Category Apollo Saturn V News Reference

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.

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

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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.