Launch Vehicle

Saturn V

The Saturn V, 363 feet tall with the Apollo spacecraft in place, generates enough thrust to place a 125-ton payload into a 105 nm circular Earth orbit or boost a smaller payload to the vicinity of any planet in the solar system. It can boost about 50 tons to lunar orbit. The thrust of the three propulsive stages range from more than 7.7 million pounds for the booster to 230,000 pounds for the third stage at operating altitude. Including the instrument unit, the launch vehicle is 281 feet tall.

First Stage

The first stage (S-IC) was developed jointly by the National Aeronautics and Space Administration's Marshall Space Flight Center, Huntsville, Ala., and the Boeing Co.

The Marshall Center assembled four S-IC stages: a structural test model, a static test version and the first two flight stages. Subsequent flight stages are being assembled by Boeing at the Michoud Assembly Facility in New Orleans. The S-IC stage destined for the Apollo 9 mission was the first flight booster static tested at the NASA-Mississippi Test Facility. That test was made on May 11, 1967. Earlier flight stages were static fired at the NASA-Marshall Center.

The S-IC stage provides first boost of the Saturn V launch vehicle to an altitude of about 37 nautical miles (41.7 statute miles, 67.1 kilometers) and provides acceleration to increase the vehicle's velocity to 9,095 feet per second (2,402 m/sec, 5,385 knots, 6,201 mph). It then separates from the S-II stage and falls to Earth about 361.9 nm (4l6.9 sm, 667.3 km) downrange.

Normal propellant flow rate to the five F-l engines is 29,522 pounds per second. Four of the engines are mounted on a ring, each 90 degrees from its neighbor. These four are gim-balled to control the rocket's direction of flight. The fifth engine is mounted rigidly in the center.

Second Stage

The second stage (S-II), like the third stage, uses high performance J-2 engines that burn liquid oxygen and liquid hydrogen. The stage's purpose is to provide second stage boost nearly to Earth orbit.

At engine cutoff, the S-II separates from the third stage and, following a ballistic trajectory, plungeB into the Atlantic Ocean about 2,412 nm (2,778.6 sm, 4,468 km) downrange from Cape Kennedy.

Five J-2 engines power the S-II. The four outer engines are equally spaced on a 17.5-foot diameter circle. These four engines may be gimballed through a plus or minus seven-degree square pattern for thrust vector control. Like the first stage, the center (number 5) engine is mounted on the stage centerline and is fixed.

The S-II carries the rocket to an altitude of 103 nm (118.7 sm, 190.9 km) and a distance of some 835 nm (961.9 sm, 1,548 km) downrange. Before burnout, the vehicle will be moving at a speed of 23<000 fps or 13,642 knots (15,708 mph, 25,291 kph, 6,619 m/sec). The J-2 engines will burn six minutes 11 seconds during this powered phase.

The Space Division of North Anerican Rockwell Corp. builds the S-II at Seal Beach, Calif. The cylindrical vehicle is made up of the forward skirt (to which the third stage attaches), the liquid hydrogen tank, the liquid oxygen tank, the thrust structure (on which the engines are mounted) and an interstage section (to which the first stage connects). The propellant tanks are separated by an insulated common bulkhead.

The S-II was static tested by North American Rockwell at the NASA-Mississippi Test Facility on Feb. 10, 1968. This Apollo 9 flight stage was shipped to the test site via the Panama Canal for the test firing.

Third Stage

The third stage (S-IVB) was developed by the McDonnell Douglas Astronautics Co. at Huntington Beach, Calif, At Sacramento, Calif., the stage passed a static firing test on Aug. 26, i960, as part of the preparation for the Apollo 9 mission. The stage was flown directly to the NASA-Kennedy Space Center.

Measuring 58 feet 4 inches long and 21 feet 8 inches In diameter, the S-IVB weighs 25,300 pounds dry. At first ignition it weighs 259,377 pounds. The interstage section weighs an additional 8,08l pounds. The staged J-2 engine burns liquid oxygen and liquid hydrogen.

The stage, with its single engine, provides propulsion three times during the Apollo 9 mission. The first burn occurs immediately after separation from the S-II. It will last long enough (112 seconds) to insert the vehicle and spacecraft into Earth parking orbit. The second burn, which begins after separation from the spacecraft, will place the stage and instrument unit into a high apogee elliptical orbit. The third burn will drive the stage into solar orbit.

The fuel tanks contain 77,675 gallons of liquid hydrogen and 19,600 gallons of liquid oxygen at first ignition, totalling 230,790 pounds of propellants. Insulation between the two tanks Is necessary because the liquid oxygen, at about 293 degrees below zero P., is warm enough, relatively, to heat the liquid hydrogen, at 423 degrees below zer F., rapidly and cause it to change into a gas.

The first reignition burn is for 62 seconds and the second reignition burn Ib planned to last 4 minutes 2 seconds. Both reignitlons will be inhibited with inhibit removal to be by ground command only after separation of the spacecraft to a Bafe distance.

Instrument Unit

The Instrument Unit (IU) Is a cylinder three feet high and 21 feet 8 Inches in diameter. It weighs 4,295 pounds and contains the guidance, navigation and control equipment which will steer the vehicle through its Earth orbits and into the final escape orbit maneuver.

The IU also contains telemetry, communications, tracking and crew safety systems, along with Its own supporting electrical power and environmental control systems.

Components making up the "brain" of the Saturn V are mounted on cooling panels fastened to the inside surface of the instrument unit skin. The "cold plates" are part of a system that removes heat by circulating cooled fluid through a heat exchanger that evaporates water from a separate supply into the vacuum of space.

The six major systems of the instrument unit are structural, thermal control, guidance and control, measuring and telemetry, radio frequency and electrical.

The Instrument unit provides navigation, guidance and control of the vehicle; measurement of vehicle performance and environment; data transmission with ground stations; radio tracking of the vehicle; checkout and monitoring of vehicle functions; Initiation of stage functional sequencing; detection of emergency situations; generation and network distribution of electric power for system operation; and preflight checkout and launch and flight operations.

A path-adaptive guidance scheme is used in the Saturn V instrument unit. A programmed trajectory is used in the initial launch phase with guidance beginning only after the vehicle has left the atmosphere. This is to prevent movements that might cause the vehicle to break apart while attempting to compensate for winds, jet streams and gusts encountered in the atmosphere.

If such air currents displace the vehicle from the optimum trajectory In climb, the vehicle derives a new trajectory. Calculations are made about once each second throughout the flight. The launch vehicle digital computer and launch vehicle data adapter perform the navigation and guidance computations.

The ST-124m inertial platform — the heart of the navigation, guidance and control system — provides space-fixed reference coordinates and measures accleration along the three mutually perpendicular axes of the coordinate system.

International Business Machines Corp., is prime contractor for the instrument unit and is the supplier of the guidance signal processor and guidance computer. Major suppliers of instrument unit components are: Electronic ComraunIcations, Inc., control computer; Bendix Corp., ST-124M Inertial platform; and IBM Federal Systems Division, launch vehicle digital computer and launch vehicle data adapter.


The 4l rocket engines of the Saturn V have thrust ratings ranging from 72 pounds to more than 1.5 million pounds. Some engines burn liquid propellants, others use solids.

The five F-l engines in the first stage burn RF-1 (kerosene) and liquid oxygen. Each engine in the first stage develops an average of 1,544.000 pounds of thrust at liftoff, building up to an average of 1,©33,900 pounds before cutoff. The cluster of five engines gives the first stage a thrust range from 7.72 million pounds at liftoff to 9*169,560 pounds just before center engine cutoff.

The F-l engine weighs almost 10 tons, is more than 18 feet high and has a nozzle-exit diameter of nearly 14 feet. The F-l undergoes static testing for an average 650 seconds In qualifying for the 150-second run during the Saturn V first stage booster phase. This run period, 800 seconds, is still far less than the 2,200 seconds of the engine guarantee period. The engine consumes almost three tons of propellants per second.

The first stage of the Saturn V for this mission has eight other rocket motors. These are the solid-fuel retro-rockets which will slow and separate the stage from the second stage. Each rocket produces a thrust of 87,900 pounds for 0.6 second.

The main propulsion for the second stage is a cluster of five J-2 engines burning liquid hydrogen and liquid oxygen. Each engine develops a mean thrust of more than 205,000 pounds at 5.0:1 mixture ratio (variable from 193,000 to 230,000 in phases of flight), giving the stage a total mean thrust of more than a million pounds.

Designed to operate In the hard vacuum of space, the 3,500-pound J-2 is more efficient than the F-l because it burns the high-energy fuel hydrogen.

The second stage also has four 21,000-pound-thrust solid fuel rocket engines. These are the ullage rockets mounted on the S-IC/S-II interstage section. These rockets fire to Bettle liquid propellant In the bottom of the main tanks and help attain a "clean" separation from the first stage, they remain with the interstage when it drops away at second plane separation. Four retrorockets are located in the S-IVB aft interstage (which never separates from the S-Il) to separate the S-II from the S-IVB prior to S-IVB Ignition.

Eleven rocket engines perform various functions on the third stage. A single J-2 provides the main propulsive force; there are two jettisonable main ullage rockets and eight smaller engines in the two auxiliary propulsion system modules.

Launch Vehicle Instrumentation and Communication

A total of 2,159 measurements will be taken in flight on the Saturn V launch vehicle: 666 on the first stage, 975 on the second stage, 296 on the third stage and 222 on the instrument unit.

The Saturn V will have 16 telemetry systems: six on the first stage, six on the second stage, one on the third stage and three on the instrument unit. A radar tracking system will be on the first stage, and a C-Band system and command system on the instrument unit. Each powered stage will have a range safety system as on previous flights.

There will be no film or television cameras on or in any of the stages of the Saturn 504 launch vehicle.

Vehicle Weights During Flight


Ignition Liftoff Mach 1 Max. Q CECO OECO

S-IC/S-II Separation S-II Ignition Interstage Jettison LET Jettison S-II Cutoff S-II/S-IVB Separation S-IVB Ignition S-IVB First Cutoff Parking Orbit Injection Spacecraft First Separation

Spacecraft Docking

Spacecraft Second Separation S-IVB First Reignition

Vehicle Weight 6,486,915 pounds 6,400,648 pounds 4,487,938 pounds 4,015,350 pounds

2.443.281 pounds 1,831,57^ pounds 1,452,887 pounds 1,452,277 pounds

1.379.282 pounds 1,354,780 pounds

461,636 pounds

357,177 pounds

357,086 pounds

297,166 pounds

297,009 pounds

232,731 pounds (Only S-IVB and IM)

291,572 pounds (S-IVB and complete spacecraft)

199,725 pounds (only S-IVB stage)

199,346 pounds


Vehicle Weight 170,344 pounds 170,197 pounds 169,383 pounds pounds 54,300 pounds 35,231 pounds 31,400 pounds

S-IVB Second Cutoff

Intermediate Orbit Injection

S-IVB Second Reignition

S-IVB Third Cutoff

Escape Orbit Injection End LOX Dump End LH2 Dump

S-IVB Restarts

The third stage (S-IVB) of the Saturn V rocket for the Apollo 9 mission will burn a total of three times In space, the last two burns unmanned for engineering evaluation of stage capability. It has never burned more than twice in space before.

Engineers also want to check out the primary and backup propellant tank pressurization systems and prove that in case the primary system fails the backup system will pressurize the tanks sufficiently for restart.

Also planned for the second restart is an extended fuel lead for chilldown of the J-2 engine using liquid hydrogen fuel flowing through the engine to chill it to the desired temperature prior to restart.

All these events — second restart, checkout of the backup pressurization system and extended fuel lead chilldown will not affect the primary mission of Apollo 9. The spacecraft will have been separated and will be sare in a different orbit from that of the spacecraft before the events begin.

Previous flights of the S-IVB stage have not required a second restart, and none of the flights currently planned have a specific need for the third burn.

The second restart during the Apollo 9 mission will come 80 minutes after second engine bum cutoff to demonstrate a requirement that the stage has the capability to restart in space after being shut down for only 80 minutes. That capability has not yet been proven because flights to date have not required as little as 80 minutes of coastings.

The first restart Is scheduled for four hours 45 minutes and 41 seconds after launch, or about four and one-half hours after first burn cutoff. The need for a second restart may-occur on future non-lunar flights. Also, the need to launch on certain days could create a need for the 80-minute restart capability.

The primary pressurization system of the propellant tanks for S-IVB restart uses a helium heater. In this system, nine helium storage spheres in the liquid hydrogen tank contain gaseous helium charged to about 3,000 psl. This helium is passed through the heater which beats and expands the gas before it enters the propellant tanks. The heater operates on hydrogen and oxygen gas from the main propellant tanks.

The backup system consists of five ambient helium spheres mounted on the stage thrust structure. This system, controlled by the fuel repressurization control module, can repressurize the tanks in case the primary system fails.

The first restart will use the primary system. If that system fails, the backup system will be used. The backup system will be used for the second restart.

The primary reason for the extended fuel lead chllldown test is to demonstrate a contingency plan in case the chilldown pumps fail.

In the extended chilldown event, a ground command will cut off the pumps to put the stage in a simulated failure condition. Another ground command will start liquid hydrogen flowing through the engine.

This is a slower method, but enough time has been allotted for the hydrogen to chill the engine and create about the same conditions as would be created by the primary system.

The two unmanned restarts in this mission are at the very limits of the design requirements for the S-IVB and the J-2 engine. For this reason, the probability of restart is not as great as in the nominal lunar missions.

Two requirements are for the engine to restart four and one-half hours after first burn cutoff, and for a total stage lifetime of six and one-half hours. The first restart on this mission will be four and one-half hours after first burn cutoff, and second restart will occur six hours and six minutes after liftoff, both events scheduled near the design limits.

Differences in Apollo 8 and Apollo 9 Launch Vehicles

Two modifications resulting from problems encountered during the second Saturn V flight were incorporated and proven successful on the third Saturn V mission. The new helium pre-valve cavity pressurization system will again be flown on the S-IC stage of Apollo 9. Also, new augmented spark igniter lines which flew on the engines of the two upper stages of Apollo 8 will again be used on Apollo 9.

The major S-IC stage differences between Apollo 8 and Apollo 9 are:

1. Dry weight was reduced from 304,000 pounds to 295,600 pounds„

2. Weight at ground ignition increased from 4,800,000 pounds to 5,026,200 pounds.

3. Instrumentation measurements were reduced from 891

to 666.

4. Camera Instrumentation electrical power system Is not installed on S-IC-4.

5. S-IC-4 carries neither a TV camera system nor a film camera system.

The Saturn V will fly a somewhat lighter and slightly more powerful second stage beginning with Apollo 9.

The changes are:

1. Nominal vacuum thrust for J-2 engines was increased from 225,000 pounds each to 230,000 pounds each. This changed the second stage thrust from a total of 1,125,000 pounds to 1,150,000 pounds.

2. The approximate empty S-11 stage weight has been reduced from 88,000 to 84,600 pounds. The S-IC/S-II interstage weight was reduced from 11,800 to 11,664 pounds.

3. Approximate stage gross liftoff weight was increased from 1,035,000 pounds to 1,069,114 pounds.

4. S-II instrumentation system was changed from research and development to a combination of research and development and operational.

5. Instrumentation measurements were decreased from 1,226 to 975.

Major differences between the S-IVB stage used on Apollo 8 and the one on Apollo 9 are s

1. S-IVB dry stage weight decreased from 26,421 pounds to 25,300 pounds. This does not include the 8,081-pound interstage section.

2. S-IVB gross stage weight at liftoff decreased from 263,204 pounds to 259,337 pounds.

3. Stage measurements evolved from research and development to operational status.

4. Instrumentation measurements were reduced from 342 to 296.

Major instrument unit differences between Apollo 8 and Apollo 9 include deletions of a rate gyro timer, thermal probe, a measuring distributor, a tape recorder, two radio frequency assemblies, a source follower, a battery and six measuring racks. Instrumentation measurements were reduced from 339 to 222.

Launch Vehicle Sequence of Events

(Note: Information presented in this press kit is based upon a nominal mission. Plans may be altered prior to or during flight to meet changing conditions.)


The first stage of the Saturn V will carry the launch vehicle and Apollo spacecraft to an altitude of 36.2 nautical miles (41.7 sm, 67.1 km) and 50 nautical miles (57.8 sm, 93 km) downrange, building up speed to 9,095.2 feet per second (2,402 m/sec, 5,385.3 knots, 6,201.1 sm) in two minutes 31 seconds of powered flight. After separation from the second stage, the first stage will continue on a ballistic trajectory ending in the Atlantic Ocean some 361.9 nautical miles (416.9 sm, 670.9 km) downrange from Cape Kennedy (latitude 30.27 degrees north and longitude 73.9 degrees west) about nine minutes after liftoff.

Second Stage

The second stage, with engines running six minutes and 11 seconds, will propel the vehicle to an altitude of about 103 nautical miles (118.7 sm, 190.9 km) some 835 nautical miles (961.7 sm, 1,547 Ion) downrange, building up to 23,040.3 feet per second (6,619 m/sec, 13,642.1 knots, 15,708.9 mph) space fixed velocity. The spent second stage will land in the Atlantic Ocean about 20 minutes after liftoff some 2,410.3 nautical miles (2,776.7 sm, 4,468.4 km) from the launch site, at latitude 31«46 degrees north and longitude 34.06 degrees west.

Third Stage First Burn

The third stage, in its 112-second initial burn, will place itself and the Apollo spacecraft into a circular orbit 103 nautical miles (119 sm, 191 km) above the Earth. Its inclination will be 32.5 degrees and the orbital period about 88 minutes. Apollo 9 will enter orfcit at about 56.66 degrees west longitude and 32.57 degrees north latitude.

Parking Orbit

The Saturn V third stage will be checked out in Earth parking orbit in preparation for the second S-IVB burn. During the second revolution, the Command/Service Module (CSM) will separate from the third stage. The spacecraft LM adapter (SLA) panels will be jettisoned and the CSM will turn around and dock with the lunar module (IM) while the LM is still attached to the S-IVB/IU. This maneuver Is scheduled to require 14 minutes. About an hour and a quarter after the docking, U4 ejection Is to occur. A three-second SM RCS burn will provide a safe separation distance at S-IVB reignition. All S-IVB reignitions are nominally inhibited. The inhibit is removed by ground command after the CSM/LM is determined to be a safe distance away.

Third Stage Second Burn

Boost of the unmanned S-IVB stage from Earth parking orbit to an intermediate orbit occurs during the third revolution shortly after the stage comes within range of Cape Kennedy (4 hours 44 minutes 42 seconds after liftoff). The second burn, lasting 62 seconds, will put the S-IVB into an elliptical orbit with an apogee of 3,052 km (1,646.3 nm, 1,896.5 sm) and a perigee of 196 km (105.7 nm, 121.8 sm). The stage will remain in this orbit about one-half revolution.

Third Stage Third Burn

The third stage is reignlted at 6 hours 6 minutes 4 seconds after liftoff for a burn lasting 4 minutes 2 seconds. This will place the stage and instrument unit into the escape orbit. Ninety seconds after cutoff, LOX dump begins and lasts for 11 minutes 10 seconds, followed 10 seconds later by the dump of the liquid hydrogen, an exercise of 18 minutes 15 seconds duration. Total weight of the stage and instrument unit placed Into solar orbit will be about 31,400 pounds.

launch Vehicle Key Events
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