Category Apollo Saturn V News Reference

The hydraulic system performs engine positioning upon command from the IU. Major components are a J-2 engine-driven hydraulic pump, two hydraulic actuator assemblies, and an accumulator-reservoir assembly.

The electrically driven auxiliary hydraulic pump is started before vehicle liftoff to pressurize the hy­draulic system. Electric power for the pump is provided by a ground source. At liftoff, the pump is switched to stage battery power. Pressurization of the hydraulic system restrains the J-2 engine in a null position with relation to the third stage eenter-

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line, preventing pendulum-like shifting from forces encountered during liftoff and boost. During power­ed flight, the J-2 engine may be gimbaled up to 7° in a square pattern by the hydraulic system upon command from the IU.

Engine-Driven Hydraulic Pump

The engine-driven hydraulic pump is a variable dis­placement type pump capable of delivering hy­draulic fluid under continuous system pressure and varying volume as required for operation of the hy­draulic actuator assemblies. The pump is driven directly from the engine oxidizer turbopump. A thermal isolator in the system controls hydraulic – fluid temperature to ensure proper operation.

Auxiliary Hydraulic Pump

The auxiliary hydraulic pump is an electrically driven variable displacement pump which supplies a constant minimum supply of hydraulic fluid to the hydraulic system at all times. The pump is also used to perform preflight engine gimbaling check­outs, hydraulically lock the engine in the null posi­tion during boost phase, maintain system hydraulic – fluid at operating temperatures during other than the powered phase, and augment the engine-driven hydraulic pump during powered flight. It also pro­vides an emergency backup supply of fluid to the system.

Hydraulic Actuator Assemblies

Two hydraulic actuator assemblies are attached directly to the J-2 engine and the thrust structure and receive IU command signals to gimba! the en­gine. The actuator assemblies are identical and interchangeable.

Accumulator-Reservoir Assembly

The accumulator-reservoir assembly is an integral unit mounted on the thrust structure. The reservoir section is the storage area for hydraulic fluid; the accumulator section supplies peak system fluid re­quirements and dampens high-pressure surges with­in the system.

The ECS cools the electronic equipment in the IU and the forward third stage skirt. Sixteen cold plates are installed in each stage.

An antifreeze-like coolant, 60 per cent methanol and 40 per cent water, from a reservoir within the IU is circulated through the cold plates. Heat gen­erated by the mounted components is transferred to the coolant by means of conduction.

Prior to liftoff a preflight heat exchanger serviced by ground support equipment transfers heat from the coolant. Approximately 163 seconds after lift­off, ECS’s sublimator-heat exchanger takes over the job of temperature control.

Some of the more complex components like the guidance computer, flight control computer, and the ST-124-M platform, have coolant fluid circulated through them to provide more efficient heat removal.

In the vacuum of space the warmed coolant, after leaving the cold plates, is routed through a device called a sublimator. Water, from an IU reservoir,

Structure Segments—Prior to splicing, mounting brackets for thermal conditioning panels can be seen on interior surface of segments. The exterior of the spring-loaded umbilical door and the access door are visible at right center.

Extremely accurate theodolites, similar to a sur­veyor’s transit, are used to align the segments in a circle prior to splicing. Metal splicing plates join the three segments, and the holes which permit the IU to be joined to mating surfaces of the launch vehicle are drilled at top and bottom edges of the structure for ease in handling. Protective rings are bolted to these edges to stiffen the structure. Vehicle antenna holes are cut after splices are bolted.

After structure fabrication is completed, module and component assembly operations begin. Tem­perature transducers are fastened to the inner skin, environmental control system (ECS) cold plates are mounted, and a cable tray is bolted to the top of the

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Splice Joint Operations—Final grinding of a splice joint ensures a smooth surface prior to splice plate assembly.

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goes to the sublimator and is exposed through a porous plate to the low temperature and pressure of outer space where it freezes, blocking the pores in the plate. The heat from the coolant, transferred to the plate, is absorbed by the ice converting it directly into water vapor (a process called sub­limation).

The system is self-regulating. The rate of heat dis­sipation varies with the amount of heat input, speed­ing up or slowing down as heat is generated. If the coolant temperature falls below a pre-set level, an

Instrument Unit Assembly in IBM Manufacturing Area – Splicing operations and assembly of the tubular cable tray are complete, the cold plates have been installed, and installation of com­ponents is underway.

Environment Control—A mobile clean room protects against contamination during assembly of environmental control system components. Gaseous nitrogen will be circulated from a ground supply through the duct partially assembled in the cable tray to purge the IU following vehicle fueling.

electronically controlled valve causes the coolant mixture to bypass the sublimator until the tem­perature rises sufficiently to require further cooling.

Nitrogen gas provides artificial pressure for both coolant solution and sublimator water reservoirs during orbit.

A coolant circulating pump along with the necessary valves and piping to control flow complete the en­vironmental control equipment.

Guidance and Control

The IU’s guidance and flight control systems nav-

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igate (determine vehicle position and velocity), guide (determine attitude correction signals), and control (determine and issue control commands to the engine actuators) the Saturn V vehicle.

Completely self-contained, these systems measure acceleration and vehicle attitude, determine velocity and position and their effect on the mission, cal­culate attitude correction signals, and determine

and issue control commands to the engine actuators.

All this is done to place the vehicle in a desired attitude to reach the required velocity and altitude for mission completion.

Major components are an inertial platform, the launch vehicle digital computer (LVDC), the launch vehicle data adapter (LVDA), an analog flight con­trol computer, and control and rate gyros.

IBM-DR-4

ST-124-M Inertia! Platform System

Prior to liftoff, launch parameters go to the LVDC.

About five seconds before liftoff, the inertial guid­ance platform and the LVDC are released from ground control. As the vehicle ascends, the guid­ance platform senses and measures vehicle accel­eration and attitude and sends these measurements to the LVDC via the LVDA.

The LVDC integrates these measurements with the time since launch to determine vehicle position relative to starting point and destination. It then computes the desired vehicle attitude, using data stored in its memory, and the difference between the desired attitude and the actual becomes the generated attitude correction signal.

This signal is sent to the analog flight control com­puter, where it is combined with information from rate gyros. Using this data, the flight control com­puter determines and issues the command to gimbal the engines and change the thrust direction.

Each mission has at least three phases: atmospheric – powered flight, boost period after initial entry into

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space, and the coasting period.

Atmospheric boost causes the greatest vehicle load because of atmospheric pressure. During this time the guidance system is primarily checking vehicle integrity and is programmed to minimize this pres­sure.

Guidance and Control—The LVDC and LVDA portion of the guidance system is shown in this block diagram. The LVDC receives information from all parts of the vehicle via the LVDA, and in turn issues commands.

The vehicle maintains liftoff orientation long enough to clear the launch equipment, and then it performs a roll maneuver to get to the flight azimuth direc­tion.

The time tilt program is applied after the roll ma­neuver. The pitch angle is regulated by the tilt program, and is independent of navigation measure­ments. However, navigation measurements and computations are performed throughout the flight, beginning at the time the platform is released (i. e., five seconds before liftoff). First stage engine cut­off and stage separation are commanded when the IU receives a signal that the tank’s fuel level has reached a predetermined point. During second stage powered flight the LVDC guides the vehicle via the best path to reach the mission objectives.

During orbit, navigation and guidance information in the LVDC can be updated by data transmission from ground stations through the IU radio com­mand system.

Approximately once every two seconds, the LVDC, using iterative or "closed loop” guidance, figures vehicle position and vehicle conditions required at the end of powered flight (velocity, altitude, etc.) and generates the attitude correction signals to gimbal the engines so that the vehicle reaches its predetermined parking orbit.

III Interior During Assembly—The large, cylindrical component simulates size and shape of the flight control computer and is used to check cable lengths and mounting arrangement.

Second stage engine cutoff comes when the IU is signaled that stage propellant has reached a pre­determined level, and then the stage is separated. By this time, the vehicle has already reached its approximate orbital altitude, and the third stage burn merely gives it enough push to reach a cir­cular parking orbit.

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.

Four fairings attach to the thrust structure and partially surround the outboard engines at the foot of the booster. They house the eight retrorockets and the actuator support structures. Fairings are shaped like cone halves and are constructed of alu­minum. Their purpose is to smooth the air flow over the engines.

The fins are airfoil attachments to the fairings. Fins are rigid and add to the vehicle’s flight stabil­ity. A titanium skin covers the fin for greatest protection against temperatures as high as 2,000 degrees Fahrenheit.

Each of the eight retrorockets generates about 86,600 pounds of thrust for two-thirds of a second

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Tube and Valve Cleaning Vat – Each stage component is treated in a cleaning solution before final assembly.

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-

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.

The liquid oxygen (LOX) tank is an ellipsoidal con­tainer 22 feet high and fabricated from ellipsoidal­shaped top and aft halves. The top half of the LOX tank is known as the common bulkhead and is actu­ally two bulkheads separated by phenolic honey­comb insulation and bonded together to form both the upper portion of the liquid oxygen tank and the lower portion of the liquid hydrogen tank.

All of the LOX tank bulkheads are formed by weld­ing together 12 high-energy-formed curved sections (gores), each approximately 20 feet long and 8 feet

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wide. When the gores are welded together, an opening is formed at the apex of the bulkhead. The apex is closed by welding the 12 gores to a circular section called a dollar section.

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.

To ensure the accuracy and reliability of guidance information, critical LVDC circuits are provided in triplicate. Known as triple modular redundancy (TMR), the system corrects for failure or inaccuracy by providing three identical circuits. Each circuit produces an output which is voted upon. In case of a discrepancy, the majority rules, and a random failure or error can be ignored. In addition, the LVDC has a duplexed memory, and if an error is found in one portion of the memory, the required output is obtained from the other and correct infor­mation read back into both memories, thus correct­ing the error.

The ST-124-M inertial platform provides signals representing vehicle attitude. Since a signal error could produce vast changes in ultimate position, component friction must be minimized. Therefore,
the platform bearings are floated in a thin film of dry nitrogen supplied at a controlled pressure and flowrate from reservoirs within the IU.

PRELAUNCH FUNCTIONS

In addition to guidance computations, other func­tions are performed by the LVDC and the LVDA. During prelaunch, the units conduct test programs. After liftoff they direct engine ignition and cutoff, direct stage separations, and conduct reasonable­ness tests of vehicle performance. During earth orbit, the computer directs attitude control, con­ducts tests, isolates malfunctions, and controls transmission of data, plus the sequencing of all events.

Instrumentation

A basic requirement for vehicle performance anal­ysis and for planning future missions is knowing what happened during all phases of flight and just how the vehicle reacted. The Ill’s measuring and telemetry equipment reports these facts. Measur­ing sensors or transducers are located throughout the vehicle monitoring environment and systems’ performance.

Measurements are made of mechanical movements, atmospheric pressures, sound levels, temperatures, and vibrations and are transformed into electrical signals. Measurements also are made of electrical signals, such as voltage, currents, and frequencies which are used to determine sequence of stage sep­aration, engine cutoff, and other flight events and to determine performance of onboard equipment.

In all, the IU makes several hundred measurements. A wide variety of sensors are used to obtain all kinds of information required: acoustic transducers monitor sound levels; resistor or thermistor trans-

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ducers monitor temperature environments; bourdon – tube or bellows transducers measure pressures; force-balance, or piezoelectric accelerometers mea­sure force levels at critical points; flow meters de­termine rates of fluid flow.

Various measuring devices produce a variety of outputs, and before these outputs can be effectively utilized, they must be standardized to some extent. Signal conditioning modules are employed to adapt transducer outputs to a uniform range of 0-5 VDC.

Different types of data require different modes of transmission, and the telemetry portion of the sys­tem provides three such modes: SS/FM, FM/FM, and PCM/FM. Each type of information is routed to the most suitable telemetry equipment; a routing is performed by the measuring racks within the IU.

To get the most out of the transmission equipment, multiplexing is employed on some telemetry chan­nels. Information originated by various measuring devices is repeatedly sampled by multiplexers, or commutators, and successive samples from dif­ferent sources are transmitted to earth.

Information sent over any channel represents a series of measurements made at different vehicle points. This time-sharing permits large chunks of data to be handled with a minimum amount of equip­ment. The LVDC also helps in data transmission.

For instance, when the vehicle is between ground receiving stations, the LVDC stores important PCM data for later transmission. Once the vehicle leaves the earth’s atmosphere, sound levels requir­ing air for continuance no longer exist. The LVDC signals a measuring distributor to switch from un­important measurements to those more critical to

Typical Saturn Measuring System 7-6

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the mission. And during stage separation retro – rocket firing, when flame attenuation distorts or destroys telemetry transmissions, signals are auto­matically recorded by an onboard tape recorder, and transmitted later.

In order to monitor vehicle performance, ground controllers must know the vehicle’s precise posi­tion at all times. The RF section of the instrumenta­tion system provides this capability, as well as linking the IU’s guidance and control equipment during flight.

TRACKING SYSTEM

Several tracking systems are used to follow ve­hicle trajectory during ascent and orbit. Consolida­tion of this data not only increases data reliability, but gives the best trajectory information.

Vehicle antennas and transponders, which increase ground-base tracking systems’ range and accuracy, make up the IU’s tracking equipment.

Saturn V Instrument Unit Command System

A pulse or series of pulses of RF energy sent by ground stations to the vehicle’s general direction will interrogate the airborne transponder. In re­sponse, the transponder produces a pulse or series of pulses. Triangulation between precisely located ground stations determines point of origin of these reply pulses and fixes location of the vehicle.

Three tracking systems are employed in conjunc­tion with the Saturn V IU: AZUSA, C-band radar, and the S-band portion of the command and com­munication system (CCS). Two C-band transpon­ders are employed to provide tracking capabilities for this system independent of vehicle attitude. A single transponder is employed with the AZUSA system.

Real-time navigation, needed to update the guid­ance system, is received in the IU by a radio com­mand link. But before it is sent, and before it is

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accepted in the IU, both ground equipment and IU instruments scrutinize update information for accu­racy. The slightest error in transmission could con­ceivably produce a greater problem than if the original data had been left alone.

The message goes from antenna to command re­ceiver for amplification and demodulation. Then it is routed to the decoder for breakout into the origi­nal pattern of digital bits.

The first validity check is here. If there is an error in a bit, or a bit is missing, the entire message is rejected. Accepted commands get further checking in the command decoder and in the LVDC.

First the vehicle address is checked in the com­mand decoder. This is important because commands for both IU instruments and the spacecraft use simi­lar command links. If the spacecraft address is rec­ognized, the IU ignores the message.

Passing this test, the message is sent to the LVDC. Upon receipt, the LVDC tests the message to deter­mine if it is proper. If it is, then the command de­coder releases a pulse via the telemetry link to the ground station verifying message acceptance. If the message fails the test, the LVDC rejects it and telemeters an error message.

Depending on the mission, several types of mes­sages can be processed. For example: commands to perform updating, commands to perform tests, com­mands to perform special subroutines or special modes of operation, a command to dump or clear certain sectors of the computer memory, or a com­mand to relay a particular address in the computer memory to the ground. Provisions have been made to expand the number of types of messages if ex­perience indicates this is necessary.

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.

When all major components of the first stage are assembled in NASA’s Michoud Assembly Facility, they are routed to the Vertical Assembly Building to be assembled.

Manipulated by an overhead crane, the components are placed in final assembly position in the single­story building rising the equivalent of 18 stories.

First the thrust structure is placed on four heavy pylons 20 feet above floor level. Meanwhile, two of the segments—the fuel and LOX tanks which are brought to the Vertical Assembly Building in seg­ments—are being completed on two tank assembly bays. Then, in building-block fashion, the thrust structure is joined by the fuel tank, intertank, LOX tank, and forward skirt. When the forward skirt is secured, the first stage stands 138 feet high.

Vertical assembly completed, the 180-ton-capacity overhead crane lifts the booster by a forward han­dling ring attached to the forward skirt and re­turns it to horizontal position on its 435,000-pound transporter.

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Vertical Assembly—Booster sections are mated in the Vertical Assembly Building. At top left the thrust structure is shown. Fuel tank, intertank assembly, LOX tank, and forward skirt are added in successive pictures.

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As assembly jobs approach completion, installation of internal systems and engines is made in prepara­tion for systems test and checkout.

Engines—One of the first stage’s F-l engines is mounted. To­gether the five will consume 4,492,000 pounds of propellants in 2.5 minutes.