NASA Project Gemini Familiarization Manual

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The Project Gemini Familiarization Manual was a document published by the McDonnell Aircraft Company as a training aid for Gemini astronauts. It is now in the public domain. This a text-only conversion of a PDF document of the August 22, 1966 revision. The images have been removed for space constraints. For now, the first seven of 15 sections are up, with more to follow.

FOREWORD
Initiated by the NASA and implemented by McDonnell Aircraft Corporation, Project Gemini is the second major step in the field of manned space exploration. Closely allied to Project Mercury in concept and utilizing the knowledge gained from the Mercury flights, Project Gemini utilizes a two man spacecraft considerably more sophisticated than its predecessor. The Gemini spacecraft is maneuverable within its orbit and is capable of rendezvous and docking with a second orbiting vehicle.
INTRODUCTION
The purpose Of this manual is to describe the Gemini spacecraft systems and major components. The manual is intended as a femiliarization-indoctrination aid and as a ready reference for detailed information on a specific system or component. The manual is sectionalized by spacecraft systems or major assemblies. Each section is as complete as is practical to minimize the need for cross-referencing.
The information contained in this manual (SEDR 300, VOL XI) is applicable to rendezvous missions only and is accurate as of 1 April 1966.
For information pertaining to long range or modified (non-rendezvous) configurations of the spacecraft, refer to SEDR 300, VOL. I.

Contents 1 Section I: Spacecraft Mission 1.1 MISSION DESCRIPTION 1.2 MISSION OBJECTIVES 1.3 SPACECRAFT DESCRIPTION 1.4 LAUNCH VEHICLE DESCRIPTION 1.5 CREW REQUIREMENTS 1.6 SPACECRAFT RECOVERY 2 Section II: Major Structural Assemblies 2.1 GENERAL INFORMATION 2.2 RE-ENTRY MODULE 2.2.1 RENDEZVOUS AND RECOVERY SECTION 2.2.2 RE-ENTRY CONTROL SYSTEM SECTION' 2.2.3 CABIN 2.3 ADAPTER 2.3.1 RETROGRADE SECTION 2.3.2 EQUIPMENT SECTION 2.4 SPACECRAFT/LAUNCH VEHICLE MATING 3 Section III: Cabin Interior Arrangement 3.1 GENERAL 3.2 CREW SEATING 3.3 SEAT DESCRIPTION 3.4 SEAT EJECTION SYSTEM 3.5 RESTRAINT SYSTEM 3.6 EGRESS KIT (Effective Spacecraft 5 and 6) 3.7 BACKBOARD ASSEMBLY 3.8 PELVIC BLOCK 3.9 BALLUTE SYSTEM 3.10 PERSONNEL PARACHUTE 3.11 PARACHUTE DROGUE MORTAR 3.12 PERS0NNEL HARNESS ASSEMBLY 3.13 SURVIVAL KIT 3.14 PYROTECHNIC DEVICES 3.15 INSTRUMENT PANELS 3.16 CABIN INTERIOR LIGHTING 3.17 STATIC SYSTEM 3.18 FOOD WATER AND EQUIPMENT STOWAGE 3.19 WASTE DISPOSAL 3.20 STOWAGE PROVISIONS 4 Section IV: Sequence System 4.1 SYSTEM DESCRIPTION 4.2 SYSTEM OPERATION 4.2.1 PRE-LAUNCH 4.2.2 LIFT-OFF 4.2.3 BOOST AND STAGING 4.2.4 SEPARATION AND INSERTION 4.2.5 PREPARE-TO-GO-RETROGRADE 4.2.6 RETROGRADE MINUS 3O SECONDS 4.2.7 RETROGRADE SEQUENCE 4.2.8 RE-ENTRY 4.2.9 ABORT MODES 4.2.10 ABORT SEQUENCE 4.3 SYSTEM UNITS 4.3.1 LEFT SWITCH/ CIRCUIT BREAKER PANEL 4.3.2 B00ST-INSERT-ABORT CORTROLS AND INDICATORS 4.3.3 SEQUENCE CONTROLS AND INDICATORS 4.3.4 RE-ENTRY VEHICLE RELAY PANELS 4.3.5 ADAPTER RELAY PANELS 4.3.6 SEPERATION SENSORS 5 Section V: Electrical Power System 5.1 SYSTEM DESCRIPTION 5.2 SYSTEM OPERATION 5.2.1 PRE-LAUNCH 5.2.2 ORBIT 5.3 RE-ENTRY 5.3.1 MONITOR AND DISPLAY 5.4 SYSTEMS UNITS 5.4.1 SILVER ZINC BATTERIES 5.4.2 POWER SYSTEM RELAY PANEL 5.4.3 POWER DISTRIBUTION RELAY PANEL 5.4.4 ADAPTER POWER SUPPLY RELAY PANEL 5.4.5 AMMETERS 5.4.6 VOLTMETER 5.4.7 POWER SYSTEM MONITOR 5.5 FUEL CELL BATTERIES 5.5.1 REACTANT SUPPLY SYSTEM 6 Section VI: Environmental Control System 6.1 SYSTEM DESCRIPTION 6.1.1 OXYGEN SUPPLY SYSTEM 6.1.2 CABIN LOOP 6.1.3 SUIT LOOP 6.1.4 WATER MANAGEMENT SYSTEM 6.1.5 SYSTEM DISPLAYS AND CONTROLS 6.2 SYSTEMS OPERATION 6.3 SERVICE AND CHECKOUT 6.4 PRE-LAUNCH 6.5 LAUNCH 6.6 ORBIT 6.7 RE-ENTRY 6.8 POST-LANDING 6.9 EMERGENCY 6.10 SYSTEM UNITS 6.10.1 DUAL SECONDARY OXYGEN RATE AND SUIT SYSTEM SHUTOFF VALVE 6.10.2 SUIT OXYGEN DEMAND REGULATOR 6.10.3 CABIN PRESSURE RELIEF VALVE 6.10.4 SUIT CIRCUIT COMPRESSOR 6.10.5 SOLIDS TRAP 6.10.6 DUAL CABIN PRESSURE REGULATOR 6.10.7 PRIMARY SUPERCRITICAL OXYGEN CONTAINER 6.10.8 SECONDARY OXYGEN CONTAINER 7 Section VII: Cooling System 7.1 SYSTEM DESCRIPTION 7.1.1 SYSTEM DISPLAYS AND CONTROLS 7.2 SYSTEM OPERATION 7.3 PRE-LAUNCH 7.3.1 LAUNCH 7.3.2 ORBIT 7.4 SYSTEM UNITS 7.4.1 PUMP PACKAGE 7.4.2 RADIATOR 7.4.3 COLDPLATES 7.4.4 HEAT EXCHANGERS 7.4.5 TEMPERATURE CONTROL VALVE 7.4.6 LAUNCH COOLING HEAT EXCHANGER 8 Section VIII: Guidance and Control System 9 Section IX: Communication System 9.1 SYSTEM DESCRIPTION 9.1.1 ANTENNAS 9.1.2 BEACONS 9.1.3 VOICE COMMUNICATION 9.1.4 TELEMTRY TRANSMITTERS 9.1.5 FLASHING RECOVERY LIGHT 9.1.6 DIGITAL COMMAND SYSTEM 9.2 SYSTEM OPERATION 9.2.1 VOICE TAPE RECORDER 9.2.2 PRE-LAUNCH 9.2.3 SPACECRAFT/LAUNCH VEHICLE SEPERATION 9.2.4 ORBIT 9.2.5 ADAPTER SEPERATION 9.2.6 RE-ENTRY 9.2.7 LANDING THROUGH RECOVERY 9.3 SYSTEM UNITS 9.3.1 ANTENNAS 9.3.2 BEACONS 9.3.3 VOICE COMMUNICATION 9.3.4 TELEMTRY TRANSMITTERS 9.3.5 FLASHING RECOVERY LIGHT AND POWER SUPPLY 9.3.6 DIGITAL COMMAND SYSTEM 10 Section X: Instrumentation System 10.1 SYSTEM DESCRIPTION 10.2 SYSTEM OPERATION 10.2.1 SEQUENTIAL SYSTEM PARAMETERS 10.2.2 ELECTRICAL POWER SYSTEM PARAMETERS 10.2.3 ENVIROMENTAL CONTROL SYSTEM PARAMETERS

Section I: Spacecraft Mission

MISSION DESCRIPTION

Fundamentally, the mission of Project Gemini is the insertion of a two man spacecraft into a semi-permanent orbit about the earth, the study of man's ability to rendezvous and dock with another orbiting vehicle, and the subsequent safe return of the spacecraft and its occupants to the earths surface. Previous missions included manned and unmanned flights to study human capabilities during extended missions in space. Rendezvous and docking with an orbiting Agena Target Vehicle or Augmented Target Docking Adapter and Extra-Vehicular Activities are planned for most missions.

MISSION OBJECTIVES

Specifically, the project will seek to:

• Demonstrate the ability of the spacecraft to perform in manual and/or automatic modes of operation.
• Evaluate the adequacy of major systems in the spacecraft.
• Verify the functional relationships of the major systems and their integration into the spacecraft.
• Determine man's requirements and performance capabilities in a space environment
• Determine man's interface problems, and develop operational techniques for the most efficient use of on-board capabilities.
• Evaluate system performance during rendezvous and docking.
• Demonstrate the ability of the pilots to perform Extra-Vehicular Activities.
• Develop operational techniques required for rendezvousing and docking with another orbiting vehicle.
• Develop controlled re-entry techniques required for landing in a predicted touchdown area.
• Develop operational recovery techniques of both spacecraft and pilots.

SPACECRAFT DESCRIPTION

• GENERAL

The Gemini Spacecraft is a conical structure 19 feet long and weighs approximately 7000 lbs. Basically it consists of a re-entry module and an adapter.

• RE-ENTRY MODULE

The re-entry module consists of the heat shield, the crew and equipment section, Re-entry Control System section and the rendezvous and recovery section. The crew and equipment section contains a pressurized area suitable for human occupation, and a number of non-pressurized compartments for housing equipment. External access doors are provided for equipment compartments. The Re-entry Control System section contains the major Re-entry Control System components. The rendezvous and recovery section contains the rendezvous radar equipment, the drogue parachute and pilot parachute assemblies, and the main parachute assembly. The rendezvous and recovery section is jettisoned after re-entry along with the drogue parachute.

• ADAPTER

The adapter consists of the launch vehicle mating ring, the equipment section and the retrograde section. The launch vehicle mating ring is bolted to the launch vehicle. A portion of the ring remains with the launch vehicle at spacecraft-launch vehicle separation. The equipment section contains major components of the Electrical, Propulsion, and Cooling Systems. The primary oxygen supply for the Environmental Control System is also located in the equipment section. The retrograde section contains the retrograde rockets and some components of the Cooling System.

LAUNCH VEHICLE DESCRIPTION

The vehicle used to launch the Gemini Spacecraft is the Gemini - Titan II, built by the Martin Company. The Titan II is modified structurally and functionally to accept the Gemini adapter and to provide for the interchange of electrical signals.
The Titan II is a two stage launch vehicle 90 feet long and 10 feet in diameter from the thrust chamber to the spacecraft adapter. The first stage is 70 feet long and develops approximately 430,000 pounds of thrust. The second stage is 20 feet long and develops about 100,000 pounds of thrust.
Titan II uses hypergolic (self-igniting when mixed) propellants. Nitrogen tetroxide is the oxidizer and unsymmetrical dimetricaldlmethylhydrazine is the fuel. The propellants can he stored within the launch vehicle indefinitely and ignite automatically when they are mixed in the propulsion chamber. The hypergolic propellants will burn (although at a very rapid rate) rather than explode, which is a significant safety advantage.

CREW REQUIREMENTS

The Gemini Spacecraft utilizes a two-man crew seated side by side. The crew member on the left is referred to as the command pilot and functions as spacecraft commander. The crew member on the right is referred to as the pilot. Crew members are selected from the NASA astronaut group.

SPACECRAFT RECOVERY

The Gemini landing module will make a water landing in a pre-determined area. A task force of ships, planes, and personnel will be standing by for locating and retrieving the spacecraft and crew. In the event an abort or other abnormal occurrence results in the spacecraft landing in a remote location, electronic and visual recovery aids and survival kits are provided in the spacecraft to facilitate spacecraft retrieval and crew survival, respectively.

Section II: Major Structural Assemblies

GENERAL INFORMATION

The Gemini Spacecraft is basically of a conical configuration consisting of a re-entry module and an adapter as the two major assemblies. Spacecraft construction is semimonocoque, utilizing titanium for the primary structure. It is designed to shield the cabin pressure vessel from excessive temperature variations, noise and meteorite penetration.

RE-ENTRY MODULE

The re-entry module is separated into three primary sections which include the Rendezvous and Recovery section (R and R), Re-entry Control System section (RCS) and the cabin section. Also incorporated in the re-entry module is the heat shield which is attached to the cabin, and a nose fairing which is attached to the forward end of the R and R section. The nose fairing is ejected during launch.

RENDEZVOUS AND RECOVERY SECTION

The (R and R) section, the forward section of the spacecraft, is semiconical in shape and is attached to the Re-entry Control System section with twenty-four bolts. Incorporated in this joint is a pyrotechnic device which severs all bolts causing the rendezvous section to separate from the RCS section on signal for parachute deployment. A drogue parachute will assist in the removal of this section. The R and R section utilizes rings, stringers and bulkheads of titanium for its primary structure. The external surface is composed of beryllium shingles, except for the nose fairing. The nose fairing is composed of fiberglass reinforced plastic laminate.

RE-ENTRY CONTROL SYSTEM SECTION'

The RCS section is located between, and mated to, the R and R and cabin sections of the spacecraft. This section is cylindrical in shape and is constructed of an inner titanium alloy cylinder, eight stringers, two rings and eight beryllium shingles for its outer skin. The RCS section is designed to house the fuel and oxidizer tanks, valves, tube assemblies, and thrust chamber assemblies for the RCS.
A parachute adapter assembly is installed on the forward face of the RCS section for attachment of the main parachute.

CABIN

The cabin, similar in shape to a truncated cone, is mated to the RCS section and the adapter. The cabin has an internal pressure vessel shaped to provide an adequate crew station with a proper water flotation attitude.
The shape of the pressure vessel also allows space between it and the outer conical shell for the installation of equipment.
The basic cabin structure consists of a fusion welded titanium frame assembly to which the side panels, small and large pressure bulkheads and hatch sill are seam welded. The side panels, small and large pressure bulkheads are of double skin construction and reinforced by stiffeners spotwelded in place. Two hatches are hinged to the hatch sill for pilot ingress and egress. For heat protection, the outer comical surface is covered with Rene' 41 shingles and an ablative heat shield is attached to the large end of the cabin section.
A spring loaded hoist loop, located near the heat shield between the hatch openings, is erected after landing to facilitate engagement of a hoisting hook for spacecraft retrieval.

• EQUIPMENT BAYS

The equipment bays are located outside the cabin pressure vessel. Two bays are located outboard of the side panels and one bay beneath the pressure vessel floor. The bays are structurally designed for mounting of the equipment.

• DOORS

To enclose the side equipment bays, two structural doors are provided on each side of the cabin. These doors provide access to the components installed in the equipment bays. The main landing gear bays, located below the left and right equipment bays, are each enclosed by one door. The landing gear is not installed but fittings are provided for the attachment of the gear for future spacecraft. 0n the bottom of the cabin, between the landing gear doors, two additional doors are installed. The forward door allows access to the lower equipment compartment and the aft door provides access to the Environmental Control System compartment which is a portion of the pressure vessel.

• HATCHES

Two large structural hatches are incorporated for sealing the cabin ingress or egress openings. The hatches are symmetrically spaced on the top side of the cabin section. Each hatch is manually operated by means of a handle and mechanical latching mechanism. Each is hinged on the outboard side. In an emergency, the hatches are opened in a three sequence operation employing pyrotechnic actuators. When initiated, the actuators simultaneously unlock and open the mechanical latches, open the hatches and supply hot gases to ignite the ejection seat rocket catapults. An external hatch linkage fitting is incorporated to allow a recovery hatch handle to be inserted for opening the hatches from the outside. The recovery hatch handle is stowed on the main parachute adapter assembly located on the forward face of the RCS section. A hatch curtain is stowed along the hinge of each hatch. After water landing, when the hatches are open, the curtains are installed to help prevent water from entering the cabin.

• WINDOWS

Each of the ingress/egress hatches incorporates a visual observation window. Each window consists of an inner and outer glass assembly. The outer assembly is a single flat pane and the inner panel assembly consists of two flat panes. The panes consist of Vycor (96% silica). The panes in the right window are optically ground for better resolution. Each surface of each pane, with the exception of the outer surface of the outer pane, is coated to lessen reflection and glare from cabin lights and to aid in impeding ultraviolet radiation into the cabin compartment.

• HEAT SHIELD

The heat shield is a dish-shaped structure composed of silicone elastomer filled, phenolic impregnated, fiberglass honeycomb. It is an ablative device, 90 inches in diameter with a spherical radius of 144 inches. The shield is designed to protect the re-entry module from extreme thermal conditions during re-entry into the atmosphere. The device is attached to the large diameter end of the cabin structure by 1/4 inch bolts.

• SHINGLES

The external surface of the cabin is made up of beaded shingles of Rene' 41. The R and R and RCS section surfaces are made up of unbeaded shingles of beryllium. The shingles protect the re-entry module structure from excessive heat and provide additional rigidity for the cabin. The shingles are black on the outer surface to control thermal radiation. The inner surface of the beryllium shingles are coated with gold to provide a low emissivity surface.

ADAPTER

The adapter functions to mate the spacecraft to the launch vehicle, to provide for mounting equipment and retrograde rockets , and to serve as a radiator for the spacecraft coolant system. The adapter is a truncated cone-shape, semimonocoque structure consisting of circmuferential aluminum rings, extruded magnesium alloy stringers, and magnesium skin. The extruded stringers are designed in a bulb-tee shape to provide a flow path for the liquid coolant which transfers heat to the adapter skin for radiation to space. The outer surface of the skin is coated with white ceramic type paint and the inner surface is covered with aluminum foil. The inner adapter surfaces of spacecraft 9 through 12 are gold plated. The forward end of the adapter is coupled to the aft end of the re-entry module by utilizing three titanium tension straps.

RETROGRADE SECTION

The retrograde section, the smaller end of the adapter, provides for installation of four retrograde rockets and six Orbital Attitude Maneuvering System thrust chamber assemblies. To provide for the installation of the retrograde rockets, the retrograde section employs an aluminum I-beam support assembly. The I beams are assembled in the form of a cruciform with one retrograde rocket mounted in each quadrant.

EQUIPMENT SECTION

The equipment section is the larger diameter end of the adapter. The section provides hard points for the attachment of structural modules for the OAMS tanks, Environmental Control System primary oxygen supply, fuel cell (batteries on spacecraft 6), coolant, electrical and electronic components, Extra Vehicular Activity (EVA) equipment on spacecraft 9 through 12, and Rendezvous Evaluation Pod on spacecraft 5 only. A honeycomb blast shield is provided above the modules to shield the equipment section and booster dome from excessive heat during retrograde rocket firing under abort conditions. Ten OAMS thrust chamber assemblies are mounted on the large diameter end of the equipment section. A gold deposited fiberglass temperature control cover protects the equipment from solar radiation through the open end of the adapter after separation from the launch vehicle.

SPACECRAFT/LAUNCH VEHICLE MATING

The spacecraft is mated to the Titan II Launch Vehicle with a machined aluminum alloy ring. This ring, 120 inches in diameter, mates with the launch vehicle mating ring. Twenty bolts secure the rings together. To provide for alignment, the launch vehicle incorporates one steel 3/16 inch diameter alignment pin located at TY and four index marks. To separate the spacecraft from the launch vehicle, a pyrotechnic charge is fired, severing the adapter section approximately 1½ inches above the launch vehicle/spacecraft mating point.

Section III: Cabin Interior Arrangement

GENERAL

The equipment within the cabin is arranged to permit the command pilot, seated to the left, and the pilot, seated to the right, to operate the controls and observe displays and instruments in full pressure suits in the restrained or unrestrained position. The cabin air outflow is regulated during launch to establish and maintain a 5.5 psi differential pressure between the cabin and outside ambient condition. The cabin is maintained at a nominal 5.1 psia throughout the flight by a cabin pressure regulator. The cabin equipment basically consists of crew ejection seats, instrument panels and controls, lighting, food, water, waste collection, and miscellaneous equipment.

CREW SEATING

The crew members are seated in the typical command pilot and pilot fashion, faced toward the small end of the re-entry module. The seats are canted 12° outboard and 8° forward to assure separation and to provide required elevation in the event an off the pad ejection is necessitated. Crew seating provisions include scats, restraint mechanisms, seat ejection devices, seat man separator, survival gear, and an egress kit assembly effective spacecraft 5 and 6 only.

SEAT DESCRIPTION

The crew seats are all metal built-up assemblies consisting of a torque box framed seat bucket, channeled backs and arm rests. The seat has lateral and vertical stiffeners, designed for a single moment of thrust. The seat is supported at a single point at the top of the seat back. At this point, the seat bolts to the rocket/catapult. Each seat is supported against fore, aft, and side movement by slide blocks mounted on the seats and retained in tee type rail assemblies attached to the large pressure bulkhead. The seats incorporate a padded contoured headrest to support the pilots helmet. Each seat also incorporates a restraint system, harness release system and a seat/man separator.

SEAT EJECTION SYSTEM

The seat ejection system provides the crew with a means of escaping from the vicinity of the spacecraft in the event of an abort or in an emergency condition during launch or re-entry. Crew member seats are ejected by means of rocket/catapults. Hot gas from each of the hatch actuators is routed to the appropriate seat catapult where dual firing pins strike dual percussion primers, thereby igniting the seat rocket/catapult main charge and ejecting the seats from the spacecraft. Hot gas from the rocket/catapult main charge ignites the sustainer rocket and the rocket provides additional separation from the spacecraft. In the event ejection becomes necessary, after deployment of main landing system parachute and while descending in the two point suspension, it is mandatory that the main landing system parachute be Jettisoned before ejecting from the spacecraft.
The ejection sequence is initiated by manually pulling either ejection control (D-ring) located on the front of the seat buckets. During the launch phase of flight each pilot erects and holds on the D-ring. This action aids in stabilizing the pilots arms and at the same time places them in a position for instant response. The D-rings are normally stowed at the front of the seat and are pinned in a downward position at the front of the seat structure. The safety pin is removed during launch and re-entry and during orbit.

RESTRAINT SYSTEM

Each pilot is restrained in his ejection seat by a restraint system consisting of personal harness, lap belt assembly, shoulder restraint, inertia reel and leg restraint. Other portions of the restraint system are part of the ejection seat. These seat restraints are the arm restraint loops, elbow restraint and foot stirrups. The restraint system provides adequate support and restraint during conditions of maximum acceleration and deceleration.

• INERTIA REEL

The inertia reel is a two position locking device, located on the rear of the backboard. Two straps connect the inertia reel and the personal harness to restrain the pilots forward movement. The inertia reel control handle is located on the front of the left arm rest and has two positions, manual lock and automatic lock. Orbital flight is accomplished with the inertia reel in the automatic lock position. Manual lock position is used during launch and re-entry. The manual lock position prevents the pilots shoulders from moving forward.
To release his shoulders, when the inertia reel is in the manual lock position, the pilot must position the control handle to the automatic position. The automatic lock a11ows the astronaut to move forward slowly a maximum of 18 inches but will lock with a 3 g deceleration. When the automatic lock has engaged, the lock will ratchet and permit movement back into the seat, but will not permit forward movement. The release of the automatic lock is accomplished by cycling the control handle to manual and back to automatic lock.

• ARM RESTRAINT

The arm restraint is a welded, 1/2 inch diameter tube assembly made up in the form of a loop. A loop is installed on each arm rest to retain the pilots arms within the ejection envelope. When the arm restraint loop is not required, it may be swung to the rear and down.

• ELBOW RESTRAINT

An elbow restraint is provided for the command pilot only. It is used to stabilize his forearm during manual re-entry.

• LEG RESTRAINT STRAP

The leg restraint consists of two straps of dacron webbing with a connecting slide buckle. One end of each strap is secured to the seat by round metal eyelets. The left strap of each leg restraint has a metal end assembly that permits the right strap to fold back on itself. Velcro tape on the right strap is used to secure the strap end in position when the strap is drawn tight over the pilots legs. During seat/man separation, the restraint strap eyelets are automatically released from the base of the seat, freeing the restraint strap.

• EJECTION SEAT FOOT STIRRUP

The ejection seat foot stirrups consist of two welded frames attached to the front of the ejection seat. Each stirrup has a short protruding platform with small vertical edges rising along the outboard side. The stirrup is so constructed that the pilots shoe heel will lock in place and prevent forward movement of the foot while the small vertical edges will prevent side movement. During seat ejection, the pilots feet will stay in place.

• LAP BELT

The lap belt is an arrangement of dacron and nylon straps, designed to restrain the pilot in the seat structure. Load carrying straps from the lap belt are fastened to the backboard and seat. The lap belt has a manual quick disconnect and a pyrotechnic release fitting near the center of the pilots lap. The manual quick disconnect can be released with one finger. Lap belt tension is adjusted by sliding excess strap through the pyrotechnic release. During ejection, the lap belt ends attached to the seat structure are released just prior to seat/man separation. During separation, the lap belt remains with the pilot. Five seconds after the backboard drogue mortar fires, the pyrotechnic lap belt release activates and allows the lap belt, backboard and seat to fall free.
A second manual release for the lap belt is also available to the pilot. It is located forward on the right arm rest and is referred to as the ditch control. Releasing the lap belt with the ditch control allows the pilot to egress from the Landing module with the backboard and seat.

EGRESS KIT (Effective Spacecraft 5 and 6)

The egress kit assembly contains the bail out oxygen for an ejected pilot. The egress kit rests in the ejection seat bucket and forms a mounting surface for the egress kit cushion. The egress kit contains an oxygen supply, for breathing and suit pressurization; a composite disconnect, which when separated closes the port and prevents escape of egress oxygen; a relief valve, to prevent pressure build up in the pressure suit; a regulator, to reduce high pressure to a controlled flow of low pressure oxygen, a pressure gage, for visually checking egress oxygen pressure; and connecting lines. Three lanyards are attached between the egress kit and the spacecraft. These lanyards pull release plns to allow the composite disconnect to separate, allow the oxygen to flow through the pressure regulator and allow the relief valve to control the pilots suit pressure. When the drogue mortar deploys the pilot parachute, a 5-second pyrotechnic time delay is initiated and at burn out the egress kit with the backboard is separated from the pilot.

• EGRESS KIT CUSHION (Effective Spacecraft 5 and 6)

The egress kit cushion has a universal type of contour and is attached to the top of the egress kit. The cushion is positioned forward of the pelvic block and up to the ejection control handle access door.

BACKBOARD ASSEMBLY

The backboard assembly is machined aluminum, designed and stressed to retain the inertia reel, ballute, ballute release and deploy mechanism, drogue mortar, parachute and survival kit. A cushion, contoured to the individual pilots body requirements, is positioned on the forward surface of the backboard. The cushion is provided to supply support and comfort to the pilots back. The inertia reel straps and lap belt secures the pilot to the backboard. The backboard accompanies the pilot through seat ejection to parachute deployment. Five seconds after parachute deployment, the backboard with the seat is separated from the pilot.

PELVIC BLOCK

The pelvic block, contoured to the lower torso of each pilot, is positioned between the backboard assembly and the seat. The block supports the pilots lower vertebra and pelvic structure. It remains with the seat structure upon seat/man separation.

BALLUTE SYSTEM

The ballute system consists of a barostat controlled pyrotechnic initiator, combined with a pyrotechnic gas generator, cutters and a packaged ballute. The ballute, located on the back and lower left side of the pilots backboard, is an aluminized nylon fabric enclosed cone. It is inflated by ram air passing through four inlets located symmetrically around the upper periphery. The ballute is connected to the backboard through an 8 inch riser, a 5 foot dual bridle, and by a one inch wide dacron webbing passing through a pyrotechnic actuated cutter. The ballute provides the pilot with a stabilized, feet into the wind, attitude for all ejections over 7,500 feet. The system is fully automatic and is actuated at seat/man separation. At altitudes below 7,500 feet, the barostat prevents deployment of the ballute.

PERSONNEL PARACHUTE

The personnel parachute is a standard 28 ft dia nylon parachute. The parachute is located on the right rear of the pilots backboard. It is deployed by the drogue mortar slug and pilot chute. The parachute risers are attached to the pilots personal harness.

PARACHUTE DROGUE MORTAR

The parachute drogue mortar is a pyrotechnic device designed to eject a 10 oz drogue slug with sufficient velocity to deploy the pilot chute of the personnel parachute. The drogue mortar is a barostat operated firing mechanism, but can be fired manually. It will fire and deploy the parachute at or below 5,700 feet plus a 2.3 seconds time delay from seat/man separation. An MDF chain is initiated by the drogue mortar and separates the backboard and seat from the pilot.

PERS0NNEL HARNESS ASSEMBLY

The personal harness assembly provides a light, strong, and comfortable arrangement to attach the personnel parachute to the pilot. The harness is constructed from nylon webbing formed into a double figure-8. The two figure-8's are joined by two cross straps, the waist strap, and the chest strap. Only the chest strap is adjustable. A quick disconnect is placed forward and below each shoulder for connection of the parachute risers and shoulder restraint straps. Below the left quick disconnect, a small ring is incorporated to attach the survival kit lanyard.

SURVIVAL KIT

The survival kit is a packaged group of specially designed equipment for the use of a downed pilot. Articles in this kit are intended to aid in preserving life under varying environmental conditions. Deployment of the survival kit is automatic if the pilot ejects and is also available to the pilot if he lands with the spacecraft.
Deployment of the survival kit during the ejection cycle takes place as the backboard and seat falls away from the parachuting pilot. As the backboard falls, the survival kit lanyard, connected to the pilots harness, pulls a pin on the life raft container. When the pin is removed, the daisy chain loops are disengaged and the life raft and rucksack are extracted from the container. The survival kit lanyard repeats the extraction process in removing the machete and water bottle from the second container. The machete and water bottle are stowed in a survival equipment container on the left front side of the backboard.
During seat/man separation, a lanyard between the seat structure and the rucksack activates the radio beacon. As the pilot descends on his parachute, the survival equipment is suspended below and the radio beacon transmits on an emergency frequency. Direction finding equipment on aircraft and aboard ship can plot the pilots position taking navigational fixes on the radio/beacon.
Survival equipment is divided into two major stowage containers. The life raft container mounted on the left rear of the backboard has the following items:

• Life Raft Container (Typical)
• 1 Life Raft
• 1 Sea anchor
• 1 4 inch x 4 inch Foam rubber pad
• 1 C02 cylinder
• 1 Sea dye marker
• 1 Sun bonnet
  
• Rucksack (Typical)
• 1 Survival light
• 1 Strobe light
• 1 Flash light
• 4 Fish hooks
• Fish line
• 2 Sewing needles and thread
• 1 Magnetic compass
• 1 Fire starter
• 4 Fire fuel
• 1 Whistle
• 1 Signal mirror
• 14 Water purification tablets
• 1 De-salter kit (less can)
• 8 De-salter tablets
• 1 Water bag
• 1 Repair kit
• 1 Medication kit (Typical)
• 6 Tablet packets
• 1 Small injector (1 CC)
• 1 Large injector (2 CC)
• 1 3 inch x 3 inch compress
• 1 12 inch x 12 inch aluminum foil
• 1 Tube zinc oxide
  
• 1 pr Sun glasses
• 1 Radio beacon

The forward survival kit, mounted on the forward surface of the backboard to the left of the pilots shoulder, contains the following:

• 1 Water container with 3 lb of water
• 1 Machete with sheath

PYROTECHNIC DEVICES

There are 18 pyrotechnic devices incorporated in the cabin all of which pertain to seat ejection, restraint release and parachute deployment. The pyrotechnic devices are 2 hatch actuators, 2 seat rocket/catapults, 2 ballute deployment and release mechanisms, 2 backboard and seat Jettison, 2 drogue mortars, 2 harness release actuators, 2 seat/man separator actuators, 2 hatch actuator initiators and 2 hatch MDF (Mild Detonating Fuse) b6vnesses. The pyrotechnic devices, except the drogue mortar, are saftied by stowing the ejection control handle (D-ring) with a safety pin through the handle into the ejection control assembly. On spacecraft 8 only, a second ejection control pyrotechnic safety pin is also inserted in the side of the ejection control assembly to completely safety the MDF manual firing mechanism.

INSTRUMENT PANELS

Instrument panels, switch and circuit breaker panels and pedestal panels are arranged to place controls and indicators within reach and convenient view of each crew member while in a full pressure suit. A swizzle stick, stowed by the overhead switch and circuit breaker panel, enables a pilot to position switches and rotate selectors on the opposite side of the cabin. With this arrangement, one pilot can control the complete spacecraft and temporarily free the second pilot of all duties.

CABIN INTERIOR LIGHTING

Cabin interior lighting is provided by three types of lights located in five separate locations, described as follows: Cabin flood lights are located aft and above the center-line stowage area. A DIM control is located under the light to control light intensity. Instrument flood lights are located at the forward inner edge of the hatches. Each instrument flood light installation contains two lamps, one lamp having a rod filter and the other a white filter projecting downward. A DIM control and a RED-WHITE-OFF switch are provided at each of the lights.
Two utility lights attached to the ends of spiral extension cords are located on the left and right side walls of the spacecraft interior. The lights stow in clips mounted on the side walls. An ON-OFF switch is located adjacent to the AUX RECEP panel on each of the spacecraft side walls. The CTR LIGHTS, BRIGHT-0FF-DIM switch and the CABIN LIGHTS switch-circuit breaker are located on the overhead switch and circuit breaker panel.

• ELECTRICAL OUTLETS

The two receptacles, powered by the spacecraft electrical system, are installed on brackets immediately aft of the left and right switch/circuit breaker panels. These receptacles are controlled by adjacent ON-OFF switches and are used for powering the utility light or other electrical equipment.

STATIC SYSTEM

The static pressure system is employed to operate the rate of descent indicator, altimeter, and to supply pressure to the static pressure transducer for instrumentation. The static system is also utilized to provide a differential pressure for the cabin pressure transducer. The static ports used for atmospheric pressure pick-up, are located in the small end of the spacecraft conical section. The static port, used for differential pressure pick-up, is located on the forward surface of the small pressure bulkhead.

FOOD WATER AND EQUIPMENT STOWAGE

Containers to left, right and aft of pilots are provided for equipment and food storage. Although minor changes in storage containers are dictated by mission requirements, the main containers are as follows: Center-line stowage box, used for larger size camera containers and EVA (Extra-Vehicular Activity) chest pack; right aft pressurized stowage box, used to stow food initially and later, body waste materials; left aft stowage box, used to stow food packages; right and left sidewall stowage boxes, used to stow small pieces of equipment; left and right fabric covered sidewall stowage boxes, used to stow lightweight head sets; hatch food pouches used to stow large quantities of food; and sidewall stowage box extensions used to stow penlight, spotmeter, exposure dial and tape recorder cartridges. Equipment stowed in the above boxes may change with each mission.
Larger pieces of equipment, emergency equipment or equipment used on every flight, have special stowage brackets or fabric pouches positioned throughout the interior of the spacecraft. Examples of specific stowage brackets are as follows : in-flight medical kit, stowed aft of abort control handle; and the optical sight, stowed under command pilots instrument panel. Without counting the food packages, stowage facilities are furnished for more than 125 pieces of equipment.
During flight, various pieces of frequently used equipment are removed from launch stowage areas and are stowed, with Velcro tape, on the spacecraft sidewalls, and on the inside surfaces of the hatch. As debris accumulates during flight, it is placed in the left aft debris area, located aft of the pilots seat. Prior to descent, the equipment is re-stowed. Only a general rule can be applied to stowage descriptions. Exposed film is placed in insulated containers, previously occupied by cameras and lens, in the center line stowage box. The left aft stowage box is filled and the remainder of the loose equipment is divided among the sidewall stowage boxes on a planned basis. The pressurized stowage box is used to store urine samples and waste containers.
A water storage container, with a 16-pound capacity, is located forward of the aft pressure bulkhead, between the seats. As the water is used from the main storage container, it is replenished by the water stowed in the adapter section. Drinking is accomplished by means of a tube and manual valve system. Food and water will be sufficient for the mission and a postlanding period of 48 hours.

WASTE DISPOSAL

Feces will be collected in a glove-like plastic bag. Urine samples are taken, and the remainder disposed of by overboard dumping. The urine samples and feces waste containers are stowed in the right aft pressurized container which allows cabin depressurization without possible boiling off of the waste materials moisture content.

STOWAGE PROVISIONS

Personal stowage facilities are provided by retaining removed portions of the pressure suit and other equipment as required. These provisions consist of floor pouches, Velcro covered areas on the walls of the pressure vessel, adjacent to the pilots and attached to the structure in usable areas. Items to be stowed utilize the hook and pile principle of mating Velcro patches.

Section IV: Sequence System

SYSTEM DESCRIPTION

The Sequence System of Gemini Spacecraft 5, 6, and 8 through 12 comprises those controls, indicators, relays, sensors and timing devices which provide semiautomatic control of the spacecraft and/or launch vehicle during the critical control times, but which are not part of other systems. The critical times are: the time from booster engine ignition through insertion into orbit; the time to prepare to go to retrograde through post-landing; and the time to abort.
The Gemini crew does not control the spacecraft during boost through Second Stage Engine Cutoff (SSEC0). The spacecraft is controlled by Radio Guidance System (RGS) and the Digital Command System (DCS), or by the Inertial Guidance System (IGS) and the on-board computer. The crew does however, monitor certain indicators to keep informed of the operation of the launch vehicle, to anticipate a crisis if one should develops, and to know if and when mission abort is mandatory. After SSEC0 the command pilot takes necessary action to separate the spacecraft from the launch vehicle and applies final thrust to place the spacecraft in the desired orbit.
During orbit, the Sequence System is in standby. The electronic timer, however, which is part of the Time Reference System, is counting down the time-to-go to retrograde.
At 4 minutes and 16 seconds before retrograde, (Tr-256 seconds), a Sequence System relay is actuated, and several Sequence System indicators illuminate amber. These Indicators provide the crew with cues for necessary operations. Again at 30 seconds before retrograde, the crew is reminded to separate the adapter equipment and arm the automatic retrograde rocket firing circuits. The Sequence System, if properly armed, will initiate retrograde automatically. The crew redundantly initiates retrograde manually as a safety precaution. During descent, altitude indicators illuminate as cues to deploy parachutes. After splash down, the main parachute is jettisoned, and all systems are shutdown.
Four abort modes comprise the abort sequence. They are: seat ejection (mode I); ride-it-out abort (mode I-II); modified re-entry (mode II); and normal re-entry (mode III). The mode selected for abort is related to the spacecraft altitude at the time the abort command is given.

SYSTEM OPERATION

To simplify explanation, the Sequence System is divided into eight stages. The eight stages are; pre-launch, lift-off, boost and staging, separation and Insertion, prepare-to-go to retrograde, retrograde, re-entry, and abort. Telemetry guidance, landing and post-landing are related to but not part of the Sequence System.
Pre-launch, lift-off, boost and staging, and separation and insertion are explained first. Prepare-to-go to retrograde, retrograde, and re-entry are discussed next. Abort is discussed last.

PRE-LAUNCH

The command pilot and the pilot ingress the Gemini cabin and take their assigned crew stations. The hatches are closed and locked. The crew checks that both D-rings are unstowed. The command pilot makes sure that the abort control handle is in the NORMAL position; the maneuver controller is stowed; the altimeter is set; and the Incremental Velocity Indicator (M) is zeroed. He verifies that the nine sequence indicators, the two ABORT indicator lights, the ATT RATE indicator light, the SEC GUIDANCE indicator light, both ENGINE I indicator lights, and the ENGINE II indicator light are extinguished. He places the top three rows of circuit breakers on the left switch/circuit breaker panel to the closed (up) position. He places the BOOST-INSERT and RETRO ROCKET SQUIB switches in the bottom bow to ABM, and the RETRO and LANDING switches to SAFE. He tests the nine sequence indicators with the SEQ LIGHTS TST switch. He selects switches for gyro run-up and platform alignment, and performs on-board computer checkout.
The pilot places the four MAIN BATTERIES switches and the three SQUIB BATTERIES switches to ON. Both pilots select and check their intercom and uhf communications. The remaining controls and indicators are also monitored or positioned as required. The crew verifies and reports all systems ready for launch.

LIFT-OFF

When the pre-launch countdown reaches zero, the first stage engine ignition signal is given from the blockhouse. Both first stage engines begin thrust chamber pressure buildup. Both ENGINE I indicators illuminate red but extinguish in about one second. When the thrust chamber pressure of these two engines exceeds 77 percent of rated pressure, a two-second time delay is initiated in the blockhouse. If all systems remain go during this delay, the hold-down-bolt fire command is given and the launch vehicle is committed to flight. First motion sensors detect vehicle ascent one and one-half inches off the pad, and energize time-zero relays in the blockhouse and in the spacecraft. A l.5-second shutdown arm time delay is initiated to prevent accidental booster engine shutdown prior to the scheduled staging time. The umbilical release command is given, disconnecting the adapter, and re-entry umbilicals. The on-hoard computer is switched from the guidance inhibit mode to the guidance initiate mode and enabled to accept acceleration data. The lift-off signal is also applied to the electronic timer the event timer. The electronic timer begins to count down the time-to-go to retrograde. The event timer begins to count up the time from lift-off.

BOOST AND STAGING

As the missile continues to climb, the crew monitor the boost sequence and ABORT Indicators. The two ENGINE I under pressure Indicators, the ATT RATE Indicator and both ABORT indicators must remain extinguished. The ENGINE II indicator illuminates amber. The STAGE I FUEL and OXIDIZER needles must indicate pressures within the required limits, and the LONGITUDANAL ACCELEROMETER must indicate an increasing acceleration within prescribed limits for the flight time indicated by the event timer. The pilots monitor their indicators and report via uhf link to the ground. Abort mode I prevails during the first 50 seconds of flight. Ground stations notify the pilot when abort mode I is no longer applicable and when abort mode I - II becomes applicable. Abort mode I-II is in effect during approximately the next 45 seconds of flight. At T+95 seconds, the crew receives and acknowledges changeover to abort mode II.
At T+145 seconds, when the acceleration has climbed to nearly 6g's, the first stage engine shutdown arm relays are energized. At approximately T+153 seconds, the thrust chamber pressure drops to less than 68 percent. The two ENGINE I indicators illuminate red, and the staging control relays are energized. The staging switches are closed. The stage I shutdown solenoids energize and both engines are shutdown. Acceleration drops sharply to approximately l.5 g's. The booster sequential system immediately ignites the second stage engine. The explosive bolts which unite stage i and stage 2 are detonated, and the stages separate. Both ENGINE I indicators are extinguished. Fuel injector pressure of the second stage engine rapidly increases above 55 percent, extinguishing the ENGINE II underpressure indicator. The LONGITUDINIAL ACCELEROMETER begins to climb slowly. The crew reports the results of the staging sequence to the ground station.
The ENGINE II underpressure indicator, the Attitude Overrate (ATT RATE) indicator, and the two ABORT indicators must remain extinguished. The STAGE 2 FUEL and OXIDIZER needles must indicate the required pressures, and the LONGITUDINAL ACCELEROMER must show the required increase.
At approximately T+310 seconds, the spacecraft has climbed above 522,000 feet and its velocity exceeds 80 percent of orbital velocity. The ground station notifies the crew that abort mode III now replaces abort mode II. Both pilots acknowledge the change of abort modes.

SEPARATION AND INSERTION

At T+330 seconds, the acceleration has climbed to almost 7g's, and the spacecraft has nearly reached orbital velocity and altitude. Approximately 337 seconds after lift-off, the blockhouse computer transmits the SSECO command tones via the Digital Command System to the launch vehicle. The SSECO solenoids energize SSECO occurs, thrust decays, and acceleration falls rapidly. The on-board computer begins to compute the delta-V required for insertion.
The command pilot waits 20 seconds for launch vehicle thrust to decay. Near the end of the thrust decay period, the command pilot depresses and releases the JETT FAIRING switch on the main instrument panel. This switch energizes nose fairing Jettison relays K3-13 and K3-17 and scanner cover jettison relays K3-18 and K3-19. These jettison relays arm the nose fairing squibs and scanner cover squibs. The squibs detonate explosive charges, which jettison the fairing and cover.
When thrust decay is complete, the command pilot, depresses and releases the SEP SPCFT switch-indicator on the mean instrument panel. When the contacts of the SEP SPCFT switch-indicator closes, squib bus number 1 power is applied through the closed BOOST-INSERT CONT 1 circuit breaker to relays K3-22, K3-24, and K3-42. K3-22 is the spacecraft shaped charge ignition relay. K3-24 is the launch vehicle/spacecraft wire guillotine relay. K3-42 is the uhf whip antenna extend relay. Redundant contacts of the SEP SPCFT switch-indicator energize redundant relays with power from squib bus number 2.
Time delays in the relays and pyrotechnics cause the separation events to occur in the following sequence. K3-24, contacts C energize the launch vehicle/spacecraft pyrotechnic switch relay K3-26. K3-26, contacts C immediately fire the pyrotechnic switch, open-circuiting the wires on the battery side of the guillotine. Next the wire guillotines are fired, severing the launch vehicle spacecraft wires at the interface. Finally the spacecraft shaped charges are ignited, breaking the structural bond between the launch vehicle and the spacecraft. The operation of all pyrotechnics mentioned in this section is explained in Section XI.
The launch vehicle may now separate from the spacecraft, or thrust from the Orbit Attitude and Maneuver System (OAMS) may be required to effect separation. When two inches of separation exist at the interface, the spacecraft separation sensors close. The spacecraft separation sensor relay K3-28 is energized when two of the three sensor switches are actuated. Contacts A of K3-28 apply main bus power through the closed SEQ LIGHTS PWR circuit and the SEQ LIGHTS BRIGHT-DIM switch to the switch-indicators. The SEP SPCFT switch-indicator illuminates green.
The command pilot observes the delta-V required for insertion which is now displayed on the IVI. He fires the aft thrusters until the IVI is nulled. The spacecraft is in the required orbit. The crew places the following switches to these positions: RETRO ROCKET SQUIB to SAFE, BOOST-INSERT SQUIB to SAFE, and MAIN BATTERIES 1, 2, 3 and 4 to OFF. For the communication switches positioned at this time, refer to Section IX.
The launch vehicle may now separate from the spacecraft, or thrust from the Orbit Attitude and Maneuver System (OAMS) may be required to effect separation. When two inches of separation exist at the interface, the spacecraft separation sensors close. The spacecraft separation sensor relay K3-28 is energized when two of the three sensor switches are actuated. Contacts A of K3-28 apply main bus power through the closed SEQ LIGHTS PWR circuit breaker and the SEQ LIGHTS BRIGHT-DIM switch to the switch-indicators. The SEP SPCFT switch-indicator illuminates green.

PREPARE-TO-GO-RETROGRADE

Approximately 30 minutes before retrofire time, the crew places the C-band beacon switch to CONT and performs platform alignment procedures. Then maneuver the spacecraft to the Blunt End Forward (BEF) position.
At Tr -256 seconds (4 minutes and 16 seconds before retrofire time), the electronic timer energizes the TR-256 second relay K8-16. The A contacts of K8-16 close and energize K8-17, K8-19 and K8-29. K8-17 is the Electrical Power System TR-256 relay, and its A contacts now close to illuminate the BTRY PWR indicator amber. K8-19 is the Re-entry Control System (RCS) amber light relay, and illuminates the RCS indicator amber. K8-29 is the indicate retrograde attitude relay, and illuminates the IND RETRO ATT indicator amber.
The amber BTRY PWR indicator reminds the pilot to turn on the main batteries by placing the four MAIN BATTERIES switches to the ON position. Relay K1-29 is energized through the ON position of the four battery switches. The BTRY PWR indicator illuminates green. Depressing the amber IND RETRO ATT switch-indicator energizes the retrograde bias relay K12-5. K12-5 extinguishes the amber lamp and illuminates the green lamp of the indicator. K12-5 also applies the retrograde attitude bias voltage to the Flight Director Indicator (FDI), and electrically places the inertial platform in the BEF mode. The FDI needles can now be used to orient the spacecraft in this attitude.
Depressing the RCS switch-indicator energizes the four RCS squib fire relays K11-7, K11-8, K11-9, and K11-IO. Relays K11-7 and K11-8 are energized from retrograde bus number 1 while K11-9, and K11-i0 are energized from retrograde bus number 2. When any of the four RCS squib fire relays energize, the RCS auxiliary relay K11-5 is latched, changing the RCS indicator from an amber to a green indication. Relays K11-7 and K11-9 both fire the package A, C, D, pressure isolation, oxidizer isolation, and fuel isolation squibs of ring B. Relays K11-8 and K11-1O fire the package A, C, D, pressure isolation, oxidizer isolation, and fuel isolation squibs of ring A. The RCS RING A and RING B switches are now placed to ACME, and the attitude controller is operated to fire and test the RCS thrusters.
02 high rate flow is initiated after the TR-256 second sequences at the option of the crew. When the CABIN FAN switch is placed to the 02 HI RATE position, the disconnect relay K7-3 is energized. K7-3 removes power from the cabin fan power supply and the two suit power supplies, and illuminates the amber 02 HI RATE indicator.
After the TR-256 sequence, re-entry communications are selected, as discussed in Section IX.

RETROGRADE MINUS 3O SECONDS

Thirty seconds prior to retrograde (Tr-30 seconds), the electronic timer initiates a contact closure. This closure energizes the retrograde TR-30 seconds relay K4-46, which illuminates the SEP OAMS LINE, SEP ELEC, SEP ADAPT, and ARM AUTO RETRO indicators amber.
As soon as the command pilot observes that the four indicators have illuminated amber, he depresses and releases the SEP 0AMS LINE switch-indicator. This switch closure energizes the OAMS propellant line guillotine relay K4-23 and the retrograde abort pyrotechnic squib relay K4-30. K4-23 changes the SEP OAMS LINE indication from amber to green, fires the OAMS propellant lines guillotine igniter 1-1, and then energizes pyrotechnic switch relays K4-25 and K4-26. Relay K4-25 and K4-26 energize pyrotechnic switches B, C, D, E, F and J.
Next, the command pilot depresses and releases the SEP ELEC switch-indicator which energizes wire guillotine relay K4-2. K4-2 ignites wire guillotine C, D and E and energizes the separate electrical latch relay K4-64. When K4-64 energizes, the SEP ELEC switch-indicator changes from amber to green. Then, the command pilot initiates the equipment adapter separation sequence by depressing and releasing the SEP ADAPT switch-indicator. Closure of the SEP ADAPT switch energizes the adapter shaped charge relay K4-3 and abort discrete relay K4-66. K4-3 detonates shaped charge igniter 2-1 and 3-1. The adapter equipment section separates, and separation is sensed by three toggle sensor switches. The switches close when the physical separation is one and one half inches. The closure of any two switches energizes the adapter separate sensor relay K4-15. K4-15 changes the SEP ADAPT switch-indicator from amber to green. The green SEP ADAPT light informs the crew that the adapter equipment section has been jettisoned from the spacecraft. K4-66 sends the abort transfer discrete to the on-board computer.
Lastly, the command pilot depresses and releases the ARM AUTO RETRO switch-indicator. The ARM AUTO RETRO switch latches the TR arm relay KM-36. This relay changes the indication from amber to green and arms the electronic timer for the TR relay contact closure. The four RETRO ROCKET SQUIB switches are now moved to the ARM position.

RETROGRADE SEQUENCE

The retrograde sequence is initiated by the TR signal from the electronic timer. The redundant sequence is initiated manually by the crew.
At Retrograde (Tr), the electronic timer latches the TR signal relay K4-34. The TR signal relay in the latched condition energizes the retrorocket automatic fire relay K4-7. K4-34 also energizes the 45-second time delay relay K4-4, initiates a 5.5-seconds, 11.0-seconds, and a 16.5-second time delay, and deactivates the IGS platform free mode. The retrorocket automatic fire relay redundantly fires retrorocket number 1 from retrograde squib bus number 1 and number 2. At the end of the 5.5-second time delay, the retrorocket automatic fire relay K4-9 is energized. K4-9 ignites retrorocket number 3 from retrograde squib bus number l and number 2. Retrorocket number 2 is redundantly ignited from retrograde squib buss number 1 and number 2 when the retrorocket automatic fire relay K4-31 energizes at the end of the 11.0-second time delay. Retrorocket automatic fire relay K4-13 is energized at the end of the 16.5-second time delay. K4-23 redundantly fires retrorocket number 4 from retrograde squib bus number 1 and number 2.
In order to assure retrograde rocket ignition, the command pilot initiates manual retrograde ignition by depressing and releasing the MAN FIRE RETRO switch-indicator approximately one second after automatic retrofire initiation. The MAN FIRE RETR0 switch latches the manual retrograde latch relay K4-37, energizes retrorocket manual fire relay K4-8, and initiates the 45-second time delay relay K4-6. This switch also initiates the 5.5-second, 11-second and 16.5-second time delays. The 5.5, 11 and 16.5-second time delays energize retrorocket manual fire relays K4-10, K4-12 and K4-14 respectively, which in turn fire retrorockets number 3, number 2, and number 4 respectively. Retrorocket number 1 is fired by K4-8. 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 command pilot moves the JETT RETRO SQUIB ARM switch on the left switch circuit breaker panel from SAFE to ARM. Forty-five seconds after retrograde ignition, K4-4 or K4-6 energizes and illuminates the JETT RETRO lamp on the main instrument panel.
As soon as the command pilot observes the JETT RETRO indicator is amber, he depresses and releases this switch-indicator. The switch energizes the retrograde separate shaped charge relay K4-17, the retrograde bias off relay K4-62, and the horizon scanner heads Jettison relay K4-38. Relay K4-I7 fires retrograde adapter shaped charge igniter 1-1, 2-1, and 3-1 and pyrotechnic switch H-1. Relay K4-62 latches the re-entry roll display relay K12-6 removing roll mix interlock from the flight director controller. K4-62 also resets two latch relays: the retrograde bias relay K-12-5 and the indicate retrograde attitude relay K8-29. Relay K8-29 extinguishes the IRD RETRO ATT indicator. K4-18 fires horizon scanner cover squib 1-1 if it was not fired previously during the boost phase. K4-38 ignites the horizon scanner head squib 1-1 through an 80-millisecond pyrotechnic time delay and jettisons the scanner head. The firing of pyrotechnic switch H-1 extinguishes the SEP ELEC, SEP ADAPT, SEP OAMS, ARM ADTO RETRO and JETT RETRO indicators.
On spacecraft 6 and 8 through 12, the JETT RETRO switch also energizes latch release relay K4-69 through the B contacts of the nose fairing jettison latch relay K3-86. K4-69 fires the release igniters of docking latches 1, 2 and 3 to jettison them. K4-69 also energizes the index bar Jettison and latch door release relay K4-73. KM-73 fires three latch door cover release igniters. These igniters release the latch doors which cover the ports left by the jettisoned docking latches. K4-73 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.

RE-ENTRY

After the retrograde adapter and horizon scanner heads have been jettisoned, the command pilot places the RETR0 PWR and RETRO JETT squib switches to SAFE. Using the attitude controller and the FDI needles, he rolls the spacecraft 180 degrees so that the horizon is visible in the upper portion of his cabin window. He changes the ATTITUDE CONTROL mode select switch on the main instrument panel from PULSE to RATE CMD (RE-ENT). 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 100,000 feet. As this altitude the FDI roll needles start to move, the computer START light illuminates, and the computer begins to calculate the point of impact. The command pilot changes the ATTITUDE CONTROL mode select switch from RATE CMD (RE-ENT) to RE-ENT. 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 100,000 feet. At this altitude, the altimeter indicator begins to come off the peg. At 80,000 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.

ABORT MODES

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. The ABORT indicator may be illuminated by three different methods. 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 15,000 feet, 75,000 feet, and 522,000 feet. The spacecraft reaches 15,000 feet approximately 50 seconds after lift-off, 75,000 feet approximately I00 seconds after lift-off, and 552,000 feet approximately 310 seconds after lift-off. Below 15,000 feet, seat ejection (mode I) is used. Between 15,000 and 75,000 feet, ride-it-out abort (mode I-II) is used. Between 75,000 and 522,000 feet, modified re-entry (mode II) is used. Above 522,000 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.

• ABORT MODE I

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. The command lights both ABORT indicators on the command pilot and pilot's panels. 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 15,000 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 two ENGINE I indicators and both ABORT indicators illuminate red. 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

Abort mode I - II is the ride-it-out abort mode. It is effective at altitudes between 15,000 and 75,000 feet approximately 50 seconds to 100 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. If an abort condition develops, and the crew elect to ride it out, the command pilot moves the abort control handle from NORMAL to SHUTDOWN. He waits 5 seconds for booster thrust to decay, then moves the handle from SHUTD0WN to ABORT.
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. The relays ignite the pyrotechnics at the equipment adapter/retrograde adapter mating line, and the vehicles separate. Then, the four retrorockets are salvo fired and the spacecraft thrusts away from the launch vehicle.
If the abort altitude is between 15,000 and 25,000 feet the retrograde adapter is jettisoned 7 seconds after retrorocket salvo fire is initiated. If the abort altitude is between 25,000 and 75,000 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 40,000 feet, the drogue parachute is deployed at 40,000 feet, and the main parachute at 10,600 feet. If the drogue parachute fails or has not been deployed before the spacecraft descents to 10,600 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 4O,000 feet, the pilots may elect to remain with the spacecraft until the operating engine has boosted them to 75,000 feet. At this altitude, abort mode I-II becomes inapplicable.

• ABORT MODE II

Abort mode II becomes effective above 75,000 feet. At approximately 100 seconds after lift-off on a normal mission, the launch vehicle has boosted the spacecraft to an altitude of 75,000 feet. Ground station computers calculate the time for changeover from abort mode I-II to abort mode II. 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 75,000 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. If a ground station sends the command to abort, both ABORT indicators illuminate red. In abort mode II, the command pilot must act. He moves the abort handle to the SHUTDOWN position. The operating engine is cutoff. Since launch vehicle destruct is imminent and escape from the fireball is urgent, he moves the ABORT handle to ABORT. The spacecraft is separated the launch vehicle at the equipment adapter/retrograde adapter mating line. The retrorockets, armed by four RETRO ROCKET SQUIB switches during pre-launch check-off, are salvo fired, propelling the spacecraft away from the launch vehicle.
Since orbital velocity could not have been reached below 522,000 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.

• Abort Mode III

At approximately 310 seconds after lift-off, the launch vehicle reaches the altitude of 522,000 feet and a velocity of approximately 21,000 feet per second. The ground station commands a change from abort mode II to abort mode III via the uhf link.
If an abort after this time should become necessary, the ABORT indicators would be illuminated red. The command pilot responds and moves the ABORT handle to the SHUTDOWN position. The shutdown command is thus given to the second stage engine. The ABORT handle remains in the SHUT0WN position. The command pilot then presses the SEP SPCFT switch-indicator on the main instrument panel. This switch fires the shaped charges and severs the wiring at the launch vehicle/spacecraft mating line as described earlier. 0AMS thrust is applied to put distance between the second stage and the spacecraft. The crew perform the Tr-256 seconds and the Tr-30 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 SEQUENCE

The abort sequence described herein occurs during abort modes II and I-II. Abort mode I, the seat ejection mode, is not covered here. The events of this mode are discussed in Section III of this Manual.
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.

• SHUTDOWN

When the command pilot moves the abort control handle to SHUTDOWN, the SHUTDOWN switch is closed. BIA common control bus power is applied to the launch vehicle engine shutdown signal relays K3-28 and K3-49. This power is also applied to the engine shutdown relays in the Titan Launch Vehicle. The operating engine(s) are cut off. As K3-48 and K3-49 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.

• ABORT INITIATE

When the command pilots moves the abort control handle to ABORT, numerous relays are energized. 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, K3-92; (2) the squib bus abort relay K3-38; (3) the Attitude Control System abort relay K3-59; (4) the retrograde abort relay K3-36; and (5) the salvo retrograde relay K3-7I.

• ABORT TELEMETRY

When the instrumentation abort relay K3-92 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 SQUIB BUS ARMING

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 K3-38. K3-38 arms the retrograde squib buses i and 2 and the retrograde common control bus.

• RE-ENTRY CONTROL SYSTEM (RCS) ACTIVATION

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.
In operation, the abort switch applies BIA squib bus power to the Attitude Control System abort relay K3-59. K3-59 applies retrograde squib bus power to RCS (ring A) squib fire relay KII-8 and the RCS (ring B) squib fire relay Kll-7. KII-8 applies retrograde squib bus power to package A, C, and D igniters of RCS ring A. The squibs thus fired open the ring A fuel and oxidizer lines and pressurize them. K11-7 applies retrograde squib bus power to similar igniter of RCS ring B with similar results.
The B contact of F13-7 and K11-8 energize the retrograde abort interlock relay K11-22. K11-25, contact A initiates the station 7.70 separation sequence.

• OAMS LINES AND LOWER WIRES GUILLOTINE

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 K3-36 which arms K4-23, the OAMS lines guillotine latch relay; K4-30, the retrograde abort pyrotechnic switch relay; and K4-74, the wire guillotine relay. When K11-25 is energized, it energizes K4-23, K4-30, and K4-74. The D contacts of K4-23 apply power to the OAMS propellant lines guillotine igniter. The guillotine now seals and cuts the lines. Pyrotechnic switch G fires, opening the launch vehicle/spacecraft interface circuits. The lower wire bundles are guillotined. The first step toward launch vehicle/ spacecraft separation has been taken.

• PYROTECHNIC SWITCH IGNITION

The second step in launch vehicle/spacecraft separation is the removal of power from the hot wires crossing station ZTO. These wires like the propellant lines, must also be guillotined, and the guillotine blade could cause a short circuit of the spacecraft power. Pyrotechnic switches B, C, D, E, F, G and J must be operated to remove power from the wires to be guillotined.
K3-36 and K11-25 apply power to launch vehicle/spacecraft pyrotechnic switch abort relay K4-30 and to wire guillotine latch relay K4-T4, initiating pyrotechnic switch ignition. K4-3O applies power to launch vehicle/spacecraft wiring pyrotechnic switch G igniter, opening pyrotechnic switch G. K4-T4 energizes pyrotechnic switch relays K4-25 and K4-26. K4-25 ignites equipment adapter pyrotechnic switches D, E end F. K4-26 ignites fuel cell wiring pyrotechnic switch B, C and S. With the operation of the pyrotechnic switches, the second step in launch vehicle/spacecraft separation has been taken.

• UPPER WIRE GUILLOTINE IGNITION

The third step in launch vehicle/spacecraft separation is the cutting of the upper wires that cross station Z70. This is accomplished by actuating the wire guillotines. Three wire guillotines igniters must be fired: the launch vehicle/spacecraft wire guillotine igniter C, the power wire guillotine igniter D, and equipment adapter wire guillotine Igniter E. When K4-25 and K4-26 energize, they apply power through the A contacts of K3-71 to wire guillotine relay K4-2. K4-2 fires the wire guillotine igniters C, D and E cutting the station ZTO wires. K4-2, contact C energizes the separate electrical latch relay K4-64, the adapter shaped charge relay K4-3 and the abort discrete relay K4-66. K4-66, contact A latches K4-2 in the energized position. K4-66 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. K4-3 prepares the way for the fourth step in the separation of the launch vehicle from the spacecraft.

• TUBING AND STRUCTURAL BOND CUTTING

The fourth and final step is to sever the adapter skin at station Z70 and breaks the launch vehicle to spacecraft structural bond. When K4-2 causes K4-3, the adapter shaped charge relay to energize, K4-3 fires the ZT0 tubing cutter igniter and the equipment adapter shaped charge igniters. The pyrotechnics complete the task of launch vehicle/spacecraft separation.

• RETROROCKET SALVO FIRE

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 5.5, 11.0, and 16.5 second time delay relays must be bypassed. Contacts C, D and E of K3-71 bypass the time delay relays. When K_-2 energizes, retrograde common bus power simultaneously energizes the retrorocket automatic fire relays K4-7, K4-9, K4-11 and K4-13. As these relays energize, retrograde squib bus power is applied to the igniters of retrorockets i, 3, 2 and 4. Salvo burn lasts approximately 5.5 seconds.

• RETROGRADE SECTION JETTISON

When the retrorocket automatic fire relays are energized by K4-2, the 45-second time delay relay K4-4 is also energized. When K4-4 energizes after 45 seconds, it illuminates the JETT RETRO indicator. The JETT RETRO switch-indicator is then pressed, and the retrograde section is Jettisoned in a mode II abort. However, in a mode I-II abort when the altitude is between 15,000 and 25,000 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.

SYSTEM UNITS

The Sequence System comprises the following units;

• Left switch/circuit breaker panel, consisting of three rows of circuit breakers and one row of switches.
• Boost and staging indicators, consisting of seven lights and three meters on the top of the command pilot and pilot's panels.
• Sequence controls, consisting of two pushbutton switches, eight switch-indicators, and one indicator are located on the left side of the main instrument panel.
• Re-entry switches and indicators, consisting of four switches on the main instrument panel center console and one switch, two lights, and two meters on the command pilot's panel.
• Abort controls, consisting of two D-rings on the ejection seats and one abort control handle on the left side of the cabin.
• Relay panels, consisting of four relay panels in the re-entry module and four in the equipment adapter and retrograde sections, and two In the rendezvous and recovery section.
• Separation sensing devices, consisting of three each in the equipment adapter section and the retrograde section.
  

The components of the Sequence System are described below:

LEFT SWITCH/ CIRCUIT BREAKER PANEL

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:

• ELECTRONIC TIMER CIRCUIT BREAKER

The electronic timer circuit breaker CB8-15 applies main bus power through contact A of lift-off relay K3-11 to start the electronic timer when the lift-off signal energizes the K3-11. The timer begins counting the time-to-go to retrograde.

• EVENT TIMER CIRCUIT BREAKER

The event timer circuit breaker CB8-14 applies main bus power through contact B of lift-off relay K3-ll to start the event timer when the lift-off signal energizes K3-11. The event counter counts the time since lift-off occurred.

• BOOST CUTOFF 1 CIRCUIT BREAKER

The boost cutoff I circuit breaker CB3-8 applies BIA common control bus power to the booster shutdown switch on the abort control and to the secondary guidance (RGS-IGS) switch. This circuit breaker arms the booster shutdown circuit and the secondary guidance manual switch-over circuit.

• BOOST CUTOFF 2 CIRCUIT BREAKER

The boost cutoff 2 circuit breaker CB3-21 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.

• RETRO AUTO CIRCUIT BREAKER

The retrograde fire automatic circuit breaker CB4-l applies retrograde common control bus power to the ARM AUTO RETRO switch. It provides power to salvo fire the retrorockets during the abort sequence. If CB4-1 is not closed, the electronic timer Tr contact closure will not automatically fire the retrorockets.

• RETRO MAN CIRCUIT BREAKER

The retrograde manual circuit breaker CB4-2 provides retrograde common control bus power for manually firing the retrorockets, and salvo firing the retrorockets with the abort control handle.

• Tr-256 CIRCUIT BREAKER

The retrograde minus 256 seconds circuit breaker CB8-16 applies common control bus power to relay contacts in the electronic timer and contacts of the TR-256 second relay. CB8-16 enables the TR-256 second signal to illuminate amber the IND RETRO ATT, BTRT PWR, and RCS indicators on the main instrument panel.

• SEQ LIGHTS POWER CIRCUIT BREAKER

The sequence lights power circuit breaker CB6-1 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.

• SEQ LIGHTS CONTROL CIRCUIT BREAKER

The sequence lights control circuit breaker CBI-13 applies common control bus power through the four MAIN BATTERIES switches to relay K1-29. When the main battery power indicator relay K1-29 is energized the BTRY PWR indicator on the main instrument panel is illuminated green. The third row of circuit breakers on the left switch/circuit breaker panel perform functions related to the Sequential System. The functions are the following:

• ATT IND CNTL RETRO CIRCUIT BREAKER

The attitude indicate control retrograde circuit breaker CB12-7 applies retrograde common control bus power to the IND RETRO ATT switch-indicator and to contacts of retrograde bias off relays K4-62 and K4-63. Power from CB12-7 energizes retrograde bias relay I(12-5when the JETT RETRO indicator is pressed.

• BOOST-INSERT CONTROL 1 CIRCUIT BREAKER

The boost-insert control 1 circuit breaker CB3-1 provides BIA squib bus number 1 power to initiate the abort sequence with the abort control handle, jettison the nose fairing and scanner cover, separate the spacecraft from the launch vehicle, sense launch vehicle/spacecraft separation, extend the uhf and diplexer whip antennas, and initiate several experiments.

• BOOST-INSERT CONTROL 2 CIRCUIT BREAKER

The boost-insert control 2 circuit breaker CB3-11 connects BIA SQUIB BUS number 2 power redundantly to the same switches to which CB3-1 connects power.

• RETRO SEQ CNTL CIRCUIT BREAKER

The retrograde sequence control 1 circuit breaker CB4-3 connects the retrograde Squib bus number 1 to the SEP OAMS LINES switch-indicator, the SEP ADAPT switch-indicator, the SEP ELEC switch-indicator, and the JETT RETR0 switch-indicator on the main instrument panel. It also arms the abort discrete relays and the equipment adapter separation sensor switches and relays.

• RETRO SEQ CNTL CIRCUIT BREAKER

The retrograde sequence control 2 circuit breaker CB4-28 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.

• SEQ LIGHTS TEST (AMBER-OFF-RED & GREEN) SWITCH

The sequence lights test switch connects main bus power to all amber-colored sequence lights and to all lights on the annunciator panel in the AMBER positions, and to all red or green sequence lights in the RED & GREEN position.

• SEQ LIGHTS (BRIGHT-DIM) SWITCH

The sequence light bright-dim switch is a single-pole, double-throw toggle switch. It connects the main bus through a diode to all sequence light circuits in the BRIGHT position. It connects the bus through a resistor to the same circuits in the DIM position.
The fourth row on the left switch/circuit breaker panel contains eight switches. These switches arm or safety the various squib buses used by the Sequential System. Their functions are as follows.

• BOOST-INSERT (ARM-SAFE) SWITCH

The boost-insert squib bus arm-safe switch is a four pole, double throw toggle switch. In the ARM position, this switch arms the BIA squib buses I and 2 and the BIA common control bus. These buses arm the SEP SPCFT switch-indicator, the BOOST C0TOFF 1 and 2 circuit breakers, the BOOST CUTOFF CNTL 1 and CNTL 2 circuit breakers, and the relay contacts which fire the nose fairing Jettison, scanner cover jettison, OAMS activate, RCS activate, spacecraft separate, guillotine and pyrotechnics.

• RETRO-POWER (ARM-SAFE) 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. Through these buses it arms the RETRO JETT ARM-SAFE switch, the RETR0 ROCKET SQUIBS ARM-SAFE 1, 2, 3, and 4 switches, the ATT IND CNTL RETRO, RETR0 SEQ. 1 and 2, and RETR0 AUTO and MAN circuit breakers on the left switch/ circuit breaker panel, and the RCS SQUIB 1 and 2 circuit breakers on the overhead switch/circuit breaker panel.

• RETRO-JETT (ARM-SAFE) SWITCH

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.

• RETRO-ROCKET SQUIB 1, 2, 3, 4 (ARM-SAFE) SWITCHES

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. When both the RETRO POWER squib arm switch and the four RETRO POWER SQUIB arm switches are placed to the ARM position, the OAMS squib buses 1 and 2 are connected redundantly to the retrorocket fire relays.

B00ST-INSERT-ABORT CORTROLS AND INDICATORS

Seven indicators, three meters and four controls are provided for the boost-insert-abort phase of the spacecraft mission.

• ENGINE I INDICATORS

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 0.91 1.25 seconds later as the pressure increases above 68 percent. Both indicators illuminate at booster engine cut-off and extinguish quickly at staging.

• ENGINE II INDICATOR

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. Approximately one second after both ENGINE I indicators extinguish, the ENGINE II indicator also extinguishes, indicating normal staging and engine 2 fuel injector pressure build up.

• ATT RATE INDICATOR

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.

• SEC GUID INDICATOR

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.

• ABORT INDICATORS

Two ABORT indicators are provided, one for each pilot. Both indicators illuminate red when the abort command is transmitted. When the ABORT indicator is illuminated, immediate and appropriate action is imperative. 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.

• STAGE 1 FUEL/OXYDIZER METERS

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.

• STAGE 2 FUEL/OXYDIZER METERS

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 30-psia mark indicates the minimum acceptable stored pressure in the tank before pressurization. After spacecraft separation, the meters indicate maximum psia.

• LONGITUDINAL ACCELEROMETER

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 l6 g's. The meter has positive and negative memory pointers. The accelerometer enables the crew to monitor the effectiveness of the