Category Apollo Saturn V News Reference

Before a booster leaves Michoud for test firing, its electrical and mechanical systems are tested ex­tensively by Boeing technicians and engineers. The Stage Test Building with four giant test cells provides the facility. Inside the building are four control rooms, four computer rooms, and two telem­etry rooms. These rooms house equipment that demonstrates the acceptability of the integrated systems of the booster. This includes telemeter calibration, continuity checks, and discrete-function monitoring. RF (radio frequency) also is evaluated.

Moving—A’completed first stage is readied for post-manufactur­ing checkout.

Mechanical, hydraulic, and pneumatic systems tests are conducted to leak-check and functionally check the propellant systems and the engine complex. Checks then are performed to demonstrate the pro­per operation of the electrical and instrumentation systems. All systems are operated and checked in­dividually and then checked as an integrated system in the automatic all-systems checkout.

Monitoring—Technicians check booster performance during a simulated flight from a stage test control room.

After the operation of the test and checkout equip­ment is verified, all electrical, pneumatic, and hy­draulic connections are made to the stage, resis­tance checks are run, and the stage undergoes physical examination.

The environmental control system is connected and checked for proper operation, and the stage’s elec­trical circuits are physically checked for resistance. Stage electrical power із applied in sequential steps and the distribution monitored. The stage instru­mentation transmission system is checked out on both coaxial hardwire and RF links. The electrical systems checkout includes checks of the power distribution circuits, heater power subsystems, destruct system, sequencing subsystem, separa­tion subsystem, and emergency detection system.

The range safety systems undergo a complete end – to-end checkout including transmittal of RF com­mands to the range safety command receiver and monitoring the arm, cutoff, and destruct signals generated by the system.

Instrumentation system testing during stage check­out includes: identification of data channels, gain adjustment of signal conditioners, and checks of measurement systems, telemetry systems, and op­erational RF systems.

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First Stage in Test Cell

Pressure and leak checks are conducted on fuel and LOX tanks and associated lines, engines, fuel and LOX delivery systems, fuel and LOX pressuriza­tion systems, and the control pressure system. Checks are made of the calibration pressure switch simulation, fill and drain operation, and prevalve operation on both fuel and LOX systems.

Propulsion system checks include checks of firing command preparation and execution, engine shut­down prior to “launch commit,” malfunction cutoff, and normal propulsion sequences.

Most of the above-mentioned tests are run for a second time prior to static testing and again during post-static checkout.

The thrust chamber body provides a combustion chamber for burning propellants under pressure and an expansion nozzle for expelling gases produced by the burned propellants at the high velocity re­quired to produce the desired thrust. The thrust chamber is tubular-walled and regeneratively fuel – cooled, and the nozzle is bell-shaped. There are four sets of outrigger struts attached to the exterior of the thrust chamber; two sets of the struts are turbo­pump mounts and the other two are attach points for the vehicle contractor’s gimbal actuators. The thrust chamber incorporates a turbine exhaust manifold at the nozzle exit and a fuel inlet manifold at the injector end which directs fuel to the fuel down tubes. Brackets and’studs welded to the re­inforcing “hatbands” surrounding the thrust cham­ber provide attach points for thermal insulation blankets.

Fuel enters the fuel inlet manifold through two diametrically opposed inlets. From the manifold, 70 per cent of the fuel is diverted through 89 alter­nate CRES “down” tubes the length of the chamber. A manifold at the nozzle exit returns the fuel to the injector through the remaining 89 return tubes. The fuel flowing through the chamber tubes pro­vides regenerative cooling of the chamber walls during engine operation. The thrust chamber tubes are bifurcated; that is, they are comprised of a primary tube from the fuel manifold to the 3:1 ex­pansion ratio area. At that point, two secondary tubes are spliced into each primary tube. This is necessary to maintain a desired cross-sectional area in each of the tubes through the large-diameter belled nozzle section.

The turbine exhaust manifold, which is fabricated from preformed sheet metal shells and which forms a torus around the aft end of the thrust chamber body, receives turbine exhaust gases from the heat
exchanger. Upon entering the manifold, the gases are distributed uniformly. As the gases are expelled from the manifold, flow vanes in the exit slots pro­vide uniform static pressure distribution in the nozzle extension. Radial expansion joints compen­sate for thermal growth of the manifold.

Thrust Chamber Nozzle Extension

The thrust chamber nozzle extension increases the expansion ratio of the thrust chamber from 10:1 to 16:1. It is a detachable unit that is bolted to the exit end ring of the thrust chamber. The interior of the nozzle extension is protected from the engine ex­haust gas environment 15800 Fahrenheit) by film cooling, using the turbine exhaust gases (1200 Fahrenheit) as the coolant. The gases enter the extension between a continuous outer wall and a shingled inner wall, pass out through injection slots between the shingles, and flow over the sur­faces of the shingles forming a boundary layer lie – tween the inner wall of the nozzle extension and the hotter exhaust gases exiting from the main en­gine combustion chamber. The nozzle extension is made of high strength stainless steel.

Hypergol Cartridge

The hypergol cartridge supplies the fluid to produce initial combustion in the thrust chamber. The car­tridge, which is cylindrical and has a burst diaphragm welded to either end, contains a hypergolic fluid consisting of 85 per cent triethylborane and 15 per cent triethylaluminum. As long as the fluid is in the hermetically sealed cartridge, it is stable, but it will ignite spontaneously upon contact with oxygen in any form. During the start phase of operation, increasing fuel pressure in the igniter fuel system ruptures the hurst diaphragms. The hypergolic fluid and the fuel enter the thrust cham­ber through a segregated igniter fuel system in the injector and contact the oxidizer. Spontaneous com­bustion occurs and thrust chamber ignition is estab­lished.

Pyrotechnic Igniter

Pyrotechnic igniters, actuated by an electric spark, provide the ignition source for the propellants in the gas generator and re-ignite the fuel-rich tur­bine exhaust gases as they exit from the nozzle ex­tension.

Thermal Insulation

The thermal insulation protects the F-l engine

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from the extreme temperature environment (2550 Fahrenheit maximum) created by radiation from the exhaust plume and backflow during clustered- engine flight operation. Two types of thermal in­sulators are used on the engine—foil-batt on com­plex surfaces and asbestos blankets on large, simple surfaces. They are made of lightweight material and are equipped with various mounting provi­sions, such as grommeted holes, clamps, threaded studs, and safetywire lacing studs.

The liquid hydrogen cylinder walls comprise the main bulk of the second stage. Five of the cylinder walls measure slightly over 8 feet in height each, while the sixth, the No. 1 cylinder, is 27 inches high. Each of the six cylindrical sections is comprised of four curved, machined aluminum skins. Numerically machine-milled into the inside of the curved skins are stringers and ring frames. Riveted to the cir­cumferential ring frames are flanged aluminum frames which extend inward for approximately 7 inches. In addition to structural rigidity, the frames act as slosh baffles for the liquid hydrogen.

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assumes its shape in the vertical assembly building of NAA’s Seal Beach facility. Assembly in the verti­cal position is based on a building-block concept. In this position, subassembly loading, circumferential

Electrical equipment in the third stage forward skirt area is thermally conditioned by a heat trans­fer system, using “cold plates” on which electronic components are mounted, and through which cool­ant fluid circulates. Coolant is pumped through the system from the IU and returned. Heat from elec­trical equipment attached to the cold plates is dis­sipated by conduction through the mounting feet and the cold plates to the fluid. Refer to the Instru­

The forward skirt area is purged with gaseous ni­trogen to minimize fire and explosion hazards while propellants are loaded or stored in the stage. Gas­eous nitrogen is supplied and remotely controlled from a ground source.

Ordnance Systems

The ordnance systems perform stage separation, retrorocket ignition, ullage control rocket ignition and jettison, and range safety functions.

The Saturn V is equipped with a myriad of equip­ment designed to detect malfunctions. Some of this equipment checks engine thrust, and monitors guidance computer status, attitude rates, angle of attack, and abort request.

This emergency detection information is flashed to the IU where it is routed to the emergency detec­tion distributor (EDS) in the electrical system. The EDS distributor is an interconnector and switching point and has the logic circuits which determine the emergency. In case of a malfunction, the equipment will turn on a light in front of the astronauts. If the spacecraft abort selector switch is in the auto­matic abort position, the abort will take place without further crew participation: the action cannot be vetoed by the astronauts. However, if the selector switch is in the manual position, the crew, consulting with NASA flight controllers, decides when to abort a mission.

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NASA ORGANIZATION

The Saturn V development program comes under the direction of the NASA Office of Manned Space Flight, Washington, D. C. That office assigned devel­opment responsibility to the Marshall Space Flight Center, one of the three Manned Space Flight field centers. Another of those field centers, the Kennedy Space Center, has been delegated the responsibility of launching the Saturn V. (Development of the Apollo spacecraft, the first “payload" for the Sat­urn V, was assigned to the Manned Spacecraft Cen­ter, the other MSF field organization.)

Marshall Center Project Management Organization

More than 125,000 prime and subcontractor em­ployees and 7,500 civil service employees are work­ing on the Saturn program. Saturn industrial activ­ities are scattered nationwide but there are three major areas of concentration:

1. the Northeast, with its grouping of electronic industries.

2. the Southeast, for production, test, and launch operations.

3. the West Coast, with its concentration of aero­space industries for design, production, and test w’ork.

In addition, various research’ projects by scientific – institutions and subcontractor production efforts contributing to the Saturn program are spread throughout the nation.

The wide dispersion makes necessary very com­prehensive and reliable management systems and control techniques to manage the program effec­tively. The geographic dispersion of the Saturn effort requires excellent communications. The Marshall Center must be aware of related programs carried out by other NASA centers—especially the Manned Spacecraft Center, managing the Apollc spacecraft program, and Kennedy Space Center, responsible for Saturn/Apollo launches.

The Marshall Center has found that one of the more effective tools for total program visibility is espe­cially constructed and outfitted rooms called Pro­gram Control Centers. The Saturn V launch vehicle program office and other major groups have such centers.

The budget for the current fiscal year at the MSFC is about SI.7 billion. The center must have a well staffed organization responsive to the many changes which can take place in a program of this magnitude.

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Program management is vested in the program manager. Technical project management, so far as NASA is concerned, occurs at the stage or project level. The program and stage managers are fully responsible for technical adequacy, reliable per­formance, and for management of all related con­tractor activity.

These program and project managers must be backed up and supported by technical competence in depth. This in-depth support is provided, to a de­gree, by a staff of competent technical and business management people in the program manager and stage manager office, and to a much larger degree, by Research and Development Operations,

There is a resident manager at each of the contractor plants to act as the “official" voice for the Marshall Center, All MSFC instructions to the contractor
are transmitted through the resident manager. Through the resident manager, MSFC maintains a direct contact with contractor operations and is kept informed of the status of all significant pro­gram events.

Marshall Center laboratory technical personnel are assigned to the resident managers’ staffs. These technical people are assigned to each resident manager’s office to provide him with assistance in resolving technical problems, and to keep the MSFC technical laboratories directly informed of field technical effort. Laboratory participation is dictated by need as determined by project management.

Many people are involved in attaining the final goal. Project management, technical, and contractor per­sonnel are tied in a close knit group capable of man­aging this country’s large launch vehicle program.

The first stage fuel system supplies RP-1 fuel to the F-l engines. The system consists of a fuel tank, fuel feed lines, pressurization system, fill and drain components, fuel conditioning system, and asso­ciated hardware to meet the propulsion system requirements.

FUEL TANK

The fuel tank, previously described, holds 203,000 gallons of kerosene and is capable of providing 1,350 gallons of fuel per second to the engines through 10 fuel-suction lines.

FUEL FILL AND DRAIN SYSTEM

The fuel tank is filled through a 6-inch duct at the bottom of the tank. Fill rate is 200 gallons per min­ute until the tank is 10 per cent full. After reaching the 10 per cent mark, filling is increased to 2,000 gallons per minute until the tank is full. Normal nonemergency drain takes place through the same duct. A ball-type valve in the fill and drain line provides fuel shutoff.

Fuel Fill and Drain

The fuel fill and drain system consists of a fill and drain line, a fill and drain valve, a fuel loading level probe, and nine temperature sensors. During fuel fill, the temperature sensors provide continuous fuel temperature information used to compute fuel density. When the fuel level in the fuel tank rises to about 102 per cent of flight requirements, the fuel loading probe indicates an overload.

After adjusting fuel to meet requirements, the fill and drain valve is closed.

The fuel tank can be drained under pressure by closing the fuel tank vent and relief valve, supply­ing a pressurizing gas to the tank through the fuel tank prepressurization system, and opening the fuel fill and drain valve.

The turbopump is a direct-drive unit consisting of an oxidizer pump, a fuel pump, and a turbine mount­ed on a common shaft. The turbopump delivers fuel and oxidizer to the gas generator and the thrust chamber. LOX enters the turbopump axially through a single inlet in line with the shaft and is discharged tangentially through dual outlets. Fuel enters the turbopump radially through dual inlets and is dis­charged tangentially through dual outlets. The dual inlet and outlet design provides a balance of radial loads in the pump.

Three bearing sets support the shaft. Matched tandem ball bearings, designated No. 1 and No. 2, provide shaft support between the oxidizer and fuel pumps. A roller bearing, No. 3. provides shaft support between the turbine wheel and the fuel pump. The bearings are cooled with fuel during pump operation. A heater block provides the outer support for No. 1 and No. 2 bearings, and is used during LOX chilldown of the oxidizer pump to pre­vent freezing of the bearings.

A gear ring installed on the shaft is used in con­junction with the torque gear housing for rotating

the pump shaft by hand, and also is used in con­junction with a magnetic transducer for monitoring shaft speed.

There are nine carbon seals in the turbopump: primary oxidizer seal, oxidizer intermediate seal, lube seal No. 1 bearing, lube seal No. 2 bearing, primary fuel seal, fuel inlet seal, fuel inlet oil seal, hot-gas secondary, and hot-gas primary seal.

The main shaft and the parts attaching directly to it are dynamically balanced prior to final assembly on the turbopump.

exactness, and station locating is benefited by the even gravitational force exerted during each as­sembly operation. Constant checks and verification

of station planes and stage alignment are main­tained during each joining procedure by special scopes, levels, and traditional plumb bobs.

Another reason for vertical assembly involves the welding of cylinders and bulkhead. If the stage were welded while in a horizontal position, temperatnre diversion over the circumference could result in harmful expansion near the top of the stage.

To facilitate movement of the huge components and of the stage itself, a motorized transfer table rolls from outside to inside the building. Essentially, the assembly sequence begins with the welding of the lower two cylinders. Then the common bulkhead is welded to that assembly. Next the uppermost cyl­inder is welded to the LHa forward bulkhead. The aft LOX bulkhead and the aft facing sheet of the common bulkhead are welded together to form the liquid oxygen tank, and the thrust structure and aft skirt are then assembled to it. The remaining cylinders are then welded to the stage, and the for­ward skirt is then mated to the stage stack. The interstage is fit-checked to the thrust structure before interstage systems are installed. Throughout the assembly and welding operations, hydrostatic, X-ray, dye penetrant, and other tests and quality control devices are performed to ensure that speci­fications are met. The liquid hydrogen portion of the second stage as well as the liquid oxygen tank are given a thorough cleaning after assembly. After each bulkhead is welded to its components, it is hy­drostatically tested. After completion of stack weld operations, the entire stage is pneumostatically tested. After completion of these tests, the liquid

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Reposiiisring–Second stage is turned horizontally for checkout operation.

Stage Complete…. Flight stage moves on transfer table from

assembly building to checkout building.

Engine Installation—J-2 engines are mounted in stage.

After assembly, the stage Is moved to a vertical checkout building for final checks on all stage sys­tems.

The stage separation system consists of a sever­able tension strap, mild detonating fuse (MDF), exploding bridgewire, (EBW), detonators and EBW firing units.

Separation System

The severable tension strap houses two redundant MDF cords in a “V” groove circumventing the stage between the aft skirt and aft interstage at the sepa­ration plane. Ignition of the MDF cords is triggered by a signal from the second stage sequencer through the EBW and EBW firing units about 3 seconds after second stage engine cutoff.

The MDF consists of a flexible metal sheath sur­rounding a continuous core of high explosive mate­rial. Once detonated, the explosive force of the MDF occurs at a rate of 23,000 feet per second.

The EBW detonator is fired to initiate the MDF explosive train. A 2,300 VDC pulse is applied to a small resistance wire and a spark gap. The high voltage electrical arc across the spark gap ignites a charge of high explosive material which in turn detonates the MDF. The high voltage pulse require­ment for ignition renders this system safe from random ground or vehicle electrical power. Upon command, each EBW firing unit supplies high volt­age and current required to fire a specific EBW detonator.