INTRODUCTIONThe S-IC stage is a cylindrical booster, 138 feet long and 33 feet in diameter, powered by five liquid propellant F-l rocket engines. These engines develop a nominal sea level thrust of 7,650,000 pounds total, and have a burn time of 167.3 seconds. The stage dry weight is approximately 293,500 pounds and total weight at ground ignition is approximately 5,030,500 pounds.
The S-IC stage provides first stage boost of the Satum V launch vehicle to an altitude of about 200,000 feet (approximately 38 miles), and provides acceleration to increase the vehicle velocity to 7,700 feet per second (approximately 4,560 knots). it then separates from the S-II stage and falls to earth about 360 nautical miles downrange.
The stage interfaces structurally and electrically with the S-II stage. It also interfaces structurally, electrically, and pneumatically with two umbilical service arms, three tail service masts, and certain electronic systems by antennae. The major systems of the stage are: structures, environmental control, propulsion, flight control, pneumatic controls, propellants, electrical, instrumentation, and ordnance.
STRUCTURE
The S-IC structure design reflects the requirements of F-l engines, propellants, control, instrumentation and interfacing systems. The structure maintains an ultimate factor of safety of at least 1.40 applied to limit load and a yield factor of safety of 1.10 on limit load. Aluminum alloy is the primary structural material. The major are the forward skirt, oxidizer tank, intertank section, fuel tank, and thrust structure.
FORWARD SKIRT
The aft end of the forward skirt is attached to the oxidizer (lox) tank and the forward end interfaces with the S-II stage. The forward skirt has accomodations for the forward umbilical plate, electrical and electronic canisters, and the venting of the lox tank and interstage cavity. The skin panels, fabricated from 7075-T6 aluminum, are stiffened and strengthened by ring frames and stringers.
OXIDIZER TANK
The 345,000 gallon lox tank is the structural link between the forward skirt and the intertank section. The cylindrical tank skin is stiffened by "integrally machined" T stiffeners. Ring baffles attached to the skin stiffeners stabilize the tank wall and sense to reduce lox sloshing. A cruciform baffle at the base of the tank series to reduce both slosh and vortex action. Support for four helium bottles is provided by the ring baffles. The tank is a 2219-T87 aluminum alloy cylinder with ellipsoidal upper and lower bulkheads. The skin thickness is decreased in eight steps from .254 inches at the aft section to .190 inches at the forward section.
INTERTANK SECTION
The intertank structure provides structural continuity between the lox and fuel tanks. This structure provides a lox fill and drain interface to the intertank umbilical. One opening vents the fuel tank. The corrugated skin panels and circumferential ring frames are fabricated from 7075-T6 aluminum.
FUEL TANK
The 216,000 gallon fuel tank (figure 4-1) provides the load carrying structural link between the thrust structure and intertank structure. The tank is cylindrical, with ellipsoidal upper and lower bulkheads. Antislosh ring baffles are located on the inside wall of the tank and antivortex cruciform baffles are located in the lower bulkhead area. Five lox ducts run from the lox tank, through the RP-I tank, and terminate at the F-l engines. The fuel tank has an exclusion riser, made of a lightweight foam material, which is bonded to the lower bulkhead of the tank to minimize unusable residual fuel. The 2219-T87 aluminum skin thickness is decreased in four steps from .193 inches at the aft section to .170 inches at the forward section.
THRUST STRUCTURE
The thrust structure assembly redistributes locally applied loads of the five F-l engines into uniform loading about the periphery of the fuel tank. It also provides support for the 5 F-l engines, engine accessories, base heat shield, engine fairings and fins, propellant lines, retrorockets, and environmental control ducts. The lower thrust ring has four holddown points which support the fully loaded Satum/Apollo (over 6,000,000 pounds) and also, as necessary, restrain the vehicle from lifting off at full F-l engine thrust. The skin segments are fabricated from 7075-T6 aluminum alloy.
The base heat shield is located at the base of the S-IC stage, forward of the engine gimbal plane. The heat shield provides thermal shielding for critical engine components and base region structural components for the duration of the flight. The heat shield panels are constructed of 15-7 PH stainless steel honeycomb, 1.00-inch thick, brazed to .010 inch steel face sheets.
Each outboard F-1 engine is protected from aerodynamic loading by a conically shaped engine fairing. The fairings also house the retrorockets and the engine actuator supports. The fairing components are primarily titanium alloy below station 115.5 and aluminum alloy above this station. Four fixed, titanium covered, stabilizing fins augment the stability of the Saturn V vehicle.
ENVIRONMENTAL CONTROL SYSTEMS
During launch preparations the environmental control systems (ECS) protect the S-IC stage and stage equipment from temperature extremes, excessive humidity, and hazardous gases. Conditioned air, from the ground support equipment environmental control unit (GSE-ECU), is forced into the forward skin and thrust structure where it is used as a temperature and humidity control medium. Approximately 20 minutes before the two upper stages are loaded with cryogenic fluids gaseous nitrogen (GN2) replaces conditioned air and is introduced into the S-IC as the conditioning mediuum. The GN2 flow terminates at umbilical disconnect since the system is not needed in flight.
FORWARD SKIRT COMPARTMENT - ECS
The environmental control system distributes air or GN2 to 18 electrical/electronic equipment module canisters located in the forward skirt. Onboard probes control the temperature of the flow medium to maintain canister temperature at 80 (+/-20) degrees F. Three phases of the conditioning/purge medium flow are provided to compensate for the environmental imbalances generated by ambient air changes, internal heat and lox load chill effects. The frost phase supplies cool, conditioned air to the canisters when onboard electrical systems are energized before cryogenic loading. The second phase occurs when relatively warm GN2 is substituted for the cool air to offset temperature differences caused by the cryogenic loading. The third phase uses a warmer GN2 flow to offset temperature decreases caused by second stage J-2 engine thrust chamber chilldown. The air or GN2 is vented from the canisters and overboard through vent openings in the forward skirt of the S-IC stage. A by-product of the use of the inert GN2 is the reduction of gaseous hydrogen or oxygen concentrations.
THRUST STRUCTURE COMPARTMENT - ECS
The envirommental control system discharges air or GN2 through 22 orificed duct outlets directly into the upper thrust structure compartment. The GSE-ECU supplies conditioned air at two umbilical couplings during launch preparations. At 20 minutes before cryogenic loading commences, the flow medium is switched to GN2 and the temperature varied as necessary to maintain the compartment temperature at 80 (+/-10) degrees F. The temperature control compensates for temperature variations caused by ambient air change, and chill effects from lox in the suction ducts, prevalves, and inter-connect ducts. The GN2 prevents the oxygen concentration in the compartment from exceeding 6 percent.
Hazardous Gas Detection- Forward and Aft
HAZARDOUS GAS DETECTION
The hazardous gas detection system monitors the atmosphere in the forward skirt and the thrust structure compartment of the S-IC. This system is not redundant; however, large leaks may be detected by propellant pressure indications displayed in the Launch Control Center.
PROPULSION
The F-1 engine is a single start, 1,530,000 pound fixed thrust, calibrated, bipropellant engine which uses liquid oxygen as the oxidizer and RP-I as the fuel. Engine features include a bell shaped thrust chamber with a 10:1 expansion ratio, and detachable, conical nozzle extension which increases the thrust chamber expansion ratio to 16:1. The thrust chamber is cooled regeneratively by fuel, and the nozzle extension is cooled by gas generator exhaust gases. Liquid oxygen and RP-I fuel are supplied to the thrust chamber by a single turbopump powered by a gas generator which uses the same propellant combination. RP-I fuel is also used as the turbopump lubricant and as the working fluid for the engine fluid power system. The four outboard engines are capable of gimbaling and have provisions for supply and return of RP-I fuel as the working fluid for a thrust vector control system. The engine contains a heat exchanger system to condition engine supplied liquid oxygen and externally supplied helium for stage propellant tank pressurization. An instrumentation system monitors engine performance and operation. External thermal insulation provides an allowable engine environment during flight operation.
ENGINE OPERATING REQUIREMENTS
The engine requires a source of pneumatic pressure, electrical power, and propellants for sustained operation. A ground hydraulic pressure source, an inert thrust chamber prefill solution, gas generator igniters, gas generator exhaust igniters, and hypergolic fluid are required during the engine start sequence. The engine is started by ground support equipment (GSE) and is capable of only one start before reservicing.
PURGE, PREFILL, AND THERMAL CONDITIONING
A gaseous nitrogen purge is applied for thermal conditioning and elimination of explosive hazard under each engine cocoon. Because of the possibility of low temperatures existing in the space between the engine and its cocoon of thermal insulation, heated nitrogen is applied to this area. This purge is manually operated, at the discretion of launch operations, whenever there is a prolonged hold of the countdown with lox onboard and with an ambient temperature below approximately 55 degrees F. In any case, the purge will be turned on five minutes prior to ignition command and continue until umbilical disconnect.
A continuous nitrogen purge is required to expel propellant leakage from the turbopump lox seal housing and the gas generator lox injector. The purge pressure also improves the sealing characteristic of the lox seal. The purge is required from the time propellants are loaded and is continuous throughout flight.
A nitrogen purge prevents contaminants from accumulating on the radiation calorimeter viewing surfaces. The purge is started at T-52 seconds and is continued during flight.
A gaseous nitrogen purge is required to prevent contaminants from entering the lox system through the engine lox injector or the gas generator lox injector. The purge system is actuated prior to engine operation and is continued until umbilical disconnect.
From T-15 hours to T-13 hours, an ethylene glycol solution fills the thrust chamber tubes and manifolds of all five engines. This inert solution serves to smooth out the combustion sequence at engine start. Flow is terminated by a signal from an observer at the engines. At approximated T-10 minutes, 50 gallons are supplied to top off the system to compensate for liquid loss that occurred during engine gimbaling.
POGO SUPPRESSION SYSTEM
The POGO supression system utilizes the lox prevalve cavities of the four outboard engines as surge chambers to suppress the POGO phenomenon. The lox prevalve cavities are pressurized with gaseous helium (GHe) at T-11 minutes from ground supply by opening the POGO suppression control valves. During the initial fill period (T-11 to T- 9 minutes), the filing of the valves is closely monitored, utilizing measurements supplied by the liquid level resistance thermometers R3 (primary) and R2 (backup). The GHe ground fill continues to maintain the cavity pressure until umbilical discomnect. Following umbilical disconnect the cavity pressure is maintained by the cold helium spheres located in the lox tank.
Status on system operation is monitored through two pressure transducers and four liquid level resistance thermometers. The pressure transducer (0-800 psia) monitors system input pressure. A second pressure transducer (0-150 psia) monitors the pressure inside the No. 1 engine lox prevalve cavity. These pressure readings are transmitted via telemetry to ground momtors. The liquid level within the prevalves is monitored by four liquid level resistance thermometers in each prevalve. These thermometers transmit a "wet" (colder than -165 degrees centigrade) and a "dry" (warmer than -165 degrees centigrade) reading to ground monitors.
ENGINE SUBSYSTEMS
The subsystems of the F-1 engine are the turbopump, checkout valve, hypergol manifold, heat exchanger, main fuel valve and main lox valve. Subsystems not shown are the gas generator, 4-way control valve, and pyrotechnic igniters.
Hypergol Manifold
The hypergol manifold consists of a hypergol container, an ignition monitor valve (IMV), and an igniter fuel valve (IFV). The hypergol solution is forced into the thrust chamber by the fuel where combustion is initiated upon mixing with the lox. The IFV prevents thrust chamber ignition until the turbopump pressure has reached 375 psi. The IMV prevents opening of the main fuel valves prior to hypergolic ignition. A positive hypergol cartridge installed indication is provided by sensors and is a prerequisite to the firing command.
Control Valve - 4-Way
The 4-way control valve directs hydraulic fluid to open and close the fuel, lox, and gas generator valves. It consists of a filter manifold, a start and stop solenoid valve, and two check valves.
Turbopump The turbopump is a combined lox and fuel pump drive, through a common shaft by a single gas turbine.
Gas Generator
The gas generator (GG) provides the gases for driving the turbopump. Its power output is controlled by orifices in its propellant feed Lines. The gas generator system consists of a dual ball valve, an injector, and a combustor. Combustion is initiated by two pyrotechnic igniters. Total propellant flow rate is approximately 170 Ib/sec at a lox/RP- 1 mixture ratio of 0.42:1. The dual ball valve must be closed prior to fuel loading and must remain closed to meet an interlock requirement for engine start.
Heat Exchanger
The heat exchanger expands lox and cold helium for propellant tank pressurization. The cold fluids, flowing through separate heating cods, are heated by the turbopump exhaust. The warm expanded gases are then routed from the heating coils to the propellant tanks.
Main Fuel Valve
There are two main fuel valves per engine. They control flow of fuel to the thrust chamber. The main fuel valve is a fast acting, pressure balanced, poppet type, hydraulically operated valve. Movement of the poppet actuates a switch which furnishes valve position signals to the telemetry system. This valve is designed to remain open, at rated engine pressures and flowrates, if the opening control pressure is lost. Both valves must be in the closed position prior to fuel loading or engine start.
Main Lox Valve
The two main lox valves on each engine control flow to the thrust chamber. These valves are fast acting, pressure balanced, poppet type, hydraulically operated valves. A sequence valve operated by the poppet allows opening pressure to be applied to the GG valve only after both main lox valve poppets have moved to a partially open position. This valve is designed to remain open, at rated engine pressures and flowrates, if the opening control pressure is lost. Both main lox valves must be in the closed position prior to lox loading or engine start.
Checkout Valve
The checkout valve directs ground supplies control fluid from the engine to ground during engine checkout. Approximately 30 seconds prior to the firing command the valve is actuated to the engine position. In this position it directs control fluid to the No. 2 turbopump inlet. An ENGINE POSITION indication is required from this valve prior to, and is interlocked with, forward umbilical disconnect command.
High Voltage Igniters
Four high voltage igniters, two in the gas generator (GG) body and two in the engine thrust chamber nozzle extension ignite the GG and the fuel rich turbopump exhaust gases. They are ignited during the F-l engine start sequence by application of a nominal 500 volts to the igniter squibs.
ENGINE OPERATION
Engine operation includes starting, mainstage and cutoff. The starting and cutoff phases are periods of transition in which a sequence of activities occurs. These phases are developed in detail in the following paragraphs.
Engine Start
The engine start and transition to mainstage is illustrated in Figure 4-6 for a typical single engine.
Engine Cutoff
The normal inflight cutoff sequence is center engine first, followed by the four outboard engines. At approximately 2 minutes, 15 seconds, the center engine is programmed by the LVDC for cutoff. This command also initiates time base No. 2 (T2 + 0.0). The LVDC provides a backup center engine shutoff signal. Outboard engine cutoff is enabled by a signal from the LVDC. Cutoff is initiated upon energization of two out of four optical type depletion sensors in either the lox or the fuel tank. (Lox depletion is most probable). The sensors start a timer which, upon expiration, energizes the 4-way control valve stop solenoid on each outboard engine. Time base No. 3 (T3 + 0.0) is initiated at this point. The remaining shutdown sequence of the outboard engines is the same as for the center engines.
Emergency Engine Cutoff
In an emergency, the engine be cut off by any of the following methods: Ground Support Equipment (GSE) Command Cutoff, Range Safety Command Cutoff. Thrust Not OK Cutoff, Emergency Detection System, Outboard Cutoff System.
GSE has the capability of initiating engine cutoff anytime until umbilical disconnect. Separate command lines are supplied through the aft umbilicals to the engine cutoff relays and prevalve close relays.
Range Safety Cutoff has the capability of engine cutoff anytime after liftoff. If it is determined during flight that the vehicle has gone outside the established corridor, the Range Safety Officer will send commands to effect engine cutoff and propellant dispersion.
Three thrust OK pressure switches are located on each F-1 engine thrust chamber fuel manifold and sense main fuel injection pressure. If the pressure level drops below the deactivation level of two of the three pressure switches, an engine cutoff signal is initiated. The circuitry is not enabled until T1 + 14.0 seconds to allow the vehicle to clear the launch umbilical tower.
Prior to its own deactivation, the Emergency Detection System (EDS) initiates engine cutoff when it is determined that two or more endnes have shutdown prematurely. When the IU receives signals from the thrust OK logic relays that two or more engines have shutdown, the IU initiates a signal to relays in the S-IC stage to shutdown the remaining engines. S-IC engine EDS cutoff is enabled at T1 + 30.0 seconds and continues until 0.8 second before center end ne cutoff or until deactivated.
Following EDS deactivation, the outboard engine cutoff system is activated 0.1 second before center engine cutoff, and is similar in function to the EDS. Whereas the EDS initiates emergency engine cutoff when any two engines are shutdown, the outboard cutoff system monitors only the outboard engines and provides outboard engine cutoff if the thrust OK pressure switches cause shutdown to two adjacent outboard engines.
NOTE Loss of two adjacent outboard engines, after center engine cutoff, could cause stage breakup.
FLIGHT CONTROL
The S-IC flight control system gimbals the four outboard engines to provide attitude control during the S-IC burn phase. See Section Vll for a detailed discussion of the Saturn V flight control.
FLUID POWER
There are five fluid power systems on the S-IC stage, one for each engine (Figure 4-8 shows a typical outboard engine fluid power system). The hydraulic pressure is supplied from a GSE pressure source during test, prelaunch checkout and engine start. After engine start, hydraulic pressure is generated by the engine turbopump. Pressure from either engine or GSE is made available to engine valves such as the main fuel and lox valves and igniter fuel valve. These valves are sequenced and controlled by the terminal countdown sequencer, stage switch selector and by mechanics or fluid pressure means. A discussion of these components may be found under the titles "PROPULSION" or "ELECTRICAL". Fluid under pressure also flows through a filter and to the two flight control servoactuators on each outboard engine.
The fluid power system uses both RJ-1 ramjet fuel and RP-1 rocket propellant as the hydraulic fluid. The RJ-1 is used by the Hydraulic Supply and Checkout Unit (GSE pressure source). RP-1 is the fuel used in the S-IC stage and as a hydraulic fluid, is pressurized by the engine turbopump. These two hydraulic fluids are separated by check valves and their return flow is directed to GSE or stage by the ground checkout valve. Drilled passages in the hydraulic components (valves and servoactuators) permit a flow of fluid to thermalIy condition the units and to bleed gases from the fluid power system.
HYDRAULIC SERVOACTUATOR
The servoactuator is the power control unit for converting electrical command signals and hydraulic power into mechanical outputs to gimbal the engines on the S-IC stage. The flight control computer (IU) receives inputs from the guidance system in the IU and sends signals to the servoactuators to gimbal the outboard endnes in the direction and magnitude required. An integral mechanical feedback varied by piston position modifies the effect of the IU control signal. A built-in potentiometer senses the servoactuator position and transmits this information to the IU for further transmission via telemetry to the ground.
The servoactuators are mounted 90 degrees apart on each outboard engine and provide for engine gimbaling at a rate of 5 degrees per second and a maximum angle of +/- 5.0 degrees square pattern.
PNEUMATIC CONTROLS
The pneumatic control system provides a pressurized nitrogen supply for command operation of various pneumatic valves. Pneumatic control of the fuel and lox fill and drain valves and the No. 2 lox interconnect valve is provided directly from GSE. Lox interconnect valves No. 1, 3 and 4 are controlled by the onboard pneumatic system.
The pneumatic control system for those valves which must be controlled during flight (fuel and lox prevalves, lox and fuel vent valves) is supplied by a GSE nitrogen source at 3200 psi. The system is charged through onboard control valves and filters, a storage bottle, and a pressure regulator which reduces the supplied pressure to 750 psi. There are direct lines from the GSE to the prevalve solenoid valves which provide pressure for emergency engine shutdown. Orifices to the fuel and lox prevalve lines control closing time of the valves.
PROPELLANTS
Propellants for the S-IC stage are RP-1 (fuel) and liquid oxygen (lox). The propellant system includes hardware for fuel fill and drain operations, tank pressurization prior to and during flight, and delivery of propellants to the engines. The system is divided into two systems, the fuel system and the lox system.
FUEL LOADING AND DELIVERY
Fuel loading starts at approximately 126 hours before liftoff and continues at a rate of 2000 gpm until 99% full and then uses a 200 gpm rate until the total mass load reaches 102% of the desired load. At T-60 minutes the propellant management GSE gives the command to begin fuel level adjustment to the prescribed flight load level. This initiates a limited drain. The fuel loading probe senses the mass level. The fuel vent and relief valve is opened during gravity drain but must be closed for pressurized drain.
RP-1 Pressurization
Fuel tank pressurization is required from engine starting through stage flight to establish and maintain a net positive suction head at the fuel inlet to the engine turbopumps. Ground supplied helium for prepressurization is introduced into the cold helium line downstream from the flow controller resulting in helium flow through the engine heat exchanger and the hot helium line to the fuel tank distributor. During flight, the source of fuel tank pressurization is helium from storage bottles mounted inside the lox tank.
Fuel tank pressure switches control the fuel vent and relief valve, the GSE pressure supply during filling operations, prepressurization before engine ignition, and pressurization during flight. The flight pressurization pressure switch actuates one of the five control valves in the flow controller to ensure a minumum pressure of 24.0 psia during flight. The other four valves are sequenced by the IU to establish an adequate helium flow rate with decreasing storage bottle pressure.
The onboard helium storage bottles are filled through a filtered fill and drain line upstream from the flow controller. The storage bottles are filled to a pressure of 1400 psi prior to lox loading. Fill is completed to 3150 psi after lox loading when the bottles are cold.
RP-1 Delivery
Fuel feed is accomplished through two 12-inch ducts which connect the fuel tank to each F-l engine. The ducts are equipped with gimbaling and sliding joints to compensate for motions from engine gimbaling and stage stresses. Prevalves, one in each fuel line, serve as an emergency backup to the main engine fuel shutoff valves. The prevalves also house flowmeters which provide flowrate data via telemetry to GSE. A fuel level engine cutoff sensor, in the bottom of the fuel tank, initiates engine shutdown when fuel is depleted if the lox sensors have failed to cut off the engines.
LOX LOADING AND DELIVERY
As the oxidizer in the bi-propellant propulsion system, lox is contained and delivered through a separate tank and delivery system. The 345,000 gallon tank is filled through two 6-inch fill and drain lines. Shortly after T-6 hours lox loading begins. Three Fiji rates are used sequentially; a 300 gpm for tank chilldown, a 1500 gpm slow fill rate to stabilize the Squid level and thus prevent structural damage, and a fast fill rate of 10,000 gpm. At approximately 95% full the rate is reduced to 1500 gpm and ceases when the lox loading level sensor automatically stops the fill mode. Lox boiloff is replenished at 500 gpm until prepressurization occurs.
Lox Drain
The lox is drained through the two fill and drain intertank lines and an aft fill and drain line in the thrust structure. The aft fill and drain line is not used until 6.5% load is reached. During lox drain, positive ullage pressure is maintained by a GSE pressure source and two vent valves which are kept closed except when overpressure occurs.
Prior to launch, boll off in the lox tank may be harmlessly vented overboard. However, excessive geysering from boiling in the lox suction ducts can cause structural damage, and high lox temperatures near the engine inlets may prevent normal engine start. The lox bubbling system eliminates geysering and maintains low pump inlet temperatures. The helium induced convection currents curculate lox through the suction ducts and back into the tank. Once established, thermal pumping is self sustaining and continues until the interconnect valves are closed just prior to launch.
Lox Pressurization System
Lox tank pressurization is required to ensure proper engine turbopump pressure during engine start, thrust buildup, and fuel bum. The pressurization gas, prior to flight (prepressurization), is helium. It is supplied to the distributor m the lox tank. Lox tank pressure is monitored by three pressure switches. They control the GSE pressure source, the lox vent valve, and the lox vent and relief valve to maintain a maximum of 25 psig ullage pressure. Prepressurization is maintained until launch commit. Gox is used for pressurizing the lox tank during flight. A portion of the lox supplied to each engine is diverted from the lox dome into the engine heat exchanger where the hot turbine exhaust transforms lox into gox. The heated gox is delivered through the gox pressurization line and a flow control valve to the distributor in the lox tank. A sensing line provides pressure feedback to the flow control valve to regulate the gox flow rate and maintain ullage pressure between 18 and 20 psia.
Lox Delivery
Lox is delivered to the engines through five suction lines. The ducts are equipped with gimbals and sliding joints to compensate for motions from engine gimbaling and stage stresses. Pressure volume compensating ducts ensure constant lox flowrate regardless of the gimbaled position of the engine. Each suction line has a lox prevalve which is a backup to the engine lox valve. The prevalve cavity is charged with helium and functions as an accumulator to absorb engine induced pulses.
ELECTRICAL
The electrical power system of the S-IC stage is made up of two basic subsystems: the operational power subsystem and the measurements power subsystem. Onboard power is supplied by two 28-volt batteries. Battery characteristics are listed in figure 4-15.
In figure 4-16, battery number 1 is identified as the operational power system battery. It supplies power to operational loads such as valve controls, purge and venting systems, pressurization systems, and sequencing and flight control. Battery number 2 is identified as the measurement power system battery. It supplies power to measurements loads such as telemetry systems, transducers, multiplexers, and transmitters. Both batteries supply power to their loads through a common main power distributor but each system is completely isolated from the other.
During the prelaunch checkout period power for and electrical loads, except range safety receivers, is supplied from GSE. The range safety receivers are hardwired to batteries 1 and 2 in order to enhance the safety and reliability of the range safety system. At T-30 seconds a ground command causes the power transfer switch to transfer the S-IC electrical loads to onboard battery power. However, power for engine ignition and for equipment heaters (turbopump and lox valves) continues to come from the GSE until terminated at umbilical disconnect.
DISTRIBUTORS
There are six power distributors on the S-IC stage. They facilitate the routing and distribution of power and also serve as junction boxes and housing for relays, diodes, switches and other electrical equipment.
There are no provisions for switching or transferring power between the operational power distribution system and the measurement power system. Because of this isolation, no failure of any kind in one system can cause equipment failure in the other system.
Main Power Distributor
The main power distributor contains a 26-pole power transfer switch, relays, and the electrical distribution busses. It serves as a common distributor for both operational and measurement power subsystems. However, each of these systems is completely independent of the other. The power load is transferred from the ground source to the flight batteries at T-30 seconds. Inflight operation of the multicontact make-before-break power transfer switch is prevented by a brake, by mechanical construction, and by electrical circuitry. Operation of the switch several times during countdown verifies performance of the brake, motor, contacts, and mechanical components.
Sequence and Control Distributor
The sequence and control distributor accepts command signals from the switch selector and through a series of magnetically latching relays provides a 28volt dc command to initiate or terminate the appropriate stage function. The input from the switch selector latches a relay corresponding to the particular command. A 28-volt dc signal is routed through the closed contacts of the relay to the stage components being commanded. The relays, one for each command function, may be unlatched by a signal from GSE. The normally closed contacts of the relays are connected in series. A 28-volt dc signal is routed through the series connected relay contacts to indicate to GSE when all sequence and control relays are in the reset state.
Propulsion Distributor
The propulsion distributor contains relays, diodes, and printed circuit boards for switching and distributing propulsion signals during launch preparation and flight.
Thrust OK Distributor
The thrust OK distributor contains relays and printed circuit assemblies which make up the thrust OK logic networks and timers required to monitor engine thrust OK pressure switches and initiate engine shutdown. Signals from two of the three thrust OK pressure switches on a particular engine will result in an output from a two-out-of-three voting network. This output activates a 0.044 second timer. If the thrust OK condition is missing longer than 0.044 seconds the timer output sends a signal to initiate engine shutdown.
Timer Distributor
Circuits to time the operation of relays, valves, and other electromechanical devices are mounted in the timer distributor.
Measuring Power Distributor
Each regulated 5-volt dc output from the seven measuring power supplies is brought to an individual bus in the measuring power distributor and then routed to the measuring and telemetry systems.
SWITCH SELECTOR
The S-IC stage switch selector is the interface between the LVDC an the IU and the S-IC stage electrical circuits. Its function is to sequence and control various flight activities such as TM calibration, retrorocket initiation, and pressurization.
A switch selector is basically a series of low power transistor switches individually selected and controlled by an eight-bit binary coded signal from the LVDC in the IU. A coded word, when addressed to the S-IC switch selector, is accepted and stored in a register by means of magnetically latching relays. The coded transmission is verified by sending the complement of the stored word back to the LVDC in the IU. At the proper time an output signal is initiated via the selected switch selector channel to the appropriate stage operational circuit. The switch selector can control 112 circuits.
LVDC commands activate, enable, or switch stage electrical circuits as a function of elapsed flight time. Computer commands include:
- Telemetry calibration.
- Remove telemetry calibration.
- Open helium flow control valve No. 2.
- Open helium flow control valve No. 3.
- Open helium flow control valve No. 4.
- Enable center engine cutoff.
- Enable outboard engine cutoff.
- Arm EBW firing unit, retrorockets, and separation system.
- Fire EBW firing unit, retrorockets, and separation system.
- Measurement switchover.
In addition, a command from the emergency detection system in the IU can shut down all S-IC stage engines.
INSTRUMENTATION
The S-IC stage instrumentation system monitors functional operations of stage systems and provides signals for vehicle tracking during the S-IC burn. Prior to liftoff, measurements are telemetered by coaxial cable to ground support equipment. During flight, data is transmitted to ground stations over RF links. The ODOP system uses the doppler principle to provide vehicle position and acceleration data during flight. Section Vll provides a detailed discussion of the telemetry system.
TELEMETRY SYSTEM
The telemetry system accepts the signals produced by the measuring portion of the instrumentation system and transmits them to ground stations The telemetry equipment includes multiplexers, subcarrier oscillators, amplifiers, modulators, transmitters, and a two element antenna system. The telemetry system uses multiplex techniques (time sharing) to transmit large quantities of measurement data over a relatively small number of basic RF links.
There are two basic types of telemetry systems in the S-lC stage. One pulse amplitude modulated/frequency modulated (PAM/FM/FM) link is used for telemtering low-to-medium frequency data such as pressure or temperature. Time multiplexed data from the PAM links are also routed through the PCM links at one third sampling rate for DDAS transmission during preflight testing and for redundant RF transmission during flight. A pulse code modulated/digital data acquisition system (PCM/DDAS) link provides for acquisition of analog and digital flight data, provides a hardwire link for obtaining PCM data and PAM time multiplexed data during test and checkout and permits the redundant monitoring of PAM data during flight. The PCM/DDAS system assembles and formats PCM/FM time shared data so it can be sent over coaxial cables for automatic ground checkout or over an RF link during flight.
The RDSM provides additional data-handling capability to the PCM telemetry system. It can accept a maximum of 100 inputs which are sampled sequentially in groups of ten. The assembly handles only digital information. The inputs and outputs are set voltage levels that represent liquid level measurements and discrete data from other sources.
MEASUREMENT SYSTEM
The measurement system senses performance parameters and feeds signals to the telemetry system. It includes transducers, signal conditioning, and distribution equipment necessary to provide the required measurement ranges and suitably scaled voltage signals to the inputs of the telemetry system. The S-IC measuring system performs three main functions
- Detection of the physical phenomena to be measured and transformation of these phenomena into electrical signals.
- Process and condition the measured signals into the proper form for telemetering.
- Distribution of the data to the proper channel of the telemetry system.
Remote Automatic Calibration System (RACS)
The RACS is used to verify measurement circuit operation and continuity by stimulating the transducer directly, or by inserting a simulated transducer signal in the signal conditioner curcuit. Measurement operation is verified at 80 percent of the maximum transducer range (high level), at 20 percent of the maximum range (low level), and at the normal run level.
ODOP
An offset doppler (ODOP) frequency measurement system is an elliptical tracking system which measures the total doppler shift in an ultra high frequency (UHF) continuous wave (CW) signal transmitted to the S-IC stage. The ODOP system uses a fixed station transmitter, a vehicle borne transponder, and three or more fixed station receivers to detemine the vehicle position. In this system the transmitter, transponder, and one receiver describe am ellipsoid whose intersection with the first ellipsoid is a line. The addition of the third receiver produces a third ellipsoid whose intersection with the line of intersection of the first and second ellipsoids is a point. This point is the transponder.
ORDNANCE
The S-IC ordnance systems include the propellant (flight termination) system and the retrorocket system.
PROPELLANT DISPERSION SYSTEM
The S-IC propellant dispersion system (PDS) provides the means of terminating the flight of the Saturn V if it varies beyond the prescribed limits of its flight path or if it becomes a safety hazard during the S-IC boost phase. The system is installed on the stage in compliance with Air Force Eastem Test Range (AFETR) Regulation 127.9 and AFETR Safety Manual 127.1.
The PDS is a dual channel, parallel redundant system composed of two segments. The radio frequency segment receives, decodes, and controls the propellant dispersion commands. The ordnance twin segment consists of two exploding bridgewire (EBW) firing units, two EBW detonators, one safety and arming (S&A) device (shared by both channels), six confined detonating fuse (CDF) assembles, two CDF tees, two CDF/flexible linear shaped charge (FLSC) connectors, and two FLSC assemblies.
The S&A device is a remotely controlled electromechanical ordnance device that is used to make safe and to arm the S-IC, S-ll, and S-IVB stage PDS's. The device can complete and interrupt the explosive train by remote control, provide position indications to remote monitoring equipment, and provide a visual position indication. It also has a manual operation capability. The S&A device consists of a rotary solenoid assembly, a metal rotor shaft with two explosive inserts, and position sensing and command switches that operate from a rotor shaft cam. In the safe mode, the longitudinal axis of the explosive inserts are perpendicular to the detonating wave path, thus forming a barrier to the explosive twin. To arm the device, the shaft is routed 90 degrees to align the inserts between the EBW detonators and the CDF adapted to form the initial part of the explosive train.
Should emergency flight termination become necessary, two coded radio frequency commands are transmitted to the launch vehicle by the range safety officer. The first command arms the EBW firing units and initiates S-IC engine cutoff. The second command, which is delayed to permit charging of the EBW firing units, discharges the EBW firing units across the exploding bridgewire in the EBW detonators mounted on the S&A device. The resulting explosive wave propagates through the S&A device inserts to the CDF assemblies and to the CDF tees. The CDF tees propagate the wave through insulated CDF assemblies to the FLSC assemblies mounted on the lox and RP-I tanks. The FLSC's provide the explosive force to longitudinally sever the propellant tanks and dispense the propellants. There are six 88-inch FLSC sections mounted on the lox tank and three 88-inch sections on the fuel tank. These sections are positioned on the propellant tanks to minimize mixing of the propellants after the tanks are severed.
RETROROCKETS
The S-IC retrorockets are mounted, in pairs, in the fairings of the F-1 engine. At retrorocket ignition the forward end of the fairing is burned and blown through by the exhausting gases. Each retrorocket is pinned securely to the vehicle support and pivot support fittings at an angle of 7.5 degrees from center line. The thrust level developed by seven retrorockets (one retrorocket out) is adequate to separate the S-IC stage a minimum of six feet from the vehicle in less than one second.
The eight retrorockets, provide separation thrust after S-IC burnout. The firing command originates in the Instrument Unit and activates redundant firing systems. Additional redundancy is provided by interconnection of the two confined detonating fuse (CDF) manifolds with CDF assemblies. The exploding bridgewire (EBW) firing unit circuits are grounded by a normally closed relay until the firing command is initiated by the instrument Unit. High voltage electrical signals are released from the two EBW firing units to the EBW detonators upon vehicle deceleration to 0.5g. The signals cause the detonator bridgewires to explode, thereby detonating the surrounding explosives. The explosion then propagates through the CDF manifold explosive and CDF assemblies into the igniter assembly. The igniter assembly located within the base of each retrorocket is then ignited, causing a buildup and release of the gases into the main grain of the retrorocket. Each retrorocket is ignited by either of two CDF initiators mounted on its aft structure. Operational ground check of the system through the firing unit is accomplished through use of pulse sensors which absorb the high voltage impulse from the firing unit and transmit a signal through the telemetry system. The pulse sensors are removed prior to launch.
Each retrorocket is a solid propellant rocket with a case bonded, twelve-point star, internal burning, composite propellant cast directly into the case and cured. The propellant is basically ammonium perchlorate oxidizer in a polysulfide fuel binder. The motor is 86 inches long by 15-1/4 inches diameter and weights 504 pounds, nominal, of which 278 pounds is propellant.
MAJOR DIFFERENCES BETWEEN SATURN V S-IC-5 AND S-IC-6 STAGES
Telemetry reduced to a system utilizing one FM/FM link, one PCM/FM link and two antennae from a system which included two SSB/FB links, three FM/FM links, one PCM/FM link, four antennae and a tape recorder.
Copyright 1997, 1998 by John
Duncan |
