Category Apollo Saturn V News Reference

F-l and J-2 engines for the Saturn V launch vehicle are manufactured at Rocketdyne’s main complex in Canoga Park, Calif. F-l static testing is conducted at the Edwards Field Laboratory located at the NASA Rocket Engine Test Site, Edwards, Calif., about 125 miles northeast of Los Angeles, and the J-2 is tested at Rocketdyne’s Santa Susana Field Laboratory located about 10 miles from Canoga Park. Rocketdyne operates the Neosho Facility (Missouri), which produces and tests subcompo­nents of the J-2 and F-l engines.

F-l Test Stands—Three of six stands for testing F-l rocket engines or components at full thrust are visible in this aerial view of NASA Rocket Engine Test Site.

F-l Test Firing… An F-l rocket engine developing 1,500,000

pounds of thrust is tested at NASA Rocket Engine Test Site. The stand is one of six in the complex.

Manufacturing of components and final assembly of both engines are carried out in eight buildings in

the Canoga complex. These facilities are equipped with general purpose machine tools for precision and heavy machining as well as some 20 numerically controlled machines for performing programmed multiple machining operations. Also included are two of the largest gas-fired brazing furnaces of their type for brazing of thrust chamber tubes and in­jectors, eight units for ultrasonic cleaning, 21 in­stallations for Gamma and X-ray inspection, more than 50 environmentally controlled areas for ultra­clean assembly operations, sheet metal prepara­tion, precision cleaning, and receiving inspection.

F-l Flight Engine Firing

An Engineering Development Laboratory pro­vides specialized facilities to support manufacturing programs. These facilities include a high-flow water test facility for checking propellant systems, 12 concrete cells for conducting hazardous tests, 28 environmental test chambers, a photo-elastic lab­oratory, two pneumatic flow benches, six vibration test rooms, and others for checking components as well as complete engines.

Research and development testing of F-l turbo­machinery, gas generators, heat-exchangers, seals, and splines is conducted on two test stands and three components test laboratories at Santa Susana.

Six large test stands, with a total of eight test posi­tions, and associated shops and support facilities at the Edwards Field Laboratory are used for testing complete F-l engines as well as injectors.

Six large engine test stand positions at the Santa Susana Field Laboratory are used for testing the J-2. One of these stands is equipped with a steam injection diffuser for altitude simulation testing. J-2 turbopumps, gas generators, valves seals, bear­ings, and other components are tested in 22 test cells in five component test laboratories in Santa Susana.

Pump Tests-Flames from gases burned during test of an F-l engine turbopump shoot more than 150 feet in air.

J-2 Testing—A hydrogen fueled J-2 rocket engine is tested under ambient altitude conditions at Santa Susana.

LOX is delivered to the engines by five 17-inch suction lines which pass through the fuel tank in five LOX tunnels. LOX suction ducts make up the lines from the LOX tank to the prevalves in the thrust structure. The ducts are equipped with gim­bals and sliding joints to counteract vibration and swelling or contraction caused by temperature. In­side the tunnels, air acts as the insulation between the LOX-wetted lines and the fuel-wetted tunnels.

LOX Delivery

LOX level engine cutoff sensors in the suction lines assure safe engine shutdown and leave a minimum amount of unused LOX in the system.

In case of emergency, LOX prevalves in each suc­tion line can stop the flow of LOX to the engines.

LOX CONDITIONING SYSTEM
it will result in gaseous oxygen (GOX). If heat is increased, the result is boiling and not temperature increase since evaporation is a cooling process. Depth in a body of LOX can increase due to the increase in hydrostatic pressure.

LOX Conditioning

The greatest chance for overheating in the LOX system is in the transmission surface of the suction lines. Also, the suction lines are too slender for maintenance of self-contained convection currents. This situation is unacceptable since intense boiling can lead to LOX geysering, which in turn can dam­age the LOX tank structurally. In addition, too high a LOX temperature near the engine inlets can cause a cavity in the LOX pumps and interfere with nor­mal engine starting. Emergency bubbling or thermal pumping is used to correct this situation.

The bubbling technique sends helium into all five suction lines to cool the LOX rapidly. Ground sup­port supplies helium through an umbilical coupling, and filter valves and orifices control the flow of helium into the suction lines. Thermal pumping is a term used to define pumping relatively cold LOX from the LOX tank into the suction lines.

The gas generator valve is a hydraulically oper­ated valve which controls and sequences entry of propellants into the gas generator. Hydraulic fuel is recirculated through a passage in the valve hous­ing to maintain seal integrity and to prevent the fuel in the fuel ball housing from freezing. Fuel is also recirculated through a passage in the piston between the opening port and the closing port to prevent the piston О-ring from freezing.

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Gas Generator Assembly Including Control Valves

Gas Generator Injector

The gas generator injector directs fuel and oxi­dizer into the gas generator combustion chamber. It is a flat-faced, multi-orificed injector incorporat­ing a dome, a plate, a ring manifold, five oxidizer rings, five fuel rings, and a fuel disc. The gas gen­erator valve and the gas generator injector fuel inlet housing tee are mounted on the injector.

Fuel enters the injector through the gas generator fuel inlet housing tee from the gas generator valve. The fuel is directed through internal passages in the plate and injected into the combustion chamber through orifices in the fuel rings and the disc. Some of the orifices in the outer fuel ring also provide a cooling film of fuel for the combustion chamber wall. Oxidizer enters the injector through the oxi­dizer inlet manifold from the gas generator valve. The oxidizer is directed from the oxidizer manifold through internal passages in the plate and is in­jected into the combustion chamber through the orifices in the oxidizer rings.

Gas Generator Combustion Chamber

The gas generator combustion chamber provides a space for burning propellants and exhausts the gases from the burning propellants into the turbopump turbine manifold. It is a single-wall chamber located between the gas generator injector and the turbo­pump inlet.

The propellant level monitoring subsystem checks the level of propellants in both tanks to provide checkpoints for the sensors used in the propellant utilization and loading subsystems and to monitor propellant levels during firing. These functions are performed by sensors mounted on continuous Stillwells adjacent or parallel to the full-length capacitance probes in each tank. There are 14 sen­sors on each Stillwell to indicate various levels in the tanks.

ULLAGE MOTORS

The solid propellant ullage motors are used to pro­vide artificial gravity by momentarily accelerating the second stage forward after first stage burnout. This moment of forward thrust is required in the weightless environment of outer space to make certain that the liquid propellant is in proper posi­tion to be drawn into the pumps prior to starting of the second stage engines.

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Eight ullage motors are utilized on the stage where they are attached around the periphery of the inter­stage structure between the first and second stages. Each ullage motor measures 12.5 inches in diameter by 89 inches long and each provides 22,500 pounds of thrust for approximately 4 seconds. The motors utilize Flexadyne solid propellant in a formulation developed specifically to provide high performance and superior mechanical properties under operat­ing conditions encountered in space. Ullage motor nozzles are canted 10 degrees to reduce exhaust impingement against the interstage structure.

THERMAL CONTROL SYSTEM

Thermal control is provided by a ground-operated system which maintains proper temperatures for the equipment containers in the forward and aft skirt areas. Tempered air is used to cool the con­tainers before propellant loading. With prepara­tion for loading, the air is changed to nitrogen for container inerting and heating. Separate thermal control systems are provided for the forward and aft skirt areas. Each of the units contains a single manifold connected to each container, individual fixed-flow orifices, and individual relief holes from each container. Container insulation and thermal inertia preclude excessive temperature changes.

J-2 ENGINE DESCRIPTION

The Rocketdyne J-2 engine is a high performance, upper stage, propulsion system utilizing liquid hy­drogen and liquid oxygen propellants and devel – opes a maximum vacuum thrust of 225,000 pounds.

All J-2 engines are identical when delivered and may be allocated to either the second or third stage. Each engine is equipped to be restarted in flight. However, the restart capability will be utilized only in the third stage.

The single J-2 engine used in the third stage is gimbal-mounted so that it can be moved in flight and used to steer the stage. Five J-2 engines are arranged in a cluster in the second stage. The four outboard engines of the five-engine cluster are gimbal-mounted to provide the vehicle with pitch, yaw, and roll control. The center engine is mounted in a fixed position.

Major systems of the J-2 engine include a thrust chamber and gimbal assembly system, propellant feed system, gas generator and exhaust system, electrical and pneumatic control system, start tank assembly system, and flight instrumentation system.

Thrust Chamber and Gimbal System

The J-2 engine thrust chamber serves as a mount for all engine components. It is composed of the following subassemblies: thrust chamber body, in­jector and dome assembly, gimbal bearing assembly, and augmented spark igniter.

Thrust is transmitted through the gimbal mounted on the thrust chamber assembly dome to the vehicle thrust frame structure. The thrust chamber injec­tor receives the propellants from a dual turbopump system (oxidizer and fuel) under pressure, mixes the propellants, and burns them to impart a high velocity to the expelled combustion gases to pro­duce thrust.

THRUST CHAMBER

The thrust chamber is constructed of stainless steel tubes of 0.012-inch wall thickness. Tubes with thin walls are required for heat transfer purposes. The thrust chamber tubes are stacked longitudi­nally and furnace-brazed to form a single unit. The chamber is bell-shaped with a 27.5 to 1 expansion area ratio for efficient operation at altitude, and is regeneratively cooled by the fuel. Fuel enters from a manifold located midway between the thrust

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J-2 Assembly—Hydrogen fueled J-2 rocket engines for upper stages of Saturn V vehicles are completed on this assembly line. J-2 develops a maximum thrust of 225,000 pounds.

New Saturn V facilities built at the Marshall Space Flight Center at Huntsville, Ala., include the first stage static test stand, an F-l engine test stand, the Saturn V launch vehicle dynamic test stand, a J-2 engine facility, and ground support and component test positions.

The Marshall Center has completed a $39 million addition to its Test Laboratory for captive testing the Saturn V booster and F-l engines. The Test Laboratory addition is called the West Test Area. The largest structure in the area is the first stage test stand. Completed in 1964, the stand has an overall height of 405 feet.

Growth at Huntsville—The growth of rocket testing facilities at the Marshall Space Flight Center is contrasted here by the size of the first Redstone Arsenal test stand, second from left, and stands at right built for the Saturn V program.

The nearby single engine test stand is being used for research and development tests of the 1.5 million pound thrust F-l engine.

Control and monitoring equipment for the first stage and F-l engine stands is located in the area’s central blockhouse. Water needed to cool the flame deflectors of the two stands is pumped from a near­by high-pressure industrial water station.

The 365-foot tall Saturn V launch vehicle was placed in another unique Marshall Center test facility — the Dynamic Test Stand. Testing of the complete three stage vehicle and its Apollo spacecraft here was done to determine its bending and vibration characteristics. Tallest of Marshall Center’s tall towers, the dynamic test stand is 98 feet square.

Several tests of the liquid hydrogen-liquid oxygen powered engine have been conducted during the past year in Marshall’s J-2 engine test facility. Tank­age for the facility is a battleship version of the Saturn V third stage. The J-2 engine stand is 156 feet tall and has a base of 34 by 68 feet. It is located in the MSFC’s East Test Area.

Vibration Version – A ground test version of the Saturn V first stage moves through the West Test Area of the Marshall Space Flight Center. The large dynamic test stage was built to undergo vibration and bending tests. Test stand at right Is a single F-l engine facility.

I

Positioning—A Saturn V first stage is placed into a test stand at Marshall Space Flight Center.

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A portion of the Kennedy Space Center “spaceport” has been created at the Marshall Center’s ground support equipment test facility to check out giant mechanical swing arms which will be used on Launch Complex 39 to connect the Apollo/Saturn V space vehicle to the launcher tower. The 18-acre facility has eight swing arm test positions and one position for testing access arms to be used by Apollo astro­nauts.

Also built at Marshall are an F-l engine turbopump position in the East Test Area and a load test facil­ity in the Propulsion and Vehicle Engineering Lab­oratory.

A new Saturn V rocket “electrical simulator” or breadboard facility at the Marshall Center dupli­cates the electrical operation of the vehicle. Ele­ments simulated include the first stage booster, second stage, third stage, and an instrument unit.

Other Saturn V facilities at the Marshall Center include a booster checkout area, two new assembly areas and a components acceptance building.

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

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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.