Halfway to Anywhere

Lifting your rocket from Terra's surface into circular orbit takes an unreasonably large amount of delta V. As a matter of fact, if your missions use Hohmann trajectories, the lift-off portion will take about the same delta V as does the Hohmann from Terra to the destination planet. As Heinlein put it:

How much delta V does it take to go from Low Terra Orbit to Mars orbit? About 5.6 kilometers per second.

How much delta V does it take to go from the surface of Terra to Low Terra Orbit? 7.6 Freaking kilometers per second, that's what! In other words it takes more delta V to travel the pathetic 360 kilometers up to Low Terra Orbit as it does to travel the 228,000,000 kilometers to Mars!

From Low Terra Orbit, where can you travel to with 7.6 km/s? Oh, only to the Planet Saturn, 1,433,000,000 kilometers towards the edge of the entire solar system.

Rob Davidoff suggests that in a rocketpunk future, people will no longer use the expression "worth its weight in gold." Instead they will say "worth its weight in upmass", referring to the outrageous cost of shipping any payload from Terra's surface into Low Terra Orbit.

But the delta V cost breakdown is interesting. Getting into orbit takes just a little bit of delta V. It is making sure you stay in space that takes a freaking lot of delta V.

A little sounding rocket can easily rise from 50 to 1,500 kilometers above Terra's surface, where outer space starts about 150 kilometers up. Then the propellant runs out, and the poor little rocket finds itself unsupported hundreds of kilometers up. So it plummets to its doom.

How do you support the sad little rocket? If it uses propellant it will eventually run out, sooner more than later. You can't build rocket legs that are hundreds of kilometers long. You can't use a helicopter blade because there is no air.

But what you can do is put the rocket in an "orbit". An orbit is a clever way to constantly fall but never hit the ground. The trouble is that entering an orbit takes a freaking lot of delta V, about 8 kilometers per second around Terra.

Of course, once you have torchships you can stop all this child's play with wimpy Hohmann transfers and start doing some big muscular Brachistochrone trajectories. Brachistochrones typically require delta Vs that are hundreds of times more than the equivalent Hohmann. So any ship that can handle a Brachistochrone is not going to even notice the delta V cost for lift-off.

But even with torchships, the real bottle-neck restricting developing space resources remains the cost to boost payloads into Earth orbit.

For some cold hard reality read When Rocket Science Meets The Dismal Science.

There are other ways besides rocket boosters and space shuttles to get payloads into orbit. These might take the form of rockets climbing rails set up the side of a mountain, a laser thermal launching facility (in THE MILLENNIAL PROJECT, Marshall Savage calls this a "Bifrost Bridge", that is, a bridge to space composed of colored light), launching loops, space fountains or the base of a Space Elevator.

Launch Sites

The two main types of orbit that launch vehicles boost payload into are equatorial orbits and polar orbits.

Polar orbits pass over both the north and south poles, with an inclination close to 90 degrees with respect to the equator. But the important point is a satellite in polar orbit will eventually pass over every single spot on Terra. Heinlein calls these "ball of yarn" orbits, since the path of the satellite resembles wrapping a strand of yarn around a yarn ball. This is why such orbits are used for Earth-mapping, Earth-observation, some weather satellites … and reconnaissance satellites aka "spy" satellites.

For communication satellites, space stations, resupply missions, space exploration, and pretty much everything else, you launch into equatorial orbits.

LAUNCH CORRIDOR

When deciding where to put a launch site, you have to plan around the Launch Corridor. This is the path the rocket will take when launching which will [1] allow the rocket to reach the desired orbit and [2] if the rocket engines fail, the rocket (or the remaining flaming rocket debris) will only fall on uninhabited areas as long as it stays inside the launch corridor. The standard practice is to arrange launch corridors to be over the ocean. Failing that, you need land areas where a rain of flaming rocket bits is unlikely to result in lawsuits or negative publicity. And of course ones that do not violate another nation's sovereign airspace.

During launch, the range safety officer will be watching the rocket like a hawk. If the rocket shows signs of failing to reach orbit, the officer will make a note to dispatch a rescue/cleanup team. If the rocket shows signs of leaving the launch corridor, the officer will hit the panic buttons. Unmanned rockets will shutdown their engines and vent their propellant. Manned rockets will have the on-board pilot take action, but if they are ineffective the range safety officer might have to shoot the rocket out of the sky.

Obviously polar launch corridors have to be along the north-south axis.

The United States uses Vandenberg AFB Space Launch Complex 6 (SLC-6 aka "Slick Six") to launch into polar orbits. Rockets launch due south so the launch corridor is thousands of miles of uninhabited Pacific ocean. The alternative is to launch due north, but that puts the launch corridor right across California, the long way.

EQUATOR BOOST

Equitorial launches have a second consideration besides the launch corridor.

When you are dealing with feeble launch vehicles using chemical propulsion you need to use every trick you can find. They have grotesque mass ratios which really cut into the payload mass. The most important trick is one to reduce the delta V the rocket needs to achieve orbit.

Since Terra is spinning on its axis, when the rocket is sitting on the ground it is actually already moving. At least it is moving relative to the desired orbit, which is the important thing. If you are standing in New York City; you, the ground, the skyscrapers, the taxi cabs, and everything else is moving at 356 meters per second. The only reason everything seems stationary is because everything is moving together. Now remember that on Terra everything is moving due east because that is the direction Terra is spinning on its axis.

The technical term is the tangential velocity of Terra's surface. It is equal to

tangentialVelocity = ((2 * π * planetRadius) / siderialRotation) * cos(latitude)

where

tangentialVelocity = tangential velocity at planet surface (m/s) (Terra = 465 m/s)
π = pi = 3.14159…
siderialRotation = siderial rotation period (seconds) (Terra = 86,164 seconds, which is actually 23 hours, 56 minutes, 4 seconds)
cos(x) = cosine of x (do not make the mistake of giving your spreadsheet or calculator "x" in degrees when it is expecting radians or something)

I gave you the entire equation in case you wanted to do the calculations for an extraterrestrial planet. If you are just trying to place launch sites on Terra, the equation is:

tangentialVelocity = 465 * cos(latitude)

The point is that the delta V the launch vehicle needs to achieve orbit is reduced by the tangential velocity of the launch site. Bottom line is the closer you can put the launch site to the equator, the better.

For Terra, the pure orbit delta V is about 9,700 m/s (would be 7,800 m/s except for air-drag, gravity-drag, and vertical acceleration). But when launching from New York the delta V is only 9,700 - 356 = 9,344 m/s. And launching from the equator it is 9,700 - 465 = 9,235 m/s. That kind of delta V reduction can buy you lots of extra payload.

Keep in mind that since Terra is spinning due east, the rocket has to launch in an easterly direction in order to take advantage of the bonus. By the same token, if the stupid rocket launches west, the bonus turns into a liability. Launching westward on Terra's equator means the rocket needs an additional 465 m/s to reach orbit.

The important point is that on Terra the equatorial launch corridor is going to point due east.

The better science fiction novels put Terran equatorial launch sites as close to the equator as possible, and where an eastward launch corridor passes over lots of ocean (i.e., on the east coast, near the equator).

• The North Maluku province of Indonesia has parts right on the equator. It has pretty much the entire Pacific Ocean to use as a launch corridor, except only scattered tiny islands in the launch corridor. Possible launch site.

• There is a part of the coast of Brazil that is right on the equator. It has pretty much the entire Atlantic Ocean to use as a launch corridor. Possible launch site.

• Parts of the Galápagos Islands are right on the equator. Unfortunately it only has 906 km of Pacific Ocean launch corridor before flaming rocket bits start raining down on Ecuador. Possible launch site.

• In ARTEMIS by Andy Weir the launch site is in Kenya, with parts right on the equator. It has pretty much the entire Indian Ocean to use as a launch corridor. However, the part closest to the equator that does not include Somalia in the launch corridor is located at 1.7° S latitude.

• In ISLANDS IN SPACE by Arthur C. Clarke the launch site is at New Guinea, with point closest to equator at about 2.6° S latitude. It has pretty much the entire Pacific Ocean to use as a launch corridor, except for the Solomon Islands.

• The real world Guiana Space Centre in French Guiana is at about 5° N latitude. It has pretty much the entire Atlantic Ocean to use as a launch corridor.

• Palmyra Atoll is at about 5° N latitude. It has pretty much the entire Pacific Ocean to use as a launch corridor. And it is a US unorganized incorporated territory. Drawbacks include it is pretty much on the opposite side of Terra from the continental US so that logistics is a nightmare, and the highest point is (currently) only 10 meters above sea level.

• The US Virgin Islands are at about 17.7° N latitude. It has pretty much the entire Atlantic Ocean to use as a launch corridor. Possible launch site.

• In High Justice by Jerry Pournelle the launch site is at Cabo San Lucas, Mexico. It is at an unhelpful 22.8° N latitude. And it only has 390 kilometers of launch corridor.

• The real world Kennedy Space Center Launch Complex 39 is at an ugly 28.5° N latitude. But the United States does not get that much closer to the equator. It has pretty much the entire Atlantic Ocean to use as a launch corridor.

• The real world Baikonur Cosmodrome is at an almost utterly worthless 45.6° N latitude. What's worse it it has to launch at a 51.6° inclination, since China takes a very dim view of being in the launch corridor. Sadly Baikonur is probably located at the best out of Russia's poor selection of launch sites.

THE CURIOUS CASE OF THE ISS INCLINATION

Now one would have expected that the International Space Station (ISS) would be in a 28.5° inclined orbit, which is the orbit you get when launching due East of Kennedy Space Center (latitude 28.5° N).

But it isn't, the ISS is instead in a 51.6° inclined orbit. Why? So that Russian cargo rockets from Baikonur Cosmodrome can reach it. Launching into a different inclination than the space port's latitude costs rocket propellant and reduces payload.

Changing the ISS planned inclination to 51.6° was in retrospect a very good decision. When NASA stupidly cancelled the Space Shuttle program before the replacement vehicle was online, they assured everybody that the replacement would be flying by 2014 at the latest. This would make a small three-year gap in NASA's ISS transport ability. Unfortunately and predictably when 2014 arrived NASA has not even started work on deciding which of the many proposals will be used, much less bending metal and cranking out functional rockets. This leaves NASA at the mercy of the Russians for access to the ISS, but without the Russians there would be no access at all and the station would have long ago burnt up in reentry like Skylab. But I digress.

Clever readers will say but wait! Baikonur Cosmodrome is at latitude 45.6°, should not that be the inclination?. In a perfect world, yes, but there is a problem. When a spacecraft is launched from Kennedy Space Center the lower stages fall into the Atlantic Ocean. And if something goes really wrong, the entire spacecraft can abort and ditch into the ocean as well. If Baikonur Cosmodrome did the same thing, large spent lower stage boosters and/or huge flaming aborting Russian spacecraft would crash into Mainland China, and the political situation would rapidly deteriorate. To avoid that unhappy state of affairs, Russian spacecraft launched from Baikonur go at a 51.6° inclination, so falling rocket bits will miss China.

The Russians already have an annoying problem with the lack of warm-water ports for seagoing vessels. They really dislike having much the same problem with respect to space launches. Therefore they are in negotiations for launch privileges at the ESA's Guiana Space Centre, which is optimally located quite near the Equator and to the West of the Atlantic Ocean.

Resuable Boosters

18-wheelers, trains, cargo aircraft, and cargo ships would all be several orders of magnitude more expensive if the vehicles could only be used once then thrown away. You cannot amortize the cost much on a single trip. But that used to be industry standard in the space booster business. The rocket hauls the satellite into orbit, lets it go, then the rocket falls back to Terra and burns up in reentry.

NASA tried to create a reusable system with its Space Shuttle, but it miserably failed to meet its promised cost and utility goals, as well as design, cost, management, and safety issues. Blasted thing was actually more expensive than a throwaway rocket: \$18,000 per kilogram delivered to LEO, while the expendable Proton could do it for \$5,000 per kilogram.

So all the rocket companies laughed and laughed when Elon Musk founded SpaceX and announced he was developing a partially reusable booster rocket. The companies sat on their expansive derrières and patiently watched as SpaceX struggled through the lengthy development process, secure in the knowledge that Elon was going to fail.

Well, they ain't laughing now. Indeed the rocket companies are panicking. SpaceX reusable launch system is a going concern, the cost per kilogram of payload is way below any technology the other companies have, and now they are about 15 years worth of development time behind SpaceX. A SpaceX Falcon-9 can be refurbished for reuse for about 10% of the cost of building an entire new rocket. There is a penalty of a 30% reduction in payload mass when a Falcon-9 is flown in re-use mode. But according to SpaceX, they break even with the second flight of a Falcon-9, and save money from the third flight on. The customers even prefer to have their payloads boosted on re-used rockets, because the rockets have been flight tested so to speak.

Along the same lines as reusable boosters are the ground-based facilities that can fling things into orbit as long as the electricity keeps running. Economically they are equivalent to reusable boosters. They include such things as Lofstrom loops and laser launchers.

Boosters: Present and Proposed

For comparison purposes, here are the masses of a few sample payloads. This is to give you a mental image of the capabilities of the following booster systems. It will also be useful if the cargo space could accommodate standard cargo cannister sizes.

GPS satellite0.8 metric ton
Communication satellite1 metric ton
Weather satellite1 metric ton
Hubble Space Telescope11 metric tons
KH-11 spy satellite13 metric tons
TransHab habitat module34 metric tons
Skylab77 metric tons
Space Station Mir124 metric tons
International Space Station287 metric tons
1 gW Solar Power Satellite1,900 metric tons
Lunar Mass Driver2,750 metric tons
Lunar Base (150 crew)17,050 metric tons
10 gW Solar Power Satellite19,000 metric tons
5 gW Solar Power Satellite (Rockwell International estimate)37,000 metric tons
2001 Space Odyssey Station V145,000 metric tons
1 tW Solar Power Satellite1,900,000 metric tons
1.5 tW Solar Power Satellite2,800,000 metric tons
L5 Colony10,000,000 metric tons

Existing Heavy Lift Launch Vehicles

"Heavy Lift" is defined as 12 metric tons or more into LEO.

Heavy Lift Launch Vehicle (HLLV)Payload mass delivered to LEOCost per payload kilogram
Zenit 2 (Ukraine)13.7 metric tons\$3,093/kg
Zenit 3SL (Sea Launch)15.9 metric tons\$5,354/kg
Japan H2B16.5 metric tons?/kg
Ariane 5G (ESA)18 metric tons\$9,167/kg
Atlas V 55118.51 metric tons?/kg
Ariane 5 ES (ESA)20 metric tons?/kg
Titan IV-B21.69 metric tons?/kg
Falcon 9 v1.2 (SpaceX)22.8 metric tons\$2,720/kg
Delta IV Heavy (ULA)28.79 metric tons?/kg
Proton-M (Russia)23 metric tons\$4,302/kg
Space Shuttle (NASA)24 metric tons\$10,416/kg
Falcon Heavy (SpaceX)63.8 metric tons\$2,968/kg
Saturn V (NASA)118 metric tons??

Proposed STO Solutions

Black Horse0.45 to 2.3 metric tons (est)\$227/kg (est)
Black Colt0.45 metric tons??
Rocketplane XS1.5 to 3.0 metric tons??
The Rocket Company DH-12.2 metric tons\$440/kg
SASSTO2.8 metric tons\$11/kg (1968 dollars)
Douglas ASTRO16.9 metric tons\$88.38/kg (1964 dollars)
Collier's space ferry25 metric tons??
ASPEN Nuclear SSTO36 metric tons??
SERV/MURP53 metric tons\$95/kg (1971 dollars)
Star-Raker91 metric tons\$22/kg to \$33/kg
Nuclear DC-X100 metric tons\$150/kg
Rombus450 metric tons\$2.30 to \$5.40/kg (1964 dollars)
Sea Dragon550 metric tons\$59/kg to \$600/kg
GCNR Liberty Ship1,000 metric tons??
Uprated GCNR Nexus1,500 metric tons??
Space Elevator x12,000 metric tons/year\$3,000/kg
Planetary Orion3,000 metric tons??
Laser Launch (HX)3,000 metric tons/year\$550/kg
Space Elevator x24,000 metric tons/year\$1,900/kg
Super Nexus4,600 metric tons??
Space Elevator x36,000 metric tons/year\$1,600/kg
Aldebaran27,000 metric tons??
Lofstrom loop small40,000 metric tons/year\$300/kg
Rocket Sled (StarTram)150,000 metric tons/year\$43/kg
Bifrost Bridge175,200 metric tons/year\$20/kg
Verne Gun280,000 metric tons??
Lofstrom loop large6,000,000 metric tons/year\$3/kg
Super Orion8,000,000 metric tons??

The Rocket Company DH-1

2.2 metric tons\$440/kg

The DH-1 is a fictional two stage to orbit re-useable rocket described in the book The Rocket Company (ISBN 1-56347-696-7). There are some sample chapters here. I recommend this book.

While the design is fictional, it would actually work. The authors have patented it. The small payload means the rocket is intended more for "space access" instead of heavy lift to orbit. The business model for the developers was more to sell the rockets (at an attractive price of \$250 million) rather than selling cargo boost services.

There are DH-1 plug-ins for the spacecraft simulation Orbiter.

SASSTO

2.8 metric tons\$11/kg (1968 dollars)
SASSTO
Gross Mass97,976 kg
Empty Mass6,668 kg
Thrust (vac)1,558,100 N
Specific
Impulse
464 s
Diameter6.6 m
Length18.8 m
EngineChemical
Plug-nozzle
LOX/LH2
Num Engines36

The Saturn Application Single-Stage-to-Orbit (SASSTO) is from Frontiers of Space by Philip Bono and Kenneth Gartland (1969)

In 1966 when winged space shuttle designs were being studied, the Douglas Aircraft Company was doing a cost-benefit analysis. They were comparing reusable space shuttle costs to throwaway two-stage ballistic boosters. Somewhere along the line they took a look at whether it was possible to make a reusable single stage ballistic booster. The SASSTO was the result. The payload was not much, but it was enough for a Gemini space capsule. A Gemini would transform the SASSTO into a space taxi or even a space fighter, capable of satellite inspection missions. Without the Gemini it could deliver supplies and propellant to space stations and spacecraft in LEO.

Bono pointed out how inoperative satellites could become space hazards (although the concept of the Kessler Syndrome would not be created until 1978). A SASSTO could deal with such satellites in LEO (Bono called this Saturn Application Retrieval and Rescue Apparatus or SARRA). Even better, such satellites could be grabbed and brought back to Terra for refurbishment and re-launch. This would be much cheaper than building an entire new satellite from scratch, which would interest satellite corporations. Only satellites in LEO though, communication satellites in geostationary orbit would be out of reach.

The interesting part was on the base. Conventional spacecraft trying to do an aerobraking landing need a large convex heat shield on the base (for example the Apollo command module.). Unfortunately a reusable spacecraft has a large concave exhaust nozzle on the bottom, exactly the opposite of what you want. Tinsley's artist conception for the "Mars Snooper" had petals that would close over the exhaust nozzle sticking out of the heat shield, but that was impractical.

Douglas' solution was to use an aerospike engine with the spike truncated (which they confusingly call a "plug nozzle", contrary to modern terminology). The truncated part became the heat shield, the untruncated part around the edge was the aerospike engine.

Douglas ASTRO

delivered to LEO
Cost per
16.9 metric tons\$88.38/kg (1964 dollars)
\$1,329.42/kg (2020 dollars)
Douglas ASTRO
Orbiter
Orbit
Delivered to
555 km
28.5°
Engines2 × RL-10
(gimballed)
1 × J-2
FuelLOX/LH2
Inert Mass14,000 kg
Dry Mass30,851 kg
Fuel Mass74,842 kg
Wet Inert Mass89,290 kg
Wet Mass106,141 kg
Length20.7 m
Span13.4m
Crewx2
Booster
Engines1 × M-1
2 × J-2
FuelLOX/LH2
Dry Mass32,558 kg
Fuel Mass269,434 kg
Wet Mass302,183 kg
Length29 m
Span18.6 m
Combined
Wet Mass407,870 kg
Length49 m
Crewx1

This was Douglas Aircraft Company's contribution to NASA's 1963 Reusable Ten Ton Orbital Carrier Vehicle study.

ASTRO is an acronym for Advance Spacecraft Truck/Trainer/Transport Reusable Orbiter

This was a Vertical-Takeoff/Horizontal-Landing (VTHL) Two-Stage-To-Orbit (TSTO) spacecraft for transporting space station crews and cargo. One of the study requirements was that off-the-shelf technology was to be used: meaning M-1, J-2, and RL-10 rocket engines (though the M-1 was never actually built because the Nova was never actually built). The RL-10, J-2, and M-1 engines had thrust levels of 67,000 N, 890,000 N, and 6,700,000 N, respectively (as mentioned by Douglas, some of the engines had higher thrust when actually built).

In its surface-to-orbit cargo transport role it was a two-stage vehicle. Conventional rocket are stacked nose-to-tail (on top of each other like a totem pole) for staging, winged STO rockets are usually stacked ventral-to-dorsal (piggyback fashion). ASTRO was unusual in that it was stacked.

Douglas figured it would be easier to make a recoverable and reusable design if they aimed for a smaller spacecraft with lower payload but capable of frequent flight schedules. Larger spacecraft with higher payloads with infrequent flight schedules are more difficult to design to be reusable, actually more difficult to design regardless of re-usability.

Both the orbiter and the booster were lifting bodies (since putting wings on a craft that can aerobrake from orbit and perform hypersonic flight is a bit of a challenge). Wings/lifting body allowed the spacecraft to be reusable, i.e., landing on a landing field instead of ditching in the ocean. Both landed on skids with steerable nose gear.

Orbiter had a pilot and co-pilot. Booster just had a pilot. The crew compartment of each are abortable, jettisoning from the spacecraft in case of a catastrophic malfunction. For missions with only pilots on-board, only the cockpit is jettisoned and the cargo compartment is not (and may not even be pressurized). For missions with more crew, some will ride in the cargo compartment. In that case both the cockpit and the cargo compartment is jettisoned and both are pressurized.

When boosting cargo both the booster and orbiter were VTHL. The orbiter alone was capable of horizontal take off and landing (HTHL), allowing it to be used as a sub-orbital trainer to educate space pilots. This solved the problem of how do you train ASTRO pilots before it is operational. They would quickly build the orbiter section and use it to train the pilots while the booster section was developed. In addition, the orbiter alone was a useful suborbital craft with a range of 8,000 kilometers and a cargo capacity of ten people or 2,100 kilograms of cargo.

In Phase II the booster prototype was made by taking the orbiter, adding a second J-2 engine, redesigning the cabin so it only holds one pilot with zero cargo, and altering the nose so that the orbiter tail bumper can perch on it. Douglas figured this early version could put one tonne and two crew into orbit.

In Phase III (final phase) the booster was beefed up, making it larger than the orbiter, and adding a freaking M-1 between the two J-2s. As previously mentioned, the M-1 was a monster rocket intended for NASA's Nova. It made the F-1 engines on the Saturn-V's first stage look like a child's bottle rocket. The Phase III could put nine tonnes and two crew into orbit. The cargo compartment has a volume of 14.9 cubic meters. If that is not enough volume (the cargo bulked-out before it grossed-out), some cargo can be carried on the skin in detachable pods. Meaning if your cargo has a mass of nine tonnes but a volume bigger than 14.9 m3 you will be forced to put some of the cargo metaphorically on the roof rack. In space slang, your cargo has "bulked-out" before it "grossed-out".

The maximum 9 tonnes of payload is only if the ASTRO is launched from the equator with a launch azimuth of 90° (due east). The maximum payload decreases as the launch site is displaced from these conditions.

But to hedge their bets, Douglas also wanted the ASTRO to be useful for any mission the military had for a space plane: defense bombardment, reconnaissance, satellite inspection, satellite interception (SAINT II), satellite logistic support (refuelling station-keeping thrusters, swapping out spy satellite camera film, etc.), recoverable space laboratory, astronaut training, maintenance, rescue, and supply.

For reconnaissance missions the payload is cameras and film. For satellite surveillance missions the payload is inspection, kill, and capture equipment. For repair, cargo transfer, and general service missions the payload is maintenance equipment including modular spares, tether reels, space pods, and the cargo. With some missions maintenance personnel are part of the cargo.

The ASTRO stack took off from a mobile launcher-erector. This was to eliminate the need for a large Cape Canaveral type launching gantry, and the need to use huge cranes to piggyback the orbiter on top of a cargo aircraft to fly the damn thing back to the launch site. NASA actually did use the piggyback method with the Space Shuttle, transporting the Shuttle from its landing site to the launch site on the back of a modified 747.

The mobile launcher erects the stack so that it points skyward (or you will not go to space today). The booster blasts off and carries the orbiter and itself to an altitude of 82 kilometers. There stage separation occurs, and the orbiter continues upward to a 555 km orbit. The poor booster pilot has to glide the corpulent booster to a dead-stick landing 830 freaking kilometers from the launch site. Get it right the first time because you can't turn around to make a second attempt.

The proposed ASTRO system would have a fleet of 12 boosters and 24 orbiters with a turnaround time between missions of less than 18 days. This would allow about 240 flight per year. The planned service life was 100 flights for each orbiter and 200 flights for each booster. The engines were rated for 50 firing before needing a major overhaul. The orbiter and booster airframes were rated for up to 300 flights.

Alas for the ASTRO, its fate was linked to the X-20 Dyna Soar. When the X-20 project was killed, ASTRO went with it.

Collier's space ferry

25 metric tons??
COLLIER'S SPACE FERRY
Study date1952
EnginesChemical
FuelNitric Acid / Hydrazine
Thrust110,300,000 N
Wet Mass6,400,000 kg
Height97 m
Diameter20 m
Apogee1,730 km
Orbit Inclination23.5°

ASPEN Nuclear SSTO

14 to 36 metric tons??

This is from ASPEN An Aerospace Plane With Nuclear Engines by R. W. Bussard (1961)

SERV/MURP

53 metric tons\$95/kg (1971 dollars)
SERV
Wet Mass2,040,816 kg
Height20.30 m
Diameter27.40 m
Apogee185 km
18 m high
Service Lifex100 flights
over 10 years
LIFT-OFF ENGINES
TypeAerospike
Num Enginesx12
Total Thrust25,795,300 N
Specific Impulse347 sec
FuelLH2
OxidizerLOX
LANDING ENGINES
TypeTurbojet
Num Enginesx28
Total Thrust111,796 N
FuelJP-4
OxidizerAir

This is from PROJECT SERV A Space Shuttle Feasibility Study, Project SERV Final Review, Astronautix, and SERV/MURP: Chrysler’s Space Truck.

SERV stands for Single-stage Earth-orbital Reusable Vehicle. This was a Chrysler study produced when NASA asked for proposals for a Space Shuttle. However, NASA made it clear that it was mostly interested in winged Shuttles. Chrysler was the only one who bothered with a non-winged proposal, and NASA returned the favor by not giving the SERV any serious consideration at all.

The SERV was shaped like an Apollo Command module magnified to a seven times larger scale. Just like the SSASTO it surrounded the aerobraking heat shield on its butt with an annular aerospike engine. Unlike the SSASTO the SERV's heat shield had hatches for the landing gear and turbofan lift engines. The aerospike engines had hatch covers, but they did not penetrate the heat shield.

The ballistic aerospike flight could aim for a landing site within a 15 kilometer diameter circle, but that was not good enough for NASA's specifications. That's whe the turbofan lift engines were added to the design. This allowed it to get within 75 meters of its aim point. The ability to use the atmosphere for oxidizer made the difference.

For uncrewed missions the SERV would carry a cylindrical payload module with a tiny nose cone on the top, and deliver it to orbit. For crewed missions the SERV would also carry on top a Manned Upper-stage Reusable Payload (MURP) spaceplane, capable of an aerobraking re-entry.

As interest in the SERV wained, Chrysler desperately invented new modules for it to carry. There was a tiny modifed Apollo Command module so the cargo version could also carry crew, a long aerodynamic spike that would lower the drag and increase the payload, and a plan to use the SERV as a sub-orbital airliner capable of carrying passengers from Heathrow to Sydney in three hours instead of twenty-two. Oh, and a solid-core nuclear upper stage suitable for a Mars mission or transporting outrageous payloads to Luna.

The MURP was based on the Northrop HL-10. It had a spray-on silicon ablative skin which was peeled off and refreshed after every mission. There were two MURP designs: the the D-10 and the D-34. Since the cylindrical cargo pod is a more efficient use of space, the D-10 has a lower mass than the D-34.

MURPs
D-10D-34
Internal cargo5 m385 m3
Cargo Pod80 m3n/a
Mass11,640 kg16,150 kg
Crew2
Passengers10

Star-Raker

91 metric tons\$22/kg to \$33/kg

Star-Raker is from a 1970's Rockwell International study, one of the many proposals on how to boost into orbit the outrageous payload requirements of a multi-kilometer solar power satellite (SPS). They were figuring on about 37,000 meric tons per SPS, and they wanted a constellation of 60 of them. For the project they estimated boosting 74,000 metric tons per year (2 SPS/year).

Star-Raker was a single-stage-to orbit airbreathing horizontal takeoff and landing craft (HTO-SSTO). The gross mass would be about 2,268 metric tons, the payload mass was about 91 metric tons, and it was claimed it would have a boost turnaround time of about a day and be really really cheap. Keeping in mind that at the time Rockwell was also claiming that the Space Shuttle would have a two-week turnaround and be really really cheap, which turned out to be somewhere between irrationally optimistic and an assurance from a used-car dealer. It was to be capable of delivering its payload into a 550 kilometer equatorial orbit.

To manage the proposed schedule of boosting the payload for two SPS per year would need about 815 flight per year, or 2.2 flights per day. This assumes a fleet of more than one Star-Raker.

Horizontal takeoff and landing, and single-stage were design choices due to the need for rapid turnaround. Having to fish stages out of the ocean, haul them to the launch site, refurbish, and re-stack them would make it impossible to have a single-day turnaround. To save mass the take-off wheels would be jettisoned at the end of the runway and recovered. For landing lighter internal landing gear is used, since by then the craft will be lighter by many metric tons of absent payload and burnt fuel.

It has a "wet-wing" design, that is, the wing is the fuel tank. The body of the craft is reserved for the payload. It was to be capable of taking off and landing on a 2,500 meter runway.

It is an air-breather using atmosphere for oxidizer up to the point where the air is too thin at thirty kilometers altitude (ten supersonic-turbofan/airturbo-exchanger/ramjet engines with a combined thrust of 6.2×107 newtons thrust). For the last portion of the boost it switches over to rocket engines (three rockets with 1.4×107 newtons thrust each). The jet engine air inlets will be closed by retractable ramps while the craft is under rocket flight and during ballistic re-entry. From zero to 1,800 m/s it will be using airbreathing propulsion, from 1,800 to 2,200 m/s it will use both airbreathing and rocket propulsion, and from 2,200 m/s to orbit it will use only rocket propulsion.

It would also be capable of making trips as a conventional cargo aircraft. For instance, from the launch site to a site where the payload had been assembled, and back to the launch site. It saves on having to ship the payload to the launch site, but I question the wisdom of risking an expensive HTO-SSTO craft when a less expensive and more expendable cargo plane would suffice. The entire nose (including crew compartment) swings open to expose the cargo hatch (which must be scary for the crew when the playload is released into orbit). This allows it to be loaded from a conventional cargo platform. Cargo floor is designed similar to a C5-A military transport aircraft.

There was another design tailored for delivering payload into polar orbits, which would reduce the payload mass. Polar orbits are expensive in terms of delta V, but are necessary for Department of Defense spy satellites.

Report can be found here.

Pan Am Space Clipper

fictional surface-to-orbit reusable shuttle featured in the movie 2001 A Space Odyssey (1968).

This is bitterly ironic, since Pan American World Airways went bankrupt in 1991, before many of our younger readers were born. For that matter, the movie was suppose to take place in the far-flung future year of 2001, Clavius moon base and all.

And what's with the name "ORION III SPACEPLANE"??!? There are too many freaking spacecraft with the name Orion.

Actually as it turns out, the Pan Am clipper was called "Orion" because originally it was going to be an honest-to-Pournelle surface launched nuclear pulse vehicle. I read that while Arthur C. Clarke was working on the movie, he was contacted by some scientists who were still angry that Project Orion had be canceled in 1964 (they were only teeny-tiny A-Bombs, honest!). They asked Clarke if an Orion drive spacecraft could be used in the movie, to promote the concept.

So the Orion III was actually going to be a real Orion. Sadly Stanley Kubrick thought that sending Dr. Floyd into orbit on a series of nuclear detonations was hard to take seriously, so the Space Clipper was downgraded to a conventional liquid hydrogen - LOX rocket. The Discovery was considered for an Orion Drive as well, but that too was vetoed.

Nuclear DC-X

100 metric tons\$150/kg

This is from a report called AFRL-PR-ED-TR-2004-0024 Advanced Propulsion Study (2004). It is a single stage to orbit vehicle using a LANTR for propulsion.

Rombus

450 metric tons\$2.30 to \$5.40/kg
(1964 dollars)
Rombus
Gross mass6,363,000 kg
Height29 m
Diameter24 m
Thrust79,769,000 N
EngineChemical
Plug-nozzle
LOX/LH2
Specific Impulse455 s
Num nozzles×36

The Reusable Orbital Module-Booster & Utility Shuttle (ROMBUS) is from Frontiers of Space by Philip Bono and Kenneth Gartland (1969). This is a reusable plug-nozzle powered booster. It used an aerospike engine with the spike truncated and turned into an aerobraking heat shield.

Bono also created a passenger carrying variant named Pegasus, and a military troop carrier called Ithacus. When the concept lost support at NASA, Philip Bono designed a more modest concept, adding an aerospike engine to a Saturn V to create the SASSTO concept.

The vehicle is staged in the sense that it jettisons external hydrogen fuel tanks during the ascent phase. The tanks have parachutes to increase the chance they can be reused.

After delivering its payload, the vehicle would typically spend 24 hours in orbit before the ground track passes close enough to the landing site. It lands using parachutes and rockets, with the final touchdown burn delivered by four engines running at 25% thrust for twelve seconds. The vehicle turnaround time would be about 76 days.

1. Payload 0.8 to 1.0 million pounds to orbit
2. Roll-control nozzle pairs
3. Vent lines for liquid hydrogen tanks (8)
4. Propellant utilization probes (8)
5. Booster centre body
6. Fuel tank support fittings (16)
7. guidance and electronic package
8. Attitude-control propellant tanks
9. Spherical oxidizer tank
10. Anti-slosh baffles
11. Fuel feed lines (18)
12. Quick-disconnect fittings (8)
13. Propellant turbopumps (18)
14. Peripherally arranged combustion chambers (36)
15. Oxidizer feed lines (18)
16. Liquid hydrogen tank for entry cooling
17. Turbine discharge lines (18)
18. Turbine discharge port
19. Oxidizer-tank-pressurization helium bottles (4)
20. Propellant tank for retro-thrust
21. Isentropic-expansion plug nozzle
22. Retractable landing legs (4)
23. Regeneration-cooling tubes
24. Liquid Oxygen Tank sump
25. Solid motors for thrust augmentation (4)
26. Liquid hydrogen manifold
27. Fuel manifold valve for liquid hydrogen tanks (8)
28. Attitude-control propellant tanks (4)
29. Centrebody recovery components
30. Cylindrical liquid hydrogen fuel tanks (8)
31. Tank recovery thermal protection (4)

Sea Dragon

delivered to LEO
Cost per
550 metric tons\$59/kg to \$600/kg
(1960 dollars)
Sea Dragon
(1963 design)
(recoverable)
Launch Cost\$300,000,000
(1960 dollars)
Height150 m
Diameter23 m
Stages2
Stage 1
OxidizerLO2
FuelRP-1
Thrust360,000,000 N
Wet Mass12,799,000 kg
Dry Mass1,333,000 kg
ΔV1,800 m/s
Max Accel4.21 g
Stage 2
OxidizerLO2
FuelLH2
Thrust62,800,000 N
Wet Mass4,823,000 kg
Dry Mass465,000 kg
ΔV5,400 m/s
Max Accel5.2 g
Mass Budget
Stage 112,799,000 kg
Stage 24,823,000 kg
Total18,121,000 kg

Details here, here, and here. Most of the illustrations here (and the data block at left) are from NASA-CR-52817 and NASA-CR-51034.

Sea Dragon was designed by Robert Truax in 1962 to be a low-cost heavy lift launch vehicle. A "big dumb booster", emphasis on "big". To reduce costs for launch pads and gantries, the vehicle was to be launched from the ocean. It would be towed out to the watery launch site, and the ballast tank in the first stage exhaust nozzle would be flooded. This would drag the tail down and the nose up, orienting the rocket into launch position.

At 150 m long and 23 m in diameter, Sea Dragon would have been the largest rocket ever built. To lower the cost of the rocket itself, it was designed to be build of inexpensive materials, specifically 8 mm steel sheeting.

The contruction techniques would be quite different than modern-day rockets. The latter are horribly damaged if they are touched by sea water, especially rocket engines. This is why SpaceX goes to the trouble of landing their reusable rockets on robot barges instead of letting them splash down in the ocean.

The design ground rules mandated a minimum payload of 450 metric tons delivered to a 600 kilometer orbit. For the reusable version of the vehicle, a 10 year useful life for the system was assumed.

The Sea Dragon project was shut down by NASA in the mid-1960's due to budget cuts.

Nuclear Thermal Turbo

delivered to LEO
Cost per
13 metric tons\$13,000
(with zero re-use)
Wet Mass 72,600 kg 35,700 kg 36,900 kg 7,260 kg 7,260 kg 13,000 19% 50.9% 1,662 sec 7,080 kg 2,270 kg

This is from The Nuclear Thermal Turbo Rocket: A Conceptual High-Performance Earth To Orbit Propulsion System by John R. Bucknell. John Bucknell was Senior Propulsion Engineer for the Raptor full-flow staged combustion methalox rocket at Spacex and is currently the Senior Propulsion Scientist for Divergent3D in Torrance, CA developing additively manufactured vehicle technologies. Slides from his talk are here.

Mr. Bucknell notes that the only practical method of dramatically bringing down the cost of boosting payloads into low Earth orbit (LEO) is to lower investment and realize a large return on that investment. The implication is you want a low dry mass Single Stage to Orbit Resuable Launch Vehicle with a high payload mass fraction. This is challenging.

Nuclear thermal rockets (NTR) have the highest specific impulse and thrust of available rockets. But the thrust-to-weight (T/W) ratio is poor since the blasted thing needs heavy radiations shielding. This really cuts into the payload fraction.

NERVA had a T/W of 5:1, particle bed had T/W of 15:1, and Miniature Reactor Engine (MITEE) managed 23:1. Unfortunately chemical LOX/RP-1 engines can achieve 150:1 easy.

Air-breathing propulsion has much higher specific impulse than NTR. But air-breathing propulsion don't work if there isn't any air. Long before LEO is reached the air pressure will drop below the level required for the air-breathing engine. Air breathers can only operate for the first 25% of the ascent, after that you need a rocket.

Therefore Mr. Bucknell's concept is to have a hybrid engine that can start in air-breathing hypersonic turbine mode and switch to NTR mode when the air runs out. This is called Nuclear Thermal Turbo Rocket (NTTR).

From Mach 0 to 8 the engine is in air-breathing subsonic ramjet mode. Combustion is subsonic. The nuclear rocket hot-hydrogen thrust is used to spin the fan rotor, driving the turbines. The hydrogen escapes via the trailing edge of the thrust fan blades. The turbine thrust fan blade vary their pitch and the variable nozzle throat geometry adapt to the changing atmospheric conditions. The turbine compresses the atmosphere from the inlet cone and the hydrogen from the thrust fan blades into the combustor, where they are burned for ramjet thrust.

From Mach 8 to 14 the engine is in air-breathing scramjet mode. Combustion is supersonic. The thrust fan blades lock into the neutral position aligned with the vehicle axis (depitches). The variable inlet cone expands, as does the PYBB variable nozzle.

From Mach 15 on up, the engine is in nuclear thermal rocket mode. The variable inlet cone contracts shut. The only thrust is rocket thrust from hot hydrogen escaping the trailing edge of the thrust fan blades.

Updated Version

Late breaking news, Mr. Bucknell has an updated paper out: The Turbo Rocket - A high performance air-breathing rocket propulsion system with nuclear and chemical variants.

Among other things the payload mass fraction calculations have been updated. The payload fraction has risen from 19% to 44.8%, for the 11 meter core version with a thrust of 1,150,000 Newtons and a mission average specific impulse of 1,695 seconds. The paper presents a sample lunar mission for comparison purposes.

The paper also discusses a totally non-nuclear version, citing the lack of available nuclear thermal propulsion hardware. Because that version has sigificantly poorer performance, and because this is the ATOMIC rocket website, I'm going to ignore it.

Improvements to the Turbo Rocket Concept

The aspects of the design that have been improved from the first paper are:

• Trajectory Optimization
• Scaling Sensitivity
• Increasing reuse through improving aerobraking performance
• Extending Airbreathing Operation

Trajectory Optimization

The first paper had plain vanilla unoptimized trajectory called Turbo Rocket Reference Trajectory MkI. This paper has the new and improved Trajectory II, which maintains inlet conditions for best air-breathing performance up to Mach 14. It also minimized airframe drag in pure rocket mode from Mach 15 to 25.

Scaling Sensitivity

The first paper had a wet-mass (GLOW or gross lift-off weight) 74,400 kg (160 klb) spacecraft with a core diameter of 3.66 meters, since with Trajectory MkI increasing the core to 5 meters reduces the payload fraction from 25.6% to 19.9%. Not good. The reduction is due to aerodynamic drag.

However, with the new and improved Trajectory MkII, increasing the core to 5 meters actually increases the payload fraction from 30.6% to 32.4%. So it is a win-win.

Now for a 445,500 kg (982 klb) GLOW spacecraft, the optimal core diameter is 11 meters. Payload fraction is 44.8%, which is fantastic! Table above includes some SpaceX boosters for comparison. SpaceX is nowhere near as good, but by the same token Elon Musk is not allowed to use nuclear rockets.

The NTTR was analyzed assuming a nuclear rocket designed around the Miniature Reactor Engine (MITEE) using highly enriched uranium (HEU, 98% Uranium 235). Actually I'm not sure that is accurate. 20%-85% U235 is highly enriched uranium. 85%-100% U235 is Weapons-Grade Uranium.

Which explains the report seriously looking into several other nuclear engines which use low enriched uranium (LEU or < 20% U235). Report says The availability of these reactors allows development with conventional nuclear fuel and doesn’t require the oversight required for highly enriched fuel. Translation: those reactors use the relatively cheap off-the-shelf commercial nuclear fuel, and you do not need an army of on-site killer SWAT teams to prevent terrorists from ripping off some HEU and making their very own terrorist nuclear bombs.

Sample Lunar Mission

A NTTR launch from Terra into LEO consumes about 42% of the GLOW, with 44% remaining for payload. Which obviously means a second NTTR could boost a complete propellant refueling load for the first ship (refuel load needs 42% GLOW of second NTTR, and it can boost 44%. 2% to spare). That would give the first ship enough delta V to go to Luna, land a large payload (68,000 kg) on the lunar surface, then lift-off and travel back to LEO (with 18,300 kg payload).

Table above has the details about the mission.

GCNR Liberty Ship

1,000 metric tons??

Anthony Tate has an interesting solution to the heavy lift problem. In his essay, he says that if we can grow up and stop panicking when we hear the N-word a reusable closed-cycle gas-core nuclear thermal rocket can boost huge amounts of payload into orbit. He calls it a "Liberty Ship." His design has a cluster of seven nuclear engines, with 1,200,000 pounds of thrust (5,340,000 newtons) each, from a thermal output of approximately 80 gigawatts. Exhaust velocity of 30,000 meters per second, which is a specific impulse of about 3060 seconds. Thrust to weight ratio of 10. Engine with safety systems, fuel storage, etc. masses 120,000 pounds or 60 short tons (54 metric tons ).

Using a Saturn V rocket as a template, the Liberty Ship has a wet mass of six million pounds (2,700,000 kilograms). Mr. Tate designs a delta V of 15 km/s, so it can has powered descent. It can take off and land. This implies a propellant mass of 2,400,000 pounds (1,100,000 kilograms). Using liquid hydrogen as propellant, this will make the propellant volume 15,200 cubic meters, since hydrogen is inconveniently non-dense. Say 20 meters in diameter and 55 meters long. It will be plump compared to a Saturn V.

Design height of 105 meters: 15 meters to the engines, 55 meters for the hydrogen tank, 5 meters for shielding and crew space, and a modular cargo area which is 30 meters high and 20 meters in diameter (enough cargo space for a good sized office building).

A Saturn V has a dry mass of 414,000 pounds (188,000 kilograms).

The Liberty Ship has seven engines at 120,000 pounds each, for a total of 840,000 pounds. Mr. Tate splurges and gives it a structural mass of 760,000 pounds, so it has plenty of surplus strength and redundancy. Add 2,400,000 pounds for reaction mass, and the Liberty Ship has a non-payload wet mass of 4,000,000 pounds.

Since it is scaled as a Saturn V, it is intended to have a total mass of 6,000,000 pounds. Subtract the 4,000,000 pound non-payload wet mass, and we discover that this brute can boost into low earth orbit a payload of Two Million Pounds. Great galloping galaxies! That's about 1000 metric tons, or eight times the boost of the Saturn V.

The Space Shuttle can only boost about 25 metric tons into LEO. The Liberty Ship could carry three International Space Stations into orbit in one trip.

Having said all this, it is important to keep in mind that a closed-cycle gas-core nuclear thermal rocket is a hideously difficult engineering feat, and we are nowhere near possessing the abilty to make one. An open-cycle gas-core rocket is much easier, but there is no way it would be allowed as a surface to orbit vehicle. Spray charges of fissioning radioactive plutonium death out the exhaust nozzle at fifty kilometers per second? That's not a lift off rocket, that's a weapon of mass destruction. However, see the Nexus.

There is an interesting analysis of the Liberty Ship on Next Big Future.

Uprated GCNR Nexus

1,500 metric tons??

This is from some fragmentary circa 1964 documents uncovered by The Unwanted Blog.

A Convair concept for an all-chemical Nexus SSTO launch vehicle with a second stage using open-cycle gas-core nuclear thermal rockets. Presumably the designers thought that the chemical stage would loft the second stage high enough so that the twin plumes of incandescent radioactive death would be diluted into plausible deniabilty.

Super Nexus

4,600 metric tons??

This is from some fragmentary circa 1964 documents uncovered by The Unwanted Blog.

This monster is the Uprated GCNR Nexus grown to three times the size. The document says that it can deliver 453 metric tons not to LEO, but to Lunar orbit. Doing some calculations on the back of an envelope with my slide rule, I estimate that it can loft 4,600 metric tons into LEO. And also with a proportional increase in radioactive exhaust.

A bit over 122 meters tall with the second stage having a diameter of 37 meters. Total wet mass of 10,900 metric tons. Second (nuclear) stage wet mass 5,900 metric tons for the Lunar orbit configuration. Dry second stage at Lunar orbit has a mass of 450 metric tons. The LEO configuration will be different.

The chemical stage has a total delta V capacity of 2.4 km/s. The gas core engines have a specific impulse rating of 2,220 seconds. The gas core stage in Lunar orbit configuration has a total delta V capacity of 21.8 km/s.

Aldebaran

27,000 metric tons??

This extreme heavy lift vehicle appears in Beyond Tomorrow by Dandridge Cole of "Macrolife" fame (Amherst Press 1965). The best place to watch lift-off is from an adjacent continent. That engine looks like it could accidentally vaporize Florida. They better work on the cargo handling system, though. Loading it crate by crate by helicopter is too much like eating a bowl of rice with tweezers one grain at a time.

Mr. Cole assumes that the economies of scale would dictate such a huge rocket to keep up with the orbital boost demands of the far-flung futurstic year 1990. The wet mass would be 50,000 tons. If the propulsion system had a specific impulse of 3,000 seconds, it would have a propellant fraction of 0.7 and a payload mass of 60 million pounds (27,000 metric tons). Or it could soft-land a smaller payload mass of 20,000 metric tons on Luna. If the propulsion system was weaker, say a specific impulse of 1,500 seconds, it would have a propellant fraction of 0.5 and a payload of 20 million pounds (9,000 metric tons). That propellant fraction doesn't make sense to me, I'll have to do the math.

The design is winged, for controlled aerodynamic Earth landing (now that would be a sight to see). Water take off and landing because there isn't a runway in the world that could survive that monster.

Orion

Planetary3,000 metric tons??
Super8,000,000 metric tons??

Thermonuclear Orion

Thermonuclear Orion
1,000 metric tons??
Performance
Engine TypeClean Fusion Orion
Engine Thrust3,000,000 N
Propellant Mass Flow10 kg/sec
Num Enginesx10
Total Thrust30,000,000 N
Total Propellant Mass Flow1,000 kg/sec
Exhaust Velocity30,000 m/s
Specific Impulse3,060 secs
(non-resuable)
ΔV8,000 m/s
Mass Ratio1.3
(resuable)
ΔV8,000 m/s
Mass Ratio1.65
Inert Mass600,000 kg
Dry Mass1,600,000 kg
Propellant Mass1,100,000 kg
Wet Mass2,700,000 kg

This is a species of Orion drive, including the useful ability to boost absurdly huge masses of payload into orbit. But with the attractive difference of not using dirty fission explosives for propulsion. It uses fusion explosions, triggered by convergent shock waves from chemical high explosives. Meaning there is zero radioactive fallout and arguably no problems from the Nuclear Test Ban Treaty. Yes, there will be some neutron radiation but you can't have everything.

The performance is very similar to the gas-core nuclear rocket Liberty Ship. But without the Liberty Ship's huge load of highly enriched uranium fuel, aka flying nuclear disaster waiting to happen. The Thermonuclear Orion's fuel would be non-radioactive deuterium and/or tritium. Both ships have approximately the same thrust (about 30,000,000 Newtons), approximately the same exhaust velocity (about 30,000 m/s, Isp around 3,060 secs), and approximately the same propellant mass flow (about 1,000 kg/sec).

Since they have the same exhaust velocity, both could manage the delta V for orbit (8,000 m/s) with a reasonable mass ratio of 1.3, or the delta V for orbit plus a powered landing (15,000 m/s) with a still reasonable mass ratio of 1.65. Which among other things means you don't have to deal with the design and maintenance nightmare called multi-staging, unlike pretty much all chemical rockets.

The amount of payload that can be carried depends upon design assumptions. As an example: the Liberty Ship was scaled to have the same mass as a Saturn V, but instead of the Saturn's top-notch payload of 118 metric tons, the Liberty Ship could boost a jaw-dropping 1,000 metric tons! Eight and a half times as much payload in one trip. And be resuable to boot. The Thermonuclear Orion's payload would be similar. Meaning a single launch could boost into orbit three International Space Stations and have enough spare payload capacity to squeeze in one Mir.

However, unlike the Liberty Ship, the Thermonuclear Orion will have severe design problems when it comes to landing the blasted thing. You see, when an Orion propulsion charge explodes in normal operation, the ship moves away from the explosion. Sadly, when landing, the ship will move into the nuclear explosion. For a conventional Orion using nuclear fission charges this would be suicide. The Thermonuke Orion might be able to get away with landing, since the fusion detonations are more like micro-explosions inside a mass of liquid hydrogen propellant.

RPL Fusion Engine

17 metric tons??

This appears to be an early version of Dr. John Slough magneto inertial fusion rocket.

Lofstrom loop

Small40,000 metric tons/year\$300/kg
Large6,000,000 metric tons/year\$3/kg

This was invented by Keith Lofstrom in 1981. Details about the mechanism of a Lofstrom loop can be found here and here, don't miss the paper here.

In science fiction, Lofstrom loops are featured in Heechee Rendezvous by Frederik Pohl, The Last Theorem by Arthur C. Clarke and Frederik Pohl, and Starquake by Robert Forward.

Rocket Sled

StarTram150,000 metric tons/year43/kg

Details about Rocket sled launch can be found here. Details about StarTram can be found here.

MagLifter

The report properly points out that NASA's Space Shuttle did many wondeful things, but lowering costs sadly was not one of them. NASA proudly predicted that the proposed shuttle could boost payload into orbit for \$260 US per kilogram. In practice the accurséd thing cost \$18,000 per kilogram of payload, which was pathetic compared to the \$5,000 per kilogram price of the non-resusable simple-as-dirt 1966-vintage Russian Proton booster.

Naturally researchers were motivated to find some alternative boost method that might lower the cost by a couple of orders of magnitude.

In theory electromagnetic acceleration should be far more efficient than using a disintegrating totem pole made of high exposives. However in the past applying electromagnetism to space launch took the form of guns, as in railguns and coilguns. Both of those are still not ready for prime-time, despite the military throwing lots of money at the project of turning them into weapons.

But the study author John Mankins said "What about magnetic levitation trains?" Good old MagLev. You know, the kind that was patented in 19-freaking-37 and which currently holds the speed record for rail vehicles? Technology that is actually being used in the real world in bullet trains is certainly mature technology.

The concept is called "MagLifter".

The bottom line is the MagLifter can provide the launch vehicle with a free 300 m/s of launch delta-V. Granted this is only about 3% of the total delta-V needed, but the cost savings are huge. It cuts the delta-V from the start of the launch, when the propellant cost per meter/sec is at its most expensive.

The paper has an analysis, comparing a sample single-stage to orbit rocket with the same rocket scaled down but using MagLifter. The scaled-down version saved 327 metric tons of wet mass, 24 metric tons of dry mass, and required only 4 rocket engines instead of 6.

Railguns and coilguns are typically short, since they have to fit on some sort of military vehicle. This means all the velocity has to be jammed into the projectile within the short length of the gun, meaning that the acceleration will be strong enough to smash an astronaut like a cockroach. It will also do nasty things to unliving cargo.

On the other hand since MagLifter is based on a railroad train, the accelerating segment can be, say, four kilometers long. This means velocity can be added at a much more leisurely pace and gentler acceleration. The advantage is that the astronauts don't die and the inert payloads do not need expensive reengineering.

Another advantage over railguns is that MagLifter does not expend lots of hardware with each launch (sabot, projectile heat shield, orbit insertion propulsion module).

And unlike the Space Shuttle, MagLifter does not require very high launch rates in order to achieve economical operations. Railgun launches are even worse, some concepts can only bring the price down to the goal by doing four launches per day.

The report estimates that current (1997) maglev train in the 300 miles-per-hour range costs about \$10 to \$20 million US per mile and \$3 to \$5 million US per train (payload of about 23 metric tons). Annual operations and maintenance cost around 1% of capital cost.

The MagLifter system has five major elements: Catapult, Structural Support Systems, Power Systems, Supporting Systems, and Launch Vehicles.

Catapult

The catapult has thee major elements: Maglev Guideway, Accelerator-Carrier Vehicle, and Accelerator-Carrier Staging Facility.

Maglev Guideway

This is the "rails" of the maglev railroad. It will be about 3 to 4 miles of maglev rails. 2.5 miles where the payload is accelerated, and 0.5 to 1.0 mile where the accelerator-carrier is frantically decelerated after the payload is launched. You have to be able to reuse the accelerator-carriers, those things are expensive. The acceleration segment is enclosed in a pressurized tube full of helium gas; since helium has low density, low drag forces, and a high speed of sound.

Accelerator-Carrier Vehicle

These are the "cars" that are accelerated by the railroad track. The launch vehicle is strapped to the accelerator-carrier with rapid precisely-controlled release mechanism. If the launch vehicle is extra-long, several accelerator-carriers will have to be linked like cars on a choo-choo train.

Each accelerator-carrier has cradles to give structural support to the launch vehicles during the acceleration phase. Mostly on the "rear" of the launch vehicle, so the carrier does not go shooting ahead while leaving the launch vehicle hovering in mid-air like Wile E. Coyote.

Accelerator-Carrier Staging Facility

This houses the operation control center, and the accelerator-carrier management center. This is where the launch vehicles are strapped to their carrier, and also contains the carrier servicing and maintenance facilities.

Structural Support Systems

This is the part that supports the maglev guideway. It is assumed to be mostly composed of a mountain, since building support towers two kilometers tall is a bit of a challenge. The guideway will either be on trestles set on the exterior of a mountain, inside a 'cut' made into the side of a mountain, or inside a tunnel in the mountain's interior.

It has three elements: Tunnel, Tunnel Environment Monitoring and Control Systems, and Launch / Exit systems.

Tunnel

The acceleration section of the maglev guideway is encased in a tunnel, to smooth things as the launch vehicle furiously accelerates. The deceleration section of the guideway has no encasing tunnel, but still needs trestles or something to support it.

Tunnel Environment Monitoring and Control Systems

The tunnel will be filled with a normal oxygen-nitrogen atmosphere at the start, but near the exit it will be filled with gaseous helium. This will provide a low-density low-drag medium as the launch vehicle exceeds Mach 1. The speed of sound in helium is also about 2.6 times what it is in ordinary atmosphere. This is a good thing because you do not want a sonic boom inside the tube.

The tunnel will need sensors and gas injectors to ensure the gaseous environment is arranged properly and the tube is clear of foreign objects.

Launch / Exit systems

This is the system that manages the separation of acceleration-carrier and launch vehicle, and their exit from the tube gaseous environment.

Power Systems

Energy Storage

This is a bank of batteries, probably a superconducting magnetic energy storage system (SMES). It will be gradually charged up from the local power grid, and used to power the launch. It would be nice to generate the required power during the launch. But since the blasted thing sucks 10 gigawats for a whopping 20 seconds, generating the power during launch is out of the question. Unless you have an antimatter power plant up your sleeve.

Power Management and Distribution

This system has to manage the massive discharge of the SMES and direct each watt to the proper component over the 30 second launch. It better not slip up or the maglev is in for tons of high-voltage fun that will make a thunderbolt look like scuffing your shoes on the carpet.

Thermal Management

The second law of thermodynamics says there will always be waste heat. If the maglev system is 80% efficient, this means the thermal management system has to deal with 40 gigaJoules of waste heat over three miles of catapult. Otherwise the entire thing turns into three miles of molten lava.

Supporting Systems

This is mostly the stuff crammed into the accelerator-carrier staging facility.

Staging Facilites

This handles staging for the launch vehicles, the payload, and the accelerator-carrier. It also handles mating the launch vehicle (including payload) with the accelerator-carrier, and performing maintenace on the launch vehicle following each flight.

Operations Control Center

The crew here control both the staging and launch operations. MagLifter operations rely upon rapid turn-around, low-cost (submarine-style) launch operations.

Installation Facilities

This is in charge of all those behind-the-scenes details vital to the operation. This includes maintenance on the access roads servicing the entire operation and temporary housing for the launch passengers.

Launch Vehicles

The small-, moderate-, and large-scale rockets that transport the payload the rest of the way to orbit. These are winged like the Space Shuttle, so they can return to the launch site and be re-used.

Prelude to Space

Prometheus Beta
(Prometheus Alpha Dry)
Wet Mass450,000 kg
EngineNuclear Ramjet/
NTR
PropellantAtomspheric gases/
Liquid Methane
Exhaust Velocity6,700 m/s atmo/
6,318 m/s methane
Specific Impulse683 sec atmo/
644 sec methane
Electromagnetic
sled velocity
220 m/s
(min ramjet 165 m/s)

This is from a science fiction novel by Sir Arthur C. Clarke. Keeping in mind that Clarke was the Chairman of the British Interplanetary Society from 1946 – 1947, and again from 1951 – 1953. The performance data for the nuclear stage was taken from the classic paper The Atomic Rocket by A.V. Cleaver and L.R. Shepherd, published in the Journal of the British Interplanetary Society in a series of articles September 1948–March 1949.

Verne Gun

280,000 metric tons??

Brian Wang has come up with an innovative concept. He mulled over a couple of his articles from his blog The Next Big Future (specifically this one and this one) and had an idea. Remember that one of the best propulsion systems for boosting huge payloads into orbit is the Orion drive; were it not for the fallout, the EMP, and the Nuclear Test Ban Treaty.

Then Mr. Wang thought about Jules Verne's novel From The Earth To The Moon, and the giant cannon Columbiad.

You set off one solitary ten megaton nuclear device in a deep underground salt dome. Perched on top is an Orion type spacecraft. All the EMP and radiation is contained in the underground cave (as has been done with historical underground nuclear tests). And 280,000 TONS of payload sails into low Earth orbit. Not pounds. Tons.

I say "sails into orbit", but of course it is more like "slammed by thousands of gs of acceleration", so this has to be unmanned (any human beings on board would instantly be converted into wall-gazpacho). But 280,000 tons? That's about one thousand International Space Stations, an entire Space Elevator (see below), an entire Lunar colony, an orbital fuel depot that would make future NASA missions ten times cheaper, a space station the size of the one in the movie 2001 A Space Odyssey, or about one-tenth of a ecologically clean 1.5 terawatt solar power station.

I know that nuclear-phobes will have a screaming fit, but this concept deserves close consideration.

Karl Schroeder analyzes the concept here.

Mass Driver

Mass Drivers are a way to use electromagnets to hurl, well, pretty much anything. But with respect to Surface To Orbit maneuvers, they can be used to accelerate spacecraft to assist their boost into orbit. They can also accelerate engine-less cannisters of cargo into orbit, if the mass driver is powerful enough.

They do have the side effect of turning a spaceport into an impromptu planetary fortress. After all, they are basically huge coil guns. This was popularized in the classic Robert Heinlein novel The Moon Is A Harsh Mistress.

The acceleration track has to be in vacuum, or air friction will do unfortunate things to the cargo cannister. Mass driver launchers on Terra have to be encased in a vacuum chamber, such a in the Bifrost Bridge. On Luna or other airless world they already have all the vacuum needed, you just have place a series of acceleration rings every few meters.

Laser Launch

Pournelle? metric tons/year\$1.9/kg plus power plant amortization
Jordin Kare HX Laser Launch3000 metric tons/year\$550/kg

Details about Laser Launch can be found here.

Matter Beam points out that the system will also work with an orbiting spacecraft equipped with a powerful laser battery, sending a beam to assist a surface-to-orbit shuttle lifting off. This can come in handy if the planet does not have a ground based laser launch facility, for instance an exploration spacecraft orbiting an uninhabited planet helping one of its landing craft return to the ship. A warship could also use its laser weapon batteries to give a boost to its fighters and missiles during a space battle, but I digress.

An important thing to keep in mind is that a laser-launch site is functionally equivalent to a planetary fortress. It can hurl projectiles and use laser beams directly at any invading spacecraft.

Laser Propulsion

This is from Laser Propulsion (1972)

This is a fairly standard laser launching setup. A high-powered laser is located at the launch pad. The spacecraft uses hydrogen propellant. Since hydrogen is regrettably transparent to most laser frequencies, it is seeded with some sort of powder to make it opaque. Otherwise the laser bolt would go sailing through the transparent hydrogen, not heating it at all, and fry the engine nozzle.

The laser energy heats the seed powder, which heats the hydrogen propellant. This is converted into high specific-impulse thrust by expanding the hydrogen through a nozzle. The report figures that the specific impulse that yields the highest payload boosted into LEO per total energy consumed lies in the range of 1,200 to 2,000 seconds.

The report figures that for a thrust-to-initial-weight ratio of 1.2 to 4.0, and with a specific impulse of 2,000 seconds, this will allow the rocket to have a whopping payload fraction of 0.20 to 0.40. This is fantastic! Most chemical rockets have payload fractions that are a miserable one-tenth of that, which implies the cost of a laser-launch vehicle per kilogram of payload could also be one-tenth of a chemical vehicle.

As with all laser-launch vehicles they require large electrical power plants to feed the hungry lasers. This will drive up the cost of the launch system, unless the power plant amortization is shared with other purposes (seawater desalinization or something valuable like that). But keep in mind that the launch vehicles are relatively low cost, each needs no expensive engine since the laser "engine" is based on the ground and can be shared by all the vehicles.

At theoretical maximums, the minimum energy required to transfer payload from Terra's surface into LEO is about 3×107 joules per kilogram. At 1972 prices this comes to about \$0.044 per kilogram. Naturally the laser launch system will be nowhere near this cheap, but there is plenty of room between that and the Falcon 9's \$2,720/kg

The hydrogen propellant along with the seed particles are injected into the exhaust nozzle through the porous walls, and are heated by the high powered laser beam. In other systems the injection system is replaced by a large slab of solid propellant, which is ablated by the laser beam into hot gas.

This system can also be used to propel spacecraft in LEO into trajectories to various destinations, or for ballistic transcontinental passenger services.

The report just looks at the performances and the simple costs of such a system. An over-all cost effectiveness study is beyond the scope of the report (translation: there was no funding for such a study).

The high-energy laser beam is directed to the propellant injection plate located inside the nozzle. The hydrogen propellant is seeded with something like carbon or natural uranium particles to render it opaque so it can absorb laser energy (this technique was previously studied for gas-core nuclear rockets). The propellant enthalpy and specific impulse are determined by dividing the beam power by the propellant mass flow rate. The hot propellant and vaporized seed material is expanded into space. A nozzle skirt directs the expansion to provide a more efficient conversion of thermal power into thrust.

Note that the laser beam does not have to be parallel to the thrust axis. The report figures that the rocket can be canted up to 45° off the laser beam and still work. The only limit is that the beam hits the propellant injection plate.

The skirt is protected from the high-temperature propellant by an opaque boundary-layer film, composed of more of the propellant+seed mix. This is injected through the porous or slotted walls of the skirt. This will reduce the specific impulse somewhat but it certainly would be a bad thing if the skirt was incinerated. The report figures that the skirt can be protected up to about a specific impulse of 5,000 seconds or so.

It is very important to keep the high-powered laser beam directed at the propellant plate on the rocket, otherwise the engine goes dead and the rocket plummets out of the sky. One solution is to have the laser beam direction slaved to tracking information sent by the rocket. The big laser has low-powered laser guidance beams parallel to the main beam, aimed at the rocket. The rocket skirt has laser detectors that can see the guidance beams. If the guidance beams start to drift off the detectors, the rocket sends radio signals to the big laser which allow it to get back on target. If the guidance beams drift too much off target, the high-powered beam will be reduced in power so it doesn't slice the rocket into bits.

Remember that laser launchers can probably be used as impromptu planetary defense weapons.

The intensity of the laser beam (kW/cm2) depends upon the desired propellant injection enthalpy and flow rate per unit area. It will probably on the order of megawatts per square centimeter. Naturally if this is beyond the capability of a single laser, there is no reason that an array of several laser cannot be used. In fact this is probably a good idea anyway, to allow redundancy and permit gracefull degradation if one laser malfunctions. Relying upon a single huge laser means if it malfunctions there is no way to prevent the rocket from doing an imitation of Icarus.

Obviously the laser frequencies will have to be ones that can penetrate Terra's atmosphere (no vacuum frequencies) and the spaceport should be located where the weather is not prone to clouds, smog, turbulance, or other atmospheric things that can attenuate the beam (and also subject to the standard spaceport location restrictions). Laser will need a line-of-sight to the rocket during boost phase.

The report figures that the thrust-producing part of the system (the part actually attached to the rocket) will be no heavier nor more complicated than a conventional rocket engine.

Analysis

The report assumes that the propellant absorbs all the laser energy. So the propellant enthalpy is

The specific impulse is related to the enthalpy by:

where they assume an overall nozzle expansion coeffcient CN ≈ 0.64 for various reason. Rearranging the equations gives:

which is plotted in Figure 2.

The laser beam power per thrust is calculated by:

and plotted in Figure 3.

The propellant mass fraction α is determined from the classic rocket equation

where

Here ΔVideal is the ideal mission velocity, e.g., 8,080 m/s for a low-orbit mission. It assumes a value of 1,070 m/s to account for atmospheric drag, and the fact the orbit will be elliptical rather than circular (because the rocket has to be visible to the ground laser during the entirety of the boost phase). The gravity drag is approximately g = 0.8 g0 which is conservative for a thrust-to-initial-weight ratios of 1.2 to 4.0.

Since tB = Ispα/k and substituting in equation (5) results in

Sove for α in terms of Isp and k gives Figure 4.

The payload mass fraction is calculated by using this propellant mass fraction and assuming a rocket structural weight fraction of 0.20. This results in:

which is plotted in Figure 5.

The laser beam energy per payload mass delivered to LEO is calculated using equations (3), (7), and (8) in the following equation:

which is plotted in Figure 6.

The electrical energy per payload mass depend upon overall beam efficiency Ee = Ebb. The value of ηb includes the ground-based electric-laser conversion efficiency and the laser beam transmission efficiency. The ground-based electric energy per payload mass is shown for various efficiencies in Figure 7.

The required electrical powerplant capacity per unit mass of payload is calculated from

and plotted in Figure 8.

The dollar cost of energy per payload mass in LEO is calculated by assuming an electrical cost of \$1.39×10-9 per joule (\$0.005/kW-hr)The energy cost per payload mass becomes

and is plotted in Figure 9.

The dollar cost of liquid hydrogen propellant per payload mass is calculated by using a projected future (+1972) cost of \$0.22 per kilogram. The propellant cost per pound of payload becomes:

Lightcraft

A Lightcraft is a type of laser launch vessel. Air enters in through vents at the waist. A laser beam is shined at the parabolic mirror on the base where it flash-heats the air there into plasma. The plasma rapidly escapes out of the base creating thrust. More air enters in through the waist vents and the cycles start again.

Since it is using atmospheric gas for propellant instead of on-board propellant, and the mass of the engine is at the spaceport instead of being on-board, most of the mass of the spacecraft will be payload. Instead of being mostly non-payload like most other booster vehicles.

Beamed Energy Propulsion (BEP)

This is from Beamed-Energy Propulsion (BEP) Study. The report looks at three different types of laser launch systems which are reasonably mature. This means the payloads are pretty small, forty to eighty kilograms as most (40 kg ` six cubesats). The payload mass will rise with technological advancement.

The hope was that using optical and millimeter wave lasers to power propulsion systems would give high exhaust velocity and high thrust. Plus the advantage that the power plant is on the ground instead of adding penalty mass to the boost vehicle.

• LASER OPTICAL: mirrored cowl intercepts and focuses laser light from a ground-based installation to heat atmosphere or water propellant.
• LASER THERMAL: propellant is circulated inside a large heat exchanger (HX). The exchanger is heated by a visible-light laser beam from a ground installation.
• MILLIMETER WAVE THERMAL: same as laser thermal, except instead of a visible light laser a microwave laser is used instead.

All the designs use two ground laser installations. The first is the "boost" beaming station, it is optimized to propel the spacecraft from the launch pad to high altitude as fast as possible. The "main" beaming station located downrange is optimized to delta-V the spacecraft up to orbital velocity and orbital height.

LASER OPTICAL

This engine operates in air-breathing laser ramjet mode from launch up to the time it reaches Mach 7 and an altitude of 35 kilometers. Then is switches to rocket mode using water propellant.

In both modes visible laser light is interceped by the mirrored base of the boost vehicle and funneled into the cowl. There the laser energy heats up either atmospheric gases or water propellant. The hot propellant exists through the bottom of the cowl, providing thrust. In ramjet mode the spacecraft forebody directs atmospheric gases into slots on the top of the cowl. The gases are compressed and injected into the laser cavity. In rocket mode the slots are closed, and water from onboard tanks is injected into the cowl.

The vehicle assembly building and the laser boost beaming station are a single building, unlike the other two concepts. This is because the laser beam has to be directed upwards into the base of the vehicle. The other two concepts direct the laser beam at the ventral side of the vehicle.

Risks and issues:

If the mirrored surface of the base and inside the cowl is damaged or degradated, the 3000 watts per square centimeter of laser energy will quickly burn through and destroy the launch vehicle. Mirror damage can come from debris impact or erosion by the propellant plasma.

If the propellant plasma comes witin a few centimeters of the mirror surface, there will be excessive heating. This is because the mirror surface is not as refective to the heat frequency from the plasma as it is to the laser beam frequency.

Laser light reflected off the mirror can possibly reach the surface of Terra, which could damage the eyesight of people on the ground watching the launch.

LASER THERMAL

A large external heat exchanger (HX) is heated by the ground-based laser installation. Liquid propellant (water and liquid hydrogen) from onboard tanks is heated inside the HX, and escapes through a conventional rocket nozzle to provide thrust.

Risks and issues:

In order to achieve the required heat transfer capabilities, the heat exchanger walls are very thin. This means the HX is very fragile. It can be broke by:

• Thermal stress due to large temperature variations during launch
• The temperature gradient across the HX during normal operation
• The high internal pressure due to the superheated and expanding hydrogen propellant

MILLIMETER WAVE THERMAL

Risks and issues:

Basically the same as the laser thermal: problems with the heat exchanger.

Bifrost Bridge

175,200 metric tons/year\$20/kg

This is a combination of a mass driver and a laser launch system. You can find details here.

Space Fountain

??

The Space Fountain utilizes fast streams of pellets that the tower structure couples to electromagnetically in order to support itself.

• Does not require materials with extreme strength
• Can be located at any point on a planet's surface instead of just the equator
• Can be raised to heights lower than the level of geostationary orbit

• Requires large constant amounts energy
• If the power is interrupted, the entire tower comes crashing down

Space Elevator

Number of
Elevators
12,000 metric tons\$3,000/kg
24,000 metric tons\$1,900/kg
36,000 metric tons\$1,600/kg

You can find details about space elevators here.

The big limitations are: it must be sited exactly on the equator, and it is absurdly vulnerable to sabotage.

You can read all about the complicated equations required to calculate the annual payload lifting capacity of a space elevator here. A baseline Edwards-Westling 20 metric ton space elevator powered by a bank of solar panels could boost about 272 metric tons a year. If powered by a large nuclear reactor it could boost about 2,720,000 metric tons a year.

Orbital Ring

An orbital ring superficially looks like a series of space elevators with the centers connected in a ring around the planet. But it is nothing of the sort.

The central station of a space elevator is 35,786 freaking kilometers from the surface of Terra, way out in geosynchronous orbit. The orbital ring is more like 300 to 600 kilometers, in LEO. The difference is that we have yet to find a material we can manufacture which will support a 35,786 km strand of itself, while a 600 km strand is orders of magnitude easier.

The ring is spinning at 8 kilometers/sec or whatever, thus preventing itself from collapsing onto the surface of Terra by centrifugal force. The "elevators" are tethers extending from the ring down to Terra's surface. The tether is attached to the ring indirectly by superconducting magnets. So the tether stays "stationary" over one spot on the ground moving at a speed of zero km/sec, while being magnetically attached to a ring moving at 8 km/sec.