NASA Project Gemini Familiarization Manual

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.

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
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. 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. 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.
 * GENERAL
 * RE-ENTRY MODULE
 * ADAPTER

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 dimethylhydrazine 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.

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. 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. 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. On 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. 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. 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. 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. 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.
 * EQUIPMENT BAYS
 * DOORS
 * HATCHES
 * WINDOWS
 * HEAT SHIELD
 * SHINGLES

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.

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. 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. 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. An elbow restraint is provided for the command pilot only. It is used to stabilize his forearm during manual re-entry. 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. 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. 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.
 * INERTIA REEL
 * ARM RESTRAINT
 * ELBOW RESTRAINT
 * LEG RESTRAINT STRAP
 * EJECTION SEAT FOOT STIRRUP
 * LAP BELT

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. 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.
 * EGRESS KIT CUSHION (Effective Spacecraft 5 and 6)

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.

PERSONNEL 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: The forward survival kit, mounted on the forward surface of the backboard to the left of the pilots shoulder, contains the following:
 * Life Raft Container (Typical)
 * 1 Life Raft
 * 1 Sea anchor
 * 1 4 inch x 4 inch Foam rubber pad
 * 1 CO2 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
 * 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-OFF-DIM switch and the CABIN LIGHTS switch-circuit breaker are located on the overhead switch and circuit breaker panel. 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.
 * ELECTRICAL OUTLETS

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 moi

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.

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 (SSECO). 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 SSECO 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-IO 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 K11-10 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. O2 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 O2 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 O2 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 OAMS 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 RETRO 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 RETRO 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 100 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. 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 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 SHUTDOWN 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 40,000 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 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. 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 SHUTOWN 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. OAMS 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 MODE I
 * Abort Mode I-II
 * ABORT MODE II
 * Abort Mode III

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. 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. 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. 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, 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 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. 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. 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. K4-30 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. 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. 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. 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. 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.
 * SHUTDOWN
 * ABORT INITIATE
 * ABORT TELEMETRY
 * ABORT SQUIB BUS ARMING
 * RE-ENTRY CONTROL SYSTEM (RCS) ACTIVATION
 * OAMS LINES AND LOWER WIRES GUILLOTINE
 * PYROTECHNIC SWITCH IGNITION
 * UPPER WIRE GUILLOTINE IGNITION
 * TUBING AND STRUCTURAL BOND CUTTING
 * RETROROCKET SALVO FIRE
 * RETROGRADE SECTION JETTISON

SYSTEM UNITS
The Sequence System comprises the following units; The components of the Sequence System are described below:
 * 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.

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: 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. 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. 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. 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. 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. 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. 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. 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. 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: 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. 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. 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. 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 RETRO switch-indicator on the main instrument panel. It also arms the abort discrete relays and the equipment adapter separation sensor switches and relays. 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. 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. 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. 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 CUTOFF 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. 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 RETRO ROCKET SQUIBS ARM-SAFE 1, 2, 3, and 4 switches, the ATT IND CNTL RETRO, RETRO SEQ. 1 and 2, and RETRO 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.
 * ELECTRONIC TIMER CIRCUIT BREAKER
 * EVENT TIMER CIRCUIT BREAKER
 * BOOST CUTOFF 1 CIRCUIT BREAKER
 * BOOST CUTOFF 2 CIRCUIT BREAKER
 * RETRO AUTO CIRCUIT BREAKER
 * RETRO MAN CIRCUIT BREAKER
 * Tr-256 CIRCUIT BREAKER
 * SEQ LIGHTS POWER CIRCUIT BREAKER
 * SEQ LIGHTS CONTROL CIRCUIT BREAKER
 * ATT IND CNTL RETRO CIRCUIT BREAKER
 * BOOST-INSERT CONTROL 1 CIRCUIT BREAKER
 * BOOST-INSERT CONTROL 2 CIRCUIT BREAKER
 * RETRO SEQ CNTL CIRCUIT BREAKER
 * RETRO SEQ CNTL CIRCUIT BREAKER
 * SEQ LIGHTS TEST (AMBER-OFF-RED & GREEN) SWITCH
 * SEQ LIGHTS (BRIGHT-DIM) SWITCH
 * BOOST-INSERT (ARM-SAFE) SWITCH
 * RETRO-POWER (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. 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.
 * RETRO-JETT (ARM-SAFE) SWITCH
 * RETRO-ROCKET SQUIB 1, 2, 3, 4 (ARM-SAFE) SWITCHES

BOOST-INSERT-ABORT CONTROLS AND INDICATORS
Seven indicators, three meters and four controls are provided for the boost-insert-abort phase of the spacecraft mission. The two ENGINE I indicators are provided on the co-pilot's panel to indicate thrust chamber underpressure of the first stage booster engines. Each indicator illuminates red when the thrust chamber pressure of the engine is 68 percent of rated pressure or less. Both indicators illuminate red at stage 1 ignition but extinguished 0.91 1.25 seconds later as the pressure increases above 68 percent. Both indicators illuminate at booster engine cut-off and extinguish quickly at staging. 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. The attitude rate indicator on the command pilot's panel indicates an evaluation of the launch vehicle attitude rates during the boost phase. The indicator is extinguishes if the attitude rates remain within acceptable limits but illuminates red if the rates exceed these 1imits. The secondary guidance indicator on the command pilot's panel indicates which guidance system is in operation. The indicator is extinguished to indicate that primary guidance is being used. The indicator illuminates amber to indicate that secondary guidance has been selected. 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. The stage 1 fuel end oxidizer meters on the command pilot's panel enable the crew to monitor the current status and progress of the boost phase, and to anticipate an abort condition if one should develop. These meters indicate the gas pressures in psia of the stage 1 fuel and oxidizer tanks. Dual indicator needles are provided for redundancy. The range of the stage 1 meters is 35 to 5 psia. A time-versus- pressure scale near the bottom of the meter shows the minimum required pressure at 20, 40, and 60 seconds after lift-off. Critical fuel tank pressure is indicated by a shaded column at the low end of the scale. After staging with no signals applied, the meters indicate maximum psia. The stage 2 fuel and oxidizer meters on the command pilot's panel indicate stage 2 fuel and oxidizer tank pressure over a 70 to 10 psia range. Redundant pointers are used. Critical fuel tank pressures are indicated by a shaded column at the low end of the scale. The S-flag at the 30-psia mark indicates the minimum acceptable stored pressure in the tank before pressurization. After spacecraft separation, the meters indicate maximum psia. The accelerometer on the command pilot's panel indicates the rate in g's at which the launch vehicle engines are changing the velocity of the spacecraft. The range of the accelerometer is minus 6g's to 16 g's. The meter has positive and negative memory pointers. The accelerometer enables the crew to monitor the effectiveness of the engines. It is a secondary indicator of staging. The guidance switch above the abort control handle permits the command pilot to manually change from primary guidance to secondary backup guidance. When back-up guidance has been selected either manually or automatically during stage I boost and the ground station determines that primary guidance is feasible during stage 2 boost, primary guidance can be selected again by momentarily placing the guidance switch to the RGS position. A D-ring is provided on the ejection seat of each pilot. These rings are pulled to initiate mode I abort at altitude below 70,000 feet. Refer to Section III of this volume for the location and operation of these devices. The abort control handle is located on the command pilot's side of the cabin. It is used for spacecraft re-entry in abort modes I-II, II and III. These modes are effective above 25,000 feet. The three positions of this handle are NORMAL, SHUTDOWN, and ABORT. In NORMAL, the handle is inoperative. When the handle is moved to SHUTDOWN, the engine cutoff command is sent to the operating launch vehicle engine. When the abort handle is moved to ABORT, an immediate spacecraft separation and retrograde sequence is performed. These sequences differs from the normal sequences in that they are performed without cues from the indicators on the main instrument panel.
 * ENGINE I INDICATORS
 * ENGINE II INDICATOR
 * ATT RATE INDICATOR
 * SEC GUID INDICATOR
 * ABORT INDICATORS
 * STAGE 1 FUEL/OXYDIZER METERS
 * STAGE 2 FUEL/OXYDIZER METERS
 * LONGITUDINAL ACCELEROMETER
 * RGS-IGS COMMAND SWITCH
 * D-RINGS
 * ABORT CONTROL HANDLE

SEQUENCE CONTROLS AND INDICATORS
The switches, indicators, and switches-indicators on the main instrument panel center console have the following nomenclature, place in the mission sequence, and functions. The jettison fairing switch is used at the end of second stage engine thrust decay, by the command pilot to jettison the nose fairing, and the horizon scanner head cover. The separate spacecraft switch-indicator is used in the separation-insertion phase of the sequence. The command pilot presses the switch-indicator approximately 20 seconds after second stage engine cutoff when the IVI displays the delta-V required for insertion. Pressing the switch-indicator causes several things to happen. Primarily, it detonates pyrotechnic devices which separate the spacecraft from the launch vehicle. Secondarily, it extends the uhf and diplexer antennas and readies the acquisition aid beacon for use. As the spacecraft moves away from the launch vehicle, separation sensors close and energize the spacecraft separation relays. The relays illuminate the Indicator green. The Indicate retrograde attitude switch-indicator is illuminated amber when the electronics timer energizes the Tr-256 second relay. The amber light cues the crew to press the switch-indicator at this time. When pressured, a bias voltage is placed on the pitch needle of the FDI, and the inertial platform is electrically placed in the BEF mode. When released the ember light is extinguished and a green light is illuminated. The battery power indicator illuminated amber by the Tr-256 second relay. The amber light cues the pilot to place the MAIN BATTERIES switch to ON, and the fuel cell switch or ADAPTER BATTERIES switch to OFF. This change must be made because the adapter section will be jettisoned at retrograde. When all of the main battery switches are on, the indicator changes from amber to green. The RCS indicator Is illuminated amber by the Tr-256 second relay. The amber light cues the command pilot to activate the RCS firing the fuel, oxidizer, and pressurant isolation squibs. Pressing the switch-indicator energizes relays which fire the squibs. The indicator changes from ember to green, indicating that the RCS has been activated. The separate OAMS lines indicator is illuminated amber by the Tr-256 second relay is the prepare-to-go to retrograde phase. The amber light cues the crew to seal and sever the OAMS lines before jettisoning the adapter. Pressing the switch-indicator energizes relays which ignite the pyrotechnics used to seal and sever the lines. The relays also fire pyrotechnic switches and wire guillotines severing some of the adapter-retrograde mating line wiring. The indicator changes from amber to green. The separate electrical indicator is also illuminated amber by the Tr-256 second relay. The amber light cues the crew to sever all the wiring at the retrograde/adapter mating line. Pressing the switch-indicator energizes the wire guillotine relay. The pyrotechnics are detonated and the wiring is cut. The indicator changes from amber to green to indicate that electrical separation has been accomplished. The separate adapter indicator is illuminated amber by the Tr-256 second relay. The amber light cues the crew to Jettison the adapter equipment section. Pressing the switch-indicator causes the adapter shaped charge and the Z7O tubing cutter pyrotechnic to be detonated, and the adapter section severed. Separation of the adapter section is sensed by the equipment adapter separation sensors. Two closed sensors energize the sensor relay and change the indicator from amber to green. The arm automatic retrofire indicator is illuminated amber by the Tr-30 second relay. The amber light cues the crew to arm the automatic retrofire circuits so that when the electronic timer closes the TR contacts at TR time, the retrorockets will fire automatically. Pressing the switch-indicator completes the patch from the retrograde common control bus to the timer TR contact, and also energizes the TR arm relay. The relay changes the light from amber to green. Contact closure at Tr time energizes the Tr signal relay. The signal relay energizes the 45-second time delay relay, fires the retro rockets at 5.5 second intervals, and puts the platform in the free mode. The manual fire retrorockets switch connects the retrograde common control bus to the manual retrograde latch relay. Contacts of this relay energizes the 45-second time delay relay, fire the retrorockets at 5.5-second intervals, and place the platform in the free mode operation. The Jettison retrograde adapter indicator is illuminated amber by the 45-second time delay relay 45 seconds after retrofire begins. The amber light cues the crew to jettison the retrograde adapter. Pressing the indicator ignites pyrotechnic switch H and other pyrotechnic devices which disconnect and guillotine the wires at the retrograde adapter section/re-entry vehicle mating line. It fires the shaped charges which sever the retrograde adapter section from the re-entry vehicle. It energizes the Horizon Sensor System scanner head jettison relays which fire the jettison squibs and Jettison the scanner head. It removes the retrograde attitude signals applied to the flight director needles at Tr-256 seconds. It switches the FDI roll channel to the mix mode for re-entry. Finally by igniting pyrotechnic switch H it extinguishes the IND RETRO ATT, SEP OAMS LINE, SEP ELEC, SEP ADAPT and ARM AUTO RETRO green indicators and the JETT RETRO amber indicator.
 * JETT FAIRING PUSHBUTTON SWITCH
 * SEP SPCFT SWITCH INDICATOR
 * IND RETRO ATT SWITCH-INDICATOR
 * BTRY PWR INDICATOR
 * RCS SWITCH-INDICATOR
 * SEP OAMS LINES SWITCH-INDICATOR
 * SEP ELEC SWITCH-INDICATOR
 * SEP ADAPT SWITCH-INDICATOR
 * ARM AUTO RETRO SWITCH-INDICATOR
 * MAN FIRE RETRO PUSHBUTTON SWITCH
 * JETT-RETRO SWITCH-INDICATOR

RE-ENTRY VEHICLE RELAY PANELS
Ten Sequence System relay panels are installed in Gemini Spacecraft 5, 6, and 8 through 12. Four relay panels are located in the re-entry vehicle, three in the retrograde section, one in the equipment section, and two in the rendezvous and recovery section. The following Sequence System relay panels are in the re-entry module. The boost-insert-abort control relay panel contains six relays to perform spacecraft separation indicator control and launch vehicle/spacecraft pyrotechnic switch firing. The necessary functions required for adapter retrograde section separation are performed by the fourteen relays of the retrograde separation relay panel. The relays perform such functions as pyrotechnic switch and shaped charge ignition, Tr-3O second indication, automatic IGS free mode selection, and arming of the contacts of the Time Reference System. Re-entry Control System squib firing, scanner cover and scanner heads jettison, abort interlock RCS amber light actuation, and RCS ring B squib firing test prior to launch are provided by the sixteen relays of the attitude control system scanner and RCS squib fire relay panel. The umbilical pyrotechnic switch relay panel contains two relays which apply landing squib bus I and 2 power to re-entry umbilical wiring pyrotechnic switch.
 * BIA CONTROL RELAY PANEL
 * RETROGRADE SEPARATION RELAY PANEL
 * ACS SCANNER AND RCS SQUIB FIRE RELAY PANEL
 * UMBILICAL PYROTECHNIC SWITCH RELAY PANEL

ADAPTER RELAY PANELS
The retrograde section contains the following three relay panels which control spacecraft separation, retrofire, and equipment section separation. The equipment section contains the Orbit Attitude Maneuver System squib fire relay panel. The spacecraft separation control relay panel contains six relays to perform the following functions: shaped charge ignition, and launch vehicle/spacecraft guillotine firing. The retrofire relay panel has twenty relays. These relays control the automatic, manual and salvo firing of the retro rockets, and time the 5.5-second firing sequence. The retrograde sequence adapter separate relay panel contains twelve relays. The relays are used for equipment adapter shaped charge ignition, propellant line guillotine, electrical wire guillotine, and retrograde abort. The OAMS squib fire relay panel contains six relays for firing the OAMS squibs and controlling the regulator valves. The Rendezvous and Recovery section contains two Sequence System relay panels: the nose fairing Jettison relay panel, and the docking relay panel. The nose fairing Jettison relay panel contains two relays which control the jettisoning of the nose fairing. The docking relay panel has eleven relays which extend the docking index bar, illuminate the MSG ACPT light, effect emergency release of the docking latches, release and jettison the locking latches at retrograde, jettison the index bar, and cover the docking latch ports.
 * SPACECRAFT SEPARATION CONTROL RELAY PANEL
 * RETROGRADE FIRE RELAY PANEL
 * RETROGRADE SEQUENCE ADAPTER SEPARATE RELAY PANEL
 * ORBIT ATTITUDE AND MANEUVER SYSTEM SQUIB FIRE RELAY PANEL
 * RENDEAVOUS AND RECOVERY SECTION RELAY PANELS
 * NOSE FAIRING JETTISON RELAY PANEL
 * DOCKING RELAY PANEL

SEPARATION SENSORS
The Sequence System contains two sets of separation sensors. These are the launch vehicle/spacecraft separation sensors and the equipment adapter/re-entry vehicle separation sensors. Separation sensors are toggle switches which are normally open before separation is initiated. The separating structure will close the sensors as it moves away from the spacecraft re-entry module. The closure of any two of a set of three sensors is sufficient to sense and indicate separation.

SYSTEM DESCRIPTION
The Electrical Power System for the Gemini Spacecraft basically consists of two fuel cell battery sections, four silver-zinc main batteries and three silver-zinc squib batteries (spacecraft 6 uses three 400 ampere/hour sliver-zinc batteries in lieu of the fuel cell batteries). No primary ac electrical power system Is provided for the spacecraft. Devices requiring ac power obtain this power from self-contained inverters within the individual systems. The Electrical Power System includes switches, circuit breakers, relay panels, ammeters, a voltmeter and telelights which provide control, distribution and monitoring for the system. Also included as an Electrical Power System subsystem is the Reactant Supply System (RSS) which provides storage and control of the reactants (hydrogen and oxygen) used for fuel cell battery operation (not applicable to spacecraft 6). Provisions are made for utilizing external power and remote monitoring of the spacecraft power buses during ground tests and pre-launch operations. The two fuel cell battery sections and four main batteries provide dc power to the spacecraft main power bus (on spacecraft 6, the three adapter module batteries and the four main batteries provide dc power to the main bus). The squib batteries provide dc power to the common control bus and the two Orbital Attitude Maneuvering System (OAMS) squib buses. The OAMS squib buses in turn distribute dc power to the Boost-Insert-Abort (BIA), retrograde, landing and agena squib buses via the individual squib bus arming switches. The fuel cell SECTION 1 and SECTION 2 POWER (PWR and CNTL switches on spacecraft 5 and 6), PURGE, X-OVER and stack control switches (IA through 2C) are located on the right instrument panel. The fuel cell PWR and CNTL switches are used to control the battery module power on spacecraft 6. The PURGE and X-OVER switches are inoperative on spacecraft 6. On spacecraft 5 and 6, a dual-vertical-readout ammeter is located on the right instrument panel. On spacecraft 5 and 8 through 12, a power system monitor consisting of: a delta pressure indicator, three dual-vertical-readout ammeters and an ac/dc voltmeter, with associated selector switches, is located on the right instrument panel. On spacecraft 5, two FCAP indicator lamps are located on the right instrument panel. On spacecraft 6, a conventional voltmeter and ammeter with associated selector switches are located on the right instrument panel. The MAIN BATTERIES switches, SQUIB BATTERIES switches, BUS TIE switches, agena BUS ARM switch, FUEL CONT, FCAP (FC PANE5 on spacecraft 5 and 6) and FC O2 and H2 (CRYO O2 and H2 on spacecraft 10, 11 and 12) regulator and heater circuit breakers are located on the right switch/circuit breaker panel. The FC O2 and H2 regulator and heater circuit breakers are inoperative on spacecraft 6. The squib bus arming switches are located on the left switch/circuit breaker panel. The BTRY PWR sequence light, FCAP telelights, O2 and H2 heater switches, O2/H2 quantity indicator (integral with Environmental Control system (ECS) O2 indicator) and selector switch are located on the center instrument panel. The O2 and H2 heater switches, quantity indicator and selector switch are identified as CRYO switches and indicator on spacecraft 1O, 11, and 12. On spacecraft 8 and 9, an O2 CROSS-FEED switch is also located on the center instrument panel. On spacecraft 10, ll and 12, this switch is identified as H2 TANK VAC. The power system relay panel, power distribution relay panel and adapter power supply relay panel are located in the left equipment area of the cabin.

PRE-LAUNCH
In order to conserve spacecraft battery power, external electrical power is utilized during the pre-launch phase of the mission. External power is supplied to the spacecraft common control, main and squib power buses through umbilical cables connected to the re-entry module and equipment adapter section umbilical receptacles. The external power source is provided by Aerospace Ground Equipment (AGE). SQUIB BATTERIES switches 1 and 2 must be placed in the umbilical (UMB) position in order to apply external power to the spacecraft squib buses. Remote control of the spacecraft squib bus arming relays, and remote monitoring of the spacecraft power buses is also accomplished through the re-entry and adapter umbilical cables. Prior to launch, all MAIN BATTRIES and SQUIB BATTERIES switches, SECTION POWER switches (SECT PWR and CNTL switches on spacecraft 5 and 6) and stack control switches (1A through 2C) are set to the ON position to insure maximum redundancy of the Electrical Power System during the launch phase of the mission. On spacecraft 5 and 8 through 12, the fuel cell batteries are activated in sufficient time prior to launch, to insure launch readiness of the fuel cell batteries and RSS. The common control bus and OAMS squib buses are switched from external power to the squib batteries in sufficient time prior to launch, to verify the squib battery circuits. The BIA buses are armed prior to launch by setting the BOOST_INSERT ARM/SAFE switch to ARM position. The re-entry and equipment adapter section umbilicals are disconnected from the spacecraft Just prior to lift-off. Normally, umbilical separation is accomplished by an electrical solenoid device. A backup method of separation is also provided by a lanyard initiated mechanism which is actuated by movement of the launch vehicle.

ORBIT
From launch time until booster separation and insertion into orbit, both the fuel cell battery sections (module batteries on spacecraft 6) and the four main batteries are connected in parallel to the main power bus. After booster separation is accomplished, the MAIN BATTERIES switches are placed in the OFF position to conserve the main battery power. Also the pilots will disarm the BIA squib buses by setting the BOOST-INSERT ARM/SAFE switch to SAFE. In the event of an abort, all squib buses required for the abort function, which are not armed prior to launch, are armed via the abort relays controlled by the Sequential System. These relays effectively bypass the applicable squib bus arming switches which normally arm these buses. The SQUIB BATTERIES switches remain in the ON position throughout the entire mission until landing is accomplished. All three squib batteries are connected to the common control bus through diodes for individual fault protection. Squib batteries 1 and 2 are connected to the two OAMS squib buses via the de-energized squib bus arming relays. The O2 CROSS-FEED switch (spacecraft 8 and 9) remains in the CLOSED position except in the event of a loss of RSS O2 tank pressure. This switch controls the O2 crossfeed valve, which provides the capability of connecting the ECS O2 supply to the fuel cell battery sections. The H2 TANK VAC switch (spacecraft 10, 11 and 12) provides the capability of venting the area between the inner and outer wall of the RSS H2 supply tank to outside vacuum in space. This switch, when in VENT position, initiates a pyrotechnic cutter device to perform this function. The H2 TANK VAC switch remains in the SAFE position until the pilots elect to perform this function. The BUS TIE switches remain in the OFF position unless the necessity arises where the pilots must use main bus power to fire the squibs. The BUS TIE switches provide a method of connecting the main bus to the OAMS squib and common control buses. The agena BUS ARM switch will be used according to the mission requirements. On spacecraft 5 and 8 through 12, a small percentage of the reactant gases must be purged from the fuel cell batteries periodically to insure that the impurities contained in the feed gases do not restrict reactant flow to the cells and to remove any accumulation of product water in the gas lines. This purging function is performed by the pilots manually actuating the O2 and H2 PURGE switches. The pilots may increase the flow of gases to the fuel cell sections for more effective purging by setting the X-OVER switch to ON position.

RE-ENTRY
At 256 seconds before retrograde time (Tr-256) the pilots will arm the retrograde squib buses by setting the RETRO PWR ARM/SAFE switch and the individual RETRO ROCKET SQUIBS No. l, 2, B or 4 ARM/SAFE switches to ARM position. The retrograde rockets are used according to the mission requirements. The MAIN BATTERIES switches must be returned to the ON position at Tr-256 seconds to insure continuity of main bus power at the time of separation of the equipment adapter section containing the RSS/fuel cell module (battery module on spacecraft 6), from the spacecraft. There is no automatic switching provided for this function. The stack control switches (IA through 2C) and the SECTION 1 and SECTION 2 POWER switches (SECT 1 and SECT 2 PWR and CNTL switches on spacecraft 5 and 6) are set to OFF position after the main batteries are properly connected to the main bus. After retrograde rocket firing has been accomplished, the pilots will set the RETRO JETT ARM/SAFE switch to ARM. This switch provides a method of arming the JETT RETRO switch (center instrument panel), and is effectively an interlock to prevent inadvertent Jettisoning of the retrograde section prior to firing of the retrograde rockets. After the equipment adapter and retrograde sections are separated from the spacecraft, the pilots will disarm the retrograde squib buses by setting the RETRO PWR ARM/SAFE switch, RETRO JETT ARM/SAFE switch and RETRO ROCEET SQUIBS ARM SAFE switches to SAFE position. At this time, the landing squib buses are armed by setting the LANDING ARM/SAFE switch to ARM position. After landing is accomplished, the pilots will disarm the landing squib buses by returning the LANDING ARM/SAFE switch to SAFE. At this time, power will be removed from the OAM3 squib buses by setting SQUIB BATTERIES switches 1 and 2 to OFF position. All unnecessary electrical equipment will be deactivated to conserve the remaining spacecraft batteries for recovery equipment operation. SQUIB BATTERIES switch 1 and the MAIN BATTERIES switches will remain in the ON position to power the main and control buses throughout the recovery phase of the mission.

MONITOR AND DISPLAY
Throughout the mission, visual displays of bus voltage and current are provided by the system voltmeter and ammeters. On spacecraft 5 and 8 through 12, a power system monitor, which consists of a delta pressure indicator, three dual ammeters and an ac/dc voltmeter is utilized. The ammeters monitor individual fuel cell stack current (IA through 2C). The voltmeter, used in conjunction with a selector switch, displays individual fuel cell stack voltage, common control bus voltage, OAMS squib bus i and 2 voltage, Inertial Guidance System (IGS) inverter output voltage (spacecraft 8 through 12 only), main bus voltage and individual main battery voltage with the selector switch in Battery Test (BT) position and a particular MAIN BATTERIES switch in TEST position. The delta pressure indicator, in conjunction with a selector switch, provides a visual display of O2 versus H2 and O2 versus H2O differential pressure in the fuel cell battery sections. In the event that the differential pressure exceeds the prescribed limits, the pilots must evaluate the fuel cell battery performance, and if a malfunction exists, shut down the malfunctioning fuel cell battery section by setting the applicable fuel cell power and stack control switches to OFF position. The delta pressure indicator is inoperative on spacecraft 5. An out of tolerance delta pressure indication is also provided by the fuel cell delta pressure (FCAP) telelights on the center instrument panel. The lights are illuminated red when a malfunction exists. On spacecraft 5 only, two FCAP indicator lamps on the right instrument panel are illuminated red when a malfunction exists. The reactant (O2 and H2) supply quantities are displayed on the ECS O2 quantity indicator(center instrument panel) when the associated selector switch is set to FC O2 or FC H2 positions. On spacecraft 10, 11 and 12, this indicator and switch are identified as CRYO O2 and H2. The BTRY PWR sequence light, located on the center instrument panel, is illuminated amber at Tr-256 seconds during the mission by action of the Tr-5 relay in the power system relay panel. This informs the pilots that they must return the MAIN BATTERIES switches to ON position to insure continuity of main bus power due to the impending separation of the equipment adapter section containing the adapter power supply (fuel cell battery sections on spacecraft 5 and 8 through 12 and silver-zinc batteries on spacecraft 6). With all main batteries properly connected to the main bus, the BTRY PWR sequence light is illuminated green. On spacecraft 5 and 6 the dual-vertical-readout section ammeter provides a display of section 1 and 2 main bus current. Section 1 includes 50 percent of the adapter power supply current (fuel cell batteries or silver-zinc batteries as applicable) plus main batteries 1 and 2 current. Section 2 includes 50 percent of the adapter power supply current plus main batteries 3 and 4 current. The stack ammeter (used for battery test ammeter on spacecraft 6) with selector switch in 1A, IB, 1C or 2A, 2B, 2C positions, displays 50 percent of adapter module battery current. With the selector switch in BT position, the ammeter displays individual main battery test current as the appropriate MAIN BATIERIES switch is set to TEST position. Displays of common control bus voltage, main bus voltage, OAMS squib bus voltage, and adapter module battery voltage is provided by the system voltmeter and selector switch. Individual main battery voltage (with the particular battery removed from the main bus) is monitored with the voltmeter selector switch in BT position and the applicable MAIN BATTERIES switch in TEST position. The FCAP telelights and reactant quantity indications are not operative on spacecraft 6.

SILVER ZINC BATTERIES
The four main batteries are 45 ampere/hour, 16 cell, silver-zinc batteries. The three squib batteries are 15 ampere/hour, 16 cell, silver-zinc batteries. The squib batteries are special high-discharge-rate batteries which will maintain a terminal voltage of 18 volts for one second under a 75 ampere load. On spacecraft 6, three 400 ampere/hour, 16 cell, silver-zinc batteries are installed in the adapter battery module. These batteries are used in lieu of fuel cell batteries. All of the silver-zinc batteries have an open circuit terminal voltage of 28.8 to 29.9 volts DC. The main and squib battery cases are made of titanium. The approximate activated (wet) weight for each squib battery is 8 lb. and each main battery 17 lbs. The adapter module battery cases (spacecraft 6) are constructed of magnesium and the approximate wet weight of each battery is 118 lbs. The battery electrolyte consists of a 40 percent solution of reagent grade potassium hydroxide and distilled water. The main and squib batteries have a vent valve in each cell designed to prevent electrolyte loss and will vent the cell to atmospheric pressure in the event a pressure in excess of 40 psig builds up within the cell. All of the silver-zinc batteries are equipped with relief valves which maintain a tolerable interior to exterior differential pressure in the battery cases. The batteries are capable of operating in any attitude in a weightless state. Prior to installation into the spacecraft, the batteries are activated and sealed at sea level pressure. All of the batteries are coldplate mounted to control battery temperature.

POWER SYSTEM RELAY PANEL
The power system relay panel contains relays necessary for controlling and sequencing power system functions. The panel contains the control relays for the fuel cell battery system and RSS, main battery power sequence light relay, Tr-5 relay and the squib bus arming relays.

POWER DISTRIBUTION RELAY PANEL
The power distribution relay panel contains the relays required for arming the retrograde squib buses in the event of an abort. These relays are controlled by the Sequential System.

ADAPTER POWER SUPPLY RELAY PANEL
The adapter power supply relay panel contains relays necessary for controlling adapter module power to the main power bus. The relay panel contains the stack power relays which connect the individual fuel cell stacks to the main bus. (On spacecraft 6 the stack power relays connect the adapter module batteries to the main bus.) The panel also contains diodes used for reverse current protection between the adapter power supply and the spacecraft main power bus.

AMMETERS
The main bus section ammeter (spacecraft 5 and 6) is a dual-edge-readout vertical reading meter having a 0-50 ampere range with a total accuracy of two percent. The NO. 1 scale displays main batteries 1 and 2 and 50 percent of the adapter power supply current. The NO. 2 scale displays main batteries 3 and 4 and 50 percent of the adapter power supply current. The ammeter is shunt connected between the main power bus and spacecraft ground. The fuel cell stack meter (used as a battery ammeter on spacecraft 6) with associated selector switch, provides a display of individual main battery test current with the selector in BT position and a particular MAIN BATIERIES switch in TEST position. With the selector switch in 1A, 1B, 1C or 2A, 2B, 2C positions, the ammeter displays 50 percent of the applicable adapter module battery current. The meter has a 0-20 ampere scale.

VOLTMETER
On spacecraft 6 the voltmeter, used in conjunction with a selector switch, displays main bus, common control bus and squib bus voltage. Individual main battery voltage may be monitored with the voltmeter selector switch set to BT position and a particular MAIN BATTERIES switch set to TEST position. The voltmeter displays applicable adapter module batteries (A, B and C) voltage when the selector switch is set to 1A, 1B, 1C or 2A, 2B, 2C positions. The voltmeter has a 0-50 vdc range.

POWER SYSTEM MONITOR
The power system monitor (not applicable on spacecraft 6) consists of five vertical reading indicators; a delta pressure indicator, three dual-readout ammeters and an ac/dc voltmeter. The delta pressure indicator (not operative on spacecraft 5) and voltmeter are used in conjunction with selector switches located just below on the instrument panel. The ammeters provide a display of individual fuel cell stack (1A through 2C) current (reading must be multiplied by 0.8 on spacecraft 5). The voltmeter, with the selector switch in appropriate position, displays individual fuel cell stack voltage, main bus, squib bus and common control bus voltage, individual main battery voltage (with a particular MAIN BATTERIES switch in TEST position) and IGS inverter output voltage with the selector switch in AC position. The AC position on the selector switch Is inoperative on spacecraft 5. The voltmeter has a 20-33 ac volt range and an 18-33 dc volt range. The delta pressure indicator has a 0-1.5 psi range with the selector switch in either H2 position and a 0-6 psi range with the selector switch in either H2O position. This indicator displays O2 versus H2 differential pressure and O2 versus H2O differential pressure for each fuel cell battery section.

FUEL CELL BATTERIES
The fuel cell batteries used in the Gemini Spacecraft are of the solid ion exchange membrane type using hydrogen (H2) for fuel and oxygen (O2) for an oxidizer. The fuel cell battery system is comprised of two separate sections which are sealed in air tight pressure containers. Each section is made up of three interconnected fuel cell stacks with plumbing for transferring hydrogen, oxygen and product water. Each fuel cell stack consists of 32 individual fuel cells. Each basic fuel cell is made up of two catalytic electrodes separated by a solid type electrolyte in laminated form. The electrolyte is composed of a sulfonated styrene polymer (plastic) approximately 0.10 inches thick. Thin films of platinum catalyst, applied to both sides of the electrolyte, act as electrodes and support ionization of hydrogen on the anode side of the cell and oxidation on the cathode side of the cell. A thin titanium screen, imbedded into the platinum catalytic electrode, reduces the internal resistance along the current flow path from the electrode to the current collector and adds strength to the solid electrolyte. On the hydrogen side of the fuel cell, a current collector is attached by means of a glass-cloth-reinforced epoxy frame which assures a tight seal around the edges of the cell, forming a closed chamber. Ribs in the collector are in contact with the catalytic electrode on the fuel cell, providing a path for current flow. The hydrogen fuel is admitted through an inlet tube in the frame of the current collector and enters each gas channel between the collector ribs by way of a series of slots in the tube. Another tube provides a purge outlet, making it possible to flush accumulated inert gases from the cell. The collector plate is made of approximately 0.003 inch thick titanium. On the oxygen side of the cell, a current collector of the same configuration and material as the hydrogen side collector is attached. Its ribs, located at right angles to those of the other collector, provide structural support to the electrolyte-electrode structure. A Dacron cloth wick, attached between the ribs, carries away the product water through capillary action, by way of a termination bar on one side of the assembly. Oxygen is admitted freely to this side of the fuel cell from the oxygen filled area of the section container. The cell cooling system consists of two separate tubes bonded in the cavity formed by the construction of the oxygen side current collector and the back side of the hydrogen current collector. Each tube passes through six of the collector ribs and has the cooling capacity to maintain operating temperature. The cooling of the oxygen current collector, which holds the product water transport wicks provides the coldplate for water condensation from the warmer oxygen electrode. The individual fuel cell assemblies are arranged in series to form a stack. When assembling the cells into a stack, the ribs of the oxygen side current collector contact the solid electrolyte of the fuel cell assembly. Titanium terminal plates are installed on the ends of the two outside cells to which connections are made for the external circuit. End plates, which are honey-comb structures of epoxy-glass laminate 0.5 inch thick, are installed on the outside of the terminal plates. Stainless steel insulated tie rods hold the stack together and maintain a compression load across the area of each cell assembly. This assures proper contact of the solid electrolyte with the ribs of each current collector. The fuel cell stacks are packaged in a pressure tight container, together with the necessary reactant and coolant ducts and manifolds, water separator for each stack, and required electrical power and instrumentation wiring. The hydrogen inlet line, hydrogen purge line, and the two coolant lines for each cell lead from their respective common manifolds running the length of the stack. The manifolds are made of an insulating plastic material and the individual cell connections are potted in place after assembly to provide a leak-tight seal. The oxygen sides of the cells are open to the oxygen environment surrounding the fuel cell assemblies within the container. An accessory pad is mounted on the outside of the fuel cell section container. It includes the gas inlet and outlet fittings, purge and shutoff valves, water valve and electrical connectors. Structurally, the container is a titanium pressure vessel consisting of a central cylinder with two end covers and two mounting brackets. Within the container, the fuel cell stacks are mounted on fiberglass impregnated epoxy rails by bolts which pass through the stack plates. These rails are in turn bolted to the mounting rings sandwiched between the two flanges on the section container. The hydrogen manifolds on each stack within a section are parallel fed with a hydrogen shutoff valve and check valve in the feed line to each stack. Oxygen is fed into the section container so that the entire free volume of the container contains oxygen at approximately 22.5 psia. The coolant reaches the fuel cell battery sections by two separate isolated lines. Any malfunction in the coolant line in one section will not affect the cooling function of the coolant line in the other section. Each stack in the section has its own water-oxygen separators which are manifolded into a single line coming out of the section container. All hydrogen, oxygen coolant, electrical and water storage pressure line connections at the section container are fastened to standard bulkhead fittings on the accessory pad. After the stacks are completely assembled within the container, all void spaces are filled with unicellular foam. The purpose of this foaming is for vibration dampening, acoustical noise deadening and minimizing free gas volume to prevent possible fire propagation. Thin plastic covers are placed over the top and bottom of each stack to manifold oxygen to the stack and to keep the foam material from entering areas around the coolant manifolds and oxygen water separator. The basic principle by which the fuel cell operates to produce electrical energy and water, is the controlled oxidation of hydrogen. This is accomplished through the use of the solid electrolyte ion-exchange membrane. On the hydrogen side of the fuel cell, hydrogen gas disassociates on the catalytic electrode to provide hydrogen ions and electrons. The electrons are provided a conducting path of low resistance by the current collector, either to an external load or to the next series-connected fuel cell. When a flow of electrons is allowed to do work and move to the oxygen side of the fuel cell, the reaction will proceed. By use and replacement, hydrogen ions flow through the solid electrolyte to the catalytic electrode on the oxygen side of the fuel cell. When electrons are available on this surface, oxygen disassociates and combines with the available hydrogen ions to form water. The oxygen current collector provides the means of distributing electrons and condensing the product water on a surface to be transported away by the wick system through capillary action. The individual cell wicks are integrated into one large wick which routes the water to an absorbent material that separates the water from the gas. By using the oxygen outlet pressure as a reference, a small pressure differential is obtained over the length of the water removal system. This pressure is sufficient to push the gas-free water toward the storage reservoir. Waste heat, generated during the fuel cell battery operation, is dissipated by means of the recirculating coolant provided by the spacecraft cooling system. In addition, the total coolant flow provides the function of preheating the incoming reactant gases. In the spacecraft, the reactant gases are supplied to the fuel cell battery sections by the RSS. This system contains the reactant supply tanks, control valves, heat exchangers, temperature sensors and heaters required for management of the fuel cell reactants.
 * CONSTRUCTION
 * OPERATION

REACTANT SUPPLY SYSTEM
The RSS is essentially a subsystem for the fuel cell battery sections. The system provides storage for the cryogenic hydrogen and oxygen, converts the reactants to gaseous form and controls the flow of the gases to the fuel cell battery sections. The RSS components are installed in the RSS/fuel cell module. On spacecraft 10, 11 and 12, the RSS O2 requirements are supplied from a central cryogenic O2 tank located in the RSS/fuel cell module. This vessel also supplies O2 to the ECS. Two tanks are utilized to separately contain the cryogenic hydrogen and oxygen required for the operation of the fuel cell battery sections. The tanks are thermally insulated to minimize heat conduction to the stored elements which would cause the homogeneous solution to revert to a mixture of gas and liquid. The tanks are capable of maintaining the stored liquids at supercritical pressures and cryogenic temperatures. The approximate total amount of liquid stored in the hydrogen vessel is 22.25 lb. for long mission configurations and 5.80 lb. for short mission configurations. The approximate total amount of liquid stored in the oxygen vessel is 180 lb. on spacecraft 5, 106 lb. on spacecraft 10, 11 and 12 (for ECS and RSS) and 46.0 lb. on spacecraft 8 and 9. The hydrogen vessel is composed of titanium alloy and the oxygen vessel Is made of a high strength nickel base alloy. Both vessels are spherical in shape and double welded. A vacuum space between the inner and outer wall (approximately one inch) provides thermal insulation from ambient heat conduction. The inner wall is supported in relation to the outer wall by an insulating material supplemented by compression loading devices. Each storage tank contains a fluid quantity sensor, a pressure sensor, a temperature sensor and an electrical heater (the O2 tank on spacecraft 10, 11, and 12 has two heaters) installed in the inner vessel in intimate contact with the stored reactants. The fluid quantity sensor is an integral capacitance unit which operates in conjunction with an indicator control unit containing a null bridge amplifier. The sensor varies the capacitance (in proportion to fluid level) in a circuit connected to the null bridge amplifier. The amplified signal is then used to drive a servo motor, which in turn operates a visual indicator for quantity indication. Power inverters supply 400 cycle, 26 vac power to the fluid quantity circuits. The temperature sensor is a platinum resistance device capable of transmitting a source signal to a balanced bridge circuit. The sensor provides cryogenic fluid temperature monitoring for telemetry and AGE. The pressure sensor is a dual resistive element, diaphragm type transducer. The sensor provides signals for cryogenic fluid pressure monitoring on a spacecraft meter. The electrical heaters provide a method of accelerating pressure build-up in the reactant supply tanks. The heaters may be operated either in a manual or automatic mode. In the automatic mode a pressure switch removes power from the heater element when the tank pressure builds up to a nominal 900 psig in the oxygen tank and a nominal 250 psig in the hydrogen tank. In the manual mode a spacecraft pressure meter indicates proper switch operation. The fill and vent valves provide a dual function in permitting simultaneous fill and vent operations. Quick disconnect fittings are provided for rapid ground service connection to both the storage tank fill check valve and the vent check valve. When fill connections are made, the pressure of the ground service connection against the fill and vent valve poppet shaft simultaneously opens both the fill and vent ports. When ground service equipment is removed, the valve poppet automatically returns to its normally spring-loaded-closed position. The vent check valve is a single poppet type, spring-loaded-closed valve which opens when system pressure exceeds 20 psig to relieve through the fill and vent valve vent ports. The supply temperature control heat exchangers are finned heat exchangers in which the supply fluid temperature is automatically controlled by absorbing heat from the recirculating coolant fluid of the spacecraft cooling system. The special doublepass design precludes freezing of the coolant and assures a reactant fluid supply at 50°F minimum and 140 ° maximum. The dual pressure regulator and relief valves are normally open poppet type regulators which control downstream pressure to the fuel cell battery sections. The regulators maintain the hydrogen pressure at approximately 21.7 psia and the oxygen pressure at approximately 2.2 psia. The oxygen side of the regulators is referenced to hydrogen pressure. The hydrogen side of the regulators is referenced to produce H2O pressure. The relief valves provide overpressurization protection for the regulated pressure to the fuel cell battery sections. This valve is precalibrated to operate at a pressure of approximately 10 psia above the normal supply level. The high pressure relief valves are single poppet type, spring-loaded-closed valves which provide system overpressurization protection. The valves vent system gas to ambient when pressure exceeds the system limits. The solenoid shutoff valves are solenoid operated latching type valves which eliminate fluid loss during the nonoperating standby periods. The valves are normally open and are closed only during fill and standby periods by applying power to the solenoids.
 * COMPONENTS
 * REACTANT SUPPLY TANKS
 * FILL AND VENT VALVES
 * HEAT EXCHANGERS
 * DUAL PRESSURE REGULATOR AND RELIEF VALVES
 * HIGH-PRESSURE RELIEF VALVES
 * SOLENOID SHUTOFF VALVES

The crossover valve is a solenoid operated latching type valve which provides the capability of selecting both dual pressure regulators to supply hydrogen and oxygen to a fuel cell battery section for the purpose of increasing flow rate for more effective purging. The crossover valve is controlled by the X-OVER switch on the right instrument panel. The O2 crossfeed reactant valve is a solenoid operated, latching type valve which provides the capability of pressurizing the RSS with O2 from the ECS oxygen supply. This provides a redundant method of supplying the proper reactant O2 pressure to the fuel cell battery sections in the event of a malfunction in the RSS oxygen supply. The crossfeed valve is controlled by the O2 CROSS-FEED switch located on the center instrument panel. During pre-launch, the two separate reactant supply tanks are serviced (using AGE equipment) with liquid hydrogen and oxygen. After the tanks are filled, in order to accelerate pressure buildup within the tanks, the internal tank heaters are operated, utilizing external electrical power. In approximately one hour the liquid is converted into a high density, homogeneous fluid at a constant pressure. During the fill operation, the solenoid shutoff valves between the storage tanks and the dual pressure regulators are closed. Once operating pressure is obtained, the solenoid shutoff valves may be opened by applying power to the coil of the valves. The high density, homogeneous fluid will then flow upon demand. The fluid flows from the supply tanks to the heat exchangers. The fluid temperature, when entering the heat exchangers is approximately -279°F for the oxygen and approximately -423°F for the hydrogen. The heat exchangers absorb heat from the recirculating coolant fluid of the spacecraft cooling system. This heat, applied to the high density fluid, raises the temperature of the reactants to approximately 50°F to 140°F. The reactants, now in gaseous form, flow through the heat exchangers, past the high pressure relief valves and AGE temperature sensors, to the supply solenoid shutoff valves. During fuel cell battery operation, if the demand on the fluid flow is inadequate to keep tank pressures within limits, the high pressure relief valves will vent, externally, the excess fluid. The AGE temperature sensors on the heat exchangers are used for pre-launch checkout only. The reactants flow through the supply solenoid shutoff valves to the dual pressure regulator and relief valves. The dual pressure regulators reduce the pressure of the reactants to approximately 21.7 psia for the hydrogen and approximately 20.5 psia for the oxygen. The gas now flows through the manual shutoff valves and is then directed to the fuel cell battery sections at a flow rate that is determined by both the electrical load applied and the frequency of purging. The flow rate of the gases may be increased for more effective purging by opening the crossover valve. After launch, the supply tank heaters are operated by spacecraft power. The heaters operate as required to maintain proper system pressures.
 * CROSSOVER VALVE
 * O2 CROSSFEED REACTANT VALVE (SPACECRAFT 8 AND 9)
 * OPERATION

SYSTEM DESCRIPTION
The Environmental Control System (ECS) may be defined as a system which provides a safe and comfortable gaseous atmosphere for the pilots. The system must perform such tasks as providing fresh oxygen, pressurization, temperature control, water removal and toxic gas removal. In addition to providing atmospheric control for the pilots, the system provides equipment cooling and regulated temperatures for certain pieces of equipment. For ease of understanding, the Environmental Control System may be separated into four systems or loops which operate somewhat independent of each other. These loops are:
 * (1) The oxygen supply system.
 * (2) The cabin loop.
 * (3) The suit loop.
 * (4) The water management system.

OXYGEN SUPPLY SYSTEM
There are three oxygen systems: Primary, Secondary and Egress. This system stores and dispenses oxygen for breathing and for suit and cabin pressurization. This system provides oxygen during the period commencing two hours prior to launch and terminating with jettison of the adapter section at retrograde. The primary oxygen supply is stored at supercritical pressure in a cryogenic spherical container in the adapter section of the spacecraft. This container is filled with liquid oxygen at atmospheric pressure. Heat is supplied by thermal leakage through the container insulation and by activation of an electric heater in order to build pressure to the critical point of 736 psia. Above this point liquid oxygen becomes a homogeneous mixture, described for simplicity as a dense supercritical fluid. This fluid is warmed, regulated and filtered before it enters the suit or cabin loop. On spacecraft 10 through 12, a long mission ECS oxygen tank will replace the short mission Reactant Supply System (RSS) oxygen tank and the ECS oxygen tank, allowing ECS breathing and the RSS oxygen to be stored in the same container. A tee fitting on the cryogenic line allows both the ECS and RSS systems to receive oxygen from the common container. The primary loop consists of the following components: primary oxygen container, pressure control switch, pressure transducer, fill and vent valves, temperature discharge sensor, pressure relief valve, pressure regulator, shutoff valve, filter, check valves, and heat exchanger. The secondary oxygen system is capable of performing the same functions as the primary oxygen system and operates when pressure in the primary system falls below 75 +- 10 psi. At retrograde, when the primary oxygen container is jettisoned with the equipment adapter, the secondary oxygen system assumes the duties of the primary oxygen system. The gaseous secondary oxygen supply is stored in two cylinders located in the re-entry module. Each tank contains 6.5 pounds of usable oxygen pressurized to 5000 psig maximum at 70°F. This oxygen supply is then regulated before it enters the suit or cabin loop. The secondary system consists of two: tanks, fill valves, transducers, pressure regulators, shutoff valves and check valves. This system provides each pilot with oxygen for breathing and for suit pressurization in the event that they initiate ejection procedures at 7%000 feet or below, during launch or re-entry. The egress oxygen is provided on spacecraft 5 and 6 only. The egress gaseous oxygen supply is stored in a tank located in each seat-mounted egress kit. Each tank contains 0.31 pound of usable pressurized oxygen. Each egress system consists of a tank, pressure regulator, pressure gage, restrictor, check valve, shutoff valve, and composite disconnects.
 * PRIMARY OXYGEN
 * SECONDARY OXYGEN
 * EGRESS SYSTEM

CABIN LOOP
Design cabin leakage at ground test conditions is 670 standard cubic centimeters per minute, (scc/min) of nitrogen at 5.0 psig. Makeup oxygen, to maintain cabin pressure at nominal 5.1 psia level, is called for by the cabin pressure regulator. In order to obtain maximum utilization of oxygen, it first passes through the suit loop before it is dumped into the cabin through the suit pressure relief valves. primary cabin components for spacecraft 5 and 6 are a cabin heat exchanger and a fan. These parts have been removed from spacecraft 8 through 12. This loop also contains a relief valve for both positive and negative pressure relief, a pressure regulator and manual valves to either dump cabin pressure or repressurize. In the latter operation, oxygen is supplied directly to the cabin.

SUIT LOOP
The pilots are provided with redundant atmospheres by having a closed pressure suit circuit within the pressurized cabin. This suit circuit provides for cooling, pressurization, purification and water removal. The suit loop is a closed system with two pressure suits operating in parallel. Circulation of oxygen through the suit is provided by a centrifugal compressor. Carbon dioxide and odors are removed by an absorber bed containing lithium hydroxide and activated charcoal. The gases are cooled in a heat exchanger by a liquid coolant, Monsanto MCS 198, to a temperature below the dew point. Water condensing within the heat exchanger is dumped overboard or routed to the water evaporator. The reconditioned Oxygen is mixed with fresh makeup oxygen. The suit circuit has two modes of operation, the normal recirculation mode which was discussed in the previous paragraph and the high rate mode which shuts off the recirculation system and pumps oxygen directly into the suit. The suit loop consists of two suit pressure demand regulator valves, four check valves, two throttle valves, two solid traps, a system shutoff and high flow rate valve, two compressors, one carbon dioxide and odor absorber, and a suit heat exchanger.

WATER MANAGEMENT SYSTEM
The purpose of the water management system is to store and dispense drinking water, collect and route unwanted water to the evaporator or dump overboard. Drinking water is stored in a tank or tanks in the adapter. Each tank contains a bladder and is pressurized to supply water to the transparent tank in the re-entry module. Spacecraft 5 utilizes two drinking water storage tanks which store both the drinking water and the fuel cell by-product water. One tank uses a combination of oxygen and fuel cell water as the pressurant while the other tank is pressurized with oxygen. Spacecraft 6 has only one storage tank and uses oxygen for the pressurant. Spacecraft 8 through 12 utilizes two storage tanks. Fuel cell by-product water is used as the pressurant for the drinking storage tank. The other tank is used to store fuel cell water. Urine and condensated water from the suit circuit heat exchanger is absorbed by the wick in the water boiler or dumped overboard. Components of the water management system, in addition to the water tanks, are a water control valve, condensate valve, water evaporator and two manual shutoff valves. The urine system is designed for total management of the crewmen’s urinary output. It samples and determines total volume of every urination and provides for chemical analysis. Volume is determined by sampling every urination and using a tracer dilution technique. Three tenths (0.3) ml of tracer chemical is added to the urine and samples are taken. During postflight operations, the amount that the tracer has been diluted in each sample determines the quantity of urine voided by the crewmen. The urine system consists of the following components; a Chemical Urine Volume Measuring System (CUVMS) with selector valve, tracer storage accumulator and collection/mixing bag, a urine receiver assembly with collection bag, a urine quick disconnect hose and urine solids trap filter, urine sampling bags and a roll-on cuff receiver assembly.
 * URINE SYSTEM

SYSTEM DISPLAYS AND CONTROLS
The displays and controls for the Environmental Control System are provided in the cabin and function as specified. A manual secondary oxygen shutoff handle is provided for each member of the flight crew for complete and positive shutoff of each secondary oxygen container. The handles are located aft of the right and left switch/circuit breaker panels. The position OPEN or CLOSED is noted. The O2 HI RATE telelight located on the instrument panel illuminates when the high oxygen rate valve is opened manually or automatically. The O2 HI RATE switch on spacecraft 5 and 6 also activates the cabin fan. The switch has three positions; CABIN FAN, O2 HI RATE, and OFF. Spacecraft 8 through 12 do not have the cabin fan, and that position of the switch is not used. The cabin air recirculation handle controls the recirculation valve which permits entry of cabin air into the suit circuit for removal of odors and carbon dioxide. This procedure will renovate cabin air without cabin decompression and reduces the possibility of carbon dioxide pockets by increasing circulation of the cabin atmosphere. This handle will control the cabin air inlet valve which provides for ventilation during landing and post-landing phases of the mission. This handle controls the operation of the cabin outflow valve to permit emergency decompression in orbit and cabin ventilation during the landing phase. This handle provides for watertight closure of the cabin pressure relief valve during a water landing. This handle provides for the manual return of the oxygen high rate valve to the closed position, thereby restoring normal oxygen flow rate. Actuation of this handle also reestablishes the capability of initiating high rate oxygen flow when necessary.
 * SECONDARY OXYGEN SHUTOFF HANDLE
 * O2 HIGH RATE TELELIGHT
 * CEN AIR RECIRC HANDLE
 * INLET SNORKEL HANDLE
 * CABIN VENT HANDLE
 * WATER SEAL HANDLE
 * O2 HIGH RATE RECOCK HANDLE

A dual indicator provides for monitoring temperatures in the suit and cabin circuits. Range markings are calibrated in degrees Fahrenheit. A dual indicator provides for monitoring cabin atmospheric pressure and the amount of carbon dioxide at the suit inlet. Cabin atmospheric pressure is calibrated in pounds per square inch. Carbon dioxide partial pressure is calibrated in millimeters of mercury. A dual indicator is provide for monitoring pressure in the individual gaseous oxygen containers in the secondary oxygen subsystem. The indicator range is from 0 to 6000 psia, divided into 500-pound increments and numbered at each l000-pound interval. Readings must be multiplied by 100 to obtain correct values. This indicator provides for monitoring quantity and pressure of cryogenic oxygen in the primary oxygen container. The quantity scale displays from 0 to 100 per cent in 2 per cent increments, numbered at 20 per cent intervals. The pressure scale ranges from 0 to l000 psia in 20-pound increments, numbered at 200-pound intervals. Red undermarkings are incorporated on the oxygen meter to indicate the point at which thermal pressurization may be discontinued by de-energizing the heaters. Spacecraft l0 through 12 reidentifies the indicator as the CRYO meter, but it performs the same function as before. This switch allows the same indicator to be used when monitoring the pressure and quantity of cryogen in any of the three cryogenic containers. The three containers are: the ECS primary oxygen supply, the RSS or fuel cell (FC) oxygen supply, and the RSS or FC hydrogen supply. The switch is located below the indicator on the center panel and has the following positions: ECS O2, FC O2, FC H2 (PX35) and OFF. Spacecraft 1O through 12 have a three position switch. The positions are; O2, H2 (PX35) and OFF. The O2 position allows the indicator to monitor the ECS-RSS oxygen supply and the H2 (PX35) position allows monitoring the RSS or FC hydrogen supply. This switch is connected to the heaters in the ECS primary oxygen container. The switch has three positions; AUTO, OFF, and ON. It is located below the flight plan display on the center panel. Spacecraft 10 through 12 reidentifies the ECS O2 HEATER switch as O2 HEATER switch and connects to heaters in the ECS-RSS oxygen container. The switch has three positions; NO 1, OFF, and NO 1 & 2. The switch is located in the upper left hand corner of the center panel. This allows suit fans number 1 and number 2 to operate together or independently. Suit fan NO 2 may be operated by placing switch in NO 1 & 2 position and placing suit fan NO 1 circuit breaker switch to OFF. The O2 CROSS FEED Switch, spacecraft 6, 8 and 9 when in the OPEN position, permits oxygen from the primary oxygen supply module for the ECS to be used in the RSS in the event of RSS oxygen module failure. The reverse is also true. A three knob panel is provided for managing, replenishing, and dumping waste water and urine overboard. Dual concentric knobs are mounted between the ejection seats for suit and cabin temperature control. These knobs control the operation of valves regulating the flow rate of primary and secondary coolant through the suit and cabin heat exchangers. Clockwise rotation results in increased temperatures. This handle is located on the console between the members of the flight crew and provides for manual control of the dual high oxygen rate and suit system shutoff valve. Actuation of the handle shall initiate the oxygen high flow rate and de-energize the suit compressor. Resumption of normal system operation shall be effected by actuation of the oxygen high rate recock handle. An individual lever is provided for each member of the flight crew for regulation of circulatory oxygen flow through the suit circuits. The levers are located on the lower section of the pedestal and shall provide any selected flow valve setting from fully open to fully closed. A detent provides an intermediate position to prevent inadvertent shutoff of suit flow. This detent may be bypassed by moving the lever outboard. A rotary handle control is provided for cabin repressurization after a decompression has occurred and for ELSS oxygen supply. The control rotates approximately 90° between fully OPEN (repressurize) and fully CLOSED (off) positions. This control is located on the center console between the suit flow control panels. This telelight, located on the annunciator panel of the center instrument panel, illuminates when the heater in the primary oxygen container has been manually activated.
 * CABIN AND SUIT TEMP INDICATOR
 * CABIN AND P CO2 PRESS INDICATOR
 * SEC O2 INDICATOR
 * ECS O2 QUANT % AND PSIA METER
 * CRYOGENIC QUANTITY SWITCH
 * ECS O2 HEATER SWITCH
 * SUIT FAN SWITCH
 * O2 CROSS FEED SWITCH
 * WASTE MANAGEMENT PANEL
 * SUIT AND CABIN TEMP CONTROLS
 * MANUAL O2 HIGH RATE HANDLE
 * SUIT FLOW CONTROL LEVERS
 * CABIN REPRESS CONTROL
 * ECS HTR TELELIGHT

SYSTEMS OPERATION
The Environmental Control System is semi-automatic in operation and provides positive control in all modes of operation. There are six operational modes: Prior to the pre-launch mode, it is necessary to service and to check the system functionally.
 * 1. Pre-Launch
 * 2. Launch
 * 3. Orbit
 * 4. Re-Entry
 * 5. Postlanding
 * 6. Emergency

SERVICE AND CHECKOUT
For this operation it is assumed that the spacecraft has been mated with the booster on the launch pad and in the unserviced condition. The primary, secondary and egress oxygen storage tanks are filled. The water boiler and drinking water supply tank are supplied with water. The cartridge in the suit loop canister is then replaced.

PRE-LAUNCH
The pilots in their suits, with face plates open, are connected to the suit circuit. The suit circuit compressor is actuated and the suit temperature control valve is adjusted to satisfy the pilot desiring the cooler temperature. The other pilot becomes comfortable by adjusting his suit flow rate control valve toward the closed position to obtain a warmer setting. A ground supply of pure oxygen is connected to the pressure suit circuit purge fitting. Flow is initiated with the face plates closed. The suit circuit gas is sampled periodically until an acceptable oxygen content is attained. A suit circuit leakage test is conducted. After satisfactorily completing the suit circuit leakage test, the primary and secondary oxygen manual shutoff valves are opened and the suit circuit purge system is disconnected and removed. The cabin hatches are closed. A ground supply of pure oxygen is connected to the cabin purge fltting2 flow is initiated and the cabin is purged. The cabin fan is actuated and the recirculation valve is opened. A cabin leakage test is conducted. After satisfactorily completing the cabin purge and leakage test, the cabin purge system is disconnected and removed and the cabin purge fitting is capped. The primary and secondary oxygen manual shutoff valves are opened. The liquid oxygen inside the primary supercritical container has been changing from a liquid to a supercritical fluid by thermal leakage and heater activation. A pressure control switch provides for automatic or manual activation of these heaters. The manual control switch is located on the center control panel. An indicator also on the center control panel indicates both pressure and quantity from a transducer and control unit that are attached to the container. The oxygen gas flows from the container and is warmed to approximately 50 F in a heat exchanger. This heat exchanger also contains a relief valve that limits maximum pressure to 1000 psig. This valve opens, permitting full flow and reseats within the range of 945-1000 psig. A discharge temperature sensor provides an indication, for telemetering only, of the temperature in the primary oxygen line downstream of the heat exchanger. The oxygen gas is regulated from 1000 psia maximum to 110 +- 10 psig. Flow capacity of 0.35 lb/min with an inlet pressure from 800 to 1000 psia and an inlet temperature of 60 F. This regulator also contains a relief feature that limits downstream pressure to 215 psig in the event of a failed-open condition. A 10-micron filter provides filtration of the primary oxygen supply before it enters the suit or cabin loop.
 * SUIT LOOP
 * CABIN LOOP
 * OXYGEN LOOP

LAUNCH
The cabin pressure relief valve opens to limit the pressure differential between cabin and ambient to 5.5 +- 0.0 psi. Oxygen is supplied to the suit loop through the suit pressure regulator. The suit pressure is controlled to between 2 and 9 inches of water above cabin pressure by the suit pressure regulator. Suit circuit oxygen from the suit circuit demand regulator recirculates through the suit compressor, the carbon dioxide and odor absorber, the suit heat exchanger and water separator, the pressure suits, and the suit circuit solids traps. There are two compressors in the circuit. One is an alternate to be used if a compressor failure occurs. The alternate compressor is activated by positioning the SUIT FAN switch on the center panel. The cartridge of lithium hydroxide and activated charcoal remove carbon dioxide and odors of an organic nature that could have any ill effects on the pilots. As suit circuit oxygen flows through the suit heat exchanger, the temperature is controlled as selected by the pilots. Solid traps, located in the oxygen outlet ducts of both pilots' suits, remove particulate solids, preventing contamination of the suit circuit system. An integral by-pass opens if the traps become choked with collected solids permitting continuous oxygen flow through the suit circuit.
 * CABIN LOOP
 * SUIT LOOP

ORBIT
Normal cabin leakage allows the cabin pressure to decay to a nominal value of 5.1 psia. The cabin pressure control valve maintains this value automatically. A dual cabin pressure regulator supplies makeup oxygen through the pilots' pressure suits to the cabin on demand, as sensed by two aneroid elements within the regulator. The regulator supplies the makeup oxygen at a controlled pressure between 5.0 to 5.3 psia. The cabin fan on spacecraft 5 and 6 circulates cabin air through the cabin heat exchanger. The cabin fan has been removed on spacecraft 8 through 12. One or both of the pilots may open their faceplates. The cabin air circulating valve is in the open position to provide for recirculation of the cabin oxygen through the suit circuit. In the event of spacecraft depressurization, whether intentionally or by spacecraft puncture, the dual cabin pressure regulator closes when cabin pressure decreases to 4.1 +0.2 -0.1 psia, preventing excessive loss of oxygen. The suit circuit demand regulator senses cabin pressure and maintains suit circuit pressure at 2.5 to 3.5 inches of water below to 2 to 9 inches of water above cabin pressure. Should cabin pressure decrease below 3.5 psia, the suit circuit demand regulators maintain the suit circuit pressure at 3.5 +0.4 -0.0 psia by constant bleed orifices and sensing aneroids within the regulator. When cabin pressure is restored to 5.1 +0.2 -0.1 psia, the suit circuit demand regulators return to normal operation. In the event of cabin and suit circuit malfunction, the suit circuit will automatically revert to the high rate of operation when suit circuit pressure decreases below 3.0 -0.0 psia. A suit circuit pressure sensing switch energizes the solenoid of the dual high flow rate and system shutoff valve. This initiates a high oxygen flowrate of 0.08 +-0.008 lb/min (total flow: 0.16 lb/min). This high flow rate flows directly into the suits bypassing the suit demand regulators. The suit recirculating system is shut off and the suit compressors are de-activated when the solenoid of the dual high flow rate and system shutoff valve has been energized. The O2 HI RATE light on the center panel illuminates when the suit circuit is on the high flow rate. There is also a manual control for the high flow rate and system shutoff valve located on the center console. When the suit circuit pressure is restored to a level above 3.0 +0.1 -0.0 psia, the high rate and system shutoff valve is reset manually by using the control marked O2 HIGH RATE RECOCK located on the center panel. This returns the suit circuit to normal operation by opening the system shutoff valve and closing the high rate valve. The suit compressor is also reactivated. The drinking water system is pressurized and manually controlled by the pilots. Water from the adapter supply is used to replenish the cabin tank water supply. The water tank drink selector valve is set in the NORM position. The pilots manually operate the drinking dispenser to provide drinking water from the cabin storage tank. The water separator removes metabolic moisture through a wicking material positioned between the plates of the suit heat exchanger. The Chemical Urine Volume Measuring System (CUVMS) has a four position selector valve labeled URINATE, SAMPLE, DUMP, and BY-PASS. Position selection is made by rotating the selector valve which is attached to a multi-ported center plug. The selector valve includes a positive displacement tracer metering pump which is activated by the handle as it passes over a plunger between the BY-PASS and URINATE positions. This supplies a quantity of tracer solution to the passages to mix with the urine in the collection/missing bag. Therefore, volume measurement can only be accomplished when the selector is in the URINATE position and the tracer chemical is added. The urine receiver assembly and collection bag provides for collecting and sampling urine or overboard dump provisions but does not provide for volume measurement. The urine quick-disconnect hose and filter assembly consists of a section of flexible hose with quick disconnect couplers on each end. This assembly connects the CUVMS or the urine receiver assembly to the water management panel through an in-line urine solids trap filter for dumping urine overboard. The urine sampling bags are plastic laminate with a valve that connects to the sampler port on the CUVMS or the urine receiver assembly for taking urine samples. The urine receiver roll-on cuff is the interface between the crewmen and the CUVMS or the urine receiver assembly. It provides an air and liquid tight seal for direct urine transfer. Selector valve positions on the CUVMS direct urine flow as follows. The SAMPLE position directs urine and tracer chemical mixture from collection/mixing bag to the sample port and into the sampling bag. The DUMP position directs the urine/tracer chemical mixture overboard. The BY-PASS position (normal purge position) directs the urine flow to the water management panel for dumping overboard. The dump selector valve on the water management panel is positioned to route the urine either to the water boiler or dumped overboard. The normal procedure is to dump. Before it is dumped the urine dump system is preheated by positioning its heater switch located on the water management panel. A urine dump heater light is also provided and located on the water management panel. This light illuminates when the heater is activated.
 * CABIN LOOP
 * SUIT LOOP
 * WATER MANAGEMENT SYSTEM
 * URINE DISPOSAL SYSTEM

RE-ENTRY
The primary oxygen system is disconnected when the adapter section is separated from the re-entry module. This removes the primly oxygen supply pressure which automatically activates the secondary oxygen supply. The system shutoff and high rate valve is positioned to the high rate position before the adapter section is Jettisoned. The pressure in the suit and cabin remains constant at 5 psia (nominal) until an altitude of approximately 27,000 feet is reached. As ambient pressure increases during descent, the cabin pressure relief valve admits ambient air into the cabin, preventing high differential pressures. The cabin pressure relief valve begins to open when the ambient pressure is 15.0 inches of water greater than cabin pressure and opens to maximum flow when the pressure differential is 20 inches of water. At an altitude of 25,600 feet, or below, the pilots manually open the cabin inflow and outflow valves to circulate external air through the cabin and suit circuit. Maximum negative pressure on the cabin should not exceed 2 psi as controlled by the cabin relief valve. Prior to re-entry the face plates should be closed. The high flow rate of oxygen is flowing directly into the suit circuit. When the cabin inflow valve is opened It activates the suit compressor and external air is circulated through the suit circuit.
 * OXYGEN SYSTEM
 * CABIN LOOP
 * SUIT LOOP

POST-LANDING
Ventilation is provided by the suit compressor as long as electrical power is available (12 hours minimum). Ambient air is drawn into the vehicle through the snorkel inflow valve, by the suit compressor, circulated through the suit circuit into the cabin, then discharged overboard through the outflow vent valve. The snorkel inlet valve functions as a water check valve. When the snorkel inlet valve is above water level t the ball check is retained freely in a wire mesh cage, permitting ambient air to enter the suit circuit. Normal oscillations of the spacecraft in the sea may result in the snorkel valve being momentarily submerged. This will cause the ball check to seat and is held there by suction from the suit compressor. Opening the cabin circulating valve allows the ball to drop from its seat. To prevent water from entering the cabin through the cabin pressure relief valve, the manual shutoff section of the valve is closed.

EMERGENCY
If cabin depressurization becomes necessary due to toxic contaminants or fire, the cabin outflow valve is opened to depressurize the cabin. The cabin regulator will close, stopping the oxygen supply to the cabin, permitting the escape of toxic contaminants and preventing oxygen assistance to combustion in the event of fire. The cabin repressurization valve permits repressurization of the spacecraft cabin. The control knob for the cabin repressurization valve is located on the lower console and is rotated counterclockwise to open the valve. It is rotated clockwise to close the valve when cabin pressure is between 4.3 and 5.3 psia. Cabin pressure is then automatically controlled at 5.1 +0.2 -0.1 psia by cabin pressure regulator valve. This system is installed on spacecraft 5 and 6 only. Operation of the egress oxygen system is initiated by three of the four lanyards which are pulled when the seat leaves the spacecraft. One lanyard pulls a pin in the composite disconnect allowing it to separate and close the normal suit circuit. Two of the remaining lanyards open the container shutoff valve and circuit relief valve activating the egress oxygen system. Each of the egress oxygen containers is pressurized to 1800 psig with gaseous oxygen. The oxygen flows from the containers through a pressure regulator, where the pressure is reduced to 40 psia. It then flows through a shutoff valve and a flow restrictor, which allows a flow of 0.052 to 0.063 lb/min, then through a check valve to the suit. After leaving the suit, oxygen flows through the shutoff and relief valve, which dumps the oxygen overboard, as well as controls the suit pressure to 3.5 -0.0 +0.6 psia if ejection occurs at an altitude above 31,500 feet, and 2 to 8.23 inches of water above ambient at an altitude below 31,000 feet.
 * CABIN LOOP
 * EGRESS OXYGEN

DUAL SECONDARY OXYGEN RATE AND SUIT SYSTEM SHUTOFF VALVE
The dual secondary oxygen rate and suit system shutoff valve provides a constant flow rate of oxygen directly to the pilot's suit during re-entry or in the event the suit circuit malfunctions during launch or orbit. The valve is designed for manual and automatic initiation. The recirculating suit oxygen circuit flows through the shutoff section of the valve, which is manually opened and is spring loaded to the closed position. The shutoff valve is held open by a 24 vdc solenoid pin, as long as the solenoid is de-energized. The secondary flow poppet valve, held closed by spring tension, remains closed whenever the shutoff valve is in the open position. When the solenoid is energized, the butterfly arm is released and rotates by spring tension, closing the suit circuit valve and mechanically opening the secondary oxygen flow rate poppet valves. Opening the poppet valves allows oxygen to flow to each pilot's suit through fixed orifices at a rate of 0.08 +-0.008 lb/min per man (total flow 0.16 lb/min.). The butterfly arm simultaneously actuates a switch that de-energizes the solenoid, turns off the suit compressor and cabin fan, and illuminates a SECONDARY FLOW RATE lamp on the pilots' center display panel. A pressure sensor switch attached to each pilot's suit circuit will energize the solenoid if the suit circuit pressure in either suit decreases below 3.0 +0.1 -0.0 psia, automatically shutting off, the suit circuit flow and initiating the secondary flow rate. A manual control is provided for resetting the valve to the normal position. The secondary flow rate is used during re-entry. Prior to retro-grade the pilots manually disengage the solenoid initiating the secondary flow rate.

SUIT OXYGEN DEMAND REGULATOR
The suit oxygen demand regulator controls the oxygen to the suit circuit from the primary or secondary oxygen system and replenishes oxygen used by the pilots or lost by leakage. Cabin pressure is sensed on one side of the diaphragm and suit pressure is sensed on the opposite side of the diaphragm. The differential pressure across this diaphragm opens or closes a poppet valve admitting or stopping oxygen flow into the suit circuit. With cabin pressure of 5.0 psia the suit regulator maintains suit pressure at 2.5 to 3.5 inches of water below cabin pressure. A resilient diaphragm type valve relieves pressure in the suit during ascent and limits excess pressure to between 2.0 and 9.0 inches of water above cabin pressure. During descent, the suit demand regulator relieves the secondary oxygen rate flow through the relief portion of the valve, maintaining suit pressure 2 to 9 inches of water above cabin pressure. A constant bleed and aneroid elements maintain the suit pressure at 3.5 +0.4 -0.0 psia. if cabin pressure decreases below this pressure. The bleed flow by-passes the tilt valve through a bleed orifice and is directed to the cabin pressure sensing side of the pressure demand diaphragm. A metering valve, controlled by an aneroid, regulates the reference pressure on the demand diaphragm. The regulator returns to normal operation when cabin pressure returns to 5.1 +0.1 -0.2 psia. In the event that cabin decompression and a ruptured relief diaphragm in the regulator occur simultaneously, an aneroid over the relief diaphragm extends to control suit pressure at 3.9 psia maximum.

CABIN PRESSURE RELIEF VALVE
The cabin pressure relief valve automatically controls the cabin-to-ambient differential pressure during launch, orbit and re-entry. Duplicate spring loaded poppet valves are controlled by servo elements within the valve. The servo elements control spring loaded metering valves which determine the pressure within the diaphragm chamber behind the poppet, controlling the poppet position. A small inlet bleed orifice admits cabin pressure to the diaphragm chamber. When the poppet opens, a large orifice permits rapid change in pressure ensuring quick closure of the poppet. During ascent the valve will relieve cabin pressure as ambient pressure decreases until cabin differential pressure is 5.5 to 6.0 psia. The valve closes maintaining differential pressure in this range. When cabin pressure decreases below 5-5 psia the servo element closes the metering valve maintaining reference pressure within the diaphragm chamber at cabin pressure. The poppet is held closed by spring force and the zero differential between the diaphragm and the cabin prevents cabin pressure release. If cabin differential pressure exceeds 5.5 psia the zero element retracts, opening the metering valves, allowing the diaphragm chamber to discharge to ambient. The discharge port being larger than the inlet bleed orifice permits the diaphragm chamber to approach external pressure. The cabin pressure reacting on the diaphragm overrides the poppet spring force, which opens permitting cabin pressure relief to ambient. During descent, as external pressure increases, ambient air is admitted to the cabin by the valve to reduce the differential pressure. As external pressure increases above the cabin pressure the metering valves are held on their seats, preventing external pressure from entering the diaphragm chamber and retaining cabin pressure in the chamber. The poppet valve senses diaphragm chamber pressure versus ambient pressure. When the ambient pressure is 15 inches of water greater than cabin pressure the poppet begins to open permitting ambient air to enter the cabin. The poppet opens fully when the differential pressure is 20 inches of water. To preclude water entering the cabin during postlanding, a manual shutoff valve is provided.

SUIT CIRCUIT COMPRESSOR
Two electric motor driven, single stage compressors are incorporated in the suit circuit. One compressor is utilized for circulation of the gases within the suit circuit, supplying both suits. The other compressor functions as a backup and is activated only by manual selection by the pilots. Either compressor can be manually selected by a switch on the center display panel, and both compressors can be selected simultaneously. When secondary oxygen flow rate is selected, the compressor is automatically deenergized. Re-entry is made using the secondary rate. At an altitude of 25,600 feet or below the manual inflow valve is opened which re-energizes the compressor. The suit compressor provides ventilation during landing and for a twelve hour postlanding period, or until the batteries fail.

SOLIDS TRAP
A solids trap is located in the oxygen outlet duct of each suit. A cylindrical 40 micron filter strains the gaseous flow in the suit circuit removing the solid matter. In the event that the trap becomes choked with collected solids, an integral by-pass opens when the differential pressure across the screen exceeds 0.50 inches of water.

DUAL CABIN PRESSURE REGULATOR
The cabin pressure regulator maintains cabin pressurization by providing makeup oxygen to the cabin on demand. The regulator contains two aneroid elements which individually sense cabin pressure. When cabin pressure decreases, the aneroids expand, forcing metering pins open and permitting oxygen flow into the cabin, maintaining cabin pressure at 5.1 +0.2 -0.1 psia. If the cabin is punctured or develops leakage greater than the flow capacity of the valve (4.79 + 0.48) 10-3 lb/min, oxygen flow to the cabin is stopped when the cabin pressure decreases to 4.0 +0.2 -0.1 psia, by the aneroids expanding enough to cause the metering to close off the oxygen.

PRIMARY SUPERCRITICAL OXYGEN CONTAINER
The primary oxygen container is a double walled tank. A dual concentric cylinder, quantity measuring devices, heaters and heat transfer spheres are internal to the container. The tank contains two heaters. The first is a 12.0 +- watt heater which is activated either manually by a switch located on the center panel 3 or automatically by a pressure switch. The pressure switch controls the activation of the heating element in the tank to automatically maintain the cryogen in a supercritical state. The switch de-energizes the heater circuit when the pressure in the tank is between 875 to 910 psig, and closes the circuit 15 to 75 psig below the opening pressure. The second heater is a 325 +50 -0 watt heater manually controlled by a switch located on the overhead switch/circuit breaker panel. The pressure relief valve maintains the oxygen pressure within the container at 1000 +0 -55 psig, and prevents overpressurization of the containers. Provisions for servicing the primary oxygen container from a ground supply source of oxygen are provided.

SECONDARY OXYGEN CONTAINER
The secondary oxygen container is a cylindrical shaped container, having a useful oxygen capacity of 6.5 pounds at an operating pressure of 5000 psig.

SYSTEM DESCRIPTION
The spacecraft cooling system consists basically of two identical temperature control circuits functioning independently of each other to provide the cooling requirements for the spacecraft. Each cooling circuit consists of a pump package, thermostatic and directional control valves, various type heat exchangers, radiators, filters, and the necessary plumbing required to provide a closed circuit. The cooling system may be operated in either the primary and/or secondary circuit, and is capable of carrying maximum heat loads in either circuit. The equipment coldplates, cabin and suit heat exchangers are located in the reentry module. The upper radiator panels are located in the retrograde section. The pump package, battery coldplates, filters, electronic equipment coldplates, ground launch cooling and regenerative heat exchangers and the lower radiator panels are located in the adapter equipment section. System manual controls are located on the pilots' pedestal console and the control switches, warning lights and indicators are located on the center panel. The cooling systems in spacecraft 8 through 12 are provided with a means of bypassing the coolant around the fuel cells rather than through them. This provision is for ground operation when fuel cells are not in use. During orbital flight, Monsanto MCS-198 coolant is supplied throughout the cooling system and thermostatic control valves regulate the coolant temperature. Temperature sensors, located in the system, provided the necessary telemetering of system temperatures to ground stations.

SYSTEM DISPLAYS AND CONTROLS
Cooling system displays and controls are located on panels in the cabin as shown in section 3 and function as specified below. Dual concentric knobs are mounted between the ejection seats for suit and cabin temperature control. These knobs control the operation of valves regulating the flow rate of primary and secondary coolant through the suit and cabin heat exchangers for spacecraft 5 and 6. Spacecraft 8 does not have the cabin heat exchanger. Clockwise rotation results in increased temperatures. A dual indicator provides for monitoring temperatures in the suit and cabin circuits. Range markings are calibrated in degrees Fahrenheit. These switches are connected to the coolant pumps power supplies, one switch for each power supply. Each switch has two positions; ON and OFF. The switches are located on the center panel. On spacecraft 8 through 12 B pump switch in each loop changes the flow rate from 183 1b/hr to 140 lb/hr. Pump lights ill-m4uate when the pumps are activated. They are located above their respective switches near the top of the center panel. The RES LO lights illuminate when the coolant level in the reservoir is low. This light illuminates when pressure in the evaporator builds up to 4.0 +0.0 -0.3 psig and is extinguished when the pressure falls to 3.1 +0.3 -0.3 psig. This switch is connected to the evaporator heater and is used to heat the water in the evaporator before dumping.
 * SUIT AND CABIN TEMP CONTROLS
 * CABIN AND SUIT TEMP INDICATOR
 * PRIMARY AND SECONDARY PUMP SWITCHES
 * PRIMARY AND SECONDARY PUMP LIGHTS
 * EVAP PRESS INDICATOR
 * EVAP PRESS HEATER SWITCH

SYSTEM OPERATION
The cooling circuit in which the cooling system operates is dependent upon the temperature loads generated by the equipment, spacecraft phase of flight and the temperature within the spacecraft cabin. Cooling is provided throughout the mission up to pre-retrograde firing. At this time the coolant pump packages are jettisoned with the adapter equipment section, terminating spacecraft cooling. Spacecraft 5 and 8 through 12 require both loops to be operated continuously. In spacecraft 6 the primary circuit operates continuously provides the required cooling during low temperature loads. The secondary circuit is used, in conjunction with the primary circuit, during phases of high temperature loads; namely launch, rendezvous, and pre-retrograde. Under normal heat loads, the number i pump in the primary circuit provides the required cooling. Under peak heat loads, the number 1 pump in the secondary circuit is used with the primary circuit number 1 pump to provide maximum cooling. In the event of a number 1 pump malfunction in either circuit, the number 2 pump in that circuit is used. In the event of both pumps failing in one circuit, both pumps of the remaining circuit can be used to provide the required cooling. (Spacecraft 6 does not have the number 2 pump in either circuit.

PRE-LAUNCH
During pre-launch an external supply of Monsanto MCS-198 coolant is circulated through the spacecraft ground cooling heat exchanger providing temperature control of the cooling system coolant. The number i pumps of the primary and secondary cooling circuits are activated, using an external power source, to provide the required cooling for spacecraft equipment and cabin. The spacecraft radiator switch, located on the center panel, is placed in the BYPASS position so the cooling system coolant by-passes the radiators and is directed through the ground cooling heat exchanger. Coolant is circulated through each coolant loop by a positive-displacement gear pump. Spacecraft 5, and 8 through 12 are provided with 2 pumps in each loop. Spacecraft 6 has only one pump in each loop. Selection of loops and number of pumps is controlled manually. The coolant is filtered as it leaves the pump, and simultaneously flows to the inlet of the battery coldplate or fuel cell temperature control valve and primary oxygen heat exchanger. The temperature control valve maintains the cooling temperature at the fuel cell or battery coldplate inlet at 75° +2° -4° F. Temperature increasing above setting will reduce by-pass flow. Coolant temperature from by-pass line varies from 80°F to 165°F. Coolant temperature from equipment lines varies from 60°F to 125°F. Coolant enters the primary oxygen heat exchanger and then is routed around the steam discharge lines in the water boiler before it passes through the regenerative heat exchanger. It then passes through the selector and pressure relief valve. This selector valve is electrically actuated and when in the radiator by-pass position allows the coolant to pass through the ground cooling heat exchanger where the external supply of coolant flowing through the ground cooling heat exchanger absorbs the heat from the spacecraft's coolant system. The ground coolant heat exchanger has an airborne flow capacity of 336 lb/hr, per coolant loop, at 125°F. It has a ground coolant flow capacity of 425 lb/hr at 40°F. The coolant is now ready to pass through the temperature control valve. This valve maintains the outlet temperature at 40° +2° -4° F. If the coolant entering the valve from the ground heat exchanger is below this range, a portion of the coolant is directed through the regenerative heat exchanger and then mixed at the valve. The coolant then flows through the water evaporator to the cabin and suit manual temperature control valves. These valves meter the coolant flow through the cabin and suit heat exchangers. The evaporator selector valve relief portion allows part of the coolant to by-pass the cabin and suit heat exchangers depending on the setting of the manual control valves. The selector portion of this valve allows the by-pass fluid to come from either downstream or upstream of the evaporator. The coolant continues through the various coldplates until it reaches the battery coldplates for spacecraft 6 or through the fuel cells on spacecraft 5, and 8 through 12. The coolant has now returned to the reservoir where the cycle is ready to be repeated. Shortly before launch, the external cooling and electrical power are disconnected.

LAUNCH
During launch the launch cooling heat exchanger goes into operation in the following sequences. The heat transfer characteristics and capabilities of the ground cooling heat exchanger no longer exist. The Monsanto MCS 198 coolant fluid now with no place to dissipate its internal heat, which is constantly being generated by and absorbed from the loop components, circulates about the temperature control valve of the heat exchanger. When the coolant temperature exceeds 46 +4° -2° F the temperature control valve opens to pressurize a donut shaped bellows which unseats the poppet valve exposing the water in the heat exchanger core to reduced pressure as altitude increases during launch. When spacecraft altitude exceeds 100,000 feet, water in the heat exchanger will boil absorbing heat from the coolant. This absorbed heat is then expelled overboard in the form of steam. When the coolant reaches a temperature of 46°F the temperature control valve repositions to relieve pressure to the donut shaped bellows holding the poppet open. As this pressure diminishes, a spring behind the poppet will reposition it to the closed position. The evaporator selector valve is positioned to allow all flow to go through the evaporator. The water boiler water reservoir is constantly replenished from the suit heat exchanger water separator, and if the need arises, from the drinking water supply tank.

ORBIT
After orbiting for approximately 30 minutes, to allow the radiator to cool after being subject to launch heating, the coolant flow is directed through the space radiators by manual selection of the radiator switch located on the center panel. This by-passes the ground cooling heat exchanger. The evaporator selector valve is also positioned so that only the flow to the suit and cabin heat exchangers pass through the evaporator. Prior to retrograde firing, the coolant pump packages, radiators, batteries and various heat exchangers are jettisoned with the adapter equipment section. Prior to adapter jettisoning and retrograde firing the number i coolant pumps for both the primary and secondary coolant circuits are activated. The suit, cabin, and equipment bays are cooled to as low a temperature as possible, before the adapter equipment section is jettisoned.

PUMP PACKAGE
The pump package for each coolant circuit incorporates two constant displacement electrical pumps, two pump inverters, an external reservoir, filters, relief and check valves. The pump package is located in the adapter equipment section. Pump selection is provided by switches on the pilots' center panel. A pump failure warning light is provided on the center panel. When a pump is activated the coolant flows from the reservoir to the pump, which circulates the coolant through the cooling circuit. The coolant returns to an external reservoir that compensates for thermal expansion, contraction, and leakage of the coolant. A 100 micron filter downstream of the pump prevents contamination of the cooling system. Check valves in the pump package prevent the operating pump from pumping coolant into the redundant pump. Flow sensing switches illuminate a pump failure lamp on the pilots' center panel in the event of pump failure.

RADIATOR
The spacecraft radiator consists of two circumferential radiator panels made of 0.25 inch diameter cooling tubes. There are four sections of tubing to each radiator panel. The tubing is manufactured as part of the spacecraft structure. Each panel incorporates two parallel cooling circuits, one for the primary cooling circuit and the other for the secondary circuit. During orbit the cooling system coolant is circulated through the radiator. The heat of the coolant radiates into space, lowering the temperature of the coolant.

COLDPLATES
The coldplates, other than the battery coldplates, are plate fin constructed units incorporating parallel coolant system passages. Coldplates are fabricated from aluminum. Battery, electrical, electronic and other heat generating components are mounted on coldplates. The coolant flowing through the coldplates absorbs the heat generated by the components2 preventing overheating of the operating equipment.

HEAT EXCHANGERS
Two types of heat exchangers are used in the spacecraft; namely, plate fin constructed and shell and tube constructed heat exchangers. The suit, cabin, water evaporator, ground cooling and regenerative heat exchangers are of plate fin construction. The primary oxygen heat exchanger is of shell and tube construction. The coolant absorbs heat from the cabin, suit and regenerative heat exchangers. The ground cooling and water evaporator heat exchangers permit heat transfer to cool the coolant. The primary oxygen heat exchanger is designed so heat transfer will heat the primary oxygen to a desired temperature.

TEMPERATURE CONTROL VALVE
Temperature control valves are provided in both the primary and secondary cooling circuits. These valves are located at the radiator outlets and at the inlets to the battery coldplates or fuel cells. The temperature control valve located in the coolant system radiator outlet automatically maintains the coolant outlet temperature at 40 +3° -4° F as long as the radiator capacity has not been exceeded. The temperature control valve located in the battery coldplate inlet automatically maintains the coolant inlet temperature at 75 +3° -4° or above. The temperature control valve contains a piston that regulates the inlet flow to the valve. The piston is spring loaded on one side. A thermostatic actuator on the opposite side of the piston determines piston movement, which in turn regulates the coolant flow through the valve. The thermostatic actuator, which is located to accurately sense mixing temperature, consists of an encapsulated wax pellet that expands or contracts as temperature varies. As temperature around the pellet increases, the wax expands exerting pressure on the diaphragm. The diaphragm moves a piston, which in turn controls the inlet flow to the valve. Temperature reduction around the wax decreases the pressure in the pellet cup allowing the spring to reposition the piston regulating the flow of coolant through the valve.

LAUNCH COOLING HEAT EXCHANGER
The launch cooling heat exchanger is located in the adapter section. Via its relief valve it can dump liquids overboard, or if the temperature control valve senses temperature greater than 50°F, it can control the outlet temperature of the primary and secondary coolants to 46 +4° -2° F. In addition, it serves as a water reservoir, storing water until it is needed for cooling. This evaporator consists of a wicking type heat exchanger and is capable of storing seven pounds of water. A temperature control valve has been set to control the outlet coolant temperature to 46° +4° -2° A relief valve opens and allows excess water to be dumped overboard at 2.75 =- 0.75 psi differential and reseats at 2.0 psi differential minimum. An electrical heater is provided in the poppet to prevent ice formation. Coolant flow capacity is 366 lb/hr at 40°F. Water flow capacity is 3 lb/min when cooling is not required from the evaporator. Maximum operating pressure in the fluid heater coolant circuits is 230 psig, and 100 psig in the core circuits. Maximum operating pressure in the water circuit is 20 psig with exit port relief valve in normal operation. The steam exit duct is continuously heated by coolant coming from the primary oxygen heat exchanger to prevent ice formation. A loss of pressure in either coolant loop will not affect the operation of the valve.

Section VIII: Guidance and Control System
REFER TO THE SEDR 300 CONFIDENTIAL SUPPLEMENT FOR INFORMATION CONCERNING THE GEMINI GUIDANCE AND CONTROL SYSTEM

SYSTEM DESCRIPTION
The Communication System is the only communication link between the ground and the Gemini Spacecraft. The system has the following capabilities: radar tracking of the spacecraft; two-way voice communications between the ground and the spacecraft, and between the crew; ground crew; ground command to the spacecraft; Instrumentation System data transmission; and postlanding and recovery aid data transmission. To make possible these various capabilities, the Communication System contains components that may be divided into the following categories: antennas, including multiplexers and coaxial switches; beacons; voice communications; telemetry transmitters; flashing recovery light; and Digital Command System. The flashing recovery light and the uhf recovery beacon are grouped together in a category called the Electronic Recovery Aids (ERA). The Communication System components are located throughout the spacecraft with the largest concentration being in the right equipment bay of the re-entry module and the electronic module of the adapter equipment section.

ANTENNAS
Eight antennas and one antenna system provide transmission and/or reception capabilities for the various Communication System components. The spacecraft Communication System contains the following antennas: uhf recovery; uhf stub; uhf descent; two uhf whips; two hf whips; C-band annular slot; and a C-band antenna system consisting of a power divider, a phase shifter, a phase shifter power supply, and three helical antennas. To achieve the most efficient antenna usage, a diplexer and a quadriplexer are used with the uhf whips and the uhf stub antenna. The multiplexers make it possible to use more than one transmitter and/or receiver with a single antenna. Five coaxial switches permit antenna and transmitter/receiver switching for best communication coverage during the various phases of the mission (launch, orbit, re-entry and recovery).

BEACONS
Four beacons in the Communication System establish the capability of locating and tracking the spacecraft during the mission. The four beacons are: An acquisition aid beacon and a recovery beacon used to locate the spacecraft, and two C-band beacons used to track the spacecraft. The acquisition aid beacon, operating on a fixed frequency, is used to determine when the spacecraft is within the range of a ground tracking station, and provides information for orientating the ground station antennas during the orbital phase of the mission. The recovery beacon is a transmitter that operates on the international distress frequency, and is used by the recovery forces to determine the spacecraft location. The C-band beacons are transponders which, when properly interrogated by a ground station, transmit signals for accurate spacecraft tracking. During the recovery phase of the mission, emergency communications may be established by connecting one of the uhf rescue beacon transceivers to the uhf recovery antenna. The rescue beacon transceivers are Government Furnished Equipment (GFE), stowed in pilot's survival kits.

VOICE COMMUNICATION
Voice communications is maintained by one hf and two uhf transmitter/receivers and the Voice Control Center (VCC). The V_C has ell the necessary controls and switches required for various keying modes, transmitter/receiver selection, squelch, volume control, and voice recording. The hf voice transmitter/receiver may also be used for Direction Finding (DF) purposes during the postlanding phase of the mission. An intercom connector is available for communication between the crew and recovery team, prior to opening the spacecraft hatches during the recovery phase of the mission. Lightweight headsets are supplied for use when the spacesuit helmets are removed during orbit, or during postlanding if the helmets or entire spacesuit is removed prior to recovery.

TELEMTRY TRANSMITTERS
Receiving inputs from the Pulse Code Modulated (PCM) programmer and the on-board tape recorder, three telemetry transmitters transmit vital spacecraft system parameters to the ground stations. The three transmitters operate on different frequencies and are identified as real-time, delayed-time, and stand-by transmitters. The stand-by transmitter is only used in case of real-time, or delayed time transmitter failure.

FLASHING RECOVERY LIGHT
The flashing recovery light, used during the postlanding phase of the mission, contains its own power supply and improves visual spacecraft location.

DIGITAL COMMAND SYSTEM
The Digital Command System (DCS) is the command link between the ground and the spacecraft. The DCS consists of two uhf receivers, a decoder, and two relay packages and Is operational from pre-launch until adapter equipment section separation. Basically, the DCS receives and decodes two types of commands: a discrete or Real-Time Command (RTC) for spacecraft equipment utilization, and Stored Program Commands (SPC) that supply digital information to various spacecraft systems. Real-time commands operate DCS relays that control power directly or energize relays in the spacecraft Electrical System that determine equipment usage. Stored program commands are received and decoded for use by the Time Reference System (TRS), or the computer.

SYSTEM OPERATION
The Communication System is semi-automatic in operation. The sequence and theory of operation of the Communication System is described in the following paragraphs. Individual components are described in System Units.

VOICE TAPE RECORDER
Voice tape recordings are made during the mission by placing the RECORD switch on the VCC to the CONT or MGM position. The TONE VOX, AUDIO & UHF T/R i and 2 circuit breakers must be in the ON position. Each tape cartridge allows approximately one hour of recording time and is easily changed. An end-of-tape light on the voice recorder illuminates for two seconds when two minutes of recording time remains on the tape. The end-of-tape light will remain on when the end of the tape is reached. A digital timing signal is applied to one channel of the tape for time correlation of the voice recording.

PRE-LAUNCH
During pre-launch the BEACONS-C circuit breaker is placed to the ON position to arm the C-RNTY and C-ADPT BEACON CORTROL switches. The C-RNTY switch is placed in the CONT position during pre-launch to enable the re-entry C-band beacon to reply when properly interrogated by a ground station. The C-ADPT switch is placed in the CMD position during pre-launch. The CMD position enables the ground station during launch, to activate the adapter C-band beacon via a DCS channel if the need arises. After the adapter C-band beacon is activated, it will reply when properly interrogated by a ground station. The C-band antenna system, used with the re-entry C-band beacon, is energized when the ANT SEL switch is placed in the RNTY position. The ART SEL switch is armed when the COAX CBTL circuit breaker is positioned to ON. The ANT SEL switch controls application of power to the phase shifter power supply in the C-band antenna system. The number 1 voice transmitter/receiver will be utilized during pre-launch unless some malfunction occurs in which case the number 2 transmitter/receiver can be selected. For operation of either uhf voice transmitter/receiver standby power is applied through the AUDIO & UHF T/R circuit breakers i and 2, which must be in the ON position. The selected transmitter/receiver will be powered by placing the UHF select switch to the number 1 or number 2 position and the MOI switch of number 2 AUDIO to the UHF position. The UHF select switch also controls coaxial switch 1 which connects the uhf transmitter/receiver to the quadriplexer. Coaxial switch power is obtained from the common control bus through the ON position of the UHF RELAY circuit breaker. The method of keying the transmitter/receiver is selected by positioning the KEYING switch on the VCC to VOX (voice operated relay), PTT (push-to-talk), or CONT INT/PTT (continuous intercom/push-to-talk transmitter keying). The desired antenna usage is obtained by placing the ANT CNTL circuit breaker to the ON position. This places coaxial switch 3 to position 1; thus connecting the quadriplexer to the IN position of the coaxial switch 5. Coaxial switch 5 was placed in position 1 when the ANT CNTL switch was placed in the ANTY position during the C-band beacon operation. With coaxial switch 5 in position 1, the uhf stub antenna is available for voice transmission and reception. Prior to umbilical release, voice communication is maintained between the spacecraft and the ground complex through a hard-line using the headset and microphone amplifiers of the VCC. After umbilical release, voice transmission to the ground complex is accomplished by means of the uhf voice transmitter/receiver. The real-time telemetry transmitter will be operating during the pre-launch phase of the mission. The real-time telemetry transmitter is powered by placing the RT AMTR circuit breaker in the ON position and placing the TM CONTROL switch to the R/T & ACQ position. The real-time telemetry transmitter uses the uhf stub antenna via the quadriplexer and coaxial switches 3 and 5, the same as the uhf transmitter/receiver. In case of real-time telemetry transmitter failure, the stand-by telemetry transmitter may be used for real-time transmission. The STBY XMTR CNTL and PWR circuit breakers must be in the ON position to operate the stand-by telemetry transmitter. If selection is made by the crew, the STBY TM CONTROL switch is placed to the R/T position. Selection can be made by a ground command via the DCS when the TM CONTROL switch is in the OFF position. When operating as the real-time telemetry transmitter, the stand-by transmitter uses the stub-antenna for transmission. The following Communication System components will be non-operational during the pre-launch phase of the mission. To assure the off condition of these components, the following switches should be in the position specified below: To assure proper sequential actuation of the various communication components, the following circuit breakers (in addition to those previously described) must be placed to the position listed prior to launch:
 * C-BAND RADAR BEACONS
 * UHF TRANSMITTER/RECIVER
 * REAL-TIME TELEMTRY TRANSMITTER
 * NON-OPERATIONAL COMPONENTS
 * HF (ON VCC)               OFF
 * BEACON-CONTROL - RESC OFF
 * HF ANT                        OFF
 * WHIP ANTENNAS        HF
 * WHIP ANTENNAS        UHF
 * WHIP ANTENNAS        DIPLEX
 * HF T/R                            ON
 * BEACONS                      ACQUIRE
 * BEACONS                      RESC
 * XMTRS        DT
 * TAPE RCDR    CNTL

SPACECRAFT/LAUNCH VEHICLE SEPARATION
Equipment usage after spacecraft/launch vehicle separation is identical to that described under Pre-Launch except for the following: Upon closure of any two of the three spacecraft separation sensors the acquisition aid beacon is energized. The uhf whip antenna solenoid actuators are powered and release the latch mechanism of the uhf whip antennas, allowing them to self extend. The acquisition aid beacon transmits via the diplexer and uhf whip antenna on the adapter equipment section. Placing the TAPE RCDR-CNTL circuit breaker to ON and the TM CORTROL switch to R/T & ACQ during pre-launch places coaxial switch 2 in position 1 which connects the acquisition aid beacon to the diplexer.

ORBIT
During orbit, operation of the telemetry transmitters and beacons will normally be controlled by ground commands via DCS channels. To operate from ground commands the C-ADPT, C-RNTY and T/M CONTROL switches must be in the CMD position. During orbit hf communications is via the hf whip antenna on the adapter retrograde section. At insertion the adapter hf whip is extended by placing the LANDING switch to the SAFE position, and the HF ANT switch to the EXT position. This will place coaxial switch in position 2 and allow hf voice transmission and reception via the adapter hf whip. After extension (approximately one minute) the HF ANT switch is returned to the OFF position. Stand-by power is applied to the hf transmitter/receiver by the SF T/R circuit breaker which was positioned to ON during pre-launch. The hf transmitter/receiver is powered by positioning the HF select switch (on the VCC) to RNTY and audio MODE switch i or 2 to the HF position. The method of keying the hf transmitter/receiver is selected by positioning the KEYING switch on the VCC to VOX, PTT, or COBT IRT/PTT. During orbit, any of the three keying modes may be selected. The uhf voice transmitter/receiver operation is identical to that described under Pre-Launch with the following exception. Preferred antenna usage during orbit for uhf transmission and reception is via the adapter retrograde uhf whip antenna. The retrograde uhf whip antenna is selected by placing the ANT SEL switch to the ADPT position which places coaxial switch 5 to position 2. Although preferred uhf transmission and reception is via the retrograde uhf whip antenna, the uhf stub antenna may be used during orbit by placing the ANT SEL switch to the RRTY position. The acquisition aid beacon operates continuously during the orbital phase of the mission except when the delayed-time telemetry transmitter is operating. When the ground station receives the acquisition aid beacon signal, it initiates a DCS command for the delayed-time telemetry transmitter to transmit data stored by the on-board recorder while the spacecraft was between ground stations. Delayed-time transmission may also be initiated by placing the T/M CONTROL switch to the R/T-D/T position. This will initiate real-time as well as delayed-time telemetry transmission. Real-time and delayed-time transmission will normally be initiated from the ground station via DCS channels. At the time the delayed-time telemetry transmitter is selected, the acquisition aid beacon is turned off and coaxial switch 2 is placed in position 2, allowing telemetry transmission via the diplexer and uhf whip antenna on the adapter equipment section. As the spacecraft goes out of range the delayed-time telemetry transmitter is turned off and the acquisition aid beacon resumes transmission. This is normally performed by the ground station but may be accomplished by placing the T/M CONTROL switch to the CMD, or the R/T & ACQ position. If the R/T & ACQ position is selected, the delayed-time transmitter is turned off and the real-time transmitter and the acquisition aid beacon begin transmitting. If the CMD position is selected, only the acquisition aid beacon will operate; however, the ground station has the capability of energizing the real-time telemetry transmitter via a DCS channel. Any of the three previously described methods of disabling the delayed-time telemetry transmitter will place coaxial switch 2 to position I, and allow acquisition aid beacon transmission via the diplexer and uhf whip antenna. The stand-by telemetry transmitter may be used for delayed-time transmission should failure of the delayed-time telemetry transmitter occur. The stand-by transmitter is switched to delayed-time transmission by a ground command via a DCS channel (if the STBY CONTROL switch is in the OFF position), or by placing the STBY TM switch to the D/T position. Delayed-time transmission via the stand-by telemetry transmitter uses the ,h_ stub or the uhf whip antenna on the retrograde adapter depending upon the setting of the ANT SEL switch. Orbital operation of the real-time telemetry transmitter is similar to that of the delayed-time telemetry transmitter in that the real-time telemetry transmitter is operated only during the period that the spacecraft is within range of a ground station. The real-time telemetry transmitter is turned on by a DCS command from the ground station or by placing the T/M CONTROL switch to the R/T & ACQ or RT-D/T position. Real-time transmission is by the uhf stub or the retrograde uhf whip antenna, depending upon the position of the ANT SEL switch. In case of failure of the real-time telemetry transmitter, the standby transmitter may be used for real-time transmission. The stand-by transmitter is switched to real-time transmission by a ground command via a DCS channel (if the STBY TM CONTROL switch is in the OFF position), or by placing the STBY TM CONTROL switch to the R/T position. The stand-by telemetry transmitter transmits via the uhf stub or the retrograde uhf whip antenna, depending upon the position of the ANT SEL switch. It should be noted that the stand-by telemetry transmitter may be used for delayed-time, or real-time transmission, but may not be used simultaneously for both. In the event that both the real-time and delayed-time transmitters fail, it is up to the ground station to determine the purpose for which the stand-by transmitter will be used. During orbit, the C-band beacons are used only while the spacecraft is within range of a ground station. Normally, the adapter C-band beacon will be used during stabilized flight and the re-entry C-band beacon used during roll maneuvers. Operation of the beacons is similar to that described under Pre-Launch. The C-RNTY and C-ADPT BEACON CONTROL switches are normally kept in the CMD position. When the spacecraft comes within range of a ground station, as determined from the acquisition aid beacon signal, power to the desired C-band beacon is applied by ground command via a DCS channel. The desired beacon may also be selected by placing the C-ADPT or C-RNTT BEACON CONTROL switch to the CONT position. After power is applied, the selected C-band beacon will transpond when properly interrogated by a ground station. When the re-entry C-band beacon is selected, the ANT SEL switch should be placed in the RNTY position to energize the phase shifter and provide optimum radiation coverage.
 * HF VOICE TRANSMITTER/RECEIVER
 * UHF VOICE TRANSMITTER/RECEIVER
 * DELAY TIME TELEMETRY TRANSMITTER
 * REAL-TIME TELEMETRY TRANSMITTER
 * C-BAND RADAR BEACONS

ADAPTER SEPARATION
Prior to adapter equipment section separation, the re-entry module antennas are selected by placing the T/M CONTROL switch to R/T & ACQ, the ART SEL switch to RNTY/,and the C-RNTY BEACON CORTROL switch to the CONT position. Transmission and reception during re-entry is via the C-band antenna system and the uhf stub antenna. The acquisition aid beacon will operate until it is jettisoned with the adapter equipment section. The hf voice communications is disabled by placing the HF select switch to the OFF position. On spacecraft 5, the hf whip on the adapter retrograde section will remain extended. On later spacecraft, the hf whip may be retracted by holding the HIP ANT switch in the RET position for approximately i.5 minutes for complete retraction. The following communications components will be jettisoned with the adapter section: This limits telemetry data transmission to real-time, voice communication to uhf, and tracking data to the re-entry C-band beacon. Following equipment section separation and retro firing, retrograde section separation will occur at which time the retrograde uhf whip and the adapter hf whip antenna will be jettisoned.
 * Digital Command System (DCS)
 * Delayed-time telemetry transmitter
 * Diplexer
 * C-band annular slot antenna
 * Adapter C-band radar beacon
 * Diplexer uhf whip antenna
 * Acquisition aid beacon
 * Coaxial switch 2

RE-ENTRY
During the re-entry phase of the mission, two short duration communication blackout periods occur. The first period, from approximately 1310 seconds after retrofire time (Tr) to 1775 seconds after Tr, is caused by an ionization shield around the spacecraft. This ionization is due to the extremely high temperatures created upon re-entry into the earth’s atmosphere. The second blackout period occurs at Rendezvous and Recovery (R & R) section separation when the uhf stub antenna is Jettisoned. This period is terminated at two-point suspension which occurs shortly after main parachute deployment. At R & R separation, energized parachute deploy time delay relays energize coaxial switch 3, placing it to position 2. This makes the uhf descent antenna available for real-time telemetry transmission and uhf voice communications. At two-point suspension the uhf recovery and uhf descent antennas are automatically extended. The uhf recovery beacon is turned on by placing the RESC BEACON CONTROL switch to the W/O LT position. Antenna usage during re-entry is as follows: Prior to R & R separation, real-time telemetry transmission and uhf voice communication is via the uhf stub antenna. After two-point suspension, the uhf descent antenna is used instead of the uhf stub. The re-entry C-band beacon and C-band antenna system is used for tracking and the uhf recovery beacon will use the uhf recovery antenna.

LANDING THROUGH RECOVERY
Upon impact the main parachute is Jettisoned by actuating the PARA JETT switch. This extends the flashing recovery light. The light is energized by engaging the RESC BEACON CONTROL switch from the W/O LT position to the ON position. The re-entry C-band beacon and real-time telemetry transmitter is turned off by placing the C-RNTY BEACON CONTROL and the T/M CGRTROL switch to the CMD position. If the stand-by telemetry transmitter was selected for real-time transmission, the stand-by transmitter will be turned off by placing the STBY TM CONTROL switch to the OFF position. The recovery hf whip antenna is extended by placing the HF ANT switch to the PST LDG position for spacecraft 5, or on later spacecraft by holding the HF ANT switch in the EXT position for approximately one minute. Voice communication via the hf transmitter/receiver is then possible by placing the HF select switch to the RRTY position and either MODE switch to HF. The hf transmitter/receiver can also be used to transmit a direction finding signal by placing either MODE switch to HF/DF. During the recovery phase of the mission the uhf rescue beacon transceiver may be connected to the uhf recovery antenna. The uhf recovery beacon can be turned off by positioning the RESC BEACON CONTROL switch to OFF. Lightweight headsets are provided to replace the spacesuit helmets if the helmets or spacesuits are removed and the crew remains inside the spacecraft. A recovery team disconnect is used for intercom conversation between the crew and recovery team prior to opening the spacecraft hatches.

ANTENNAS
Purpose: The uhf descent antenna is used for simultaneous transmission of the real-time and stand-by telemetry transmitters, and transmission and reception for the uhf voice transmitter/receiver. The uhf recovery antenna provides transmission capability for the uhf recovery beacon. The two antennas are used from two-point suspension of the main parachute through final recovery of the spacecraft. Physical Characteristics: Both antennas are mounted in the parachute cable trough where they are stowed until main parachute two point suspension during the landing phase of the mission. The antenna element consists of two one-half inch wide gold plated steel blades bolted together at two places. The uhf descent antenna is approximately 18 inches long. The uhf recovery antenna is approximately 18 inches long. Mechanical Characteristics: For rigidity, the antenna element is shaped in a 0.5 inch wide arc having a radius of 1.5 inches. The two laminations of steel blades, compounding a single antenna element, are rigidly secured at the lower half of the antenna. To allow a slight displacement of the two laminations with respect to each other during stowage and deployment, two nuts and bolts placed through elongated holes secure the two laminations together at the upper half of the antenna element. The antennas are bent towards the small end of the spacecraft for stowage and are held in place by a retaining strap. The strap is broken when the Landing System shifts from single point to two point suspension, allowing the antennas to extend. Each of the two antennas have a radiation pattern which is identical to that of a quarter wave stub. Purpose: The uhf stub antenna allows simultaneous transmission of the real-time and stand-by telemetry transmitters, transmission and reception for the uhf voice transmitter/receivers, and reception for DCS receiver number 2. The antenna may be used from pre-launch until separation of the R & R section during re-entry, but is normally used from pre-launch to insertion and from re-entry preparation to R & R section separation. Physical Characteristics: The uhf stub antenna is mounted in the nose of the R & R section. The antenna protrudes forward from the R & R section and is covered by a nose fairing during the boost phase of the mission. The antenna consists of a mast and base which weighs approximately 1.1 pounds. The mast is constructed of 3/4 inch cobalt protection during re-entry. The antenna is approximately 13.5 inches long including the connector, and 1.25 inches in diameter over the ablation material. The mast consists of two sections. The front section is mounted on a cobalt steel ball Joint and retained to the rear section by a spring loaded cable. Electrical contact between the mast sections is made through the ball joint and the spring loaded cable assembly. The ball joint allows the front section of mast to be deflected to approximately 90 degrees in any direction around the antenna axis. The spring of the cable assembly is pre-loaded to approximately 45 pounds to cause the front section, when deflected, to return to the erected position. The rf connector is press fitted into a socket and makes contact to the mast through the socket and sleeve, which are the same material as the mast. The shell of the rf connector is mounted to the base which is isolated from the mast by a Teflon spacer and sleeve. Mechanical Characteristics: The uhf stub is a quarter wave length antenna. The radiating length of the antenna is approximately 11.2 inches. Purpose: Two identical uhf whip antennas supply the required ,uhf transmission and reception facilities during orbit. One of the uhf antennas is located on the adapter equipment section and serves the DCS receiver number 1, and the acquisition aid beacon or delayed-time telemetry transmitter. The second uhf antenna, mounted on the adapter retrograde section, serves the real-time and standby telemetry transmitters, the uhf voice transmitter/receivers, and DCS receiver number 2. Physical Characteristics: The uhf whip antenna is self extendable and requires no power other than that required for initial release. The antenna element is a tubular device made from a 2 inch wide beryllium copper strip processed in the form of a tube. The antenna, when fully extended, forms an element that is approximately 12 inches long and i/2 inch in diameter. During stowage, the tube is opened flat, wound inside of a retaining drum, and latched in position. Upon release of the latch by a solenoid, the extension of the antenna depends entirely on the energy stored in the rolled strip material. This energy is sufficient to erect the antenna at a rate of 5 feet/second into its tubular form. In the stored condition, the antenna is flush with the outer skin of the spacecraft. Mechanical Characteristics: The antenna element is retained inside the housing by a metal lid. A metal post Is attached to the lid and passes through the center of the coiled antenna. The bottom of the post is grooved to accept a forked latch which holds the catch post assembly firmly in position prior to release. The forked latch is attached to a miniature pull-solenoid which is spring loaded in the extended position to ensure that launch shock and vibration loads will not cause inadvertent antenna extension. Then a voltage from the Sequence System is applied to the antenna solenoid, the latch will be withdrawn allowing the antenna cap to eject and the antenna to extend. As the catch post assembly is ejected, a microswitch in series with the solenoid coil opens the circuit to the coil to prevent further current drain from the power source. The two antennas are Jettisoned with the corresponding adapter section. Purpose: The hf whip antennas provide transmission and reception for the hf voice transmitter/receiver during the orbital and postlanding phases of the mission. The recovery hf whip antenna is mounted on the small pressure bulkhead, outside the pressurized area of the spacecraft re-entry module. The other hf whip antenna is located on the adapter retrograde section. The antenna mechanism housing, approximately 6.25 inches wide and 22.5 inches high, completely encloses all parts of the antenna, including storage space for the antenna elements. The recovery hf whip antenna contains six elements which, when fully extended, comprises a single antenna mast approximately 13 feet 3 inches long. The adapter hf whip antenna contains three elements which, when fully extended comprise a single antenna mast approximately 16 feet long on spacecraft 5 and 6, and approximately 13 feet long on later spacecraft. The mast is one inch in diameter on all spacecraft. Two connectors, supported by the antenna body, provide a means of applying power and connecting the antenna to the rf connector on the hf voice transmitter/receiver. The recovery hf whip antenna weighs approximately 9.0 pounds. The 16-foot version of the adapter hf whip antenna weighs approximately 7.5 pounds and the 13-foot version 6.0 pounds. The main supporting structure of the antenna mechanism housing is the antenna body consisting of a thin fiberglass shell. The outer shell is made in two sections which mate together and form a completely sealed envelope around all moving parts. The antenna mast elements are heat treated stainless steel strips and are stored in adc motor driven cassette. Mechanical Characteristics : The strip material comprising the antenna elements is heat treated into a material circular section in such a manner that the edges of the material overlap approximately 180 degrees. When the antenna is retracted, the tubular elements are continuously transformed by guide rollers into a flattened condition, and stored in a strained manner in a cassette. Extension and retraction of the antenna is accomplished by a motor which, by means of a chain, drives the storage cassette core. Because of the natural physical shape of the antenna elements, the antenna has a tendency to self-extend; thus giving an extension time of approximately 25 seconds. Retraction time is approximately 40 seconds. The antenna is stopped within its desired limits by two microswitches, one for extension and one for retraction, which automatically cut the power applied to the motor at the time of extreme limits of the antenna are reached. The rf connection to the antenna is obtained by a wiper arm sliding on the cassette core drive shaft. On spacecraft 5 the hf whip antennas are operated as follows : Spacecraft control bus voltage is supplied through the WHIP ANTENNAS -HF circuit breaker to the HF ANT switch. The adapter hf whip antenna is extended during orbit by positioning the HF ANT switch to EXT. The adapter hf whip antenna is not retracted during orbit, but is Jettisoned in the extended position with the retrograde section. After landing, the recovery hf whip antenna is extended by positioning the HF ANT switch to PST LDG, and is retracted by positioning the HF ANT switch to EXT. On spacecraft 6 through 12, extension of the hf whip antennas is controlled through the HF ANT switch and LANDING switch. The hf antennas are operated as follows: Spacecraft control bus voltage is supplied through the WHIP ANTENNAS - HF circuit breaker to the HF ART switch, which has momentary type contacts. During orbit, the LANDING switch is in the SAFE position and adapter hf whip antenna can be extended or retracted by holding the HF ANT switch in the EXT or RET position respectively. During re-entry, the LANDING switch is placed in the ARM position. After landing, the recovery hf whip antenna can be extended or retracted by holding the HF ANT switch in the EXT or RET position respectively. The HF ANT switch should be held in the EXT position for approximately one minute for full extension of the antennas, and in the SET position for approximately 1.5 minutes for full retraction. Purpose: The C-band Annular slot antenna serves the adapter C-band radar beacon and is normally used during stabilized flight. Physical Characteristics: The C-band annular slot antenna is mounted on the equipment section of the adapter. The physical construction is such that the antenna is flush with the outer skin of the spacecraft. The antenna is approximately 1.5 inches in diameter, 1.35 inches long and weighs 8 ounces maximum. Mechanical Characteristics: _he antenna radiation pattern is identical to that of a quarter wave stub on a ground plane. The antenna is used for both reception and transmission of the adapter C-band radar beacon during the orbital phase of the mission. The antenna is jettisoned with the equipment section of the adapter. Purpose: The C-hand ante_ system, consisting of a power divider, a phase shifter, and three helical antennas, provides transmission and reception capability for the re-entry C-band radar beacon. The power divider supplies equal transmission power to the three helical antennas. A phase shifter is in series with one of the antennas to compensate for areas of low or no radiation coverage between lobes of the three individual radiation patterns. A phase shifter power supply supplies the phase shifter with 26 vac 453 cps power. The antenna system gives the circular radiation pattern around the spacecraft longitudinal axis required for ascent, descent and roll spacecraft attitudes. Physical Characteristics: The power divider, phase shifter, and phase shifter power supply are mounted on the small pressure bulkhead, outside the pressurized area of the spacecraft. The power divider measures approximately 3.86 inches over the connectors, 4.0 inches over the tuning knobs and weighs approximately 6.5 ounces. The phase shifter is approximately 5.8 inches long, 2.8 inches wide at the large end, 1.4 inches high, has a diameter at the small end of about 1.5 inches, and weighs approximately 12 ounces. The phase shifter power supply measures approximately 1.5 inches wide, 1.75 inches high, 3.5 inches long over the connector and weighs approximately 8 ounces. The three C-baud helical antennas are mounted flush with the outside skin of the spacecraft and spaced approximately 120 degrees apart. Each antenna unit is approximately 3.4 inches long, 1.8 inches wide, has a depth of 2.21 inches over the connector and weighs approximately 3.5 ounces. Electrical Characteristics : The power divider, phase shifter, and helical antennas comprise an antenna system that satisfies the transmission and reception requirements for the re-entry C-band radar beacon during the launch and re-entry phases of the mission. The power divider is basically a cavity type power splitter. During beacon transmission, power is delivered to the power divider where it is divided equally among the C-band helical antennas. The power divider compensates for loss of power due to the phase shifter in series with the right antenna. The power divider contains a double stub tuner to compensate for mismatch between the re-entry C-band beacon, the C-band helical antennas, and the phase shifter. Tuning is accomplished by means of a self-locking tuning shell located underneath each tuning stub cap. The phase shifter has its own ac power supply. The input to the phase shifter, is half wave rectified and applied across a coil wound around a ferrite material. Due to the characteristics of the ferrite material, the rf signal from the power divider is delayed 0 to 180 degrees +- 20 degrees at the rate of 453 cycles per second. The changing phase shift of the rf power on one of the C-band helical antennas with respect to the other two, shifts the lobe of that antenna by approximately +- 45 degrees; thus giving the effect of an almost ideal circular radiation pattern around the longitudinal axis of the spacecraft. The combination of the three antenna elements gives a radiation pattern which extends in all directions except forward and aft of the spacecraft. The phase shifter power supply is a dc-ac inverter which supplies a nominal 26 vac, 453 cps power to operate the phase shifter. The power supply is a hermetically sealed solid-state unit consisting of a voltage regulator, single-stage oscillator, buffer stage, and a push-pull output stage with transformer coupled output. The power supply provides a minimum output of 21 volts cps at 453 +- 17 cps with an input voltage range from 20 to 30 vdc. Input voltage is applied from the spacecraft main bus via the BEACON-C circuit breaker, C-RNTY BEACON CONTROL switch and the RNTY position of the ANT 8EL switch. Maximum input current is 370 m111iamperes. Purpose: The uhf diplexer provides isolation between DCS receiver number 1, and the acquisition aid beacon or the delayed-time telemetry transmitter operating into a common antenna. The uhf quadriplexer provides isolation between the standby telemetry transmitter, the real-time telemetry transmitter, a uh voice transmitter/ receiver, and DCS receiver number 2 operating into a common antenna via coaxial switches. Physical Characteristics: The diplexer is located on the electronic module of the adapter equipment section. The quadriplexer is located forward of the small pressurized bulkhead outside the pressurized area of the cabin. The diplexer is approximately 4.5 inches wide, 4 inches high, and 2.7 inches deep; contains two input and one output connectors, and weighs approximately 1.25 pounds. The uhf quadriplexer is approximately 5.75 inches wide, 5.5 inches deep, and 4.1 inches high; weighs approximately 2.75 pounds, and has four input and one output connectors. Electrical Characteristics: Each channel consists of a high Q cavity, tuned to the corresponding operating frequency. All channels are isolated from each other without appreciably attenuating the rf signals passing through it. Each channel can be re-tuned if the assigned operating frequency is changed. The diplexer isolates DCS receiver number l, and the acquisition aid beacon or the delayed-time telemetry transmitter, depending upon the position of coaxial switch number 2. The diplexer operates into the uhf whip antenna on the adapter equipment section. The uhf quadriplexer isolates the real-time telemetry transmitter, the stand-by telemetry transmitter, one of the two uhf voice transmitter/receivers, and DCS receiver number 2. The quadriplexer operates into one of the following three uhf antennas, depending on the position of the coaxial switches in series with the antennas: uhf stub antenna, uhf descent antenna, or the uhf whip antenna on the adapter retrograde section. Purpose: Five coaxial switches are used to perform the following functions : (1) select the acquisition aid beacon or the delayed-time telemetry transmitter output as the input to the diplexer; (2) select one of the two uhf voice transmitter/receiver outputs as the input to the quadriplexer; (3) connect the hf voice transmitter/receiver to the adapter hf whip antenna on the retrograde section, or to the recovery hf whip antenna on the re-entry module; (4) connect the output of the quadriplexer to the uhf descent antenna or through coaxial switch 5 to the uhf stub or the retrograde adapter uhf whip antenna. Physical Characteristics: The location of the switches is as follows: Each switch contains a power connector, an input connector, two output connectors, and weighs approximately 0.5 pounds. The dimensions of each switch are approximately 2.65 inches long, 1.82 inches high, and 1 inch wide. Electrical Characteristics: The five coaxial switches are identical and may be used interchangeably. Basically, the coaxial switches supply single pole double throw switching action. The switch, having a 20 millisecond maximum operation time, operates on 3 amperes at 28 vdc and uses a latching solenoid break-before-make switching action. The coaxial switches are designed to operate from 15 mc to 500 mc, and from 5500 mc to 5900 mc. Pins D and E of each switch are brought out to AGE test points to permit monitoring of the switch positions prior to lift-off. Pins A and B of each switch are utilized to accomplish the switching action.
 * UHF DESCENT AND UHF RECOVERY ANTENNAS
 * UHF STUB ANTENNA
 * UHF WHIP ANTENNAS
 * HF WHIP ANTENNAS
 * C-BAND ANNULAR SLOT ANTENNA
 * C-BAND ANTENNA SYSTEM
 * MULTIPLEXERS (UHF DIPLEXER AND UHF QUADRIPLEXER
 * COAXIAL SWITCHES
 * Coaxial switch I: approximately five inches from the small end of the cabin, in the fourth quadrant.
 * Coaxial switch 2: approximately 10 inches from the forward (small) end of the adapter equipment section, in the third quadrant.
 * Coaxial switch 3: approximately i0 inches from the small end of the cabin, in the third quadrant.
 * Coaxial switch 4: located adjacent to coaxial switch 1
 * Coaxial switch 5: approximately 7 inches from the small end of the cabin, in the third quadrant.

BEACONS
Purpose : The re-entry C-band radar beacon provides tracking capability of the spacecraft from lift-off to insertion and from retrograde to landing. The re-entry C-band beacon may be used during roll maneuvers or in the event of adapter C-band beacon failure. Physical Characteristics : The re-entry C-band radar beacon is a sealed unit which measures approximately 7.64 x 6.14 x 3.02 inches and weighs about 8.3 pounds. Located on the rear of the beacon are various adjustments for transmitter, preselector, and local oscillator tuning. Solid-state modular circuitry is used throughout the beacon with the exception of the transmitter magnetron and the local oscillator. The beacon is mounted on the right forward equipment bay, and uses the C-band antenna system for reception and transmission. Electrical Characteristics: The re-entry C-band radar beacon is a transponder which upon reception of a properly coded interrogation signal from a ground radar tracking station, transmits a pulse modulated signal back to the tracking station. By measuring the elapsed time between transmission and reception at the tracking stations, and compensating for the time delay of the beacon, the position of the spacecraft can be determined. The signal arriving at the antenna is routed through the directional coupler to one half of a dual ferrite circulator. The ferrite circulator isolates the transmitter from the receiver, allowing a single antenna system to be used for both reception and transmission. The beacon utilizes a superhetrodyne receiver which is tunable, by means of a three stage preselector, over a range of 5600 mc to 5800 mc. The assigned receiver center frequency is 5690 me. The output of the preselector is combined with the local oscillator frequency in the crystal mixer to produce an output intermediate frequency of 80 inc. The local oscillator is of the metal-ceramic triode cavity type. The mixer contains a ferrite circulator for isolation between the local oscillator, mixer and preselector. The output of the mixer is amplified by three tuned intermediate frequency amplifier stages, followed by a video detector and a video preamplifier. Additional amplification is obtained by a pulse amplifier whose output is supplied to the decoder. The purpose of the decoder is to initiate triggering of the transmitter after a correctly coded signal has been received. The system delay, in conjunction with the delay variation correction circuitry, provides for a constant fixed delay used in determining the exact position of the spacecraft. The beacon incorporates a cw immunity circuit that prevents the transmitter from being triggered by random noise. The noise level is reduced below the triggering level of the transmitter by controlling the gain of the pulse amplifier. The transmitter uses a magnetron and provides a one kilowatt peak pulse modulated signal at a frequency of 5765 mc to the power divider. The beacon is powered by a dc-dc converter employing a magnetic amplifier and silicon controlled rectifiers. The converter provides voltage regulation for input voltage variations between 18 and 32.5 vdc. The input to the converter is filtered by a pi-type filter to minimize any line voltage disturbances. Purpose: The adapter C-band radar beacon provides tracking capability of the spacecraft during the orbital phase of the mission and is jettisoned with the adapter equipment section. Physical Characteristics : The adapter C-band beacon is a sealed unit and measures approximately 9.34 x 8.03 x 3.26 inches. The adapter beacon has a power and test connector, an antenna connector, and a crystal current test point connector. The beacon contains external adjustments for local oscillator, preselector (rf filter), and transmitter tuning; switches for selecting the desired interrogation code, and one of two preset transponder fixed delay times. These adjustments and switches are accessible by removing pressure sealing screws. The beacon employs solid-state circuitry, except for the transmitter magnetron and receiver local oscillator. The adapter beacon is located on the electronic module of the adapter equipment section and uses the C-band annular slot antenna for reception and transmission. Electrical Characteristics: The adapter C-band radar beacon is a transponder, which employs the same basic operating principles as the re-entry C-band beacon to provide spacecraft location data upon receipt of a properly coded interrogation signal. The interrogation signal is fed from the antenna to the duplexer. The duplexer is a ferrite circulator which couples the received signal to the rf filter preselector and also isolates the receiver from the transmitter to permit use of a common antenna for reception and transmission. The superhetrodyne receiver frequency is tunable from 5395 me to 5905 inc. The assigned operating center frequency is 5690 mc and is selected by adjustment of the rf filter. The rf filter is a three-stage preselector, employing three separately tuned coaxial resonator cavities to provide adequate rf selectivity and to protect the mixer crystal from damage due to transmitter power reflected by the antenna. The output of the preselector is combined with the local oscillator output in the mixer stage to provide a 60 mc output to the intermediate frequency amplifier. The mixer consists of a coaxial directional coupler and a mixer crystal. The directional coupler isolates the local oscillator output from the antenna and directs it to the mixer crystal. The local oscillator is a re-entrant cavity type employing a planar triode to generate the cw signal required to operate the mixer. The intermediate frequency amplifier is a high gain amplifier composed of an input stage, five amplifier stages, and a video amplifier. The amplified video output is fed to the pulse form restorer circuits which prevent a ranging error due to variations in receiver input signal levels, and also provides a standard amplitude pulse to the decoder for each input signal exceeding its triggering threshold. The decoder determines when a correctly coded signal is received and supplies an output to the modulator driver. The type code to he accepted is selected by the CODE switch. Single pulse, two pulse or three pulse codes may be selected. The modulator driver and control circuits initiate and control triggering of the transmitter modulator. The modulator driver supplies two fixed values of overall system delay. The desired delay is selected by the position of the DLY switch. An alternate value of maximum delay is available by removing an internal jumper lead. The modulator control furnishes the trigger and turn-off pulse for the modulator and limits modulator triggers to prevent the magnetron duty cycle from being exceeded, regardless of the interrogating signal frequency. The modulator circuit employs silicon controlled rectifiers which function similar to a thyratron, but require a much shorter recovery time. The associated modulator Pulse Forming Network (PFN) and transformer provide the necessary pulse to drive the transmitter magnetron. The desired pulse width is selected by the internal contains made to the PFN. The transmitter magnetron frequency is tunable from 5400 mc to 5900 mc. The assigned transmitter center frequency is 5765 inc. A minimum of 500 watts peak pulse power is supplied to the antenna under all conditions of rated operation. The transponder power supply consists of input line filters, a series regulator, and a dc-dc converter. The power supply furnishes the required regulated output voltages with the unregulated input voltage between 21 and 30 vdc. The converter employs a multivibrator aM full wave rectifier circuits. Purposes: Unlike the C-band beacons that supply accurate tracking data, the acquisition aid beacon is merely a transmitter used to determine when the spacecraft comes within range of a ground tracking station. When the spacecraft comes within the range of a ground tracking station, the acquisition aid beacon is disabled and remains off until the spacecraft is again out of range. Physical Characteristics: The acquisition aid beacon is cylindrical, having a diameter of approximately 2.6 inches, and a height of approximately 3.5 inches. The beacon contains a power connector, a coaxial antenna connector and weighs approximately 17 ounces. Electrical Characteristics: The acquisition aid beacon consists of a transmitter, dc-dc voltage regulator, and a low pass output filter. The transmitter is an all transistorized unit, containing a push-pull output stage to obtain a minimum output of 200 milliwatts at a frequency of 246.3 mc. The transmitter frequency is derived by taking the basic frequency of an oscillator and multiplying it through a series of tripler and doubler stages. The transmitter is powered by a de-de voltage regulator. The regulator is completely transistorized and supplies a regulated output voltage of 28 vdc. To reduce the probability of obtaining a spurious output signal, a band pass filter is placed in the output circuit. Purpose: The uhf recovery beacon, operating on the international distress frequency of 243 me, serves as a recovery aid by providing information regarding location of the spacecraft. Physical Characteristics: The beacon is mounted on the aft right equipment bay of the spacecraft re-entry module. The beacon is approximately 9.0 inches long, 4.0 inches wide, 2.5 inches high, and weighs 3.9 pounds maximum. The beacon contains one multipin power connector and one coaxial connector. Electrical Characteristics. The uhf recovery beacon consists of a spike eliminator, a regulator, a dc-dc converter, a pulse coder, a modulator, and a transmitter. Spacecraft main bus voltage is fed to the switching type regulator through the spike eliminator filter. The voltage regulator provides a dc regulated output voltage of 12 vdc to the dc-dc converter, the transmitter tube filaments, and the pulse coder. The dc-dc converter is a solid-state device providing two high voltage outputs to the transmitter and modulator. The pulse coder, a solid-state device, operates with the modulator to apply correctly coded high voltage pulses to the transmitter for plate modulation of the power amplifiers. The transmitter consists of an oscillator stage, a doubler stage, and a power amplifier. The transmitter power amplifier provides a uhf pulse coded output having a peak power of at least 50 watts to the uhf recovery antenna. An external rf band-pass filter is installed between the transmitter output and the antenna to reduce spurious rf radiation, especially at the uhf voice transmitter frequencies.
 * RE-ENTRY C-BAND RADAR BEACON
 * ADAPTER C-BAND RADAR BEACON
 * ACQUISITION AID BEACON
 * UHF RECOVERY BEACON

VOICE COMMUNICATION
Purpose: The Voice Control Center contains switches and controls for selecting the type of voice communication and the desired operating mode. The VCC also contains microphone and headset amplifiers, an alarm tone generator, and voice actuated transmitter keying circuitry. Five connectors located on the rear of the unit provide connection to the other voice communication system components and test connectors. The switches and controls of the VCC are located on the front panel. The number 1 and number 2 audio MODE switches are for selection of UHF, INT, HF, or HF/DF transmission. Below the MODE switches are three thumb-wheel-type multidetent volume controls, one for each of the above mentioned modes. In the center is the KEYING switch, a HF select switch, a UHF select switch, and thumb-wheel-type squelch controls for uhf and hf circuitry. The KEYING switch provides for selection of PTT, VOX, or CONT INT/PTT for the voice transmitters. The UHF and HF select switches provide capability of selecting the desired transmitter/receiver. The ADPT position of the HF select switch is not used. The record switch, lower right, permits recordings to be made in any mode of operation. Continuous (CONT) or Momentary (MOM) recording can be selected. The SILENCE switch, lower left, is to permit uninterrupted sleep during extended spacecraft missions. The NO. 1 position allows reception for both pilots. The NO. 1 position removes power from the command pilot's headset amplifiers and the NO. 2 position removes power from the pilot's headset amplifiers; thus, making reception impossible. Electrical Characteristics: The VCC contains two headset and two microphone amplifiers for each of the audio channels. An audio signal, from the microphone in the helmets or lightweight headsets, is amplified by two microphone amplifiers and then applied to the MODE switch. With the MODE switch in the HF position, the output of the microphone amplifiers is applied to the hf transmitter. When the MODE switch is in the INT position, the output of the microphone is applied to the four headset amplifiers, via the two INT volume controls. With the MODE switch in the HF position, the output of the microphone amplifiers is applied to the uhf transmitter. The uhf switch selects uhf transmitter number 1 or number 2 and also operates coaxial switch 1 to connect the selected transmitter output to the uhf quadriplexer. The desired keying mode is selected by a common KEYING switch. Three methods may be selected to key the voice transmitters. The VOX position enables keying of the selected transmitter at the instant the microphone has an output signal. The PTT position enables keying of the transmitter when either push-to-talk switch, on the suit disconnect cables of the attitude control handle, is depressed. The CONT INT/PTT position gives continuous intercommunication between the crew and push-to-talk keying for transmission from the spacecraft to the ground station. The VCC also controls the power supplies of the transmitter/receivers by means of ground switching. With the MODE switch in a position other than HF and the HF select switch in the RNTY position, a ground is supplied to the hf transmitter/receiver auxiliary power supply to power the hf receiver. With the HF select switch in RNTY and the MODE switch in the HF position, a ground is supplied to the hf transmitter/receiver main power supply to power the hf receiver and transmitter. The uhf circuitry operates on the same principle as the hf. The UHF select switch supplies power ground for the selected receiver. The MODE switch (UHF position) together with the UHF select switch, supplies a power return for the uhf transmitter and receiver. The HF/DF position of the MODE switch is used for direction finding purposes. With the MODE switch in HF/DF and the HF select in the RNTY position, the hf transmitter is modulated by a 1,000 cps tone which is utilized to determine spacecraft location. Purpose: Two uhf voice transmitter/receivers are provided for redundant line-of-sight voice communication between the spacecraft and the ground. Physical Characteristics: Both transmitter/receivers are identical and are mounted side by side in the forward right equipment bay of the re-entry module. Each transmitter/receiver is a modular constructed, hermetically sealed unit approximately 7.7 inches long, 2.8 inches wide, 2.4 inches deep and weighs approximately 3.0 pounds. Each unit has a multipin audio and power connector, and a coaxial connector. Electrical Characteristics: The uhf voice transmitter/receiver consists of a transmitter, receiver, and power supply. The transmitter consists of a crystal controlled oscillator, two rf amplifiers, a driver, and a push-pull power amplifier. All stages except the driver and power amplifier are transistorized. The transmitter is fixed-tuned at 296.8 mc and is capable of producing an rf power output of 3.0 watts into a 50 ohm resistive load. The transmitter is amplitude modulated by a transistorized modulator stage. The am superhetrodyne receiver is fully transistorized, is fixed-tuned at a frequency of 296.8 mc, and contains a squelch circuit for noise limiting. The squelch threshold is manually controlled. An automatic volume control stage is also incorporated to provide a constant audio output with input signal variations. The uhf voice transmitter/receiver is powered by two dc-dc converters comprising an auxiliary and a main power supply. Operating power for the two power supplies is limited by two circuit breakers located on the left switch/circuit breaker panel. One circuit breaker is provided for each unit. Actuation of the power supplies is accomplished by ground return switching through the Voice Control Center. If the UHF select switch is in the NO. 1 or NO. 2 position and the MODE switch is in a position other than UHF, a ground is supplied to the auxiliary power supply only, placing the transmitter/receiver into a receive condition. With the MODE switch in the UHF position, a ground is supplied to the main power supply, placing the selected uhf voice transmitter/receiver into a receive and transmit condition. It should be noted that when the uhf transmitter is keyed, the uhf receiver is disabled and uhf voice transmissions from the ground station can not be received. Purpose: The hf voice transmitter/receiver is provided to enable beyond the line-of-sight voice communication between the spacecraft and the ground. Physical Characteristics: The approximate location of the hf voice transmitter/receiver is in the forward right equipment bay of the re-entry module. The unit weighs approximately 62 ounces, is approximately 8.5 inches long, 3.3 inches wide, and 2.9 inches deep. One multipin audio connector and one rf connector are provided. Electrical Characteristics: Basically, the hf voice transmitter/receiver is electrically identical to the uhf transmitter/receiver except for the operating frequency and power output. The hf transmitter and receiver are fixed tuned to a frequency of 15.016 mc and the hf transmitter provides an rf power output of 5 watts. Actuation of the hf receiver and transmitter is accomplished through the VCC. If the HF select switch is in RNTY and the MODE switch is in a position other than HF, the hf transmitter/receiver is in a receive condition. With the MODE switch in the HF position, the hf transmitter/receiver is placed in a receive and transmit condition. When the hf transmitter is keyed, the hf receiver is disabled. Purpose: The voice tape recorder is provided so recordings can be made during the spacecraft mission. Physical Characteristics: The physical construction and approximate location of the voice tape recorder is shown in Figure 9-19. The voice tape recorder is located inside the cabin in a vertical position between the pilot's seat and the right-hand side wall on spacecraft 5 and 6. On spacecraft 8 through 12 the recorder is located on the left-hand side wall aft of the abort handle. The voice tape recorder assembly consists of the recorder, tape cartridge, and shock absorber mounting plate and is supplied as GFE equipment. The recorder is approximately 6.25 inches long, 2.87 inches wide, one inch thick, and weighs 30 ounces maximum without the tape cartridge. The shock absorber mounting plate is approximately 6.3 inches long, three inches wide, and weighs 20 ounces maximum. The tape cartridge is approximately 2.25 inches square, 3/8 inch thick, and weighs two ounces. The recorder contains a power connector and a signal connector located on the end. The recorder is retained in the shock mount by guides and two Allen-head bolts for easy removal. The door contains a red plastic lens so that light from the end-of-tape bulb is visible. A safety latch prevents accidental opening of the door. The door is opened by pressing down on the latch and sliding it sideways. When the latch is released, the spring loaded hinge causes the door to open, exposing the cartridge tab. Flat pressure springs on the door hold the inserted cartridge in place and maintains tape contact with the recorder head and end-of-tape contact. The tape cartridge is guided into the recorder by step rails on each side of the cartridge. When the recorder door is opened, a heavy tab on the cartridge springs up to provide easy removal. The cartridge contains approximately 180 feet of magnetic tape, a supply reel, make-up reel, and associated gears and clutches. Electrical Specifications: The recorder is a two-channel transistorized unit consisting of the cartridge hold-down mechanism, voltage regulator, voice amplifier, time signal amplifier, bias oscillator, motor drive circuit synchronous drive motor, speed reduction unit, capstan, magnetic record head, and end-of-tape circuit. When the tape cartridge is inserted and secured in the tape recorder, the pressure roller in the cartridge contacts the capstan and the tape is pressed against the record head and the end-of-tape contact. The voice tape recorder is energized by spacecraft main bus power applied through the TONE VOX circuit breaker and the CONT or MCN position of the RECORD switch on the VCC. The voltage regulator supplies 15 vdc to the motor drive circuits bias oscillator and amplifiers. With the VCC and recorder energized, voice signals from the microphones are applied through microphone amplifiers in the VCC to the recorder voice amplifier. The voice signal is amplified and applied to the lower record head for recording on the magnetic tape. The time channel receives a digital timing signal from a time correlation buffer in the TRS. The timing signal is amplified by the recorder time signal amplifier and applied to the upper record head for recording on the magnetic tape. Simultaneously with the voice or timing signal, a 20 kc bias current from the bias oscillator is applied to the recorder heads to make a linear recording. The motor drive circuit consists of a 133 cps oscillator, a driver and push-pull output stage used to drive the synchronous motor. Phase-shift capacitors are connected to one motor winding for self-starting. The motor speed of 8000 rpm is reduced through the speed reduction unit to a capstan speed of 122 rpm. The end-of-tape circuit is energized by conductive foil on the tape contacting the recorder head and end-of-tape contact, causing the end-of-tape light to illuminate. The end-of-tape light will illuminate for two seconds when two minutes of recording time remains on the tape. The light will remain on when the end-of-tape is reached. Recordings cannot be made when the light is illuminated. The pilot may remove the used tape cartridge, insert another cartridge and continue recording. Each cartridge provides approximately one hour of recording. The tape speed is approximately 0.6 inches per second.
 * VOICE CONTROL CENTER
 * UHF VOICE TRANSMITTERS/RECIVERS
 * HF VOICE TRANSMITTER/RECIVER
 * VOICE TAPE RECORDER

TELEMTRY TRANSMITTERS
Purpose: The three telemetry transmitters provide a radio frequency (rf) link from the spacecraft to ground communication facilities for transmission of various data obtained by the Instrumentation System. Physical Characteristics: The three telemetry transmitters are identical except for the operating frequency. The transmitters are approximately 2.75 inches high, 2.25 inches wide, 6.5 inches long, and weigh approximately 41 ounces. Each transmitter contains a dc power connector, an rf output power connector, and a video connector. Two of the transmitters are located in the right forward equipment bay of the re-entry module, the third is located on the electronic module in the adapter equipment section. Electrical Characteristics: The three telemetry transmitters are classified by their operating frequency or by their function. The real-time (low-frequency) telemetry transmitter operates at 230.4 mc. The delayed-time (mid-frequency) telemetry transmitter operates at a frequency of 246.3 mc. The stand-by (high-frequency) transmitter, operating at 259.7 mc, may be used for real-time or delayed-time transmission in case one of the transmitters fails. The telemetry transmitters are solid-state fm transmitters. After a 30 second warm-up, the transmitters are capable of continuous uninterrupted operation for 500 hours. Information is transmitted to the ground in digital format by deviating the carrier frequency to the higher frequency deviation limit to transmit a 1, and to the lower deviation limit to transmit a 0. The transmitters receive Non-Return to Zero (NRZ) PCM pulse trains from the PCM programmer and voice tape recorder. The real-time transmitter provides the ground monitoring stations with current real-time data at a rate of 51.2 kilobits per second. The delayed time transmitter provides the ground monitoring station with data stored on the tape recorder while the spacecraft was between ground stations. The delayed-time data is transmitted at a rate of 112.6 kilobits per second. The stand-by transmitter is used as backup for the real-time or delayed-time transmitters in event of a failure in either transmitter. Transmission of the real-time and delayed-time data provide essentially full-time coverage throughout the spacecraft mission. The transmitters can be energized by a command from the ground station via the DCS or by controls on the instrument panel. The oscillator-modulator consists of a video amplifier, crystal controlled oscillator, phase shift networks and buffer amplifiers. The oscillator frequency is modulated by the video amplifier output. The phase shift networks provide impedance matching of the crystal oscillator to improve signal linearity for large deviations of frequency. The buffer amplifiers increase signal levels and isolate the crystal circuit from the frequency multipliers. The times 12 multiplier and power amplifier consists of a buffer amplifiers, times 4 multiplier, power amplifier and times 3 multiplier which increase the carrier frequency and power to the desired output values. The power amplifier develops 6 to 7 watts of power at a frequency from 75 to 80 mc into the output tripler circuit. The bandpass filter is used to minimize spurious radiation at the output of the transmitter. The real-time, stand-by and delayed-time transmitters each contains filter with a different frequency bandpass. The filter has a minimum 3 dB bandwidth of 16 mc, an impedance of 50 ohms and a vswr of less than 1.5 to I. The rf output connector J3 is an integral part of the bandpass filter. The line filter prevents noise on the input power bus from affecting transmitter operation and prevents transients generated within the transmitter from feeding back to the input power bus. The multipin power connector J2 is an integral part of the line filter. The dc-dc converter is a completely encapsulated unit employing transistors, diodes and a transformer to provide regulated outputs of 30 vdc and 70 vdc from an unregulated input voltage of 18 to 30.5 vdc. The converter is a constant power input type, thus minimizing the heat dissipation caused by high voltage inputs.

FLASHING RECOVERY LIGHT AND POWER SUPPLY
Purpose: The flashing recovery light and power supply provide visual spacecraft location information. Physical Characteristics: The light is self-extended by a torsion spring. The plug applying power to the light is kept in place by a compression spring. The recovery light will be automatically extended at the time the main parachute is jettisoned. The flashing recovery light power supply is mounted in the cabin, aft of the ejection seats. The power supply is approximately 7 inches long, 4 inches wide, 3 inches deep and contains one connector. The flashing recovery light is approximately 1.25 inches wide, 0.75 inches thick, and 3.25 inches high, excluding tube and erecting mechanism. The overall length of the light and erecting mechanism is approximately 6.5 inches. Electrical Characteristics: The recovery light is automatically extended at main parachute jettison. The extended recovery light is energized by positioning the RESC BEACON CONTROL switch to ON. The power supply consists of a battery pack and converter. The battery pack consists of several mercury cells to comprise a power source of 6.75 vdc to a dc-dc converter whose output is fed to a voltage doubler and a capacitive network. The 450 vdc output of the voltage doubler is used to power the flashing light while the capacitive network in conjunction with a thyratron, provides trigger pulses to accomplish switching or flashing action of the light. The trigger pulses occur at a rate of 15 triggers per minute.

DIGITAL COMMAND SYSTEM
The DCS provides a discrete command link and a digital data updating capability for the computer and TRS. The discrete command link enables the ground to control radar tracking beacons, selection of telemetry transmitters, instrumentation data acquisition, and abort indications. The capability of digital data updating enables the mission control center to update the computer and _RS to bring about a controlled re-entry at a predetermined point, and allows timed shutdown of equipment controlled by DCS relays. The DCS consists of a receiver/decoder package and two relay boxes. The three components are located in the electronic module of the adapter equipment section. The receiver/decoder package is approximately 8 inches high, 8 inches wide, and 12 inches long. Both relay boxes are identical. Each relay box is approximately 2.25 inches wide, 5 inches high, and 3 inches deep. The combined weight of the receiver/decoder package and the two relay boxes is approximately 28 pounds. The receiver/decoder package contains two uhf receivers and a decoder while each of the two relay boxes contain eight relays. The DCS receives Phase Shift Keyed (PSK) signals composed of a reference and an information signal. The information signal is in phase with the reference for a logical 1 and 180 degrees out of phase with the reference for a logical 0; thus establishing the necessary requirements for digital data. The DCS receives two types of digital commands: Real Time Commands (RTC) and Stored Program Commands (SPC). RTC causes relays within the DCS to be actuated. Nine of the 16 relays available for RTC are utilized to perform the following functions: The remaining seven relays are not utilized and perform no mission function. DCS channel assignments for the nine functions listed above may be different on each spacecraft. When the spacecraft goes out of range of the ground station, equipment controlled by DCS channels may be shutdown by a signal applied from the TRS to reset the DCS relays. This condition is known as salvo. The DCS relays in one relay box may be reset by momentarily positioning the TAPE PLY BK switch to RESET. The ground station transmits a 30-bit message for SPC and a 12-bit message for RTC. Each bit consists of five sub-bits. The five sub-bits are coded to represent a logical 1 or 0. The first three bits of each message designate the vehicle address. If the vehicle address is not correct, the DCS will reset itself and will not accept the message. If the vehicle address is accepted the sub-bit code will be automatically changed for the remainder of the message to reduce the probability of accepting an improper message. The second three bits of each message designate the system address and identify the remainder of the message as being a RTC or one of the following SPC : computer update, TRS time to go (TTG) to Tr, or TRS TTG to equipment reset (Tx) If the message is a SPC, the last 24 bits will be a data word. If the SPC is a TRS TTG TO Tx command, the last eight bits are ignored by the TRS. In case of a computer message, six bits of the data word contains the internal computer address and the remaining 18 contains information. Since a RTC consists of 12 bits, the six bits following the system address contain a 5-bit relay number and a 1-bit relay set-reset discrete. The PSK modulation signals are i kc reference and a 2 kc information signal. The receiver output is the composite audio of the I kc and the 2 kc signals. The composite audio output is filtered to recover the 1 kc and the 2 kc signals. The phase comparator compares the 2 kc to the 1 kc signal. The output of the phase comparator is used to trigger a flip-flop to produce either a logical 1 or 0 sub-bit. The i kc reference signal is used to synchronize the DCS. The audio outputs of the two receivers are linearly summed in an emitter follower of the sub-bit detector module. The sub-bit detector converts the audio to sub-bits. The 5-stage shift register provides buffer storage for the output of the sub-bit code. When a proper sub-bit code exists in the shift register, the bit detector produces a corresponding 1 or 0 bit. The output of the bit detector is applied to the 24 stage shift register. The operation for RTC and SPC is identical up to the input to the 24 stage shift register. The sub-bit sync counter produces a bit sync output for every five sub-bits. The bit sync is used to gate the 24 stage shift register. When a message is received, the vehicle address is inserted into the first three stages of the 24 stage shift register. If the vehicle address is correct, the vehicle address decoder circuit will produce an output to the bit detector which changes the acceptable sub-bit code for the remainder of the message. The next three bits of the message, the system address, are inserted into the first three stages of the 24 stage shift register, displacing the vehicle address to the next three stages. The system address decoder circuit identifies the specific address and sets up the DCS to handle the remainder of the message. When the system address is recognized to be a RT% the message is inserted into the first six stages of the 24 stage shift register and the system address and vehicle address are shifted into the next six stages. The RTC selection circuit recognizes the first stage of the 24 stage shift register to be a relay set or reset function and will apply a positive voltage to all set or reset relay coils, as applicable. The RTC selection gates select the proper relay from the relay number stored in the 24-stage shift register and provides an output which applies a power return to the coil of the selected relay. When the system address is a SPC, the six address bits in the 24 stage shift registers are cleared and the remaining 24 bits of the message are placed into the register. Assuming that the system address recognizes a TRS TTG to Tr message, the data flow would be as follows: The TRS Tr isolation amplifier, in the interface circuit, will apply a READY pulse to the TRS. The READY pulse sets up the TRS to transfer TRS TTG to Tr data from the DCS. When the TRS is ready to accept the data, it sends 24 shift pulses, at the TRS data rate, to the TRS input of the DCS. The data in the 24 stage shift register is then shifted out of the register through the DCS data isolation amplifier to the TRS. The DCS operations for computer updating and TRS TTG to Tr messages are similar to TRS TTG to Tr operations. Salvo occurs when TRS TTG to Tx reaches zero. At Tx = 0, the TRS applies a signal to the TRS Tx input line of the DCS which causes the RTC selection circuits to reset the DCS relays. After a SPC or RTC has been carried out by the DCS, a verification signal is supplied to the telemetry system for transmission to a ground station. The DCS indicator, on the instrument panel, illuminates when a SPC is transferred to the appropriate system. Upon completion of data transfer or if the system to which the data was transferred fails to respond within 100 milliseconds, the DCS will reset in preparation for the next message. The DCS will also reset in the event of a timing error in transmission of data, or if the DCS power supply voltages become out of tolerance.
 * PHYSICAL CHARACTERISTICS
 * GENERAL DESCRIPTION
 * TYPES OF COMMANDS
 * (1) Select the stand-by telemetry transmitter for real-time transmission.
 * (2) Select the stand-by telemetry transmitter for delayed-time transmission.
 * (3) Select the real-time telemetry and acquisition aid beacon transmission.
 * (4) Select real-time and delayed-time telemetry transmission.
 * (5) Actuate the adapter C-band radar beacon.
 * (6) Actuate the re-entry C-band radar beacon.
 * (7) Illuminate the abort indicators.
 * (8) Actuate the playback tape recorder.
 * (9) Initiate calibration voltage for the PCM programmer.
 * MESSAGE FORMAT AND MODULATION
 * OPERATIONAL DESCRIPTION

SYSTEM DESCRIPTION
The Instrumentation System provides a means of data acquisition with respect to the performance and operation of the spacecraft throughout its mission. Data acquisition is defined as the sensing of specific conditions or eventson board the spacecraft, displaying the derived data from these inputs to the crew and ground operation personnel, and recording and later processing this data for use in post flight reports and analysis. In this respect the data acquisition function is shared by all spacecraft systems, the ground operational support system, and the data processing facility. Basically, the instrumentation Parameters are divided into two categories : operational and nonoperational. Operational parameters are those which are necessary for determing the progress of the mission, assessing spacecraft status, and making decisions concerning flight safety. Nonoperational parameters are those which are required for post mission ana]jsis and evaluation. The basic components comprising the Instrumentation System are: sensors, signal conditioners, Multiplexer-Encoder System, and transmitters. Because the system is used to sense Parameters of every spacecraft system, its components are located throughout the spacecraft.

SYSTEM OPERATION
The purpose of the Instrumentation System is data acquisition with respect to the progress and condition of the spacecraft, necessitating its operation throughout the mission. The Instrumentation System provides the capability of data acquisition and transmission to the ground stations. The data is supplied by all spacecraft systems. The basic operations by which the system fulfills its purpose are: to sense the various conditions and functions; convert them to proportional electrical signals (if applicable); condition the resulting signal (when necessary) to make it compatible with the encoding and multiplexing equipment; display pertinent data in the cabin; record data for delayed time (datadump) transmission; and provide signals for real-time transmission to the ground station. The system senses the prescribed parameters through the use of sensors which may be contained within the Instrumentation System or which may be an integral part of another system. Typical sensors include pressure transducers, accelerometers, and temperature sensors. Signals may also be obtained from such functions as switch and relay actuations, and from electronic package monitor points. The majority of the signals acquired are usable for the spacecraft cabin indicators and/or the encoding equipment without alteration. Some of them, however, are routed to signal conditioning packages (instrumentation assemblies) where their characteristics and/or amplitudes are changed. The resulting signals, as well as those from the other sensors, are of four basic types: low-level (0-20 my tic), high-level (0-5 vdc), bi-level (0 or 28 vdc), and hi-level pulse (28 or 0 vdc). Signals of selected parameters are supplied to the cabin indicators, while signals of all parameters are supplied to the Multiplexer/Encoder System. The Multiplexer/Eneoder System converts the various spacecraft analog and digital signals to a serial binary-coded digital signal for presentation to the data-dump tape recorder and the real-time telemetry transmitter. The tape recorder records a portion of the real-time data from the programmer at a tape speed of 1 7/8 inches per second and, upon command, will play back the data for transmission to a ground station, at a speed of 41.25 ips (22 times the recording speed). Four physiological functions are monitored for each pilot. All of the measurements are supplied as real-time data, while only one is supplied as delayed-time data. In addition, most of the measurements are recorded by two special (biomed) tape recorders. During pre-launch operations, data acquisition is accomplished by use of hardlines attached to the spacecraft umbilical and by telemetry. Between launch and orbital insertion, data acquisition is via the real-time telemetry transmitter. While the spacecraft is in orbit, data is acquired via the realtime telemetry transmitter for the period while the spacecraft is within range of a ground station. Data during the period while the spacecraft is out of range of a ground station is recorded On the PCM recorder and played back via the delayed-time telemetry transmitter while the spacecraft is within range of a ground station. A more detailed description of the telemetry transmitters is given in Section IX. The paragraphs to follow, present a brief description of all instrumentation parameters. The parameters are described in groups identified by their applicable data source system. It should be noted that although most of the parameters are applicable to all spacecraft, the following parameters is for spacecraft 8 specifically.

SEQUENTIAL SYSTEM PARAMETERS
The Instrumentation System monitors 41 sequential events and Sequential System parameters. Each parameter is described below individually, or as part of a group of related parameters. The Time Reference System (TRS) supplies three 24-bit digital words to the 24-bit shift register of the PCM progr_er. These three signals are: time since liftoff (AA01, AA02) and time to retrograde (AA03). Time since lift-off is referenced to the launch vehicle lift-off signal and provides time correlation for the data tape recorders. Time to retrograde (MOB) indicates the time remaining before retrofire initiation by the TRS. This signal is used to verify that the correct retrofire time has been inserted into the TRS by ground command or by the pilots. Launch vehicle second stage cut-off (ABOI) is monitored for ground station indication of this event. This parameter is provided by a signal from the spacecraft IGS computer to a bi-level channel of the programmer. Launch vehicle/spacecraft separation (AB03) is indicated to the ground station when any two of the three spacecraft/launch vehicle limit switches close, energizing the spacecraft separation relays. Actuation of any two of the three relays applies 28 vdc to a bi-level channel of the programmer. Rendezvous Radar prlmarypower (AGO6) is a high-level signal applied to the reentry high-level multiplexer. This signal originates when the Rendezvous Radar primary power switch is energized. Docking catch release (AC03) originates during the separation sequence after docking has occured. Actuation of the release catch mlcroswitch energizes the docking catch release relay in the instrumentation relay panel, and provides a bi-level signal to the programmer. Equipment section separation (AD02) is monitored to indicate a safe condition for retrograde prior to manual initiation or ground command of retrofire as a backup to the automatic system. This signal is originated when any two of the three separation sensors close, energizing the equipment section separation relays. Actuation of two of the three relays applies 28 vdc to a bi-level channel of the programmer. The retrorocket ignition commands are monitored by ground stations to obtain data for calculation of expected re-entry trajectory. Automatic (ADOB) and manual (ADO6) ignition commands are monitored. Parameters are obtained from the ignition command of the four retrorockets individually; ADO9, rocket 2; ADO8; rocket 3; ADIO, rocket 4. The manual and automatic retrofire commands indicate retrorocket I fire. The signals, 28 vdc, are applied to the re-entry hlgh-level multiplexer. Channel i0 of the Digital Command System is used by the ground station to relay the abort command to the spacecraft. Verification of ABORT light lllumination is by (AF06) parameter. Indication that the pilot actuated abort (AFO1) is supplied to the ground station. The signal is originated when the abort handle is moved to the ABORT position actuating a limit switch which energizes the instrumentation abort relays. Actuation of one of the relays applies a signal to a hi-level channel of the programmer. In case of pilot ejection during an abort, left (AFOB) and right (AFO2) ejection seat gone signals are relayed to the ground station. The signals are originated at the time the ejection seats leave the spacecraft closing the corresponding limit switch and applying the signals to the bi-level channels of the programmer. Confirmation of salvo retrofire is given to the ground station in case of an abort. A signal is applied to a hi-level channel of the re-entry high-level multiplexer when the salvo retrograde relay is energized. Indication of booster cut-off command(AB04) is given to the ground station when pilots move the ABORT handle to the SHUTDOWN position, actuating a limit switch. This energizes a relay applying 28 vdc to a bi-level channel of the programmer. Ground indication of pilot parachute deployment (AE02) is provided via a bilevel channel of the programmer. The signal is originated when a lanyard from the parachute actuates a toggle switch, energizing the pilot parachute deployed instrumentation relay. The parachute Jettisoned (AE13) signal is initiated when the pilot depresses the CHUTE JETT switch energizing redundant main parachute jettison relays. The relays apply a 28-vdc signal to a bi-level channel of the re-entry highlevel multiplexer. Platform mode selection (AG05) is indicated to a ground station. Any position other than OFF on the PLATFORM mode switch will apply a signal to a bi-level channel of the programmer. Primary (AGI6) or secondary (AG17) horizon scanner operation can be monitored by the ground station via bi-level channels of the programmer. Primary pitch (AG02), roll (AG03), and yaw (AG04) and secondary pitch (AG13), roll (AGI4), and yaw (AGIS), rate gyro operation is monitored to indicate an on or off condition. Each signal is applied to a signal conditioner whose output is applied to a hi-level channel of the high-level multiplexer. Pitch (AG1O), roll (AG11), and yaw (AG12) rate gyro (primary or secondary depending which is operational) outputs are applied to three signal conditioners. Each of the signal conditioners is a transistor switch providing no output for an input of 0-0.325 volts and a 16. 5 volt output for an input greater than 0.325 volts. The conditioned signals are applied to bi-level input channels of the programmer. Bio-medical tape recorder on-off signals (AGIS, AGI9) are used for time correlation of the recorded bio-medical data with the telemetry data. An on-off indication is provided to the playback recorder and to telemetry by a hi-level signal to the programmer (AGI9) and re-entry high-level multiplexer (AG18). Drogue parachute deployment (AE27) and drogue release (AE28) can be verified by the ground station via hi-level channels of the programmer. The signals are initiated when the HI-ALT DROGUE switch is depressed. The selected cryogenic quantity switch position is indicated to the ground station by AG21 (Reactant Supply System oxygen), AG22 (Reactant Supply System hydrogen), and AG2B (Environmental Control System oxygen) to allow the ground station to identify the reading of CA09 described under Environmental Control System.

ELECTRICAL POWER SYSTEM PARAMETERS
Approximately 24 Electrical Power System parameters are monitored by the Instrumentation System. The parameters are listed and described in the following subparagraphs. Fuel cell oxygen (BA02) and hydrogen (BA04) tank pressures are monitored by dual potentiometer pressure transducers installed as part of the fuel cell system. Each dual transducer provides one output to the adapter hlgh-level multiplexer and the other output drives an indicator on the instrument panel in the cabin. To evaluate proper operation of the fuel cell, stack 1A (BDO1), IB (BD02), 2A (HE01), 2B (HE02) and section 1 (BHOI) and 2 (BH02) currents are monitored and transmitted to the ground station. Stack C currents are obtained mathematically by subtracting section A and B currents from the corresponding section current. The signals being monitored originate from 50 millivolt shunts. The shunts are installed at the main buses for the section, and in the lines from stacks A and B to the main buses for stack A and B currents. Each of these signals is conditioned to a 0 to 20 millivolt signal which is directly proportional to the input current and then applied to the re-entry low-level multiplexer. The following parameters relate to the ground station information regarding spacecraft main, squib and control bus voltages: BGOI (main), BG02 (squib l), BG03 (squib 2), BG04 (control bus). Each of these parameters is conditioned and then applied to the re-entry high-level multiplexer. The Reactant Supply System (RSS) O2 (BA05) and H2 (BA06) supply bottle temperatures are monitored by means of two temperature sensors located on each supply bottle. The output of the sensors is applied to the adapter low-level multiplexer. Fuel cell section 1 O2 to H2 (BC05), section 1 O2 to H2O (BBO7), section 2 O2 to H2 (BC06), and section 2 O2 to H2O (BBO8) differential pressures are monitored by a pressure-sensitlve switch installed within the fuel cell to provide for safe operation monitoring capability of the fuel cell by the ground station. The outputs of the pressure switch is applied to hi-level channels of the adapter high-level multiplexer. Oxygen (BB05) and hydrogen (BC03) temperatures at the outlet of the heat exchanger are monitored and relayed to the ground station via the adapter low-level multiplexer. To provide an aid in evaluating fuel cell operation by the ground station, section 1 O2 (BD04), section 2 O2 (BE04), section 1 H2 (BD06), and section 2 H2 (BE06) purging is monitored. The signals are actuated by the pilot by placing the corresponding section purge switch to the H2 or O2 position. The signals are applied to the hi-level channels of the programmer.

ENVIROMENTAL CONTROL SYSTEM PARAMETERS
Twenty-eight parameters and RSS/ECS qu-ntities associated with the ECS are monitored by the Instrumentation System and relayed to the ground station for analysis. The primary oxygen tank pressure (CA02) is telemetered to the ground station and displayed in the spacecraft cabin. The signals originate from a dual potentiometer pressure transducer installed as part of the ECS. The signal is relayed to the ground station via the adapter high-level multiplexer. A differential pressure transducer is used to sense cabin to forward compartment pressure differential (CB01). The transducer has a dual output used for cabin indications and for transmission to the ground station via the re-entry high-level multiplexer. Left (CC01) and right (CC02) suit to cabin differential pressure is displayed in the spacecraft cabin and telemetered to the ground station. Dual potentiometer pressure transducers serve as the signal source. The output of each transducer is applied to the cabin indicator and to the re-entry high-level multiplexer.