Successful SpaceX launch

On Monday, April 11, 2016 at 1:09:46 PM UTC+12, Sylvia Else wrote:
On 11/04/2016 9:15 AM, William Mook wrote:
On Sunday, April 10, 2016 at 9:50:23 PM UTC+12, Sylvia Else wrote:
On 10/04/2016 7:27 PM, William Mook wrote:
Alright, I've stated my prejudices against the Skylon's approach
to build a LACE propulsion system. SO, let's look at this more
closely and see if my prejudice is justified.

A Liquid Air Cycle Engine (LACE) is a type of spacecraft
propulsion engine that attempts to increase its efficiency by
gathering part of its oxidizer from the atmosphere. A liquid air
cycle engine uses liquid hydrogen (LH2) fuel to liquefy the air.

At its atmospheric boiling point, the specific
(constant-pressure) heat capacity of

liquid hydrogen is about 14.4 J/(g K). liquid oxygen is about 1
J/(g K). liquid nitrogen is about 1 J/(g K).

The heat of fusion of LH2: 58.5 J/(g K)

The heat of vaporization of LH2 is 452 J/g The heat of
vaporization of LOX is 216 J/g The heat of vaporization of LN2 is
199 J/g

Boiling point: LH2: 20.28 K LOX: 90.19 K LN2: 77.00 K

Now, at room temperature (293 K) a pure oxygen atmosphere will
require 203 Joules to reduce a gram of gas to the boiling point
of LOX. It will take another 216 J per g to liquefy the oxygen.
A total of 419 J per g of oxygen.

Now a solid block of hydrogen at 14 K will absorb 76x14.4 J =
1,094 J rising to 90.19 K, and will also absorb an added 58.5 J
along with another 452 J for each gram. A grand total of 1,605
J. So, each gram will have sufficient capacity to absorb energy
to liquefy 3.83 grams of oxygen at room temperature.

To obtain 5.5 grams of oxygen for each gram of hydrogen which is
the ideal O:F ratio, we must absorb 2,095 J if we are to liquefy
all the oxygen. This means we must raise the temperature to
room temperature. Of course doing this, allows the oxygen to
boil and then to rise in temperature to room temperature, which
reduces the energy required to that lost in the process.

So, how is this supposed to work then?

Liquid hydrogen runs through a heat exchanger which cools
incoming air. That air eventually liquefies the oxygen, but not
the nitrogen which boils at a lower temperature. The cold
nitrogen gas is used to chill the air, so that the hydrogen
doesn't have to. The purified oxygen, also chills the air, so
the hydrogen doesn't have to.

So, by boiling away a gram of solid hydrogen and raising it to
room temperature, 2,095 Joules of energy is absorbed in the
process. This liquefies 5.5 grams of oxygen starting at room
temperature, assuming the nitrogen is recovered.

With air, 78% is nitrogen and 21% oxygen. By liquefying the
oxygen in the air, the nitrogen remains gaseous. So, it can flow
through a heat exchanger and exit the craft, reducing the work
the liquid hydrogen has to do. The lox in a similar fashion can
absorb heat to reduce the load on the liquid hydrogen process.

km/sec R..... K...... J/gram grams grams

0.30 500 277.78 403.59 1.01 0.47 0.60 750 416.67 542.48
1.36 0.82 0.90 1000 555.56 681.37 1.70 1.16 1.20 1500 833.33
959.14 2.40 1.86 1.50 2000 1,111.11 1,236.92 3.09 2.55 1.80
3000 1,666.67 1,792.48 4.48 3.94 2.10 4000 2,222.22 2,348.03
5.87 5.33 2.40 5000 2,777.78 2,903.59 7.26 6.72

At 5.5 to 1 and 60 atmosphere pressure, the flame temperature is
3400 K.

I don't see how this can work in practice. At high speeds there
isn't enough time for the air to be chilled. If the air is
stopped and held until chilled, the drag and temperatures become
too high.

Evaporating liquid hydrogen in a heat exchanger could cool
ambient air at a certain rate to extract LOX seems doable. Yet,
looking at LOX production plants it takes about 800 kWh to
produce a ton of LOX from 5 tons of air. That's 2880 Joules per
gram. That's the amount of energy absorbed by bringing 1 gram of
solid hydrogen up to room temperature.

https://books.google.co.nz/books?id=...%20air&f=false





So, I don't see how it can work realistically.



. On Sunday, April 10, 2016 at 2:41:49 PM UTC+12, Sylvia Else
wrote:
On 10/04/2016 6:34 AM, Jeff Findley wrote:
In article ,
ess says...

On 9/04/2016 8:24 AM, Jeff Findley wrote:

I'm tired of people saying "chemical propulsion is too
expensive" when the fuel is so damn cheap! Fuel costs
are *not* the problem. Skylon/Sabre is a solution to a
problem which *does not exist*!

Skylon/Sabre is not about reducing the amount of fuel
consumed.

Sorry, I was being sloppy and lumping the mass of the
oxidizer in with the fuel.

How about this. LOX is one of the cheapest fluids used in
the aerospace industry. It's literally made from air.
Trying to reduce LOX consumption is quite counter-intuitive
if reducing launch costs is the goal.

Jeff


The point of Sabre is not to reduce the amount of LOX that is
consumed, the cost of LOX being, as you point out, negligible
in this context.

The goal is to build an SSTO vehicle, with the economic
advantages that brings. Sabre is a means to that end, because
it reduces the amount of LOX that has to be *lifted*, and
allows atmospheric nitrogen to be used as reaction mass during
the air-breathing phase.

Sylvia.

Sabre is not a LACE. The air is only cooled to the vapour phase
boundary, not liquified. The cycle would be less efficient if the
air were liquified (even in part).

The system involves cooling the air at the rate it arrives. If this
is not achieved, then the engine won't work, but you can't a priori
say that it cannot be achieved.

Sylvia/

Here's what I got from online sources;

The RB545 is covered by the official secrets act, so not much can be
said about it. This is a barrier right at the outset.

The SABRE design is neither a conventional rocket engine nor jet
engine, but a hybrid that uses air from the environment at low
speeds/altitudes, and stored liquid oxygen (LOX) at higher altitude.

The SABRE engine "relies on a heat exchanger capable of cooling
incoming air to -150 °C (-238 °F), to provide oxygen for mixing with
hydrogen and provide jet thrust during atmospheric flight before
switching to tanked liquid oxygen when in space.

At the front of the engine, a simple translating axisymmetric shock
cone inlet slows the air to subsonic speeds using two shock
reflections.

Part of the air then passes through a precooler into the central
core, with the remainder passing directly through a ring of bypass
ramjets.

The central core of SABRE behind the precooler uses a
turbo-compressor run off the same gaseous helium loop Brayton cycle
which compresses the air and feeds it into four high pressure
combined cycle rocket engine combustion chambers.

The oxygen is also fed to the combustion unit, using a turbopump.

***

So, I don't get how you end up with oxygen without liquefying it. I
do get that using liquid hydrogen's larger specific heat its possible
to cool air, and in a sense compress it unconventionally. This seems
to me critical. Since it appears nowhere in any of the literature
available, and it is a fundamental feature, that suggests to me one
of the essential features hidden by the official secrets act around
the RB545.

There's quite a lot of information available in the technical documents
section of the Reaction Engines web site.

http://www.reactionengines.co.uk/tech_docs.html

I don't see a detailed entropy temperature graph of the oxygen and hydrogen flow through the engine, but I'll keep looking.


Importantly, there is no separation of oxygen from the air during the
air breathing ascent. The cooled compressed air is fed into to the
rocket engines.

That's an improvement on the LACE concept certainly.

In 2011, hardware testing of the heat exchanger technology "crucial
to [the] hybrid air- and liquid oxygen-breathing [SABRE] rocket
motor" was completed, demonstrating that the technology is viable.
The tests validated that the heat exchanger could perform as needed
for the engine to obtain adequate oxygen from the atmosphere to
support the low-altitude, high-performance operation.

* * *

Not clear how this helps with high altitude high speed operation!
lol.

Regardless of the speed through the air, the inlet reduces the air speed
to subsonic.

This is a problem in that the stagnation temperature gets really high at higher speeds.

Beyond mach 5.5, the craft operates as a pure rocket.

Yes, that's 1.87 km/sec. A little slower than I calculated. That means you take more oxygen along.

The rest of your analysis appears to relate to something rather
different from the sabre engine and skylon.

Just being as generous to the concept as possible so that the advantages were maximised for air breathing. I still think something like the Delta Heavy or the Falcoln Heavy with three common core boosters, make a lot more sense at the present time.

Sylvia.

I will modify my commentary to this extent. It might be made to work, though I think LACE is totally unworkable as I understand it.

On 11/04/2016 11:49 AM, William Mook wrote:
On Monday, April 11, 2016 at 1:09:46 PM UTC+12, Sylvia Else wrote:
On 11/04/2016 9:15 AM, William Mook wrote:
On Sunday, April 10, 2016 at 9:50:23 PM UTC+12, Sylvia Else wrote:
On 10/04/2016 7:27 PM, William Mook wrote:
Alright, I've stated my prejudices against the Skylon's approach
to build a LACE propulsion system. SO, let's look at this more
closely and see if my prejudice is justified.

A Liquid Air Cycle Engine (LACE) is a type of spacecraft
propulsion engine that attempts to increase its efficiency by
gathering part of its oxidizer from the atmosphere. A liquid air
cycle engine uses liquid hydrogen (LH2) fuel to liquefy the air.

At its atmospheric boiling point, the specific
(constant-pressure) heat capacity of

liquid hydrogen is about 14.4 J/(g K). liquid oxygen is about 1
J/(g K). liquid nitrogen is about 1 J/(g K).

The heat of fusion of LH2: 58.5 J/(g K)

The heat of vaporization of LH2 is 452 J/g The heat of
vaporization of LOX is 216 J/g The heat of vaporization of LN2 is
199 J/g

Boiling point: LH2: 20.28 K LOX: 90.19 K LN2: 77.00 K

Now, at room temperature (293 K) a pure oxygen atmosphere will
require 203 Joules to reduce a gram of gas to the boiling point
of LOX. It will take another 216 J per g to liquefy the oxygen.
A total of 419 J per g of oxygen.

Now a solid block of hydrogen at 14 K will absorb 76x14.4 J =
1,094 J rising to 90.19 K, and will also absorb an added 58.5 J
along with another 452 J for each gram. A grand total of 1,605
J. So, each gram will have sufficient capacity to absorb energy
to liquefy 3.83 grams of oxygen at room temperature.

To obtain 5.5 grams of oxygen for each gram of hydrogen which is
the ideal O:F ratio, we must absorb 2,095 J if we are to liquefy
all the oxygen. This means we must raise the temperature to
room temperature. Of course doing this, allows the oxygen to
boil and then to rise in temperature to room temperature, which
reduces the energy required to that lost in the process.

So, how is this supposed to work then?

Liquid hydrogen runs through a heat exchanger which cools
incoming air. That air eventually liquefies the oxygen, but not
the nitrogen which boils at a lower temperature. The cold
nitrogen gas is used to chill the air, so that the hydrogen
doesn't have to. The purified oxygen, also chills the air, so
the hydrogen doesn't have to.

So, by boiling away a gram of solid hydrogen and raising it to
room temperature, 2,095 Joules of energy is absorbed in the
process. This liquefies 5.5 grams of oxygen starting at room
temperature, assuming the nitrogen is recovered.

With air, 78% is nitrogen and 21% oxygen. By liquefying the
oxygen in the air, the nitrogen remains gaseous. So, it can flow
through a heat exchanger and exit the craft, reducing the work
the liquid hydrogen has to do. The lox in a similar fashion can
absorb heat to reduce the load on the liquid hydrogen process.

km/sec R..... K...... J/gram grams grams

0.30 500 277.78 403.59 1.01 0.47 0.60 750 416.67 542.48
1.36 0.82 0.90 1000 555.56 681.37 1.70 1.16 1.20 1500 833.33
959.14 2.40 1.86 1.50 2000 1,111.11 1,236.92 3.09 2.55 1.80
3000 1,666.67 1,792.48 4.48 3.94 2.10 4000 2,222.22 2,348.03
5.87 5.33 2.40 5000 2,777.78 2,903.59 7.26 6.72

At 5.5 to 1 and 60 atmosphere pressure, the flame temperature is
3400 K.

I don't see how this can work in practice. At high speeds there
isn't enough time for the air to be chilled. If the air is
stopped and held until chilled, the drag and temperatures become
too high.

Evaporating liquid hydrogen in a heat exchanger could cool
ambient air at a certain rate to extract LOX seems doable. Yet,
looking at LOX production plants it takes about 800 kWh to
produce a ton of LOX from 5 tons of air. That's 2880 Joules per
gram. That's the amount of energy absorbed by bringing 1 gram of
solid hydrogen up to room temperature.

https://books.google.co.nz/books?id=...%20air&f=false





So, I don't see how it can work realistically.



. On Sunday, April 10, 2016 at 2:41:49 PM UTC+12, Sylvia Else
wrote:
On 10/04/2016 6:34 AM, Jeff Findley wrote:
In article ,
ess says...

On 9/04/2016 8:24 AM, Jeff Findley wrote:

I'm tired of people saying "chemical propulsion is too
expensive" when the fuel is so damn cheap! Fuel costs
are *not* the problem. Skylon/Sabre is a solution to a
problem which *does not exist*!

Skylon/Sabre is not about reducing the amount of fuel
consumed.

Sorry, I was being sloppy and lumping the mass of the
oxidizer in with the fuel.

How about this. LOX is one of the cheapest fluids used in
the aerospace industry. It's literally made from air.
Trying to reduce LOX consumption is quite counter-intuitive
if reducing launch costs is the goal.

Jeff


The point of Sabre is not to reduce the amount of LOX that is
consumed, the cost of LOX being, as you point out, negligible
in this context.

The goal is to build an SSTO vehicle, with the economic
advantages that brings. Sabre is a means to that end, because
it reduces the amount of LOX that has to be *lifted*, and
allows atmospheric nitrogen to be used as reaction mass during
the air-breathing phase.

Sylvia.

Sabre is not a LACE. The air is only cooled to the vapour phase
boundary, not liquified. The cycle would be less efficient if the
air were liquified (even in part).

The system involves cooling the air at the rate it arrives. If this
is not achieved, then the engine won't work, but you can't a priori
say that it cannot be achieved.

Sylvia/

Here's what I got from online sources;

The RB545 is covered by the official secrets act, so not much can be
said about it. This is a barrier right at the outset.

The SABRE design is neither a conventional rocket engine nor jet
engine, but a hybrid that uses air from the environment at low
speeds/altitudes, and stored liquid oxygen (LOX) at higher altitude.

The SABRE engine "relies on a heat exchanger capable of cooling
incoming air to -150 °C (-238 °F), to provide oxygen for mixing with
hydrogen and provide jet thrust during atmospheric flight before
switching to tanked liquid oxygen when in space.

At the front of the engine, a simple translating axisymmetric shock
cone inlet slows the air to subsonic speeds using two shock
reflections.

Part of the air then passes through a precooler into the central
core, with the remainder passing directly through a ring of bypass
ramjets.

The central core of SABRE behind the precooler uses a
turbo-compressor run off the same gaseous helium loop Brayton cycle
which compresses the air and feeds it into four high pressure
combined cycle rocket engine combustion chambers.

The oxygen is also fed to the combustion unit, using a turbopump.

***

So, I don't get how you end up with oxygen without liquefying it. I
do get that using liquid hydrogen's larger specific heat its possible
to cool air, and in a sense compress it unconventionally. This seems
to me critical. Since it appears nowhere in any of the literature
available, and it is a fundamental feature, that suggests to me one
of the essential features hidden by the official secrets act around
the RB545.

There's quite a lot of information available in the technical documents
section of the Reaction Engines web site.

http://www.reactionengines.co.uk/tech_docs.html

I don't see a detailed entropy temperature graph of the oxygen and hydrogen flow through the engine, but I'll keep looking.


Importantly, there is no separation of oxygen from the air during the
air breathing ascent. The cooled compressed air is fed into to the
rocket engines.

That's an improvement on the LACE concept certainly.

In 2011, hardware testing of the heat exchanger technology "crucial
to [the] hybrid air- and liquid oxygen-breathing [SABRE] rocket
motor" was completed, demonstrating that the technology is viable.
The tests validated that the heat exchanger could perform as needed
for the engine to obtain adequate oxygen from the atmosphere to
support the low-altitude, high-performance operation.

* * *

Not clear how this helps with high altitude high speed operation!
lol.

Regardless of the speed through the air, the inlet reduces the air speed
to subsonic.

This is a problem in that the stagnation temperature gets really high at higher speeds.

Indeed, it's at least part of the reason for the precoooler.


Beyond mach 5.5, the craft operates as a pure rocket.

Yes, that's 1.87 km/sec. A little slower than I calculated. That means you take more oxygen along.

The rest of your analysis appears to relate to something rather
different from the sabre engine and skylon.

Just being as generous to the concept as possible so that the advantages were maximised for air breathing. I still think something like the Delta Heavy or the Falcoln Heavy with three common core boosters, make a lot more sense at the present time.

Sylvia.

I will modify my commentary to this extent. It might be made to work, though I think LACE is totally unworkable as I understand it.


Reaction Engines seem to be of the same view, which is why Sabre is not
a LACE. Amongst other things, liquifying oxygen in a heat exchanger
involves an unavoidable increase in entropy, because at the phase
boundary, the air temperature cannot reduce until the oxygen is all
liquid, whereas the coolant on the other side of the exchanger does
change in temperature. So there's a significant temperature gradient
across the heat-exchanger walls, and thus an increase in entropy,
representing a permanent loss of the ability of the transferred energy
to do work.

Sylvia.

In article ,
ess says...

On 10/04/2016 6:34 AM, Jeff Findley wrote:
In article ,
ess says...

On 9/04/2016 8:24 AM, Jeff Findley wrote:

I'm tired of people saying "chemical propulsion is too expensive" when
the fuel is so damn cheap! Fuel costs are *not* the problem.
Skylon/Sabre is a solution to a problem which *does not exist*!

Skylon/Sabre is not about reducing the amount of fuel consumed.

Sorry, I was being sloppy and lumping the mass of the oxidizer in with
the fuel.

How about this. LOX is one of the cheapest fluids used in the aerospace
industry. It's literally made from air. Trying to reduce LOX
consumption is quite counter-intuitive if reducing launch costs is the
goal.

The point of Sabre is not to reduce the amount of LOX that is consumed,
the cost of LOX being, as you point out, negligible in this context.

The goal is to build an SSTO vehicle, with the economic advantages that
brings. Sabre is a means to that end, because it reduces the amount of
LOX that has to be *lifted*, and allows atmospheric nitrogen to be used
as reaction mass during the air-breathing phase.

By the time Sabre/Skylon flies it may very well have to compete with a
next generation fully reusable TSTO. A fully reusable, LOX/methane,
TSTO is what SpaceX plans on pursuing as their Mars launch vehicle.

I welcome the competition, but I just don't see the point of wings and
intakes on a vehicle that is trying to go to LEO.

Jeff
--
All opinions posted by me on Usenet News are mine, and mine alone.
These posts do not reflect the opinions of my family, friends,
employer, or any organization that I am a member of.

On 11/04/2016 8:18 PM, Jeff Findley wrote:
In article ,
ess says...

On 10/04/2016 6:34 AM, Jeff Findley wrote:
In article ,
ess says...

On 9/04/2016 8:24 AM, Jeff Findley wrote:

I'm tired of people saying "chemical propulsion is too expensive" when
the fuel is so damn cheap! Fuel costs are *not* the problem.
Skylon/Sabre is a solution to a problem which *does not exist*!

Skylon/Sabre is not about reducing the amount of fuel consumed.

Sorry, I was being sloppy and lumping the mass of the oxidizer in with
the fuel.

How about this. LOX is one of the cheapest fluids used in the aerospace
industry. It's literally made from air. Trying to reduce LOX
consumption is quite counter-intuitive if reducing launch costs is the
goal.

The point of Sabre is not to reduce the amount of LOX that is consumed,
the cost of LOX being, as you point out, negligible in this context.

The goal is to build an SSTO vehicle, with the economic advantages that
brings. Sabre is a means to that end, because it reduces the amount of
LOX that has to be *lifted*, and allows atmospheric nitrogen to be used
as reaction mass during the air-breathing phase.

By the time Sabre/Skylon flies it may very well have to compete with a
next generation fully reusable TSTO. A fully reusable, LOX/methane,
TSTO is what SpaceX plans on pursuing as their Mars launch vehicle.

I welcome the competition, but I just don't see the point of wings and
intakes on a vehicle that is trying to go to LEO.

Jeff


An important issue with these capital intensive machines is their level
of utilisation. While they're sitting on the ground, they're not earning
money, but they're still costing money, in the form of interest on the
capital invested in them.

Skylon lands in one piece back at its launch site. All that's required
is a standard aircraft tug to tow it back to the hangar.

With a TSTO, after a mission you have the extra time involved in putting
it back together. Further, if it's a vertical stack, loading it for the
next mission is not so straight forward. I also suspect it will be a
long time before SpaceX are allowed to land their rockets anywhere near
where the reintegration and loading work is done. By contrast, Skylon is
just a glider at that point in its mission, and is no more (arguably
less) of a threat than an airliner.

I expect SpaceX will succeed in making a reusable TSTO, if they decide
to. They're clearly very competent in their art.

But if Skylon works, I wouldn't want to bet that SpaceX will be able to
undercut it, or even compete with it.

Sylvia.

Jeff Findley wrote:
I welcome the competition, but I just don't see the point of wings
and intakes on a vehicle that is trying to go to LEO.

Because slapping a PanAm logo on a Falcon second stage just isn't the
same

rick jones
--
Wisdom Teeth are impacted, people are affected by the effects of events.
these opinions are mine, all mine; HPE might not want them anyway...
feel free to post, OR email to rick.jones2 in hpe.com but NOT BOTH...

On Tuesday, April 12, 2016 at 4:06:44 AM UTC+12, Rick Jones wrote:
Jeff Findley wrote:
I welcome the competition, but I just don't see the point of wings
and intakes on a vehicle that is trying to go to LEO.

Because slapping a PanAm logo on a Falcon second stage just isn't the
same

rick jones
--
Wisdom Teeth are impacted, people are affected by the effects of events.
these opinions are mine, all mine; HPE might not want them anyway...
feel free to post, OR email to rick.jones2 in hpe.com but NOT BOTH...

An air breathing first stage, lifting a rocket second stage is also a possibility.

A two stage LOX/LH2 setup with a 5.5 to 1.0 oxidizer/fuel ratio, and a 4.3 km/sec exhaust speed in the first stage and 4.5 km/sec in the second stage, with 4.6 km/sec delta vee in each stage and 4.65% structure fraction is;

Payload..: 1.0000
Structu 0.1484
Hydrogen: 0.3144
Oxygen...: 1.7290

Total.......: 3.1918

Payload..: 3.1918
Structu 0.5004
Hydrogen: 1.0876
Oxygen...: 5.9820

Total.......: 10.7620


By taking oxygen out of the air during the operation of the first stage, we have the potential to reduce the entire weight at take off by 55.5% !!

Now if we take the propellant in the first stage, divide it by the propellant in the second stage, and divide that by three we obtain 1.1532x - which means that we can build a four element launcher with cross-feeding all made of the same element. Equipping each element with an aerospike engine that performs well at low altitude as well as high, we have a pretty standardised component to lift stuff into LEO.

For each ton of payload in orbit we have four elements where each is;

Hydrogen: 0.3625
Oxygen...: 1.9940
Structu 0.1668

Total.......: 2.5233

Whilst for the airbreathing device we have a single upper stage of approximately this size,

Hydrogen: 0.3144
Oxygen...: 1.7291
Structu 0.1484

Total......: 2.1919

and a first stage with scramjet engines of some sort that;

Hydrogen: 1.0876
Oxygen...: 0.0000 (ideally)
Structu 0.5004

Total......: 1.5880


A fully reusable three stage vehicle with

Payload...: 1.0000
Structure.: 0.1013
Oxygen...: 0.9116
Hydrogen: 0.1658

Total.......: 2.1787

Payload...: 2.1787
Structure.:0.2285
Oxygen...: 2.1226
Hydrogen:0.3859

Total.......: 4.9157 (divide by 2 - for common core)

Payload...: 4.9157
Structure.: 0.5157
Oxygen...: 4.7892
Hydrogen: 0.8708 (divide by 4 - for common core)

Total........: 11.0914

So, a common core booster that's 1.1339x larger than the payload with the following;

Payload...: 1.0000

Seven ELements each;

Structure.: 0.0527
Hydrogen: 0.1664
Oxygen...: 0.9148


http://www.buran.ru/images/jpg/gk175-1.jpg
http://goo.gl/T4iOap
http://www.russianspaceweb.com/superheavy.html

The weight of an external tank is 730 tons. Using this as representative, we can see the following;

Payload: 670.0 tons

Tank Empty: 35.4 tons
Hydrogen..: 111.5 tons
Oxygen.....: 613.1 tons

A tremendous capacity! Total cost should be around $2.1 billion to build - assuming existing shuttle hardware. Rebuilding the hardware, adds another $5.4 billion to the total. A fleet of highly reusable vehicles, with added infrastructure to suppor higher launch rates, adds another $5.0 billion. Of course, then you have the payload supply chain. What are you launching that's that large that often? The answer is, power satellites! With 22 kW per kg you have 14.74 billion watts of power. At $0.18 per kWh a single satellite produces $23.26 billion per year in revenue;

14.74 billion watts -- 14.74 million kW -- x8766 hours/yr x $0.18 per kWh = $23,257 million/year

Now, a revenue of this magnitude, discounted at 8.5% per year over 20 years, is worth $220.12 billion the day it switches on! Discounting this value at venture capital rates of return, over a five year development cycle, obtains $38.94 billion valuation for the first satellite, before construction starts. So, arranging the sale of power ahead of time, and selling it on the market, provides a means to raise the money needed.

Year 1: $595.24 million --- $3,364.63 million 1.53% $1 -- $5.66
Year 2: $1190.48 million --- $4759.03 million 2.16% $1 -- $4.00
Year 3: $2380.95 million --- $6731.30 milllion 3.06% $1 -- $2.83
Year 4: $4761.90 million --- $9520.93 million 4.33% $1 -- $2.00
Year 5: $3571.42 million --- $5050.00 million 2.29% $1 -- $1.41

Totals: $12,500 million ---- $29,425.90 million 13.37% $5 -- $15.90


On the other end; we have smaller rockets adapted to smaller vehicles;

http://www.yang.gatech.edu/publicati...20Casiano).pdf

A single RL-10 pumpset, connected to a toroidal aerospike engine, a 1/10th scale engine similar to the J2-toroidal aerospike engine, is capable of 4.3 km/sec exhaust speed at low altitudes rising to 4.6 km/sec exhaust speeds at high altitudes. Masses 130 kg and produces 11,360 kgf thrust. This is capable of putting 6,953 kg into LEO (15,297 lbs) and each flight element is 7,884.5 kg mass.

Each element

Structure..: 366.4 kg
Hydrogen.: 1,157.1 kg
Oxygen....: 6,361.0 kg

Each element is 10.23 m (33.54 ft) long by 1.83 m (6.02 ft) in diameter.

The vehicle should be around $65 million to build, including infrastructure..

7 tons puts up a very powerful communications satellite. With 12 launches, carried out over 365 days, a global wireless telecommunications network can be placed on orbit.

Thin film concentrating photovoltaics - 10 megawatt - 450 kg
Inflatable phased array microwave - 100 meters diam - 450 kg
two open optical data links - inflatable optics - 3 meters diameter - 900 kg

Sub total: 1800 kg

5.20 tons - remaining for router and other...

Uplink downlink 444.000 GHz 524.000 GHz four 20 GHz channels worldwide - colouring cells from orbit - doppler corrected. From GEO, they paint a large number 10 km diameter cells, forming 981,524 cells per satellite, 11,778,291 cells world wide. Approximately 10 Watts per cell.

http://ieeexplore.ieee.org/xpl/login...er% 3D4131291

http://www.eecs.berkeley.edu/~bora/J...014/JSSC14.pdf

There are 4.2 billion wireless handsets world wide capable of broadband. Accessing those markets access $1,330 billion per year.

An iPhone type wallet case

http://www.amazon.com/Best-Sellers-C...ess/9375357011

that links via bluetooth or micro usb, and has extended battery capacity, that recharges the telephone, sporting a flexible phased array antenna in the case, with satellite decoder, and GUI that operates as an app on the telephone - to provide Skype like unlimited broadband satellite service for a fixed fee of say $20 per month. Sign up for two years and get the telephone wallet free.

http://www.newsweek.com/elon-musk-wa...ty-mars-300523

If the question is whether Skylon will liquify the air, numerous sources say
it will not; it will just cool the air:

http://www.gravityloss.com/2010/10/w...lon-and-sabre/

http://www.islandone.org/Propulsion/LACE.html

Bob Clark


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Finally, nanotechnology can now fulfill its potential to revolutionize
21st-century technology, from the space elevator, to private, orbital
launchers, to 'flying cars'.
This crowdfunding campaign is to prove it:

Nanotech: from air to space.
https://www.indiegogo.com/projects/n...ce/x/13319568/
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"William Mook" wrote in message
...

On Monday, April 11, 2016 at 1:09:46 PM UTC+12, Sylvia Else wrote:
On 11/04/2016 9:15 AM, William Mook wrote:
On Sunday, April 10, 2016 at 9:50:23 PM UTC+12, Sylvia Else wrote:
On 10/04/2016 7:27 PM, William Mook wrote:
Alright, I've stated my prejudices against the Skylon's approach
to build a LACE propulsion system. SO, let's look at this more
closely and see if my prejudice is justified.

A Liquid Air Cycle Engine (LACE) is a type of spacecraft
propulsion engine that attempts to increase its efficiency by
gathering part of its oxidizer from the atmosphere. A liquid air
cycle engine uses liquid hydrogen (LH2) fuel to liquefy the air.

At its atmospheric boiling point, the specific
(constant-pressure) heat capacity of

liquid hydrogen is about 14.4 J/(g K). liquid oxygen is about 1
J/(g K). liquid nitrogen is about 1 J/(g K).

The heat of fusion of LH2: 58.5 J/(g K)

The heat of vaporization of LH2 is 452 J/g The heat of
vaporization of LOX is 216 J/g The heat of vaporization of LN2 is
199 J/g

Boiling point: LH2: 20.28 K LOX: 90.19 K LN2: 77.00 K

Now, at room temperature (293 K) a pure oxygen atmosphere will
require 203 Joules to reduce a gram of gas to the boiling point
of LOX. It will take another 216 J per g to liquefy the oxygen.
A total of 419 J per g of oxygen.

Now a solid block of hydrogen at 14 K will absorb 76x14.4 J =
1,094 J rising to 90.19 K, and will also absorb an added 58.5 J
along with another 452 J for each gram. A grand total of 1,605
J. So, each gram will have sufficient capacity to absorb energy
to liquefy 3.83 grams of oxygen at room temperature.

To obtain 5.5 grams of oxygen for each gram of hydrogen which is
the ideal O:F ratio, we must absorb 2,095 J if we are to liquefy
all the oxygen. This means we must raise the temperature to
room temperature. Of course doing this, allows the oxygen to
boil and then to rise in temperature to room temperature, which
reduces the energy required to that lost in the process.

So, how is this supposed to work then?

Liquid hydrogen runs through a heat exchanger which cools
incoming air. That air eventually liquefies the oxygen, but not
the nitrogen which boils at a lower temperature. The cold
nitrogen gas is used to chill the air, so that the hydrogen
doesn't have to. The purified oxygen, also chills the air, so
the hydrogen doesn't have to.

So, by boiling away a gram of solid hydrogen and raising it to
room temperature, 2,095 Joules of energy is absorbed in the
process. This liquefies 5.5 grams of oxygen starting at room
temperature, assuming the nitrogen is recovered.

With air, 78% is nitrogen and 21% oxygen. By liquefying the
oxygen in the air, the nitrogen remains gaseous. So, it can flow
through a heat exchanger and exit the craft, reducing the work
the liquid hydrogen has to do. The lox in a similar fashion can
absorb heat to reduce the load on the liquid hydrogen process.

km/sec R..... K...... J/gram grams grams

0.30 500 277.78 403.59 1.01 0.47 0.60 750 416.67 542.48
1.36 0.82 0.90 1000 555.56 681.37 1.70 1.16 1.20 1500 833.33
959.14 2.40 1.86 1.50 2000 1,111.11 1,236.92 3.09 2.55 1.80
3000 1,666.67 1,792.48 4.48 3.94 2.10 4000 2,222.22 2,348.03
5.87 5.33 2.40 5000 2,777.78 2,903.59 7.26 6.72

At 5.5 to 1 and 60 atmosphere pressure, the flame temperature is
3400 K.

I don't see how this can work in practice. At high speeds there
isn't enough time for the air to be chilled. If the air is
stopped and held until chilled, the drag and temperatures become
too high.

Evaporating liquid hydrogen in a heat exchanger could cool
ambient air at a certain rate to extract LOX seems doable. Yet,
looking at LOX production plants it takes about 800 kWh to
produce a ton of LOX from 5 tons of air. That's 2880 Joules per
gram. That's the amount of energy absorbed by bringing 1 gram of
solid hydrogen up to room temperature.

https://books.google.co.nz/books?id=...%20air&f=false





So, I don't see how it can work realistically.



. On Sunday, April 10, 2016 at 2:41:49 PM UTC+12, Sylvia Else
wrote:
On 10/04/2016 6:34 AM, Jeff Findley wrote:
In article ,
ess says...

On 9/04/2016 8:24 AM, Jeff Findley wrote:

I'm tired of people saying "chemical propulsion is too
expensive" when the fuel is so damn cheap! Fuel costs
are *not* the problem. Skylon/Sabre is a solution to a
problem which *does not exist*!

Skylon/Sabre is not about reducing the amount of fuel
consumed.

Sorry, I was being sloppy and lumping the mass of the
oxidizer in with the fuel.

How about this. LOX is one of the cheapest fluids used in
the aerospace industry. It's literally made from air.
Trying to reduce LOX consumption is quite counter-intuitive
if reducing launch costs is the goal.

Jeff


The point of Sabre is not to reduce the amount of LOX that is
consumed, the cost of LOX being, as you point out, negligible
in this context.

The goal is to build an SSTO vehicle, with the economic
advantages that brings. Sabre is a means to that end, because
it reduces the amount of LOX that has to be *lifted*, and
allows atmospheric nitrogen to be used as reaction mass during
the air-breathing phase.

Sylvia.

Sabre is not a LACE. The air is only cooled to the vapour phase
boundary, not liquified. The cycle would be less efficient if the
air were liquified (even in part).

The system involves cooling the air at the rate it arrives. If this
is not achieved, then the engine won't work, but you can't a priori
say that it cannot be achieved.

Sylvia/

Here's what I got from online sources;

The RB545 is covered by the official secrets act, so not much can be
said about it. This is a barrier right at the outset.

The SABRE design is neither a conventional rocket engine nor jet
engine, but a hybrid that uses air from the environment at low
speeds/altitudes, and stored liquid oxygen (LOX) at higher altitude.

The SABRE engine "relies on a heat exchanger capable of cooling
incoming air to -150 °C (-238 °F), to provide oxygen for mixing with
hydrogen and provide jet thrust during atmospheric flight before
switching to tanked liquid oxygen when in space.

At the front of the engine, a simple translating axisymmetric shock
cone inlet slows the air to subsonic speeds using two shock
reflections.

Part of the air then passes through a precooler into the central
core, with the remainder passing directly through a ring of bypass
ramjets.

The central core of SABRE behind the precooler uses a
turbo-compressor run off the same gaseous helium loop Brayton cycle
which compresses the air and feeds it into four high pressure
combined cycle rocket engine combustion chambers.

The oxygen is also fed to the combustion unit, using a turbopump.

***

So, I don't get how you end up with oxygen without liquefying it. I
do get that using liquid hydrogen's larger specific heat its possible
to cool air, and in a sense compress it unconventionally. This seems
to me critical. Since it appears nowhere in any of the literature
available, and it is a fundamental feature, that suggests to me one
of the essential features hidden by the official secrets act around
the RB545.

There's quite a lot of information available in the technical documents
section of the Reaction Engines web site.

http://www.reactionengines.co.uk/tech_docs.html

I don't see a detailed entropy temperature graph of the oxygen and hydrogen
flow through the engine, but I'll keep looking.


Importantly, there is no separation of oxygen from the air during the
air breathing ascent. The cooled compressed air is fed into to the
rocket engines.

That's an improvement on the LACE concept certainly.

In 2011, hardware testing of the heat exchanger technology "crucial
to [the] hybrid air- and liquid oxygen-breathing [SABRE] rocket
motor" was completed, demonstrating that the technology is viable.
The tests validated that the heat exchanger could perform as needed
for the engine to obtain adequate oxygen from the atmosphere to
support the low-altitude, high-performance operation.

* * *

Not clear how this helps with high altitude high speed operation!
lol.

Regardless of the speed through the air, the inlet reduces the air speed
to subsonic.

This is a problem in that the stagnation temperature gets really high at
higher speeds.

Beyond mach 5.5, the craft operates as a pure rocket.

Yes, that's 1.87 km/sec. A little slower than I calculated. That means you
take more oxygen along.

The rest of your analysis appears to relate to something rather
different from the sabre engine and skylon.

Just being as generous to the concept as possible so that the advantages
were maximised for air breathing. I still think something like the Delta
Heavy or the Falcoln Heavy with three common core boosters, make a lot more
sense at the present time.

Sylvia.

I will modify my commentary to this extent. It might be made to work,
though I think LACE is totally unworkable as I understand it.

---

Combustion with LOX Methane

vacuum expansion, ε=40, chamber pressure 6.89 MPa
max. Isp 368.9 s ( Ve=3,617.6 m/s)
mixture ratio 3.45:1.00
chamber temperature 3563 K
bulk density 830 kg/m3
characteristic velocity 1838 m/s

Boiling point: Methane: −161.6°C (111.55 K) 511 kJ/kg heat of vaporisation.
Boiling point: Oxygen: -183.0°C ( 90.19 K) 214 kJ/kg heat of vaporisation.


To place 100 tons into LEO with a seven element launcher, consists of seven tanks that are 22.7 meters (74.5 ft) long and 4.1 m (13.4 ft) in diameter each carrying 197.63 tonnes of LOX/LNG and massing 9.19 tonnes empty. The entire system weighs 1547.73 tonnes take off weight and produce 2,100 tonnes force of thrust (300 tonnes force per element).

The aerospike engine on each element consists of a Raptor pump set and a modified Merlin pumpset. The Raptor feeds 30 Super Draco injectors and nozzles and the Merlin feeds 10 Super Draco injectors spaced around the zero height aerospike nozzle, that is equipped with an ablative heat sheild.

Both pump sets operate during ascent. The Merlin pumpset operates during descent. There is a collector line that operates the Merling driven engines during landing.

http://www.fireflyspace.com/assets/i...1014092659.png

The first four tanks accelerate to an idealised 2.586 km/sec, with gravity and air drag, is 1.6 km/sec. The next two tanks are dropped after an idealised 5.464 km/sec, with gravity and air drag is 5.2 km/sec. The final tank is emptied after an idealised 9.200 km/sec which after air drag and gravity losses add up to 7.9 km/sec.

One launch launches a 9.19 tonne tank and 90.81 tonnes propellant. A second launch puts up a similar payload. A third launch puts up 16.01 tonnes propellant and 74.80 tonnes manned payload, 9.19 tonne propellant tank.. Propellant from the two half empty tanks is transferred to the manned system, providing 197.63 tonnes of propellant.

The delta vee of this system is 3.095 km/sec. This is less than escape, but sufficient to project a payload to the moon totalling 74.80 tonnes, or the same amount to nearly escape velocity.

Entering lunar orbit requires a delta vee of 0.73 km/sec. Leaving lunar orbit requires another delta vee 0.73 km/sec. So, for a 74.80 tonne payload with an exhaust velocity of 3.6176 km/sec exhaust speed, 24.84 tonnes of propellant are needed to enter and leave low lunar orbit. This leaves 50.00 tonnes in LEO.

Now, a 100 kg payload landing on the lunar surface, orbiting at 1.56 km/sec and with a total delta vee of 3.12 km/sec requires 136.9 kg for each 100 kg astronaut with rocket belt. Now 350 kg is required for 40 days surival, and adding another 150 kg of propellant, is a total of 500 kg. So, 50 tonnes support the transport of 100 persons to the moon and back. Reducing the number to 25, requires only 12.5 tonnes, and taking up the difference of 37.5 tonnes as propellant allows 11 take offs and landings on the moon for each of the 25.

So, this is a quite exciting journey for the 25 persons! Each of which visits the moon a dozen times over a dozen days, and 8 more days to travel there and back.

Sending 500 tonnes to Low Lunar Orbit, and using 176 tonnes to land 8 BA330 inflatable habitats on the lunar surface, with recovery of the carrier craft only, is sufficient to provide space for 48 persons. So, 48 persons can be sent to a lunar hotel. Producing propellant from water found at the lunar poles using sunlight, provides a means to visit several spots on the moon via suborbital rocket belt.

A 100 ton payload is sufficient to orbit a 2200 MW solar power satellite, or a dozen 8 tonne comsats. Either of which supply tremendous revenue.

By adding 1.72 km/sec to the 10.995 km/sec transfers the 74.80 tonnes to Mars. Mars transfer requires 28.3 tonnes of propellant leaving 46.50 tonnes of payload, to Mars. Aerobraking into Mars orbit orbiting at 3.56 km/sec. Astronauts can enter the Mars atmosphere and use aerobraking to land, using rocket belts to slow the final 0.15 km/sec. Another 3.56 km/sec is required to return to orbit.

Here's an article about SpaceX that appeared FIVE YEARS AGO. Long before SpaceX successfully landed a recoverable Falcon on a platform.

Monday, April 18, 2011
Building Less Expensive Rocket Launchers

The title of the April 15th, 2011 Aviation Week article "China Great Wall Confounded By SpaceX Prices" seems self explanatory but the article makes the point explicit by stating that "executives at China Great Wall Industry Corp. are finding it hard to believe that California-based Space Exploration Technologies Inc. (SpaceX) is offering lower launch prices than they can."

Space-X Falcon 9.

The article then goes on to state that "Chinese officials say they find the published prices on the SpaceX website very low for the services offered, and concede they could not match them with the Long March series of launch vehicles even if it were possible for them to launch satellites with U.S. components in them."

According to the SpaceX website, a Falcon 9 rocket launch, with an advertised lift capacity of 10,450 kg. (23,000 lb.) from Cape Canaveral, costs between $54 million - $59.5 million dollars.

What is the secret of Space-X?

According to the June 18th, 2010 article "Elon Musk on Why His Rockets Are Faster, Cheaper and Lighter Than What You've Seen Before" from the Private Equity website (peHUB), there are a number of good reasons for the Space-X success including vertical integration, less outsourcing and some good old fashioned and very basic common sense.

The article quotes Space-X CEO Elon Musk as stating:

Some of what we've done is really just common sense -- for example,

(1) using the same propellant in the upper and lower stages means that operationally, you only need to have one set of fuel tanks. If you can imagine a situation where you have a kerosene first stage, hydrogen upper stage, and solid rocket side boosters, you've just tripled your cost right there.

(2) Also, the upper stage of Falcon 9 is simply a short version of the first stage. That may seem pretty obvious, but nobody else does that. They tend to create upper stage in a totally different way than they create the first stage.

(3) The Merlin engine -- we used it on the upper stage of Falcon 9, on the main stage of Falcon 9 and on the first stage of Falcon 1. So we get economies of scale in use of the Merlin engine.

(4) Our tanks are frictionl welded aluminum skin and stringer as opposed to machined aluminum, giving us a 20 fold advantage in the cost of materials, and our stage ends up being lighter ...because geometrically, we can have deeper stringers.

Others are slowly beginning to put together more detailed assessments of the secrets behind the Space-X pricing.

Several specific choices are credited with keeping costs down including:

(1) The decision to burn kerosene (RP-1) and liquid oxygen (LOX) for fuel instead of liquid hydrogen and LOX. Hydrogen, the lightest element, is difficult to cool and store as a liquid and not terribly dense even then so the fuel tank has to be quite large and complex. Kerosene is denser and much easier to store since it can be stored at room temperature. While NASA generally prefers hydrogen for it's greater specific impulse, most of the storage and cooling problems go away with kerosene, which provides substantial cost savings. LOX and kerosene are used in the lower stages of most Russian and Chinese boosters plus the first stages of the Saturn V and the Atlas V. LOX and liquid hydrogen are used in the upper stages of the Atlas V and Saturn V, the newer Delta IV rocket, the H-IIA rocket, and most stages of the European Ariane rockets.

(2) The decision not to use solid fuel rockets strapped to the side of the liquid fuel launchers, as is the case with the space shuttle is the second key to lowering costs. Solid rockets are useful when there is a need for long-term storage (and solid fuel rockets make great intercontinental ballistic missiles) but are less useful when you simply need something light and powerful, which is the prime criteria for satellite launchers. A second, solid fueled system added to an existing liquid fueled system also adds an additional level of complexity, which also adds to costs.

Space-X Manager of Business Development Josh Brost went into a little more detail on his firms methodologies, including;

(1) Rockets assembled and integrated horizontally, not vertically, which Space-X considers less expensive and less hazardous.

(2) Cross-training of technicians in multiple areas to insure that there are no narrow skilled specialists waiting around for specific work.

(3) An intentional decision to build more than 70% of vehicle components in-house at one location, which assists with the development of a direct feedback loop between on-site engineers and technicians as they work together on projects.

(4) Strong in-house and on-site capabilities in multiple areas including precision machining & inspection, tank fabrication, variable polarity plasma arc and friction stir welding, composites, precision assembly, tooling design/ fabrication, avionics laboratory & environmental testing, propulsion assembly and numerical control tube bending.

There are a lot of local political concessions built into most rocket programs and the real key to cost control is minimizing the various locations where the work is being done by doing as much as possible in-house and at one location.

This is the real secret of how they avoid the massive duplication of roles and the consequent cost overruns typical in aerospace development.

* * *

This was before a successful landing and reuse of the stages, and before switching to LOX/LNG combination, which offers lower costs and higher performance still.

The cost of natural gas is $3.50 per 1000 cubic feet. This is $0.168 per kg. According to the GAO NASA bought oxygen at 67 cents per gallon for use on the Space Shuttle. A gallon of liquid oxygen weighs 4.322 kg, so they paid $0.16 per kg for liquid oxygen.

So, for a Falcon 9 analogue, we have a 508,646 kg take off weight with a 30,557 kg first stage structure carrying 317,012 kg of propellant. We also have a 5,367 kg second stage structure carrying 55,667 kg propellant, and a 13,000 kg payload. A total of 372,679 kg of propellant at $59,629 per launch. This is $4.58 per kg to LEO. 35,954 kg of inert structure x $1500 per kg is $54,000,000. Reusing this 1000x reduces the cost to $54,000 per use. $114,000 per launch. Less than $10 per kg.

To place 100 tons into LEO with a seven element launcher described previously, consists of seven tanks that are 22.7 meters (74.5 ft) long and 4.1 m (13.4 ft) in diameter each carrying 197.63 tonnes of LOX/LNG and massing 9.19 tonnes empty. So, it costs $32,000 to fill up one tank, and $224,000 to fill up all seven tanks. At $1,500 per kg for the inert weight each system costs $14 million. $98 million to build all seven tanks. At 1000 reuses the cost is $98,000 per launch. $322,000 per launch. Divided by 100,000 kg this is $3.22 per kg. This is about 1/3 the previous number primarily because we're using 4.65% structure fractoin instead of 7.25% structure fraction and we're using 3.6 km/sec exhaust speed instead of 3.3 km/sec exhaust speed.

The entire system seven element system weighs 1547.73 tonnes take off weight and produce 2,100 tonnes force of thrust (300 tonnes force per element) to put 100 tonnes into LEO (220,000 lbs).

A 500 meter diameter sphere that's 1 micron thick made of structured silicene occupies 785.4 litres of volume and efficiently folds into a 523 mm diameter sphere that weights 942.5 kg. Placed into a polar sunrise sunset orbit, in constant sunlight with a 1 tonne launcher, it inflates to 500 meter diameter using solar power.

It collects 268.6 megawatts of power on orbit around Earth. It ejects material from its surface at 120 km/sec It absorbs light energy efficiently from the sun, and fires up an array of ion engines that eject 37.3 grams/sec of propellant producing 456.6 kgf of thrust, accelerating the satellite at 0..58 gees.

The satellite flies to GEO from LEO and takes up a stationary orbit around Earth. There is beams 268.6 watts to a large number of points on Earth in various amounts simultaneously, for $0.18 per kWh. Each satellite earns $423.8 million per year and is worth $4.2 billion the day its turned on.

Three satellites located at the prime meridian above the equator, (London) and at East 120 degrees (Hangzhou) and at West 120 degrees (Los Angeles). This network delivers power to wherever its needed on Earth, up to 805.8 MW and at $0.18 per kWh is worth $12.6 billion the day it is operational.

Using similar technology, a 5 km diameter sphere that's 1 micron thick made of structured silicene occupies 78.54 cubic meters of space and efficiently folds into a 5,320 mm diameter sphere that weighs 94.25 tonnes. Placed into LEO polar orbit, in constant sunlight, with a 100 tonne launcher, it inflates to 5 km diameter using solar power.

It collectes 26.86 billion watts on orbit around Earth, and ejects material from its surface at 120 km/sec. . It absorbs light energy efficintly from the sun across the range of visible colours, and fires up an array ion engines that eject 3.73 kg/sec of propellant producing 45.65 tonnes force of thrust, accelerating the satellite at 0.58 gees.

Sending an inflatable optical probe from Earth to Jupiter, to use gravity assist to enter a highly elliptical orbit above the solar poles, which at perihelion is circularised using solar sail technology, is the first step in creating a highly efficient solar powered system.

A solar pumped laser at 1/50th AU, operating at 3.42 MW per square meter, is 2500x solar intensity, which is outlined in my solar energy patent, and routinely achieved in the 1990s in my shop.

http://www.google.com/patents/US20050051205

US 7081584

An emitter operating at 250 nm wavelength over a distance of 800 AU can form a spot efficiently that's 5 km in diameter using an objective that's 5 km in diameter.

http://scitation.aip.org/content/aip...0.1063/1.93625
http://news.mit.edu/2013/chips-that-...eer-light-0109
http://www.deepspace.ucsb.edu/wp-con...aper_R05.p df


It takes 6.24 km/sec from LEO to enter a transfer orbit from Earth to Jupiter. It takes 18.28 minutes to boost from LEO to Jupiter transfer, and 3.98 tonnes. It takes 2.736 years to fly from Earth to Jupiter, and from there it is tossed into an orbit with a perihelion of 0.02 AU. This takes another 2.118 years.

At perihelion this sphere produces 67.15 terawatts of power and can form a beam with a spot size 9.1 meters in diameter at 1 AU distance! So, a very efficient power link is made between Earth and the Sun orbiting power satellite. 22.38 terawatts of power are now transmitted simultaneously through the three smaller satellites which now act to redirect the laser energy arriving from the Sun.

The surface absorbs the full spectrum of sunlight, and powers an array of UV lasers operating at the 350 nm wavelength, can efficiently send energy to a similar sphere sent to 1 AU distance, equipped to efficiently absorb and generate 350 nm wavelength energy.

At $0.01 per KWh, 67.15 billion kW generate 588.6 trillion kWh per year. At $0.01 per kWh this is $5.886 trillion per year. $0.01 per kWh is $16.95 per barrel of crude. At $5.19 per watt 67.15 trillion watts creates a value of 4.48x larger than the economy of today. This is $404.1 trillion per year. $54,756 per person per year.

A laser sustained rocket with a 9.2 km/sec exhaust speed, is capable ofputting up 497.5 tonnes at 21.05 billion watts. It masses 311.15 tonnes at take off, weighs 14.47 tonnes empty, carries 196.7 tonnes propellant and carries 100 tonnes to orbit. 3,190 ships like this could be flying simultaneously throughout the world.