Category Apollo Saturn V News Reference

This volume has been prepared by the five Saturn V major contractors: The Boeing Company; Douglas Aircraft Company: Space Division of North Amer­ican Aviation, Inc.; Rocketdyne Division of North American Aviation, Inc.; and International Business Machines Corporation in cooperation with the Na­tional Aeronautics and Space Administration.

It is designed to serve as an aid to newsmen in pres­ent and future coverage of the Saturn V in its role in the Apollo program and as a general purpose large launch vehicle. Every effort has been made to present a comprehensive overall view of the vehicle and its capabilities, supported by detailed

The Boeing Company P. 0. Box 29100 New Orleans, La. 70129 Attention: William W. Clarke

Douglas Aircraft Company Missile & Space Systems Division Space Systems Center 5301 Bolsa Avenue Huntington Beach, Calif. 92647 Attention: Larry Vitskv

International Business Machines Corporation

Federal Systems Division

150 Sparkman Drive

Huntsville, Ala. 35807

Attention: James F. Harroun

information on the individual stages and all major systems and subsystems.

Weights and measurements cited throughout the book apply to the AS-501 vehicle, the first flight version of the Apollo/Saturn V.

All photographs and illustrations in the book are available for general publication. The first letter in each photo number is a code identifying the or­ganization holding that negative: В for Boeing; R for Rocketdyne Division of North American; D for Douglas; IBM for IBM; S for Space Division of North American; H for NASA, Huntsville, Ala.; and К for NASA, Kennedy Space Center, Fla.

s are:

Rocketdyne Division

North American Aviation, Inc.

6633 Canoga Avenue Canoga Park, Calif. 91304 Attention: R. K. Moore

National Aeronautics and Space Administration George C. Marshall Space Flight Center Public Affairs Office Huntsville, Ala. 35812 Attention: Joe Jones

National Aeronautics and Space Administration Public Affairs Office Kennedy Space Center, Fla. 32931 Attention: Jack King

Space Division

North American Aviation, Inc. Seal Beach, Calif. 90241 Attention: Richard E. Barton

^ SATURN V NEWS REFERENCE

SATURN V FACT SHEET

‘SINCE INDIVIDUAL STAGE DIMENSIONS OVERLAP IN SOME CASES, OVERALL VEHICLE

LENGTH IS NOT THE SUM OF INDIVIDUAL STAGE LENGTHS

“INCLUDES AFT INTERSTAGE WEIGHT

PROPULSION SYSTEMS

FIRST STAGE —Five bipropellant F-l engines developing 7,500,000 lb. thrust

RP-1 Fuel-203,000 gal. (1,359,000 lb.), LOX-331,000 gal. (3,133,000 lb.)

SECOND STAGE Five bipropellant J-2 engines developing more than 1,000,000 lb. thrust LH2—260,000 gal. (153,000 lb.), LOX-83,000 gal. (789,000 lb.)

THIRD STAGE —One bipropellant J-2 engine developing up to 225,000 lb. thrust LH2—63,000 gal. (37,000 lb.), LOX—20,000 gal. (191,000 lb.)

CAPABILITY

FIRST STAGE —Operates about 2.5 minutes to reach an altitude of about 200,000 feet (38 miles) at burnout

SECOND STAGE Operates about 6 minutes from an altitude of about 200,000 feet to an altitude of 606,000 feet (114.5 miles)

THIRD STAGE —Operates about 2.75 minutes to an altitude of about 608,000

feet (115 miles) before second firing and 5.2 minutes to translunar injection

PAYLOAD —250,000 lb. into a 115 statute-mile orbit

I SATURN V NEWS REFERENCE

An unusual but convenient type of fluid power or hydraulic system is in use on the Saturn V first stage. It incorporates the same types of fuels —RP-1 and RJ-1 (kerosene)—that are used in the stage fuel system. Ordinarily a different and weaker type of fluid is used for hydraulics. This system elimi­nates the use of a separate pumping system.

The fluid power system provides ground and flight fluid power for valve actuation and thrust vector­ing. It gives power primarily to the engine start system and the engine gimbaling system. Its source is the fuel system. RJ-1 is provided from the ground before liftoff, and RP-1 is supplied from the fuel tank during flight.

The ground supply of RJ-1 is routed to all five en­gines at 1,500 psig and eventually back to the ground supply. After ignition, RP-1 is routed from the high pressure fuel duct to the servoactuators for hy­draulic power to position the engines.

The center engine, which has no thrust vectoring system, directs its hydraulic fluid through the feed line and 4-way hydraulic control valve to supply pressure to the closing ports of the gas generator, main fuel valves, and main LOX valves. The fuel passes through orifices and then is ducted through the ground checkout valve and back to ground supply through the return line.

The four outboard engines direct RJ-1 through the servoactuators to the ground checkout valve where it is returned through a coupling to ground supply.

Three pressure switches, mounted on a single mani­fold located on the thrust chamber fuel manifold, sense fuel injection pressure. These thrust-OK pres­sure switches are used in the vehicle to indicate that all five engines are operating satisfactorily. И pressure in the fuel injection cavity decreases, the switches deactuate, breaking the contact and interrupting the thrust-OK output signal.

PRESSURIZATION SYSTEM

The pressurization system heats GOX and helium for vehicle tank pressurization. The pressurization system consists of a heat exchanger, a heat ex­changer check valve, a LOX flowmeter, and various heat exchanger lines. The LOX source for the heat exchanger is tapped from the thrust chamber oxi­dizer dome, and the helium is supplied from the vehicle. LOX flows from the thrust chamber oxi­dizer dome through the heat exchanger check valve, LOX flowmeter, and the LOX line to the heat ex­changer.

Heat Exchanger

The heat exchanger heats GOX and helium with hot turbine exhaust gases, which pass through the heat exchanger over the coils. The heat exchanger consists of four oxidizer coils and two helium coils installed within the turbine exhaust duct. The heat exchanger is installed between the turbopump manifold outlet and the thrust chamber exhaust manifold inlet. The shell of the heat, exchanger contains a bellows assembly to compensate for thermal expansion during engine operation.

Heat Exchanger Check Valve

The heat exchanger check valve prevents GOX or

vehicle prepressurizing gases from flowing into the oxidizer dome. It consists of a line assembly

and a swing check valve assembly. It is installed between the thrust chamber oxidizer dome and the heat exchanger LOX inlet line.

LOX Flowmeter

The LOX flowmeter is a turbine-type, volumetric, liquid-flow transducer incorporating two pickup coils. Rotation of the LOX flowmeter turbine gen­erates an alternating voltage at the output ter­minals of the pickup coils.

Heat Exchanger Lines

LOX and helium are routed to and from the heat exchanger through flexible lines. The GOX and helium lines terminate at the vehicle connect inter­face. The LOX line connects the heat exchanger to the heat exchanger check valve.

Ground support operations play an important part in getting the second stage ready for operation. Among the vital operations in this area are check­out (performed mostly with complex electronic equipment and computerized routines which stimu­late stage systems and analyze responses), leak detection and insulation purge, and engine com­partment conditioning.

Leak Detection and Insulation Purge

The purpose of this system is to detect hydrogen, oxygen, or air leaks; to dilute and remove leaking gases; and to prevent air from liquifying during tanking operations.

Any operation involving liquid hydrogen can be. extremely hazardous; liquid hydrogen in the pres­ence of oxygen can explode or create a fire. The low-temperature atmosphere of liquid hydrogen causes air to liquify and solidify against the hy­drogen tank wall if there is any leak in the tank insulation. The organic portion of the insulation will become impact-sensitive when drenched in liquid air or oxygen; insulation saturated with cryopumped air will add weight to the stage and
could cause damage during draining because of a pressure buildup created by the liquified air re­turning to a gas. For these reasons, detection, con­trol, and elimination of any hydrogen leaks from the stage and ground equipment are of great importance. The leak detection system checks out the liquid hydrogen tank, tank insulation, and the common bulkhead. The areas to be checked are divided (tank wall, forward bulkhead, and common bulk­head), each with inlet and outlet taps. A gas ana­lyzer determines the concentration of hydrogen in the purge gas (helium) after it has been forced through the insulation, and thus indicates any leak­age.

From the start of hydrogen loading until launch, the insulation and core of the common bulkhead are continuously purged of hazardous gases. Vacuum equipment is used for evacuation to pre­vent pressure buildup in the insulation and bulk­heads by removing trapped gases. The insulation purge prevents air from entering the insulation in the event of damage during cryogenic operations.

Engine Compartment Conditioning

The purpose of this system is to purge the engine and interstage areas of explosive mixtures and to maintain proper temperature in critical regions of the aft compartment of the second stage. The com­partment is purged before tanking and while the propellants are loaded.

The system consists of a 13-inch diameter feed line, manifold, ducts, and a series of vents surrounding the engine compartment and skirt area. The system provides temperature control for the hydraulic systems and certain components on the J-2 engines. The purge gas is forced through orifices in the mani­fold to the following areas requiring warming: the area between the thrust structure and the liquid oxygen tank, the bottom of the thrust structure including the lower surface of the thrust cone, the aft skirt and interstage, and the top surface of the heat shield.

The vent holes are located under the supporting hat sections on the outside of the aft skirt; this prevents wind, rain, and dust from entering the engine com­partment. The vents are located so that the flow pattern provides good thermal control and expels hazardous gases.

The aft skirt and interstage are purged with warm (80 to 250 degrees) nitrogen. The nitrogen is sent through the feed line into the manifold, and then through ducts to the temperature-sensitive areas. By maintaining a 98 per cent nitrogen atmosphere in the engine compartment, desired temperatures are maintained and the danger of fire or explosion resulting from propellant leaks are minimized.

j SATURN V NEWS REFERENCE

THIRD STAGE FACT SHEET

WEIGHT: 34,000 lb. (dry) including 7,700-lb. aft interstage

262,0 lb. (loaded)

DIAMETER: 21 ft. 8 in.

HEIGHT: 58 ft. 7 in.

BURN TIME: 1st burn—2.75 min. (approx.)

2nd burn….. 5.2 min. (approx.)

VELOCITY: 1st burn—17,500 miles per hour at burnout (approx.)

2nd burn—24,500 miles per hour (approx, typical lunar mission escape velocity)

ALTITUDE AT BURNOUT: 115 miles after 1st burn and into a translunar injection on 2nd burn MAJOR STRUCTURAL COMPONENTS

AFT INTERSTAGE THRUST STRUCTURE COMMON BULKHEAD

AFT SKIRT PROPELLANT TANK FORWARD SKIRT

MAJOR SYSTEMS

PROPULSION: One bipropellant J-2 engine Total Thrust: 225,000 lb. (maximum)

Propellants: LH2—63,000 gal. (37,000 lb.)

LOX—20,000 gal. (191,000 lb.)

HYDRAULIC: Power for gimbaling J-2 engine

ELECTRICAL: One 56 VDC and three 28 VDC batteries, providing basic power for all electrical functions

TELEMETRY AND INSTRUMENTATION: Five modulation subsystems, providing transmission of flight data to ground stations ENVIRONMENTAL CONTROL: Provides temperature-controlled environment for components in aft skirt, aft interstage, and forward skirt ORDNANCE: Provides explosive power for stage separation, retrorocket ignition, ullage rocket ignition and jettison, and range safety requirements FLIGHT CONTROL: Provides stage attitude control and propellant ullage control

The main fuel valve is a butterfly-type valve, spring – loaded to the closed position, pneumatically oper­ated to the open position, and pneumatically assisted to the closed position. It is mounted between the fuel high-pressure duct from the fuel turbopump and the fuel inlet manifold of the thrust chamber assembly. The main fuel valve controls the flow of fuel to the thrust chamber. Pressure from the igni­tion stage control valve on the pneumatic control package opens the valve during engine start. As the gate starts to open, it allows fuel to flow to the fuel inlet manifold.

MAIN OXIDIZER VALVE

The main oxidizer valve (MOV) is a butterfly-type valve, spring-loaded to the closed position, pneu­matically operated to the open position, and pneu­matically assisted to the closed position. It is mounted between the oxidizer high-pressure duct from the oxidizer turbopump and the oxidizer inlet on the thrust chamber assembly.

Pneumatic pressure from the normally closed port of the mainstage control solenoid valve is routed to both the first and second stage opening actuators of the main oxidizer valve. Application of opening pressure in this manner, together with controlled venting of the main oxidizer valve closing pressure through a thermal-compensating orifice, provides a controlled ramp opening of the main oxidizer valve through all temperature ranges. A sequence valve, located within the MOV assembly, supplies pneu­matic pressure to the opening control part of the gas generator control valve and through an orifice to the closing part of the oxidizer turbine bypass valve.

PROPELLANT UTILIZATION VALVE

The propellant utilization (PU) valve is an electri­cally operated, two-phase, motor-driven, oxidizer

6-4

SATURN V NEWS REFERENCE

pressure-actuated to the closed position. Both pro­pellant bleed valves are mounted to the bootstrap lines adjacent to their respective turbopump dis­charge flanges.

The valves allow propellant to circulate in the pro­pellant feed system lines to achieve proper operat­ing temperature prior to engine start. The bleed valves are engine controlled. At engine start, a he­lium control solenoid valve in the pneumatic con­trol package is energized allowing pneumatic pres­sure to close the bleed valves, which remain closed during engine operation.

Saturn V vehicles are assembled, checked out, and launched at Launch Complex 39 at Kennedy Space Center. Complex 39 embodies a new mobile concept of launch operations which includes superior re­liability and time savings offered by assembly and checkout in a protected environment and reduc­tion of actual pad time as much as 80 per cent with

a consequent increase in launch rate capability. The ability to adapt economically to future program requirements is another advantage. For example, the service platforms used in the Saturn/Apollo program could be used for other vehicles of similar configuration, and the area can accommodate space boosters with thrusts up to 40 million pounds.

Facilities

The major components of Launch Complex 39 in­clude: (1) the Vehicle Assembly Building, where the space vehicle is assembled and prepared; (2) the mobile launcher, upon which the vehicle is erected for checkout, and from which, later, it is launched; (3) the crawlerway, upon which the fully assembled vehicle is carried by transporter to the launch site; (4) the mobile service structure, which provides external access to the vehicle at the launch site; (5) the transporter which carries the launch vehicle, mobile launcher, and mobile service struc­ture to various positions at the launch complex; and (6) the launch area from which the space vehicle is launched.

THE VEHICLE ASSEMBLY BUILDING

The Vehicle Assembly Building (VAB) consists of a high bay area 525 feet tall, a low bay area 210 feet tall, and a four-story launch control center iLCC) connected to the high bay by an enclosed bridge. The VAB, with 130 million cubic feet, is the world’s largest building in volume. It covers eight acres of land. There are four assembly and checkout bays in the high bay area. The low bay area contains eight stage preparation and checkout cells equipped with systems to simulate stage inter­face. The launch control center houses display, monitoring, and control equipment for checkout and launch operations. There are four firing rooms in the LCC, one for each high bay and checkout area. Work platforms, mounted on opposite walls in the high bay area, are designed to enclose various work areas around the launch vehicle. Platforms extend or retract in less than 10 minutes. Twenty – ton hydraulic jacks are used to align platforms.

The Saturn V, after prelaunch checkout on its mo­bile launcher, is carried by the transporter from the VAB through a door shaped like an inverted “T”. The door is 456 feet high. The base of the door is 149 feet wide and 113 feet high; the remainder is 76 feet wide. There are four such doors in the VAB, one for each of its four high bays. In keeping with the protective environment of the building, doors were designed to withstand winds of 125 miles per hour.

There are 141 lifting devices in the VAB, ranging from one-ton mechanical hoists to two 250-ton high-

Checkout Vehicle—The Saturn V facilities vehicle begins its journey from the Vehicle Assembly Building to the launch pad. Its purpose was to check out facilities, train launch crews, and verify procedures at KSC.

INTRODUCTION

When the United States made the decision in 1961 to undertake a manned lunar landing effort as the focal point of a broad new space exploration pro­gram, there was no rocket in the country even approaching the needed capability. There was a sort of “test bed” in the making, a multi-engine vehicle now known as Saturn I. It had never flown. And it was much too small to offer any real hope of sending a trio to the moon, except possibly through as many as a half dozen separate launchings from earth and the perfection of rendezvous and docking techniques, which had never been tried.

That was the situation that brought about the an­nouncement on Jan. 10, 1962, that the National Aeronautics and Space Administration would de­velop a new rocket, much larger than any previously attempted. It would be based on the F-l rocket en­gine, the development of which had been underway since 1958, and the hydrogen-fueled J-2 engine, upon which work had begun in 1960.

The Saturn V, then, is the first large vehicle in the U. S. space program to be conceived and de­veloped for a specific purpose. The lunar landing task dictated the make-up of the vehicle, but it was not developed solely for that mission. As President Kennedy pointed out when he issued his space chal­lenge to the Congress on May 25, 1961, the overall objective is for “this Nation to take a clearly lead­ing role in space achievement which in many ways may hold the key to our future on earth.” He said of the lunar landing project: “No single space pro­ject in this period will be more exciting, or more impressive to mankind, or more important for the long-range exploration of space: and none will be so difficult or expensive to accomplish…”

The Saturn V program is the biggest rocket effort undertaken in this country. Its total cost, including the production of 15 vehicles between now and early 1970, will be above $7 billion.

NASA formally assigned the task of developing the Saturn V to the Marshall Space Flight Center on Jan. 25, 1962. Launch responsibility was committed to the Kennedy Space Center. (The Manned Space­craft Center, the third center in manned space flight, is responsible for spacecraft development, crew training, and inflight control.)

DESCRIPTION

Marshall Center rocket designers conceived the Saturn V in 1961 and early 1962. They decided that
a three-stage vehicle would best serve the immedi­ate needs for a lunar landing mission and would serve well as a general purpose space exploration vehicle.

One of the more important decisions made early in the program called for the fullest possible use of components and techniques proven in the Saturn I program. As a result, the Saturn V third stage (S-IVB) was patterned after the Saturn I second stage (S-IV). And the Saturn V instrument unit is an outgrowth of the one used on Saturn I. In these areas, maximum use of designs and facilities already avail­able was incorporated to save time and costs.

Many other components were necessary, including altogether new first and second stages (S-IC and S-II). The F-l and J-2 engines were already under development, although much work remained to be done. The guidance system was to be an improve­ment on that of the Saturn I.

Saturn V, including the Apollo spacecraft, is 364 feet tall. Fully loaded, the vehicle will weigh some

6.1 million pounds.

The 300,000-pound first stage is 33 feet in diameter and 138 feet long. It is powered by five F-l engines generating 7.5 million pounds thrust. The booster will burn 203,000 gallons of RP-1 (refined kerosene) and 331,000 gallons of liquid oxygen (LOX) in 2.5 minutes.

Saturn V’s second stage is powered by five J-2 engines that generate a total thrust of a million pounds. The 33-foot diameter stage weighs 95,000 pounds empty and more than a million pounds loaded. It burns some 260,000 gallons of liquid hydrogen and

83.0 gallons of liquid oxygen during a typical 6- minute flight.

Third stage of the vehicle is 21 feet and 8 inches in diameter and 58 feet and 7 inches long. An inter­stage adapter connects the larger diameter second stage to the smaller upper stage. Empty weight of the stage is 34,000 pounds and the fueled weight is

262.0 pounds. A single J-2 engine developing up to 225,000 pounds of thrust powers the stage. Typi­cal burn time is 2.75 minutes for the first burn and

5.2 minutes to a translunar injection.

The vehicle instrument unit sits atop the third stage. The unit, which weighs some 4,500 pounds, contains the electronic gear that controls engine ig­nition and cutoff, steering, and all other commands necessary for the Saturn V mission. Diameter of the instrument unit is 21 feet and 8 inches, and height is 3 feet.

Directly above the instrument unit in the Apollo

configuration is the Apollo spacecraft. It consists of the lunar module, the service module, the com­mand module, and the launch escape system. Total height of the package is about 80 feet.

The electrical power and distribution system of the

first stage provides power for controlling and mea­suring functions of the vehicle. The system operates during static firing, launch preparation and check­out. launch, and flight.

The electrical system consists of two batteries, a main power distributor, a sequence and control distributor, propulsion distributor, timer distrib­utor. measuring distributors, thrust OK distributor, and measuring power distributor.

Two independent 28-volt DC power systems are installed on the stage. System No. 1, the main power battery, energizes the stage controls. The battery has a 640-ampere-minute rating, weighs about 22 pounds, and is used to control various solenoids. Battery No. 2, the instrumentation battery, ener­gizes the flight measurement system and gives power to redundant systems for greater mission reliability. It has a 1,250-ampere-minute rating and weighs approximately 55 pounds. The range safety system can be operated by either battery.

Preflight power is supplied from ground equipment through umbilical connections. The supply for each system is 28 volts. Ground sources supply power for heaters, ignitors, and valve operators that are not operated during flight.

The distributors subdivide the electrical circuits and serve as junction boxes. Both electrical sys­tems share the same distributors. The main power distributor houses relays, the power transfer switch, and electrical distribution buses. The relays con­trol circuits that must be time-programmed. The motor-operated, multi-contact, power transfer switch transfers the stage load from the ground supply to the stage batteries. The transfer is tried several times during countdown to verify opera­tion. Power is distributed by the main buses.

The switch selector, actuated by the instrument unit (IU), commands the sequence and control dis­tributor, which in turn amplifies the signals re­ceived. The sequence and control distributor then energizes the various circuit relays required to implement the flight program. The switch selector is an assembly of redundant low power relays and transistor switches, which control the sequence and control distributor. It is activated by a coded signal from the instrument unit computer.

The propulsion distributor contains the monitor and control circuits for the propulsion system.

The thrust OK distributor contains the circuits that shut down the engines when developed thrust is inadequate. Two of the three thrust OK switches must operate or the engine will be shut down.

The timer distributor houses the circuits to delay the operation of relay valves and other electro­mechanical devices. The programmed delays are essential for optimum performance and safety.

The measuring power distributor contains electrical buses, and the measuring distributors route data from measuring racks, serve as measurement sig­nal junction boxes, and switch data between the hardwire and telemetry.

The engine interface panel, mounted above the turbopump LOX and fuel inlets, provides the ve­hicle connect location for electrical connectors be­tween the engine and the vehicle. It also provides the attachment point for the vehicle flexible heat – resistant curtain. The panel is fabricated from heat – resistant stainless-steel casting made in three sec­tions and assembled by rivets and bolts.

ELECTRICAL SYSTEM

The electrical system consists of flexible armored wiring harnesses for actuation of engine controls and the flight instrumentation harnesses.

HYDRAULIC CONTROL SYSTEM

The hydraulic control system operates the engine propellant valves during the start and cutoff se­quences. It consists of a hypergol manifold, a check­out valve, an engine control valve, and the related tubing and fittings.

Hypergol Manifold

The hypergol manifold directs hypergolic fluid to the separate igniter fuel system in the thrust cham­ber injector. It consists of a hypergol container, an ignition monitor valve, a position switch, and an igniter fuel valve. The hypergol container, position switch, and igniter fuel valve are internal parts of the hypergol manifold.

A spring-loaded, cam-lock mechanism incorporated in the hypergol manifold prevents actuation of the
ignition monitor valve until after the upstream hypergol cartridge diaphragm bursts. The same mechanism actuates a position switch that indicates when the hypergol cartridge is installed. The igniter fuel valve is a spring-loaded, cracking check valve that opens and allows fuel to flow into the hypergol container. The hypergol cartridge diaphragms are ruptured by the resultant pressure surge when the igniter fuel valve opens.

Ignition Monitor Valve

The ignition monitor valve is a pressure-actuated, three-way valve mounted on the hypergol mani­fold. It controls the opening of the fuel valves and permits them to open only after satisfactory com­bustion has been achieved in the thrust chamber.

When the hypergol cartridge is installed in the hypergol manifold, a cam-lock mechanism prevents the ignition monitor valve poppet from moving from the closed position. The ignition monitor valve has six ports: a control port, an inlet port, two outlet ports, a return port, and an atmospheric reference port. The control port receives pressure from the thrust chamber fuel manifold. The inlet port re­ceives hydraulic fuel pressure for opening the fuel valves. When the ignition monitor valve poppet is in the deactuated position, hydraulic fuel from the inlet port is stopped at the poppet seat. When the hypergol cartridge diaphragm bursts, the spring- loaded cam-lock retracts to permit the ignition moni­tor valve poppet unrestricted motion. When thrust chamber pressure (directed to the control port from the thrust chamber fuel manifold I increases, the ignition monitor valve poppet moves to the open (actuated) position and hydraulic fuel is directed through the outlet ports to the fuel valves.

STAGE DESCRIPTION

Basically, the Saturn V third stage, the S-IVB, is an aluminum air-frame structure powered by a single, J-2 engine, which burns liquid oxygen and liquid hydrogen. The engine has a maximum thrust of 225.000 pounds. The structure has a bipropellant capacity of 228,000 pounds of fuel and oxidizer.

STAGE FABRICATION AND ASSEMBLY

The third stage structure consists of a forward skirt assembly, propellant tank assembly, thrust structure assembly, aft skirt assembly, and aft in­terstage assembly. The propellant tank assembly consists of a single tank separated by a common bulkhead into a fuel compartment and an oxidizer compartment.

Forward Skirt Assembly

The forward skirt is a cylindrical aluminum skin and stringer structure that provides a hard attach point for the instrument unit. In addition, the for­
ward skirt provides an interior mounting structure for electrical and electronic equipment that requires environmental conditioning, as well as range safety and telemetry antennas mounted around the ex­terior periphery. Environmental conditioning for electronic equipment is provided by cold plates which utilize a coolant supplied from the IU thermo­conditioning system.

Propellant Tank Assembly

Structural elements of the propellant tank assembly are a cylindrical tank section, common bulkhead, aft dome, and forward dome. Seven segments are machined from aluminum alloy plate to form the tank section. A waffle pattern is then machine – milled into each segment to reduce weight and pro­vide shell stiffness. The formed segments are joined into a complete cylinder by single-pass internal weld on a Pandjiris welding machine.

Aft and forward domes are made by forming "orange peel” segments on a stretch press. Orange peel segments are then joined in a dome welder. Each

Third Stage Production Sequence

6-і

Slosh Baffle…. Horizontal rings are installed inside LH2 tank for

propellant stabilization during flight.

Third stage vehicles reach end of assembly sequence with final assembly and checkout in 115-foot vertical towers.

Engine I nstalled—J-2 engine is attached to stage in final assembly tower at Huntington Beach.

Final installation of various subsystem components is performed in a checkout tower, along with the installation and alignment of the,1-2 engine. The stage is in a vertical position in the tower where a complete stage checkout of subsystems and systems is conducted except for actual ignition of engine. After satisfactory checkout, the stage is removed from the tower, placed on a dolly, and ground sup­port rings are installed at each end of the stage. It is then painted, weighed, and prepared for ship­ment to the Douglas Sacramento Test Center for simulated and static firing of APS engines and J-2 engine.