A single-stage-to-orbit (or SSTO) vehicle reaches orbit from the surface of a body without jettisoning hardware, expending only propellants and fluids. The term usually, but not exclusively, refers to reusable vehicles.  No Earth-launched SSTO launch vehicles have ever been constructed. To date, orbital launches have been performed either by multi-stage fully or partially expendable rockets, or by the Space Shuttle which was multi-stage and partially reusable.
Launch costs for Low Earth Orbit (LEO) range from $4500 to $8500 per pound of payload. Reusable SSTO vehicles offer the promise of reduced launch expenses by eliminating recurring costs associated with hardware replacement inherent in expendable launch systems. However, the nonrecurring 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.
It is considered to be marginally possible to launch a single stage to orbit 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 the earth's gravity, especially in the early stages of flight; and flight within the Earth's atmosphere, which limits speed in the early stages of flight and influences engine performance. The marginality of SSTO can be seen in the launch of the space shuttle. The shuttle and main tank combination successfully orbits after booster separation from an altitude of 45 kilometres (28 mi) and a speed of 4,828 kilometres per hour (1,341 m/s; 3,000 mph). This is approximately 12% of the gravitational potential energy and just 3% of the kinetic energy needed for orbital velocity (4% of total energy required).
Notable single stage to orbit research spacecraft include Skylon, the DC-X, the X-33, and the Roton SSTO. However, despite showing some promise, none of them has come close to achieving orbit yet due to problems with finding the most efficient propulsion system.
Single-stage-to-orbit has been achieved from the Moon by both the Apollo program's Lunar Module and several robotic spacecraft of the Soviet Luna program; the lower lunar gravity and absence of any significant atmosphere makes this much easier than from Earth.
- 1 History
- 2 Approaches
- 3 Design challenges inherent in SSTO
- 4 Dense versus hydrogen fuels
- 5 One engine for all altitudes
- 6 Airbreathing SSTO
- 7 Launch assists
- 8 Nuclear propulsion
- 9 Beam-powered propulsion
- 10 Comparison with the Shuttle
- 11 Examples
- 12 Current development
- 13 Alternative approaches to inexpensive spaceflight
- 14 See also
- 15 References
- 16 External links
|This section requires expansion. (June 2011)|
- Early rocket pioneers believed that single stage to orbit was impossible.
- In the 1960s some people (Ex. Philip Bono) began to investigate SSTOs.
- From 1965 Robert Salked investigated various single stage to orbit spaceplane concepts.
- Around 1985 the NASP project was intended to create a scramjet vehicle to reach orbit, but this had its funding stopped and was cancelled.
- The HOTOL tried to use precooled jet engine technology, but failed to show significant advantages over rocket technology.
- Around 1992 the Skylon spaceplane concept was created.
- 1999-2001 Rotary Rocket attempted to build a SSTO called the Roton.
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.
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.
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.
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. It has a sufficiently efficient engine to achieve orbit, but without carrying much payload.
Design challenges inherent in SSTO
The design space constraints of SSTO vehicles were described by rocket design engineer Robert Truax:
The Tsiolkovsky rocket equation expresses the maximum change in velocity any single rocket stage can achieve:
- (delta-v) is the maximum change of velocity of the vehicle,
- is the propellant specific impulse,
- is the Standard Gravity,
- is the vehicle mass ratio,
- refers to the natural logarithm function.
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:
- 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. 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. This growth factor sensitivity is shown parametrically for both SSTO and two-stage-to-orbit (TSTO) vehicles for a standard LEO mission. 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, reusable SSTO vehicles are typically an impractical choice even when using the highest performance propellants available.
Dense versus hydrogen fuels
Hydrogen 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.
Hydrogen has the following advantages:
- 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:
- 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.
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 end result is the thrust/weight ratio of hydrogen-fueled engines is 30–50% lower than comparable engines using denser fuels.
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.
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.
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.
One engine for all altitudes
Some SSTO vehicles 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 optimal. Engines operating in the lower atmosphere have shorter bells than those designed to work in vacuum. Having a bell not optimized for the height makes the engine less efficient.
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 used in the X-33 design.
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 due to their excess weight. Some SSTO vehicles simply 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.
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.
Some of the issues with this approach are:
- 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). 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, 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 whereas since acceleration is a vector, 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.
Similar issues occur with single-stage vehicles attempting to carry conventional jet engines to orbit- the weight of the jet engines is not compensated by the reduction in propellant sufficiently.
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.
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.
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.
Proposed launch assists include:
- sled launch (rail, maglev including Bantam, MagLifter, and StarTram, etc.)
- aircraft tow
- in-flight fueling
- Lofstrom launch loop/space fountains
And on-orbit resources such as:
- Space tether
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. No current program is attempting nuclear propulsion from Earth's surface.
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 relatively undeveloped, and current technology falls far short of this.
Comparison with the Shuttle
The high cost per launch of the Space Shuttle sparked interest throughout the 1980s in designing a cheaper successor vehicle. Several official design studies were done, but most were basically smaller versions of the existing Shuttle concept.
Most cost analysis studies of the Space Shuttle have shown that workforce is by far the single greatest expense. Early shuttle discussions speculated airliner-type operation, with a two-week turnaround. However, senior NASA planners envisioned no more than 10 to 12 flights per year for the entire shuttle fleet. The absolute maximum flights per year for the entire fleet was limited by external tank manufacturing capacity to 24 per year.
Very efficient (hence complex and sophisticated) main engines were required to fit within the available vehicle space. Likewise the only known suitable lightweight thermal protection was delicate, maintenance-intensive silica tiles. These and other design decisions resulted in a vehicle that requires great maintenance after every mission. The engines are removed and inspected, and prior to the new "block II" main engines, the turbopumps were removed, disassembled and rebuilt. While Space Shuttle Atlantis was refurbished and relaunched in 53 days between missions STS-51-J and STS-61-B, generally months were required to repair an orbiter for a new mission.
Many in the aerospace community[who?] concluded that an entirely self-contained, reusable single-stage vehicle could solve these problems. The idea behind such a vehicle is to reduce the processing requirements from those of the Shuttle.
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 achieved deorbit to a soft landing, and return to lunar orbit, each with a single stage for descent and ascent.
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. 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 unmanned 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 crashed on the fourth flight after transferring management from the Strategic Defense Initiative Organization to NASA.
The Aquarius Launch Vehicle was designed to bring bulk materials to orbit as cheaply as possible.
The British Government partnered with the ESA in 2010 to promote a single-stage to orbit spaceplane concept called Skylon. This design was pioneered by Reaction Engines Limited (REL), a company founded by Alan Bond after HOTOL was canceled. The Skylon spaceplane has been positively received by the British government, and the British Interplanetary Society. 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. 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.
On 1 June 2012, Romanian organization ARCA announced that they are constructing an expendable rocket, named Haas 2C that will attempt to reach orbit in one stage. The rocket has 520 kilograms (1,150 lb) empty weight and can carry 15.5 tons of fuel. It will use kerosene as fuel and liquid oxygen as oxidizer. In Spring 2012 they have successfully tested a lightweight composite kerosene fuel tank. The liquid oxygen tank is being designed and it will also be made of composite materials. The launch is expected to take place in Spring 2013.
Alternative approaches to inexpensive spaceflight
Many studies have shown that regardless of selected technology, the most effective cost reduction technique is economies of scale. Merely launching a large total quantity reduces the manufacturing costs per vehicle, similar to how the mass production of automobiles brought about great increases in affordability.
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.
A related idea is to obtain economies of scale from building simple, massive, multi-stage rockets using cheap, off-the-shelf parts. The vehicles would be dumped into the ocean after use. This strategy is known as the "big dumb booster" approach.
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. These nations' launches are significantly cheaper than their Western counterparts.
An alternative to scale is to make the discarded stages practically reusable: this is the goal of the SpaceX reusable launch system development program and its Grasshopper demonstrator.
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