No Earth-launched SSTO launch vehicles have ever been constructed. Current orbital launches are either performed by multi-stage fully expendable rockets, or by the Space Shuttle which is multi-stage and partially reusable. Several research spacecraft have been designed and partially or completely constructed, including the DC-X, the X-33, and the Roton SSTO. However, none of them has come close to achieving orbit.
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 programme; the lower lunar gravity and absence of any significant atmosphere makes this much easier than from Earth.
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 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.
The goals of fully reusable SSTO vehicles are lower operating costs, improved safety, and better reliability than current launch vehicles. The ultimate goal for an SSTO vehicle would be airliner-like operations.
However, even a non-reusable single-stage vehicle might be worth building, since it would have a much lower part count, and may therefore be cheaper to design and build.
For pure rocket approaches Tsiolkovsky's rocket equation shows that dead weight will prevent reaching orbit unless the ratio of propellant to structural mass (called mass ratio) is very high — between about 10 and 25 (i.e. 24 parts propellant weight to 1 part structural weight; depending on propellant choice). By contrast, an airliner has a mass ratio of about 2 (1 part fuel to 1 part structural weight). High performance aircraft (such as stunt planes and jet fighters) tend to have a mass ratio between 2.5 (40% structure) and 4 (25% structure). However, the Rutan Voyager aircraft had a mass ratio of over 10.
It is extremely difficult to design a structure which is strong, safe, very light, and economical to build. Designers often liken the task to designing and building an egg shell. The problem originally seemed insuperable, and drove all early designers to multistage rockets.
Multistage rockets are able to reach orbital velocity because they discard structural weight during launch. Thus a single-stage rocket is at a disadvantage because it must carry its entire vehicle mass to orbit, which in turn reduces payload capacity. On the other hand, a single-stage vehicle doesn't have to carry a second stage, so the vehicle is easier to make lightweight.
Alternatively, since expendable multistage rockets entail discarding costly structure and engines, if the stages could be reused, this could permit much cheaper operation since the parts costs would be amortized over many flights.
One problem with multistage reusable rockets is the difficulty of reusing even the first stage, and the development cost of such a large device. Analysis shows the optimum staging velocity (the speed at which the first stage is dropped) is very high — possibly 3.65 km/s (12,000 feet per second). This means after separation, the large first stage is at high altitude and headed downrange very fast, which makes it difficult to turn around and get back to the launch point. The stage also must reenter without damage from a speed as high as Mach 10.
The reusable first stage would be very large, nearly the size of a Saturn V to lift an orbiter the size of the current shuttle. Because development cost of aerospace vehicles is related to weight, it would be extremely expensive to develop.
Some approaches envisioned parachutes to gently lower a reusable first stage. However, for most US launches the trajectory is over the Atlantic ocean, and complex liquid-fueled stages are easily damaged by a salt water landing.
These problems with the multistage approach drive the design path toward SSTO.
All these complications drove designers to consider a single reusable stage as this:
If an SSTO vehicle were combined with reliable systems and lower maintenance design of a more automated nature, it could greatly reduce operational costs.
Single-stage rockets were once thought to be beyond reach, but 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.
It is now clear an expendable SSTO vehicle is achievable, but it is less clear if a reusable SSTO carrying a useful payload can be built and operated for a feasible cost with current or near-term projected technologies.
Hydrogen has the following advantages:
However, hydrogen also has these disadvantages:
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 meters per second. 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.
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 jet engines that collect oxidiser from the atmosphere to reduce the take-off weight of the vehicle.
Some of the issues with this approach are:
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 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.
Proposed launch assists include:
And on-orbit resources such as:
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.
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 are required to repair an orbiter for a new mission. Given that there are 25,000 people working on Shuttle operations, the payroll alone is the Shuttle's single biggest operating cost.
Many in the aerospace community 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.
The first stage of the Titan II had the mass ratio required for single-stage-to-orbit capability with a small payload. A rocket stage is not a complete launch vehicle, but this demonstrates that an expendable SSTO was probably achievable with 1962 technology.
The Apollo Lunar Module was a true SSTO vehicle, albeit on the moon. It achieved lunar orbit using a single stage.
A detailed study into SSTO vehicles was prepared by Chrysler Corporation's Space Division in 1970–1971 under NASA contract NAS8-26341. Their proposal was an enormous vehicle with more than 50,000 kg of payload, utilizing jet engines for (vertical) landing. While the technical problems seemed to be solvable, the USAF required a winged design (for cross range) 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 tiny crew 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.
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 70’s, early 80’s in West Germany with the Democratic Republic of the Congo-based OTRAG rocket and could have been successful if the project was not killed following political pressure from France and the Soviet Union.
A related idea is to simply 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.