S-IB STAGE DESCRIPTIONThe S-IB stage consists, basically, of a cluster of eight H-1 rocket engines (4 fixed inboard and 4 steerable outboard), a tail unit assembly, nine propellant containers, a spider beam unit assembly, and eight fin assemblies. The nine propellant containers attach to the tail unit assembly at the lower end and to the spider beam unit assembly at the top.
S-IB STAGE FABRICATION AND ASSEMBLY
Construction of the S-IB stage begins at the MSFC Michoud Assembly Facility with the fabrication of the tail unit assembly and the spider beam unit assembly. The tail unit assembly, propellant containers, and spider beam unit assembly are then brought together in a major assembly operation called clustering. After clustering, the eight engines and various pneumatic, mechanical, and electrical systems are installed to complete the assembly of the stage.
Tail Unit Assembly
The tail unit assembly consists of four radial thrust support outriggers and four radial fin support outriggers, all of which are attached to a barrel assembly core. The ends of the outriggers are joined by shroud panels which form the periphery of the tail unit. A lower shroud panel assembly encloses the engines and forms the engine compartment. The forward end of the assembly is closed by fire walls and the aft end is closed by flame shields. When assembled into the stage, four of the engines are attached to the barrel assembly and four are attached to the thrust outriggers.
The barrel assembly is comprised of an upper and a lower thrust ring, four shear web assemblies, and four skin panels.
The thrust support outrigger assemblies consist of two shear panels, a thrust beam, an actuator support beam, an outboard engine mounting pad, three webs, a bulkhead, a shroud support plate, and various angles and channels. The fin support and thrust support outriggers are similar. The fin support outriggers have no thrust support beam and no actuator support beam.
The water quench system, calorimeter purge system, and fire detection system, along with various lines, components, and electrical equipment that are parts of other stage systems are installed to complete the tail unit assembly.
Propellant Containers
The cylindrical section of each propellant container is built up of skin-milled, butt-welded, aluminum alloy segments internally reinforced with rings to form a monocoque type construction. Container wall thickness varies from top to bottom in relation to stress concentrations. Hemispherical bulkheads are welded to each end of the cylindrical section, and a sump is welded to the aft bulkhead. The forward bulkhead of the 105 inch diameter LOX container is fitted with a pressurization and vent manifold; the 70-inch diameter containers have openings for fuel or LOX vent manifold connections. A cylindrical skirt reinforced with longerons is attached to both the forward and aft bulkheads to complete the basic container. The units are cleaned, painted, pressure tested, and calibrated for precise volume before shipment to the Michoud Assembly Facility.
Internal and external equipment is installed to modify the basic container. Electrical equipment is installed in the aft skirt areas of all four fuel containers, and in the instrument compartments located in the forward skirts of fuel containers F-1 and F-2. A 20-cubic foot high-pressure sphere is installed into the forward skirt of fuel containers F-3 and F-4.
Spider Beam Unit Assembly
The spider beam unit assembly is assembled in a special fixture. A hub assembly is placed in the center of the fixture and the upper and lower splice plates are attached to the hub. Eight radial beams are attached to the hub assembly at 46-degree intervals and the outer ends of the radial beams are joined by cross-beams fastened with splice plates. When the basic structure is completed, hardware is installed that will he used to attach the propellant containers during the clustering operation.
Holes are drilled for use in attaching the S-IVB stage to the spider beam at the launch site. A special handling ring is attached to the spider beam and an alignment cheek is made in preparation for the clustering operation.
Installing Fuel Container Pressurization Sphere
Clustering
The S-IB stage is clustered in a horizontal position. Multiple-level platforms at the ends of the stage allow simultaneous assembly operations.
The assembly fixture for the S-IB stage is a truss structure with a front and a rear cradle which is supported by four adjustable leveling stands that allow vertical and horizontal movement for the alignment of components. Rollers in the front and rear cradles allow the stage to be rotated during assembly.
Positioning Tail Unit Assembly
Positioning Spider Beam Assembly
A large bridge crane is used to place the tail unit assembly in the rear cradle. The 106-inch diameter center LOX container is positioned so that the aft skirt can be fastened to the top end of the barrel assembly. The spider beam unit assembly is then attached to the forward end of the center LOX container. The tail unit assembly, center LOX container, and spider beam unit assembly are rotated as a unit to facilitate attachment of the outer containers.
The 70-inch diameter outer LOX containers are installed first. The containers are installed in an opposite-pair sequence to keep the assembly within the balance requirements of the assembly fixture. LOX containers are attached to the tail unit assembly and to the spider beam unit assembly with retaining bolts and eyebolt assemblies.
The fuel containers are installed in a similar manner to complete the container clustering operation. Fuel containers are attached to the tail unit assembly by a ball and socket arrangement, and to the spider beam unit assembly by a sliding pin and socket arrangement. This allows for the variable distance between the spider beam unit assembly and the thrust structure assembly caused by contraction of the LOX tanks at cryogenic temperatures.
Assembly operations are completed by installing lines, manifolds, cables, electrical and pneumatic components, engines, and instruments. The fin assemblies are attached to the stage as the vehicle is erected at the launch site.
H-1 ENGINE INSTALLATION
Before being installed in the stage, the H-1 engines undergo modification, consisting primarily of the installation of static test and flight measurement instrumentation and cables on inboard and outboard engines, and the installation of the hydraulic system on outboard engines.
The inboard engines are installed first, using a special handling fixture to insert each engine into the tail unit assembly. Fuel and LOX suction lines are installed and, when the last inboard engine is installed, the flame shield is installed and a series of operations connecting the inboard engines into other stage systems is performed. These operations include the installation of engine drain lines, turbine exhausts, GOX lines, purge systems, LOX supply lines to heat exchangers, and electrical cables.
The outboard engines are installed in approximately the same manner as the inboard engines. The engines are connected into the stage systems, and the electrical cables that monitor and control all eight engines are installed. Engine installation is considered complete after the flexible flame curtains, the heat shield panels, and other heat protective equipment has been installed.
Heat Shield Panel Installation
S-IB Fin Undergoing Vibration Test
FIN ASSEMBLIES
Each of eight fin assemblies is built around a front and rear spar, a hold down fitting, and two diagonal tubes which strengthen the structure. A leading edge is attached to the front spar and a trailing edge to the rear spar. A heat shield is attached to the trailing edge to protect the fin from engine exhaust. Skin panels are riveted to the fin structure to form a smooth aerodynamic surface.
The basic fin assembly is completed by installing instrumentation such as calorimeters, temperature gages, and strain gages which are part of the measuring program.
Electrical Fabrication and Assembly
The S-IB booster stage is much more than eight rocket engines and a cluster of propellant containers. In addition to thousands of feet of tubing and numerous mechanical and electromechanical valves and regulators, the S-IB booster stage contains 53 miles of wire terminating in approximately 73,000 electrical connections, that tie together almost 1,700 electrical and electronic components.
All of the electrical cables are fabricated by Chrysler. Automatic magnetic-tape programmed instruments check the assemblies for continuity between pins and perform high-voltage leakage tests.
Components of some of the major electronic units, such as a telemetry assembly, a multiplexer, or a timer assembly, are purchased from vendors. The components are assembled by Chrysler, qualified as a unit, functionally tested, and installed on the stage. Other units such as measuring racks, liquid level adapter racks, and distributors are fabricated by Chrysler.
S-IB Stage Checkout
Chrysler technicians and engineers conduct tests on the propulsion, electrical, control, instrumentation, and telemetry systems of the S-IB stage in accordance with the final acceptance test procedures. The tests are performed just prior to shipment of the stage to MSFC, Huntsville for static fire and again just prior to shipment of the stage to KSC, Cocoa Beach, Florida, for launch.
Dual checkout stations located adjacent to the final assembly area permit two stages to be checked out simultaneously. Each checkout station consists of a checkout bay and a control room.
A central computer complex and two telemeter ground stations support the operation of the checkout stations.
The components and systems within each stage are subjected to a series of tests to demonstrate the acceptability of each system and family of systems. A summary of these tests is presented in the following tabulation. The order of presentation is not necessarily the order in which the tests are made since many tests are run simultaneously.
- Antennas: Antennas, power dividers, and transmission lines are checked for insulation resistance, continuity, attenuation, and voltage standing wave ratio.
- Power Distribution: Resistance checks of the checkout station equipment and stage busses are performed before power is applied. Buss electrical isolation and distribution is verified with the application of power to each buss.
- Instrumentation Two categories of instrumentation are made:
1. Indirect measurements utilizing signal conditioners for evaluation of acoustic, temperature, pressure, vibration, liquid level, current, strain, and turbine rpm measurements.
2. Direct measurements using a direct output in the flight mode to evaluate acceleration, pressure, and signal levels.
- Mechanical Systems: Fuel, oxidizer, propul- sion, pneumatic, and hydraulic systems are pressure tested to verify system integrity and to assure that leakage is within the allowable specifications. While these systems are pres- surized, other system characteristics such as pressure switch actuation and de-actuation levels, valve timing characteristics, vent valve cracking and reseating pressures, and critical orifice flow rate are checked.
- Electrical Networks: Station equipment and stage redundant cutoff circuits are checked for the capability to initiate engine cutoff at any time a malfunction is detected. The flight sequence operation is checked by verifying circuit operations that occur during prelaunch, liftoff, and separation functions. Engine cut off is also verified by simulating the conditions of an engine failure, failure to launch due to loss of liftoff circuitry, and a malfunction after liftoff causing emergency engine cutoff.
- RF Systems: Tests are made to assure proper operation of the telemetry systems and the ODOP (offset doppler) tracking system. The tests assure that power output, voltage standing wave ratio, frequency, insertion loss and doppler shift are within specifications.
- Radiated Compatibility and Interference: Tests are performed to verify that radiated frequencies of the telemetry and television systems do not interfere with the operation of the command system.
- Control System: The outboard engines are tested to assure proper function of hydraulic actuators, control system linearity, and proper reaction to control accelerometer signals. During gimbaling, checks are made to assure proper clearance between the stage tail section and engines.
- Instrumentation Compatibility: This test assures compatibility between the stage measuring instruments and the telemetry systems. Each telemeter channel is tested to assure that a linear output response of proper level can be obtained.
- Simulated Flight: All vehicle systems are evaluated as a unit. The stage is programmed through a complete preflight and flight sequence during which time the umbilicals are retracted, the hydraulic system is activated, engine gimbaling tests are performed, and inboard and outboard engine cutoff sequences are verified. In addition, the Range Safety command system is activated and checked for proper operation.
S-IB STAGE SYSTEMS DESCRIPTIONS
Fuel System
The stage fuel system receives RP-1 fuel from a ground source, stores the fuel, and then supplies it to the eight H-1 engines. The system consists of four fuel containers, pressurization components, distribution manifolds, control valves, switches, sensors, piping, interconnect lines and the connecting hardware required to fill or drain the containers, bubble the fuel before flight, pressurize the containers, and supply the fuel to the engines.
Equal pressurization of the containers and uniform distribution of fuel to the engines is maintained through interconnect lines at the top and bottom of the containers. In the event of an engine failure, the fuel normally consumed by the inoperative engine is supplied to the operating engines.
Each container supplies fuel to one outboard and one inboard engine through suction lines connected to the container sump. Two engine cutoff fuel sensors generate a signal when the fuel is decreased to their level that will initiate inboard engine shutdown. Similar engine cutoff sensors in the LOX system will initiate inboard engine shutdown if LOX depletion occurs prior to fuel depletion.
Outboard engine shutdown occurs approximately 6 seconds after inboard engine shutdown. The outboard engines are normally shut down when engine thrust decay, resulting from LOX depletion, causes the outboard engine thrust OK pressure switches to deactuate. The outboard engines are shut down simultaneously since the thrust OK pressure switches on all outboard engines are interconnected at this time. However, if fuel depletion occurs prior to LOX depletion, the outboard engines will be shut down when fuel reaches the level of fuel depletion sensors located in the container sump.
Fifteen discrete level sensors and one continuous level sensor in each of the four fuel containers permit telemetric monitoring of fuel level during flight.
FUEL FILL
Prior to filling the containers with fuel, normally closed vent valves and a fuel fill and drain valve are opened. The vent valves are actuated by control pressure from a ground source supplied through the vent valve control quick-disconnect coupling and a solenoid valve. The fill and drain valve is actuated by control pressure from the ground source supplied through the opening control quick-disconnect coupling. Fuel is then pumped through the fill and drain nozzle into the sump of container F-1. The sumps of the containers are interconnected to ensure equal distribution of fuel. The fuel containers are initially filled to a predetermined level based upon a nominal density.
The rate of flow to the containers is controlled by a fuel tanking computer in the ground control station. The computer shuts off the supply when the containers are filled to the predetermined level. If the computer should malfunction and not shut off the fuel supply, an overfill sensor will initiate a signal to stop the fill sequence.
FUEL LEVELING
Fuel in the containers is maintained at the required level by adding or draining. The level is adjusted according to calcalations based on pressure differentials in container F-4 and on fuel density calculations based on temperature measurements from each container. The fuel density computer and the fuel tanking computer are connected to container F-4 through separate quick-disconnect couplings. Both computers are also connected to container F-4 through a common quick-disconnect coupling. Temperature sensors monitor the temperature of the fuel in each container for the density calculations and electrically transmit the results to the fuel density computer.
FUEL DRAIN
In the event that fuel must be drained from the system, the containers are pressurized from a ground source. The vent valves are closed, the fuel fill and drain valve is opened, and fuel under pressure flows from the containers through the open valve and the fill and drain nozzle to the ground storage tank. Container pressure for the draining operation is maintained between 3 psig and 10 psig.
SUCTION AND INTERCONNECT LINES
Eight suction lines conduct fuel from the container sumps to the engine turbopumps. Normally open fuel prevalves, located near the top of each suction line, may be closed by pneumatic pressure supplied by the stage control pressure system or by pneumatic pressure supplied from a ground source and through an orifice.
The fuel containers are interconnected at the top through the fuel container pressurization manifold. The fuel vent valves are connected into this manifold. The fuel containers are also interconnected at the sumps through interconnect lines. An antivortex device and screen is located at the aft bulkhead of each container.
FUEL BUBBLING
Pressurized GN2 from a ground source is bubbled through each fuel suction line to agitate the fuel and thereby maintain a uniform fuel temperature throughout each container. Fuel bubbling begins just before LOX fill and continues until the start of fuel container pressurization.
FUEL CONTAINER PRESSURIZATION
The fuel containers are pressurized with helium starting from approximately two minutes and thirty-three seconds prior to launch and continuing until the S-IB flight is completed. The container pressure maintains a pressure head for the engine fuel pumps and provides structural integrity by preventing the formation of a vacuum in the containers as fuel is depleted during flight. Major components of the pressurizing system are high-pressure helium storage spheres, solenoid valves, pressure switches, a sonic nozzle, and distribution lines.
Prior to launch, two 20-cubic-foot, high-pressure storage spheres are pressurized to 3,000 psig. Helium is supplied from the launch facility through the fuel container pressurization quick-disconnect coupling connected to one of the launch facility umbilicals. The helium passes through a filter and a cheek valve before it enters the storage spheres. From the storage spheres, helium flows through two normally open solenoid valves, a sonic nozzle, the distribution line, and into the fuel containers. The sonic nozzle restricts the flow of helium. Prior to flight, container pressure is maintained at 17 psig by a control pressure switch located at the top of container F-2. The switch controls the operation of the normally open solenoid valves and shuts off the helium supplied to the container when the pressure reaches 17 psig. If container pressure should exceed 21 psig, the vent control pressure switch actuates; the vent valves open, and the containers are vented Until normal pressure is maintained again. Further overpressurization protection is provided through the relief capabilities of the vent valves using pressure sensing lines from the containers to the relief section of the vent valve if valve control pressure is not available.
A high pressure switch located on the storage spheres monitors sphere pressure and is a part of the engine ignition interlock. The switch must indicate a minimum pressure of 2,835 psig or the countdown is stopped.
At liftoff, the electrical circuit to the solenoid valves and switches is disconnected and uninterrupted pressure from the spheres flows through the open solenoid valves and sonic nozzle into the containers. Engine fuel consumption is such that the system cannot be overpressurized and the pressure will fluctuate between a 17 psig upper limit and an 11 psig lower limit until stage burnout.
LOX System
The LOX system receives LOX from a ground source, stores the LOX in containers, and then supplies LOX to the eight H-1 engines. Major components of the system are the four outer LOX containers 0-1, 0-2, 0-3, and 0-4, the center LOX container 0-C, control valves, interconnect lines, suction lines, piping, switches, manifolds, sensors, vent and relief valves, and the GOX flow regulator. These components provide the means of filling or draining the containers, replenishing LOX, bubbling LOX, pressurizing the containers, and supplying LOX to the H-1 engines.
Equal pressurization of the containers and uniform distribution of LOX to the engines is maintained through interconnect lines at the top and bottom of the containers. In the event of an engine failure, the LOX normally consumed by the inoperative engine is supplied to the operating engines.
The center container sump is connected to the sumps of the outer containers. Each outer container supplies LOX to one outboard and one inboard engine through suction lines connected to the container sump. When LOX falls to the level of the engine cutoff LOX sensors located in the bottom of containers 0-2 and 0-4, a signal is generated that initiates the inboard engine shutdown sequence, provided that similar sensors in the fuel system have not already initiated the sequence. Outboard engine shutdown occurs approximately six seconds after inboard engine shutdown. The outboard engines are normally shut down when engine thrust decay, resulting from LOX depletion, causes the outboard engine thrust OK pressure switches to deactuate.
The outboard engines are shut down simultaneously since the "thrust O.K." pressure switches on all outboard engines are electrically interconnected at this time. However, if fuel depletion occurs prior to LOX depletion, the outboard engines will be shut down when fuel reaches the level of two depletion sensors located in the sumps of fuel containers F-2 and F-4. Fifteen discrete level sensors and one continuous level sensor in each LOX container permit telemetric monitoring of the LOX level.
LOX FILL
Before filling the tanks with LOX, the normally closed vent valve and two vent and relief valves are opened. Gaseous nitrogen control pressure from the stage control pressure system opens the three valves.
After the vent valves are opened, the normally closed fill and drain valve is opened by control pressure supplied by the ground control station. LOX is then pumped through the fill and drain nozzle and the open fill and drain valve into the sump of container 0-3. The sumps of the containers are interconnected to ensure equal distribution of LOX. The LOX containers are initially filled to a level based on nominal fuel density.
The flow into the containers is controlled by a ground station LOX tanking computer. When the containers are filled to the predetermined level, the computer shuts off the LOX supply. If the computer should malfunction and not shut off the LOX supply, an overfill sensor will generate a signal to stop the fill sequence.
LOX REPLENISHING
LOX in the containers must be replenished until containers are pressurized to compensate for boiloff losses and changes in fuel density. Corrections are applied to the LOX tanking computer and the containers are replenished from the ground LOX storage tank. A normally closed replenishing valve is opened by control pressure applied from the launch facility. The LOX flows into the sump of container O-4. Differential pressure is sensed by the top and bottom sensing lines in container 0-C and routed to the tanking computer through the LOX sensing line quick- disconnect couplings.
LOX DRAIN
In the event that LOX must be drained from the system, the vent valves and vent and relief valves are closed, the containers are pressurized, and the fill and drain valve is opened. LOX under pressure then flows from the containers through the open valve and fill and drain nozzle to the ground storage tank.
SUCTION AND INTERCONNECT LINES
Eight suction lines conduct LOX from the outer container sumps to the engine turbopumps. Normally open LOX prevalves are located near the top of each suction line and may be closed by pneumatic pressure from the stage control pressure system or the launch facility should the need arise prior to launch. The prevalves also provide a backup to the main LOX valve in the H-1 engines for LOX shutoff.
The LOX containers are interconnected at the top by lines between the LOX pressurization and vent manifold and each outer container. Interconnect lines connect the sump of the center container to the sumps of each outer container.
LOX BUBBLING
Pressurized helium is applied the inlet of each engine LOX pump and bubbles up through the LOX suction lines into the container. The helium rising through the LOX stabilizes the temperature at the pump inlet to prevent pump cavitation. LOX bubbling is initiated prior to fuel container pressurization and continues until LOX container pressurization.
LOX CONTAINER PRESSURIZATION
The LOX containers are pressurized with helium during preflight operations to provide a pressure head at the inlet of the engine LOX pumps. During flight, container pressure also provides structural integrity to the containers. GOX converted from LOX by the heat exchanger on the H-1 engines is used as the pressurant for inflight operation.
To pressurize the LOX containers before engine ignition, the vent and relief valves are closed. Helium then flows from the launch facility through the LOX container pressurization quick-disconnect coupling, a check valve, and the pressurant diffuser into container 0-C. The helium is distributed from container 0-C to the other containers through the interconnecting lines to equally pressurize all containers. The vent and relief pressure switch monitors container pressure and causes the launch facility helium supply to shut off when the pressure reaches 60.5 psia. The pressure switch also energizes a solenoid valve which permits control pressure to open the vent valve and vent and relief valve No.. 1. If the switch should fail, an emergency vent pressure switch will actuate at 67.6 psia and energize a solenoid valve which will permit pneumatic control pressure to open vent and relief valve No. 2.
At liftoff, the helium from the launch facility is disconnected and GOX from the heat exchangers is used as the pressurant. The GOX output from all eight heat exchangers flows into a manifold, then through a GOX flow regulator, the pressurant diffuser in container 0-C, and through interconnecting lines to the outer containers. The GOX flow regulator maintains LOX container pressure at approximately 50 psia. The regulator is controlled by a feedback pressure sensing line from container 0-C.
Pressure relief after liftoff plus 30 seconds is provided through the emergency vent pressure switch. The switch signals a solenoid valve to open vent and relief valve No. 2 if an overpressure condition develops. The vent and relief valves have a mechanical relief capability independent of pneumatic control, which is set to open at 75 psig. Whether the operation will be pneumatic or mechanical is determined by the altitude at which the containers become overpressurized.
Control Pressure System
Major components of the system are a storage sphere, a pressure regulator, a manifold, a relief valve, switches, solenoid-operated valves, filters, and monitoring devices.
The control pressure system receives pressurized GN2 from the launch facility, stores it, and then supplies it at a reduced and regulated pressure to operate pneumatic valves in the fuel and LOX systems, open LOX vent valves, purge calorimeters, pressurize the engine gearboxes, and purge the LOX seal area of the engine turbopumps.
The high-pressure storage sphere is pressurized from the launch facility through the sphere pressurization quick- disconnect coupling, a filter, a check valve, and the vent valve and- manifold assembly. The 1.0-cubic foot storage sphere is charged to 3,000 psig; a pressure switch and pressure transducer monitor sphere pressure. When necessary, the system is vented through the vent valve and manifold assembly.
High-pressure GN2 from the storage sphere through the vent valve and manifold assembly and the filter to a pressure regulator. The pressure regulator reduces and regulates the pressure to 750 psig and supplies the gas to a control pressure manifold. Pressure in the manifold is monitored by a pressure switch and pressure transducer. The manifold is protected from overpressurization by a relief valve.
Control pressure GN2 is routed from the manifold through two solenoid valves for operation of the LOX container vent valve and the vent and relief valves. Control pressure GN2 is also routed to eight solenoid valves to close the LOX and fuel prevalves during static test and at stage burnout.
Control pressure GN2 is also used to pressurize the engine gearboxes and purge the LOX seals on the turbopumps; a manual valve provides shutoff control. Nitrogen is also used to purge the windows of four calorimeters so that exhaust products do not accumulate on the windows and cause the calorimeters to give erroneous indications.
ENGINE PURGE AND GEARBOX PRESSURIZATION SYSTEM
Several GN2 purges are started during launch preparations to prevent contamination of the engine.
LOX DOME PURGE
The dome purge utilizes GN2 from a common source.
The gas is supplied by the vehicle control pressure system; the supply is routed through a ring-line manifold with branch lines to each engine gearbox.
The LOX pump seal purge and gearbox pressurization starts when the S-IB stage control pressure system is pressurized and continues throughout launch preparations, engine starting, and S-IB flight. If the launch is cancelled, purging continues until all LOX within the engine LOX pump has boiled off.
Low-pressure GN2 purge is applied to the space between the LOX and lube seals in the turbopump to isolate LOX leakage from the lubricant leakage. LOX and lubricant leakage into the area between the seals is drained overboard through separate drains.
Gearbox pressurization is accomplished with GN2 also reduced in pressure. The check valve and the relief valve operate together to maintain a pressure within the gearbox to prevent the lubricant from foaming at high altitudes. The relief valve also functions to bleed spent lubricant (RP-1 fuel and Oronite 262 additive) to the atmosphere.
The pressure within the gearbox also forces any fuel that leaks past the fuel seals and into the gear box to drain overboard through the gearbox lube drain. The leakage can be visually monitored prior to engine ignition.
The LOX dome of each H-1 engine is purged by both a low-level purge and a full-flow purge. The low-level purge maintains a slight positive GN2 pressure in the LOX dome to prevent contaminants from entering the thrust chamber nozzle and flowing to the injector plate and the LOX dome. It also prevents moisture from condensing in the area. The low-level purge is started prior to propellant loading and continues until shortly before engine ignition; GN2 pressure and flow rate are then increased to the full-flow level. The full-flow purge continues until LOX pressure in the LOX dome is greater than purge pressure. If a launch is cancelled, the full-flow purge resumes as LOX pressure decays below purge pressure. The GN2 purge then expels LOX from the LOX dome and the LOX bootstrap line. After a short interval, the full-flow purge rate is reduced to the low-level rate.
Ground source GN2 for both purge levels flows through the LOX dome purge check valve, through a manifold on the main LOX valve, and then into the LOX dome. The GN2 is vented through the thrust chamber nozzle.
THRUST CHAMBER FUEL INJECTOR MANIFOLD PURGE
The thrust chamber fuel injector manifold purge prevents LOX from entering the fuel injector manifold during engine ignition. The purge is started just before engine ignition. Ground source GN2 flows through a ring-line manifold around the engine compartment and then into a purge manifold of each of the engines. Each purge manifold distributes the GN2 into three thrust chamber fuel injector manifold purge check valves. The GN2 flows through the thrust chamber fuel injector manifold, through the injector plate, and out the thrust chamber nozzle. After the engine starting sequence begins, the purge is stopped when fuel pressure builds up in the fuel injector manifold and closes the three check valves.
LIQUID PROPELLANT GAS GENERATOR LOX INJECTOR MANIFOLD PURGE
The liquid propellant gas generator LOX injector manifold purge prevents the combustion products from the solid propellant gas generator from contaminating the liquid propellant gas generator LOX injector manifold. The purge is started just before engine ignition and is stopped by LOX pressure buildup in the manifold. If a launch is cancelled, the purge commences immediately following engine shutdown and continues until the solid propellant gas generator has been removed.
Ground source GN2 flows through branch lines leading to each engine, and through a check valve into the LOX injector manifold. The GN2 purge then flows through the liquid propellant gas generator, the gas turbine, and the heat exchanger. The GN2 purge is exhausted through the aspirator on outboard engines and through the turbine exhaust duct on inboard engines.
Hydraulic System
Each outboard engine has an independent, closed-loop hydraulic system that gimbals the engine for vehicle flight control, moving the engine in proportion to the magnitude of an electrical input signal. Movement is provided by two hydraulic actuators that may be extended or retracted independently or simultaneously.
Major components of each system are a main hydraulic pump, an auxiliary hydraulic pump and motor, an accumulator-reservoir and manifold assembly, and the hydraulic actuators.
The accumulator is filled with GN2 from a ground source through the GN2 charging valve. Hydraulic fluid pumped from the ground source flows through the high-pressure quick-disconnect coupling and the main filter into the reservoir and the system. Fluid level in the reservoir is indicated by a potentiometer. Any excess hydraulic fluid returns to the ground source through the low-pressure quick-disconnect coupling during servicing.
An electric auxiliary pump is used for ground checkout of the hydraulic system operation. A check valve protects the main pump from high pressure fluid during auxiliary pump operation. After engine ignition, a check valve protects the auxiliary pump from high pressure during main pump operation.
The main pump is driven by the turbopump after engine ignition. Hydraulic fluid is drawn from the reservoir portion of the accumulator-reservoir and manifold assembly, and pumped through a check valve and the main filter into the accumulator portion of the accumulator-reservoir and manifold assembly. Pressure is increased to approximately 3,200 psig on the high pressure side of the main pump. Fluid flows from the accumulator-reservoir and manifold assembly to the hydraulic actuators.
An electro-hydraulic servovalve in each actuator directs high-pressure fluid against the required side of the actuator piston, when activated by an electrical signal. The piston is connected to the actuator arms, which extend or retract, and provide gimbaling action to the engine. Fluid flows from the hydraulic actuators hack to the reservoir.
Instrument Compartment Environmental Conditioning System
The instrument compartments located in the top skirt of fuel containers F-1 and F-2 are environmentally conditioned prior to flight. Conditioned air or GNU supplied from the launch facility environmental conditioning system is ducted into the instrument compartments to dissipate heat generated by the electronic instruments, and reduce the possibility of fire during launch.
Compartment conditioning is initiated when the electronic equipment is turned on at approximately T minus 19.5 hours. Air is used until LH2 tanking is initiated in the S-IVB stage; GN2 is then used for the remainder of the conditioning period until liftoff. At liftoff, the precooling check valves close and a bleed orifice prevents an increase in compartment pressure during flight.
Tail Unit Conditioning and Water Quench System
The tail Unit conditioning and water quench system is provided to direct conditioned air, conditioned GN2 or water into the engine compartment. The conditioned air or GN2 is supplied by the launch complex to provide a temperature controlled atmosphere in the tail unit area. In the event of fire in the engine compartment, or any malfunction which causes engine shutdown, a cold GN2 deluge and, if necessary, a water quench is supplied to the tail unit area. All mediums flowing through the system use the same vehicle network of valves, pipes, and dispersal manifolds.
At liftoff, the system is separated from the ground source; therefore, tail unit conditioning and fire protection are not provided during flight.
TAIL UNIT FIRE DETECTION SYSTEM
A fire detection system is provided to detect any fire that may occur in the engine compartment prior to liftoff. The fire detection system consists of 32 thermocouple sensors located on the aft thrust structure and the firewall substructure. The sensors are connected in four loops of eight sensors. The eight sensors in each loop are positioned at a similar location near each engine; thus, the temperature rise-rate in four critical areas of each engine is monitored until liftoff.
If a sensor detects a fire prior to engine ignition, a signal is transmitted to the launch control center to initiate a cold GN2 deluge of the engine compartment. If a fire is detected during the time between engine ignition and launch commit, the signal initiates the engine cutoff sequence and the cold GN2 deluge.
If the cold GN2 deluge does not extinguish the fire, a water quench operation is manually initiated from the launch control center. The GN2 deluge uses the same dispersal manifold network as the tail unit conditioning and water quench system.
Range Safety System
Each vehicle launched from the ETR must have two separate and independent methods of emergency flight termination for use if the vehicle should become a safety hazard during powered flight. Air Force Missile Test Center safety regulations require missiles using liquid propellants to have systems that accomplish zero thrust and propellant dispersion.
The ETR Safety Officer decides whether or not a safety hazard exists and bases his decision on (1) the imminence of an explosion, and (2) deviation of the vehicle from the programmed flight path.
The S-IB stage range safety system consists of four receiver antennas, two command receivers, two command controllers, two electronic bridgewire firing units, two EBW detonators, and a safe-and-arm device. The safe-and-arm device connects the system to the shroud fuze assembly and nine linear-shaped charges. Reliability is further increased by supplying power to each of the dual sets from separate batteries.
A ground-based command transmitter sends the frequency-modulated audio-tone coded signals that initiate engine cutoff, charge the capacitor in the electronic bridgewire firing unit, and trigger the propellant dispersion system. Four receiver antennas mounted in sets of two on opposite sides of the stage, receive the signals. The signals are demodulated within the command receivers. The resulting audio tones are applied to the decoder, which separates the tones according to frequency, and energizes relays that generate the engine cutoff signal, the arming command, and the destruct command.
The command controller directs signals from the command receiver to the proper units and relays supervisory signals to the ground station. Electrical connectors are provided for attaching a nodestruct-delay plug or a delay timer to the controller. The no-destruct-delay plug completes the electrical circuit immediately when no time delay is desired. A delay timer will delay the destruct command to allow a payload to be jettisoned before the vehicle is destroyed. The destruct command, transmitted approximately 3 seconds after the arming command, is coupled to the electronic bridgewire firing unit to fire the EBW detonator.
The EBW detonator is coupled to the shroud fuze assembly through the safe-and-arm device. The shroud fuze assembly interconnects the nine shaped charges on the nine propellant tanks and connects to the fuze system in the S-IVB stage.
When the system is actuated, the linear-shaped charges fire and the propellant tanks are split open, resulting in complete propellant dispersion and destruction of the vehicle.
The safe-and-arm device provides safety to personnel during the installation of the EBW detonators and the fuse assembly. The device is a rotary mechanical block that, when in the safe position, blocks the flame path between the EBW detonators and the shroud fuse assembly. Prior to liftoff, the safe-and-arm device can be armed or disarmed electrically or manually. Prior to liftoff, the device is actuated to the armed position.
Electrical System
Two independent 28-volt bus systems distribute the power throughout the stage. Generally, constant loads are supplied from one bus and variable loads are supplied from the other bus. However, exceptions to this general rule are frequently made to form redundant circuits in the interest of reliability.
Prior to launch, primary power is supplied by the launch complex through umbilical connections. After launch, two wet-cell storage batteries located in instrument compartment No. 2 furnish 38 vdc. Actual power transfer occurs approximately 25 seconds before engine ignition. Two squib-actuated switch assemblies are actuated at liftoff to complete power transfer circuits which parallel the normal transfer circuits. This ensures that the stage is switched from ground power to battery power, compensates for possible contractor bounce, and makes the transfer non- reversible and virtually failure-proof.
The 28 vdc power is switched and distributed to the inflight subsystems through the power distributor, also located in instrument compartment No. 2.
Many of the transducers in the measurement system require 5 vdc for excitation. Three master measuring voltage supplies are installed on the stage to convert 28 vdc to precisely regulated 5 vdc.
CONTROL COMPONENTS
A switch selector is installed on the S-IB stage to connect and disconnect the various flight sequence commands during S-IB stage powered flight.
The switch selector installed in instrument compartment No. 2 controls time-referenced events in the S-IB stage. It receives digital-coded commands from the data adapter in the Instrument Unit, decodes the commands, and transmits them to the main distributor.
Two control accelerometers are attached to the spider beam. The accelerometers sense the lateral acceleration of the vehicle, perpendicular to the vehicle longitudinal axis in the pitch and yaw planes. The output of the accelerometers is used in the vehicle control system to reduce structural loading and control vehicle deflections due to lateral winds.
Flight Measurement Program
The requirements for preflight and inflight performance measurements of the Saturn vehicles differ substantially from those of conventionally guided missile systems. A large number of measurements must be obtained to meet the stringent demands of the Saturn research and development program. The various physical events and environmental conditions which prevail throughout the vehicle before and during flight must be made available to ground stations in a precise, real time format.
Almost 500 measurements will be made and telemetered during the first flight of the S-IB stage. Before launch, approximately 175 measurements will be transmitted to the blockhouse by hardwire connections.
Many of the source signals are not suitable for direct transmission by the telemetry system; therefore, signal conditioning devices are required to modify the signals. The conditioning devices are replaceable modules installed in 12 measuring racks in the tail unit area and in 6 measuring racks in instrument compartment No. 2.
Measuring distributors are junction boxes which connect the measurement signals to the telemetry systems and provide points for checkout, maintenance, and modification of the systems. Three distributors are located in the tail unit area and one is located in instrument compartment No. 2.
TELEMETRY SYSTEMS
Stage performance measuring signals, when grouped according to frequency and accuracy requirements, can be most effectively transmitted by using several types of telemeters. Four telemeter systems are required to transmit the S-IB stage measuring signals. Most of the components of the telemetry systems are located in instrument compartment No. 1; two multiplexers, however, are installed in the aft skirt of two of the fuel containers. The telemeter transmits data through a common antenna system.
Telemeters F1 and F2
Telemeters F1 and F2 are identical systems which transmit narrow band, frequency-type data such as that generated by strain gages, temperature gages, and pressure gages. The system can handle 234 measurements on a time-sharing basis and 14 measurements transmitted continuously. Data may be sampled 120 times per second, or 12 times per second.
Telemeter S1
Telemeter S1 transmits wide hand frequency-type data generated by vibration sensors. The S-IB stage measuring program requires 25 data sources transmitted on a 25 per cent duty cycle, 8 data sources transmitted on a 50 per cent duty cycle, and 4 data sources transmitted continuously.
Telemeter P1
Telemeter P1 transmits pulse code modulated, or "hang-hang" type data. This type of data is generated by limit switches, pressure-actuated switches, valves, and relays. Three multiplexers supply data to the telemeter; two handling 100 hits of digital data each while the third has a capacity of 234 data inputs.
Telemetering Calibrator
A telemetering calibrator installed in instrument compartment No. 1 improves the accuracy of the telemetry systems. The calibrator supplies known voltages to the telemeters periodically during the S-IB stage operation. Their reception at tracking stations provides a valid reference for data reduction.
Tape Recorder
The effects of retrorocket firing attenuation can seriously degrade the telemetry transmission during stage separation; therefore, a tape recorder installed in instrument compartment No. 1 records data for delayed transmission. The commands for tape recorder operation originate in the flight programmer located in the Instrument Unit.
Tracking System
The S-IB stage carries a transponder to facilitate ground tracking. The transponder, installed in instrument compartment No. 2, is part of the ODOP tracking system. The ODOP is an elliptical tracking system that measures the sum of the ranges between the stage and three ground stations. The range sum is determined by measuring the total Doppler shift in frequency of a continuous wave radio frequency. Since the transponder is phase-coherent, the Doppler shift is determined primarily by the range and velocity of the stage.
Recoverable Camera System
Two motion picture cameras installed in the stage provide a permanent visual record of the S-IB/SIVB separation sequence, operation of the S-IVB stage ullage roekets, and ignition of the S-IVB stage engine. The cameras are housed in separate capsules that are ejected from the spent S-IB stage and recovered from the ocean. Major components of the system are the two cameras, two camera ejection cylinders, a high pressure storage sphere, and two solenoid valves.
Each camera capsule is held in an ejection tube attached to a radial member of the spider beam. One capsule is at fin position 2 and the other at fin position 6. The ejection tubes protrude through the spider beam seal plates and are canted at an angle towards the stage centerline to provide collision-free ejection paths for the capsules.
The capsule consists of a waterproof aluminum forward housing and an expendable fiber-composition aft housing. Eight stainless steel flaps attached to the aft housing extend during capsule descent. In addition to the camera, the capsule contains a power pack, a compressed gas storage bottle, shark repellent, a paraballoon, dye marker, a radio beacon, and a parachute.
The cameras are started approximately 5 seconds before S-IB/S-IVB stage separation by a signal from the switch selector and operate until they are ejected 26 seconds after stage separation. The stage separation signal activates a 26-second timer that generates the ejection command at the expiration of the time interval. Premature ejection of the capsule could damage the S-IVB stage. An additional timer, operated in parallel with the 26 second timer, keeps the ejection circuit deactivated for 8 seconds after stage separation in the event that the 26-second timer should operate too soon.
After the capsule leaves the ejection tube, spring-loaded flaps extend to stabilize and slow the descent. The capsules re-enter the atmosphere at approximately Mach 10 (7,900 mph). At an altitude of approximately 14,000 feet, a barometrically actuated valve releases compressed GN2 into an 18-inch paraballoon. The force of the expanding balloon causes the two capsule housings to separate. The aft housing then falls away and the parachute is released. The combined drag of the parachute and the paraballoon reduces the capsule impact velocity to prevent capsule damage.
The paraballoon prevents the capsule from sinking and serves as a location aid. Alternate panels of the balloon are painted Day-Glo International Orange for visual detection in sunlight and a white glass-beaded reflective surface to aid detection by searchlight at night.
Upon contact with salt water, a yellow-green fluorescent dye is metered out over a period of 4 to 6 hours. At the same time, a plug of shark repellent is dissolved in the water to ward off fish that could damage the capsule and paraballoon. The action of the repellant lasts 2 to 4 hours, depending on water temperature.
A radio beacon attached to the top of the paraballoon transmits a signal for approximately 20 hours to aid recovery teams.
Copyright 1997-2005 by John
Duncan |
