inert/dry mass = 294,200 lbs = 133.447 tonnes
Second stage
LXO: 828,114 lbs = 375.626 tonnes
LH2: 158,231 lbs = 71.772 tonnes
Third stage
LXO: 190,785 lbs = 86.539 tonnes
LH2: 43,452 lbs = 19.710 tonnes
-------------------------------------
Total fuel load = 2,220.877 tonnes
According to yet another official source; Apollo-17 launch mass:
2,923,387 kg = 2,923.387 tonnes, as opposed to the previous 2.7 million
kg, which beggs to ask why these mass related numbers are all over the
map.
Inert/dry mass = 2,923.387 - 2,220.877 = 702.51 tonnes
This otherwise gives us their impressive payload ratio = 2,923.387 /
52.74 = 55.43:1
I'm to guess that because 55.43:1 is soooo gosh darn impressive that
folks simply don't seem to want to discuss this achievement. Perhaps
it's because within the past 4 decades there has been absolutely
nothing developed that comes close after taking better than twice that
amount for accomplishing a moon shot. Even the very latest and bestest
capability of a fully loaded 790 tonne Ariane-5 that utilizes nifty
SRBs only has a capability of getting 9.6 tonnes into GSO, and even
achieving that much is 82.3:1.
Here's some basic oxidiser/fuel formulation info.
LXO/RP-1 Isp = LOX/Kerosene. Isp: 353.00 vac. Isp: 300.00 sl.
H2O2/RP-1 Isp = H2O2-95%/RP-1. Isp: 319.00 vac. Isp: 273.00 sl.
98% H2O2/RP-1 has a maximum Isp of 362, slightly better than LOX/RP-1
of 353.
What's wrong with this pictu
LXO/RP-1 Optimum Oxidiser to Fuel Ratio: 2.56:1, which beggs yet
another question as to why the Saturn-V first stage took such a ratio
of 7.8:1, as that seems a wee bit LOX rich and/or RP-1 lean. If
anything, you'd think running a bit RP-1 rich would have been the case,
or perhaps that's where the smoke and mirrors of how they'd managed to
outperform SRBs by such a huge factor.
http://www.friends-partners.org/part...s/h2oosene.htm
As a comparison; H2O2-98%/Kerosene has an optimum Oxidiser to Fuel
Ratio of 7.07
Density: 1.31 g/cc. Temperature of Combustion: 2,975.00 deg K. Ratio of
Specific Heats: 1.20. Characteristic velocity c: 1,665 m/s. Isp
Shifting: 276. Isp Frozen: 270. Pp Isp Shifting: 362. Mol: 22.00 M.
At any rate, if we are to take the 55.43:1 as being gospel, and applied
that formula as to getting microsatellites efficiently off and running
their missions around our moon, as such that's actually not half bad
(key word being "half"), whereas in fact it's offering a downright
impressive 3-stage overall ratio that shouldn't be ignored. Of course,
all cloak and dagger kidding aside, perhaps using first stage SRBs
should actually be even better.
There's also a bit more interesting H2O2 usage as extra bang/kg to
being had:
H2O2/propargyl alcohol adds 40% to the H2O2/RP-1 payload capability
H2O2/methylacetylene adds 16% to the H2O2/RP-1 payload capability
Of course H2O2 is essentially made primarily from plain old water plus
good amount of applied electrical energy for adding that extra amount
of Oxygen, of which we all know that Earth has way more than it's fair
share of water, especially ever since our artificial induced pollution
and subsequent global warming has been running our environment amuck.
Methyl acetylene alcohol is made from just about anything that grows,
including all sorts of weeds and certainly the remains of food stocks
that are so fiber based that cows can't hardly digest, but we certainly
could just as well make such remainders into methyl acetylene alcohol.
-
Brad guth