Category Apollo Saturn V News Reference

Uprated Saturn I (Saturn IB)

The space agency, using the building-block approach, conceived the Uprated Saturn I as the quickest, most reliable, and most economical means of provid­ing a vehicle with greater payload than the Saturn

I. This vehicle was planned for orbital missions with the Apollo spacecraft before the Satum V vehicle would be available.

The Uprated Satum I is based on a blending of existing elements of Saturn I and Saturn V. A re­designed Saturn I booster (designated the S-IB stage), and an S-IVB upper stage and instrument unit from the Satum V are used on this launch vehicle.

Maximum use of designs and facilities available from the earlier approved Satum programs saved both time and costs. .

The Satum I first stage was redesigned in several areas by NASA and the Chrysler Corporation, the stage contractor, for the expanded role as the Up­rated Saturn I booster. Basically, it retained the same shape and size, but required some modifica­tion for mating with the upper stage, which has a greater diameter and weight than the Satum I upper stage.

Stage weight was cut by more than 20,000 pounds to increase payload capacity. The Rocketdyne H-l engine was uprated to 200,000 pounds of thrust, compared with 188,000 pounds of thrust for each engine in the final Saturn I configuration. The en­gines will be improved again to 205,000 pounds beginning with the SA-206.

For the Uprated Saturn I, a guidance computer used in the early Satum I was replaced by another IBM computer of completely new design which in­corporates the added flexibility and extreme re­liability necessary to carry out the intended Uprated Satum I missions.

The Uprated Satum I, topped by the Apollo spaee-


craft, stands approximately 224 feet tall, and is about 21.7 feet in diameter. Total empty weight is about 85 tons, and liftoff weight fully fueled is approximately 650 tons.

Several uprated Saturn I vehicles have been launched since the original SA 201 launch on Feb. 26, 1966.


The onboard control pressure system consists of a high-pressure nitrogen storage bottle, an umbilical coupling and tubing assembly for filling the storage bottle, a manifold assembly, and control valves at the terminal ends of various nitrogen distribution lines. In some cases, two valves are paired with other associated equipment and block-mounted to form a control assembly.

The nitrogen onboard storage bottle has 2,200- cubic inch capacity and is made of titanium alloy. It is designed for a maximum proof pressure of 5,000 psig. It is filled and discharged through a port in the single boss. During flight launch preparation, the bottle is filled from a ground supply first to a pressurization of 1,600 psig well in advance of final countdown. This weight pressure is adequate for any prelaunch operational use. The second step occurs in the last hour of the launch countdown and


Control Pressure System

Подпись: SATURN V NEWS REFERENCE brings the storage bottle pressure up to its normal capacity of 3,250 plus or minus 50 psig. The manifold assembly serves as a gaseous nitrogen central receiving and distributing center as well as a mounting block for filters, shutoff solenoid valves, a pressure regulator, a relief valve, and pressure transducers. Ported manifolds provide tubing as­sembly connections to the storage bottle, umbilical coupling, and various tubing assembly distribution lines to control valves throughout the stage.


Heat exchanger LOX inlet temperature Heat exchanger GOX outlet temperature Heat exchanger helium outlet temperature Fuel pump inlet No. 2 temperature Heat exchanger LOX inlet flowrate

Primary and Auxiliary Junction Box

There are two electrical junction boxes in the flight instrumentation system. The primary junction box has provisions for eight electrical connectors, and the auxiliary junction box for five. Both junction boxes are welded closed and pressurized with an inert gas to prevent possible entry of contaminants and moisture.


The engine requires a source of pneumatic pres­sure, electrical power, and propellants for sustained engine operation. A ground hydraulic pressure source, thrust chamber prefill, gas generator and turbine exhaust igniters, and hypergolic fluid are required to start the engine.

When the start button is actuated, the checkout valve moves to transfer the hydraulic fuel return from the ground line to the turbopump low-pressure fuel inlet. The high-level oxidizer purge is initiated to the gas generator and thrust chamber LOX dome.

The gas generator and turbine exhaust gas igniters fire, and the engine control valve start solenoid is energized. Hydraulic pressure is directed to the opening port of the oxidizer valves. The oxidizer valves are part way open, and the hydraulic pres­sure is directed to the gas generator valve opening port. The gas generator valve opens, propellants under tank pressure enter the gas generator com­bustion chamber, and the propellant mixture is ig­nited by the gas generator igniters. The exhaust gas is ducted through the turbopump turbine, the heat exchanger, and the thrust chamber exhaust manifold into the nozzle extension walls where the fuel-rich mixture is ignited by the turbine exhaust gas igniters. As the turbine accelerates the fuel and the oxidizer pumps, the pump discharge pressures increase and propellants at increasing flowrates are supplied to the gas generator. Turbopump acceler-





ation continues and, as the fuel pressure increases, the igniter fuel valve opens and allows fuel pres­sure to huild up against the hypergol cartridge hurst diaphragm. The hypergol diaphragms burst under the increasing fuel pressure. Hypergolic fluid, followed by the ignition fuel, enters the thrust chamber. When hypergolic fluid enters the thrust chamber and contacts the oxidizer, spontaneous combustion occurs, establishing thrust chamber ignition. Thrust chamber pressure is transmitted through the sense line to the diaphragm of the ignition monitor valve. When the thrust chamber pressure increases, the ignition monitor valve ac­tuates and allows hydraulic fluid flow to the open­ing port of the fuel valves. The fuel valves open and fuel is admitted to the thrust chamber.

Fuel enters the thrust chamber fuel inlet manifold and passes through the thrust chamber tubes for cooling purposes and then through the injector into the thrust chamber combustion zone. As the thrust chamber pressure increases, the thrust-OK pres­sure switches are actuated indicating the engine is operating satisfactorily. The thrust chamber pres­sure continues to increase until the gas generator reaches rated power, controlled by orifices in the propellant lines feeding the gas generator. When engine fuel pressure increases above the ground-
supplied hydraulic pressure, the hydraulic pres­sure supply source is transferred to the engine. Hydraulic fuel is circulated through the engine com­ponents and then returned through the engine con­trol valve and checkout valve into the turbopump fuel inlet. The ground hydraulic source facility shutoff valve is actuated to the closed position when the fuel valves open. This allows the engine hy­draulic system to supply the hydraulic pressure during the cutoff sequence.


The primary function of the PU system is to assure simultaneous depletion of propellants by controlling the LOX flowrate to the J-2 engine. It also provides propellant mass information for controlling the fill and topping valves during propellant loading opera­tions. The system consists of mass sensors, an elec­tronics assembly, and an engine-mounted mixture ratio valve.


Propellant Utilization System

During loading operations, the mass of propellants loaded is determined within one per cent by the mass sensors. Tank over-fill sensors act as a backup system in the event the loading system fails to ter­minate fill operations.

Continuous LH2 and LOX residual readout signals are provided throughout third stage powered flight. Differences between the fuel and oxidizer mass indications, as sensed by the mass sensors, are continually analyzed and are then used to control the oxidizer pump bypass flowrate, which changes the engine mixture ratio correspondingly. The static inverter/converter supplies the analog volt­ages necessary to operate the PU system. It is commanded “on" and “off” by a switch selector and sequencer combination.

Engine Operation


Start sequence is initiated by supplying energy to two spark plugs in the gas generator and two in the augmented spark igniter for ignition of the propellants. Next, two solenoid valves are actuated: one for helium control, and one for ignition phase control. Helium is routed to hold the propellant bleed valves closed and to purge the thrust cham­ber LOX dome, the LOX pump intermediate seal, and the gas generator oxidizer passage. In addition, the main fuel valve and ASI oxidizer valve are opened, creating an ignition flame in the ASI cham­ber that passes through the center of the thrust chamber injector.

After a delay of 1, 3, or 8 seconds, during which time fuel is circulated through the thrust chamber to condition the engine for start, the start tank dis­charge valve is opened to initiate turbine spin. The length of the fuel lead is dependent upon the length of the Saturn V first stage boost phase. When the J-2 engine is used in the second stage of the Saturn V vehicle, a one-second fuel lead is necessary. The third stage of the Saturn V vehicle, on the other hand, utilizes a three-second fuel lead for its initial start and an eight-second fuel lead for its restart.

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After an interval of 0.450 second, the start tank discharge valve is closed and a mainstage control solenoid is actuated to: 1) turn off gas generator and thrust chamber helium purges; 2) open the gas generator control valve (hot gases from the gas generator now drive the pump turbines); 3) open the main oxidizer valve to the first position (14 degrees) allowing LOX to flow to the LOX dome to burn with the fuel that has been circulating through the injector; 4) close the oxidizer turbine bypass valve (a portion of the gases for driving the oxi­dizer turbopump were bypassed during the igni­tion phase); 5) gradually bleed the pressure from the closing side of the oxidizer valve pneumatic actuator controlling the slow opening of this valve for smooth transition into mainstage. Energy in the spark plugs is cut off and the engine is operating at rated thrust. During the initial phase of engine opera­tion, the gaseous hydrogen start tank will be re­


charged in those engines having a restart require­ment. The hydrogen tank is repressurized by tapping off a controlled mixture of liquid hydrogen from the thrust chamber fuel inlet manifold and warmer hy­drogen from the thrust chamber fuel injection mani­fold just before entering the injector.


During mainstage operation, engine thrust may be varied between 175,000 and 225,000 pounds by actuating the propellant utilization valve to increase or decrease oxidizer flow as described in the sec­tion “PU Valve”. This is beneficial to flight trajec­tories and for overall mission performance to make greater payloads possible.


When the engine cutoff signal is received by the electrical control package, it de-energizes the main – stage and ignition phase solenoid valves and ener­gizes the helium control solenoid de-energizer timer.

This, in turn, permits closing pressure to the main fuel valve, main oxidizer valve, gas generator con­trol valve, and augmented spark igniter valve. The oxidizer turbine bypass valve and propellant bleed valves open and the gas generator and LOX dome purges are initiated.



Final preparation of the space vehicle for launching, including propellant and ordnance loading, final checkout, and countdown takes place in the launch area.



Aerial of Pad 39A with VAB in Background

There are presently two launch areas on Complex 39. Each area is polygon-shaped with the linear distance from side to side at approximately 3,000 feet. The launch sites are 8,730 feet apart to allow operations on the pads to be handled independently for safety reasons.

Liquid oxygen, RP-1, and liquid hydrogen are stored near the perimeter of the launch sites. Heli­um and nitrogen gases are stored at 10,000 psi near the center.

An elevated steel and concrete hardstand is located in the center of each area. Steel support fittings for the mobile launcher and mobile service struc­ture are anchored to the hardstand. The exhaust flame trench runs through the center of the hard­stand. Prior to the launch, the wedge-shaped flame deflector is moved along rails into the trench.

The liquid oxygen system consists of a 900,000- gallon LOX storage facility and transfer system.

The RP-1 system consists of a storage area con­taining three 86,000-gallon tanks and a transfer system. The tanks have a carbon steel shell and a bonded stainless steel lining.

Gaseous nitrogen and helium are stored under­ground in vessels near the launch pad at pressures of 6,000 psi.

Automation of vehicle prelaunch checkout is ex­pected to uprate mission confidence and to increase launch rate capability. The heart of this automatic checkout system is the computer complex.



Cutaway Illustration of Pad 39A


While a major effort of this country’s space commit­ment was to explore the moon, the broader target was to build a capability—people, launch vehicles, propulsion, spacecraft, production, testing, and launching sites —to explore a vast new frontier and develop a long-range spacefaring capability that would establish continuing national preeminence.

The questions facing national space planners in 1961 and 1962 were complex. Although the use of a Saturn I for a manned lunar landing was theoreti­cally possible, it would have been extremely difficult. About six Saturn I launches would have been re­quired, their payloads being assembled in earth orbit to form a moon ship. No space rendezvous and docking had taken place at that time.

During the first half of 1962, two paramount de­cisions were announced: to develop a new general purpose launch vehicle in the middle range of sev­eral under consideration, and to conduct the manned lunar landing by use of a lunar orbit rendezvous (LOR) technique.

The Saturn V, as the chosen vehicle was named, was given the go-ahead in January, 1962.

It was to be composed of three propulsive stages and a small instrument unit to contain guidance and control. It could perform earth orbital missions through the use of the first two stages, while all three would be required for lunar and planetary ex­peditions. The ground stage was to be powered by five F-l engines, each developing 1.5 million pounds of thrust, and the stage would have five times the power of the Saturn I booster then under develop­ment. The upper stages would use the J-2 hydrogen/ oxygen engine, five in the second stage and one in the third. Each would develop up to 225,000 pounds of thrust. Such a rocket would be capable of placing 120 tons into earth orbit or dispatching 45 tons to the moon. (The numbers have been uprated now to about 125 and 47-1/2.)

During its assembly, checkout, and launch, the Saturn V7 would use a new mobile launch concept. It would be assembled in a huge Vehicle Assembly Building, and then transported in an upright posi­tion to a launch pad several miles away.

Propulsion development decisions preceded those for the vehicles.

The need for a building-block rocket engine in the million-pound-thrust class was apparent even as ARPA was ordering work to begin on the first stage cluster of engines for the Saturn I. In January, 1959, NASA contracted with North American Aviation’s Rocketdyne Division for development of the F-l.

Late in 1959, the Silverstein Committee recom­mended the development of a new high-thrust hydro­gen engine to meet upper stage requirements. In June, 1960, Rocketdyne was selected to develop the J-2 engine after evaluation of competitive proposals by NASA.

Three proposed Apollo modes which were considered in detail were: the direct flight mode, using a very large launch vehicle called "Nova”; the earth orbital rendezvous (EORI mode, requiring separate Saturn launches of a tanker and a manned spacecraft; and the lunar orbital rendezvous mode, requiring a single launch of the manned spacecraft and the lunar module.

Selected was the LOR mode, in which the injected spacecraft weight would be reduced from 150,000 pounds to approximately 80,000 pounds by eliminat­ing the requirement for the propulsion needed to soft-land the entire spacecraft on the lunar surface.

A small lunar excursion module, or LEM, now re­ferred to as the lunar module, would be detached after deboost into lunar orbit. The lunar module would carry two of the three-man Apollo crew to a soft landing on the moon and would subsequently be launched from the moon to rendezvous with the third crew member in the “mother ship.” The entire crew would then return to earth aboard the com­mand module.

NASA concluded that LOR offered the greatest assurance of successful accomplishment of the Apollo objectives at the earliest practical date.

Members of NASA’s Manned Space Flight Man­agement Council recommended LOR unanimously in 1962 because it:

1. Provided a higher probability of mission suc­cess with essentially equal mission safety;

2. Promised mission success some months earlier than did other modes;

3. Would cost 10 to 15 per cent less than the other modes; and

4. Required the least amount of technical develop­ment beyond existing commitments while ad­vancing significantly the national technology.

As a part of the Saturn V decision, it was deter-



mined that elements of the existing Saturn I ve­hicle and the planned Saturn V would be combined to form a new mid-range vehicle, the uprated Saturn I (Saturn IB), The Uprated Saturn I would have a payload capability 50 per cent greater than the Saturn I and would make possible the testing of the Apollo spacecraft in earth orbit about one year earlier than would be possible with the Saturn V.

By the end of 1962, all elements of the new pro­gram were under way, with the Marshall Space Flight Center directing the work for NASA. The Boeing Company; Space Division of North Amer­ican Aviation, Inc.; and Douglas Aircraft Company were acting as prime contractors for the Saturn V first, second, and third stages, respectively. Engines were being developed by the Rocketdyne Division of North American. MSFC designed the instrument unit and awarded a production contract to Inter­national Business Machines Corp. (Chrysler Corp. had been selected to produce the first stage of the Uprated Saturn I.)

A large network of production, assembly, testing, and launch facilities was also being prepared by the end of 1962. Aside from the provision of various facilities at contractor plants and the augmentation of the Marshall Space Flight Center resources, three new government operations were established: the launch complex in Florida operated by the NASA – Kennedy Space Center and two new elements of MSFC—Michoud Assembly Facility in New Orleans, La., for the production of boosters, and Mississippi Test Facility, Bay St. Louis, Miss., for captive firing of stages.

Four years after its establishment, the Saturn V program was progressing on schedule, pointing toward the launch of the first vehicle in 1967 and fulfillment of the manned lunar landing before the end of the decade.


Following are highlights of the Saturn V develop­ment program:


Aug. 24 NASA announced the selection of the 88,000-acre site at Merritt Island, Fla., adjacent to Kennedy Space Center, then Cape Canaveral, for the assembly, check­out, and launch of the Saturn V.

Sept. 7 NASA selected the government-owned Michoud plant, New Orleans, as production site for Saturn boosters. It became a part of the Marshall Space Flight Center.

Sept. 11 NASA selected North American Aviation, Inc., to develop and build the second stage
for an advanced Saturn launch vehicle (as yet undefined) for manned and unmanned missions. One month later the Marshall Center directed NAA to design the second stage using five J-2 engines. A prelimi­nary contract was signed in February, 1962.

Oct. 6 NASA selected the Picayune-Bay St. Louis, Miss., area for its Mississippi Test Facility — an arm of the Marshall Center —for use in static testing of rocket stages and en­gines.

Dec. 15 The Boeing Company was selected as prime contractor for the first stage of the ad­vanced Saturn vehicle —not yet fully de­fined. A preliminary contract was signed in February, 1962, with the work to be con­ducted at the Michoud Assembly Facility.

Dec. 21 NASA selected the Douglas Aircraft Com­pany to negotiate a contract to develop the third stage (S-IVB) of the advanced Saturn, based on the Saturn I’s S-IV stage. A sup­plemental contract for production of 11 third stages was signed in August, 1962.


Jan. 10 Announcement was made that the advanced Saturn vehicle would have a first stage powered by five F-l engines, a second stage powered by five J-2 engines, and for lunar missions a third stage with one J-2 engine.

Jan. 25 NASA formally assigned development of the three-stage Saturn C-5 (Saturn V be­came the name in February, 1963) to MSFC.

April 11 NASA Headquarters gave the Apollo/ Saturn I/Saturn V highest national priority.

May 26 Rocketdyne Division of NAA conducted the first full-thrust, long-duration F-l en­gine test.

July 11 It was announced that the Saturn IB (Up­rated Saturn I) would be developed and that the lunar orbit rendezvous method of accomplishing a lunar landing had been selected.

Decern – The U. S. Army Corps of Engineers

ber awarded a contract for the design of the Vehicle Assembly Building (VAB) at the Florida launch complex.


Feb. 27 The first contract for the Mississippi Test Facility (MTF) Saturn V test facilities was awarded.

May The J-2 engine was successfully fired for the first time in a simulated space altitude of 60,000 feet.

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Oct. 31 The Marshall Center received the first production model of the F-l engine.


Nov. 12 NASA contracted for the first Saturn V launch pad at the Kennedy Space Center.


March IBM was awarded an instrument unit con­tract for the digital computer and data adapter by the Marshall Center. IBM be­came the prime IU contractor in May.

Oct. 9 The Edwards AFB test facility was ac­cepted as the F-l test complex, amounting to a cost of $34 million.

Dec. 1 The first mainstage shakedown firing of the third stage battleship was accomplish­ed, lasting 10 seconds.

Dec. 23 First full-duration firing of the third stage battleship occurred.


April 16 All five engines of the S-IC-T, first stage test vehicle, were fired at the Marshall Center for 6.5 seconds.

April 24 The first cluster ignition test of the second stage battleship was successfully com­pleted.

Aug. 5 The first full-duration firing of the first

stage was conducted successfully at the Marshall Center.

Aug. 8 Third stage flight readiness test of 452 seconds, fully automated, was accomplish­ed at Sacramento.

Aug. 13 The IU was qualified structurally and man­
rated for Saturn V use by withstanding a 140 per cent load limit.

Aug. 17 The third stage battleship was tested in Saturn V configuration for full duration (start-stop-restart).

Dee. 16 The S-IC-T static firings were completed at the Marshall Center with a total of 15 firings—three of full duration.


Feb. 17 The S-IC-1 underwent static firing at the

& 25 Marshall Center and required no more static firings.

Mar. 30 The S-IU-500F was mated to the three stages of the Saturn V facilities vehicle at the Kennedy Space Center’s VAB.

May 20 First full-duration firing of the second stage flight stage was conducted at MTF.

May 25 The Apollo/Saturn V facilities vehicle, AS – 50Q-F, was transported to Pad A at Launch Complex 39, KSC, on the crawler.

May 26 Full-duration acceptance firing of the S – IVB-501, the first flight version of the third stage for Saturn V, was accomplished.

Septem – The F-l and J-2 engines were qualified for

ber manned flights.

Dec. 1 Initial static firing of the first flight ver­sion of the second stage occurred at MTF.

Nov. 15 The first flight version of the first stage was static fired at MSFC.








WEIGHT: 300,000 lb. (dry)

4,792,0 lb. (loaded)

DIAMETER: 33 ft.

HEIGHT: 138 ft.

BURN TIME: About 2.5 min.

VELOCITY: 6,000 miles per hour at burnout (approx.)





PROPULSION: Five bipropellant F-l engines Total thrust: 7.5 million lb.

Propellant: RP-1- 203,000 gal. or 1,359,000 lb.

LOX—331,000 gal or 3,133,000 lb.

Pressure: Control 1.27 cubic feet of gaseous nitrogen at 3,250 psig

Fuel pressurization – 124 cubic feet or 636 lb. of gaseous helium at 3,100 psig LOX pressurization gaseous oxygen converted from 6,340 pounds of LOX by the engines HYDRAULIC: Power primarily for engine start and for gimbaling four outboard engines ELECTRICAL: Two 28 VDC batteries, basic power for all electrical functions INSTRUMENTATION: Handles approx. 900 measurements TRACKING: ODOP Transponder



The ground control pressure system provides a direct ground pressure supply for some of the first stage pneumatically actuated valves. The valves are involved with propellant fill and drain and emer­gency engine shutdown system operations. Direct ground control assures a backup system in case of emergency and conserves the onboard nitrogen supply.


The onboard purge pressure system consists of three high-pressure nitrogen storage bottles iden­tical to the onboard control pressure storage bottle, an umbilical coupling and tubing for filling the bot­tles, and a manifold assembly and tubing for re­ceiving and delivering the gas to the engine and calorimeter purge systems. These purge systems expel propellant leakage and are necessary from the time of loading throughout flight.

Environmental Control System

The 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




Temperature-controlled air is provided by a ground air conditioning unit from approximately 14 hours before launch to approximately 6 hours before launch. At this time, gaseous nitrogen from an auxiliary nitrogen supply unit is introduced into the system and used to purge and condition the for­ward skirt and thrust structure areas until umbilical disconnect at launch.

A distribution manifold vents air and gaseous ni­trogen through orifices into the thrust structure to maintain proper temperature. Air and nitrogen are supplied from the ground.

The system also distributes air and gaseous nitro­gen to instrumentation canisters mounted in the forward skirt. Temperatures in the canisters are held to meet requirements of electrical equipment. From the canisters, the conditioning gas is vented into the forward skirt compartment.

Visual Instrumentation

Visual instrumentation, presently planned to be installed on two flight stages, is designed to monitor critical stage functions prior to and during static test and flight conditions.


When the cutoff signal is initiated, the LOX dome operational oxidizer purge comes on. and the en­gine control valve stop solenoid is energized. Hy­draulic pressure holding open the gas generator valves, the oxidizer valves, and the fuel valves is routed to return. Simultaneously, hydraulic pres­sure is directed to the closing ports of the gas gen­erator valve, the oxidizer valves, and the fuel valves. The checkout valve is actuated and, as propellant pressures decay, the high level oxidizer purge be­gins to flow: then the igniter fuel valve and the igni­tion monitor valve close. Thrust chamber pressure will reach the zero level at about the same time the oxidizer valves reach full-closed.




WEIGHT: 95,000 lb. (dry)

1,037,0 lb. (loaded)

DIAMETER: 33 ft.

HEIGHT: 81 ft. 7 in.

BURN TIME: 6 min. approx, (actually 395 sec.)

VELOCITY: 15,300 miles per hour at burnout (approx.)






Thrust: More than 1,000,000 lb. (225,000 maximum each engine)

Propellant: LH2~-260,000 gal. (153,000 lb.)

LOX 83,000 gal. (789,000 lb.)

ELECTRICAL: 6 electrical bus systems, four 28-volt DC flight batteries, and motor-operated power transfer switches ORDNANCE: Provides, in operational sequence, ignition of eight ullage motors before ignition of five main engines, explosive separation of second stage interstage skirt, explosive separation of second stage from third stage, and ignition of four retrorockets to decelerate second stage for complete separation MEASUREMENT: Instrumentation, telemetry, and radio frequency subsystems THERMAL CONTROL: A ground-operated system that provides proper temperature control for equipment containers in the forward and aft skirt FLIGHT CONTROL: Gimbaling of the four outboard J-2 engines as required for thrust vector

control, accomplished by hydraulic-powered actuators which are electrically controlled from signals initiated in the flight control computer of the instrument unit (atop the Saturn V third stage)



The pneumatic control system provides GHe (gas­eous helium) pressure to operate all third stage pneumatically operated valves with the exception of those provided as components of the J-2 engine. GHe is supplied from an ambient helium sphere and pressurized from a ground source before propel­lant fill operations at 3,100 ± 100 psia at 70° Fahren­heit for valve operation. The sphere is located on the thrust structure and is pre-conditioned to above 70° Fahrenheit from the environmental control system before liftoff.

The pneumatic control system provides regulated pressure at 475 ± 25 psig for operation of the LH, and LOX vent-relief valves during propellant load­ing, LH2 directional control valve, LOX and LH, fill and drain valves during loading, and the GH2 engine start system vent-relief valve. It also pro­vides operating pressures for the LH, and LOX turbopump turbine purge module, LOX chilldown pump purge module control, LOX and LH, pre­valves, LOX and LH, chilldown shutoff valves, and the LH, continuous propulsive vent control module.

The pneumatic control subsystem is protected from overpressure by a normally open solenoid valve controlled by a downstream pressure-sensing switch. At pressures greater than 535 + 15, -10 psia, the pressure switch actuates and closes the valve. At pressures below 450 + 15, -10 psia, the pressure switch drops out and the solenoid opens, thus acting as a backup regulator.