Launch system that only uses one rocket stage
A
single-stage-to-orbit
(
SSTO
) vehicle reaches
orbit
from the surface of a body using only propellants and fluids and without expending tanks, engines, or other major hardware. The term exclusively refers to
reusable vehicles
.
[1]
To date, no Earth-launched SSTO launch vehicles have ever been flown; orbital launches from Earth have been performed by either fully or partially
expendable
multi-stage rockets
.
The main projected advantage of the SSTO concept is elimination of the hardware replacement inherent in expendable launch systems. However, the non-recurring costs associated with design, development, research and engineering (DDR&E) of reusable SSTO systems are much higher than expendable systems due to the substantial technical challenges of SSTO, assuming that those technical issues can in fact be solved.
[2]
SSTO vehicles may also require a significantly higher degree of regular maintenance.
[3]
It is considered to be marginally possible to launch a single-stage-to-orbit
chemically fueled
spacecraft from Earth. The principal complicating factors for SSTO from Earth are: high orbital velocity of over 7,400 metres per second (27,000 km/h; 17,000 mph); the need to overcome Earth's gravity, especially in the early stages of flight; and flight within
Earth's atmosphere
, which limits speed in the early stages of flight due to drag, and influences engine performance.
[4]
Advances in rocketry in the 21st century have resulted in a substantial fall in the cost to launch a kilogram of payload to either
low Earth orbit
or the
International Space Station
,
[5]
reducing the main projected advantage of the SSTO concept.
Notable single stage to orbit concepts include
Skylon
, which used the hybrid-cycle SABRE engine that can use oxygen from the atmosphere when it is at low altitude, and then using onboard liquid oxygen after switching to the closed cycle rocket engine at high altitude, the McDonnell Douglas
DC-X
, the
Lockheed Martin X-33
and
VentureStar
which was intended to replace the Space Shuttle, and the
Roton SSTO
, which is a helicopter that can get to orbit. However, despite showing some promise, none of them have come close to achieving orbit yet due to problems with finding a sufficiently efficient propulsion system and discontinued development.
[1]
Single-stage-to-orbit is much easier to achieve on extraterrestrial bodies that have weaker gravitational fields and lower atmospheric pressure than Earth, such as the Moon and Mars, and has been achieved from the
Moon
by the
Apollo program
's
Lunar Module
, by several robotic spacecraft of the Soviet
Luna program
, and by China's
Chang'e 5
.
History
[
edit
]
Early concepts
[
edit
]
Before the second half of the twentieth century, very little research was conducted into space travel. During the 1960s, some of the first concept designs for this kind of craft began to emerge.
[6]
One of the earliest SSTO concepts was the expendable One stage Orbital Space Truck (OOST) proposed by
Philip Bono
,
[7]
an engineer for
Douglas Aircraft Company
.
[8]
A reusable version named ROOST was also proposed.
Another early SSTO concept was a reusable launch vehicle named
NEXUS
which was proposed by
Krafft Arnold Ehricke
in the early 1960s. It was one of the largest spacecraft ever conceptualized with a diameter of over 50 metres and the capability to lift up to 2000 short tons into Earth orbit, intended for missions to further out locations in the Solar System such as
Mars
.
[9]
[10]
The
North American Air Augmented VTOVL
from 1963 was a similarly large craft which would have used ramjets to decrease the liftoff mass of the vehicle by removing the need for large amounts of liquid oxygen while traveling through the atmosphere.
[11]
From 1965, Robert Salkeld investigated various single stage to orbit winged
spaceplane
concepts. He proposed a vehicle which would burn
hydrocarbon fuel
while in the atmosphere and then switch to
hydrogen fuel
for increasing efficiency once in space.
[12]
[13]
[14]
Further examples of Bono's early concepts (prior to the 1990s) which were never constructed include:
- ROMBUS (Reusable Orbital Module, Booster, and Utility Shuttle), another design from Philip Bono.
[15]
[16]
This was not technically single stage since it dropped some of its initial hydrogen tanks, but it came very close.
- Ithacus, an adapted ROMBUS concept which was designed to carry soldiers and military equipment to other continents via a sub-orbital trajectory.
[17]
[18]
- Pegasus, another adapted ROMBUS concept designed to carry passengers and payloads long distances in short amounts of time via space.
[19]
- Douglas SASSTO
, a 1967 launch vehicle concept.
[20]
- Hyperion, yet another Philip Bono concept which used a sled to build up speed before liftoff to save on the amount of fuel which had to be lifted into the air.
[21]
Star-raker
: In 1979
Rockwell International
unveiled a concept for a 100-ton payload heavy-lift multicycle airbreather ramjet/
cryogenic rocket engine
, horizontal takeoff/horizontal landing single-stage-to-orbit spaceplane named
Star-Raker
, designed to launch heavy
Space-based solar power
satellites into a 300 nautical mile Earth orbit.
[22]
[23]
[24]
Star-raker would have had 3 x LOX/LH2 rocket engines (based on the
SSME
) + 10 x turboramjets.
[22]
Around 1985 the
NASP
project was intended to launch a scramjet vehicle into orbit, but funding was stopped and the project cancelled.
[25]
At around the same time, the
HOTOL
tried to use
precooled jet engine
technology, but failed to show significant advantages over rocket technology.
[26]
DC-X technology
[
edit
]
The DC-X, short for Delta Clipper Experimental, was an uncrewed one-third scale vertical takeoff and landing demonstrator for a proposed SSTO. It is one of only a few prototype SSTO vehicles ever built. Several other prototypes were intended, including the DC-X2 (a half-scale prototype) and the DC-Y, a full-scale vehicle which would be capable of single stage insertion into orbit. Neither of these were built, but the project was taken over by
NASA
in 1995, and they built the DC-XA, an upgraded one-third scale prototype. This vehicle was lost when it landed with only three of its four landing pads deployed, which caused it to tip over on its side and explode. The project has not been continued since.
[
citation needed
]
Roton
[
edit
]
From 1999 to 2001 Rotary Rocket attempted to build a SSTO vehicle called the Roton. It received a large amount of media attention and a working sub-scale prototype was completed, but the design was largely impractical.
[27]
Approaches
[
edit
]
There have been various approaches to SSTO, including pure rockets that are launched and land vertically, air-breathing
scramjet
-powered vehicles that are launched and land horizontally,
nuclear-powered
vehicles, and even
jet-engine
-powered vehicles that can fly into orbit and return landing like an airliner, completely intact.
For rocket-powered SSTO, the main challenge is achieving a high enough mass-ratio to carry sufficient
propellant
to achieve
orbit
, plus a meaningful
payload
weight. One possibility is to give the rocket an initial speed with a
space gun
, as planned in the
Quicklaunch
project.
[28]
For air-breathing SSTO, the main challenge is system complexity and associated
research and development
costs,
material science
, and construction techniques necessary for surviving sustained high-speed flight within the atmosphere,
and
achieving a high enough mass-ratio to carry sufficient propellant to achieve orbit, plus a meaningful payload weight. Air-breathing designs typically fly at
supersonic
or
hypersonic
speeds, and usually include a rocket engine for the final burn for orbit.
[1]
Whether rocket-powered or air-breathing, a reusable vehicle must be rugged enough to survive multiple round trips into space without adding excessive weight or maintenance. In addition a reusable vehicle must be able to reenter without damage, and land safely.
[
citation needed
]
While single-stage rockets were once thought to be beyond reach, advances in materials technology and construction techniques have shown them to be possible. For example, calculations show that the
Titan II
first stage, launched on its own, would have a 25-to-1 ratio of fuel to vehicle hardware.
[29]
It has a sufficiently efficient engine to achieve orbit, but without carrying much payload.
[30]
Dense versus hydrogen fuels
[
edit
]
Hydrogen fuel
might seem the obvious fuel for SSTO vehicles. When burned with
oxygen
, hydrogen gives the highest
specific impulse
of any commonly used fuel: around 450 seconds, compared with up to 350 seconds for
kerosene
.
[
citation needed
]
Hydrogen has the following advantages:
[
citation needed
]
- Hydrogen has nearly 30% higher specific impulse (about 450 seconds vs. 350 seconds) than most dense fuels.
- Hydrogen is an excellent coolant.
- The gross mass of hydrogen stages is lower than dense-fuelled stages for the same payload.
- Hydrogen is environmentally friendly.
However, hydrogen also has these disadvantages:
[
citation needed
]
- Very low density (about
1
⁄
7
of the density of kerosene) ? requiring a very large tank
- Deeply
cryogenic
? must be stored at very low temperatures and thus needs heavy insulation
- Escapes very easily from the smallest gap
- Wide combustible range ? easily ignited and burns with a dangerously invisible flame
- Tends to condense oxygen which can cause flammability problems
- Has a large
coefficient of expansion
for even small heat leaks.
These issues can be dealt with, but at extra cost.
[
citation needed
]
While kerosene tanks can be 1% of the weight of their contents, hydrogen tanks often must weigh 10% of their contents. This is because of both the low density and the additional insulation required to minimize boiloff (a problem which does not occur with kerosene and many other fuels). The low density of hydrogen further affects the design of the rest of the vehicle: pumps and pipework need to be much larger in order to pump the fuel to the engine. The result is the thrust/weight ratio of hydrogen-fueled engines is 30?50% lower than comparable engines using denser fuels.
[
citation needed
]
This inefficiency indirectly affects
gravity losses
as well; the vehicle has to hold itself up on rocket power until it reaches orbit. The lower excess thrust of the hydrogen engines due to the lower thrust/weight ratio means that the vehicle must ascend more steeply, and so less thrust acts horizontally. Less horizontal thrust results in taking longer to reach orbit, and gravity losses are increased by at least 300 metres per second (1,100 km/h; 670 mph). While not appearing large, the mass ratio to
delta-v
curve is very steep to reach orbit in a single stage, and this makes a 10% difference to the mass ratio on top of the tankage and pump savings.
[
citation needed
]
The overall effect is that there is surprisingly little difference in overall performance between SSTOs that use hydrogen and those that use denser fuels, except that hydrogen vehicles may be rather more expensive to develop and buy. Careful studies have shown that some dense fuels (for example liquid
propane
) exceed the performance of hydrogen fuel when used in an SSTO launch vehicle by 10% for the same dry weight.
[31]
In the 1960s
Philip Bono
investigated single-stage, VTVL
tripropellant rockets
, and showed that it could improve payload size by around 30%.
[32]
Operational experience with the
DC-X
experimental rocket has caused a number of SSTO advocates to reconsider hydrogen as a satisfactory fuel. The late Max Hunter, while employing hydrogen fuel in the DC-X, often said that he thought the first successful orbital SSTO would more likely be fueled by propane.
[
citation needed
]
One engine for all altitudes
[
edit
]
Some SSTO concepts use the same engine for all altitudes, which is a problem for traditional engines with a bell-shaped
nozzle
. Depending on the atmospheric pressure, different bell shapes are required. Engines designed to operate in a vacuum have large bells, allowing the exhaust gasses to expand to near vacuum pressures, thereby raising efficiency.
[33]
Due to an effect known as
Flow separation
, using a vacuum bell in atmosphere would have disastrous consequences for the engine. Engines designed to fire in atmosphere therefore have to shorten the nozzle, only expanding the gasses to atmospheric pressure. The efficiency losses due to the smaller bell are usually mitigated via staging, as upper stage engines such as the
Rocketdyne J-2
do not have to fire until atmospheric pressure is negligible, and can therefore use the larger bell.
One possible solution would be to use an
aerospike engine
, which can be effective in a wide range of ambient pressures. In fact, a linear aerospike engine was to be used in the
X-33
design.
[34]
Other solutions involve using multiple engines and other
altitude adapting designs
such as double-mu bells or
extensible bell sections
.
[
citation needed
]
Still, at very high altitudes, the extremely large engine bells tend to expand the exhaust gases down to near vacuum pressures. As a result, these engine bells are counterproductive
[
dubious
–
discuss
]
due to their excess weight. Some SSTO concepts use very high pressure engines which permit high ratios to be used from ground level. This gives good performance, negating the need for more complex solutions.
[
citation needed
]
Airbreathing SSTO
[
edit
]
Some designs for SSTO attempt to use
airbreathing jet engines
that collect oxidizer and reaction mass from the atmosphere to reduce the take-off weight of the vehicle.
[35]
Some of the issues with this approach are:
[
citation needed
]
- No known air breathing engine is capable of operating at orbital speed within the atmosphere (for example hydrogen fueled
scramjets
seem to have a top speed of about Mach 17).
[36]
This means that rockets must be used for the final orbital insertion.
- Rocket thrust needs the orbital mass to be as small as possible to minimize propellant weight.
- The thrust-to-weight ratio of rockets that rely on on-board oxygen increases dramatically as fuel is expended, because the oxidizer fuel tank has about 1% of the mass as the oxidizer it carries, whereas air-breathing engines traditionally have a poor thrust/weight ratio which is relatively fixed during the air-breathing ascent.
- Very high speeds in the atmosphere necessitate very heavy thermal protection systems, which makes reaching orbit even harder.
- While at lower speeds, air-breathing engines are very efficient, but the efficiency (
Isp
) and thrust levels of air-breathing jet engines drop considerably at high speed (above Mach 5?10 depending on the engine) and begin to approach that of rocket engines or worse.
- Lift to drag ratios
of vehicles at hypersonic speeds are poor, however the effective lift to drag ratios of rocket vehicles at high g is
not dissimilar
.
Thus with for example scramjet designs (e.g.
X-43
) the mass budgets do not seem to close for orbital launch.
[
citation needed
]
Similar issues occur with single-stage vehicles attempting to carry conventional jet engines to orbit?the weight of the jet engines is not compensated sufficiently by the reduction in propellant.
[37]
On the other hand, LACE-like
precooled airbreathing
designs such as the
Skylon spaceplane
(and
ATREX
) which transition to rocket thrust at rather lower speeds (Mach 5.5) do seem to give, on paper at least, an improved orbital
mass fraction
over pure rockets (even multistage rockets) sufficiently to hold out the possibility of full reusability with better payload fraction.
[38]
It is important to note that mass fraction is an important concept in the engineering of a rocket. However, mass fraction may have little to do with the costs of a rocket, as the costs of fuel are very small when compared to the costs of the engineering program as a whole. As a result, a cheap rocket with a poor mass fraction may be able to deliver more payload to orbit with a given amount of money than a more complicated, more efficient rocket.
[
citation needed
]
Launch assists
[
edit
]
Many vehicles are only narrowly suborbital, so practically anything that gives a relatively small delta-v increase can be helpful, and outside assistance for a vehicle is therefore desirable.
[
citation needed
]
Proposed launch assists include:
[
citation needed
]
And on-orbit resources such as:
[
citation needed
]
Nuclear propulsion
[
edit
]
Due to weight issues such as shielding, many nuclear propulsion systems are unable to lift their own weight, and hence are unsuitable for launching to orbit. However, some designs such as the
Orion project
and some
nuclear thermal
designs do have a
thrust to weight ratio
in excess of 1, enabling them to lift off. Clearly, one of the main issues with nuclear propulsion would be safety, both during a launch for the passengers, but also in case of a failure during launch. As of February 2024, no current program is attempting nuclear propulsion from Earth's surface.
[
citation needed
]
Beam-powered propulsion
[
edit
]
Because they can be more energetic than the potential energy that chemical fuel allows for, some laser or microwave powered rocket concepts have the potential to launch vehicles into orbit, single stage. In practice, this area is not possible with current technology.
[
citation needed
]
Design challenges inherent in SSTO
[
edit
]
The design space constraints of SSTO vehicles were described by rocket design engineer
Robert Truax
:
Using similar technologies (i.e., the same propellants and structural fraction), a two-stage-to-orbit vehicle will always have a better payload-to-weight ratio than a single stage designed for the same mission, in most cases, a very much better [payload-to-weight ratio]. Only when the structural factor approaches zero [very little vehicle structure weight] does the payload/weight ratio of a single-stage rocket approach that of a two-stage. A slight miscalculation and the single-stage rocket winds up with no payload. To get any at all, technology needs to be stretched to the limit. Squeezing out the last drop of specific impulse, and shaving off the last pound, costs money and/or reduces reliability.
[40]
The
Tsiolkovsky rocket equation
expresses the maximum change in velocity any single rocket stage can achieve:
where:
(
delta-v
) is the maximum change of velocity of the vehicle,
is the vehicle
mass ratio
,
The mass ratio of a vehicle is defined as a ratio the initial vehicle mass when fully loaded with propellants
to the final vehicle mass
after the burn:
where:
is the initial vehicle mass or the
gross liftoff weight
,
is the final vehicle mass after the burn,
is the structural mass of vehicle,
is the propellant mass,
is the payload mass.
The
propellant mass fraction
(
) of a vehicle can be expressed solely as a function of the mass ratio:
The structural coefficient (
) is a critical parameter in SSTO vehicle design.
[41]
Structural efficiency of a vehicle is maximized as the structural coefficient approaches zero. The structural coefficient is defined as:
The overall structural mass fraction
can be expressed in terms of the structural coefficient:
An additional expression for the overall structural mass fraction can be found by noting that the payload mass fraction
, propellant mass fraction and structural mass fraction sum to one:
Equating the expressions for structural mass fraction and solving for the initial vehicle mass yields:
This expression shows how the size of a SSTO vehicle is dependent on its structural efficiency. Given a mission profile
and propellant type
, the size of a vehicle increases with an increasing structural coefficient.
[42]
This growth factor sensitivity is shown parametrically for both SSTO and
two-stage-to-orbit
(TSTO) vehicles for a standard LEO mission.
[43]
The curves vertically asymptote at the maximum structural coefficient limit where mission criteria can no longer be met:
In comparison to a non-optimized TSTO vehicle using
restricted staging
, a SSTO rocket launching an identical payload mass and using the same propellants will always require a substantially smaller structural coefficient to achieve the same delta-v. Given that current materials technology places a lower limit of approximately 0.1 on the smallest structural coefficients attainable,
[44]
reusable SSTO vehicles are typically an impractical choice even when using the highest performance propellants available.
Examples
[
edit
]
It is easier to achieve SSTO from a body with lower gravitational pull than Earth, such as the
Moon
or
Mars
. The
Apollo Lunar Module
ascended from the lunar surface to lunar orbit in a single stage.
[45]
A detailed study into SSTO vehicles was prepared by
Chrysler Corporation
's Space Division in 1970?1971 under NASA contract NAS8-26341. Their proposal (
Shuttle SERV
) was an enormous vehicle with more than 50,000 kilograms (110,000 lb) of payload, utilizing
jet engines
for (vertical) landing.
[46]
While the technical problems seemed to be solvable, the
USAF
required a winged design that led to the Shuttle as we know it today.
The uncrewed
DC-X
technology demonstrator, originally developed by
McDonnell Douglas
for the
Strategic Defense Initiative
(SDI) program office, was an attempt to build a vehicle that could lead to an SSTO vehicle. The one-third-size test craft was operated and maintained by a small team of three people based out of a trailer, and the craft was once relaunched less than 24 hours after landing. Although the test program was not without mishap (including a minor explosion), the DC-X demonstrated that the maintenance aspects of the concept were sound. That project was cancelled when it landed with three of four legs deployed, tipped over, and exploded on the fourth flight after transferring management from the
Strategic Defense Initiative Organization
to NASA.
[
citation needed
]
The
Aquarius Launch Vehicle
was designed to bring bulk materials to orbit as cheaply as possible.
[
citation needed
]
Current development
[
edit
]
Current and previous SSTO projects include the Japanese
Kankoh-maru
project,
ARCA Haas 2C
,
Radian One
and the Indian
Avatar
spaceplane.
[
citation needed
]
Skylon
[
edit
]
The British Government partnered with the
ESA
in 2010 to promote a
single-stage to orbit
spaceplane
concept called
Skylon
.
[47]
This design was pioneered by
Reaction Engines Limited (REL)
,
[48]
[49]
a company founded by
Alan Bond
after
HOTOL
was canceled.
[50]
The Skylon spaceplane has been positively received by the British government, and the
British Interplanetary Society
.
[51]
Following a successful propulsion system test that was audited by ESA's propulsion division in mid-2012, REL announced that it would begin a three-and-a-half-year project to develop and build a test jig of the
Sabre engine
to prove the engines performance across its air-breathing and rocket modes.
[52]
In November 2012, it was announced that a key test of the engine precooler had been successfully completed, and that ESA had verified the precooler's design. The project's development is now allowed to advance to its next phase, which involves the construction and testing of a full-scale prototype engine.
[52]
[53]
Starship
[
edit
]
Elon Musk, CEO of SpaceX, has claimed that the
upper stage of the prototype "Starship" rocket
, currently in development in
Starbase (Texas)
, has the capability to reach orbit as an SSTO. However he concedes that if this was done, there would be no appreciable mass left for a
heat shield
, landing legs, or fuel to land, much less any usable payload.
[54]
Alternative approaches to inexpensive spaceflight
[
edit
]
Many studies have shown that regardless of selected technology, the most effective cost reduction technique is
economies of scale
.
[
citation needed
]
Merely launching a large total number reduces the manufacturing costs per vehicle, similar to how the
mass production
of automobiles brought about great increases in affordability.
[
citation needed
]
Using this concept, some aerospace analysts believe the way to lower launch costs is the exact opposite of SSTO. Whereas reusable SSTOs would reduce per launch costs by making a reusable high-tech vehicle that launches frequently with low maintenance, the "mass production" approach views the technical advances as a source of the cost problem in the first place. By simply building and launching large quantities of rockets, and hence launching a large volume of payload, costs can be brought down. This approach was attempted in the late 1970s, early 1980s in
West Germany
with the
Democratic Republic of the Congo
-based
OTRAG rocket
.
[55]
This is somewhat similar to the approach some previous systems have taken, using simple engine systems with "low-tech" fuels, as the
Russian
and
Chinese space programs
still do.
[
citation needed
]
An alternative to scale is to make the discarded stages practically
reusable
: this was the original design goal of the
Space Shuttle
phase B studies, and is currently pursued by the
SpaceX reusable launch system development program
with their
Falcon 9
,
Falcon Heavy
, and
Starship
, and
Blue Origin
using
New Glenn
.
See also
[
edit
]
Further reading
[
edit
]
- Andrew J. Butrica:
Single Stage to Orbit - Politics, Space Technology, and the Quest for Reusable Rocketry.
The Johns Hopkins University Press, Baltimore 2004,
ISBN
9780801873386
.
References
[
edit
]
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a
b
c
Richard Varvill & Alan Bond (2003).
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(PDF)
.
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(PDF)
on 15 June 2011
. Retrieved
5 March
2011
.
- ^
Dick, Stephen and Lannius, R., "Critical Issues in the History of Spaceflight," NASA Publication SP-2006-4702, 2006.
- ^
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.
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Bibcode
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.
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Archived
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2021
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- ^
Toso, Federico.
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(PDF)
.
Centre for Future Air Space Transportation Technologies
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(PDF).
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- ^
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- ^
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- ^
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.
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"
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.
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.
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.
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www.astronautix.com
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a
b
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.
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.
- ^
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.
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.
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- ^
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External links
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Concepts
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Physical propulsion
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Chemical propulsion
| State
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Propellants
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Power cycles
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Intake mechanisms
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Electrical propulsion
| Electrostatic
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Electromagnetic
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Electrothermal
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Other
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Nuclear propulsion
| Closed system
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Open system
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External power
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Related concepts
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