Category Apollo Saturn V News Reference

Ten fuel suction lines (two per engine) supply fuel from the fuel tank to the five F-l engines. The suc­tion line outlets attach directly to the F-l engine fuel pump inlets.

Each suction line has a pneumatically controlled fuel prevalve which normally remains open. This

FUEL-CONDITIONING (BUBBLING) SYSTEM

The fuel-conditioning system bubbles gaseous ni­trogen through the fuel feed lines and fuel tank to prevent fuel temperature stratification prior to launch. A wire mesh filter in the nitrogen supply line prevents discharge of contaminants into the conditioning system.

Fuel Conditioning

A check valve in the outlet of each fuel-conditioning line prevents fuel from entering the nitrogen lines when the fuel-conditioning system is not operating.

An orifice located near each fuel-conditioning check valve provides the proper nitrogen flow into each fuel duct.

The oxidizer pump supplies oxidizer to the thrust chamber and gas generator at a flowrate of 24.811 gpm. The pump consists of an inlet, an inducer, an impeller, a volute, bearings, seals, and spacers. Oxidizer is introduced into the pump through the inlet which is connected by duct to the oxidizer tank. The inducer in the inlet increases the pressure of the oxidizer as it passes into the impeller to pre­vent cavitation. The impeller accelerates the oxi­dizer to the desired pressure and discharges it through diametrically opposed outlets into the high-pressure oxidizer lines leading to the thrust chamber and gas generator.

The oxidizer inlet, which attaches to a duct leading to the vehicle oxidizer tank, is bolted to the oxi­dizer volute. Two piston rings seated between the inlet and the volute expand and contract with tem­perature changes to maintain an effective seal between the high and low pressure sides of the inlet. Holes in the low-pressure side of the inlet allow leakage past the ring seals to flow into the suction side of the inducer, thus maintaining a low – pressure.

The oxidizer volute is secured to the fuel volute with pins and bolts which prevent rotational and axial movement. The primary oxidizer seal and spacer located in the oxidizer volute prevent fuel from leaking into the primary oxidizer seal drain cavity. The oxidizer intermediate seal directs a purge

Fuel Pump

The fuel pump supplies fuel to the thrust chamber and gas generator at a flowrate of 15,471 gpm. The pump consists of an inlet, an inducer, an impeller, a volute, bearings, seals, and spacers. Fuel is intro­duced into the pump from the vehicle fuel tank through the inlet. The inducer in the inlet increases the pressure of the fuel as it passes into the impeller to prevent cavitation. The impeller accelerates the fuel to the desired pressure and discharges it through two diametrically opposed outlets into the high-pressure fuel lines leading to the thrust chamber and gas generator.

The fuel volute is bolted to the inlet and to a ring, which is pinned to the oxidizer volute. A wear-ring installed on the volute mates against the impeller. The cavity formed between the volute and the impeller is called the balance cavity. Pressure in the balance cavity exerts a downward force against the fuel impeller and counterbalances the upward force of the oxidizer impeller to control the amount of shaft axial force applied to the No. 1 and No. 2 bearings. Leakage between the impeller inlet and the discharge is controlled by a wear-ring, which mates with the impeller and acts as an orifice. The fuel volute provides support for the bearing retainer, which supports the No. 1 and No. 2 bearings and houses the bearing heater. The No. 3 seal, which is installed between the oxidizer intermediate seal and the No. 1 bearing, prevents lubricating fuel for the bearings from contacting the oxidizer. If fuel should pass the seal, purge flow from the oxidizer intermediate seal will expel the fuel overboard. On the fuel side of the No. 2 bearing, the No. 4 lube seal contains the lubricant within the bearing cavity. The remaining seal in the fuel volute is the primary seal and contains fuel under pressure in the balance cavity, maintains the desired balance cavity pres­sure, and keeps high-pressure fuel out of the low – pressure side.

The propellant system is composed of seven sub­systems : purge, fill and replenish, venting, pres­surization, propellant feed, recirculation, and pro­pellant management..

Purge Subsystem

The purge subsystem uses helium gas to clear the propellant tanks of contaminants before they are loaded. The important contaminants art* oxygen in the liquid hydrogen tank (liquid hydrogen will freeze oxygen which is impact-sensitive) and moisture in the liquid oxygen tank.

The tanks are purged with helium gas from ground storage tanks. The tanks are alternately pressur­ized and vented to dilute the concentration of con­taminants. The operation is repeated until samples of the helium gas emptied from the tanks show that contaminants have been removed or reduced to a safe level.

Fill and Replenish Subsystem

Filling of the propellant tanks on the second stage is a complex and precise task because of the nature of thd cryogenic liquids and the construction of the stage.

Because the metal of the stage is at normal outside temperature, it must be chilled gradually before pumping the ultra-cold propellants into the tanks. The filling operation thus starts with the introduc­tion of cold gas into the tanks, lines, valves, and other components that will come into contact with the cryogenic fluids. The cold gas is circulated until

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Channel Installed—Feed line from IH2 tank to one of the five engines is installed.

the metal has become chilled enough to begin pump­ing in the propellants. The filling and replenishing subsystem operation consists of five phases:

Chilldown – Propellants are first pumped into the tank at the rate of 500 gallons per minute for LQX and 1,000 gallons per minute for LHa. Despite the preliminary chilling by cold gas, the tanks are still so much warmer than the propellants that much of the latter boils off (converts to gaseous form) when it first goes into the tank. Filling con­tinues at this rate until enough of the propellants remain liquid so that the tanks are full to the five per cent level.

Fast Fill —As soon as tank sensors report that the liquid has reached the five per cent level, the fill­ing rate is increased to 5,000 gallons per minute for LOX and 10,000 gallons per minute for LH2. This rate continues until the liquid level in the tank reaches the 98 per cent level.

Slow-Fill—Propellant tanks are filled at the rate of 1,000 gallons per minute for both LOX and LH2 until the 100 per cent level is reached.

Replenishment—Because filling begins many hours before a scheduled liftoff and the cryogenic liquids are constantly boiling off, filling continues almost up to liftoff (160 seconds before liftoff for LOX and 70 seconds before liftoff for LH2). Tanks

are filled at the rate of up to 200 gallons per min­ute for LOX and up to 500 gallons per minute for LH,, depending on signals from sensors in the tanks on the liquid level.

101 Per Cent Shutdown—A sensor in each tank will send a signal to indicate that the 101 per cent level (over the proper fill level) has been reached; this signal causes immediate shutdown of filling operations.

Filling is accomplished through separate connec­tions, lines, and valves. The ground part of the con­nections is covered by special shrouds in which he­lium is circulated during filling operations. This provides an inert atmosphere around the coupling between the ground line and the tanks.

The coupling of the fill line and the tanks is engaged manually at the start of filling operations; it is nor­mally disengaged remotely by applying pneumatic pressure to the coupling lock and actuating a push – off mechanism. A backup method involves a remotely attached lanyard in which the vertical rise of the vehicle will unlock the coupling. The fill valves are designed so that loss of helium pressure or electrical power will automatically close them.

Liquid oxygen is the first propellant to be loaded onto the stage. It is pumped from ground storage tanks. Liquid hydrogen is transferred to the stage by pressurizing the ground storage tanks with gaseous hydrogen. The liquid hydrogen tank is chilled before the liquid oxygen is loaded to avoid structural stresses.

After filling is completed, the fill valves and the liquid oxygen debris valves in the coupling are closed, but the liquid hydrogen debris valve is left open. The liquid oxygen fill line is then drained and purged with helium. The liquid hydrogen line is purged up to the coupling. When a certain signal is received (first stage thrust-commit), the liquid hydrogen debris valve is closed and the coupling is separated from the stage.

The tanks can be drained by pressurizing them, opening the valves, and reversing the filling opera­tion.

Four solid propellant retrorockets are mounted equidistant around the aft interstage assembly, and when ignited, assure clean separation of the third stage from the second stage by decelerating or braking the spent booster. Each retrorocket is rated for a nominal thrust of 35,000 pounds, weight of 384 pounds, and burn time of about 1.5 seconds.

A signal from the second stage initiates two EBW firing units located on the aft interstage. The EBW firing units ignite two detonator manifolds, which in turn ignite the retrorockets through redundant pairs of confined detonating fuse (CDF) and py­rogen initiators.

Ullage Rocket System

ULLAGE CONTROL ROCKET IGNITION AND JETTISON SYSTEM

Two solid propellant ullage rockets, located on the third stage aft skirt just forward of the stage sepa­ration plane, are ignited on signal from the stage sequencer by EBW initiators.

After firing, the burned-out ullage rocket casings and fairings are jettisoned to reduce stage weight. Upon command from the stage sequencer, two forward and aft frangible nuts, which secure each rocket motor and its fairing to the stage, are det­onated by confined detonating fuse (CDF), to free the entire assembly from the vehicle.

Three IBM-owned buildings at Huntsville comprise the Space Systems Center where component test­ing, fabrication, assembly, and systems checkout of the instrument unit are completed. Assembly and the majority of the testing activity take place in a 130,000-square-foot building located in Hunts­ville’s Research Park.

As units are received, they are inspected and then moved to one of the testing laboratories where they are subjected to detailed quality and reliability testing. From component testing, the parts move

Ш Assembly and Test—All instrument unit assembly work and the majority of testing are done in this IBM-owned building in Huntsville’s Research Park. The rear of the building is the high – bay area where assembly operations take place.

Following assembly operations, the IU is moved to one of two systems checkout stands—one for uprated Saturn I vehicles, the other for Saturn V.

Automatic Checkout-IBM technicians monitor systems checkout tests as another technician optically adjusts the inertial guidance platform, prior to a simulated mission.

A complete systems checkout is performed auto­matically. Hooked by underground cables, two digital checkout computer systems examine the IU. Each of the IU’s six subsystems is tested before the IU is tested as an integrated unit. With indepen­dent computers, systems tests for two instrument units can be conducted simultaneously.

Simulation Laboratory Saturn V flight guidance and navigation programs as well as launch computer programs are tested in IBM’s Engineering Building at Huntsville. Here a technician checks a computer readout of a simulated mission.

A cutoff sensor mounted on the bottom of the fuel tank provides signal voltages to shut off fuel after a predetermined level of depletion is reached. The fuel is measured during flight by four fuel slosh probes and a single liquid level measuring probe. Fuel levels are detected electronically and reported through the stage telemetry system. Telemetry

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signals are transmitted to ground support either by radio frequency or, before launch, by coaxial cable. The cutoff sensor, mounted in the lower fuel tank bulkhead, initiates engine cutoff as fuel level falls below two sensing points on the probe. Engine cutoff will normally be initiated by sensors in the LOX system. The cutoff capability is provided as a backup system should fuel be depleted before LOX.

Fuel Level Sensing and Engine Cutoff

FUEL PRESSURIZATION SYSTEM

The fuel pressurization system maintains enough pressure in the fuel tank to provide proper suction at the fuel turbopumps to start and operate en­gines. The system consists of a helium supply, a helium flow controller, helium fill and drain com­ponents, a prepressurization subsystem, a fuel tank vent and relief valve, and associated ducts.

Four 31-cubic-foot, high pressure storage bottles in the LOX tank store the helium required for in­flight pressurization of the fuel tank ullage. A high pressure line is used for filling the bottles and rout­ing the helium to the flow controller. A solenoid dump valve is installed for emergencies. The helium flow controller uses five solenoid valves mounted parallel in a manifold to control helium flow to the fuel tank ullage. The cold helium duct routes helium from the flow controller to the cold helium mani­fold. From there, it is distributed to the heat ex­changers on the five F-l engines. The hot helium manifold receives the heated, expanded helium from the engine heat exchangers and routes it to the hot helium duct which then carries it through the he­lium distributor and on to the fuel tank ullage.

SATURN V NEWS REFERENCE

It is replenished between the periods of loading and prepressurization through the fill and drain line.

Before LOX drain can be performed, the helium cylinders in the LOX tank must have their pressure decreased from about 3,100 psig to about 1,660 psig. Fill and drain valves are opened to complete drainage of the LOX tank although total evacua­tion of LOX from the tank requires draining the engines or waiting for boil-off of residual LOX. LOX drain can be speeded with the aid of a pres­surizing gas, usually nitrogen.

The turbine, producing 55,000 brake horsepower, drives the fuel and oxidizer pumps. It is a two-stage, velocity-compounded turbine consisting of two ro­tating impulse wheels separated by a set of stators. The turbine mounts on the fuel pump end of the turbopump so that the two elements of the turbo­pump having the greatest operating temperature extremes (1500 Fahrenheit for the turbine and -300

Hot gas from the gas generator enters the turbine at a flowrate of 170 pounds per second through the inlet manifold and is directed through the first-stage nozzle onto the 119-blade first-stage wheel. The hot gas then passes through the second-stage stators onto the 107-blade second-stage wheel, and then into the heat exchanger. This flow of hot gas rotates the turbine, which in turn rotates the propellant pumps. Turbine speed during mainstage operation is 5,550 rpm.

Bearing Coolant Control Valve

This valve, which incorporates three 40-micron fil­ters, three spring-loaded poppets, and a restrictor, performs two functions. Its primary function is to control the supply of coolant fuel to the turbopump bearings. Its secondary function is to provide a means of preserving the turbopump bearings be­tween static firings or during engine storage. During engine firing, the coolant poppet opens and delivers filtered fuel to the turbopump bearing coolant jets, and the restrictor provides the proper turbopump bearing jet pressure.

GAS GENERATOR SYSTEM

The gas generator system provides the hot gases for driving the velocity-compounded turbine, which drives the fuel and oxidizer pumps. The system con­sists of a gas generator valve, an injector, a com­bustion chamber, and propellant feed lines connect­ing the No. 2 turbopump fuel and oxidizer outlet lines to the gas generator. The propellants are supplied to the gas generator from the No. 2 turbo­pump fuel and oxidizer outlet lines. The gas genera­tor mixture ratio, relative to the engine mixture ratio, is fuel-rich. This provides a lower combustion temperature in the uncooled gas generator and in the turbine.

Propellants enter the gas generator through the valve and injector and are ignited in the combustion chamber by dual pyrotechnic igniters. The gas generator valve is hydraulically operated by fuel pressure from the hydraulic control system.

The venting subsystem is used during loading and flight operations. While the propellant tanks are being loaded, the vent valves (two for each tank) are opened by electrical signals from ground equipment to allow the gas created by propellant boil-off to leave the tanks. The valves are spring-loaded to be normally closed, but a relief valve will open them if pressure in the tanks reaches an excessive level. Each valve is capable of venting enough gas to

The LH2 recirculation subsystem pumps the propel­lants through the feed lines and valves and back to the LIL tank through a single return line. The pumps are powered from a 56-volt DC battery sys­tem located in the interstage; the batteries are ejected with the interstage approximately 30 sec­onds after first plane separation. Refore liftoff, power for the LH2 recirculation subsystem is sup­plied by ground equipment.

The LOX recirculation system works on the basis of a thermal syphon; heat entering the system is used to provide pumping action by means of fluid density differences across the system. Helium gas is used to supplement the density differences and thereby improve the pumping action.

Recirculation of oxygen begins at the start of tank­ing; LIL recirculation begins just before launch. The propellants continue to circulate through first stage firing and up until just before the first stage and second stage separate. While the subsystems are operating, the LH2 prevalves which lead to the combustion chambers are closed; as soon as the re­circulation subsystem stops, the LH2 prevalves open and the engines ignite.

Propellant Management System

The propellant management system controls load­ing, flow rates, and measurement of the propel­lants. It includes propellant utilization, propellant loading, propellant mass indication, engine cutoff, and propellant level monitoring subsystems.

PROPELLANT UTILIZATION SUBSYSTEM

The propellant utilization subsystem controls the flow rates of liquid hydrogen and liquid oxygen in such a manner that both will be depleted simulta­neously. It controls the mixture ratio so as to min­imize propellant residuals (propellant left in the tanks) at engine cutoff. Propellant utilization bypass valves at the liquid oxygen turbopump outlets con­trol flow of liquid oxygen in relation to the liquid hydrogen remaining. Control of the engine mixture ratio increases the stage’s payload capability. The propellant utilization subsystem is interrelated with the propellant loading subsystem and uses some of the same tank sensors and ground checkout equip­ment.

PROPELLANT LOADING SUBSYSTEM

The loading subsystem is used to control propellant loading and maintain the quantity of propellants
in the tanks. Capacitance probes (sensors) running the full length of the propellant tanks sense liquid mass in the tanks and send signals to an airborne computer, which relays them to a ground computer to control loading. They also send signals to an airborne computer for the propellant utilization subsystem’s control of flow rates.

PROPELLANT MASS INDICATION SUBSYSTEM

The propellant mass indication subsystem is in­tegrated with the propellant loading subsystem and is used to send signals to the flight telemetry sys­tem for transmission to the ground. It utilizes pro­pellant loading sensors to determine propellant levels.

ENGINE CUTOFF SUBSYSTEM

The main function of the engine cutoff subsystem is to signal the depletion point of either propellant. It is an independent subsystem and consists of five sensors in each propellant tank and associated electronics. The sensors will initiate a signal to shut down the engines when two out of five sensors in the same tank signal that propellant is depleted.

The range safety system terminates vehicle flight upon command of the range safety officer. Redun­dant systems are used throughout to provide max­imum reliability.

Four antennas, mounted around the periphery of the third stage forward skirt assembly, feed two redundant secure range receivers located in the for­ward skirt assembly. Both receivers have separate power supplies and circuits. A unique combination of coded signals must be transmitted, received, and decoded to energize this destruct system.

A safety and arming device prevents inadvertent initiation of the explosive train by providing a posi­tive isolation of the EBW detonator and explosive train until arming is commanded. Visual and remote indications of SAFE and ARMED conditions are displayed at all times at the firing center. Upon proper command, EBW firing units activate EBW detonators.

A CDF, detonated by the safety and arming device, explodes a flexible linear-shaped charge which cuts through the tank skin to disperse both fuel and oxidizer.

I SATURN V NEWS REFERENCE

J-2 ENGINE FACT SHEET

LENGTH

WIDTH

NOZZLE EXIT DIAMETER THRUST (altitude)

SPECIFIC IMPULSE RATED RUN DURATION FLOWRATE: Oxidizer Fuel

MIXTURE RATIO CHAMBER PRESSURE (Pc)

WEIGHT, DRY, FLIGHT CONFIGURATION

EXPANSION AREA RATIO

COMBUSTION TEMPERATURE

Note: J-2 engines will be uprated to a maximum

11 ft. 1 in.

6 ft. SV2 in.

6 ft. 5 in.

225,0 lb.

424 sec. (427 at 5:1 mixture ratio)

500 sec.

449 lb sec (2,847 gpm)

81.7 lb sec (8,365 gpm)

5.5:1 oxidizer to fuel 763 psia 3,480 lb.

27.5:1

5,750°F

of 230,000 pounds of thrust for later vehicles.

I SATURN V NEWS REFERENCE

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

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

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

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

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

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

F-l Flight Engine Firing

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

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

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

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

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

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