isp of LOX/Acetylene or LOX/Cyanogen or LOX/carbon subnitride

On Wednesday, April 16, 2014 3:17:54 PM UTC+12, Orval Fairbairn wrote:
In article ,

William Mook wrote:



Beryllium metal be treated as a fuel grain in a hybrid liquid rocket, similar

to this;



http://upload.wikimedia.org/wikipedi...ybrids_big.png



Liquid oxygen and beryllium metal form very high exhaust speed rockets with

moderately high densities.



Liquid oxygen and aluminium metal form high exhaust velocities with high

propellant densities.



LOX/Be/H2 tripropellant has an even higher Isp, but how do you meter teh

Be?

To approaches,

(1) Nanoparticles of Beryllium colloidally suspended in the hydrogen,
(2) Hybrid rocket that has LOX/H2 injected normally, and Be is the fuel grain

http://upload.wikimedia.org/wikipedi...ybrids_big.png

Checking out the Lithium/Fluorine/Hydrogen combination;

The hydrogen was added to reduce temperature so the engine wouldn't burn out.

http://www.flightglobal.com/pdfarchi...0-%200069.html

Lithium-Fluoride by itself has -616 kJ/mole - Li has an atomic weight of 7 and Fluorine has an atomic weight of 19 - a total of 26 - so this is 23.69 MJ/kg.

All of this converted to kinetic energy implies an exhaust velocity of 6.88 km/sec. How much gets converted depends on expansion ratio and that depends on chamber pressure versus exist pressure. 85% efficiency at sea level, 92% efficiency in vacuum implies 6.35 km/sec at sea level and 6.60 km/sec in vacuo.

Lithium is 7/26th of the total propellant weight and Fluorine is 19/26th of the total propellant weight in a stoichiometric mixture. This means that for each kilogram of mixture 269.2 grams is Lithium with a density of 0.53 kg/litre whilst 730.8 grams of Fluorine make up the balance at 1.50 kg/litre. This means that 508 cc of lithium along with 487.2 cc of fluorine are combined for each kilogram of propellant. A total of 995.2 cc yielding an averaged density of 1.005 kg/litre. Slightly higher than water.

The addition of hydrogen radically reduces this density, and the energy per unit mass flow - which reduces the exhaust velocity accordingly.

The addition of 4 H2O molecules to the exhaust means that 4x 285.8 kJ/mole is released and 4 moles mass 4x 18 grams = 72 grams so, we have a specific energy of 15.87 MJ/kg of propellant. This is 5.60 km/sec ideally, 4.78 km/sec at sea level, and 5.18 km/sec in vacuum.

Here we have 2/18th of the total as hydrogen and 16/18th of the total as oxygen. The density of hydrogen is 0.07 kg/litre whilst the density of oxygen in liquid form is 1.14 kg/litre. So, the volume of hydrogen to make 2/18th kg is 1.59 litres whilst the volume of oxygen to make 16/18th kg is 0.78 litres. A total of 2.37 litres per kg a density of 0.42 kg/litre. Actual engines run at 5.5 to 1 instead of stoichiometric ratios because of temperature issues again at pressures that are reasonable with expansion ratios needed to be efficient at sea level.

Running at a 5.5 to 1 Oxygen Fuel ratio means that 312.5 grams of excess hydrogen is carried along and exhausted while 97.49 MJ is released for every 6.5 kg of propellant - a specific energy of 15 MJ/kg. These translate to an idealized exhaust velocity of 4.65 km/sec and 5.04 km/sec

Now, these exhaust velocities are higher than reported. That's because efficiencies are less than 85% and 92% respectively for macro-engines. This is due to unfavourable scaling of large engines wrt heat transfer.

However, for micro-engines, these efficiencies are ROUTINELY attained due to their rapid cooling and small size. Since microengines are made on wafers that cost $15 per square inch, and thrust is 50 lbs per square inch, whilst massing 0.8 ounces per square inch, cost is $0.30 per pound of lift and thrust to weight is 1,000 to 1!! This means that we can over-build engine capacity which gives us far higher reliability on these sort of thrust surfaces.

http://uclarocketproject.com/?p=1

A kilogram of Beryllium Oxide made by burning 9/25th of a kilogram of Beryllium at 1.848 kg/litre with 16/25th kilogram of oxygen at 1.14 kg/litre, releases 28.43 MJ/litre whilst occupying 756.2 cc of volume. 1.322 kg/litre of propellant volume. ( 561.4 cc of O2, 194.8 cc of Be - creating a thick colloidal suspension - or sizing the solid rocket 'grain')

Magnesium Fluoride releases 1123.4 kJ/mole - that's MgF2 so 23.405 amu/molecule for magnesium and 19.00 amu/molecule for Flourine - a total of 18.29 MJ/kg - LESS than the Beryllium Oxide reaction - so, given the difficulty of Fluorine handling - the BeO is preferred.

However, Magnesium Oxide releases 601.7 kJ/mole - this produces MgO. So, the same 23.405 for Mg and 16.00 amu/molecule for oxygen - combines to total 39.405 grams per mole. So, this is 15.27 MJ/kg. Only slightly less than hydrogen oxygen, but lookie here - the propellant density is waay higher! Mg is 1.738 kg/litre and lox is 1.14 kg/litre. Mg is 23.405/39.405 of the total and LOX is 16/39.405 of the total. 341.7 cc of Mg is consumed along with 356.2 cc of LOX for each kg propellant. A total of 697.9 cc of propellant per kg yielding a density average of 1.433 kg/litre!

Magnesium is a lot easier to handle than hydrogen has a lot higher density, and produces nearly the same results - in MEMS rockets - superior results - to hydrogen! Magnesium isn't as powerful as beryllium, but its a lot cheaper and easier to handle.

830,000 tonnes per year of the metal is produced, mostly from China (though the USA could produce far more than it currently does). Vast quantities of magnesium can also be produced from sea water, providing the cost of energy drops (something that will occur with the creation of low-cost power satellites efficiently beaming energy to Earth)

http://photos1.blogger.com/blogger/2...r_platform.jpg

At $2,900 per tonne, Magnesium and $210 per tonne for Liquid Oxygen, a four stage rocket that went to the moon, and returned all the pieces, with a 7% structure fraction, due to MEMS technology, we obtain with a 4.7 km/sec exhaust speed the following;

u = 1 - 1 / exp( 4.0 / 4.7 ) = 0.573039523
p =(1 - u - .07) = 0.356960477

prop mg lox structure
1.00000000
2.801430591 1.6053304496 0.9535023264 0.6518281232 0.1961001414
7.8480133561 4.4972218302 2.6711705858 1.8260512443 0.5493609349
21.985664694 12.5986548094 7.4830989929 5.1155558165 1.5389965286
61.5913136368 35.2942569882 20.9633824339 14.3308745543 4.3113919546

TOTALS 53.9954640773 32.071154339 21.9243097383 6.5958495594

COSTS $2,900 $210 $2,000,000

COSTS/LAUNCH $93,006.35 $4,604.11 $13,191,699.12

500 reuses $26,383.40

TOTAL COST/TONNE (500x) $123,993.85

So, we can send people to the moon and return them safely to the Earth at $124 per kilogram using this technology. Not too shabby. At 145 kg per passenger - using a long-duration biosuit - and MEMS based life support, etc., We send 20 people to the moon with a 3 tonne payload. 16 paying passengers at $25,000 each - covers the cost of the trip. Double that to pay for biosuit and so forth. $50,000 - at a cost of say $5 million per passenger, the development of a fleet of ships, one for every 16 clients - is possible. 96 tonnes of magnesium is consumed on each flight.

In article ,
William Mook wrote:

On Wednesday, April 16, 2014 3:17:54 PM UTC+12, Orval Fairbairn wrote:
In article ,

William Mook wrote:



Beryllium metal be treated as a fuel grain in a hybrid liquid rocket,
similar

to this;



http://upload.wikimedia.org/wikipedi...ybrids_big.png



Liquid oxygen and beryllium metal form very high exhaust speed rockets
with

moderately high densities.



Liquid oxygen and aluminium metal form high exhaust velocities with high

propellant densities.



LOX/Be/H2 tripropellant has an even higher Isp, but how do you meter teh

Be?

To approaches,

(1) Nanoparticles of Beryllium colloidally suspended in the hydrogen,
(2) Hybrid rocket that has LOX/H2 injected normally, and Be is the fuel
grain

http://upload.wikimedia.org/wikipedi...ybrids_big.png

Checking out the Lithium/Fluorine/Hydrogen combination;

The hydrogen was added to reduce temperature so the engine wouldn't burn out.


http://www.flightglobal.com/pdfarchi...0-%200069.html

Lithium-Fluoride by itself has -616 kJ/mole - Li has an atomic weight of 7
and Fluorine has an atomic weight of 19 - a total of 26 - so this is 23.69
MJ/kg.

All of this converted to kinetic energy implies an exhaust velocity of 6.88
km/sec. How much gets converted depends on expansion ratio and that depends
on chamber pressure versus exist pressure. 85% efficiency at sea level, 92%
efficiency in vacuum implies 6.35 km/sec at sea level and 6.60 km/sec in
vacuo.

Lithium is 7/26th of the total propellant weight and Fluorine is 19/26th of
the total propellant weight in a stoichiometric mixture. This means that for
each kilogram of mixture 269.2 grams is Lithium with a density of 0.53
kg/litre whilst 730.8 grams of Fluorine make up the balance at 1.50 kg/litre.
This means that 508 cc of lithium along with 487.2 cc of fluorine are
combined for each kilogram of propellant. A total of 995.2 cc yielding an
averaged density of 1.005 kg/litre. Slightly higher than water.

The addition of hydrogen radically reduces this density, and the energy per
unit mass flow - which reduces the exhaust velocity accordingly.

The addition of 4 H2O molecules to the exhaust means that 4x 285.8 kJ/mole is
released and 4 moles mass 4x 18 grams = 72 grams so, we have a specific
energy of 15.87 MJ/kg of propellant. This is 5.60 km/sec ideally, 4.78
km/sec at sea level, and 5.18 km/sec in vacuum.

Here we have 2/18th of the total as hydrogen and 16/18th of the total as
oxygen. The density of hydrogen is 0.07 kg/litre whilst the density of
oxygen in liquid form is 1.14 kg/litre. So, the volume of hydrogen to make
2/18th kg is 1.59 litres whilst the volume of oxygen to make 16/18th kg is
0.78 litres. A total of 2.37 litres per kg a density of 0.42 kg/litre.
Actual engines run at 5.5 to 1 instead of stoichiometric ratios because of
temperature issues again at pressures that are reasonable with expansion
ratios needed to be efficient at sea level.

Running at a 5.5 to 1 Oxygen Fuel ratio means that 312.5 grams of excess
hydrogen is carried along and exhausted while 97.49 MJ is released for every
6.5 kg of propellant - a specific energy of 15 MJ/kg. These translate to an
idealized exhaust velocity of 4.65 km/sec and 5.04 km/sec

Now, these exhaust velocities are higher than reported. That's because
efficiencies are less than 85% and 92% respectively for macro-engines. This
is due to unfavourable scaling of large engines wrt heat transfer.

However, for micro-engines, these efficiencies are ROUTINELY attained due to
their rapid cooling and small size. Since microengines are made on wafers
that cost $15 per square inch, and thrust is 50 lbs per square inch, whilst
massing 0.8 ounces per square inch, cost is $0.30 per pound of lift and
thrust to weight is 1,000 to 1!! This means that we can over-build engine
capacity which gives us far higher reliability on these sort of thrust
surfaces.

http://uclarocketproject.com/?p=1

A kilogram of Beryllium Oxide made by burning 9/25th of a kilogram of
Beryllium at 1.848 kg/litre with 16/25th kilogram of oxygen at 1.14 kg/litre,
releases 28.43 MJ/litre whilst occupying 756.2 cc of volume. 1.322 kg/litre
of propellant volume. ( 561.4 cc of O2, 194.8 cc of Be - creating a thick
colloidal suspension - or sizing the solid rocket 'grain')

Magnesium Fluoride releases 1123.4 kJ/mole - that's MgF2 so 23.405
amu/molecule for magnesium and 19.00 amu/molecule for Flourine - a total of
18.29 MJ/kg - LESS than the Beryllium Oxide reaction - so, given the
difficulty of Fluorine handling - the BeO is preferred.

However, Magnesium Oxide releases 601.7 kJ/mole - this produces MgO. So, the
same 23.405 for Mg and 16.00 amu/molecule for oxygen - combines to total
39.405 grams per mole. So, this is 15.27 MJ/kg. Only slightly less than
hydrogen oxygen, but lookie here - the propellant density is waay higher! Mg
is 1.738 kg/litre and lox is 1.14 kg/litre. Mg is 23.405/39.405 of the total
and LOX is 16/39.405 of the total. 341.7 cc of Mg is consumed along with
356.2 cc of LOX for each kg propellant. A total of 697.9 cc of propellant
per kg yielding a density average of 1.433 kg/litre!

Magnesium is a lot easier to handle than hydrogen has a lot higher density,
and produces nearly the same results - in MEMS rockets - superior results -
to hydrogen! Magnesium isn't as powerful as beryllium, but its a lot cheaper
and easier to handle.

830,000 tonnes per year of the metal is produced, mostly from China (though
the USA could produce far more than it currently does). Vast quantities of
magnesium can also be produced from sea water, providing the cost of energy
drops (something that will occur with the creation of low-cost power
satellites efficiently beaming energy to Earth)

http://photos1.blogger.com/blogger/2...r_platform.jpg

At $2,900 per tonne, Magnesium and $210 per tonne for Liquid Oxygen, a four
stage rocket that went to the moon, and returned all the pieces, with a 7%
structure fraction, due to MEMS technology, we obtain with a 4.7 km/sec
exhaust speed the following;

u = 1 - 1 / exp( 4.0 / 4.7 ) = 0.573039523
p =(1 - u - .07) = 0.356960477

prop mg lox structure
1.00000000
2.801430591 1.6053304496 0.9535023264 0.6518281232 0.1961001414
7.8480133561 4.4972218302 2.6711705858 1.8260512443 0.5493609349
21.985664694 12.5986548094 7.4830989929 5.1155558165 1.5389965286
61.5913136368 35.2942569882 20.9633824339 14.3308745543 4.3113919546

TOTALS 53.9954640773 32.071154339 21.9243097383 6.5958495594

COSTS $2,900 $210 $2,000,000

COSTS/LAUNCH $93,006.35 $4,604.11 $13,191,699.12

500 reuses $26,383.40

TOTAL COST/TONNE (500x) $123,993.85

So, we can send people to the moon and return them safely to the Earth at
$124 per kilogram using this technology. Not too shabby. At 145 kg per
passenger - using a long-duration biosuit - and MEMS based life support,
etc., We send 20 people to the moon with a 3 tonne payload. 16 paying
passengers at $25,000 each - covers the cost of the trip. Double that to pay
for biosuit and so forth. $50,000 - at a cost of say $5 million per
passenger, the development of a fleet of ships, one for every 16 clients - is
possible. 96 tonnes of magnesium is consumed on each flight.

As the man said: "One slight problem!" The y both pollute like a drunken
dragon at a Mexican buffet.

Both fluorine and beryllium are either highly toxic or create highly
toxic combustion products or all of the above, so we shouldn't use them
anywhere near the Earth or anywhere that man willland.

In article orfairbairn-B4C163.00065917042014@70-3-168-
216.pools.spcsdns.net, says...

As the man said: "One slight problem!" The y both pollute like a drunken
dragon at a Mexican buffet.

Both fluorine and beryllium are either highly toxic or create highly
toxic combustion products or all of the above, so we shouldn't use them
anywhere near the Earth or anywhere that man willland.


Mook seems to be oblivious to these sorts of things. The cost of the
environmental impact doesn't enter into his equations. Unfortunately
for these propellants, the EPA would have something to say.

I'm not sure what he has against LOX/kerosene. Both SpaceX and Russian
LOX/kerosene engines seem to work just fine, and are reasonably
affordable when compared to other existing alternatives.

Jeff
--
"the perennial claim that hypersonic airbreathing propulsion would
magically make space launch cheaper is nonsense -- LOX is much cheaper
than advanced airbreathing engines, and so are the tanks to put it in
and the extra thrust to carry it." - Henry Spencer

"Fred J. McCall" wrote in message
...

Jeff Findley wrote:

In article orfairbairn-B4C163.00065917042014@70-3-168-
216.pools.spcsdns.net, says...

As the man said: "One slight problem!" The y both pollute like a drunken
dragon at a Mexican buffet.

Both fluorine and beryllium are either highly toxic or create highly
toxic combustion products or all of the above, so we shouldn't use them
anywhere near the Earth or anywhere that man willland.


Mook seems to be oblivious to these sorts of things. The cost of the
environmental impact doesn't enter into his equations. Unfortunately
for these propellants, the EPA would have something to say.


Once you get beyond basic arithmetic Mook seems pretty oblivious.


I'm not sure what he has against LOX/kerosene. Both SpaceX and Russian
LOX/kerosene engines seem to work just fine, and are reasonably
affordable when compared to other existing alternatives.


He can use a bunch of fancy arithmetic with theoretical fuels without
having reality smack him in the face.


He reminds me of a guy I knew in college that proved to me I could not see
Long Island from the CT side of the sound. Despite that I had, numerous
times. For this guy, the math trumped reality.

Much like Mook.


--
Greg D. Moore http://greenmountainsoftware.wordpress.com/
CEO QuiCR: Quick, Crowdsourced Responses. http://www.quicr.net

On Wednesday, April 16, 2014 3:17:54 PM UTC+12, Orval Fairbairn wrote:
In article ,

William Mook wrote:



Beryllium metal be treated as a fuel grain in a hybrid liquid rocket, similar

to this;



http://upload.wikimedia.org/wikipedi...ybrids_big.png



Liquid oxygen and beryllium metal form very high exhaust speed rockets with

moderately high densities.



Liquid oxygen and aluminium metal form high exhaust velocities with high

propellant densities.



LOX/Be/H2 tripropellant has an even higher Isp, but how do you meter teh

Be?

There are two ways;

(1) colloidal suspension of metal nanoparticles in liquid hydrogen, or LOX,

This has been done quite successfully with hydrogen peroxide and graphite. With cryogens the system can be made hypergolic, so all you have to do is heat the mixture to its ignition point.

http://www.rocketlab.co.nz/propulsio...onopropellant/

(2) hybrid liquid-solid rocket, with the liquids pumped past a solid grain,

With LOX/LH2 combination, you put a conventional LOX/LH2 rocket engine combustion chamber, without a delaval nozzle or expansion chamber, ahead of the grain and run it oxygen rich, past your solid grain.

http://www.freepatentsonline.com/6679049-0-large.jpg

By shaping the grain pattern, you control the thrust profile during ascent, but lack any ability to modify it (other than putting it out by starving it of oxygen, and restarting it.

http://digitalvideo.8m.net/rocketry/...lantgrains.gif

A variant is to have a metal wire feed into the engine at a rate consistent with the liquid flow rate. This gives you the ability to vary thrust as in a liquid rocket engine,and as in the first case.

This works well for a main engine, or engines, that are gimballed or directed in the usual manner. When you have a lot of tiny engines in an array, as in a MEMS engine array, having wire feeds dedicated to each engine doesn't let you use all propellant efficiently especially if you have a generic propulsive skin approach which I've described elsewhere. However, there are ways to work around this.

For example, you can have wire feeds, but each wire is a short segment - a rod - similar to the lead in a mechanical pencil. These are stacked together in a holding chamber and fed into any of the engines in a group of four engines. One normal to the surface, three at angles to the surface, and at right angles to each other pointed along the edges of a corner of a cube. That is 35.26 degrees away from the engine normal to the surface, with the thrust line making an angle 54.74 degrees with the surface - each of the 3 pointing outward radially from the normal facing engine located 120 degrees from each other when the angle is measured around the center of the normal engine along the plane of the surface at that point.

On Friday, April 18, 2014 10:35:27 AM UTC+12, Greg (Strider) Moore wrote:
"Fred J. McCall" wrote in message

...



Jeff Findley wrote:



In article orfairbairn-B4C163.00065917042014@70-3-168-

216.pools.spcsdns.net, says...



As the man said: "One slight problem!" The y both pollute like a drunken

dragon at a Mexican buffet.



Both fluorine and beryllium are either highly toxic or create highly

toxic combustion products or all of the above, so we shouldn't use them

anywhere near the Earth or anywhere that man willland.





Mook seems to be oblivious to these sorts of things. The cost of the

environmental impact doesn't enter into his equations. Unfortunately

for these propellants, the EPA would have something to say.





Once you get beyond basic arithmetic Mook seems pretty oblivious.





I'm not sure what he has against LOX/kerosene. Both SpaceX and Russian

LOX/kerosene engines seem to work just fine, and are reasonably

affordable when compared to other existing alternatives.





He can use a bunch of fancy arithmetic with theoretical fuels without

having reality smack him in the face.





He reminds me of a guy I knew in college that proved to me I could not see

Long Island from the CT side of the sound. Despite that I had, numerous

times. For this guy, the math trumped reality.



Much like Mook.





--

Greg D. Moore http://greenmountainsoftware.wordpress.com/

CEO QuiCR: Quick, Crowdsourced Responses. http://www.quicr.net

Magnesium is a lot easier to handle than hydrogen has a lot higher density, and produces nearly the same results - in MEMS rockets - superior results - to hydrogen! Magnesium isn't as powerful as beryllium, but its a lot cheaper and easier to handle.

Anyone who is actually paying attention to what was said would see I developed a detailed design for a four stage moonship based on Magnesium and liquid oxygen.

Magnesium and liquid oxygen is a better choice than either beryllium or fluorine.

So, let's go over our magnesium/lox system - instead of 4 stages, we have 7 similar elements and a single lunar lander.

The lunar lander lands on the moon and returns to Earth. This means it must go through a delta vee of about 4.7 km/sec. We do this so we can retrieve all elements for reuse. If we designed a fourth stage with a 4.0 km/sec delta vee, we would have to use the third stage to slow down at the moon, separate from that stage during boost, and continue landing on the moon with the fourth stage, keeping enough propellant to return to Earth. This wastes the third stage which falls to the moon. So, by reducing the fourth stage payload from 3,000 kg to 2,500 kg, and adding 500 kg to the propellant mass, we save the third stage, which does a lunar free return flight, and returns to Earth for recovery and reuse. The lunar lander lands, and returns as well.

Of course since we're not doing as much work with the third stage, it can be smaller, and all the way down the line, the launcher is smaller.

Summing the propellant in the first three stages, and dividing by 7 we obtain a common flight unit;

Total weight: 25,195.6 kg

Propellant: 22,452.9 kg 15,670.4 litres

Mg: 13,336.1 kg 7,673.3 litres
LOX: 9,116.8 kg 7,997.1 litres

Structu 2,742.7 kg

Seven of these in a hexagonal close pack arrangement (HCP). When viewed from above, the 7 element system looks like this;

(1) (2)
(3) (4) (5)
(6) (7)

With colloidal suspension of magnesium particles in the LOX, the system is an advanced high density monopropellant, similar to Rocketlab's graphite and hydrogen peroxide monopropellant. In this case the stages may be equipped with cross feeding. This permits the assembly to launch with all engines running, and more efficient use of engines.

When equipped with hybrid solid liquid engines cross feeding of magnesium is problematical. In that case engines 1,2,6 & 7 are running at take off, and then engines 3 and 5 are lit at stage separation, and finally stage 4 is lit.

With cross-feeding of an advanced high-density monopropellant the system drains propellant from 1,2,6 & 7 to feed all engines, then drains propellant from 3 and 5 to feed all engines, and then operates from propellant in element 4.

With 1.5 gees at lift off each element must generate 37.8 metric tons of thrust when cross-feeding is possible. Without cross-feeding the first four elements must lift the entire vehicle and so generate 66.15 metric tons. The second two elements must generate only 1 gee on the assembly and so must generate 75.6 metric tons of thrust and so each elements must generate 37.8 metric tons of thrust. The last element also generates an acceleration of 1 gee, and so requires only 25.6 metric tons of thrust.

With cross feed: 37.8 x 7 = 264.6 metric tons thrust

Without cross feed: 66.15 x 4 = 264.6 metric tons thrust
37.80 x 2 = 75.6 metric tons thrust
25.60 x 1 = 25.6 metric tons thrust

total: 365.8 metric tons thrust

This is a 38.25% increase in overall thrust, which means a similar increase in rocket engine structure weight and cost. Of course with MEMS based rockets that cost less than $1,000 per metric ton of lift, and have a 1,000 to 1 thrust to weight ratio, this isn't as big a deal as it was back in the 20th century.

Here's the revised commercial moonship - using seven similar elements;

Payload.... 2.503

Lunar Lander 8.404

Element 25.196

Propellant 22.453

Stages..... Stage 1 Stage 2 Stage 3

Total...... 184.774 83.991 33.600 metric tons
Propellant. 89.812 44.906 22.453 metric tons
u.......... 0.486 0.535 0.668 fraction
Ve......... 4.700 4.700 4.700 km/sec
Vf......... 3.130 3.600 5.190 km/sec (ideal)

Total Vf... 3.130 6.720 11.910 km/sec (ideal)


This seven element system is interesting since it is used also as a commercial launcher. For example, a single element with this low structural fraction is capable of SSTO operation with a payload of 950 kg

A three element system places 7,250 kg into the same orbits as the space shuttle.

A seven element system used to launch to LEO places 19,250 kg into these same orbits.

At 145 kg per passenger, a 2,503 kg payload is sufficient to blast 17 people to the moon. Subtracting out a crew of two we have 15 paying passengers.

Each element is shaped similarly to the old Space Shuttle's External Tank.

http://upload.wikimedia.org/wikipedi...ternaltank.jpg

Excepting the volume if far smaller. Instead of 2,433,180.1 litres (including intertank) this element carries only 15,670.4 litres of propellant. So, this element is 8,725.9 millimeters long and 1,562.9 millimeters in diameter.

Rounding up, 9 meters long and 1.6 meters in diameter.

This same shape also houses the lunar lander stage perched atop element (4) above. The landing gear doubles as an inter-stage support during launch.

Total propellant volume for the lunar lander is 3,361.2 litres with 1,645.9 litres dedicated to the 2,860.5 kg of magnesium and 1,715.3 litres dedicated to the 1,955.5 kg of lox.

The total volume of a bullet tank inside the tail of this element, with spherical end caps, that's 1,562.9 mm in diameter, with a cylindrical section 710.1 mm long, has a total volume of 3,361.2 litres needed.

So, an advanced high-density monopropellant fills a tank this size for the lunar lander. This is at the base of the lunar lander, and the MEMS array coats the rear facing hemisphere.

A hybrid liquid-solid rocket requires a spherical oxygen tank 1,485.2 mm in diameter (inner diameter) and magnesium grain section 1,715.8 mm tall, since the grain fills half the available area - and given the geometry of a well designed grain, consists of 7 engines packed together inside the airframe of conventional macro design. Using more advanced MEMS based systems, produces a surface which has a thickness of 1,715.8 with thousands of engines in an array.

Seats are two across, forward facing, with an access hatch for each. Each traveller is equipped with a long-duration mechanical counter pressure biosuit. Each seat has its own supply of air, water, power and food. 9 rows of two each. The 9th row has only 1 seat and is equipped with spares and common equipment.

Each suit has liquid oxygen and liquid hydrogen supply with its own cryocooler.

To power the suit a 418.7 Watt fuel cell array is used that also produces 73.9 Watts of thermal energy. This requires the use of 300.2 grams of hydrogen along with 2,401.6 grams of oxygen each day. In addition another 840 grams of oxygen each day is breathed by each astronaut. The astronaut also exhales 1,000 grams of carbon-dioxide. Another 181.8 grams of hydrogen is used to react this CO2 into 363.6 grams of methane and another 818.2 grams of water. The 818.2 grams of water is combined with the 2,701.8 grams of water produced by the fuel cell to provide 3,870 grams of potable water per day. That eventually ends up as 3,870 grams of liquid waste along with 110 grams of solid waste each day. This waste along with the methane is evaporated into the vacuum of space, which cools the astronaut. Another 620 grams per day of food is carried along. So the totals are;

840.0 O2 - breathing
2,401.6 O2 - electricity/heat
3,241.6 O2 - TOTAL

181.8 H2 - CO2 sorbent (breathing)
300.2 H2 - electricity/heat
482.0 H2 - TOTAL

620.0 FOOD - TOTAL

4,343.6 CONSUMABLE TOTAL/DAY (grams)

47.8 kg - CONSUMABLE - 11 days.

PRODUCED ON BOARD

3,870.0 H2O - potable water produced per day (grams)
418.7 Watts - electricity (continuous average)
73.9 Watts - heat from fuel cell (hot water)
110.0 solid waste
3,870.0 liquid waste
363.6 methane gas

The methane water and solid waste evaporation carries away up to 500 Watts per person. This includes the 61.3 watts from the astronaut themselves plus the 73.9 watts from the fuel cell and electronics.

"William Mook" wrote in message
...


You know, if you spent 1/2 as much time building any of this stuff as you do
writing about it, you'd be on the Moon by now.


--
Greg D. Moore http://greenmountainsoftware.wordpress.com/
CEO QuiCR: Quick, Crowdsourced Responses. http://www.quicr.net

In article ,
says...

"William Mook" wrote in message
...


You know, if you spent 1/2 as much time building any of this stuff as you do
writing about it, you'd be on the Moon by now.

No, he wouldn't because none of what he talks about is economically
viable (the exotic propellants). Heck, some of it isn't even physically
possible to build in the way he envisions (the MEMS based rocket engines
being built into stages with mass fractions which are laughable).

He'd do better writing this stuff into bad sci-fi books.

Jeff
--
"the perennial claim that hypersonic airbreathing propulsion would
magically make space launch cheaper is nonsense -- LOX is much cheaper
than advanced airbreathing engines, and so are the tanks to put it in
and the extra thrust to carry it." - Henry Spencer

On Saturday, April 19, 2014 2:09:25 AM UTC+12, Fred J. McCall wrote:
William Mook wrote:





Anyone who is actually paying attention to what was said would see





That you did a bunch of arithmetic and acted as if it actually meant

something.



--

"Some people get lost in thought because it's such unfamiliar

territory."

--G. Behn

Fred, its called analysis, you might want to look into it.

On Friday, April 18, 2014 11:21:42 PM UTC+12, Jeff Findley wrote:
In article ,

says...



"William Mook" wrote in message

...





You know, if you spent 1/2 as much time building any of this stuff as you do

writing about it, you'd be on the Moon by now.



No, he wouldn't because none of what he talks about is economically

viable (the exotic propellants). Heck, some of it isn't even physically

possible to build in the way he envisions (the MEMS based rocket engines

being built into stages with mass fractions which are laughable).



He'd do better writing this stuff into bad sci-fi books.



Jeff

--

"the perennial claim that hypersonic airbreathing propulsion would

magically make space launch cheaper is nonsense -- LOX is much cheaper

than advanced airbreathing engines, and so are the tanks to put it in

and the extra thrust to carry it." - Henry Spencer

Obviously those who haven't read any AIAA papers on MEMS rocketry in the past 15 years are clueless as to what's possible.

Thrust to weight - 1,000 to 1 - not 40 to 1. Why? Thrust varies by surface area, weight varies by volume. So an RL10 has a dry weight of 277 kg and produces 11,363 kg thrust has a thrust to weight of 41 and a diameter of 2..3 meters. Reduce this engine to 94.2 mm diameter - reduces its weight far faster than the area is reduced - and so, thrust to weight rises. In this case close to 1,000 to 1.

Cost per unit thrust - the thrust per unit area of exhaust is 2,735 kgf per square meter. Modern HDTV and other sophisticated MEMS structures of large areas cost $1,000 per sq m. So, this translates to $0.37 per kgf lift!

Low Structural Fractions - combined with the very low engine weight, advanced materials make very low structure weight possible - and at low cost.

http://www.physicscentral.org/explor...crolattice.cfm

Structure fractions using 1960s technology were below 7% for the S-IVB upper stage. In estimating 7% for a fully reusable stage, using modern materials, I am being conservative - as others who have published in AIAA and elsewhere, are projecting even lower structural fractions than this.

Another aspect I didn't relate fully is the use of magnesium with oxygen in the air, to make use of that resource during early stages of lift off. This is a minor, but important detail that provides measurable improvement in overall performance.