单级入轨

DC-X技术

The maiden flight of the DC-X

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.

Roton

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.[22]

方法

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.[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.

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.[23] It has a sufficiently efficient engine to achieve orbit, but without carrying much payload.[24]

Airbreathing SSTO

Skylon spaceplane

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).[27] 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.

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.[28]

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.[29]

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.

单级入轨的设计难题

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.[31]

The Tsiolkovsky rocket equation expresses the maximum change in velocity any single rocket stage can achieve:

${\displaystyle \Delta v=I_{\text{sp}}\cdot g_{0}\ln(MR)}$

where:

${\displaystyle \Delta v}$ (delta-v) is the maximum change of velocity of the vehicle,

${\displaystyle I_{\text{sp}}}$ is the propellant specific impulse,

${\displaystyle g_{0}}$ is the standard gravity,

${\displaystyle MR}$ is the vehicle mass ratio,

${\displaystyle \ln }$ 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 ${\displaystyle \left(m_{i}\right)}$ to the final vehicle mass ${\displaystyle \left(m_{f}\right)}$ after the burn:

${\displaystyle MR={\frac {m_{i}}{m_{f}}}={\frac {m_{p}+m_{s}+m_{\text{pl}}}{m_{s}+m_{\text{pl}}}}}$

where:

${\displaystyle m_{i}}$ is the initial vehicle mass or the gross liftoff weight ${\displaystyle \left(GLOW\right)}$,

${\displaystyle m_{f}}$ is the final vehicle mass after the burn,

${\displaystyle m_{s}}$ is the structural mass of vehicle,

${\displaystyle m_{p}}$ is the propellant mass,

${\displaystyle m_{\text{pl}}}$ is the payload mass.

The propellant mass fraction (${\displaystyle \zeta }$) of a vehicle can be expressed solely as a function of the mass ratio:

${\displaystyle \zeta ={\frac {m_{p}}{m_{i}}}={\frac {m_{i}-m_{f}}{m_{i}}}=1-{\frac {m_{f}}{m_{i}}}=1-{\frac {1}{MR}}={\frac {MR-1}{MR}}}$

The structural coefficient (${\displaystyle \lambda }$) is a critical parameter in SSTO vehicle design.[32] Structural efficiency of a vehicle is maximized as the structural coefficient approaches zero. The structural coefficient is defined as:

Comparison of growth factor sensitivity for Single-Stage-to-Orbit (SSTO) and restricted stage Two-Stage-to-Orbit (TSTO) vehicles. Based on a LEO mission of Delta v = 9.1 km/s and payload mass = 4500 kg for range of propellant Isp.

${\displaystyle \lambda ={\frac {m_{s}}{m_{p}+m_{s}}}={\frac {m_{s}}{m_{i}-m_{\text{pl}}}}={\frac {\frac {m_{s}}{m_{i}}}{1-{\frac {m_{\text{pl}}}{m_{i}}}}}}$

The overall structural mass fraction ${\displaystyle \left({\frac {m_{s}}{m_{i}}}\right)}$ can be expressed in terms of the structural coefficient:

${\displaystyle {\frac {m_{s}}{m_{i}}}=\lambda \left(1-{\frac {m_{\text{pl}}}{m_{i}}}\right)}$

An additional expression for the overall structural mass fraction can be found by noting that the payload mass fraction ${\displaystyle \left({\frac {m_{\text{pl}}}{m_{i}}}\right)}$, propellant mass fraction and structural mass fraction sum to one:

${\displaystyle 1={\frac {m_{\text{pl}}}{m_{i}}}+{\frac {m_{p}}{m_{i}}}+{\frac {m_{s}}{m_{i}}}={\frac {m_{\text{pl}}}{m_{i}}}+\zeta +{\frac {m_{s}}{m_{i}}}}$

${\displaystyle {\frac {m_{s}}{m_{i}}}=1-\zeta -{\frac {m_{\text{pl}}}{m_{i}}}}$

Equating the expressions for structural mass fraction and solving for the initial vehicle mass yields:

${\displaystyle m_{i}=GLOW={\frac {m_{\text{pl}}}{1-\left({\frac {\zeta }{1-\lambda }}\right)}}}$

This expression shows how the size of a SSTO vehicle is dependent on its structural efficiency. Given a mission profile ${\displaystyle \left(\Delta v,m_{\text{pl}}\right)}$ and propellant type ${\displaystyle \left(I_{\text{sp}}\right)}$, the size of a vehicle increases with an increasing structural coefficient.[33] This growth factor sensitivity is shown parametrically for both SSTO and two-stage-to-orbit (TSTO) vehicles for a standard LEO mission.[34] The curves vertically asymptote at the maximum structural coefficient limit where mission criteria can no longer be met:

${\displaystyle \lambda _{\text{max}}=1-\zeta ={\frac {1}{MR}}}$

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,[35] reusable SSTO vehicles are typically an impractical choice even when using the highest performance propellants available.

案例

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.

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（110,000英磅） of payload, utilizing jet engines for (vertical) landing.[36] 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.

The Aquarius Launch Vehicle was designed to bring bulk materials to orbit as cheaply as possible.

廉价空间飞行的其他方法

Many studies have shown that regardless of selected technology, the most effective cost reduction technique is economies of scale. 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.

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.[44]

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.

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 their Falcon 9, Falcon Heavy, and Starship. A similar approach is being pursued by Blue Origin, using New Glenn.