Category Apollo Saturn V News Reference

FLIGHT INSTRUMENTATION 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

FLIGHT INSTRUMENTATION SYSTEM

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LH2 Continuous Propulsive Vent System

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.

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LH, Feed System —Prior to vehicle liftoff and prior

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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.

PRIMARY PACKAGE

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.

FLAME DEFLECTOR

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

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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.

Uprated Saturn I (Saturn IB)

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.

ONBOARD CONTROL PRESSURE SYSTEM

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

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Control Pressure System

Подпись: SATURN V NEWS REFERENCE brings the storage bottle pressure up to its normal capacity of 3,250 plus or minus 50 psig. 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.

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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.

PROPELLANT UTILIZATION SYSTEM

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.

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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.

Engine Operation

START SEQUENCE

Start sequence is initiated by supplying energy to two spark plugs in the gas generator and two in the augmented spark igniter for ignition of the propellants. Next, two solenoid valves are actuated: one for helium control, and one for ignition phase control. Helium is routed to hold the propellant bleed valves closed and to purge the thrust cham­ber LOX dome, the LOX pump intermediate seal, and the gas generator oxidizer passage. In addition, the main fuel valve and ASI oxidizer valve are opened, creating an ignition flame in the ASI cham­ber that passes through the center of the thrust chamber injector.

After a delay of 1, 3, or 8 seconds, during which time fuel is circulated through the thrust chamber to condition the engine for start, the start tank dis­charge valve is opened to initiate turbine spin. The length of the fuel lead is dependent upon the length of the Saturn V first stage boost phase. When the J-2 engine is used in the second stage of the Saturn V vehicle, a one-second fuel lead is necessary. The third stage of the Saturn V vehicle, on the other hand, utilizes a three-second fuel lead for its initial start and an eight-second fuel lead for its restart.

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After an interval of 0.450 second, the start tank discharge valve is closed and a mainstage control solenoid is actuated to: 1) turn off gas generator and thrust chamber helium purges; 2) open the gas generator control valve (hot gases from the gas generator now drive the pump turbines); 3) open the main oxidizer valve to the first position (14 degrees) allowing LOX to flow to the LOX dome to burn with the fuel that has been circulating through the injector; 4) close the oxidizer turbine bypass valve (a portion of the gases for driving the oxi­dizer turbopump were bypassed during the igni­tion phase); 5) gradually bleed the pressure from the closing side of the oxidizer valve pneumatic actuator controlling the slow opening of this valve for smooth transition into mainstage. Energy in the spark plugs is cut off and the engine is operating at rated thrust. During the initial phase of engine opera­tion, the gaseous hydrogen start tank will be re­

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charged in those engines having a restart require­ment. The hydrogen tank is repressurized by tapping off a controlled mixture of liquid hydrogen from the thrust chamber fuel inlet manifold and warmer hy­drogen from the thrust chamber fuel injection mani­fold just before entering the injector.

FLIGHT MAINSTAGE OPERATION

During mainstage operation, engine thrust may be varied between 175,000 and 225,000 pounds by actuating the propellant utilization valve to increase or decrease oxidizer flow as described in the sec­tion “PU Valve”. This is beneficial to flight trajec­tories and for overall mission performance to make greater payloads possible.

CUTOFF SEQUENCE

When the engine cutoff signal is received by the electrical control package, it de-energizes the main – stage and ignition phase solenoid valves and ener­gizes the helium control solenoid de-energizer timer.

This, in turn, permits closing pressure to the main fuel valve, main oxidizer valve, gas generator con­trol valve, and augmented spark igniter valve. The oxidizer turbine bypass valve and propellant bleed valves open and the gas generator and LOX dome purges are initiated.

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THE LAUNCH AREA

Final preparation of the space vehicle for launching, including propellant and ordnance loading, final checkout, and countdown takes place in the launch area.

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Aerial of Pad 39A with VAB in Background

There are presently two launch areas on Complex 39. Each area is polygon-shaped with the linear distance from side to side at approximately 3,000 feet. The launch sites are 8,730 feet apart to allow operations on the pads to be handled independently for safety reasons.

Liquid oxygen, RP-1, and liquid hydrogen are stored near the perimeter of the launch sites. Heli­um and nitrogen gases are stored at 10,000 psi near the center.

An elevated steel and concrete hardstand is located in the center of each area. Steel support fittings for the mobile launcher and mobile service struc­ture are anchored to the hardstand. The exhaust flame trench runs through the center of the hard­stand. Prior to the launch, the wedge-shaped flame deflector is moved along rails into the trench.

The liquid oxygen system consists of a 900,000- gallon LOX storage facility and transfer system.

The RP-1 system consists of a storage area con­taining three 86,000-gallon tanks and a transfer system. The tanks have a carbon steel shell and a bonded stainless steel lining.

Gaseous nitrogen and helium are stored under­ground in vessels near the launch pad at pressures of 6,000 psi.

Automation of vehicle prelaunch checkout is ex­pected to uprate mission confidence and to increase launch rate capability. The heart of this automatic checkout system is the computer complex.

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Cutaway Illustration of Pad 39A