Category Apollo Saturn V News Reference

Liquid Oxygen Tank

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.

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Second Stage LOX 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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Tank Fabrication—Workmen close out dollar section of propel­lant tank.

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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.

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.

TRIPLE RELIABILITY

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.

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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

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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.

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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.

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.

Vertical Assembly

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.

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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.

Oxidizer Dome

The oxidizer dome serves as a manifold for dis­tributing oxidizer to the thrust chamber injector, provides a mounting surface for the gimbal bearing, and transmits engine thrust forces to the vehicle structure. Oxidizer at a volume flowrate of 2-1.811 gpm enters the dome through two inlets positioned 180 degrees apart (to maintain even distribution of the propellant!.

Thrust Chamber Injector

The thrust chamber injector directs fuel and oxi­dizer into the thrust chamber in a pattern which ensures efficient and satisfactory combustion. The injector is multi-orificed with copper fuel rings and copper oxidizer rings forming the face (combustion side! of the injector and containing the injection orifice pattern. Assembled to the face are radial and circumferential copper baffles which extend down-

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ward and compartmentalize the injector face. The baffles and rings, together with a segregated ig­niter fuel system, are installed in a stainless steel body.

Oxidizer enters the injector from the oxidizer dome. Fuel enters the injector from the thrust chamber fuel inlet manifold, and in order to facilitate the engine start phase and to reduce pressure losses, part of the flow is introduced directly into the thrust chamber. The remaining fuel (controlled by ori­fices) flows through alternate tubes which run the length of the thrust chamber body to the nozzle exit. There, it enters a return manifold and flows back to the injector through the remaining tubes.

AFT LOX BULKHEAD

The aft LOX bulkhead, like the aft facing sheet of the common bulkhead, is composed of 12 thin aluminum gores welded to mechanically milled waffle panels. The waffle panels are sheets into which diagonal ribs are machined to form a series of diamonds. The waffle panels are used around the middle (widest part of the LOX tank) to provide structural strength. Baffles adjacent to the aft facing sheet of the com­mon bulkhead prevent wave action (sloshing) dur­ing flight. At the lower apex of the LOX tank, anti­vortex baffles, consisting of a 14-foot cruciform (four fins arranged in a cross) and 12 smaller baffles, are installed over the sump and engine supply line con­nections. The smaller baffles are essentially thin metal plates extending from the center of the cruci­form, three between each pair of fins.

COMMON BULKHEAD

The common bulkhead may be likened to two giant domes, one placed inside the other, open end down, with a layer of insulation sandwiched between. The top dome is called the forward facing sheet and the

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Bulkhead-Common bulkhead shows aft facing sheet (in sling preparatory to mating).

bottom, the aft facing sheet. The forward facing sheet has a J-shaped periphery, which is welded to the No. 1 liquid hydrogen tank cylinder. In final assembly, a 15-inch, 12-section bolting ring is bolted to the aft skirt and the No. 1 liquid hydrogen tank
cylinder. A total of 636 bolts attach the bolting ring to the liquid oxygen tank.

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Gore Section of Aft Facing Sheet of Common Bulkhead Before Assembly

Insulating and joining the forward and aft facing sheets into a common bulkhead is a process of sev­eral operations. First the aft facing sheet is placed on a bonding fixture and numerous sections of hon­eycomb phenolic insulation are fitted and tapered to exact but varying thicknesses. Then the insula­tion is cemented to the aft facing sheet in a multi­stage bonding operation which includes chemical processing of the aft facing sheet, application of adhesive, and pressurizing and curing in the auto­clave. After mating the forward facing sheet over the insulated aft facing sheet, impression checks are made to assure a perfect fit. The forward facing sheet is then chemically processed, the insulation placed on the exposed top of the aft facing sheet is prepared with adhesive, and the entire bulkhead assembly is joined and placed in the autoclave for pressurizing and curing. In both bonding operations, checks are performed with ultrasonic equipment to ensure that the adhesive has completely covered the surface.

Basic PCM Digital Data Acquisition System PAM/FM/FM SYSTEMS

Transducer input signals constitute the PAM input. The PAM systems use an electronically switched network that samples up to 30 channels of trans­ducer inputs at 120 times a second. Deviations in transducer input voltages are represented as out­put pulses of varying amplitude for subsequent evaluation.

Подпись: SATURN V NEWS REFERENCE SS/FM SYSTEM ment Unit section for a complete description of the IU environmental conditioning system.

Подпись:The SS/FM system is reserved for pertinent re­search requirements. Vibration and acoustical data needed for manned flight development will be trans­mitted by this system.

PCM/FM SYSTEM

The PCM/FM system (DDAS) is used during auto­matic checkout to provide data for the ground check­out computer. The system is also used to provide precise information concerning stage environment and performance of systems during flight.

Environmental Control Systems

AFT SKIRT AND INTERSTAGE THERMOCONDITIONING AND PURGE

The thermoconditioning and purge system purges the aft skirt and aft interstage of combustible gases andj distributes temperature controlled air or gas­eous nitrogen around electrical equipment in the aft skirt during the vehicle countdown.

The purging gas, supplied from a ground source through the umbilical, passes over the electrical equipment and flows into the aft interstage area. Some of the gas is directed through each of the auxiliary propulsion modules and exhausts into the interstage. A duct from the skirt manifold directs air or GN, to a thrust structure manifold. From the thrust structure manifold supply duct, a portion of air or GN, is directed to a shroud covering the hydraulic accumulator reservoir.

Temperature control is accomplished by two dual­element thermistor assemblies located in the gas­eous exhaust stream of each of the auxiliary propul­sion modules. Elements are wired in series to sense average temperature. Two series circuits are formed, each circuit utilizing one element from each thermistor assembly. One series is used for tem­perature control, the other for temperature record­ing.

SWITCH SELECTOR

All stages, and the IU, are equipped with a switch selector. This unit has electronic and electrome­chanical components which decode LVDC/LVDA sequence commands and switches them to the prop­er circuits en each stage. This system has several advantages: reduction of stage interface lines, in­creased flexibility with respect to timing and se­quencing, and conserving the discrete output circui­try in the LVDC/LVDA. Sequencing commands can come as fast as every 100 milliseconds.

Stage power isolation is maintained in the switch selector by using relays as the input circuit. The relays are driven by IU power, while the decoding circuitry and driver output are powered from the parent stage. Input and output are coupled through relay contacts. These contacts drive a diode matrix used in decoding the 8-bit input code to select the output driver, producing the switch selector output.

There also is a check and proceed system built into the switch selector. After the switch selector re­lays have been “picked,” the complement of the received message is fed back to the LVDA/LVDC where it is checked. If the feedback is good, a read command is issued. If there is disagreement, a new message is sent which accomplishes the same func­tion. (Note: For redundancy, two messages’ codes are assigned for each switch selector output).

Electrical System

The electrical system powers the IU’s equipment. As with most of the IU’s systems, the electrical system is divided into two sectors: prelaunch and flight. Ground sources provide power through the umbilical lines before launch. At approximately 25 seconds prior to liftoff, power is transferred to the four 28 VDC IU batteries. Each battery has a 350 ampere hour capacity, and loads are equally distributed to drain.

Two special power supplies are provided: a 5-volt master measuring voltage supply converts 28 VDC main supply to a highly regulated 5 VDC for refer­ence and supply voltage to the measuring com­ponents, and a 56-volt power supply for operation of the guidance system’s ST-124-M inertial platform and the platform’s AC power supply.

In order to get the most out of the battery stored power during flight, the LVDC and LVDA turn off unused or unimportant circuits in favor of more important applications as the mission progresses.

TESTING AT LAUNCH SITE

At the launch pad, the transporter moves the mobile launcher into position, lowers and locks it onto another set of mount mechanisms. The transporter then moves to the mobile service structure parking area, picks up the service tower, and positions it beside the Saturn V to provide vehicle access for pad operations.

The digital data link, communications circuitry, pneumatic supply lines, propellant lines, environ­mental controls, and electrical power supply lines are connected.

Power again is applied to the vehicle and the con-

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trol and monitor links are verified. Pad testing is held to a minimum. The high hay from which the vehicle was moved remains empty during pad oper­ations.

A spacecraft systems verification test is performed, followed by a space vehicle cutoff and malfunction test. Radio frequency compatibility is established and preparations are made for a final flight readi­ness test, which involves sequence tests paralleling the actual countdown and inflight operations. Com­patibility with the stations of the Eastern Test Range and the Integrated Mission Control Center in Houston, Tex., are verified at this time. Following an evaluation of the flight readiness test, all systems are reconfigured for launch, and all plugs reverified. A countdown-demonstration test is then performed as the final test prior to launch. The countdown-demonstration test con­sists of an actual launch countdown, complete with propellant loading, astronaut embarkation, etc., with the exception of actual ignition. This test exercises all systems, the launch crew, and the astronauts, and prepares the “team” for the actual operation to follow. This "dress rehearsal" is used to divulge any last minute problems and affords the mission a better chance of success.

Upon completion of the countdown demonstration test, the space vehicle is recycled to pre-count sta­tus, and preparations are made for the final count­down phase of launch operations. Normal recycle time between completion of the countdown dem­onstration test and beginning of launch countdown is 48 to 72 hours.

Propellant loading of the Apollo spacecraft is performed prior to launch day. Aerozine 50 is the fuel and nitrogen tetroxide, the oxidizer. Also prior to launch day, hypergolics for the third stage reaction control system are loaded and ordnance connected. Loading of the cryogenic propellants for the launch vehicle begins on launch day at approxi­mately T-7 hours. (The kerosene is loaded one day before launch.)

Liquid oxygen loading is begun first. The tanks are precooled before filling. Precool of one tank can be accomplished concurrently with the fill of another. Loading is started with the second stage to 40 per cent, followed by the third stage to 100 per cent. The second stage is then brought to a full 100 per cent followed by loading the first stage to 100 per cent. This procedure allows time for the liquid oxygen leak checks to be performed prior to full loading of the second stage. Liquid oxygen is pumped at a flowrate of 1,000 gallons per minute for the third stage. For the second stage, the tank rate is 5,000 gallons per minute, and the first stage tank flowrate is 10,000 gallons per minute.

Liquid hydrogen loading is initiated next, begin­ning with the second stage to 100 per cent. Load­ing of the third stage liquid hydrogen is last. Liquid hydrogen is pumped to the second stage at a rate of 10,000 gallons per minute, and to the third stage at a rate of 4,000 gallons per minute. Topping of cryogenic tanks of the launch vehicle continues until launch. Total cryogenic loading time from start, to finish is 4 hours and MO minutes.

At approximately T-90 minutes, after propellants are loaded, the astronauts enter the spacecraft from the mobile launcher over the swing arm walkway.

LAUNCH

During the remainder of the countdown, the final systems checks are conducted.

Launch vehicle propellant tanks are then pressur­ized. and the first stage engines ignited. During the thrust buildup of the F-l engines, the operation of each of these engines will be automatically check­ed. Upon confirmation of thrust OK condition, the launch commit signal is given to the holddown arms and liftoff occurs.

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Assembled Vehicle—The Saturn V facilities vehicle, the 500F, arrives at Launch Complex 39A.

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