Diameter 21.7 ft Height 58.4 ft Weight (empty) 23,000 lb Weight (loaded) 253,000 lb Burn Time 470 sec (approx)
Major Structural Components
- AFT INTERSTAGE
- AFT SKIRT
- THRUST STRUCTURE
- PROPELLANT TANK
- COMMON BULKHEAD
- FORWARD SKIRT
Propulsion: One bipropellant J-2 engine
Total Thrust: 200,000 lb
- Liquid Hydrogen 38,000 lb (64,000 gal)
- Liquid Oxygen 191,000 lb (20,000 gal)
Basically, the S-IVB stage is an aluminum airframe structure powered by a single 200,000 pound-thrust J-2 engine which burns liquid oxygen and liquid hydrogen. The structure has a bipropellant tank capacity of approximately 230,000 pounds of usable fuel and oxidizer.
STAGE FABRICATION AND ASSEMBLY
The S-IVB stage structure consists of an aft interstage assembly, aft skirt assembly, thrust structure assembly, propellant tank assembly, and forward skirt assembly. The propellant tank assembly consists of a single tank separated by a common bulkhead into a fuel compartment and an oxidizer compartment.
Aft Interstage Assembly
The aft interstage is a cylindrical structure fabricated of stringer-stiffened aluminum skin panels. The aft interstage supports the S-IVB stage and provides a structural interface with the S-IB first stage. Vehicle design is such that the aft interstage remains with the S-IB stage when the two stages are separated in flight. Four solid propellant retrorocket motors are located equidistant around the base of this structure to brake the S-IB stage during separation.
S-IVB Stage Exploded View
Aft Skirt Assembly
The aft skirt is a cylindrical structure fabricated of aluminum, stringer-stiffened, skin panels and provides structural interface between the aft interstage and propellant tank assembly. After first stage burnout, the S-IB/S-IVB separation plane forms a part of the aft skirt structure. Two auxiliary propulsion system engine modules, three ullage rocket modules, and the aft umbilical connector plate are located on the aft skirt assembly.
Aft interstage /Aft skirt
Thrust Structure Assembly
The thrust structure is a truncated, cone-shaped structure fabricated of aluminum skin and stringers. It is bolted to the aft dome of the propellant tank assembly to provide an attach point for the J-2 engine. The thrust structure distributes J-2 engine thrust over the entire tank circumference. In addition, hydraulic system components, propellant feed lines, pneumatic components, and miscellaneous components which support engine operation are mounted on the thrust structure assembly.
Thrust Structure /Propellant Tank
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 dome assembly rotates in the fixture under a stationary welding head which is automatically positioned by a servo-controlled sensing element. To complete the hemisphere, a 30-inch "dollar" segment is welded in the top center of the dome. Subsequently, all fittings for various tank connections are installed by machine weld.
The common bulkhead, which forms the physical separation between the LOX and hydrogen tanks, is a 130-inch- radius hemisphere consisting of two aluminum domes separated and insulated by a fiberglass honeycomb core. The honeycomb core is bonded between the two domes under heat and pressure. Edges of two peripheral tees are welded together to provide a seal for the core, and structurally join the top and bottom domes.
Joining the common bulkhead and the aft dome completes the LOX tank subassembly. A slosh-baffle located within the LOX tank breaks up any wave action of the oxidizer during flight. The baffle is made up of four rings supported by the "A" frames.
Forward Skirt Assembly
The forward skirt is a cylindrical aluminum skin and stringer structure that provides a hard attach point for the IU. In addition, the forward 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 thermo conditioning system.
Forward Skirt Assembly
Final assembly of the S-IVB stage propellant tank structural components is accomplished in the assembly and welding tower. The assembly is then removed from the tower and transported to the insulation chambers building where the LH2 tank insulating tiles are fitted and installed. a glass cloth liner placed on the insulation and a sealant added. Propulsion system components, internally mounted in the LH2 tank, are installed following the completion of tank insulating. The structure is then returned to the assembly tower where the foward and aft skirts and thrust structure are installed.
Final installation of various subsystem components is performed in a checkout tower, along with the installation and alignment of the J-2 engine. The stage is in a vertical position in the tower where a complete stage checkout is conducted. 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.
Assembly Tower No. 2
S-IVB STAGE SYSTEMS
Major systems required for S-IVB stage operation are the propulsion system, flight control system, electrical power and distribution system, instrumentation and telemetry system, environmental control system, and ordnance system.
The propulsion system consists of the J-2 engine, propellant system, pneumatic control system and propellant utilization system. The J-2 engine burns LOX as an oxidizer and LH2 as fuel at a nominal mixture ratio of 5:1. Both fuel and oxidizer systems have a vent and relief capability to protect the propellant tanks from overpressurization. The pneumatic control system regulates and controls both the oxidizer and fuel systems. The propellant utilization (PU) system assures simultaneous and precise fuel and oxidizer depletion by controlling engine mixture ratio.
The engine system consists of the J-2 engine, propellant feed system, start system, gas generator system, control system, and a flight instrumentation system. The propellant feed system utilizes independently driven, direct drive fuel and oxidizer turbopumps to supply propellants at the proper mixture ratio to the engine combustion chamber. Additional information on the J-2 engine system may be found in the J-2 Engine section.
The propellant system consists of related stage subsystems to support propulsion of the S-IVB, and includes the oxidizer system, fuel system, pressurization system, tank venting system, and chilldown recirculation system.
LOX is loaded into the LOX tank at a temperature of -297 degrees F. The LOX tank capacity is 2,828 cubic feet which provides tankage for approximately 191,000 pounds (20,000 gallons) of usable LOX. The tank is pressurized with gaseous helium at 37 to 40 psia, and is maintained at this pressure during liftoff, boost, and stage engine operation.
Propulsion System Components
Fill and Drain: The LOX filling operation consists of purging and chilldown of the tank, and filling in four stages: slow fill, fast fill, replenish (topping), and pressurization. The ground controlled combination vent and relief valve is pneumatically opened at the start of the fill operation. During slow fill LOX is loaded at a rate of 300 gpm until a 5 percent of full level is attained, then fast fill at 1,000 gpm is initiated. When 93 per cent of the LOX has been loaded, the fill rate is reduced to a rate of 0 to 270 gpm. If for any reason, the LOX tank becomes overpressurized during fill, a vent malfunction occurs, or if there is an excessive LOX fill flow, a pressure switch signals the LOX ground fill valve to close.
The LOX tank is capable of being unloaded by reversing the flow through the fill system under tank pressure and/or from gravity effect. Drain capacity is at 500 gpm at 33 psia.
LOX Tank Pressurization: The LOX tank is pressurized at 39.5 +/- 0.5 psia from a ground supply of cold helium regulated to -360 degrees F. Following liftoff, the LOX tank pressure is maintained from helium storage spheres located in the LH2 fuel tank that have been charged to 3,000 +/- 100 psi at -360 degrees F. The engine heat exchanger heats and expands a portion of the helium flow before it is fed into the LOX tank. An ullage tank pressure switch controls inflight pressurization by opening or closing the cold helium flow valve as required. In case of pressure switch failure inflight, a pressure switch and plenum chamber act as a backup pressure regulator.
LOX Tank Vent-Relief System: The LOX tank vent-relief system consists of a tee assembly with a pneumatically-operated vent valve and a backup relief valve. Pneumatic operation is provided by the LOIS vent actuation module using helium gas from the pneumatic control system. The vent-relief valve is opened during the ground fill operation and closed prior to pressurization. During fill operations, the vent valve is capable of venting all LOX vapor.
The relief valve backup system automatically relieves at 45 psia and reseats at 42 psia. During liftoff and non- powered stage flight, pressure relief of venting is not anticipated. However, the vent system becomes operational in the event of LOX tank overpressurization.
LOX Feed System: Prior to vehicle liftoff, all LOX feed system components of the J-2 LOX turbopump assembly must be "chilled" to operating temperature for proper operation. Chilldown of the LOX system is accomplished by a closed loop, forward flow recirculation system. On command from the IU, a prevalve in the LOX feed duct closes and a shutoff valve opens. An auxiliary electrically driven centrifugal chilldown pump, mounted in the LOX tank, starts and LOX chilldown circulation begins.
LOX is circulated from the LOX tank, through the low pressure feed duct, to the J-2 engine LOX pump and bleed valve, then back to the LOX tank through return lines. The pump is capable of delivering a minimum flow rate of 31 gpm at 25 psia. Recirculation chilldown continues through the boost phase and up to the time for J-2 engine ignition. In the event of an emergency, the chilldown system shutoff valve closes upon command from the IU.
A low pressure supply duct supplies LOX from the tank to the engine at a nominal flow rate of 391 pounds-per-second at -297 degrees F at 25 psia and up.
The main LOX feed valve is a 4-inch butterfly type valve and opens in two distinct steps; the first, a partially opened position; the second, a fully opened position. The LOX feed valve is solenoid controlled.
A signal from the engine sequencer energizes the LOX feed valve, as required, to obtain steady-state operation. During steady-state operation, LOX feed is regulated by a propellant utilization valve which controls the oxidizer flow to the engine. A complete description of engine operation may be found in the J-2 Engine section.
LH2 is loaded into the insulated fuel tank at a temperature of -422.9 degrees F. The tank capacity is 10,446 cubic feet, ensuring approximately 38,000 pounds (64,000 gallons) of usable fuel. The tank is pressurized from a ground source of helium at 30.5 +/- .5 psia. During liftoff, boost and stage engine operation pressure is maintained in the fuel tank at 28 to 31 psi.
Fill and Drain: The LH2 loading operation consists of purging, chilldown of the tank, and filling in four stages: slow fill, fast fill, replenish (topping), and pressurization.
Immediately prior to LH2 input into the tank, a combination vent and relief valve is pneumatically opened. Loading is then initiated into the tank at 500 gpm until 5 per cent of full level is reached, then fast fill begins. During fast fill, LH2 is supplied to the tank at 3,000 gpm. When 93 per cent of the load is completed, a replenish or topping loading rate is initiated at between 0 to 270 gpm. Filling is automatically terminated at 100 per cent full. During the final topping operation, the fuel tank venting system is closed and the tank is simultaneously pressurized from the ground source of helium.
If overpressurization of the tank should occur during fill, or during the boost phase, a relief valve, which is spring loaded to open at 43 psia and close at 40 psia, is actuated to relieve excess pressure.
The LH2 tank is capable of being unloaded through the fill system. LH2 unloading is accomplished by reversing the flow through the fill system under tank pressure and/or from gravity effect.
Fuel Tank Pressurization: During initial tank pressurization, an external tank connection is made to a ground supply of helium. The helium is supplied to the fuel tank at -360 degrees F at 600 psig. When the tank ullage pressure reaches a maximum of 30.5 +/- .5 psia, a pressure switch sends a signal to deactivate the ground pressurization valve indicating that a satisfactory liftoff pressure has been attained, and pressurization is discontinued.
During liftoff and prior to J-2 engine start, additional pressurization is not required, as tank ullage pressure will be maintained from fuel boiloff.
At the initiation of J-2 engine start, GH2 is bled from the J-2 engine at 750 psia, -260 degrees F to provide ullage pressure during fuel depletion. The pressure bled from the engine into the fuel tank is controlled by a fuel tank pressurization control module.
Fuel Tank Vent-Relief System: Venting of the LH2 tank is accomplished by a vent and relief system capable of relieving all excess pressure accumulated from overpressurization or fuel boiloff during fill and flight operation. During fill, vaporization is vented through a self-sealing disconnect located in the forward skirt. During liftoff and flight, the gases are vented overboard through a non-propulsive exhaust.
The venting system consists of an actuation control module, vent valve, relief valve, directional control valve, and a non-propulsive overboard exhaust. Actuation of the vent valve is commanded from an external ground signal during fill operations, and from the flight sequencer during liftoff and flight. The vent valve is designed to open in a maximum of 0.1 second upon command.
The relief valve, which provides a backup capability in case of vent valve failure, opens at 42 psia and reseats at 39 psia, and has a flow/relief capability of 2 pounds/second at sea level.
A directional control valve directs excessive pressures through the ground disconnect during fill, and directs excessive pressures through the nonpropulsive vent during liftoff and flight. The non-propulsive vent system extends from the directional control valve into two 4-inch vent lines that terminate into two non-propulsive exhaust ports. The ports are located 180ø apart, in the forward skirt area. The ports are arranged to direct the exhaust for total thrust cancellation.
LH2 Feed System: Prior to vehicle liftoff, all LH2 feed system components of the J-2 turbopump assembly must be "chilled" to assure proper operation. Chilldown of the LH2 system is accomplished by a closed loop, forward-flow, recirculation system. On command from the IU, the pre-valve in the LH2 feed duct closes and the chilldown shutoff valve opens. An auxiliary electrically-driven LH2 chill-down pump mounted in the LH2 tank, circulates the LH2 within the system and is capable of a minimum flow rate of 135 gpm at 6.1 psi.
LH2 is circulated from the LH2 tank through the low pressure feed duct, through the J-2 engine fuel pump, the fuel bleed valve, and back to the tank through a return line. Recirculation chilldown continues through the boost phase and up to J-2 engine ignition. In the event of an emergency shutdown requirement, the chilldown system shutoff valve is closed upon command from the IU.
LH2 is supplied to the J-2 engine through a vacuum-jacketed, low-pressure duct at a flow rate of 81.26 pounds per second at -423 degrees F., 28 psia. The duct is located in the fuel tank side wall above the common bulkhead joint and is equipped with bellows to compensate for thermal motion. Signals from the engine sequencer energizes the LH2 feed valve, as required to obtain steady-state operation. A complete description of engine operation may be found in the J-2 Engine section.
PROPELLANT UTILIZATION SYSTEM
The primary function of the PU system is to assure simultaneous depletion of propellants by controlling the LOX flow rate to the J-2 engine. It also provides propellant mass information for controlling the fill and topping valves during propellant loading operations. The system consists of mass sensors, an electronics assembly, and an engine-mounted mixture ratio valve.
During loading operations, the mass of propellants loaded is determined within 1 per cent by the mass sensors. Tank over-fill sensors act as a backup system in the event the loading system fails to terminate fill operations.
Continuous LH2 and LOX residual readout signals are provided throughout S-IVB powered flight. The difference between the fuel and oxidizer mass indications, as sensed by the mass sensors, are continually analyzed and are then used to control the oxidizer pump bypass flow rate, which changes the engine mixture ratio correspondingly. The static inverter/converter supplies the analog voltages necessary to operate the PU system. It is commanded "on" and "off" by a switch selector and sequencer combination.
PNEUMATIC CONTROL SYSTEM
The pneumatic control system provides GHe pressure to operate all S-IVB stage pneumatically operated valves with the exception of those provided as components of the J-2 engine. GHe is supplied from an ambient helium sphere, pressurized from a ground source before propellant fill operations, at 3,000 +/- 100 psia at 70 degrees F. for valve operation. The sphere is located on the thrust structure and is pre-conditioned to above 70 degrees F from the environmental control system before liftoff.
The pneumatic control system provides regulated pressure at 475 +/-25 psig for operation of the LH2 and LOX vent-relief valves during propellant loading, LH2 directional control valve, LOX and LH2 fill and drain valves during loading, and the GH2 engine start system vent-relief valve. It also provides operating pressures for the LH and LOX turbopump turbine purge module, LOX chilldown pump purge module control, LOX and LH2 prevalves, and the LOX and LH2 chilldown shutoff valves.
The pneumatic control subsystem is protected from overpressure by a normally open solenoid valve controlled by- a don nstream pressure-sensing switch. At pressures greater than 535 +15, -10 psia, the pressure switch actuates and closes the valve. At pressures below 450 +15, - 10 psia, the pressure switch drops out and the solenoid opens, thus acting as a backup regulator.
Flight Control System
The flight control system provides stage thrust vector steering and attitude control. Steering is achieved by gimbaling the J-2 engine during powered flight. Hydraulic actuator assemblies provide J-2 engine deflection rates proportional to steering signal corrections supplied by the IU.
Stage roll attitude during powered flight is controlled by firing the auxiliary propulsion system (APS) attitude control engines.
J-2 Engine Hydraulic System Components
The hudraulic system performs engine positioning upon command from the IU. Major components are a J-2engine- driven hydraulic pump, an electrically driven auxiliary hydraulic pump, two hydraulic accumulator assemblies, and an accumulator-reservoir assembly.
The electrically driven auxiliary hydraulic pump is started before vehicle liftoff to pressurize the hydraulic system. Electric power for the pump is provided by a ground source. At liftoff, the pump is switched to stage battery power. Pressurization of the hydraulic system restrains the J-2 engine in a null position with relation to the S-IVB stage centerline, preventing pendulum-like shifting from forces encountered during liftoff and boost. During powered flight, the J-2 engine may be gimbaled up to +/- 7 degrees in a square pattern by the hydraulic system upon command from the IU.
Engine-Driven Hydraulic Pump
The engine-driven hydraulic pump is a variable displacement type pump capable of delivering hydraulic fluid under continuous system pressure and varying volume as required for operation of the hydraulic actuator assemblies. The pump is driven directly from the engine oxidizer turbopump. A thermal isolator in the system controls hydraulic fluid temperature to ensure proper operation.
Auxiliary Hydraulic Pump
The auxiliary hydraulic pump is an electrically driven variable displacement pump which supplies a constant minimum supply of hydraulic fluid to the hydraulic system at all times. The pump is also used to perform preflight engine gimbaling checkouts, hydraulically lock the engine in the null position during boost phase, maintain system hydraulic fluid at operating temperatures during other than the powered phase, and augment the engine- driven hydraulic pump during powered flight. It also provides an emergency backup supply of fluid to the system.
Hydraulic Actuator Assemblies
Two hydraulic actuator assemblies are attached directly to the J-2 engine and the thrust structure, and receive IU command signals to gimbal the engine. The actuator assemblies are identical and interchangeable.
The accumulator-reservoir assembly is an integral unit mounted on the thrust structure. The reservoir section is the storage area for hydraulic fluid; the accumulator section supplies peak system fluid requirements and dampens high-pressure surges within the system.
AUXILIARY PROPULSION SYSTEM
The APS includes modules that provide three-axis attitude control. Two APS modules are mounted 180 degrees apart on the aft skirt assembly. Each APS module contains three, 150-pound-thrust engines. Three solid propellant ullage rocket motors are mounted 120 degrees apart on the aft skirt assembly.
APS module engines are fired in short bursts for three-axis attitude control during coast. Minimum engine firing pulse duration is approximately 70 milliseconds.
Each APS module contains an individual oxidizer system, fuel system, and pressurization system. The modules are self-contained and easily detached for separate checkout and envirommental testing.
The individual engines are approximately 15 inches long with exit cones approximately 6.5 inches in diameter. Engine cooling is accomplished by an ablative process.
Ignition is unnecessary because fuel and oxidizer are hypergolic (self-igniting). Nitrogen tetroxide (N2O4), the oxidizer, is stable at room temperature.
Approximately 37.1 pounds of usable oxidizer is stored in the upper section of a common oxidizer/ fuel tank of the expulsion bellows type. Stored high-pressure helium is used for pressurizing oxidizer and fuel tanks.
The fuel, monomethyl hydrazine (CH3N2H3), is stable to shock and extreme heat or cold. Approximately 22.9 pounds of usable fuel is stored in the lower section of the common oxidizer/fuel tank.
Three solid propellant Thiokol TX-280 rocket motors, each rated at 3,390 pounds of thrust, are ignited during separation of the S-IB and S-IVB stages for ullage control. This thrust produces additional positive stage acceleration during separation, and position LOX and LH2 propellants toward the aft end of their tanks to cover outlets. In addition, propellant boiloff vapors are forced to the forward end where they are safely vented overboard. Tank outlets are covered to insure a net positive suction head (NPSH) to the propellant pumps, thus preventing possible pump cavitation during J-2 engine start.
Electrical Power and Distribution System
Four battery-powered systems provide electrical requirements for S-IVB stage operation. Forward Power System No. 1 includes a 28 vdc battery and power distribution equipment for telemetry, range safety system No. 1, forward battery heaters, and a power switch selector located in the forward skirt area.
Forward Power System No. 2 includes a 28 vdc battery and power distribution equipment for the PU assembly, inverter-converter, and range safety system No. 2.
Aft Power System No. 1 includes a 28 vdc battery and power distribution equipment for the J-2 engine, pressurization systems, APS modules, TM signal power, aft battery heaters, hydraulic system valves, and stage battery sequencer.
Aft Power System No. 2 includes a 56 vdc battery and power distribution equipment for the auxiliary hydraulic pump, oxidizer chilldown inverter, and fuel chilldowm inverter.
Silver-oxide, zinc batteries used for electrical power and distribution systems are manually activated. The batteries are "one-shot" units, and not interchangeable due to different load requirements.
Electrical power and distribution systems are activated by command through the aft umbilical prior to liftoff.
Telemetry and instrumentation System
Radio telemetry is used for transmission of stage instrumentation information to ground receiving stations. Five transmitters, using two separate antenna systems, are capable of returning information on 60 continuous output data channels and 81 sampled data channels during S-IVB flight. Three different modulation systems are utilized: Pulse Amplitude Modulated/FM/FM (PAM/FM/ FM); Single Sideband/FM (SS/FM); and Digital Data Acquisition System (DDAS).
A DDAS airborne tape recorder stores sampled data normally lost during over-the-horizon periods of orbital missions, and plays back information when in range of ground stations.
Transducer input signals constitute the PAM input. The PAM systems use an electronically switched mixing network that samples up to 30 channels of transducer inputs at 120 times a second. Deviations in transducer input voltages are represented as output pulses of varying amplitude for subsequent evaluation.
The SS/FM system is reserved for pertinent research requirements. Sonic vibration and acoustical data needed for manned flight development will be transmitted by this system.
DIGITAL DATA ACQUISITION SYSTEM
The DDAS is used during automatic checkout on the ground, and for information playback from the airborne tape recorder when the stage is within communication range (line of sight) of ground stations. Redundant data is recorded in parallel from the PAM inputs and played back at high speed upon ground command. Continuous data flow charts are then pieced together at ground stations.
Environmental Control Systems
AFT SKIRT AND INTERSTAGE CONTROL THERMOCONDITIONING
The thermoconditioning and purge system purges the aft skirt and aft interstage of oxygen and combustible gases and distributes temperature controlled air or gaseous nitrogen around electrical equipment in the aft skirt.
The purging gas, supplied from a ground source through the umbilical, passes over the electrical equipment and flows into the aft interstage area. Some of the gas is directed through each of the auxiliary propulsion modules and exhausts into the interstage. A duct from the skirt manifold directs air or GN2 to a thrust structure manifold. From the thrust structure manifold supply duct, a portion of air or GN2 is directed to a shroud covering the hydraulic accumulator reservoir.
Temperature control is accomplished by two dualelement thermistor assemblies located in the gaseous exhaust stream of each of the auxiliary propulsion modules. Elements are wired in series to sense average temperature. Two series circuits are formed, each circuit utilizing one element from each thermistor assembly. One series is used for temperature control, the other for temperature recording.
FORWARD SKIRT THERMOCONDITIONING
Electrical equipment in the S-IVB forward skirt area is thermally conditioned by a heat transfer system, using "cold plates" on which electronic components are mounted, and through which coolant fluid circulates. Coolant is pumped through the system from the IU and returned. Heat from electrical equipment attached to the cold plates is dissipated by conduction through the mounting feet and the cold plates to the fluid. Refer to the Instrument Unit section for a complete description of the IU environmental conditioning system.
FORWARD SKIRT AREA PURGE
The forward skirt area is purged with gaseous nitrogren to minimize fire and explosion hazards while propellants are loaded or stored in the stage. Gaseous nitrogren is supplied and remotely controlled from a ground source.
The Ordnance systems perform stage separation, retrorocket ignition, ullage control rocket ignition and jettison, and range safety functions.
STAGE SEPARATION SYSTEM
The stage separation system consists of a severable tension strap, mild detonating fuse (MDF), exploding bridgewire detonators (EBW), and EBW firing units.
The severable tension strap houses 2 redundant MDF cords in a "V" groove circumventing the stage between the aft skirt and aft interstage at the separation plane. Ignition of the MDF cords is triggered by a signal from the S-IB stage sequencer about 1 second after engine cutoff. The exploding bridgewire detonators ( EBW ), and EBW firing units.
The MDF consists of a flexible metal sheath surrounding a continuous core of high explosive material. Once detonated, the explosive force of the MDF occurs at a rate of 23,000 feet per second.
The EBW detonator is electrically activated for initiating the MDF explosive train. A 2,300 vdc, 1,500 ampere pulse is applied to a small resistance wire and a spark gap. The high voltage electrical arc across the spark gap ignites a charge of high explosive material which in turn detonates the MDF. The high voltage and current requirement for ignition renders this system safe from random ground or vehicle electrical power. Upon command, each EBW firing unit supplies high voltage and current required to fire a specific EBW detonator.
RETROROCKET IGNITION SYSTEM
Four solid propellant retrorockets are mounted equidistant around the aft interstage assembly, and when ignited, assure clean separation of the S-IVB stage from the S-IB stage by decelerating or braking the spent booster. Each retrorocket is rated for a nominal thrust of 35,000 pounds, weight of 384 pounds, and burn time of about 1.5 seconds.
A signal from the S-IB stage initiates two EBW firing units located on the aft interstage. The EBW firing units ignite two detonator manifolds, which in turn ignite the retrorockets.
ULLAGE CONTROL ROCKET ENGINE AND JETTISON SYSTEM
Three solid propellant ullage rockets, located on the S-IVB aft skirt just forward of the stage separation plane, are ignited on signal from the stage sequencer by EBW initiators. The ullage rockets provide a positive "G" force to settle propellants in the tanks before ignition of the J-2 engine.
After firing, the burned-out ullage rocket casings and associated EBW firing units are jettisoned to reduce stage weight. Upon command from the stage sequencer, two forward and aft frangible nuts, which secure each rocket motor and its fairing to the stage, are detonated by confined detonating fuze (CDF), to free the entire assembly from the vehicle.
RANGE SAFETY SYSTEM
The range safety system terminates vehicle flight upon command of the range safety officer. Redundant systems are used throughout to provide maximum reliability.
Four antennae, mounted around the periphery of the S-IVB forward skirt assembly, feed two redundant range safety receivers located in the forward skirt assembly. Both receivers have separate power supplies and circuits. A unique combination of coded signals must be transmitted, received, and decoded to energize this destruct system.
A safety and arming device prevents inadvertent , initiation of the explosive train by providing a positive isolation of the EBW detonator and explosive train until arming is commanded. Visual and remote indications of SAFE and ARMED conditions are displayed at all times at the firing center. Upon proper command, EBW firing units activate EBW detonators.
A CDF; detonated by the safety and arming device, explodes a flexible linear-shaped charge which cuts through the tank skin to disperse both fuel and oxidizer.
Copyright 1997-2005 by John