Одноступенчатые к орбите (или SSTO ) транспортного средства достигает орбиты от поверхности тела с использованием только ракетного топлива и жидкости и без расходуя резервуаров, двигателей, или других крупных аппаратных средств. Этот термин обычно, но не исключительно, относится к многоразовым транспортным средствам . [1] На сегодняшний день запускаемые с Земли ракеты-носители SSTO никогда не запускались; орбитальные запуски с Земли выполнялись полностью или частично одноразовыми многоступенчатыми ракетами .
Основное предполагаемое преимущество концепции SSTO - исключение аппаратной замены, присущей одноразовым пусковым системам. Однако единовременные расходы, связанные с проектированием, разработкой, исследованиями и проектированием (DDR & E) многоразовых систем SSTO, намного выше, чем у расходных систем, из-за существенных технических проблем, связанных с SSTO, при условии, что эти технические проблемы действительно могут быть решены. [2]
Считается маловероятным запуск с Земли одноступенчатого орбитального космического корабля на химическом топливе . Основными осложняющими факторами для SSTO с Земли являются: высокая орбитальная скорость более 7 400 метров в секунду (27 000 км / ч; 17 000 миль в час); необходимость преодоления земного притяжения, особенно на ранних этапах полета; и полет в атмосфере Земли , который ограничивает скорость на ранних этапах полета и влияет на характеристики двигателя. [ необходима цитата ]
Достижения в области ракетостроения в 21 - м веке привели к существенному снижению стоимости для запуска килограмма полезного груза либо низкой околоземной орбиту или Международной космической станции , [3] уменьшение основного прогнозируемого преимущества концепции SSTO.
Известные концепции одноступенчатого вывода на орбиту включают Skylon , DC-X , Lockheed Martin X-33 и Roton SSTO . Однако, несмотря на некоторые обещания, ни один из них еще не приблизился к достижению орбиты из-за проблем с поиском достаточно эффективной двигательной установки. [1]
Одноступенчатый к орбите гораздо легче достичь на внеземных тел , которые имеют более слабые гравитационные поля и снизить атмосферное давление , чем Земли, таких как Луна и Марс, и было достигнуто с Луны по программе Apollo «s Lunar Module , несколькими космическими роботами советской программы «Луна» и китайским « Чанъэ 5» .
История
Ранние концепции
До второй половины двадцатого века космические путешествия проводились очень мало. В течение 1960-х годов начали появляться некоторые из первых концептуальных проектов такого рода судов. [4]
Одна из самых ранних концепций SSTO был расходным Одноэтапным Orbital Space Truck (OOST) , предложенный Филипп Боно , [5] инженер для Douglas Aircraft Company . [6] Также была предложена многоразовая версия под названием ROOST.
Другой ранней концепцией SSTO была многоразовая ракета-носитель под названием NEXUS, предложенная Краффтом Арнольдом Эрике в начале 1960-х годов. Это был один из крупнейших космических кораблей, когда-либо созданных, с диаметром более 50 метров и способностью поднимать до 2000 коротких тонн на околоземную орбиту, предназначенный для миссий в более отдаленные районы Солнечной системы, такие как Марс . [7] [8]
Североамериканский воздуха дополненной VTOVL с 1963 был аналогичным большим корабль , который использовал бы ПВРД по снижению Liftoff массы транспортного средства за счет устранения необходимости в больших количествах жидкий кислород при путешествии через атмосферу. [9]
С 1965 года Роберт Салкельд исследовал различные концепции одноступенчатых крылатых космических самолетов. Он предложил транспортное средство, которое будет сжигать углеводородное топливо, находясь в атмосфере, а затем переключаться на водородное топливо для повышения эффективности в космосе. [10] [11] [12]
Другие примеры ранних концепций Боно (до 1990-х годов), которые так и не были созданы, включают:
- ROMBUS (многоразовый орбитальный модуль, ускоритель и служебный шаттл), еще одна разработка Филипа Боно. [13] [14] Технически это не была одноступенчатая установка, поскольку она сбросила некоторые из своих первоначальных резервуаров с водородом, но она подошла очень близко.
- Ithacus, адаптированная концепция ROMBUS, которая была разработана для перевозки солдат и военной техники на другие континенты по суборбитальной траектории. [15] [16]
- Pegasus, еще одна адаптированная концепция ROMBUS, предназначенная для перевозки пассажиров и грузов на большие расстояния за короткие промежутки времени в космосе. [17]
- Дуглас САССТО , концепция ракеты-носителя 1967 года. [18]
- Hyperion, еще один концепт Филипа Боно, в котором использовались сани для набора скорости перед взлетом, чтобы сэкономить количество топлива, которое нужно было поднять в воздух. [19]
Star-raker : в 1979 году Rockwell International представила концепцию 100-тонного многоциклового воздушно-реактивного прямоточного воздушно- реактивного двигателя большой грузоподъемности / криогенного ракетного двигателя с одноступенчатым выводом на орбиту горизонтального взлета / горизонтальной посадки под названием Star-Raker , предназначенного для запуска тяжелых космических аппаратов. спутников на солнечной энергии на орбиту Земли 300 морских миль. [20] [21] [22] Star-raker должен был иметь 3 ракетных двигателя LOX / LH2 (на основе SSME ) + 10 турбореактивных двигателей. [20]
Примерно в 1985 году проект NASP был предназначен для запуска на орбиту ГПВРД, но финансирование было прекращено, и проект был отменен. [23] Примерно в то же время HOTOL попытался использовать технологию реактивного двигателя с предварительным охлаждением , но не смог показать существенных преимуществ перед ракетной технологией. [24]
Технология DC-X
DC-X, сокращение от Delta Clipper Experimental, представлял собой беспилотный демонстратор вертикального взлета и посадки в масштабе одной трети для предлагаемого SSTO. Это один из немногих когда-либо построенных прототипов автомобилей SSTO. Планировалось несколько других прототипов, в том числе DC-X2 (полуразмерный прототип) и DC-Y, полномасштабный аппарат, который можно было бы одноступенчато выводить на орбиту. Ни один из них не был построен, но проект был передан НАСА в 1995 году, и они построили DC-XA, модернизированный прототип в масштабе одной трети. Этот автомобиль был потерян, когда приземлился с развернутыми только тремя из четырех посадочных площадок, в результате чего он перевернулся на бок и взорвался. С тех пор проект не получил продолжения. [ необходима цитата ]
Ротон
С 1999 по 2001 год Rotary Rocket пыталась построить автомобиль SSTO под названием Roton. Он привлек большое внимание средств массовой информации, и был завершен рабочий прототип, но его конструкция была в значительной степени непрактичной. [25]
Подходы
Существуют различные подходы к SSTO, в том числе чистые ракеты, которые запускаются и приземляются вертикально, воздушно- реактивные ГРП, которые запускаются и приземляются горизонтально, аппараты с ядерными двигателями и даже аппараты с реактивными двигателями, которые могут летать на орбиту. и возвратная посадка, как у авиалайнера, полностью цела.
Для SSTO с ракетными двигателями основная задача заключается в достижении достаточно высокого отношения масс, чтобы нести достаточно топлива для выхода на орбиту , а также значительный вес полезной нагрузки . Одна из возможностей - придать ракете начальную скорость с помощью космической пушки , как и планировалось в проекте Quicklaunch . [26]
Для SSTO с воздушным дыханием основной проблемой является сложность системы и связанные с ней затраты на исследования и разработки , материаловедение и методы строительства, необходимые для выживания в длительном высокоскоростном полете в атмосфере и достижения достаточно высокого отношения масс, чтобы нести достаточное количество топлива. достичь орбиты, плюс значительный вес полезной нагрузки. Конструкции с воздушным дыханием обычно летают на сверхзвуковых или гиперзвуковых скоростях и обычно включают в себя ракетный двигатель для окончательного выхода на орбиту. [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.[27] It has a sufficiently efficient engine to achieve orbit, but without carrying much payload.[28]
Dense versus hydrogen fuels
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 end 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.[29]
In the 1960s Philip Bono investigated single-stage, VTVL tripropellant rockets, and showed that it could improve payload size by around 30%.[30]
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
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 optimal. Engines operating in the lower atmosphere have shorter bells than those designed to work in vacuum. Having a bell that is only optimal at a single altitude lowers the overall engine efficiency.[citation needed]
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.[citation needed]
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 ] 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
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.[citation needed]
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).[31] 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.[32]
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.[33]
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
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]
- sled launch (rail, maglev including Bantam, MagLifter, and StarTram, etc.)[34]
- air launch or aircraft tow
- in-flight fueling
- Lofstrom launch loop/space fountains
And on-orbit resources such as:[citation needed]
- Space tether
- tugs
Nuclear propulsion
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.[citation needed]
Beam-powered propulsion
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]
Проблемы проектирования, присущие SSTO
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.[35]
The Tsiolkovsky rocket equation expresses the maximum change in velocity any single rocket stage can achieve:
where:
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:
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.[36] 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.[37] This growth factor sensitivity is shown parametrically for both SSTO and two-stage-to-orbit (TSTO) vehicles for a standard LEO mission.[38] 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,[39] reusable SSTO vehicles are typically an impractical choice even when using the highest performance propellants available.
Примеры
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.[citation needed]
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.[40] 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
Current and previous SSTO projects include the Japanese Kankoh-maru project, ARCA Haas 2C, and the Indian Avatar spaceplane.[citation needed]
Skylon
The British Government partnered with the ESA in 2010 to promote a single-stage to orbit spaceplane concept called Skylon.[41] This design was pioneered by Reaction Engines Limited (REL),[42][43] a company founded by Alan Bond after HOTOL was canceled.[44] The Skylon spaceplane has been positively received by the British government, and the British Interplanetary Society.[45] 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.[46] 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.[46][47]
Альтернативные подходы к недорогим космическим полетам
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.[48]
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 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.
Смотрите также
- Aerospike engine
- Bristol Spaceplanes
- British Aerospace HOTOL
- Kankoh-maru
- Launch loop
- Lockheed Martin X-33
- Mass fraction
- NASA X-43
- Orbital ring
- Rockwell X-30
- Roton
- Scramjet
- Space elevator
- Spacecraft propulsion
- Three-stage-to-orbit
- Two-stage-to-orbit
- VentureStar
- XS-1 (spacecraft)
дальнейшее чтение
- 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.
Рекомендации
- ^ a b c Richard Varvill & Alan Bond (2003). "A Comparison of Propulsion Concepts for SSTO Reusable Launchers" (PDF). JBIS. Archived from the original (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.
- ^ Harry W. Jones (2018). "The Recent Large Reduction in Space Launch Cost" (PDF). ICES. Retrieved 12 December 2018.
- ^ Gomersall, Edward (20 July 1970). A Single Stage To Orbit Shuttle Concept. Ames Mission Analysis Division Office of Advanced Research and Technology: NASA. p. 54. N93-71495.
- ^ Philip Bono and Kenneth William Gatland, Frontiers of Space, ISBN 0-7137-3504-X
- ^ Wade, Mark. "OOST". Encyclopedia Astronautica. Archived from the original on 10 October 2011. Retrieved 18 October 2015.
- ^ "Aerospace projects Review". 3 (1). Cite journal requires
|journal=
(help) - ^ "SP-4221 The Space Shuttle Decision". NASA History. Retrieved 18 October 2015.
- ^ "Encyclopedia Astronautica - North American Air Augmented VTOVL". Retrieved 18 October 2015.
- ^ "Salkeld Shuttle". astronautix.com. Retrieved 13 June 2015.
- ^ "ROBERT SALKELD'S". pmview.com. Retrieved 13 June 2015.
- ^ "STS-1 Further Reading". nasa.gov. Retrieved 13 June 2015.
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Внешние ссылки
- A Single-Stage-to-Orbit Thought Experiment
- Why are launch costs so high?, an analysis of space launch costs, with a section critiquing SSTO
- The Cold Equations Of Spaceflight A critique of SSTO by Jeffrey F. Bell.
- Burnout Velocity Vb of a Single 1-Stage Rocket