User:WDGraham/List of unflown orbital launch systems

Throughout the history of spaceflight many orbital launch systems have been proposed, designed and entered development, only to be cancelled before development was completed, without making any orbital launch attempts. Some launch systems were concept designs or studies, or proposals which were never taken further, whilst others such as the Ares I reached the stage of hardware construction and partial flight tests before cancellation.

This list contains orbital launch systems which were cancelled before making an orbital launch attempt. It includes rockets which made atmospheric and suborbital tests prior to cancellation, such as Ares I, however it does not include rockets cancelled after failed orbital launch attempts, such as the N1, Europa and Pilot.

A

• Ares I - United States. NASA design to launch Orion as part of Project Constellation, first stage to have been built by ATK, second by Boeing. Capable of placing 25.6 tonnes (25.2 long tons; 28.2 short tons) of payload into low Earth orbit. Designed by Exploration Systems Architecture Study in 2005, cancelled in 2010 along with Constellation programme
• Ares IV - United States. Proposed alternative to Ares I using the same first stage as Ares V. Capable of placing 40 tonnes (39 long tons; 44 short tons) to TLI.
• Ares V - United States. NASA design to launch Altair and Earth Departure Stage as part of Project Constellation, boosters to have been built by ATK. Capable of placing 160 to 188 tonnes (157 to 185 long tons; 176 to 207 short tons) into low Earth orbit and 63 to 71 tonnes (62 to 70 long tons; 69 to 78 short tons) to TLI. Designed by Exploration Systems Architecture Study in 2005, cancelled in 2010 along with Constellation programme
• Ares V Lite - United States. Proposed unified launch system for Project Constellation, combining the roles of the Ares I and Ares V into two launches of a single vehicle. Payload of 140 tonnes (140 long tons; 150 short tons) to Low Earth Orbit.

F

• Falcon 5 - United States. SpaceX rocket with five first stage engines but otherwise identical to the Falcon 9. Cancelled in 2007
• Falcon 9S5 - United States. SpaceX rocket consisting of a Falcon 9 with two Falcon 5 first stages being used as boosters. Cancelled along with Falcon 5 in 2007
• Falcon X - United States. SpaceX concept design capable of delivering 38 tonnes (37 long tons; 42 short tons) into low Earth orbit. Heavy configuration capable of 125 tonnes (123 long tons; 138 short tons) to LEO.
• Falcon XX - United States. SpaceX concept design capable of delivering 140 tonnes (140 long tons; 150 short tons) into low Earth orbit.

J

• Jupiter - United States. Designed in 2006 as part of the DIRECT programme, an unofficial NASA design study which proposed alternatives to the Ares family of rockets for Project Constellation (started in 2006) to replace Space Shuttle. Many variants were proposed, such as the Jupiter-130 and Jupiter-246, with claimed lift capacities exceeding 60 and 90 tonnes (59 and 89 long tons; 66 and 99 short tons) to LEO respectively; increasing to 100 tonnes (98 long tons; 110 short tons) with an upper stage, and 120 tonnes (120 long tons; 130 short tons) with five-segment solid rocket boosters as well.

M

• Magnum - United States. Designed by NASA between 1996 and 2004, Magnum was intended to launch a manned mission to Mars, with a payload capacity of 55 to 94 tonnes (54 to 93 long tons; 61 to 104 short tons) to Low Earth orbit, with 80 tonnes (79 long tons; 88 short tons). The rocket never made it past the preliminary design phase.

• MLV-SDV-1a: 55 tonnes (54 long tons; 61 short tons) of payload to a 407 kilometres (253 mi) circular low Earth orbit inclined at 28.5 degrees. Two reusable RSRM boosters on first stage.
• MLV-SDV-1b: 94 tonnes (93 long tons; 104 short tons) of payload to a 407 kilometres (253 mi) circular low Earth orbit inclined at 28.5 degrees. Four reusable RSRM boosters on first stage, core stage air-lit 100 seconds after launch.
• MLV-SDV-2: 80 tonnes (79 long tons; 88 short tons) of payload to a 407 kilometres (253 mi) circular low Earth orbit inclined at 28.5 degrees. First stage augmented by two reusable RSRM boosters and two liquid fuelled boosters each powered by two SSMEs.
• MLV-SDV-3: 91 tonnes (90 long tons; 100 short tons) of payload to a 407 kilometres (253 mi) circular low Earth orbit inclined at 28.5 degrees. First stage augmented by two liquid-fuelled boosters each powered by three RD-180 engines.
• MLV-SDV-4: 64 tonnes (63 long tons; 71 short tons) of payload to a 407 kilometres (253 mi) circular low Earth orbit inclined at 28.5 degrees. First stage augmented by two liquid-fuelled boosters each powered by four unspecified engines burning RP-1 and liquid oxygen.
• MLV-LFBB: 93 tonnes (92 long tons; 103 short tons) of payload to a 407 kilometres (253 mi) circular low Earth orbit inclined at 28.5 degrees. First stage augmented by two reusable fly-back boosters powered by RD-180 engines.

N

• N1-MOK - Soviet Union. Cryogenically-fuelled single-stage to orbit rocket based on the N1 first stage. Proposed in 1974, it would have had a payload capacity of 90 tonnes (89 long tons; 99 short tons) into a circular low Earth orbit at an altitude of 450 kilometres (280 mi) and an inclination of 97.5 degrees.
• N1 Nuclear A - Soviet Union. A variant of the N1 studied by Sergei Korolev in 1963, the N1 Nuclear A would have used a third stage powered by a nuclear thermal engine, and been capable of placing 270 tonnes (270 long tons; 300 short tons) of payload into a circular low Earth orbit at an altitude of 220 kilometres (140 mi), and an inclination of 51.6 degrees.
• N1 Nuclear AF - Soviet Union. Derivative of the proposed N1 Nuclear A, which would have used less efficient engines delivering higher thrust, and could have placed 300 tonnes (300 long tons; 330 short tons) of payload into a circular low Earth orbit at an altitude of 220 kilometres (140 mi), and an inclination of 51.6 degrees.
• N1 Nuclear V - Soviet Union. Derivative of the N1 rocket with nuclear-powered second and third stages, capable of placing 420 tonnes (410 long tons; 460 short tons) of payload into a circular low Earth orbit at an altitude of 220 kilometres (140 mi), and an inclination of 51.6 degrees.
• N1 Nuclear V-B - Soviet Union. A variant of the N1 Nuclear V, the N1 Nuclear V-B would have had additional radiation shielding, allowing it to launch manned spacecraft. It could have placed 360 tonnes (350 long tons; 400 short tons) into a circular low Earth orbit at an altitude of 220 kilometres (140 mi), and an inclination of 51.6 degrees.
• N1F - Soviet Union. An N1 derivative with upgraded engines, larger second and third stages, and incorporating changes made as a result of the four launch failures between 1969 and 1972. The 1965 design would have been capable of placing 100 tonnes (98 long tons; 110 short tons) of payload into a circular low Earth orbit at an altitude of 220 kilometres (140 mi) and an inclination of 51.6 degrees. By the time development was cancelled in 1975, this had increased to 105 tonnes (103 long tons; 116 short tons). Use of Blok S and Block R upper stages was also considered for launching L3M lunar missions.

• N1FV-II-III: Proposed variant with cryogenic second and third stages capable of placing 150 tonnes (150 long tons; 170 short tons) into a circular low Earth orbit at an altitude of 220 kilometres (140 mi) and an inclination of 51.6 degrees.
• N1FV-III: Proposed variant with Blok V-III third stage, capable of placing 125 tonnes (123 long tons; 138 short tons) into a circular low Earth orbit at an altitude of 220 kilometres (140 mi) and an inclination of 51.6 degrees.
• N1F/Sr: A derivative of the N1F, with a Blok Sr cryogenic upper stage in place of the N1F's Blok G and Blok D fourth and fifth stages.

• N1M - Soviet Union. Designed in 1965 as the heaviest N1 derivative, with the payload capacity of the first stage doubled. It would have been capable of placing 155 tonnes (153 long tons; 171 short tons) into a circular low Earth orbit at an altitude of 220 kilometres (140 mi), and an inclination of 51.6 degrees, subsequently redesigned a standard first stage and cryogenic upper stages to launch the LEK spacecraft on missions to the Moon. Development abandoned in 1971.

• N1MV-II-III: Proposed variant using cryogenic second and third stages to deliver a payload of 230 tonnes (230 long tons; 250 short tons) to a circular low Earth orbit at an altitude of 220 kilometres (140 mi) and an inclination of 51.6 degrees.
• N1MV-III: Proposed variant with Blok V-III third stage, capable of placing 185 tonnes (182 long tons; 204 short tons) into a circular low Earth orbit at an altitude of 220 kilometres (140 mi) and an inclination of 51.6 degrees.

• N1U - Soviet Union. Proposed production version of the N1 with improvements to increase reliability, particularly its engines. Capable of placing 95 tonnes (93 long tons; 105 short tons) into a circular low Earth orbit at an altitude of 220 kilometres (140 mi) and an inclination of 51.6 degrees.
  • N1UV-III: Derivative with Blok V-III third stage, capable of placing 115 tonnes (113 long tons; 127 short tons) into a circular low Earth orbit at an altitude of 220 kilometres (140 mi) and an inclination of 51.6 degrees.
• Nova - United States. Family of super-heavy launch systems studied by NASA for manned missions beyond Earth orbit. Some Nova-rockets, including the Nova C-8 and Nova 8L, were intended for direct ascent lunar missions, which were cancelled in favour of lunar orbit rendezvous with the Saturn V rocket. These rockets had a payload capacity of 24 to 75 tonnes (24 to 74 long tons; 26 to 83 short tons) tons to TLI, compared to the Saturn V's capacity of 45 tonnes (44 long tons; 50 short tons). Larger variants designed for missions to Mars could have placed up to 455 tonnes (448 long tons; 502 short tons) into Low Earth orbit.

S

• Saturn V/4-260 - United States. Late 1960s Saturn V derivative studied by Boeing; using four 6.6-metre (260 in) diameter solid rocket motors, which would also have incorporated liquid propellant tanks for the first stage. Capable of placing 362 tonnes (356 long tons; 399 short tons) of payload into a circular low Earth orbit at an altitude of 426 kilometres (265 mi) and an inclination of 28 degrees.
• Saturn V-23(L) - United States. Saturn V derivative studied in 1967 by Boeing, with stretched first and third stages, and four liquid fuelled boosters, each powered by two F-1 engines. Capable of placing 262.67 tonnes (258.52 long tons; 289.54 short tons) into a circular low Earth orbit at an altitude of 185 kilometres (115 mi) and an inclination of 28 degrees.
• Saturn V-24(L) - United States. Saturn V derivative studied in 1967 by Boeing, with all three stages stretched, first stage engines replaced with F-1As and second and third stage engines replaced with HG-3s. Augmented by four F-1A-powered liquid boosters. Capable of delivering 435.3 tonnes (428.4 long tons; 479.8 short tons) of payload into a circular low Earth orbit at an altitude of 185 kilometres (115 mi) and an inclination of 28 degrees.
• Saturn V-25(S)B - United States. Saturn V derivative studied in 1967 by Boeing, with stretched first and third stages, and four 3.96-metre (156 in) solid rocket motors. Capable of placing 223.5 tonnes (220.0 long tons; 246.4 short tons) into a circular low Earth orbit at an altitude of 185 kilometres (115 mi) and an inclination of 28 degrees.

• Saturn V-4X(U): A derivative using four clustered first and second stages to give a payload capacity of 527 tonnes (519 long tons; 581 short tons) to
• Saturn V-25(S)U: Derivative with NERVA nuclear-powered third stage, intended for assembling spacecraft in Earth orbit capable of sending men to Mars. Payload capacity of 248.663 tonnes (244.736 long tons; 274.104 short tons) to a circular low Earth orbit at an altitude of 426 kilometres (265 mi) and an inclination of 28 degrees, or 160 tonnes (160 long tons; 180 short tons) to TLI.

• Sea Dragon - United States. A two-stage reusable sea-launched rocket studied in 1962, which would have been capable of placing 550 tonnes (540 long tons; 610 short tons) into LEO.
• SERV - United States. Designed by Chrysler as their proposal for the Space Transportation System in the 1970s, SERV would have used the existing Saturn V launch complex, and could be used to launch the MURP spaceplane. It was capable of placing 11 to 57 tonnes (11 to 56 long tons; 12 to 63 short tons) into low Earth orbit, depending on the configuration, and whether MURP was being carried.
• Shuttle-C - United States. Proposed Shuttle-Derived Launch Vehicle which would have used then-existing Shuttle-components and infrastructure to place an unmanned payload of 60 to 80 tonnes (59 to 79 long tons; 66 to 88 short tons) into orbit. The configuration consisted of a Space Shuttle external tank and set of solid rocket boosters, but with an SSME-powered unmanned cargo vehicle in place of the orbiter.
• Shuttle-Derived Heavy Lift Launch Vehicle - United States. Shuttle-Derived Launch Vehicle proposed as an alternative launch system for Project Constellation, capable of placing up to 104 tonnes (102 long tons; 115 short tons) of cargo into low Earth orbit. Other configurations would have been able to place 66 to 83 tonnes (65 to 82 long tons; 73 to 91 short tons) into LEO.
• Superraket - Soviet Union. 1950s study of a rocket incorporating a nuclear-powered second stage. Capable of placing 150 tonnes (150 long tons; 170 short tons) into low Earth orbit at an altitude of 185 kilometres (115 mi).

U

• UR-700 - Soviet Union. Designed by Vladimir Chelomey in the 1960s, the UR-700 was based on the UR-500 rocket and designed for manned lunar missions using the direct ascent method. The N1 was developed instead, using the lunar orbit rendezvous method. UR-700 development continued until 1968 with different variants considered, with payloads between 70 and 175 tonnes (69 and 172 long tons; 77 and 193 short tons) to LEO (151 tonnes (149 long tons; 166 short tons) for the original design). Using an 11D54 upper stage, it would have been able to deliver 185 tonnes (182 long tons; 204 short tons) to a circular low Earth orbit at 200 kilometres (120 mi) altitude and 51.5° inclination. Using an upper stage powered by three RD-350 engines this could have been increased to 215 tonnes (212 long tons; 237 short tons) to an orbit at 200 kilometres (120 mi) and 51.0° inclination, With an upper stage powered by seven RO-31 engines it could have placed 230 to 270 tonnes (230 to 270 long tons; 250 to 300 short tons) into low Earth orbit.
• UR-700M - Soviet Union. Designed by Vladimir Chelomey in 1969 for Project Aelita, a Soviet programme to land a man on Mars in 1969. The design was based on the UR-700 rocket, and had a payload capacity of 750 tonnes (740 long tons; 830 short tons) to a circular low Earth orbit at 200 kilometres (120 mi) altitude and 51.0° inclination.
• UR-900 - Soviet Union. Derivative of the UR-700 designed by Vladimir Chelomey, and proposed in 1969 for manned missions to Mars. Payload capacity of 240 tonnes (240 long tons; 260 short tons) to a circular low Earth orbit at 200 kilometres (120 mi) altitude and 51.0° inclination. Cancelled after a Proton launch failure raised concerns over the safety of its hypergolic propellants

V

• Vulkan-Herkules - Soviet Union. Energia derivative capable of placing 200 tonnes (200 long tons; 220 short tons) of payload into low Earth orbit.

Unsorted

• Saturn V-A, a NASA-study in 1968 essentially identical to Saturn INT-20. Payload to LEO 60,000 kg.
• Saturn V-C, a NASA-study in 1968 extending the Saturn V-A and Saturn INT-20-studies. Payload to LEO 81,600 kg.
• Saturn V-Centaur, another NASA-study in 1968 extending the Saturn V-A and Saturn INT-20-studies. Payload to LEO 118,000 kg.
• Saturn V-D, a NASA-study of 1968 extending the Boeing-study of 1967 to develop a Saturn V-based rocket. Payload to LEO 326,500 kg.
• Saturn V ELV, a NASA study of 1966 to develop Saturn V-based rocket. Payload to LEO 200,000 kg.
• Saturn C-3B, a launcher studied in the USA in 1961. Cancelled after the Saturn C-5 was selected for Apollo program(Saturn C-5-rocket later evolved into Saturn V). Would have had the payload capacity of 78,000 kg to LEO.
• Saturn C-3BN, a launcher studied in the USA in 1961. Cancelled after the Saturn C-5 was selected for Apollo program; the Saturn V-rocket was based in the Saturn C-5. Otherwise similar to Saturn C-3B, but would have utilized a nuclear upper stage. Payload capacity of 94,000 kg to LEO.
• Saturn C-4, an American launch vehicle studied for the lunar orbit rendezvous-method of lunar exploration. Lost competition for the launcher of the Apollo program to Saturn C-5 (Saturn C-5 was modified slightly during the 1960s to produce the Saturn V-rocket) because Saturn C-5 had reserve capacity that the Moon mission designers wanted. Payload to LEO 99,000 kg.
• Saturn C-4B, the last variant of Saturn C-4 before Saturn C-5 was chosen for the Moon landing in 1961 (Saturn C-5 was modified slightly during the 1960s to produce the Saturn V-rocket) and the development of other Saturn C-series rockets was halted. Payload 95,000 kg to LEO.
• Saturn C-5, the rocket that was chosen for the Apollo program in 1961. Saturn C-5's development was continued after it was chosen to be the American Moon rocket, and the result was Saturn V. The difference between Saturn C-5 and Saturn V (albeit small) is that the upper stages of Saturn V were enlarged in relation to the C-5. The Saturn C-5-configuration of 1961 had payload capacity to LEO 120,000 kg.
• Saturn C-5N, was a conceptual version of the Saturn V launch vehicle which would have had a nuclear third stage. Payload to LEO 155,000 kg.
• Saturn C-8, the largest of Saturn-variants to be considered. Was intended for direct landing method of lunar exploration, like the Nova's. Was abandoned after the Saturn C-5 was selected for Apollo program(Saturn C-5 developed into the Saturn V). Payload to LEO 210 ton.
• Saturn INT-18, a conceptual study in 1966 to build a rocket utilizing various Saturn V-components. Numerous version were studied, with payload capacity between 21,300 - 66,590 kg to LEO (two heaviest variant had payload capacities of 51,700 kg and 66,400-66,590 kg to LEO.)
• Saturn INT-20, a proposed launcher in the 1960s-1970s using the Saturn V-components. Three variants were studied with the heaviest (the five-engine variant) having payload capacity of 60,500 kg to LEO, and the second heaviest (the four-engine variant) having payload capacity of 60,000 kg to LEO.
• Saturn INT-21, described in a study of the 1970s to develop a smaller launcher based on Saturn V. It was expected to be composed of Boeing S-IC and modified North American S-II with payload capacity of 75,000 kg to LEO. Also heavier variants with payload to LEO 84,000 kg, 89,000 kg, 101,000 kg, 112,000 kg and 116,000 kg were studied (the heavier variants had successively more engines).
• Saturn MLV-V-1, a NASA-study of an improved Saturn V-rocket in 1965. Payload to LEO 137,250 kg.
• Saturn MLV-V-1A, a NASA-study of an improved Saturn V-rocket in 1965. Payload to LEO 145,000 kg.
• Saturn MLV-V-2, a NASA-study to develop the Saturn V-rocket in 1965. Payload to LEO 137,250 kg
• Saturn MLV-V-3, a NASA-study in 1965 to improve the Saturn V. Payload to LEO 160,400 kg

• Saturn V-3B, a Boeing-studied variation of a Saturn MLV-V-3-study. Studied in 1967. Rocket was based in the Saturn V-rocket. Payload capacity 166,600 kg to LEO.

• Saturn MLV-V-4(S), rocket NASA studied in 1965. A developed version of Saturn V. Payload 118,000 kg to LEO.
• Saturn MLV-V-4(S)-A, a development of the Saturn V studied by NASA in 1965. Payload to LEO 160,880 kg.
• Saturn MLV-V-4(S)-B, a rocket studied by Boeing in 1967. Was based on Saturn V. Payload to LEO 171,990 kg.