Category Apollo Saturn V News Reference

Pressurizing gases used in the LOX tank are he­lium, gaseous oxygen, and nitrogen. These gases are used in prepressurization, flight pressurization, and storage pressurization.

LOX cannot exceed -297 degrees Fahrenheit or

Prepressurization is necessary 45 seconds prior to

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engine ignition to give sufficient tank ullage pres­sure for engine start and thrust buildup. Helium, used as the pressurizing gas to reduce flight weight, is supplied by ground support through the helium ground connection. It proceeds up the gaseous oxygen line into the LOX tank through the GOX distributor. The flow of helium is monitored by the pressure duct and stopped at 26 pounds per square inch absolute (psia) maximum and is resumed when the pressure drops to 24.2 psia during engine start. Ground-supplied helium is available until liftoff. GOX is added to the LOX tank for pressurization during flight. Each engine contributes to GOX pressurization. A portion of LOX—6,340 pounds — passing through the engine is diverted from the LOX dome into the engine heat exchanger where hot gases exhausted from each engine turbine trans­form LOX into GOX. The GOX flows from each heat exchanger into the GOX line manifold through the flow control valve, up the GOX line, and into the LOX tank through the GOX distributor. The GOX flow is approximately 40 pounds per second to main­tain a LOX tank ullage pressure of 18 to 23 psia.

LOX Pressurization

While the booster is being stored or transferred from one location to another, a slight positive ni­trogen pressure is maintained for cleanliness and low humidity conditions. The external nitrogen pressure source is removed during flight operations.

The propellant feed system transfers LOX and fuel from the propellant tanks into the pumps which dis­charge into the high-pressure ducts leading to the gas generator and the thrust chamber. The system consists of two oxidizer valves, two fuel valves, a bearing coolant control valve, two oxidizer dome purge check valves, a gas generator and pump seal purge check valve, turbopump outlet lines, orifices, and lines connecting the components. High-pressure fuel is supplied from the propellant feed system of the engine to the vehicle-contractor-supplied thrust vector control system.

Propellant Feed…. The main LOX valve and high-pressure line

are shown at left. At right are the main fuel valve and high- pressure line.

Oxidizer Valves

Two identical oxidizer valves, designated No. 1 and No. 2, control LOX flow from the turbopump to the thrust chamber oxidizer dome and sequence the hydraulic fuel to the opening port of the gas genera­tor valve. When the valves are in the open position at rated engine pressures and flowrates, neither will dose if the hydraulic fuel opening pressure is lost. Each of the oxidizer valves is a hydraulically actuated, pressure-balanced, poppet type, and con­tains a mechanically actuated sequence valve. A spring-loaded gate valve permits reverse flow for recirculation of the hydraulic fluid with the pro­pellant valves in the closed position, but prevents fuel from passing through until the oxidizer valve is open 1*3.4 per cent. As the oxidizer valve reaches this position, the piston shaft opens the gate, allow­ing fuel to flow through the sequence valve, which in turn opens the gas generator valve.

LOX Distribution -Oxidizer is distributed by the LOX dome (lower center). Main LOX valves are shown at left and right with the engine interface panel above.

A position indicator provides relay logic in the en­gine electrical control circuit and provides instru­mentation for recording movement of the oxidizer valve poppet.

The two oxidizer dome purge check valves, mounted on each of the oxidizer valves, allow purge gas to enter the oxidizer valves, but prevent oxidizer from

entering the purge system.

Fuel Valves

Two identical fuel valves, designated No. 1 and No. 2, are mounted 180 degrees apart on the thrust chamber fuel inlet manifold and control the flow of fuel from the turbopump to the thrust chamber. When the valves are in the open position at rated engine pressures and flowrates, they will not dose if hydraulic fuel pressure is lost.

Position indicators in the fuel valves provide relay logic in the engine electrical control circuit and instrumentation for recording movement of the

valve poppets.

Flight control of the second stage is maintained by gimbaling the four rocket thrust engines for thrust vector (direction) control. These are the four out­

board engines; the fifth J-2 engine located in the center of the cluster is stationary.

Each outboard engine has a separate engine actua­tion system to provide the force to position the en­gine. Gimbaling is achieved by hydraulic-powered actuators controlled by electrical signals generated through a flight control computer located in the instrument unit just above the third stage. Hydrau­lic power for operating each of the gimbaling actu­ators is supplied by individual engine-driven hy­draulic pumps. Each system is self-contained and operates under a pressure of 3,500 psi. The compon­ents of each hydraulic system are attached to the thrust structure above each of the outboard engines. The main hydraulic pump is driven by the liquid oxygen turbopump on the respective engine. Two servoactuators that control each engine programmed for gimbaling are located on the engine outboard side. One is on the pitch plane, and the other on the yaw plane. Each actuator will gimbal the en­gine plus or minus 7 degrees in pitch or yaw and plus or minus 10 degrees in combination to correct for roll errors at a minimum rate of 8 degrees per second.

During flight, the guidance system continuously determines an optimum vehicle steering command based on the vehicle’s position, velocity, and accel­eration. This system, located in the instrument unit, has a guidance signal processor which de­livers attitude correction signals to the flight con­trol computer in the instrument unit. These signals are shaped, scaled, and summed electronically. These summed error signals are then directed to the servoactuator amplifiers, which, in turn, drive their respective servoactuators in the second stage. These signals cause the servoactuators to position the engines.

MEASUREMENT SYSTEM

A wide variety of transducers and signal condi­tioners is used in the instrumentation system, which feeds signals to a high-level telemetering system for transmission to the ground. The various instru­mentation sensors monitor pressure, temperature, and propellant flow rates within the tanks. Other sensors record the amount of vibration and noise, and flight position and acceleration.

Tied into the measurement system are telemetry and radio frequency subsystems which transmit the performance signals to ground receiving sta­tions for immediate (real-time) and postflight ve­hicle performance evaluation. Antennas which serve the telemetry and radio frequency subsystems are flush-mounted on the forward skirt and are omni­directional in coverage.

Approximately 10 seconds before second stage pro­pellant depletion, a signal activates the separation system which will sever the second stage from the third. An interstage connecting the second and third stage has four retrorockets which are fired to decel­erate the second stage.

The augmented spark igniter (ASI) is mounted to the injector face. It provides the flame to ignite the propellants in the thrust chamber. When engine start is initiated, the spark exciters energize two spark plugs mounted in the side of the igniter cham­ber. Simultaneously, the control system starts the initial flow of oxidizer and fuel to the spark igniter. As the oxidizer and fuel enter the combustion cham­ber of the ASI, they mix and are ignited.

Mounted in the ASI is an ignition monitor which in­dicates that proper ignition has taken place. The ASI operates continuously during entire engine fir­ing, is uncooled, and is capable of multiple reigni­tions under all environmental conditions.

Propellant Feed System

The propellant feed system consists of separate fuel and oxidizer turbopumps, main fuel valve, main oxidizer valve, propellant utilization valve, fuel and oxidizer flowmeters, fuel and oxidizer bleed valves, and interconnecting lines.

FUEL TURBOPUMP

The fuel turbopump, mounted on the thrust cham-

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ber, is a turbine-driven, axial flow pumping unit consisting of an inducer, a seven-stage rotor, and a stator assembly. It is a high-speed pump operating at 27,000 rpm, and is designed to increase hydrogen pressure from 30 psia to 1,225 psia through high – pressure ducting at a flowrate which develops 7,800 brake horsepower.

Power for operating the turbopump is provided by

a high-speed, two-stage turbine. Hot gas from the gas generator is routed to the turbine inlet mani­fold which distributes the gas to the inlet nozzles where it is expanded and directed at a high velocity into the first stage turbine wheel.

After passing through the first stage turbine wheel, the gas is redirected through a ring of stator blades and enters the second stage turbine wheel. The gas leaves the turbine through the exhaust ducting.

Three dynamic seals in series prevent the pump fluid and turbine gas from mixing. Power from the turbine is transmitted to the pump by means of a one-piece shaft.

OXIDIZER TURBOPUMP

The oxidizer turbopump is mounted on the thrust chamber diametrically opposite the fuel turbopump.

It. is a single-stage centrifugal pump with direct

J-2 Major Component Breakdown

SATURN V NEWS

above the fuel inlet manifold. Power from the tur­bine is transmitted by means of a one-piece shaft to the pump. The velocity of the liquid oxygen is increased through the inducer and impeller. As the liquid oxygen enters the outlet volute, velocity is converted to pressure and the liquid oxygen is dis­charged into the outlet duct at high pressure.

Bearings in the liquid hydrogen and liquid oxygen turbopumps are lubricated by the fluid being pumped because the extremely low operating temperature of the engine precludes use of lubricants or other fluids.

The General Electric Co., under a prime contract with NASA, operates and maintains the facility, providing site services, technical systems, and test support to NASA and to stage contractors and other tenants.

North American Aviation, Inc., through its Space Division, is the prime contractor to NASA for devel­opmental and acceptance testing of second stages. SD personnel conduct the tests within the second stage test complex.

The Boeing Company is the prime contractor to NASA for developmental and acceptance testing of first stages. Stages manufactured by Boeing will be tested by the company in the first stage test complex.

The U. S. Army Corps of Engineers is NASA’s agent for land acquisition, design engineering, and construction.

Static Firing—The Marshall Space Flight Center captive fired all five F-l engines of the Saturn V S-IC-T for 16!4 seconds on May 6,1965. Later they were fired for 41 seconds.

MISSISSIPPI TEST FACILITY

NASA has developed the Mississippi Test Facility, a field organization of the Marshall Space Flight Center, as a testing site for the Saturn V launch vehicle’s two lower stages.

Acceptance testing of first and second stages will be conducted at the $300 million facility. In addi­tion, limited repair and modification of J-2 engines will be performed at MTF on behalf of all NASA operations in the Southeast.

First Stage Test Stand at MTF

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Stage Hoisted—The all-systems test version of the second stage is lifted into its test stand at MTF.

Management, operational, and support personnel engaged at the Mississippi Test Facility after its current construction and development work is complete in 1967 will number approximately 3,000.

Checkout—Engineers and technicians of North American are shown in the second stage Test Control Center at the Mississippi Test Facility during final preparation for static firing of all­systems test model of the stage.

The MTF site was selected from 34 areas considered mainly because of its accessibility to water routes and its nearness (45 miles by water) to the Michoud Assembly Facility in New Orleans. The govern­ment-owned fee area comprises 13,424 acres and is surrounded by an acoustic buffer zone involving an additional 128,526 acres in Hancock and Pearl River counties and Saint Tammany Parish.

Static Firing—A giant plume of vapor billows skyward during the first static firing test at MTF. The Saturn second stage, built for NASA by North American Aviation, Inc., burned for 15 seconds April 23, 1966.

MTF is composed of three principal complexes including approximately 60 buildings and structures. Among predominant features are the three huge test stands in the Saturn V complex. There are two separate stands for testing second stages. The first stage test stand is a dual-position structure which, with overhead crane, towers over 400 feet. The Laboratory and Engineering Complex houses engineering, administrative, and technical per­sonnel. The Industrial Complex has facilities for equipment and personnel necessary for site and test support maintenance.

The relatively small force of NASA personnel as­signed to MTF has overall management and super­visory responsibilities in overseeing the work of the contractors. NASA personnel are also respon­sible for final evaluation of static firings and is­suance of flightworthiness certificates to stage con­tractors.

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

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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.

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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

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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.