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The Saturn II was a series of American expendable launch vehicles, studied by North American Aviation under NASA contract in 1966, derived from the Saturn V rocket used for the Apollo lunar program.[1] The intent of the study was to eliminate production of the Saturn IB, and create a lower-cost heavy launch vehicle based on Saturn V hardware. North American studied three versions with the S-IC first stage removed: the INT-17, a two-stage vehicle with a low Earth orbit payload capability of 47,000 pounds (21,000 kg); the INT-18, which added Titan UA1204 or UA1207 strap-on solid rocket boosters, with payloads ranging from 47,000 pounds (21,000 kg) to 146,400 pounds (66,400 kg); and the INT-19, using solid boosters derived from the Minuteman missile first stage.
For this study, the Boeing company also investigated configurations designated INT-20 and INT-21 which employed its S-IC first stage, and eliminated either North American's S-II second stage, or the Douglas S-IVB third stage. Budget constraints led to cancellation of the study and exclusive use of the Space Shuttle for orbital payloads.
Concept[modifica | modifica wikitesto]
There was a large payload gap between the Saturn IB's 46 000-libbra (21 000 kg) low Earth orbit capacity and the Saturn V's 310 000-libbra (140 000 kg) capability. In the mid-1960s NASA's Marshall Space Flight Center (MSFC) initiated several studies to extend the capabilities of the Saturn family. NASA specified a LEO of 100 miglia nautiche (185 km), 28° inclination for payload calculations, and the studies examined a number of Modified Launch Vehicle (MLV) configurations based on the Saturn IB and Saturn V launch vehicles as well as Intermediate Payload (INT) launch vehicles based on modified Saturn V stages (MS-IC, MS-II, and MS-IVB). Martin Marietta (builder of Atlas and Titan rockets), Boeing (builder of S-IC first stages), and North American Aviation (builder of the S-II second stage) were three of the companies that provided responses.
North American considered the best way to fill the gap was to use the Saturn V's second stage, the S-II, as the first stage of an intermediate launch vehicle. The basic concept of the Saturn II was to save money by ceasing production of the Saturn IB rocket, and replacing it with launch vehicles built entirely with current Saturn V components. This would allow closing down Chrysler Space Division production lines for the S-IB first stage, and would allow for more efficient integration of rocket systems.
Design[modifica | modifica wikitesto]
The baseline for the Saturn II was a Saturn V, without the Boeing-built S-IC first stage. The Saturn V's second stage S-II became the first stage, and the non-restartable S-IVB-200 used on the Saturn IB became the second stage. Such a vehicle could not fly without modification, because the S-II was designed to operate in the near-vacuum of high altitude space. Atmospheric thrust suppression reduced the five Rocketdyne J-2 engines' 1 000 000 libbre forza (4 400 kN) of vacuum thrust to 546 500 libbre forza (2 431 kN) at sea level,[1] insufficient to lift the 1 364 900-libbra (619 100 kg) weight of the two stages, even without a payload, off of the ground. This required that the S-II be either refit with higher thrust engines, augmented with solid rocket boosters, or both. Another design variable was the amount of the full 1 005 500 libbre (456 100 kg) propellant load carried in the S-II, and 241 300 libbre (109 500 kg) in the S-IVB stage.
Before any version could be put into production, work on all Saturn variants was stopped in favor of launching all future payloads from the Space Shuttle.
Saturn INT-17[modifica | modifica wikitesto]
The Saturn INT-17 was the first version of the Saturn II to be considered. It replaced the first stage's five J-2 engines with seven higher thrust HG-3-SL engines, giving 1 334 000 libbre forza (5 930 kN) of sea level thrust. It would burn a reduced S-II propellant load of 986 000 libbre (447 000 kg) in 200 seconds. The vehicle had a LEO payload capability of 92 000 libbre (42 000 kg) with a gross weight of 1 112 000 libbre (504 000 kg). The reduced payload permitted a savings of 660 libbre (300 kg) in structural weight, and omitting the S-IVB restart capability saved 1 500 libbre (700 kg).[2]
This configuration was dropped when it was determined that the HG-3-SL could not compete with the J-2 in terms of overall performance, reliability, and cost-effectiveness.[2] This required the addition of booster stages in order to provide more takeoff thrust.
Saturn INT-18[modifica | modifica wikitesto]
The Saturn INT-18 would have used the standard S-II with J-2 engines, augmented by two or four Titan SRBs. The UA1204 and UA1207 boosters were considered, with the highest total impulse configuration using four UA1207 boosters, capable of placing 146,000 libbre (66,224 kg) of payload into low Earth orbit. Designers considered changing the amount of fuel loaded into the rocket, and whether to ignite the S-II stage on the ground, or whether to launch using the solids, and start the main stage in flight. Two versions omitted the S-IVB stage.
The following configurations were studied:[3]
| Liftoff Mass | Boosters | S-II propellant | S-IVB propellant | Payload |
|---|---|---|---|---|
| 2 496 000 lb (1 132 000 kg) | 4 UA1204 | 474 900 lb (215 400 kg) | 177 000 lb (80 300 kg) | 47 000 lb (21 300 kg) |
| 2 496 000 lb (1 132 000 kg) | 4 UA1204 | 474 900 lb (215 400 kg) | 173 100 lb (78 500 kg) | 50 900 lb (23 100 kg) |
| 2 271 600 lb (1 030 400 kg) | 2 UA1207 | 560 000 lb (254 000 kg) | 177 900 lb (80 700 kg) | 60 400 lb (27 400 kg) |
| 2 496 500 lb (1 132 400 kg) | 2 UA1207 | 769 900 lb (349 200 kg) | 175 900 lb (79 800 kg) | 78 000 lb (35 400 kg) |
| 2 388 000 lb (1 083 000 kg) | 2 UA1205 | 951 500 lb (431 600 kg) | 170 600 lb (77 400 kg) | 89 300 lb (40 500 kg) |
| 3 462 400 lb (1 570 500 kg) | 4 UA1205 | 970 900 lb (440 400 kg) | 170 600 lb (77 400 kg) | 114 000 lb (51 700 kg) |
| 4 178 200 lb (1 895 200 kg) | 4 UA1207 | 984 800 lb (446 700 kg) | 166 900 lb (75 700 kg) | 146 400 lb (66 400 kg) |
| 3 254 500 lb (1 476 200 kg) | 4 UA1205 | 984 800 lb (446 700 kg) | No S-IVB | 86 000 lb (39 000 kg) |
| 3 923 300 lb (1 779 600 kg) | 4 UA1207 | 984 800 lb (446 700 kg) | No S-IVB | 97 000 lb (44 000 kg) |
Saturn INT-19[modifica | modifica wikitesto]
The Saturn INT-19 would have used smaller solid boosters, derived from the first stage of the Minuteman missile, to supplement the thrust of the S-II. Eleven configurations were studied, using between four and twelve solids, with some being started at lift-off, and some being started in flight, and varying propellant loads in the Saturn stages. The S-II stage would have been modified by refitting the J-2–SL engines with reduced expansion ratio nozzles, to increase sea level thrust to 174 400 libbre forza (776 kN) per engine. The highest total impulse configuration would have used twelve boosters, with eight started at launch and four started after the first group had been jettisoned. It would have been capable of lofting a payload of 75 400 libbre (34 200 kg).[4]
The following configurations were studied:[5]
| Liftoff mass | Boosters, liftoff | Boosters, round 1 | Boosters, round 2 | S-II propellant | S-IVB propellant | Payload |
|---|---|---|---|---|---|---|
| 723 800 libbre (328 300 kg) | 0 | 0 | 0 | 414 900 libbre (188 200 kg) | 170 000 libbre (77 100 kg) | 12 100 libbre (5 500 kg) |
| 1 021 800 libbre (463 500 kg) | 2 | 2 | 0 | 479 900 libbre (217 700 kg) | 177 000 libbre (80 300 kg) | 29 100 libbre (13 200 kg) |
| 1 277 800 libbre (579 600 kg) | 4 | 2 | 0 | 612 000 libbre (277 600 kg) | 168 900 libbre (76 600 kg) | 44 300 libbre (20 100 kg) |
| 1 277 800 libbre (579 600 kg) | 4 | 4 | 0 | 521 800 libbre (236 700 kg) | 161 000 libbre (73 000 kg) | 39 900 libbre (18 100 kg) |
| 1 593 700 libbre (722 900 kg) | 6 | 2 | 0 | 810 900 libbre (367 800 kg) | 168 900 libbre (76 600 kg) | 60 000 libbre (27 200 kg) |
| 1 593 700 libbre (722 900 kg) | 6 | 4 | 0 | 702 000 libbre (318 400 kg) | 172 000 libbre (78 000 kg) | 59 100 libbre (26 800 kg) |
| 1 618 600 libbre (734 200 kg) | 6 | 4 | 2 | 649 900 libbre (294 800 kg) | 179 000 libbre (81 200 kg) | 50 900 libbre (23 100 kg) |
| 1 593 700 libbre (722 900 kg) | 6 | 6 | 0 | 603 800 libbre (273 900 kg) | 173 900 libbre (78 900 kg) | 56 000 libbre (25 400 kg) |
| 1 910 700 libbre (866 700 kg) | 8 | 4 | 0 | 905 900 libbre (410 900 kg) | 177 900 libbre (80 700 kg) | 63 500 libbre (28 800 kg) |
| 1 910 700 libbre (866 700 kg) | 8 | 4 | 0 | 905 900 libbre (410 900 kg) | 166 900 libbre (75 700 kg) | 74 300 libbre (33 700 kg) |
| 1 910 700 libbre (866 700 kg) | 8 | 4 | 0 | 905 900 libbre (410 900 kg) | 165 800 libbre (75 200 kg) | 75 400 libbre (34 200 kg) |
See also[modifica | modifica wikitesto]
Notes[modifica | modifica wikitesto]
- ^ "J-2", Astronautix
- ^ a b "Saturn INT-17",Encyclopedia Astronautica
- ^ "Saturn INT-18", Astronautix Archiviato il 24 settembre 2011 in Internet Archive.
- ^ Errore nelle note: Errore nell'uso del marcatore
<ref>: non è stato indicato alcun testo per il marcatoreD5-13183 - ^ "Saturn INT-19", Astronautix