The titan ii handbook

Each missile carried a single warhead—the largest in U. The missiles were deployed at basing facilities in Arkansas, Arizona, and Kansas and remained in active service for over twenty years. Since military deactivation in the early s, the Titan II has served as a reliable satellite launch vehicle.

This is the richly detailed story of the Titan II missile and the men and women who developed and operated the system. David K. Stumpf uses a wide range of sources, drawing upon interviews with and memoirs by engineers and airmen as well as recently declassified government documents and other public materials.

Over drawings and photographs, most of which have never been published, enhance the narrative. The three major accidents of the program are described in detail for the first time using authoritative sources. Titan II will be welcomed by librarians for its prodigious reference detail, by technology history professionals and laymen, and by the many civilian and Air Force personnel who were involved in the program—a deterrent weapons system that proved to be successful in defending America from nuclear attack.

The data presented was derived both from a literature survey and from a test program conducted at Bell Aerosystems Company and at the U.

This switch also initiates the 5. The 5. Retrorocket number 1 is fired by K As in automatic retrorocket fire, each retrorocket is fired from retrograde squib bus number i and number 2. Twenty-two seconds after retrofire is initiated, the last retrorocket ceases firing. The switch energizes the retrograde separate shaped charge relay K, the retrograde bias off relay K, and the horizon scanner heads Jettison relay K Relay K4-I7 fires retrograde adapter shaped charge igniter , , and and pyrotechnic switch H Relay K latches the re-entry roll display relay K removing roll mix interlock from the flight director controller.

K also resets two latch relays: the retrograde bias relay K and the indicate retrograde attitude relay K K fires horizon scanner cover squib if it was not fired previously during the boost phase. K ignites the horizon scanner head squib through an millisecond pyrotechnic time delay and jettisons the scanner head. K fires the release igniters of docking latches 1, 2 and 3 to jettison them.

K also energizes the index bar Jettison and latch door release relay K KM fires three latch door cover release igniters. These igniters release the latch doors which cover the ports left by the jettisoned docking latches.

K also jettisons the docking index bar. If the bar was not extended previously, it is first extended and then Jettisoned. These functions are not a part of the retrograde sequence during an abort if the abort occurs prior to nose fairing jettison. Using the attitude controller and the FDI needles, he rolls the spacecraft degrees so that the horizon is visible in the upper portion of his cabin window.

The command pilot uses attitude control and maneuvering electronics and the attitude controller to control the roll attitude during approximately the next 10 minutes in which the altitude diminishes to , feet.

The computer computes the roll attitude for optimum reentry lift and also automatically controls the roll attitude.

During approximately the next 10 minutes, the altitude decreases to , feet. At this altitude, the altimeter indicator begins to come off the peg. At 80, feet, the computer commands the spacecraft to assume the best attitude for drogue parachute deployment. Then the command pilot places all guidance and electronic switches to OFF. An abort is an unscheduled termination of the spacecraft mission.

An abort may be initiated at any time during the spacecraft mission. In all cases the actual abort sequence has to be initiated by the crew after an abort command has been received. An abort indication consists of illumination of the ABORT indicators located on the command pilot and pilot's panels. During pre-launch prior to umbilical disconnect, the ABORT indicator may be illuminated from the blockhouse via hard-line through the launch vehicle tail plug connector. After umbilical release, the ABORT indicator may be illuminated by ground command to the spacecraft via a channel of the DCS or by ground command to the launch vehicle to shutdown the booster.

The abort sequence is part of the Sequence System. The abort sequence comprises the abort indicators, controls, relays, and pyrotechnics. The part of the abort sequence which the crew make use of is determined by the abort mode in effect at the time when the abort command is received or the decision to abort is made.

The abort mode to be used at any time during the mission is determined by calculations made on the ground and depends on the altitude and velocity attained by the spacecraft. The critical abort altitudes are 15, feet, 75, feet, and , feet. The spacecraft reaches 15, feet approximately 50 seconds after lift-off, 75, feet approximately seconds after lift-off, and , feet approximately seconds after lift-off. Below 15, feet, seat ejection mode I is used. Between 15, and 75, feet, ride-it-out abort mode I-II is used.

Between 75, and , feet, modified re-entry mode II is used. Above , feet normal re-entry mode III is used, except that the spacecraft electronic timer does not illuminate the sequential indicators amber when the time to press them occurs, unless the timer is updated by ground command. When an abort becomes necessary during pre-launch, it is accomplished by using abort mode I. The abort command is given from the blockhouse by hard-line through the launch vehicle tall plug connector.

When the pilots see this display, they immediately pull the D-rings attached to their ejection seats. When one D-ring is pulled, both ejection systems are energized. One-half seconds later, the hatches are open, and one-half second after that the seats have been ejected.

Sensors detect the ejection of the seats and notify the blockhouse that the pilots are out of the spacecraft. One-quarter second after the seats are ejected, a sustainer rocket under each seat is fired, which extend the distance between the pilots add the launch vehicle. Then a pyrotechnic ignites and separates the ejection seat from the pilots. Two seconds after sustainer ignition, the main parachutes have opened and the pilots are lowered safely to the ground. After normal lift-off, and before the Gemini-Titan reaches an altitude of 15, feet, an abort condition could develop.

The crew monitor their booster indicators so that they are aware at all times of the manner in which the flight is proceeding. Booster operation data is telemetered to the ground for analysis and interpretation.

The range safety officer, the booster systems engineer, the flight director, or the flight dynamics officer, who are on the ground, any decide that danger is imminent and an abort mandatory. A channel of the DCS is used to send the abort command to the spacecraft and ground commands are sent to the launch vehicle to shutdown the booster engines. Then the engine shutdown tones are received, the destruct switches of the launch vehicle are armed. The command pilot and pilot evaluate these displays and pull the D-rlngs.

The hatches open and the pilots in their seats are ejected. Refer to Section III for a description of the remainder of this sequence. Abort mode I - II is the ride-it-out abort mode. It is effective at altitudes between 15, and 75, feet approximately 50 seconds to seconds after lift-off. Abort mode I - II is used when a mode I abort is inadvisable and when a delay to permit entry into the mode II conditions is impractical. The crew however has the option to eject or to ride-it-out depending upon their assessment of the abort conditions.

Therefore the D-rings are not stowed during the I - II mode. Abort mode I - II begins during stage I boost approximately 50 seconds after liftoff. The retrograde abort relays and the retrograde abort interlock relays are energized. These relays arm the buses needed for abort action. The retrograde common control bus is armed from the common control bus.

Retrograde squib buses number 1 and number 2 are armed from OAMS squib buses number 1 and number 2. On spacecraft 5 only, spacecraft separation squib buses number 1 and number 2 are armed from Boost Insert Abort BIA squib buses number 1 and number 2. Two parallel circuits are used for redundancy. This arming of buses by means of relays eliminates the motion of the switch ordinarily required to arm the buses.

Then, in rapid succession, wire guillotine relays, pyrotechnic switch relays, and shaped charge igniter relays are energized. Then, the four retrorockets are salvo fired and the spacecraft thrusts away from the launch vehicle. If the abort altitude is between 15, and 25, feet the retrograde adapter is jettisoned 7 seconds after retrorocket salvo fire is initiated.

If the abort altitude is between 25, and 75, feet, the retrograde adapter is jettisoned 45 seconds after salvo fire. After retrograde adapter Jettison, the spacecraft is maneuvered to the re-entry attitude.

If the abort altitude is above 40, feet, the drogue parachute is deployed at 40, feet, and the main parachute at 10, feet. If the drogue parachute fails or has not been deployed before the spacecraft descents to 10, feet, the emergency main parachute switch is used to deploy the main parachute.

If one of the two first stage engines should fall and the launch vehicle is above 40, feet, the pilots may elect to remain with the spacecraft until the operating engine has boosted them to 75, feet. At this altitude, abort mode I-II becomes inapplicable.

Abort mode II becomes effective above 75, feet. At approximately seconds after lift-off on a normal mission, the launch vehicle has boosted the spacecraft to an altitude of 75, feet. The ground station notifies the crew via the uhf communications link of the change to abort mode II. Both the command pilot and pilot acknowledge the change via the same link, and stow the ejection seat handles D-ring. Initiation of abort mode I above 75, feet could be disastrous. Abort mode II begins during stage 1 boost before booster engine cutoff and ends during stage 2 boost before second stage engine cutoff.

The crew continues to monitor the booster indicators. If they should notice an abort situation developing, they analyze it. The decision to abort may be theirs or it may come from the ground. In abort mode II, the command pilot must act. The operating engine is cutoff. Since orbital velocity could not have been reached below , feet, the spacecraft immediately begins a re-entry trajectory.

The spacecraft is maneuvered to the retrograde blunt end forward attitude, the retrograde section is jettisoned, and normal landing procedures are initiated. At approximately seconds after lift-off, the launch vehicle reaches the altitude of , feet and a velocity of approximately 21, feet per second. The shutdown command is thus given to the second stage engine. OAMS thrust is applied to put distance between the second stage and the spacecraft. The crew perform the Tr seconds and the Tr seconds procedures, using the main instrument panel switch-indicators.

After retrofire has been initiated manually, normal re-entry, landing, and postlanding procedures are followed. Abort mode I, the seat ejection mode, is not covered here. Abort mode III is executed by performing a launch vehicle engine shutdown, a spacecraft separation sequence and a retrograde sequence.

Separation and retrograde in abort mode III differs from normal separation and retrograde in that the abort sequence is performed without cues from the indicators on the main instrument panel. BIA common control bus power is applied to the launch vehicle engine shutdown signal relays K and K This power is also applied to the engine shutdown relays in the Titan Launch Vehicle.

The operating engine s are cut off. As K and K energize, common control bus power is applied through their B contacts to the spacecraft instrumentation programmer. The programmer encodes the voltage from this bus as the booster cutoff command signal for telemetry transmission to the ground tracking station. However five of these relays are key relays in that they control the principal abort operations. These operations are: 1 telemetry of the abort action to the ground; 2 arming of the retrograde buses; 3 activation of the RCS; 4 separation of the spacecraft from the launch vehicle; and 5 salvo firing of the retro rockets.

The relays which control those operations are: i the instrumentation abort relay, K; 2 the squib bus abort relay K; 3 the Attitude Control System abort relay K; 4 the retrograde abort relay K; and 5 the salvo retrograde relay KI.

When the instrumentation abort relay K is energized by the abort switch, its B contacts connect common control bus power to the spacecraft instrumentation programmer. The programmer encodes this signal as the pilot actuated abort signal for telemetry transmission to the ground. Abort, if it occurs, requires that power for the circuits used in the retrograde phase of the mission become immediately available. When the abort switch is closed, squib bus power is applied to K K arms the retrograde squib buses i and 2 and the retrograde common control bus.

Re-entry immediately and automatically follows an abort. Re-entry requires the use of the RCS for control of the spacecraft during this phase.

Hence the RCS is activated. Activation involves opening and pressurizing the RCS fuel and oxidizer lines. This is done by firing the squibs of the fuel, oxidizer, and pressurant packages.

The squibs thus fired open the ring A fuel and oxidizer lines and pressurize them. K applies retrograde squib bus power to similar igniter of RCS ring B with similar results.

The B contact of F and K energize the retrograde abort interlock relay K K, contact A initiates the station 7. Since the retrorockets are to be fired in the abort modes controlled by the abort switch, the spacecraft must separate from the launch vehicle at station ZTO. ZTO is on the mating line between the spacecraft and the equipment adapter section. To make separation complete, the OAMS propellant lines which cross this station must be sealed and guillotined.

The abort switch energizes the retrograde abort relay K which arms K, the OAMS lines guillotine latch relay; K, the retrograde abort pyrotechnic switch relay; and K, the wire guillotine relay.

When K is energized, it energizes K, K, and K The guillotine now seals and cuts the lines. The lower wire bundles are guillotined. These wires like the propellant lines, must also be guillotined, and the guillotine blade could cause a short circuit of the spacecraft power. K4-T4 energizes pyrotechnic switch relays K and K K ignites equipment adapter pyrotechnic switches D, E end F. K ignites fuel cell wiring pyrotechnic switch B, C and S.

This is accomplished by actuating the wire guillotines. When K and K energize, they apply power through the A contacts of K to wire guillotine relay K K, contact C energizes the separate electrical latch relay K, the adapter shaped charge relay K and the abort discrete relay K K, contact A latches K in the energized position.

K changes the computer from the ascent mode to the re-entry mode and enables the computer to accept re-entry data and solve the re-entry problem. K prepares the way for the fourth step in the separation of the launch vehicle from the spacecraft. The fourth and final step is to sever the adapter skin at station Z70 and breaks the launch vehicle to spacecraft structural bond. When K causes K, the adapter shaped charge relay to energize, K fires the ZT0 tubing cutter igniter and the equipment adapter shaped charge igniters.

The retrorockets are salvo fired at the same time that the tubing and structural bond is cut. To salvo fire the retrorockets, power must be applied simultaneously to the retrorocket automatic fire relays and thus to the retrorockets. Therefore the 5. Contacts C, D and E of K bypass the time delay relays. As these relays energize, retrograde squib bus power is applied to the igniters of retrorockets i, 3, 2 and 4.

Salvo burn lasts approximately 5. When the retrorocket automatic fire relays are energized by K, the second time delay relay K is also energized. However, in a mode I-II abort when the altitude is between 15, and 25, feet, the switch-indicator is pressed seven seconds after the retrorockets begin firing.

After the retrograde section has been jettisoned, normal re-entry and landing procedures are initiated. The switches and circuit breakers on the left switch and circuit breaker panel perform important functions in the operation of the Sequence System.

The top tow of circuit breakers however pertain largely to communications. The second row of circuit breakers perform functions related to the operation of the Sequence System. Their functions are as follows:. The electronic timer circuit breaker CB applies main bus power through contact A of lift-off relay K to start the electronic timer when the lift-off signal energizes the K The timer begins counting the time-to-go to retrograde.

The event timer circuit breaker CB applies main bus power through contact B of lift-off relay K3-ll to start the event timer when the lift-off signal energizes K The event counter counts the time since lift-off occurred.

This circuit breaker arms the booster shutdown circuit and the secondary guidance manual switch-over circuit. The boost cutoff 2 circuit breaker CB applies BIA common control bus power redundantly to the booster shutdown switch, and supplies power for the second stage engine cutoff signal input to the computer. It provides power to salvo fire the retrorockets during the abort sequence. If CB is not closed, the electronic timer Tr contact closure will not automatically fire the retrorockets.

The retrograde manual circuit breaker CB provides retrograde common control bus power for manually firing the retrorockets, and salvo firing the retrorockets with the abort control handle. The retrograde minus seconds circuit breaker CB applies common control bus power to relay contacts in the electronic timer and contacts of the TR second relay. The sequence lights power circuit breaker CB applies main bus power to the sequence light BRIGHT-DIM switch and to open contacts on the barostat switch arm relay and the message acceptance pulse relay.

The functions are the following:. It also arms the abort discrete relays and the equipment adapter separation sensor switches and relays.

The retrograde sequence control 2 circuit breaker CB connects the retrograde squib bus number 2 redundantly to the same switches to which the retrograde sequence control 1 circuit breaker connects power and arms the same circuits.

The sequence light bright-dim switch is a single-pole, double-throw toggle switch. It connects the bus through a resistor to the same circuits in the DIM position.

These switches arm or safety the various squib buses used by the Sequential System. Their functions are as follows. The boost-insert squib bus arm-safe switch is a four pole, double throw toggle switch.

The retrograde power squib bus arm-safe switch is a four-pole, double-throw switch. In the ARM position, it arms retrograde squib bus 1 and 2 and the retrograde common control bus. The retrograde Jettison squib bus arm-safe switch is a two-pole double-throw toggle switch. In the ARM position, it arms retrograde Jettison squib buses number 1 and number 2. From these buses, the retrograde jettison relays get the power to fire the retrograde adapter shaped charges and retrograde pyrotechnic switch H.

The four retrograde rocket squib arm switches apply the voltages which ignite the four retrofire rockets to open contacts of the retro rocket automatic and manual fire relays. In the safe position of these four switches, the ignition voltage is removed from the relays. Seven indicators, three meters and four controls are provided for the boost-insert-abort phase of the spacecraft mission.

The two ENGINE I indicators are provided on the co-pilot's panel to indicate thrust chamber underpressure of the first stage booster engines. Each indicator illuminates red when the thrust chamber pressure of the engine is 68 percent of rated pressure or less. Both indicators illuminate red at stage 1 ignition but extinguished 0. Both indicators illuminate at booster engine cut-off and extinguish quickly at staging. The ENGINE II indicator on the command pilot's panel illuminates amber to indicate the fuel injector underpressure or off condition of the second stage engine.

The critical pressure for engine 2 is 55 percent of rated value. The indicator illuminates when the first stage engine is ignited and stays amber through first stage boost. The attitude rate indicator on the command pilot's panel indicates an evaluation of the launch vehicle attitude rates during the boost phase. The indicator is extinguishes if the attitude rates remain within acceptable limits but illuminates red if the rates exceed these 1imits. The secondary guidance indicator on the command pilot's panel indicates which guidance system is in operation.

The indicator is extinguished to indicate that primary guidance is being used. The indicator illuminates amber to indicate that secondary guidance has been selected. Both indicators illuminate red when the abort command is transmitted. The indicator signals the crew to initiate immediately the abort mode appropriate for the altitude and velocity of the spacecraft.

These modes are described under Sequence System Operation. During the boost phases the crew has been reminded via the uhf communications link of the abort mode in effect. The stage 1 fuel end oxidizer meters on the command pilot's panel enable the crew to monitor the current status and progress of the boost phase, and to anticipate an abort condition if one should develop.

These meters indicate the gas pressures in psia of the stage 1 fuel and oxidizer tanks. Dual indicator needles are provided for redundancy. The range of the stage 1 meters is 35 to 5 psia. A time-versus- pressure scale near the bottom of the meter shows the minimum required pressure at 20, 40, and 60 seconds after lift-off.

Critical fuel tank pressure is indicated by a shaded column at the low end of the scale. After staging with no signals applied, the meters indicate maximum psia. The stage 2 fuel and oxidizer meters on the command pilot's panel indicate stage 2 fuel and oxidizer tank pressure over a 70 to 10 psia range. Redundant pointers are used. Critical fuel tank pressures are indicated by a shaded column at the low end of the scale.

The S-flag at the psia mark indicates the minimum acceptable stored pressure in the tank before pressurization. After spacecraft separation, the meters indicate maximum psia. The accelerometer on the command pilot's panel indicates the rate in g's at which the launch vehicle engines are changing the velocity of the spacecraft. The range of the accelerometer is minus 6g's to 16 g's.

The meter has positive and negative memory pointers. The accelerometer enables the crew to monitor the effectiveness of the engines. It is a secondary indicator of staging. The guidance switch above the abort control handle permits the command pilot to manually change from primary guidance to secondary backup guidance. When back-up guidance has been selected either manually or automatically during stage I boost and the ground station determines that primary guidance is feasible during stage 2 boost, primary guidance can be selected again by momentarily placing the guidance switch to the RGS position.

A D-ring is provided on the ejection seat of each pilot. These rings are pulled to initiate mode I abort at altitude below 70, feet. Refer to Section III of this volume for the location and operation of these devices. The abort control handle is located on the command pilot's side of the cabin. These modes are effective above 25, feet. When the abort handle is moved to ABORT, an immediate spacecraft separation and retrograde sequence is performed.

These sequences differs from the normal sequences in that they are performed without cues from the indicators on the main instrument panel. The switches, indicators, and switches-indicators on the main instrument panel center console have the following nomenclature, place in the mission sequence, and functions. The jettison fairing switch is used at the end of second stage engine thrust decay, by the command pilot to jettison the nose fairing, and the horizon scanner head cover.

The separate spacecraft switch-indicator is used in the separation-insertion phase of the sequence. The command pilot presses the switch-indicator approximately 20 seconds after second stage engine cutoff when the IVI displays the delta-V required for insertion. Pressing the switch-indicator causes several things to happen. Primarily, it detonates pyrotechnic devices which separate the spacecraft from the launch vehicle.

Secondarily, it extends the uhf and diplexer antennas and readies the acquisition aid beacon for use. As the spacecraft moves away from the launch vehicle, separation sensors close and energize the spacecraft separation relays. The relays illuminate the Indicator green. The Indicate retrograde attitude switch-indicator is illuminated amber when the electronics timer energizes the Tr second relay.

The amber light cues the crew to press the switch-indicator at this time. When pressured, a bias voltage is placed on the pitch needle of the FDI, and the inertial platform is electrically placed in the BEF mode. When released the ember light is extinguished and a green light is illuminated.

The battery power indicator illuminated amber by the Tr second relay. This change must be made because the adapter section will be jettisoned at retrograde. When all of the main battery switches are on, the indicator changes from amber to green. The amber light cues the command pilot to activate the RCS firing the fuel, oxidizer, and pressurant isolation squibs. Pressing the switch-indicator energizes relays which fire the squibs.

The indicator changes from ember to green, indicating that the RCS has been activated. The separate OAMS lines indicator is illuminated amber by the Tr second relay is the prepare-to-go to retrograde phase. The amber light cues the crew to seal and sever the OAMS lines before jettisoning the adapter. Pressing the switch-indicator energizes relays which ignite the pyrotechnics used to seal and sever the lines.

The relays also fire pyrotechnic switches and wire guillotines severing some of the adapter-retrograde mating line wiring. The indicator changes from amber to green. The separate electrical indicator is also illuminated amber by the Tr second relay.

Pressing the switch-indicator energizes the wire guillotine relay. The pyrotechnics are detonated and the wiring is cut. The indicator changes from amber to green to indicate that electrical separation has been accomplished.

The separate adapter indicator is illuminated amber by the Tr second relay. The amber light cues the crew to Jettison the adapter equipment section.

Pressing the switch-indicator causes the adapter shaped charge and the Z7O tubing cutter pyrotechnic to be detonated, and the adapter section severed.

Separation of the adapter section is sensed by the equipment adapter separation sensors. Two closed sensors energize the sensor relay and change the indicator from amber to green. The arm automatic retrofire indicator is illuminated amber by the Tr second relay. The amber light cues the crew to arm the automatic retrofire circuits so that when the electronic timer closes the TR contacts at TR time, the retrorockets will fire automatically.

Pressing the switch-indicator completes the patch from the retrograde common control bus to the timer TR contact, and also energizes the TR arm relay. The relay changes the light from amber to green. Contact closure at Tr time energizes the Tr signal relay. The signal relay energizes the second time delay relay, fires the retro rockets at 5. The manual fire retrorockets switch connects the retrograde common control bus to the manual retrograde latch relay.

Contacts of this relay energizes the second time delay relay, fire the retrorockets at 5. The Jettison retrograde adapter indicator is illuminated amber by the second time delay relay 45 seconds after retrofire begins. The amber light cues the crew to jettison the retrograde adapter. It fires the shaped charges which sever the retrograde adapter section from the re-entry vehicle. It energizes the Horizon Sensor System scanner head jettison relays which fire the jettison squibs and Jettison the scanner head.

It removes the retrograde attitude signals applied to the flight director needles at Tr seconds. It switches the FDI roll channel to the mix mode for re-entry. Ten Sequence System relay panels are installed in Gemini Spacecraft 5, 6, and 8 through Four relay panels are located in the re-entry vehicle, three in the retrograde section, one in the equipment section, and two in the rendezvous and recovery section.

The following Sequence System relay panels are in the re-entry module. The necessary functions required for adapter retrograde section separation are performed by the fourteen relays of the retrograde separation relay panel. The relays perform such functions as pyrotechnic switch and shaped charge ignition, Tr-3O second indication, automatic IGS free mode selection, and arming of the contacts of the Time Reference System.

Re-entry Control System squib firing, scanner cover and scanner heads jettison, abort interlock RCS amber light actuation, and RCS ring B squib firing test prior to launch are provided by the sixteen relays of the attitude control system scanner and RCS squib fire relay panel. The umbilical pyrotechnic switch relay panel contains two relays which apply landing squib bus I and 2 power to re-entry umbilical wiring pyrotechnic switch.

The retrograde section contains the following three relay panels which control spacecraft separation, retrofire, and equipment section separation. The equipment section contains the Orbit Attitude Maneuver System squib fire relay panel.

The retrofire relay panel has twenty relays. These relays control the automatic, manual and salvo firing of the retro rockets, and time the 5. The retrograde sequence adapter separate relay panel contains twelve relays.

The relays are used for equipment adapter shaped charge ignition, propellant line guillotine, electrical wire guillotine, and retrograde abort. The Rendezvous and Recovery section contains two Sequence System relay panels: the nose fairing Jettison relay panel, and the docking relay panel. The nose fairing Jettison relay panel contains two relays which control the jettisoning of the nose fairing.

The docking relay panel has eleven relays which extend the docking index bar, illuminate the MSG ACPT light, effect emergency release of the docking latches, release and jettison the locking latches at retrograde, jettison the index bar, and cover the docking latch ports. The Sequence System contains two sets of separation sensors. Separation sensors are toggle switches which are normally open before separation is initiated. The separating structure will close the sensors as it moves away from the spacecraft re-entry module.

The closure of any two of a set of three sensors is sufficient to sense and indicate separation. No primary ac electrical power system Is provided for the spacecraft. Devices requiring ac power obtain this power from self-contained inverters within the individual systems.

The Electrical Power System includes switches, circuit breakers, relay panels, ammeters, a voltmeter and telelights which provide control, distribution and monitoring for the system.

Also included as an Electrical Power System subsystem is the Reactant Supply System RSS which provides storage and control of the reactants hydrogen and oxygen used for fuel cell battery operation not applicable to spacecraft 6. Provisions are made for utilizing external power and remote monitoring of the spacecraft power buses during ground tests and pre-launch operations. The two fuel cell battery sections and four main batteries provide dc power to the spacecraft main power bus on spacecraft 6, the three adapter module batteries and the four main batteries provide dc power to the main bus.

On spacecraft 5 and 6, a dual-vertical-readout ammeter is located on the right instrument panel. On spacecraft 5, two FCAP indicator lamps are located on the right instrument panel. The corrosion data which are presented will apply to storing, handling, and control equipment outside of missiles and to missile components excluding combustion chamber.

The compatibility of materials with reaction products in combustion chambers, nozzles, etc. Included in the summary are data for many nonmetallic materials. The memorandum is subdivided into. The book has proved to be a sought-after and widely used source of reference material to help people avoid or analyse engineering failures, design and manufacture for greater safety and economy, and assess operating, maintenance and fitness-for-purpose procedures.

In the last three years, Engineering Failure Analysis has continued to build on its early success as. Final Handbook Summarized are the physical properties, materials compatibility, handling techniques, flammability and explosivity, and procedures for storing, cleaning, and flushing of the Titan II propellants.