Category Apollo Saturn V News Reference

The first stage instrumentation system measures and reports information on stage systems and com­ponents and provides data on internal and external environments. It keeps abreast of approximately 900 measurements on the stage, such as measure­ments of valve positions, propellant levels, tem­peratures, voltages, and pressures. The measure­ments are telemetered by coaxial cable to ground support equipment and by radio frequency trans­mission to ground stations.

The instrumentation system consists of a measure­ment system, a telemetry system, and the Offset Doppler tracking system. A remote automatic cali­bration system provides remote rapid checkout of the measurements and telemetry systems.

MEASUREMENT

The measurement system reports environmental situations and how the first stage reacts to them. Making use of transducers, signal conditioners, measuring rack assemblies, measuring distributors, and the onboard portion of the remote automatic calibration system, this system involves many phases of stage operation. Included are measure­ments of acceleration, acoustics, current, flow, flight

TELEMETRY

Telemetry is a method of remote monitoring of flight information accomplished by means of a radio link. The first stage telemetry system is composed of six radio frequency links.

Most of the components of the telemetry systems are located in the thrust structure; RF assemblies and a tape recorder are located in the forward skirt. The telemeter transmits data through two common antenna systems.

Links FI, F2, and F3 are identical systems which transmit narrow-band, frequency-type data such as that generated by strain gages, temperature gages, and pressure gages. The system can handle 234 measurements on a time-sharing basis and 14 mea­surements transmitted continously. Data may be sampled either 120 times per second or 12 times per second.

Links SI and S2 transmit wide-band, frequency – type data generated by vibration sensors. Each link provides 15 continuous channels or a maximum of 75 multiplexed channels depending on the specific measuring program.

Telemeter PI transmits cither pulse code modu­lated or digital type data. Five multiplexers, four analogs, and one digital supply data to the PCM assembly. This provides the most accurate data and is used for ground checkout as well.

A telemetering calibrator is used to improve the accuracy of the telemetry systems. The calibrator supplies known voltages to the telemeters periodi­
cally during the stage operation. Their reception at tracking stations provides a valid reference for data reduction.

The effects of ullage and retrorocket firing attenu­ation can seriously degrade the telemetry trans­mission during stage separation; therefore, a tape recorder installed in the forward skirt records data for delayed transmission. The commands for tape recorder operation originate in the digital computer located in the instrument unit.

ODOP SYSTEM (Offset Doppler Tracking System)

The ODOP system is an elliptical tracking system that measures the rate of motion at which the ve­hicle is moving away from or toward a tracking station. The total Doppler shift in the frequency of a continuous wave, ultra-high frequency signal transmitted from the ground to the first stage is measured. The signal is received by the transponder at the stage, modified, and then retransmitted back to the ground. Retransmitted signals are received simultaneously by three tracking stations. Separate antennas on the stage are used for receiving and retransmitting the signals.

The checkout valve consists of a ball, a poppet, and an actuator. The checkout valve provides for ground checkout of the ignition monitor valve and fuel valves and prevents the ground hydraulic return fuel, used during checkout, from entering the en­gine system and consequently the vehicle fuel tank.

When performing the engine checkout or servicing, the checkout valve ball is positioned so fuel enter­ing the engine hydraulic return inlet port will be directed through the ball and out the GSL return port. For engine static firing or flight, the ball is positioned so fuel entering the engine hydraulic re­turn inlet port will be directed through the ball and out the engine return outlet port.

Engine Control Valve

(Hydraulic Filter and Four-Way Solenoid Valve Manifold)

The engine control valve incorporates a filter mani­fold, a four-way solenoid valve, and two swing check valves.

The filter manifold contains three filters. One filter is in the supply system and one each in the opening and closing pressure systems. The filters prevent entry of foreign matter into the four-way solenoid valve or the engine. Two swing check valves are "teed” into the supply system filter. The check valves permit hydraulic system operation from the ground supplied hydraulic fluid for checkout and servicing procedures or engine supplied hydraulic fluid for normal engine operation.

The four-way solenoid valve is comprised of a main spool and sleeves to achieve two-directional control of the fluid flow to the main fuel, main oxidizer, and gas generator valve actuators. The spool is pressure-positioned by two three-way slave pilots. Each slave pilot has a solenoid-controlled, normally open, three-way primary pilot.

The de-energized position of the engine control valve provides hydraulic closing pressure to all engine propellant valves. Momentary application of 28 VDC to the start solenoid will initiate control valve actuations that culminate in the positioning of the main spool so that hydraulic pressure is applied to the opening port, and the pressure previously applied to the closing port is vented to the return port.

An internal passage in the housing maintains com­mon pressure applied between the opening port and start solenoid poppet. This pressure, after start solenoid de-energization, holds the main spool in its actuated position thereby maintaining the pres­sure directed to the opening port without further application of the start solenoid electrical signal. Momentary application of 28 VDC to the stop so­lenoid will initiate control valve actuations that culminate in positioning the main spool so that pres­sure is vented from the opening port and applied to the closing port. The override piston may be actuated at any time by a remote pressure supply, which, in the event of an electrical power loss, would re­position the main spool and apply hydraulic pres­sure to the closing port. If electrical power and hydraulic pow-er are both removed, the valve will return to the de-energized position by spring force. If hydraulic pressure is then reapplied, pressure will be applied to the closing port. If an electrical signal is simultaneously sent to the start and stop solenoids, the stop solenoid will override the start and return the valve to a deactuated position.

Swing Check Valve

There are two identical swing check valves installed on the engine control valve. They allow – the use of ground hydraulic fuel pressure during engine start­ing transient and engine hydraulic fuel pressure during engine mainstage and shutdown. One check valve is installed in the engine hydraulic fuel supply inlet port, the other in the ground hydraulic fuel supply inlet port.

Major systems required for third stage operation

are the propulsion system, flight control system, electrical power and distribution system, instru­mentation and telemetry system, environmental control system, and ordnance systems.

Propulsion System

The propulsion system consists of the J-2 engine, propellant system, pneumatic control system, and propellant utilization system. The J-2 engine burns LOX as an oxidizer and LH2 as fuel at a nominal mixture ratio of 5:1. Both fuel and oxidizer sys­tems utilize tank pressurization systems and have vent and relief capabilities to protect the propel­lant tanks from overpressurization. The pneumatic control system regulates and controls both the oxi­dizer and fuel systems. The propellant utilization (PU) system assures simultaneous and precise fuel and oxidizer depletion by controlling engine mix­ture ratio.

Propulsion System Components

J-2 ENGINE

The engine system consists of the J-2 engine, pro­pellant feed system, start system, gas generator system, control system, and a flight instrumenta­tion system. The propellant feed system utilizes independently driven, direct-drive fuel and oxi­dizer turbopumps to supply propellants at the prop­er mixture ratio to the engine combustion chamber. Additional information on the J-2 engine system may be found in the J-2 Engine section.

The electrical sequence controller is a completely self-contained, solid-state system, requiring only DC power and start and stop command signals.

Pre-start status of all critical engine control func­tions is monitored in order to provide an “engine ready” signal. Upon obtaining “engine ready” and “start” signals, solenoid control valves are ener­gized in a precisely timed sequence as described in the “Engine Operation” section to bring the en­gine through ignition, transition, and into main – stage operation. After shutdown, the system auto­matically resets for a subsequent restart.

start Tank Assembly System

This system is made up of an integral helium and hydrogen start tank, which contains the hydrogen and helium gases for starting and operating the en­gine. The gaseous hydrogen imparts initial spin to the turbines and pumps prior to gas generator com­bustion, and the helium is used in the control system to sequence the engine valves.

HELIUM AND HYDROGEN TANKS

The spherical helium tank is positioned inside the hydrogen tank to minimize engine complexity. It holds 1,000 cubic inches of helium. The larger spher­ical hydrogen gas tank has a capacity of 7,257.6 cubic inches. Both tanks are filled from a ground source prior to launch and the gaseous hydrogen tank is refilled during engine operation from the thrust chamber fuel inlet manifold for subsequent restart in third stage application.

Flight Instrumentation System

The flight instrumentation system is composed of a primary instrumentation package and an auxiliary package.

Propellant facilities at Launch Complex 39 include a LOX system, the RP-1 system, the liquid hydrogen system, the propellant tanking computer system, the spacecraft support system, and the data trans­mission system.

The propellant tanking computer system provides a means of monitoring amounts during the fueling operations. It also accurately controls fuel level during the final phase of tank fill and replenish.

The data transmission system provides an accurate method for the transmission of propellant and en­vironmental control system electrical signals from the launch site to the LCC.

The liquid oxygen system provides oxidizer fill and drain for the three stages of the Saturn V. The sys­tem includes a storage tank, a vaporizer, two re­plenishing pumps, transfer lines, vent lines and drain basin, and electric circuitry for monitoring and actuating the pneumatic control system.

The round liquid oxygen storage tank holds 900,000 gallons and is situated 1,450 feet from the launch site. It has a stainless steel inner wall 62 feet 9 inches in diameter. The space between this inner sphere and the outer wall is filled with gaseous nitrogen and perlite for insulation.

To load liquid oxygen, a command originates in the LCC at the LOX control panel. The signal is trans­mitted to the mobile launcher by the data trans­mission system and then to the LOX storage area.

The electrical signals are converted to pneumatic pressure to operate the vaives, and the flow of LOX from the storage tank into the vaporizer begins. The vaporizer converts the liquid oxygen into gas­eous oxygen, which then is fed back into the tank to pressurize it to the 10 psig needed to begin the flow. The pumps are started and the LOX is pumped through the transfer lines to the vehicle.

The RP-1 system provides fuel fill, drain, and filter­ing capabilities for the first stage. The system in­cludes three storage tanks each with a capacity of 86,000 gallons, transfer lines, a launch site facility, and electric circuitry.

The liquid hydrogen system provides fueling and draining for the second and third stages. It includes a storage tank with a capacity of 850,000 gallons, a vaporizer, transfer lines, and a burn pond in which excess propellant is burned.

The double walled storage tank, 1,450 feet from the launch site, has a stainless steel inner wall with a diameter of 61 feet 6 inches. The space between the inner and outer walls is filled with perlite.

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.

SATURN V NEWS REFERENCE

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.

SATURN V NEWS

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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REFERENCE

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.