Category Apollo Saturn V News Reference

TESTING

INTRODUCTION

The expense of the Saturn V makes it imperative that no effort be spared to assure that it will per­form as expected in flight. The magnitude of the Saturn V ground test program, therefore, is un­precedented. To qualify for flight, all components and systems must meet standards deliberately set much higher than actually required. This margin of safety is built into all manrated space hardware.

Compared with earlier rocket programs the ground testing on Saturn V is more extensive and the flight testing is shorter. The ground test programs con­ducted on the F-l and J-2 engines, which power the three stages, offer an example of the thoroughness of this testing effort. The J-2 has been fired some 2,500 times on the ground, for a total running time of more than 63 hours. The F-l has been fired more than 3,000 times for a running time of more than 43 hours.

Further, in earlier rocket programs such as Red­stone, Thor, and Jupiter, 30 to 40 R&D flight tests were standard. In the Saturn I program, where more emphasis was placed on ground testing prior to the flight phase, 10 R&D flight tests were planned. The vehicle was declared operational after the first six firings met with success.

The Uprated Saturn I (Saturn IB) —an improve­ment on the basic Saturn I —was manrated after three flights. On the Saturn V, only two flights are planned prior to the attainment of a "manned con­figuration.”

The inspection to which flight hardware is subjected is thorough. Following are examples of many steps which are taken to inspect the Saturn V vehicle:

1. X-rays are used to scan fusion welds, 100 cast­ings, and 5,000 transistors and diodes.

2. A quarter mile of welding and 5 miles of tubing are inspected with the use of a sound technique (ultrasonics). The same type of inspection is given to adhesive bonds, which are equivalent in area to an acre.

3. An electrical current inspection method is used on 6 miles of tubing, and dye penetrant tests are run on 2.5 miles of welding.

Each contractor has his own test program that is patterned to a rather basic conservative approach. It begins with research to verify specific principles to be applied and materials to be used. After pro­duction starts each contractor puts flight hardware through qualification testing, reliability testing, development testing, acceptance testing, and flight testing.

QUALIFICATION TESTING

Qualification testing of parts, subassemblies, and assemblies is performed to assure that they are capable of meeting flight requirements. Tests under the conditions of vibration, high-intensity sound, heat, and cold are included.

RELIABILITY TESTING

Reliability analysis is conducted on rocket parts and assemblies to determine the range of failures or margins of error in each component. Reliability information is gathered and shared by the rocket industry.

DEVELOPMENT TESTING

A battleship test stage constructed more solidly than a flight stage is often used to prove major design parameters within a stage. Such a vehicle verifies propellant loading, tank and feed operation, and engine firing techniques.

Battleship testing is followed by all-systems test­ing. For example, one of four ground test stages of the first stage completed 15 firings at Marshall Space Flight Center in Huntsville. The firings proved that the design and fabrication of the complete booster and of its subsystems were adequate.

The entire Apollo/Saturn V vehicle, consisting of the three Saturn V propulsive stages, the instru­ment unit, and an Apollo spacecraft, was assembled in the Dynamic Test Stand at the Marshall Center. This is the only place, aside from the launch site, where the entire Saturn V vehicle has been assem­bled. The purpose of dynamic testing was to deter­mine the bending and vibration characteristics of the vehicle to verify the control system design. The 364-foot assembly was placed on a hydraulic bearing or “floating platform”. Electromechanical shakers caused the vehicle to vibrate, simulating the response expected from flight forces.

TYPICAL LUNAR LANDING MISSION

The jumping-off place for a trip to the moon is NASA’s Launch Complex 39 at the Kennedy Space Center. After the propellants are loaded, the three astronauts will enter the spacecraft and check out their equipment.

While the astronauts tick off the last minutes of the countdown in the command module, a large crew in the launch control center handles the complicated launch operations. For the last two minutes, the countdown is fully automatic.

At the end of countdown, the five F-l engines in the first stage ignite, producing 7.3 million pounds of thrust. The holddown arms release the vehicle, and three astronauts begin their ride to the moon.

Turhopumps, working together with the strength of 31) diesel locomotives, force 15 tons of fuel per second into the engines. Steadily increasing accel­
eration pushes the astronauts back into their couch­es as the rocket generates 1-1, If times the force of earth gravity.

After 2,5 minutes, the first stage has burned its

4,192,0 pounds of propellants and is discarded at about 38 miles altitude. The second stage’s five,1-2 engines are ignited. Speed at this moment is 5,330 miles per hour.

The second stage’s five 4-2 engines burn for about (5 minutes, pushing the Apollo spacecraft to an altitude of nearly 115 miles and near orbital velocity of 15,300 miles per hour. After burnout the second stage drops away and retrorockets slow it for its fall into the Atlantic Ocean west of Africa.

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The single 4-2 engine in the third stage now ignites and burns for 2.75 minutes. This brief burn boosts the spacecraft to orbital velocity, about 17,500 miles an hour. The spacecraft, with the third stage still attached, goes into orbit about 12 minutes after liftoff. Propellants in the third stage are not depleted when the engine is shut down. This stage stays with the spacecraft in earth orbit, for its en­gine will be needed again.

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Подпись: K-107P-66P-258 Throughout the launch phase of the mission, telem­etry systems are transmitting continously, track­ing systems are locked on, and voice communica­tions are used to keep in touch with the astronauts. All stage separations and engine thrust terminations are reported to the Mission Control Center at Houston.

The astronauts are now in a weightless condition as they circle the earth in a “parking orbit” until the timing is right for the next step to the moon.

The first attempt at a lunar landing is planned as an “open-ended” mission with detailed plans at every stage for mission termination if necessary. A comprehensive set of alternate flight plans will be laid out and fully rehearsed for use if such a de­cision should prove necessary. For example, a de­cision might be made in the earth parking ortm not to continue with the mission. At every stage of the mission, right up to touchdown on the moon, this termination decision can be made and an earth flight plan initiated.

During the one to three times the spacecraft circles the earth, the astronauts make a complete check of the third stage and the spacecraft. When the precise moment comes for injection into a trans­lunar trajectory, the third stage J-2 engine is re­ignited. Burning slightly over 5 minutes, it acceler­ates the spacecraft from its earth orbital speed of 17,500 miles an hour to about 24,500 miles an hour in a trajectory which would carry the astronauts around the moon. Without further thrust, the space­craft would return to earth for re-entry.

If everything is operating on schedule, the astro­nauts will turn their spacecraft around and dock with the lunar landing module. After the docking maneuver has been completed, the lunar module will be pulled out of the forward end of the third stage, which will be abandoned. Abandonment com­pletes the Saturn V’s work on the lunar mission.

Checkout Valve

The checkout valve consists of a ball, a poppet, and an actuator. The checkout valve provides for ground checkout of the ignition monitor valve and fuel valves and prevents the ground hydraulic return fuel, used during checkout, from entering the en­gine system and consequently the vehicle fuel tank.

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When performing the engine checkout or servicing, the checkout valve ball is positioned so fuel enter­ing the engine hydraulic return inlet port will be directed through the ball and out the GSL return port. For engine static firing or flight, the ball is positioned so fuel entering the engine hydraulic re­turn inlet port will be directed through the ball and out the engine return outlet port.

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Engine Control Valve

(Hydraulic Filter and Four-Way Solenoid Valve Manifold)

The engine control valve incorporates a filter mani­fold, a four-way solenoid valve, and two swing check valves.

The filter manifold contains three filters. One filter is in the supply system and one each in the opening and closing pressure systems. The filters prevent entry of foreign matter into the four-way solenoid valve or the engine. Two swing check valves are "teed” into the supply system filter. The check valves permit hydraulic system operation from the ground supplied hydraulic fluid for checkout and servicing procedures or engine supplied hydraulic fluid for normal engine operation.

The four-way solenoid valve is comprised of a main spool and sleeves to achieve two-directional control of the fluid flow to the main fuel, main oxidizer, and gas generator valve actuators. The spool is pressure-positioned by two three-way slave pilots. Each slave pilot has a solenoid-controlled, normally open, three-way primary pilot.

The de-energized position of the engine control valve provides hydraulic closing pressure to all engine propellant valves. Momentary application of 28 VDC to the start solenoid will initiate control valve actuations that culminate in the positioning of the main spool so that hydraulic pressure is applied to the opening port, and the pressure previously applied to the closing port is vented to the return port.

An internal passage in the housing maintains com­mon pressure applied between the opening port and start solenoid poppet. This pressure, after start solenoid de-energization, holds the main spool in its actuated position thereby maintaining the pres­sure directed to the opening port without further application of the start solenoid electrical signal. Momentary application of 28 VDC to the stop so­lenoid will initiate control valve actuations that culminate in positioning the main spool so that pres­sure is vented from the opening port and applied to the closing port. The override piston may be actuated at any time by a remote pressure supply, which, in the event of an electrical power loss, would re­position the main spool and apply hydraulic pres­sure to the closing port. If electrical power and hydraulic pow-er are both removed, the valve will return to the de-energized position by spring force. If hydraulic pressure is then reapplied, pressure will be applied to the closing port. If an electrical signal is simultaneously sent to the start and stop solenoids, the stop solenoid will override the start and return the valve to a deactuated position.

Swing Check Valve

There are two identical swing check valves installed on the engine control valve. They allow – the use of ground hydraulic fuel pressure during engine start­ing transient and engine hydraulic fuel pressure during engine mainstage and shutdown. One check valve is installed in the engine hydraulic fuel supply inlet port, the other in the ground hydraulic fuel supply inlet port.

ELECTRICAL SEQUENCE CONTROLLER

The electrical sequence controller is a completely self-contained, solid-state system, requiring only DC power and start and stop command signals.

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Pre-start status of all critical engine control func­tions is monitored in order to provide an “engine ready” signal. Upon obtaining “engine ready” and “start” signals, solenoid control valves are ener­gized in a precisely timed sequence as described in the “Engine Operation” section to bring the en­gine through ignition, transition, and into main – stage operation. After shutdown, the system auto­matically resets for a subsequent restart.

start Tank Assembly System

This system is made up of an integral helium and hydrogen start tank, which contains the hydrogen and helium gases for starting and operating the en­gine. The gaseous hydrogen imparts initial spin to the turbines and pumps prior to gas generator com­bustion, and the helium is used in the control system to sequence the engine valves.

HELIUM AND HYDROGEN TANKS

The spherical helium tank is positioned inside the hydrogen tank to minimize engine complexity. It holds 1,000 cubic inches of helium. The larger spher­ical hydrogen gas tank has a capacity of 7,257.6 cubic inches. Both tanks are filled from a ground source prior to launch and the gaseous hydrogen tank is refilled during engine operation from the thrust chamber fuel inlet manifold for subsequent restart in third stage application.

Flight Instrumentation System

The flight instrumentation system is composed of a primary instrumentation package and an auxiliary package.

FUEL LEVEL SENSING AND ENGINE CUTOFF SYSTEMS

A cutoff sensor mounted on the bottom of the fuel tank provides signal voltages to shut off fuel after a predetermined level of depletion is reached. The fuel is measured during flight by four fuel slosh probes and a single liquid level measuring probe. Fuel levels are detected electronically and reported through the stage telemetry system. Telemetry

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signals are transmitted to ground support either by radio frequency or, before launch, by coaxial cable. The cutoff sensor, mounted in the lower fuel tank bulkhead, initiates engine cutoff as fuel level falls below two sensing points on the probe. Engine cutoff will normally be initiated by sensors in the LOX system. The cutoff capability is provided as a backup system should fuel be depleted before LOX.

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Fuel Level Sensing and Engine Cutoff

FUEL PRESSURIZATION SYSTEM

The fuel pressurization system maintains enough pressure in the fuel tank to provide proper suction at the fuel turbopumps to start and operate en­gines. The system consists of a helium supply, a helium flow controller, helium fill and drain com­ponents, a prepressurization subsystem, a fuel tank vent and relief valve, and associated ducts.

Four 31-cubic-foot, high pressure storage bottles in the LOX tank store the helium required for in­flight pressurization of the fuel tank ullage. A high pressure line is used for filling the bottles and rout­ing the helium to the flow controller. A solenoid dump valve is installed for emergencies. The helium flow controller uses five solenoid valves mounted parallel in a manifold to control helium flow to the fuel tank ullage. The cold helium duct routes helium from the flow controller to the cold helium mani­fold. From there, it is distributed to the heat ex­changers on the five F-l engines. The hot helium manifold receives the heated, expanded helium from the engine heat exchangers and routes it to the hot helium duct which then carries it through the he­lium distributor and on to the fuel tank ullage.

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It is replenished between the periods of loading and prepressurization through the fill and drain line.

Before LOX drain can be performed, the helium cylinders in the LOX tank must have their pressure decreased from about 3,100 psig to about 1,660 psig. Fill and drain valves are opened to complete drainage of the LOX tank although total evacua­tion of LOX from the tank requires draining the engines or waiting for boil-off of residual LOX. LOX drain can be speeded with the aid of a pres­surizing gas, usually nitrogen.

Venting Subsystem

The venting subsystem is used during loading and flight operations. While the propellant tanks are being loaded, the vent valves (two for each tank) are opened by electrical signals from ground equipment to allow the gas created by propellant boil-off to leave the tanks. The valves are spring-loaded to be normally closed, but a relief valve will open them if pressure in the tanks reaches an excessive level. Each valve is capable of venting enough gas to

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The LH2 recirculation subsystem pumps the propel­lants through the feed lines and valves and back to the LIL tank through a single return line. The pumps are powered from a 56-volt DC battery sys­tem located in the interstage; the batteries are ejected with the interstage approximately 30 sec­onds after first plane separation. Refore liftoff, power for the LH2 recirculation subsystem is sup­plied by ground equipment.

The LOX recirculation system works on the basis of a thermal syphon; heat entering the system is used to provide pumping action by means of fluid density differences across the system. Helium gas is used to supplement the density differences and thereby improve the pumping action.

Recirculation of oxygen begins at the start of tank­ing; LIL recirculation begins just before launch. The propellants continue to circulate through first stage firing and up until just before the first stage and second stage separate. While the subsystems are operating, the LH2 prevalves which lead to the combustion chambers are closed; as soon as the re­circulation subsystem stops, the LH2 prevalves open and the engines ignite.

Propellant Management System

The propellant management system controls load­ing, flow rates, and measurement of the propel­lants. It includes propellant utilization, propellant loading, propellant mass indication, engine cutoff, and propellant level monitoring subsystems.

PROPELLANT UTILIZATION SUBSYSTEM

The propellant utilization subsystem controls the flow rates of liquid hydrogen and liquid oxygen in such a manner that both will be depleted simulta­neously. It controls the mixture ratio so as to min­imize propellant residuals (propellant left in the tanks) at engine cutoff. Propellant utilization bypass valves at the liquid oxygen turbopump outlets con­trol flow of liquid oxygen in relation to the liquid hydrogen remaining. Control of the engine mixture ratio increases the stage’s payload capability. The propellant utilization subsystem is interrelated with the propellant loading subsystem and uses some of the same tank sensors and ground checkout equip­ment.

PROPELLANT LOADING SUBSYSTEM

The loading subsystem is used to control propellant loading and maintain the quantity of propellants
in the tanks. Capacitance probes (sensors) running the full length of the propellant tanks sense liquid mass in the tanks and send signals to an airborne computer, which relays them to a ground computer to control loading. They also send signals to an airborne computer for the propellant utilization subsystem’s control of flow rates.

PROPELLANT MASS INDICATION SUBSYSTEM

The propellant mass indication subsystem is in­tegrated with the propellant loading subsystem and is used to send signals to the flight telemetry sys­tem for transmission to the ground. It utilizes pro­pellant loading sensors to determine propellant levels.

ENGINE CUTOFF SUBSYSTEM

The main function of the engine cutoff subsystem is to signal the depletion point of either propellant. It is an independent subsystem and consists of five sensors in each propellant tank and associated electronics. The sensors will initiate a signal to shut down the engines when two out of five sensors in the same tank signal that propellant is depleted.

NORTH AMERICAN ROCKETDYNE FACILITIES

F-l and J-2 engines for the Saturn V launch vehicle are manufactured at Rocketdyne’s main complex in Canoga Park, Calif. F-l static testing is conducted at the Edwards Field Laboratory located at the NASA Rocket Engine Test Site, Edwards, Calif., about 125 miles northeast of Los Angeles, and the J-2 is tested at Rocketdyne’s Santa Susana Field Laboratory located about 10 miles from Canoga Park. Rocketdyne operates the Neosho Facility (Missouri), which produces and tests subcompo­nents of the J-2 and F-l engines.

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F-l Test Stands—Three of six stands for testing F-l rocket engines or components at full thrust are visible in this aerial view of NASA Rocket Engine Test Site.

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F-l Test Firing… An F-l rocket engine developing 1,500,000

pounds of thrust is tested at NASA Rocket Engine Test Site. The stand is one of six in the complex.

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Manufacturing of components and final assembly of both engines are carried out in eight buildings in

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the Canoga complex. These facilities are equipped with general purpose machine tools for precision and heavy machining as well as some 20 numerically controlled machines for performing programmed multiple machining operations. Also included are two of the largest gas-fired brazing furnaces of their type for brazing of thrust chamber tubes and in­jectors, eight units for ultrasonic cleaning, 21 in­stallations for Gamma and X-ray inspection, more than 50 environmentally controlled areas for ultra­clean assembly operations, sheet metal prepara­tion, precision cleaning, and receiving inspection.

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F-l Flight Engine Firing

An Engineering Development Laboratory pro­vides specialized facilities to support manufacturing programs. These facilities include a high-flow water test facility for checking propellant systems, 12 concrete cells for conducting hazardous tests, 28 environmental test chambers, a photo-elastic lab­oratory, two pneumatic flow benches, six vibration test rooms, and others for checking components as well as complete engines.

Research and development testing of F-l turbo­machinery, gas generators, heat-exchangers, seals, and splines is conducted on two test stands and three components test laboratories at Santa Susana.

Six large test stands, with a total of eight test posi­tions, and associated shops and support facilities at the Edwards Field Laboratory are used for testing complete F-l engines as well as injectors.

Six large engine test stand positions at the Santa Susana Field Laboratory are used for testing the J-2. One of these stands is equipped with a steam injection diffuser for altitude simulation testing. J-2 turbopumps, gas generators, valves seals, bear­ings, and other components are tested in 22 test cells in five component test laboratories in Santa Susana.

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Pump Tests-Flames from gases burned during test of an F-l engine turbopump shoot more than 150 feet in air.

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J-2 Testing—A hydrogen fueled J-2 rocket engine is tested under ambient altitude conditions at Santa Susana.

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F-l ENGINE

ENGINE DESCRIPTION

The F-1 engine is a single-start. 1,500.000-pound fixed-thrust, bipropeliant rocket system. The en­gine uses liquid oxygen as the oxidizer and RP-1 (keroseneI as fuel. The engine is bell-shaped, with an area expansion ratio—the ratio of the area of the throat to the base—of Hi:L RP-1 and LOX are com­bined and burned in the engine’s thrust chamber as­sembly. The burning gases are expelled through an expansion nozzle to produce thrust. The five-engine cluster used on the first stage of the Saturn V pro­duces 7.500,000 pounds of thrust. All of the engines are identical with one exception. The four outboard engines gimbal; the center engine does not.

The major engine systems are the thrust chamber assembly, the propellant feed system, the turbo-

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Assembly Thrust chambers of the F-i rocket engine—the most powerful engine under development by the United States – are assembled in this manufacturing line.

pump, the gas generator system, the propellant tank pressurization system, the electrical system, the hydraulic control system, and the flight instru­mentation system.

THRUST CHAMBER ASSEMBLY

The thrust chamber assembly consists of a gimbal bearing, an oxidizer dome, an injector, a thrust chamber body, a thrust chamber nozzle extension, and thermal insulation. The thrust chamber as­sembly receives propellants under pressure sup­plied by the turbopump, mixes and burns them, and imparts a high velocity to the expelled combus­tion gases to produce thrust. The thrust chamber assembly also serves as a mount or support for all engine hardware.

Gimbal Bearing

The gimbal bearing secures the thrust chamber assembly to the vehicle thrust frame and is mounted on the oxidizer dome. The gimbal is a spherical, universal joint consisting of a socket-type bearing with a bonded Teflon-fiberglass insert which pro­vides a low-friction bearing surface. It permits a maximum pivotal movement of <i degrees in each direction of both the X and Zaxes (roughly analogous to pitch and yaw! to facilitate thrust vector control. The gimbal transmits engine thrust to the vehicle and provides capability for positioning and thrust alignment.

AUXILIARY PROPULSION SYSTEM

The APS provides auxiliary propulsive thrust to the stage for three-axis attitude control and for ullage control. Two APS modules are mounted 180c apart on the aft skirt assembly. Two solid pro­pellant rocket motors are mounted 180° apart be­tween the APS modules on the aft skirt assembly and provide additional thrust for ullage control.

APS Modules

Each APS module contains three 150-pound-thrust

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The attitude control engines are fired upon com­mand from the IU in short duration bursts for atti­tude control of the stage during the orbital coast phase of flight. Minimum engine-firing pulse-dura­tion is approximately 70 milliseconds. The attitude control engines are approximately 15 inches long with exit cones approximately 6.5 inches in diam­eter. Engine cooling is accomplished by an ablative process.

The ullage control engines are fired also upon com­mand from the IU during the transition between J-2 engine first burn and the coast phase of flight to prevent undesirable propellant movement within the tanks. Firing continues for approximately 50 seconds until activation of the LH, continuous pro­pulsive vent system. The ullage engines are again fired at the end of the third stage coast phase of flight and prior to J-2 engine restart to assure pro­per propellant positioning at inlets to the propellant feed lines during propellant tank repressurization.

The ullage control engines are similar to the atti­tude control engines and are approximately 15 inches long wuth an exit cone approximately 5.75 inches in diameter. Engine cooling is accomplished by an ablative process.

Each APS module contains an oxidizer system, fuel system, and pressurization system. The modules are self-contained and easily detached for separate checkout and environmental testing.

An ignition system is unnecessary because fuel and oxidizer are hypergolic (self-igniting). Nitrogen tetroxide lN,04), the oxidizer, is stable at room temperature.

Separate fuel and oxidizer tanks of the expulsion bellows type are mounted within the APS module along with a high-pressure helium bottle, which provides pressurization for both the propellant tanks and the associated plumbing and control systems.

The fuel, monomethyl hydrazine (CH. NTH.,), is stable to shock and extreme heat or cold. The APS module carries approximately 115 pounds of usable fuel and about 150 pounds of usable oxidizer.

Ullage Control

Two solid propellant Thiokol TX-280 rocket motors, each rated at 3,390 pounds of thrust, are ignited during separation of the second and third stages for ullage control approximately 4 seconds before J-2 ignition. This thrust produces additional positive stage acceleration during separation and positions LOX and LH2 propellants toward the aft end of the tanks. In addition, propellant boil-off vapors are forced to the forward end where they are safely vented overboard. Tank outlets are covered to en­sure a net positive suction head (NPSH) to the pro­pellant pumps, thus preventing possible pump cavi­tation during J-2 engine start. Ullage rockets ig­nite upon command from the stage sequencer and fire for approximately 4 seconds. At about 12 sec­onds from ignition, the complete rocket motor as­semblies, including bracketry, are jettisoned from the stage, upon command from the stage sequencer.

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Electrical Power and Distribution System

Four battery-powered systems provide electrical requirements for third stage operation. Forward Power System No. 1 includes a 28 VDC battery and power distribution equipment for telemetry, secure range receiver No. 1, forward battery heaters, and a power switch selector located in the forward skirt area.

Forward Power System No. 2 includes a 28 VDC battery and power distribution equipment for the PU assembly, inverter-converter, and secure range receiver No. 2.

Aft Power System No. 1 includes a 28 VDC battery and power distribution equipment for the J-2 en­gine, pressurization systems, APS modules, TM signal power, aft battery heaters, hydraulic system valves, and stage sequencer.

Aft Power System No. 2 includes a 56 VDC battery and power distribution equipment for the auxiliary hydraulic pump, oxidizer chilldowm inverter, and fuel chilldowm inverter.

Silver-oxide, zinc batteries used for electrical power and distribution systems are manually activated. The batteries are “one-shot” units, and not inter­changeable due to different load requirements.

Electrical power and distribution systems are switched from ground power to the batteries by­command through the aft umbilical prior to liftoff.

Telemetry and Instrumentation System

Radio frequency telemetry systems are used for transmission of stage instrumentation information to ground receiving stations. Five transmitters, using two separate antenna systems, are capable of returning information on 45 continuous output data channels during third stage flight. The telem­etry transmission links consist of five systems using three basic modulation schemes: Pulse Amplitude Modulated/FM/FM (PAM/FM/FM); Single Side – band/FM (SS/FM); and Pulse Code Modulated/FM (PCM/FM). There are three separate systems using PAM/FM/FM modulation.

A Digital Data Acquisition System (DDAS) air­borne tape recorder stores sampled data normally – lost during staging and over-the-horizon periods of orbital missions, and plays back information w-hen in range of ground stations.

ACCEPTANCE TESTING

Finished work undergoes functional checkout to insure it meets operational requirements. Tests range from continuity and compatibility of wiring to all-systems ground testing. Fluid-carrying com­ponents are subjected to pressures beyond normal operating requirements, and structural components receive visual and X-ray inspections. Instruments simulate flight conditions to evaluate total per­formance of electrical and mechanical equipment.

Rocket engines are static-fired before delivery to the stage contractor. Such tests demonstrate per-

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ASSEMBLY AND CHECKOUT

Saturn V stages are shipped to the Kennedy Space Center by ocean-going vessels or by specially de­signed aircraft. Apollo spacecraft modules are transported by air and delivered to the Manned Spacecraft Operations Building at Kennedy Space Center for servicing and checkout before mating with the Saturn V.

Saturn V stages go into the Vehicle Assembly Building low bay area where preparation and check­out begins. Receiving inspection and the low bay checkout operations are first performed before stages are erected within a high bay.

After being towed into the high bay area and posi­tioned under the 250-ton overhead bridge crane, slings are attached to the first stage and hooked to the crane. The stage is positioned above the launch platform of the mobile launcher and lowered into place. Then it is secured to four holddown/support arms. These support the entire space vehicle dur­ing launch preparation and provide holddown dur­ing thrust buildup prior to launch.

Next, engine fairings are installed on the stage and fins are moved into position and installed in line with the four outboard engines.

Mobile launcher electrical ground support equip­ment is connected to the ‘launch control center (LCC) via the high speed data link, and the test pro­gram is started with the actual launch control equip­ment.

Prior to and during this time, all low bay testing is completed and the upper stages are prepared for mating. The mating operation consists of stacking the stages. Umbilical connection begins immediately and continues during the mating operation on a noninterference basis. The vertical alignment of the vehicle is performed after each stage is mated.

When the launch vehicle is ready, the Apollo space­craft is brought to the VAB and mated.

Checkout of all systems is performed concurrently in the high bay. The first tests provide power and cooling capability to the vehicle, validate the con­nections, and establish instrumentation. When this is completed, systems testing begins. The systems tests are controlled and monitored from the LCC wherever practical and “break-in” tests are held to a minimum. Following the validation of each stage, a data review7 is held and the vehicle is pre­pared for combined systems tests.

Illustration of Vehicle Assembly Building Interior at Kennedy Space Center

The combined systems tests verify the flight-readi­ness of the overall vehicle. These tests include a malfunction sequence test, an overall test of the launch vehicle, an overall test of the spacecraft, and a simulated flight test. Prior to the simulated flight test, final ordnance installation is completed. After the test, vertical alignment is checked, a data review is held, and the vehicle is prepared for trans­fer to the pad. These preparations include discon­necting pneumatic, hydraulic, and electrical lines from the mobile launcher to the VAB.

After the lines are disconnected, the transporter is moved into position beneath the mobile launcher. Hydraulic jacks engage the fittings on the mobile launcher and raise it approximately 3 feet so that it clears its mount mechanisms. Then the transport­er moves out of the VAB, over the crawlerway, to the launch pad.