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THE SATURN FAMILY
Saturn I: (1961) Originally known as Juno V. "Cluster's last stand" - 8 Redstone tanks
around Jupiter tank core, eight Jupiter engines. Intended as launch vehicle for Apollo CM
circumlunar flights. Developed so early, no payloads available for it.
Saturn IB: (1966) Improved Saturn I, with uprated first stage and Saturn IVB stage
(common with Saturn V) replacing Saturn IV. Used for earth orbit flight tests of Apollo CSM.
Saturn V: (1967) Lunar landing booster. Design frozen before landing mode selected;
could be used for either EOR or LOR methods. Size dictated by ceiling height at Michoud factory
selected for first stage manufacture.
SATURN FAMILY RECORD
First launch: 27-Oct-1961
Last launch: 15-Jul-1975
Number launched: 33
Launch sites: Cape Canaveral pads 34, 37 & 39A
Vehicle success rate: 97.0%
SATURN I SPECIFICATIONS
First launch: 27-Oct-1961
Last launch: 30-Jul-1965
Number launched: 10
Launch sites: Cape Canaveral pad 37
Principal uses: intended as launch vehicle for Apollo CM circumlunar flights
Vehicle success rate: 100.0%
Performance: LEO (185 km, 28.5o): 9,000 kg; translunar trajectory:
2,200 kg
Cost: $107.3 million
Number of stages: 3
Overall length: 45.8 m
Principal diameter: 6.5 m
Launch mass: 498,900 kg
Launch thrust: 6,690 kN sea level
SATURN I STAGE 1
Designation: Saturn 1
Engines: eight Rocketdyne H-1 liquid bipropellant; four outboard engines are
gimbaled for steering
Length: 24.5 m
Diameter: 6.5 m
Dry mass: 45,270 kg
Oxidizer: liquid oxygen
Fuel: RP-1 hydrocarbon
Propellant mass: 387,400 kg
Thrust: 6,690 kN SL; 7,582 kN vac
Burn time: 150 s
SATURN I STAGE 2
Designation: Saturn IV
Engines: six P&W RL-10 liquid bipropellant
Length: 12.2 m
Diameter: 5.5 m
Dry mass: 5,217 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 45,360 kg
Thrust: 400.2 kN vac
Burn time: 482 s
SATURN I STAGE 3
Designation: Centaur C
Engines: two P&W RL-10A-1 liquid bipropellant
Overall length: 9.1 m
Principal diameter: 3.1 m
Dry mass: 1,996 kg dry
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 13,600 kg
Thrust: 133.4 kN vacuum
Burn time: 430 s
SATURN IB SPECIFICATIONS
First launch: 26-Feb-1966
Last launch: 15-Jul-1975
Number launched: 10
Launch sites: Cape Canaveral pads 34, 37 & 39A
Principal uses: earth orbit flights of Apollo CSM
Vehicle success rate: 90.0%
Performance: LEO (185 km, 28.5o): 18,600 kg
Cost: $101.8 million
Number of stages: 2
Overall length: 42.3 m
Principal diameter: 6.6 m
Launch mass: 568,600 kg
Launch thrust: 7,295 kN sea level
SATURN IB STAGE 1
Designation: Saturn 1B
Engines: eight Rocketdyne H-1B liquid bipropellant; four outboard engines are
gimbaled for steering
Length: 24.5 m
Diameter: 6.5 m
Dry mass: 41,590 kg
Oxidizer: liquid oxygen
Fuel: RP-1 hydrocarbon
Propellant mass: 407,100 kg
Thrust: 7,295 kN SL; 8,242 kN vac
Burn time: 155 s
SATURN IB STAGE 2
Saturn IB S-IVB stages and instrument unit (serial numbers beginning with 200) were nearly
identical in appearance to those in the Saturn V (serial numbers beginning with 500). On Saturn
IB, the J-2 engine was configured for a single-use capability; externally, the two Auxiliary
Propulsion System (APS) pods were smaller than those fitted on the Saturn V. Three ullage
rockets were fitted on the base of the 200-series S-IVB, only two were fitted on the 500-series
S-IVBs. The 200-series instrument units deleted instrumentation normally carried for monitoring
and controlling Saturn V systems.
Designation: Saturn IVB
Engines: one gimbaled Rocketdyne J-2 liquid bipropellant
Length: 17.8 m
Diameter: 6.6 m
Dry mass: 13,310 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 106,600 kg
Thrust: 1,001 kN vac
Burn time: 475 s
Instrument unit: length 0.91 m; diameter 6.6 m
SATURN V SPECIFICATIONS
Saturn V was developed at Marshall Space Flight Center as the space launch vehicle for the
Apollo lunar landing missions. The first two vehicles (SA-501, 502) were unmanned test flights.
The first manned Apollo mission was launched by an upgraded version of Saturn V (SA-503), and
the following vehicles incorporated further upgrades, including more powerful F-1 engines
(SA-504 through 509). Apollo 8 was a circumlunar orbital CSM test, Apollo 9 an earth orbital
CSM/LM test, Apollo 10 a lunar orbital CSM/LM test, Apollo 11 the first manned lunar landing,
and Apollo 12-14 lunar explorations. Apollos 11-14 were classified as H-missions. Subsequent
lunar landings, called J-missions, included an uprated Apollo CSM and LM for extended lunar
stay time and the addition of a lunar rover. This increased payload capacity necessitated a
further Saturn V upgrade (SA-510 through 513) which was used to launch the Apollo 15-17
advanced lunar missions. The last of the Saturn Vs launched the unmanned Skylab space station.
First launch: 09-Nov-1967
Last launch: 14-May-1973
Number launched: 13
Launch sites: Cape Canaveral pad 39A
Principal uses: Apollo lunar missions, Skylab-1
Vehicle success rate: 100.0%
Performance: LEO (185 km): 118,000 kg; translunar trajectory: 47,000 kg
Cost: $739.8 million
Number of stages: 3
Overall length: 84.7 m no payload; 110.65 m with Apollo; 104.86 m with Skylab
Principal diameter: 10.06 m
Launch mass: varies, approximately 2,820 to 2,965 tonne
Launch thrust: 33,362 kN SL for vehicles SA-501-503; later uprated to 33,851 kN SL
for vehicles SA-504-513
(Masses are not typical for every mission, figures given are median.)
SATURN V STAGE 1
Designation: Saturn 1C
Engines: five Rocketdyne F-1 liquid bipropellant; four outboard engines are
gimbaled for steering
Length: 42.07 m
Diameter: 10.06 m
Dry mass: 130,570 kg dry
Oxidizer: liquid oxygen
Fuel: RP-1 hydrocarbon
Propellant mass: 2,149,500 kg
Other mass: 2,450 kg
Thrust: 33,362 kN SL; later uprated to 33,851 kN SL
Burn time: about 162 s for outboard engines, center engine cutoff at 135.5 s
(velocity at burnout ~9,900 km/h, altitude 67 km)
Separation: jettison 0.6 s after engine cutoff by eight 394 kN solid motors firing
for 2/3 s
Pressurization: RP-1 tank by 288 kg (3,510 liters) of gaseous helium at 211 atm; LOX
tank by gaseous oxygen converted from 2,880 kg of LOX by the engines; control system by 36
liters of gaseous nitrogen at 221 atm
Hydraulic system: power primarily for engine start and for gimballing of four
outboard engines
Electrical system: two 28 VDC batteries, basic power for all electrical functions
Instrumentation: handles ~900 measurements
Tracking: ODOP transponder
Environmental control: a ground-operated environmental control system protects stage
equipment from temperature extremes in both the forward skirt and thrust structure areas and
provides a nitrogen purge during prefiring and firing operations
SATURN V STAGE 2
Designation: Saturn II
Engines: five Rocketdyne J-2 liquid bipropellant; four outboard engines are
gimbaled for steering
Length: 24.84 m (includes aft interstage)
Diameter: 10.06 m
Dry mass: 36,395 kg dry + 5,195 kg aft interstage
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 451,650 kg
Other mass: 505 kg
Thrust: 5,004 kN vac maximum; later uprated to 5,116 kN vac max
Burn time: about 390 s for outboard engines, center engine cutoff at 296.5 s
(velocity at burnout ~24,950 km/h, altitude 188 km for SA-503-509; ~25,150 km/h, 174 km
for SA-510-512)
Ordnance: provides, in operational sequence, ignition of eight ullage motors
before ignition of five main engines, explosive separation of stage 2 interstage skirt,
explosive separation of stage 2 from stage 3, and ignition of four retrorockets to decelerate
stage 2 for complete separation
Electrical system: 6 electric bus systems, four 28 VDC batteries, and motor operated
transfer switches
Measurement: instrumentation, telemetry, and radio frequency subsystems
Environmental control: a ground-operated system provides proper temperature control
for equipment containers in the forward and aft skirt
Flight control: gimballing of the four outboard J-2 engines required for TVC,
accomplished by hydraulic-powered actuators which are electrically controlled from signals
initiated in the flight control computer of the instrument unit forward of stage 3
SATURN V STAGE 3
Designation: Saturn IVB
Engines: one gimbaled Rocketdyne J-2 restartable liquid bipropellant
Length: 17.87 m (includes aft interstage)
Diameter: 6.60 m
Dry mass: 11,380 kg + 3,650 kg aft interstage
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 106,940 kg
Other mass: 755 kg
Thrust: 1,001 kN vac maximum; later uprated to 1,023 kN vac max (about 905 kN vac
average)
Burn time: first burn about 145 s during launch (velocity at burnout ~28,050 km/h,
altitude 191 km for SA-503-509; ~28,090 km/h, 172 km for SA-510-512); second burn about 345 s for
translunar injection (velocity at burnout ~39,000 km/h)
Ordnance: provides explosive power for stage separation, retrorocket ignition, ullage
rocket ignition and jettison, and range safety requirements
Hydraulic system: power for gimballing J-2 engine
Electrical system: one 56 VDC and three 28 VDC batteries, basic power for all
electrical functions
Telemetry/instrumentation: pulse code modulated/frequency modulator (PCM/FM)
subsystem, providing transmission of flight data to ground stations
Environmental control: provides temperature-controlled environment for components in
aft skirt, aft interstage, and forward skirt
INSTRUMENT UNIT
The IBM Instrument Unit (IU), located forward of stage 3, contains the flight control
computers and instrumentation for monitoring and controlling the Saturn V systems.
Length: 0.91 m
Diameter: 6.6 m
Mass: 2,030 kg
ROCKETDYNE / F-1
The largest liquid engine rocket motor ever developed, it represented an incredible
order-of-magnitude development from the 670 kN thrust class motor developed for Navaho, Jupiter,
Thor, Atlas, Titan I, and Saturn I. Severe combustion stability problems were solved
during development and it never failed in flight.
Application: Saturn IC
First flown: 1967
Dry mass: 8,390 kg
Length: 5.8 m
Maximum diameter: 3.7 m
Oxidizer: liquid oxygen, delivered at 1,789 kg/s
Fuel: RP-1, delivered at 788 kg/s
Mixture ratio: 2.27
Turbopump: 5,550 rpm, 41,000 kW single turbine powered by gas generator requiring 77 kg/s propellants
Thrust: 6,672 kN SL; later uprated to 6,770 kN SL
Specific impulse: 265 s SL, 304 s vac
Expansion ratio: 16:1 with nozzle extension, 10:1 without
Combustion chamber pressure: 65.7 atm
Combustion chamber temperature: 3,300oC
Burn time: rated at 150 s
ROCKETDYNE / H-1
Application: Saturn 1
First flown: 1961
Dry mass: 635 kg
Length: 2.1 m
Maximum diameter: 0.8 m
Oxidizer: liquid oxygen
Fuel: RP-1
Thrust: 836.2 kN SL, 947.7 kN vac
Specific impulse: 255 s SL, 289 s vac
Expansion ratio: 8:1
Combustion chamber pressure: 40 atm
Burn time: 150 s
ROCKETDYNE / H-1B
Application: Saturn 1B, Saturn 1B-15, Saturn 1B-11
First flown: 1966
Dry mass: 988 kg
Length: 2.2 m
Maximum diameter: 0.8 m
Oxidizer: liquid oxygen
Fuel: RP-1
Thrust: 911.9 kN SL, 1030.2 kN vac
Specific impulse: 262 s SL, 296 s vac
Expansion ratio: 8:1
Combustion chamber pressure: 48 atm
Burn time: 155 s
ROCKETDYNE / J-2
Application: Saturn IVB stage in Saturn IB and Saturn V
First flown: 1967
Dry mass: 1,578 kg
Length: 3.4 m
Maximum diameter: 2.0 m
Oxidizer: liquid oxygen, delivered at 204 kg/s
Fuel: liquid hydrogen, delivered at 37.1 kg/s
Mixture ratio: 5.5
Oxidizer Turbopump: 8,600 rpm, 1,640 kW, 73.5 atm discharge pressure
Fuel Turbopump: 27,000 rpm, 5,820 kW, 83.4 atm discharge pressure
Thrust: 1,001 kN vac maximum; later uprated to 1,023 kN vac max
Specific impulse: 424 s vac (427 s vac at 5:1 mixture ratio)
Expansion ratio: 27.5:1
Combustion chamber pressure: 51.9 atm
Combustion chamber temperature: 3,180oC
Burn time: rated at 500 s
P&W / RL-10
Application: Nova A-3, Nova B-3, Saturn IV
First flown: 1964
Dry mass: 131 kg
Maximum diameter: 0.9 m
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Thrust: 66.70 kN vac
Specific impulse: 410 s vac
Expansion ratio: 40:1
Combustion chamber pressure: 24 atm
Burn time: 482 s
P&W / RL-10A-1
Application: Centaur C
First flown: 1962
Dry mass: 131 kg
Maximum diameter: 1.5 m
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Thrust: 66.70 kN vacuum
Specific impulse: 425 s vac
Expansion ratio: 47:1
Combustion chamber pressure: 24 atm
Burn time: 430 s