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  #1  
Old November 20th 13, 04:14 AM posted to sci.space.policy
William Mook[_2_]
external usenet poster
 
Posts: 3,840
Default 3D Printed Rocket

3D Printed Rocket

http://www.youtube.com/watch?v=kFFeqmMprV4
http://www.youtube.com/watch?v=1uZHUkeIkpk
http://www.youtube.com/watch?v=VFXIN3ZvJYw

Sciaky Additive Manufacturing...
http://www.youtube.com/watch?v=A10XEZvkgbY

19 ft long, 4 ft wide, 4 ft high... off the shelf,
38 ft long, 8 ft wide, 8 ft high... available through special order.

Trumpf Additive Manufacturing...
http://www.youtube.com/watch?v=iLndYWw5_y8

5-axis CNC machining center
http://www.paragondie.com/worlds_largest_fidia.shtml

60 ft long, 12 ft wide, 12 ft high... off the shelf (can be combined with Trumpf additive manufacturing system)

Multiple feeds through special order with increases in length to 120 ft.

A 120 ft long, 24 ft wide, 24 ft high... available through special order.

* * *

The space shuttle external tank was 153.8 ft long and 27.6 ft in diameter. It carried 1.385 million pounds of LOX and 0.235 million pounds of LH2 and massed 1.672 million pounds at lift off.

A small version only 47.5% the mass and 78.0% the size of the larger tank would be 120 ft long and 21.6 ft in diameter. It would mass 0.794 million pounds at lift off and carry 0.658 million pounds of LOX and 0.112 million pounds of LH2.

With an array of small engines feeding an aerospike nozzle at the base of the vehicle it would be capable of 455 sec Isp from sea level to vacuum which translates to an exhaust speed of 14,650 ft/sec. The system produces 1.111 at million pounds of thrust at 111% rated value and may be throttled back to 1% full thrust. Thus it can produce up to 1.40 gees at lift off.

With a take-off weight of 0.794 million pounds and a combined propellant weight of 0.769 million pounds the self-propelled tank is capable of

14,650 * ln( 0.794 / (0.794 - 0.769) ) = 50,378 ft/sec = Vf

This is with zero payload. A speed of 34,380 mph. Greater than escape speed.

A payload of 86,851 pounds may be accelerated through a speed of 30,171 ft/sec. That's a speed of 20,571 mph. Sufficient to achieve the same orbits as the space shuttle! Gee forces at lift off is 1.26 gees. Similar again to the old space shuttle.

Three tanks operating in parallel, with the two outboard tanks equipped with cross-feeding, similar to the space shuttle external tank, to feed propellant to the central tank engine during lift off and first stage operation.

This creates in effect a two-stage system with the outboard tanks forming the first stage and the central tank forming the second stage.

Assuming the tanks are identical, we can know that take-off-weight is

3 * 0.794 million pounds

and that the propellant in the first stage is

2 * 0.769 million pounds

so we can calculate that the ideal delta vee of the first stage is;

14,650 * ln((3 * 0.794)/(3 * 0.794 - 2 * 0.769)) = 15,200 ft/sec

That's 10,363 mph - less air drag and gravity losses.

we can also calculate that the ideal delta vee of the second stage is the same as before namely 50,378 ft/sec or 34,380 mph.

Adding these two velocities together obtains 44,743 mph!!

This launcher is capable of carrying 302,500 lbs through a delta vee of 20,570 mph thus achieving the same orbits as the old space shuttle with this load. The first stage separates at 8,500 mph less air drag and gravity losses, and the second stage adds 12,070 mph less air drag and gravity losses.

A 55 ft long 21.6 ft diameter section is added above the aerospike nozzle and below the aft bulkhead on the tank. This payload bay lengthens the central tank from 120 ft to 175 ft. It carries a 302,500 lb 'interplanetary' stage. This stage consists of two parts. One that does trans-lunar injection adding 9,675 ft/sec to the orbiting mass. (6,600 mph). This takes both stages around the moon in a lunar free-return trajectory. When oriented toward interplanetary use, the stages have an apogee of 170,000 miles, and return to Earth.

On a lunar trip the booster stage loops around the Moon and returns to Earth, where it re-enters the atmosphere and is recovered near the launch centre.

The lunar stage executes a 7,480 ft/sec (5,100 mph) burn to come safely to rest on the lunar surface. It then executes another 7,480 ft/sec burn to travel from the Moon to the Earth. At Earth is re-enters the atmosphere and is recovered near the launch centre.


u = 1 - 1/exp(9,675/14,650) = 0.48336

and

0.48336 * 302,500 = 146,217 lbs.

Allowing 0.05000 for the structure fraction obtains

0.05000 * 302,500 = 15,125 lbs

for the inert weight of this stage.

Subtracting these figures from the total obtains a stage weight for the lunar lander stage of;

302,500 - 146,217 - 15,125 = 141,158 lbs

and for the lunar lander and return stage we have a propellant fraction of;

u = 1 - 1/exp(14,960/14,650) = 0.63983

so we have a propellant weight of;

0.63983 * 141,158 = 90,317 lbs

allowing the same 0.05000 for the structure fraction obtains this stage's inert weight budget;

0.05000 * 141,158 = 7,057 lbs

This leaves for the net payload on the moon (and back);

141,158 - 90,317 - 7,057 = 43,784 lbs

Allowing 350 lbs per person this is sufficient to carry 125 persons to the moon and back! With a crew of 5, 120 paying passengers may be carried to the moon and back aboard this ship.

With a diameter of 21.6 ft the stage and a 32 inch aisle down the center, seat width of 18 inches and seat pitch 36 inches this is sufficient for 74 seats and a central aisle

There are two rows of 7 seats on either side of the aisle, two rows of 6 seats on either side of the aisle, two rows of 4 seats on either side of the aisle and two rows of 3 seats at either end of the aisle.

This is economy seating.

First-class seating on the same plan provides 46 seats. Five crew members are located on the control deck and lounge, located between the economy and first class decks.

A total length of 22.5 ft for the cabins with a diameter of 21.6 ft as already stated.

There are up to 18 propellant tanks each 7.2 ft in diameter arrayed in three rings of 6 tanks each. 16 of these are typically used per mission. Two may be replaced with payload containers.

Each tank carries up to 15,060 lbs of propellant and has an inert mass of only 300 lbs each. The rings are arrayed around a central engine that these tanks feed. A fold away landing gear and inflatable aeroshell for atmospheric re-entry is also present.

http://www.space.com/16615-nasa-infl...-saturday.html

The trans-lunar insertion burn involves emptying 10 of the 16 tanks, leaving a single ring of tanks attached to the lunar lander. The 10 tanks, forming a ring of 6 atop a ring of 4. These detach prior to the lunar burn, and return to Earth for a landing at the launch centre.

The lander slows around the far-side of the moon, and slows to a descent trajectory. There it lands at the desired location on the moon. This burns through four of the six tanks. The empties may be left on the moon, along with the excess payload destined on a one way journey. The two remaining tanks possess sufficient propellants to carry the ship and its payload back to Earth.

* * *

302,500 lbs at Low Earth Orbit (LEO) is sufficient to place a sizable power satellite in space.

Using a 0.5 mil thick biaxially oriented PET film coated transparent on one side and coated with aluminium on the other, a total of 381.7 square feet of film may be covered per pound. And 2.35 million cubic feet may be filled per pound of hydrogen gas at low pressure.

http://www.grc.nasa.gov/WWW/RT/RT199...490tolbert.jpg

280,000 lb thin film optical system is 8,428 ft in diameter with an 8,000 ft diameter active area. It concentrates the sun to an image 80 ft in diameter where 6.25 GW of solar energy are concentrated. 17,983 hexagons cut from 200 mm diameter wafers form an active thin disk laser element all of which operate together to create a solar pumped laser array. Each element produces 275,000 watts of laser output with a slope efficiency approaching 80% - producing 4.97 GW of laser power beamed to Earth.

http://www.opticsinfobase.org/ao/abs...=ao-51-26-6382

Some wafers are MEMS based ion engines that are equipped to receive laser energy attached to the inflated optical system. These use the hydrogen gas as a propellant in a laser based ion engine - so that the satellite may maintain orientation and boost from Low Earth Orbit to Geosynchronous Equatorial Orbit. Sufficient hydrogen is maintained on board with solar powered cyrogenic MEMS based refrigeration, to last 30 years allowing for leakage in the optical system and for attitude control using the ion engines. All following the initial boost.

http://web.mit.edu/aeroastro/labs/spl/research_ieps.htm

http://iopscience.iop.org/0960-1317/...1E0ABD3A48C.c2

60 MW ground stations using a compact receiver of band-gap match photocells, produces hydrogen and oxygen from water when the satellite is visible in the sky. The hydrogen is then used in a fuel cell to produce electricity on demand at power levels up to 100 MW. Efficiency of throughput is 72%. Capital utilization of the receiver ranges from 65% to 95%. Average continuous output ranges from 28.08 MW to 41.04 MW for each station on the ground, generating 43.2 MW when illuminated.

Each satellite supports 100 ground stations of this type. It charges $0.04 per kWh and produces 42 billion kWh per year generating $1.68 billion per year in revenue per satellite.

With 100 MW peak and charging $1.25 per peak watt along with $0.11 per kWh of electricity used, each station has a $125 million capital cost and annual operating costs ranging from $27.1 million per year to $39.6 million per year. This translates to $500 million per ground station valuation for this revenue. Buyers obtain pollution free power at a fixed rate for 30 years, anywhere in sight of the satellite.

So, 100 ground stations and one satellite is worth at start-up $12.5 billion in capital payments and $50.0 billion in revenue capitalization. A combined valuation of EACH SATELLITE of $72.5 billion!!

A six year program to develop the supply chain for the launcher and satellite will cost $21 billion.

Paying 41.42% interest per year, compounded, against a convertible debenture will result in the sale of this asset along the following lines (flowing first to the early-adopters who pay for their ground stations)

In Millions of USD$ (2013AD) -

Total: $21,000.00 Project Cost:

FIRST SATELLITE:

At Risk Return Pct Total

Year 1: $ 807.69 $ 5,435.31 7.50%
Year 2: $1,615.38 $ 7,686.22 10.60%
Year 3: $3,230.77 $10,869.29 14.99%
Year 4: $6,461.54 $15,370.56 21.20%
Year 5: $4,846.15 $ 8,150.97 11.24%
Year 6: $4,038.46 $ 4,802.71 6.62%

Totals:$21,000.00 $52,315.08 72.16%

Present Value: $72,500.00 at start up - FIRST SATELLITE.

To replace all the the coal fired power plants in the world with these ground stations requires the deployment of 385 satellites. The recurring cost of each deployment is more than covered by the CAPEX. A small fleet launches 3 satellites per week. In 30 months all satellites are deployed and $1.27 trillion per year in power sales are received.

This is sufficient to support industrial development of the Moon, Mars, and the Main Belt Asteroids.

As well as the development of advanced laser propulsion systems, including personal ballistic transport on Earth.

Options on the FIRST satellite of $8.08 million per ground station, times 100 ground stations, completes the first round of production and testing to demonstrate all the elements described here, providing venture capital rates of return. We then sell over the next two years, the rights to the ground station to utilities. The final two years we finance through a variety of instruments against the off-take contracts.


  #2  
Old November 21st 13, 01:39 AM posted to sci.space.policy
William Mook[_2_]
external usenet poster
 
Posts: 3,840
Default 3D Printed Rocket

On Wednesday, November 20, 2013 4:14:08 PM UTC+13, William Mook wrote:
3D Printed Rocket



http://www.youtube.com/watch?v=kFFeqmMprV4

http://www.youtube.com/watch?v=1uZHUkeIkpk

http://www.youtube.com/watch?v=VFXIN3ZvJYw



Sciaky Additive Manufacturing...

http://www.youtube.com/watch?v=A10XEZvkgbY



19 ft long, 4 ft wide, 4 ft high... off the shelf,

38 ft long, 8 ft wide, 8 ft high... available through special order.



Trumpf Additive Manufacturing...

http://www.youtube.com/watch?v=iLndYWw5_y8



5-axis CNC machining center

http://www.paragondie.com/worlds_largest_fidia.shtml



60 ft long, 12 ft wide, 12 ft high... off the shelf (can be combined with Trumpf additive manufacturing system)



Multiple feeds through special order with increases in length to 120 ft.



A 120 ft long, 24 ft wide, 24 ft high... available through special order.



* * *



The space shuttle external tank was 153.8 ft long and 27.6 ft in diameter.. It carried 1.385 million pounds of LOX and 0.235 million pounds of LH2 and massed 1.672 million pounds at lift off.



A small version only 47.5% the mass and 78.0% the size of the larger tank would be 120 ft long and 21.6 ft in diameter. It would mass 0.794 million pounds at lift off and carry 0.658 million pounds of LOX and 0.112 million pounds of LH2.



With an array of small engines feeding an aerospike nozzle at the base of the vehicle it would be capable of 455 sec Isp from sea level to vacuum which translates to an exhaust speed of 14,650 ft/sec. The system produces 1..111 at million pounds of thrust at 111% rated value and may be throttled back to 1% full thrust. Thus it can produce up to 1.40 gees at lift off.



With a take-off weight of 0.794 million pounds and a combined propellant weight of 0.769 million pounds the self-propelled tank is capable of



14,650 * ln( 0.794 / (0.794 - 0.769) ) = 50,378 ft/sec = Vf



This is with zero payload. A speed of 34,380 mph. Greater than escape speed.



A payload of 86,851 pounds may be accelerated through a speed of 30,171 ft/sec. That's a speed of 20,571 mph. Sufficient to achieve the same orbits as the space shuttle! Gee forces at lift off is 1.26 gees. Similar again to the old space shuttle.



Three tanks operating in parallel, with the two outboard tanks equipped with cross-feeding, similar to the space shuttle external tank, to feed propellant to the central tank engine during lift off and first stage operation..



This creates in effect a two-stage system with the outboard tanks forming the first stage and the central tank forming the second stage.



Assuming the tanks are identical, we can know that take-off-weight is



3 * 0.794 million pounds



and that the propellant in the first stage is



2 * 0.769 million pounds



so we can calculate that the ideal delta vee of the first stage is;



14,650 * ln((3 * 0.794)/(3 * 0.794 - 2 * 0.769)) = 15,200 ft/sec



That's 10,363 mph - less air drag and gravity losses.



we can also calculate that the ideal delta vee of the second stage is the same as before namely 50,378 ft/sec or 34,380 mph.



Adding these two velocities together obtains 44,743 mph!!



This launcher is capable of carrying 302,500 lbs through a delta vee of 20,570 mph thus achieving the same orbits as the old space shuttle with this load. The first stage separates at 8,500 mph less air drag and gravity losses, and the second stage adds 12,070 mph less air drag and gravity losses..



A 55 ft long 21.6 ft diameter section is added above the aerospike nozzle and below the aft bulkhead on the tank. This payload bay lengthens the central tank from 120 ft to 175 ft. It carries a 302,500 lb 'interplanetary' stage. This stage consists of two parts. One that does trans-lunar injection adding 9,675 ft/sec to the orbiting mass. (6,600 mph). This takes both stages around the moon in a lunar free-return trajectory. When oriented toward interplanetary use, the stages have an apogee of 170,000 miles, and return to Earth.



On a lunar trip the booster stage loops around the Moon and returns to Earth, where it re-enters the atmosphere and is recovered near the launch centre.



The lunar stage executes a 7,480 ft/sec (5,100 mph) burn to come safely to rest on the lunar surface. It then executes another 7,480 ft/sec burn to travel from the Moon to the Earth. At Earth is re-enters the atmosphere and is recovered near the launch centre.





u = 1 - 1/exp(9,675/14,650) = 0.48336



and



0.48336 * 302,500 = 146,217 lbs.



Allowing 0.05000 for the structure fraction obtains



0.05000 * 302,500 = 15,125 lbs



for the inert weight of this stage.



Subtracting these figures from the total obtains a stage weight for the lunar lander stage of;



302,500 - 146,217 - 15,125 = 141,158 lbs



and for the lunar lander and return stage we have a propellant fraction of;



u = 1 - 1/exp(14,960/14,650) = 0.63983



so we have a propellant weight of;



0.63983 * 141,158 = 90,317 lbs



allowing the same 0.05000 for the structure fraction obtains this stage's inert weight budget;



0.05000 * 141,158 = 7,057 lbs



This leaves for the net payload on the moon (and back);



141,158 - 90,317 - 7,057 = 43,784 lbs



Allowing 350 lbs per person this is sufficient to carry 125 persons to the moon and back! With a crew of 5, 120 paying passengers may be carried to the moon and back aboard this ship.



With a diameter of 21.6 ft the stage and a 32 inch aisle down the center, seat width of 18 inches and seat pitch 36 inches this is sufficient for 74 seats and a central aisle



There are two rows of 7 seats on either side of the aisle, two rows of 6 seats on either side of the aisle, two rows of 4 seats on either side of the aisle and two rows of 3 seats at either end of the aisle.



This is economy seating.



First-class seating on the same plan provides 46 seats. Five crew members are located on the control deck and lounge, located between the economy and first class decks.



A total length of 22.5 ft for the cabins with a diameter of 21.6 ft as already stated.



There are up to 18 propellant tanks each 7.2 ft in diameter arrayed in three rings of 6 tanks each. 16 of these are typically used per mission. Two may be replaced with payload containers.



Each tank carries up to 15,060 lbs of propellant and has an inert mass of only 300 lbs each. The rings are arrayed around a central engine that these tanks feed. A fold away landing gear and inflatable aeroshell for atmospheric re-entry is also present.



http://www.space.com/16615-nasa-infl...saturday..html



The trans-lunar insertion burn involves emptying 10 of the 16 tanks, leaving a single ring of tanks attached to the lunar lander. The 10 tanks, forming a ring of 6 atop a ring of 4. These detach prior to the lunar burn, and return to Earth for a landing at the launch centre.



The lander slows around the far-side of the moon, and slows to a descent trajectory. There it lands at the desired location on the moon. This burns through four of the six tanks. The empties may be left on the moon, along with the excess payload destined on a one way journey. The two remaining tanks possess sufficient propellants to carry the ship and its payload back to Earth.



* * *



302,500 lbs at Low Earth Orbit (LEO) is sufficient to place a sizable power satellite in space.



Using a 0.5 mil thick biaxially oriented PET film coated transparent on one side and coated with aluminium on the other, a total of 381.7 square feet of film may be covered per pound. And 2.35 million cubic feet may be filled per pound of hydrogen gas at low pressure.



http://www.grc.nasa.gov/WWW/RT/RT199...490tolbert.jpg



280,000 lb thin film optical system is 8,428 ft in diameter with an 8,000 ft diameter active area. It concentrates the sun to an image 80 ft in diameter where 6.25 GW of solar energy are concentrated. 17,983 hexagons cut from 200 mm diameter wafers form an active thin disk laser element all of which operate together to create a solar pumped laser array. Each element produces 275,000 watts of laser output with a slope efficiency approaching 80% - producing 4.97 GW of laser power beamed to Earth.



http://www.opticsinfobase.org/ao/abs...=ao-51-26-6382



Some wafers are MEMS based ion engines that are equipped to receive laser energy attached to the inflated optical system. These use the hydrogen gas as a propellant in a laser based ion engine - so that the satellite may maintain orientation and boost from Low Earth Orbit to Geosynchronous Equatorial Orbit. Sufficient hydrogen is maintained on board with solar powered cyrogenic MEMS based refrigeration, to last 30 years allowing for leakage in the optical system and for attitude control using the ion engines. All following the initial boost.



http://web.mit.edu/aeroastro/labs/spl/research_ieps.htm



http://iopscience.iop.org/0960-1317/...1E0ABD3A48C.c2



60 MW ground stations using a compact receiver of band-gap match photocells, produces hydrogen and oxygen from water when the satellite is visible in the sky. The hydrogen is then used in a fuel cell to produce electricity on demand at power levels up to 100 MW. Efficiency of throughput is 72%. Capital utilization of the receiver ranges from 65% to 95%. Average continuous output ranges from 28.08 MW to 41.04 MW for each station on the ground, generating 43.2 MW when illuminated.



Each satellite supports 100 ground stations of this type. It charges $0.04 per kWh and produces 42 billion kWh per year generating $1.68 billion per year in revenue per satellite.



With 100 MW peak and charging $1.25 per peak watt along with $0.11 per kWh of electricity used, each station has a $125 million capital cost and annual operating costs ranging from $27.1 million per year to $39.6 million per year. This translates to $500 million per ground station valuation for this revenue. Buyers obtain pollution free power at a fixed rate for 30 years, anywhere in sight of the satellite.



So, 100 ground stations and one satellite is worth at start-up $12.5 billion in capital payments and $50.0 billion in revenue capitalization. A combined valuation of EACH SATELLITE of $72.5 billion!!



A six year program to develop the supply chain for the launcher and satellite will cost $21 billion.



Paying 41.42% interest per year, compounded, against a convertible debenture will result in the sale of this asset along the following lines (flowing first to the early-adopters who pay for their ground stations)



In Millions of USD$ (2013AD) -



Total: $21,000.00 Project Cost:



FIRST SATELLITE:



At Risk Return Pct Total



Year 1: $ 807.69 $ 5,435.31 7.50%

Year 2: $1,615.38 $ 7,686.22 10.60%

Year 3: $3,230.77 $10,869.29 14.99%

Year 4: $6,461.54 $15,370.56 21.20%

Year 5: $4,846.15 $ 8,150.97 11.24%

Year 6: $4,038.46 $ 4,802.71 6.62%



Totals:$21,000.00 $52,315.08 72.16%



Present Value: $72,500.00 at start up - FIRST SATELLITE.



To replace all the the coal fired power plants in the world with these ground stations requires the deployment of 385 satellites. The recurring cost of each deployment is more than covered by the CAPEX. A small fleet launches 3 satellites per week. In 30 months all satellites are deployed and $1.27 trillion per year in power sales are received.



This is sufficient to support industrial development of the Moon, Mars, and the Main Belt Asteroids.



As well as the development of advanced laser propulsion systems, including personal ballistic transport on Earth.



Options on the FIRST satellite of $8.08 million per ground station, times 100 ground stations, completes the first round of production and testing to demonstrate all the elements described here, providing venture capital rates of return. We then sell over the next two years, the rights to the ground station to utilities. The final two years we finance through a variety of instruments against the off-take contracts.


Another seat layout for a 21.6 ft diameter cabin involves 40 seats facing inward with four spaces leading to a door. This creates 4 arcs of 10 seats covering 90 degrees of arc.

Then, on a circle 32 inches inward - the seat pitch of a typical airliner, another 32 seats facing outward four aisles aligned with the doors - broken into four arcs of 8 seats each. These centres of these seats are aligned with the armrests of the outer circle and oriented so that both aisles have tremendous leg room. The circular aisle itself is wide as well.

32 inches inward again four arcs with 5 seats facing outward totalling another 20 seats with four aisles to the center.

This is a total of 92 seats on one deck.

At the centre is a 67.2 inch diameter space that has a 40 inch hatch centered above on the ceiling and below on the floor to allow travel between decks along a ladder. A pressure door is fitted in each hatch.

There is storage above each seat, and autostereoscopic displays along the exterior behind the outermost ring of seats and interior walls behind the innermost ring of seats, provides views of the outside and the lighting gives a feeling of spaciousness to the cabin.

Along the hatchway at the centre of the cabin are stores for food and drink.. At the four hatches at the periphery there is a fold-out tent like structure that forms a zero gee toilet - allowing four toilets for the 92 passengers.

184 passengers and 10 service staff and 6 flight crew is 200 persons. With a payload capacity of 43,784 pounds this is 218.2 pounds per person. Allowing 187 pounds per person for body weight and luggage, leaves 31.2 pounds per person for food drink and air. A ten day trip leaves 3.12 pounds per person per day on average.

So, this is quite possible.

Restocking the ship on the moon, and refuelling the ship on the moon, increases provisions immeasurably.

Still, a fully and highly reusable launcher that has very low operating costs, combined with this sort of payload, would carry out a mission to the moon and back for $20 million per flight. This is a cost of $100,000 per passenger.

According to Credit Suisse, as of October 2012 there were;

25,613,500 people with $1 million to $5 million in the bank
1,921,000 people with $5 million to $10 million in the bank
928,000 people with $10 million to $50 million in the bank
84,500 people with $50 million or more in the bank

These are world-wide totals.

Now someone with $1 million in the bank likely has an income that allows them to support a $1 million purchase. Its not likely they will commit that level of resource, but they can do it.

http://depts.washington.edu/amath/co...dale_wolfe.pdf


RESPONSE POPULATION ADOPTERS ASSET CLASS (Cash on Deposit)

3.00% 84,500 2,534 - $50 million+
1.68% 928,000 15,644 - $10 million to $50 million
0.95% 1,921,000 18,224 - $ 5 million to $10 million
0.53% 25,613,500 136,643 - $ 1 million to $ 5 million

TOTAL per year 173,045

Persons/flight 200
Flights/year 865

Flights/day 3

Equipment will take 10 days per flight cycle, so a market likely exists at these price levels ($1 million per trip) to support thirty (30) lunar landers.

Now, the booster has a three day turn around - and so only 10 boosters will support 30 landers - to support this level of tourism.

A profit of $500,000 per flight times 173,000 per year translates to $86.5 billion per year EBITDA. A market-cap to earnings of 23 to 1 implies a market cap of $1.99 trillion for the 30 landers and 10 boosters (20 flight elements).

The inert weight of the lander is 7,057 pounds. The flight elements mass 15,125 pounds. At $8,200 per pound this translates to $124 million for each booster and $57.9 million for each lander. So, 10 boosters (of 3 elements each) is $3.75 billion and 30 landers $1.75 billion - a total CAPEX of $5.00 billion. The launch centre, development cost and supply chain cost another $3.50 billion based on other parameters.

An $8 billion investment turning into a $2 trillion asset in six years!

Awesome!

A flight leaving Earth every 8 hours, and one arriving every 8 hours, with a seven day turn-around - (3 day out, 3 day back, 1 day on lunar surface) and three day turn-around at Earth, means 21 in flight and 9 on the ground. This gives the scale of the programme.

Over 40,000 pounds to the moon per trip. That's a capacity of 120,000 pounds per day. 5,000 pounds per hour. That's 43.8 million pounds per year. A person consumes 3 pounds per day, air, water, food. With recycling and locally supplied water source, this is radically reduced. However, wealthy individuals consume 30 pounds per day of imported stuff anyway. So, with this range, a fleet of 30 cargo ships and 10 additional three element boosters support 40,000 people minimally and 4,000 who live very well on the moon.. With recycling and local production these figures will be even higher.

  #3  
Old November 22nd 13, 04:14 AM posted to sci.space.policy
William Mook[_2_]
external usenet poster
 
Posts: 3,840
Default 3D Printed Rocket

Here's an interesting article:

http://iopscience.iop.org/1367-2630/...5011/fulltext/
http://scitation.aip.org/content/aip...1063/1.3520417
http://www-pub.iaea.org/MTCD/publica...apers/p-1p.pdf

The off-axis-parabolic-mirror (OAPM) that reflects energetic particles efficiently is a step up from the pusher plate concepts of the 1940s and 50s.

A microscopic fission free Jetter Cycle Fusion pulse unit operating at the focal point of a parabolic mirror that efficiently reflects alpha particles produced by the pulse unit efficiently generates collimated beams of alpha particles.

Consider the parabolic formula;

y = x^2 / 4

The focal point is at x=0, y=1. Draw the parabola from x=-7 to x=+7 and you get a nice shape that rises to y=12.25 at the lip. The focal point is 11.25 y-units back from that lip and at the center x=7 units.

This means the hypotenus is r=sqrt(11.25^2 + 7^2) = 13.25 units

This also means that the opening covers a 31.89 degree half angle;

theta = asin(7/13.25)*180/pi() = 31.89

The surface area of a sphere 13.25 units in radius is given by;

A = 4*pi()*r^2 = 2,206.18 units

The surface area of the curved end cap not occluded by the parabola of rotation is

A = 2*pi()*h*r = 166.50 units

This means the the parabola directs all but

166.50 / 2,206.18 = 0.075472

of the plasma burst. Of course that plasma burst that is not redirected into a collimated beam is spread across a 31.89 degree half angle. Which means that integrating over this range that 0.84906 (or around 85%) of the blast momentum propels the ship forward from plasma exiting this region. This is 7.5472% of the total, so

1.0000 - 0.075472 * 0.84906 = 0.93592 ~ 93.5%

of the momentum from the plasma exhaust is producing thrust. The other 6.5% can actually be used to direct the thrust by offsetting the blast point slightly from the centre line near the focus.

http://en.wikipedia.org/wiki/Blast_wave

A mm diameter fusion dot that consists of 434.59 nanograms of lithium-6 deuteride and 50 micrograms of detonator. This releases 105,170 Joules of energy when detonated. Equivalent to 2.74 cc of petrol. The plasma velocity is 2,051 km/sec.

A thousand detonations per second produce 50e-6 kilograms per second which yield

F = mdot * Ve = 1000 * 50e-9 * 2.051e6 = 102.55 N -- 10.46 kg.

Of course only 93.5% of this thrust is useful so this nets out at

0.935 * 10.46 = 9.78 kg

So, the thrust detonation rate relation is;

F(kg) = 102.227 * Rate

A nozzle 2.8 cm in diameter and 2.5 cm tall that produces 1,000 fusion detonations per second of 50 microgram sized fusion dots produces 9.78 kgf of thrust.

The lunar lander cabin described previously equipped with six 7 foot spheres containing fusion dots, detonators and delivery mechanism.

The delivery mechanism is interesting.

The dots move on their own, powered by a microwave beam flooding the containment area. The self-select and insert themselves in the dot ejector unit.

The average density of the dot is 3.5 g/cc. Packing of dots is limited to 74% of the available volume due to the way things pack together. So, the average density is 2.6 g/cc. So a 7 ft diameter sphere described previously contains 48.27 metric tons of fusion dots. Six of these spheres contain a total of 289.6 metric tons of fusion dots. That's 637,120 pounds. The vehicle and payload weighs less than 1/10th of this.

Increasing the weight to 220,000 pounds - or 100 metric tons - allows ten decks along the 120 ft length of the 21.6 ft diameter ship. This increases the nature of the cabins for 200 persons, or increases the number of persons to 2,000 persons! lol.

You could also make a space yacht.

With 637,000 pounds of propellant and 220,000 pounds of payload and ship, and an exhaust speed of 2,051 km/sec (6.7273 million ft/sec) - this ship is capable of attaining 2,789 km/sec (almost 1% of light speed!)

The vehicle is capable of 79 hours of constant one gee boost.

A trip to the moon takes place in just a few hours (3.75) with this ship, and the crew have a one gee environment throughout the flight.

This ship can travel up to 200 million km at one gee - this spans the inner solar system from Earth in 79 hours. The ship can travel up to 2,000 million km at one tenth gee in 790 hours. This spans the outer solar system.

So, this is quite a handy little ship. The base is large enough to accommodate an array of 78,000 thrusters of the type described above which is sufficient to operate at 2 gees maximum thrust at 1,100 pulses per second each.

Looking something like this when landed.

http://www.matterhorn1959.com/blog1/douglas1.jpg

The nose has four mobile airlocks that ride on tracks down the side of the vehicle when landed and operate like airlock and elevator which take people down to the surface.

An industry standard 3500 pound capacity elevator utilizing 42" center opening doors with an outside platform dimension of 7'-0" wide by 6'-2" deep and utilizing an interior wall finish thickness of 1 3/4" on each wall, yield an inside car dimension of approximately 6'-6" wide by 5'-2" deep.

26 people in normal dress can be carried by each elevator. 15 people in space suits.

The elevators are stowed during atmospheric transit. They are used when landed or when in free space.
  #4  
Old December 14th 13, 12:48 PM posted to sci.space.policy
William Mook[_2_]
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http://www.gizmag.com/micro-laser-si...30115/pictures

http://3dmicroprint.com/

http://www.eos.info/eos_pressemitteilung_3dmicorprint

Submicron resolution over 450 mm cube - 18 inch cube.

A sphere 450 mm in diameter enclosing another sphere 284.4 mm in diameter containing 12.04 litres of LOX massing 13.73 kg. The volume between the two spheres is the difference between 47.71 litres and 12.04 litres or 35.67 litres. This contains 2.50 kg of liquid hydrogen. A total of 15.23 kg of propellant.

The sphere masses 1.05 kg and consists of an array of micromachinery capable of performing a wide array of functions, including that of a propulsive skin, phased array antenna, super camera array to create a panoramic picture in all directions that is easily explored on a standard PC, like this;

http://www.youtube.com/watch?v=Th5zlUe6gOE#t=56

http://jonaspfeil.de/ballcamera


A 1.05 kg empty weight and a 16.28 kg full weight with hydrogen oxygen propellant in an efficient nozzle design implies a final speed of a single stage of;

Vf = 4.5 * ln(16.28 / 1.05) = 12.33 km/sec.

Not bad! This system is capable of leaving the Earth altogether and flying around the moon and return to Earth, even accounting for air drag and gravity losses. This system is capable of flying around the moon and returning to Earth, but not capable of landing on the moon after leaving Earth.

Of course, this could also fly to any point on Earth, take photos, hover for a bit, and return to its starting point - causing comment among viewers at the recon point perhaps.

http://www.youtube.com/watch?v=lDsXl7hAv3Q

A 450 mm diameter sphere covered with a high efficiency solar panel array operating at 55% overall efficiency produces 119.66 watts travelling through space in sunlight near 1 AU. The system using fuel cells to produce power at 100 watts continuously uses 47.6 milligrams of hydrogen and 333.2 milligrams of oxygen to produce 380.8 milligrams of water per minute. The ability to store water on board that is electrolyzed by solar powered fuel cells, and generate electricity on demand using those same fuel cells, independently of lighting conditions. 100 grams of water storage permit 4 hours and 25 minutes of operation without loss of propellant, with propellant restored using solar power.

Two spheres, one lifting the other and then separating, provides an initial boost of 2.83 km/sec. The second sphere starting at this speed attains an additional 12.33 km/sec giving a total delta vee of 15.16 km/sec.

This is sufficient to land on the moon, and give an additional 0.81 km/sec delta vee. Sufficient to land, navigate around the landing site, and visit another landing site via suborbital boost. Thus, several Apollo sites could be visited and thoroughly photographed by this system, after a soft landing.

Three spheres, one lifting two, and the second lifting the last, achieves an additional 1.68 km/sec - a total of 16.84 km/sec. Sufficient for the last sphere to land on the moon and take off again with 0.19 km/sec delta vee capability while on the lunar surface - again enough to do significant navigation around the lunar landing site.

http://baseball.physics.illinois.edu/WattsFerrerAJP.pdf

A spinning sphere is capable of inducing a L/D of 0.6 - this suggests that a 1.05 kg sphere, with 0.05 kg of excess hydrogen in the depleted tank, that is combined with atmospheric oxygen for propulsion via MEMS based hydrogen fuelled jet, would be capable of returning to the launch point when operating as either the first or second stages in the system thus described. So, all three components are fully recovered in these operations.
  #5  
Old January 13th 14, 12:31 AM posted to sci.space.policy
William Mook[_2_]
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A 450 mm (17.7 inch) diameter sphere containing a smaller sphere 284.4 mm (11.2 inches) equipped so that the small sphere carries 13.73 kg of LOX and the larger sphere contains 2.50 kg of LH2 and the sphere's themselves mass 1.05 kg. So we have a single sphere capable of a final velocity of;

Vf = 4.5 * ln(16.28 / 1.05) = 12.33 km/sec.

So, a take of weight for a sphere of 24.09 kg containing 15.23 kg of propellant (13.73 + 2.50) propels 8.86 kg through 4.5 km/sec. With 1.05 kg of 'structure' we have a structure fraction of 1.05 / 24.09 = 4.36% and propellant fraction of 15.23 / 24.09 = 63.23% leaving 100% - 63.23% - 4.36% = 32.41% payload.

Now, if we say a sphere carries a payload through a delta vee of 4.5 km/sec and we have four stages, two stages to orbit Earth, one stage to go from Earth orbit to the vicinity of the Moon, and finally, one stage to go from the Moon, to the Moon's surface, and back to Earth.

With a biosuit, and supplies for 14 days, an astronaut leaving Earth, massing 85 kg (on average) in a long-duration mechanical counter-pressure suit and supplies bringing the total to 140 kg. That's the payload.

Working backwards we have

140 kg / 0.3241 = 431.97

and so

431.97 * 0.0436 / 1.05 = 17.93 ~ 18

Eighteen spheres take a fully equipped astronaut with supplies for 12 days - which mean about 18 kilos of samples - from the lunar surface to Earth's surface.

We then do the same thing again using the 18 spheres + 140 kg as payload

18 * 16.28 + 140 = 433.04

and so;

433.04 / 0.3241 = 1,336.14

therefore;

1,336.14 * 0.0436 / 1.05 = 55.48 ~ 56

spheres... to take the return spheres and the astronaut from Earth orbit to the vicinity of the moon.

56 * 16.28 + 433.04 = 1,344.72

Using this as payload, we calculate the upper stage of the launcher... in terms of spheres.

1,344.72 / 0.3421 = 3930.79 kg

which translate to;

3930.79 * 0.0436 / 1.05 = 163.22 ~ 164 spheres

and so, to calculate the first stage and take off weight;

164 * 16.28 + 3930.79 = 6600.71

6600.71 * 0.0436 / 1.05 = 274.09 ~ 275

275 Stage 1
164 Stage 2
56 Stage 3
18 Stage 4

513 Total

513 * 16.28 + 140 = 8,491.64 kg TAKE OFF WEIGHT.

So, a disk of hexagonal close packed array of spheres 8.55 meters (28.0 ft) in diameter topped by another disk of hcp array of spheres 7.83 meters (25..7 ft) in diameter, is sufficient to take an astronaut to the moon and back, and return ALL the spheres to be reused to the launch point.
  #6  
Old January 13th 14, 07:32 AM posted to sci.space.policy
William Mook[_2_]
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Posts: 3,840
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Alright, so we have a hexagon assembled as follows;

19 spheres in a line. Two lines on either side 18 spheres long. Two lines 17 spheres long. 16,15,14,13,12,11,10,9 - so 10 lines on either side ranging from 17 spheres long to 9 spheres long.

18 + 9 = 27
17 + 10 = 27
16 + 11 = 27
15 + 12 = 27
14 + 13 = 27

On each side of the 19 sphere line...

10*27+19 = 289 spheres

in the first disk. So, if we had 275 spheres in this layer, we would have 14 spares - places for landing gear and other equipment.

Now we stack a disk of spheres on top of this one, between the spheres in hexagonal close pack fashion...

18 spheres in a line. Two lines on either side 17 spheres long. Two lines on either side 16 spheres long. 16,15,14,13,12,11,10,9 - so 9 lines on either side ranging from 17 to 9 spheres in length. This forms a hexagon 9 sphere diameters on a side.

17 + 9 = 26
16 + 10 = 26
15 + 11 = 26
14 + 12 = 26
13

On either side - so we have 9 * 26 + 18 = 252

Well, we have 164 + 56 + 18 = 230

So, there are plenty of locations... 22 locations to spare.

The last two stages are 56 + 18 = 74

So, a line of spheres 10 long, with five rows on either side, 9,8,7,6,5 total

2x 9+5=14, 8+6=14, 7 + 10 = 80

which is 6 more locations than needed.

This is both the third and fourth stages. The last stage is only 18. Here we have a line of 5 with 4 on either side and 3 on either side. A hexagon that's 3 units long on a side and 6 units in diameter.

5 + 2* (4+3) = 19

Which is 1 more space than needed.

Here's an earlier design using larger spheres

http://www.scribd.com/doc/40549127/Disk-Moonship

http://www.scribd.com/doc/40623446/Disk-Moonship-2

The current proposal uses 0.45 m diameter spheres. The project costs $22 million per person to the moon. The project time is five years - that's 2019 - the fiftieth anniversary of Apollo.

Budget (thousands)

$1,048 Test Complete
$2,095 Sphere on Orbit
$4,190 Person Ballistic
$8,381 Person on Orbit
$6,286 Person on Moon

$22,000

  #7  
Old January 13th 14, 09:08 PM posted to sci.space.policy
William Mook[_2_]
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Posts: 3,840
Default 3D Printed Rocket

This is an interesting document;

http://www.foia.af.mil/shared/media/...090218-169.pdf

Aircraft 9532
Power Plant 10450
Cockpit 1060
Total Empty 21050
Fuel 7750
Crew 200
Total 29000

With today's materials and power plants these weights may be cut by half, leaving a 10,000 pound payload in this same airframe. The fuel contains 19.45 MJ/lb. Thus 7750 lbs contains 150.8 GJ of energy. Replacing the JetA fuel with liquid hydrogen which contains 64.45 MJ/lb reduces that weight by 5410 lbs to 2340 lbs. At 0.56 lbs per gallon this increases the volume of the fuel tank to 4,179 gallons from 1,142 gallons. So, even allowing for increased fuel tank volume and mass, total payload is further increased to 15,000 lbs.

The range is intercontinental at 4,000 mph (3,470 knots) at 90,000 ft.

Compare this to a Learjet 60


Crew: 2
Capacity: 8 passengers
Length: 58 ft 8 in (17.88 m)
Wingspan: 43 ft 9 in (13.34 m)
Height: 14 ft 8 in (4.47 m)
Wing area: 264.5 ft² (24.57 m²)
Empty weight: 14,640 lb (6,641 kg)
Max. takeoff weight: 23,500 lb (10,660 kg)
Powerplant: 2 × Pratt & Whitney Canada PW305A turbofan, 4,600 lbf (20..46 kN) each

Performance

Maximum speed: 522 mph (453 knots, 839 km/h) (max cruise)
Cruise speed: 484 mph (Fast Cruise 536 mph) (420 knots, 778 km/h, Mach 0.74) (long-range cruise)
Range: 2,773 mi (2,409 nmi, 4,461 km)
Service ceiling: 51,000 ft (15,545 m)
Rate of climb: 4,500 ft/min (22.9 m/s)

Ten captain's chairs, facing outward, spaced around an 8 foot diameter circle, providing seating. Retractable canopies over each seat provides direct access to the outside, over the radial wing, which extends to form the disk shown in the document. The cabin is 5.7 feet tall. There is an 8 foot hemisphere atop a 5.7 foot tall cylinder and another 8 foot hemisphere below - to create a tank containing the liquid hydrogen. Smaller more efficient turbo jets in larger number provide power around the rim of the radial wing.

Two of the ten seats are equipped with flight controls and a heads up display on the canopy.
  #8  
Old January 13th 14, 10:44 PM posted to sci.space.policy
Rick Jones
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Posts: 685
Default 3D Printed Rocket

William Mook wrote:
This is an interesting document;


http://www.foia.af.mil/shared/media/...090218-169.pdf


Isn't that the failed Avrocar?

rick jones

Aircraft 9532
Power Plant 10450
Cockpit 1060
Total Empty 21050
Fuel 7750
Crew 200
Total 29000


With today's materials and power plants these weights may be cut by half, l=
eaving a 10,000 pound payload in this same airframe. The fuel contains 19.=
45 MJ/lb. Thus 7750 lbs contains 150.8 GJ of energy. Replacing the JetA f=
uel with liquid hydrogen which contains 64.45 MJ/lb reduces that weight by =
5410 lbs to 2340 lbs. At 0.56 lbs per gallon this increases the volume of =
the fuel tank to 4,179 gallons from 1,142 gallons. So, even allowing for i=
ncreased fuel tank volume and mass, total payload is further increased to 1=
5,000 lbs. =20


The range is intercontinental at 4,000 mph (3,470 knots) at 90,000 ft. =20


Compare this to a Learjet 60



Crew: 2
Capacity: 8 passengers
Length: 58 ft 8 in (17.88 m)
Wingspan: 43 ft 9 in (13.34 m)
Height: 14 ft 8 in (4.47 m)
Wing area: 264.5 ft=B2 (24.57 m=B2)
Empty weight: 14,640 lb (6,641 kg)
Max. takeoff weight: 23,500 lb (10,660 kg)
Powerplant: 2 =D7 Pratt & Whitney Canada PW305A turbofan, 4,600 lbf (20=
.46 kN) each


Performance


Maximum speed: 522 mph (453 knots, 839 km/h) (max cruise)
Cruise speed: 484 mph (Fast Cruise 536 mph) (420 knots, 778 km/h, Mach =
0.74) (long-range cruise)
Range: 2,773 mi (2,409 nmi, 4,461 km)
Service ceiling: 51,000 ft (15,545 m)
Rate of climb: 4,500 ft/min (22.9 m/s)


Ten captain's chairs, facing outward, spaced around an 8 foot diameter circ=
le, providing seating. Retractable canopies over each seat provides direct=
access to the outside, over the radial wing, which extends to form the dis=
k shown in the document. The cabin is 5.7 feet tall. There is an 8 foot h=
emisphere atop a 5.7 foot tall cylinder and another 8 foot hemisphere below=
- to create a tank containing the liquid hydrogen. Smaller more efficient=
turbo jets in larger number provide power around the rim of the radial win=
g.


Two of the ten seats are equipped with flight controls and a heads up displ=
ay on the canopy.


--
No need to believe in either side, or any side. There is no cause.
There's only yourself. The belief is in your own precision. - Joubert
these opinions are mine, all mine; HP might not want them anyway...
feel free to post, OR email to rick.jones2 in hp.com but NOT BOTH...
  #9  
Old January 17th 14, 12:24 PM posted to sci.space.policy
William Mook[_2_]
external usenet poster
 
Posts: 3,840
Default 3D Printed Rocket

On Tuesday, January 14, 2014 10:44:05 AM UTC+13, Rick Jones wrote:
William Mook wrote:

This is an interesting document;




http://www.foia.af.mil/shared/media/...090218-169.pdf




Isn't that the failed Avrocar?



rick jones



Aircraft 9532


Power Plant 10450


Cockpit 1060


Total Empty 21050


Fuel 7750


Crew 200


Total 29000




With today's materials and power plants these weights may be cut by half, l=


eaving a 10,000 pound payload in this same airframe. The fuel contains 19.=


45 MJ/lb. Thus 7750 lbs contains 150.8 GJ of energy. Replacing the JetA f=


uel with liquid hydrogen which contains 64.45 MJ/lb reduces that weight by =


5410 lbs to 2340 lbs. At 0.56 lbs per gallon this increases the volume of =


the fuel tank to 4,179 gallons from 1,142 gallons. So, even allowing for i=


ncreased fuel tank volume and mass, total payload is further increased to 1=


5,000 lbs. =20




The range is intercontinental at 4,000 mph (3,470 knots) at 90,000 ft. =20




Compare this to a Learjet 60






Crew: 2


Capacity: 8 passengers


Length: 58 ft 8 in (17.88 m)


Wingspan: 43 ft 9 in (13.34 m)


Height: 14 ft 8 in (4.47 m)


Wing area: 264.5 ft=B2 (24.57 m=B2)


Empty weight: 14,640 lb (6,641 kg)


Max. takeoff weight: 23,500 lb (10,660 kg)


Powerplant: 2 =D7 Pratt & Whitney Canada PW305A turbofan, 4,600 lbf (20=


.46 kN) each




Performance




Maximum speed: 522 mph (453 knots, 839 km/h) (max cruise)


Cruise speed: 484 mph (Fast Cruise 536 mph) (420 knots, 778 km/h, Mach =


0.74) (long-range cruise)


Range: 2,773 mi (2,409 nmi, 4,461 km)


Service ceiling: 51,000 ft (15,545 m)


Rate of climb: 4,500 ft/min (22.9 m/s)




Ten captain's chairs, facing outward, spaced around an 8 foot diameter circ=


le, providing seating. Retractable canopies over each seat provides direct=


access to the outside, over the radial wing, which extends to form the dis=


k shown in the document. The cabin is 5.7 feet tall. There is an 8 foot h=


emisphere atop a 5.7 foot tall cylinder and another 8 foot hemisphere below=


- to create a tank containing the liquid hydrogen. Smaller more efficient=


turbo jets in larger number provide power around the rim of the radial win=


g.




Two of the ten seats are equipped with flight controls and a heads up displ=


ay on the canopy.




--

No need to believe in either side, or any side. There is no cause.

There's only yourself. The belief is in your own precision. - Joubert

these opinions are mine, all mine; HP might not want them anyway...

feel free to post, OR email to rick.jones2 in hp.com but NOT BOTH...


No, its the secret program that was denied by the USAF until declassified last year.

The 'failed' VZ-9AV was an 18 foot disk that had a central lifting fan powered by 3 Continental J-69-T9 turbojets at 660 lbs each.

This vehicle was unstable out of ground-effect and lacked positive control in wind tunnel tests done by the USAF at Wright Pat and NASA Moffett Field.

There apparently was a 33 ft diameter version with 120 peripheral nozzle assemblies located around the rim of the disk for attitude control propelled by four Armstrong Siddeley Viper 8 jet engines.

There was also a 40 ft diameter version with similar arrangement of peripheral nozzles optimized for thrust at high speeds with eight Viper 8 jets.

This was part of the System 606A program for the PV70H aircraft, according to the report.

These configurations were tested in both supersonic and subsonic wind tunnels and found to be highly controllable and capable of smooth transitions from vertical to horizontal flight. They also tested ramjet operation and a radial afterburner capability to improve thrust.

http://www.secretsdeclassified.af.mi...121113-019.pdf

I'm looking at making a small drone version of this aircraft powered by a collection of small jet engines -

http://www.jetjoe.com/main.php

Each drone engine is 2 inches in diameter 5.2 inches long, weighs half a pound and produces 3.5 pounds of thrust. Four of these operate a drone that's 52 inches in diameter and weighs 37.6 pounds at take off.

The original was powered by four, six or eight Viper 8 engines that were 25 inches long and 64 inches in length, weighs 549 lbs and produces 2700 lbs thrust, within airframes that ranged from 33 ft to 40 ft in diameter and massed up to 29,000 lbs at take off.

* * *

An electric version is possible - which has been worked out in great detail....

http://acl.mit.edu/papers/Cutler_Masters12.pdf

The trajectory control described starting on page 54 using four rotors with four electric motors (two CW, two CCW turning) of this thesis, is easily adapted to a four jet system with rim mounted nozzles.

Another possibility is to use the exhaust of the tiny JetJoe engine to propel a tip jet rotor, to multiply thrust, creating a thermal engine driven quad rotor. The rotor is supported by a film of compressed air or magnetic bearing, or a combination of both.

With an Arcam 3D printer I'm thinking of making tiny lightweight tip jet rotor blades that travel along a magnetic or compressed air track at the rim of a disk shaped aircraft. Here the tiny jet engines blow their exhaust into hollow rotor blades affixed to the track which is exhausted at the tip of the blades. Two trains of blades, one CW the other CCW rotate around the rim of the aircraft. Shutters direct the flow of air produced by rotors to direct it in flight.

http://www.tecaeromex.com/ingles/RH-i.htm

Another possibility is to use hydrogen peroxide in a tip-rotor arrangement to produce a disk type lifting craft.

Here's a tip rotor with an electric blower, made of plastic parts 3D printed

https://www.youtube.com/watch?v=CSZqFoOiyF4

Here's a 1955 technical summary of a similar concept

https://www.youtube.com/watch?v=RRrYdIqrnWY

and of course, little henry

https://www.youtube.com/watch?v=qs7WlNrHRmA

A tip jet driven quad rotor with the exhaust of the tiny engine blown through the center of each of four rotors, would provide substantial capacity. If one thinks of a disk shaped airframe made of foamed composite - as a duct for a ducted fan, with nozzles...

this is what I'm thinking. Instead of a single rotor at the center of a disk, I'm thinking of four rotors, driven tip jet fashion, encased in a duct at each 'corner' or quadrant of the disk airframe. An intake near the center of the airframe feeds the encased rotors, and a fan exhaust near the rim of the airframe bleeds the encased rotors.

This is what I'm thinking.

The 3.5 lb x 4 = 14 lb total thrust with a 12 to 1 multiplier with the tip jet rotor (as in Little Henry's case 600 lbs lift with two 10 lb motors that produce 25 lbs each) - achieving the same leverage with a quadrotor system (encased inside a disk airframe) produces 168 lbs lift - with the tiny JoeJet

Here's a five stage turbojet engine from the Bladon Brothers.

http://www.youtube.com/watch?v=PFnZJeHc-KY

This one is 4.5 inches in diameter 12 inches long and produces 100 lbs of thrust - so four of these produce 400 lbs of static thrust, and when multiplied by tip jet driven rotors in a quad rotor type arrangement - produce 4,800 lbs of total lift. So, a four passenger 2,400 lb vehicle, is possible with two gee take off thrust - straight up - and speeds up to 580 mph should be possible. With VTOL capabilities, this is awesome!

Four passenger centrally mounted seats with iPad touchscreen controls that use GPS and Google Earth to guide the aircraft point to point autonomously. A typical helicopter control system is also provided for manual control.

https://www.youtube.com/watch?v=NevgqMqWf5Y

 




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