Category Apollo Saturn V News Reference

Saturn I

Studies which led to the Saturn family of rockets were started by the Wernher von Braun organiza­tion in April of 1957. The aim of the program was to create a 1.5 million-pound-thrust booster by cluster­ing previously developed and tested engines.

On Aug. 15, 1958, the Advanced Research Projects Agency I ARP A) formally initiated what was to be­come the Saturn project. The agency, a separately organized research and development arm of the Department of Defense, authorized the Army Ballis­tic Missile Agency to conduct a research and devel­opment program at Redstone Arsenal for a 1.5

Test Vehicle… The first assembled Apollo Saturn V vehicle

approaches the launch pad at Kennedy Space Center. It was used to verify launch facilities, train launch crews, and develop test and checkout procedures at KSC. It was roiled out on

May 25, 1966.

million-pound-thrust vehicle booster. A number of available rocket engines were to be clustered and tested by a full-scale static firing by the end of 1959.

The program objectives were expanded by ARPA in October of 1958 to include a multi-stage carrier vehicle capable of performing advanced space mis­sions. Concurrent with the development of a multi­stage vehicle, static test facilities at Redstone Arsenal and launch complex facilities at Cape Canaveral—now Cape Kennedy—were being con­structed.

The proposed large vehicle project was officially renamed Saturn on Feb. 3, 1959, by ARPA memo­randum. The space agency assumed technical di­rection of the Saturn project in Sate 1959. The pro­ject was transferred officially on Mar. 16. I960, and the Army development group at Huntsville was transferred to NASA and became the nucleus of the new Marshall Space Flight Center. The first static­firing of a Saturn I booster was conducted April 29, 1960.

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The NASA Saturn Vehicle Evaluation Committee (Silverstein Committee) on Dec. 15, 1959, recom­mended a long-range development program for a Saturn vehicle with upper stage engines burning liquid hydrogen and liquid oxygen. The initial ve­hicle, identified as Saturn C-l and now as Saturn I, was to be a stepping stone to a larger vehicle. A building-block concept was proposed that would yield a variety of Saturn configurations, each using previously proven developments as far as possible.

Early in I960 the Satum program was given the highest national priority, and a 10-vehicle research and development program was approved.

The two-stage Saturn I vehicle with the Apollo spacecraft was about 188 feet tall and weighed some

1,125,0 pounds at liftoff.

While plans for the lunar mission were progressing, the Saturn I project made history. On Oct. 27, 1961, the first Saturn I booster was flight tested success­fully from Cape Kennedy. The first flight booster with dummy upper stages was called SA-1. This vehicle was followed by successful flights of SA-2 on April 25, 1962, SA-3 on Nov. 16, 1962, and SA-4 on Mar. 28, 1963.

The SA-5 vehicle, combining the first stage (S-l) with the second stage (S-IV), was successfully launch­ed on Jan. 29, 1964, with both stages functioning

perfectly to place a 37,700-pound payload into earth orbit. SA-6, launched on May 28, 1964, and SA-7, launched on Sept. 18,1964, each placed “unmanned” boilerplate configurations of Apollo spacecraft into earth orbit.

SA-9, launched on Feb. 19, 1965, was the first Saturn I vehicle to launch a Pegasus meteoroid tech­nology satellite into earth orbit.

The SA-8 and SA-10 Satum I vehicles were success­fully launched on May 25, 1965, and July 30, 1965, respectively, also placing a Pegasus satellite into earth orbit to complete the test and launch pro­gram with an unprecedented 100 per cent record of success.

A redundant initiation system actuates the separa­tion of the first stage from the second stage. A command signal for arming and another for firing the initiation systems are programmed by the in­strument unit computer.

After LOX depletion, the computer signals operate relays in the switch selector and sequence and con­trol distributor to control the exploding bridgewire firing units. When armed, the firing units store a high voltage electrical charge. When fired, the electrical charge actuates the ordnance.

Two firing units are installed on the first stage for the eight retrorockets, and two are installed on the second stage for the separation ordnance.

Range Safety System

The function of the range safety system is to provide ground command with the capability of flight termi­nation by shutting off the engines, blowing open the stage propellant tanks, and dispersing the fuel in event of a flight malfunction.

The system is redundant, consisting of two identical, independent systems, each made up of electronic and ordnance subsystems.

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Flight termination by way of the range safety sys­tem goes into effect upon receipt of the proper radio frequency commands from the ground. A frequency-modulated RF signal transmitted from the ground range safety transmitter is received by the antennas and transmitted by way of a hybrid ring to the range safety command receiver. There, the signal is conditioned, demodulated, and decoded.

Range Safety System

The resulting signal simultaneously causes arming of the exploding bridgewire firing unit and shut­down of the stage engines. A second command sig­nal transmitted by the ground range safety trans­mitter ignites the explosive train (detonating fuses and shaped charges) to blow open the stage pro­pellant tanks.

Control Pressure System

The control pressure system supplies pressurized gaseous nitrogen for the pneumatic actuation of propellant system valves and purging of various F-l engine systems.

The complete integrated system is made up of an onboard control pressure system, a ground control

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pressure system, and an onboard purge pressure system. The object in each system is to deliver an actuating or purge medium to an interfacing stage system.

The flight instrumentation system consists of pres­sure transducers, temperature transducers, posi­tion indicators, a flow measuring device, power dis­tribution junction boxes, and associated electrical harnesses, and permits monitoring of engine per­formance. The basic flight instrumentation system is composed of a primary and an auxiliary system. The primary instrumentation system is critical to all engine static firings and subsequent vehicle launches; the auxiliary system is used during re­search, development, and acceptance portions of the engine static test program and initial vehicle flights. The flight instrumentation system compo­nents, including both the primary and auxiliary systems, are listed below:

Primary Instrumentation

Fuel turbopump inlet No. 1 pressure Fuel turbopump inlet No. 2 pressure Common hydraulic return pressure Oxidizer turbopump bearing jet pressure Combustion chamber pressure Gas generator chamber pressure Oxidizer turbopump discharge No. 2 pressure Fuel turbopump discharge No. 2 pressure Oxidizer pump bearing No. 1 temperature Oxidizer pump bearing No. 2 temperature Turbopump bearing temperature Turbopump inlet temperature Turbopump speed

Auxiliary Instrumentation

Oxidizer turbopump seal cavity pressure

Turbine outlet pressure

Heat exchanger helium inlet pressure

Heat exchanger outlet pressure

Oxidizer turbopump discharge No. 1 pressure

Heat exchanger LOX inlet pressure

Heat exchanger GOX outlet pressure

Fuel turbopump discharge No. 1 pressure

Engine control opening pressure

Engine control closing pressure

The continuous vent system is used to provide a thrust force required to position propellants at the aft end of each tank during coast. The system con­sists of a vent line originating at the vent-relief valve, terminating at two low’ thrust nozzles located 180° apart, and facing aft on the forward skirt. Continuous venting is controlled and regulated by a pneumatically operated continuous propulsive vent module.

At the completion of the first burn engine cutoff, APS ullage engines are activated to settle the liquid propellants in the aft end of the tanks during the shutdowm phase. LHZ tank pressure is then vented through the continuous propulsive vent system, providing a continuous propulsive thrust to the stage. This maintains control of the propellants within the tanks. The APS engines are shut off after the transition is complete and the propulsive venting continues throughout the coast phase. The continuous propulsive vent module controls vent­ing from a maximum of 45 pounds to a minimum of approximately 7 pounds.

LH, Feed System —Prior to vehicle liftoff and prior

to engine restart, all LH2 feed system components of the J-2 turbopump assembly must be chilled to assure proper operation, Chilldown of the LH2 sys­tem is accomplished by a closed loop, forward-flow recirculation system. On command from the IU, the prevalve in the LH, feed duct closes and the chill – down shutoff valve opens. An auxiliary, electrically driven LH, chilldown pump, mounted in the LH2 tank, circulates the LH, within the system and is capable of a minimum flowrate of 135 gpm at 6.1 psi.

LH, is circulated from the LH2 tank through the low pressure feed duct, through the J-2 engine fuel pump, the fuel bleed valve, and back to the tank through a return line. Recirculation chilldown con­tinues through the boost phase and up to J-2 engine ignition. In the event of an emergency shutdown requirement, the chilldown system shutoff valve is closed upon command from the IU. LH, is sup­plied to the J-2 engine through a vacuum-jacketed, low-pressure duct at a flowrate of 81 pounds per second at -423° Fahrenheit, 28 psia. The duct is located in the fuel tank side wall above the common bulkhead joint and is equipped with bellows to compensate for thermal motion. Signals from the engine sequencer energize the LH2 feed valve, as required to obtain steady-state operation, A com­plete description of engine operation may be found in the J-2 Engine section.

The primary package instrumentation measures those parameters critical to all engine static firings and subsequent vehicle launches. These include some 70 parameters such as pressures, tempera­tures, flows, speeds, and valve positions for the engine components, with the capability of trans­mitting signals to a ground recording system or a telemetry system, or both. The instrumentation system is designed for use throughout the life of the engine, from the first static acceptance firing to its ultimate vehicle flight.

AUXILIARY PACKAGE

The auxiliary package is designed for use during early vehicle flights. It may be deleted from the basic engine instrumentation system after the pro­
pulsion system has established its reliability during research and development vehicle flights. It con­tains sufficient flexibility to provide for deletion, substitution, or addition of parameters deemed nec­essary as a result of additional testing. Eventual deletion of the auxiliary package will not interfere with the measurement capability of the primary package.

To dissipate the rocket exhaust from the F-l en­gines, a flame deflector, a flame trench, and a water deluge system are used in the launch area.

The inverted V-shaped steel flame deflector fea­tures a replaceable ceramic-coated leading edge. Exhaust from the outer engines strikes the point of the inverted V. At the same time, the deflector is exposed to water deluge during and after liftoff.

The center engine exhaust impinges on the ceramic

View of Pad 39A East Side at KSC and Flame Trench from North End

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leading edge. The heat resistant ceramic surfaces erode slowly in the blast. As they do, the great thermal energy generated is carried away in super­heated particles. All exhaust and particles are de­flected through a flame trench where their energy is dissipated harmlessly into the atmosphere.

The mobile deflector weighs 700,000 pounds and is moved to its position beneath the launch pedestal along a rail system. Two deflectors are available for each launch area, although only one is required per launch.

The space agency, using the building-block approach, conceived the Uprated Saturn I as the quickest, most reliable, and most economical means of provid­ing a vehicle with greater payload than the Saturn

I. This vehicle was planned for orbital missions with the Apollo spacecraft before the Satum V vehicle would be available.

The Uprated Satum I is based on a blending of existing elements of Saturn I and Saturn V. A re­designed Saturn I booster (designated the S-IB stage), and an S-IVB upper stage and instrument unit from the Satum V are used on this launch vehicle.

Maximum use of designs and facilities available from the earlier approved Satum programs saved both time and costs. .

The Satum I first stage was redesigned in several areas by NASA and the Chrysler Corporation, the stage contractor, for the expanded role as the Up­rated Saturn I booster. Basically, it retained the same shape and size, but required some modifica­tion for mating with the upper stage, which has a greater diameter and weight than the Satum I upper stage.

Stage weight was cut by more than 20,000 pounds to increase payload capacity. The Rocketdyne H-l engine was uprated to 200,000 pounds of thrust, compared with 188,000 pounds of thrust for each engine in the final Saturn I configuration. The en­gines will be improved again to 205,000 pounds beginning with the SA-206.

For the Uprated Saturn I, a guidance computer used in the early Satum I was replaced by another IBM computer of completely new design which in­corporates the added flexibility and extreme re­liability necessary to carry out the intended Uprated Satum I missions.

The Uprated Satum I, topped by the Apollo spaee-

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craft, stands approximately 224 feet tall, and is about 21.7 feet in diameter. Total empty weight is about 85 tons, and liftoff weight fully fueled is approximately 650 tons.

Several uprated Saturn I vehicles have been launched since the original SA 201 launch on Feb. 26, 1966.

The onboard control pressure system consists of a high-pressure nitrogen storage bottle, an umbilical coupling and tubing assembly for filling the storage bottle, a manifold assembly, and control valves at the terminal ends of various nitrogen distribution lines. In some cases, two valves are paired with other associated equipment and block-mounted to form a control assembly.

The nitrogen onboard storage bottle has 2,200- cubic inch capacity and is made of titanium alloy. It is designed for a maximum proof pressure of 5,000 psig. It is filled and discharged through a port in the single boss. During flight launch preparation, the bottle is filled from a ground supply first to a pressurization of 1,600 psig well in advance of final countdown. This weight pressure is adequate for any prelaunch operational use. The second step occurs in the last hour of the launch countdown and

Control Pressure System

The manifold assembly serves as a gaseous nitrogen central receiving and distributing center as well as a mounting block for filters, shutoff solenoid valves, a pressure regulator, a relief valve, and pressure transducers. Ported manifolds provide tubing as­sembly connections to the storage bottle, umbilical coupling, and various tubing assembly distribution lines to control valves throughout the stage.

Heat exchanger LOX inlet temperature Heat exchanger GOX outlet temperature Heat exchanger helium outlet temperature Fuel pump inlet No. 2 temperature Heat exchanger LOX inlet flowrate

Primary and Auxiliary Junction Box

There are two electrical junction boxes in the flight instrumentation system. The primary junction box has provisions for eight electrical connectors, and the auxiliary junction box for five. Both junction boxes are welded closed and pressurized with an inert gas to prevent possible entry of contaminants and moisture.

ENGINE OPERATION

The engine requires a source of pneumatic pres­sure, electrical power, and propellants for sustained engine operation. A ground hydraulic pressure source, thrust chamber prefill, gas generator and turbine exhaust igniters, and hypergolic fluid are required to start the engine.

When the start button is actuated, the checkout valve moves to transfer the hydraulic fuel return from the ground line to the turbopump low-pressure fuel inlet. The high-level oxidizer purge is initiated to the gas generator and thrust chamber LOX dome.

The gas generator and turbine exhaust gas igniters fire, and the engine control valve start solenoid is energized. Hydraulic pressure is directed to the opening port of the oxidizer valves. The oxidizer valves are part way open, and the hydraulic pres­sure is directed to the gas generator valve opening port. The gas generator valve opens, propellants under tank pressure enter the gas generator com­bustion chamber, and the propellant mixture is ig­nited by the gas generator igniters. The exhaust gas is ducted through the turbopump turbine, the heat exchanger, and the thrust chamber exhaust manifold into the nozzle extension walls where the fuel-rich mixture is ignited by the turbine exhaust gas igniters. As the turbine accelerates the fuel and the oxidizer pumps, the pump discharge pressures increase and propellants at increasing flowrates are supplied to the gas generator. Turbopump acceler-

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ation continues and, as the fuel pressure increases, the igniter fuel valve opens and allows fuel pres­sure to huild up against the hypergol cartridge hurst diaphragm. The hypergol diaphragms burst under the increasing fuel pressure. Hypergolic fluid, followed by the ignition fuel, enters the thrust chamber. When hypergolic fluid enters the thrust chamber and contacts the oxidizer, spontaneous combustion occurs, establishing thrust chamber ignition. Thrust chamber pressure is transmitted through the sense line to the diaphragm of the ignition monitor valve. When the thrust chamber pressure increases, the ignition monitor valve ac­tuates and allows hydraulic fluid flow to the open­ing port of the fuel valves. The fuel valves open and fuel is admitted to the thrust chamber.

Fuel enters the thrust chamber fuel inlet manifold and passes through the thrust chamber tubes for cooling purposes and then through the injector into the thrust chamber combustion zone. As the thrust chamber pressure increases, the thrust-OK pres­sure switches are actuated indicating the engine is operating satisfactorily. The thrust chamber pres­sure continues to increase until the gas generator reaches rated power, controlled by orifices in the propellant lines feeding the gas generator. When engine fuel pressure increases above the ground-
supplied hydraulic pressure, the hydraulic pres­sure supply source is transferred to the engine. Hydraulic fuel is circulated through the engine com­ponents and then returned through the engine con­trol valve and checkout valve into the turbopump fuel inlet. The ground hydraulic source facility shutoff valve is actuated to the closed position when the fuel valves open. This allows the engine hy­draulic system to supply the hydraulic pressure during the cutoff sequence.

The primary function of the PU system is to assure simultaneous depletion of propellants by controlling the LOX flowrate to the J-2 engine. It also provides propellant mass information for controlling the fill and topping valves during propellant loading opera­tions. The system consists of mass sensors, an elec­tronics assembly, and an engine-mounted mixture ratio valve.

Propellant Utilization System

During loading operations, the mass of propellants loaded is determined within one per cent by the mass sensors. Tank over-fill sensors act as a backup system in the event the loading system fails to ter­minate fill operations.

Continuous LH2 and LOX residual readout signals are provided throughout third stage powered flight. Differences between the fuel and oxidizer mass indications, as sensed by the mass sensors, are continually analyzed and are then used to control the oxidizer pump bypass flowrate, which changes the engine mixture ratio correspondingly. The static inverter/converter supplies the analog volt­ages necessary to operate the PU system. It is commanded “on" and “off” by a switch selector and sequencer combination.