The control system included the attitude control and maneuver electronics and the horizon sensors. It was used in conjunction with the propulsion system and associated guidance systems to provide spacecraft orientation about its three major axes and for translational maneuvering. The orbit attitude and maneuver thrusters were employed to assist in spacecraft/launch vehicle separation and for attitude control and maneuvering prior to adapter section jettison. During retrograde and reentry, control thrusting was provided by the reentry control system thrusters.

Attitude control and maneuver electronics (ACME) included the attitude control electronics (ACE), orbit attitude and maneuver electronics (OAME), power inverter and two rate gyro packages. Input signals to the attitude control and maneuver electronics were supplied by the IGS, the horizon sensors, or by crew manipulation of the attitude control handle or the maneuver controller, depending on the operational mode the astronauts had selected.

The attitude control electronics accepted input signals from the inertial guidance system, from the attitude control handle, from the rate gyros, and from horizon sensors. Input signals were converted into drive commands for the reentry control systems solenoids and to logic commands for the orbit attitude and maneuvering electronics subsystem. Honeywell, Inc. was subcontractor for ACME, ACE, and OAME.

The orbit attitude and maneuver electronics accepted input signals from the attitude control electronics and the maneuver controller for conversion to drive commands for the OAME solenoids.

Rate gyros sensed angular rates about the pitch, roll and yaw axes of the spacecraft. Rate signals were supplied to the ACE in the rate command mode plus all other modes where rate damping was employed. The rate gyro packages also provided inputs to attitude displays and to the telemetry system.

Attitude control and maneuver electronics modes provided for maneuver control to effect spacecraft/launch vehicle separation, for translational maneuvering during the catch-up mode operation, and for maneuvering in all planes by attitude maneuvers and forward/aft thruster operations. Forward or aft displacement of the maneuver controller from the neutral position produced a direct command to the respective solenoid valve driver. The astronauts could select any one of six control modes for spacecraft orientation during orbit and reentry. These modes were rate command, direct, pulse, reentry rate command, horizon scan and reentry.

Rate command mode: Spacecraft angular rates were proportional to rate command signals initiated by flight crew displacement of the attitude control handle. The rate command signals were compared in the attitude control electronics with rate gyro outputs and when the difference between the two signals exceeds the damping dead zone in the system, proper reaction control jets fire for attitude control. This mode was utilized during manually initiated attitude control operation and during retrofire and other velocity change maneuvers. In this mode it was possible to maintain attitude within +/- 1 degree using the flight director indicator for reference.

In the direct mode, switched on the attitude hand controller directly control the ACME solenoid drivers. Control switches were actuated when controller displacement was greater than approximately 1/4 of its full travel.

The pulse mode permitted the astronauts to make fine attitude corrections of the spacecraft about its three major axes. In this mode, angular rates were incrementally changed by single thrust commands of fixed duration. Each displacement of the hand controller by the astronaut triggers a single pulse generator in the attitude control electronics and sends a single pulse command to the proper reentry control system or orbit attitude and maneuvering electronics solenoid valve drivers. The astronaut must return the hand controller to the neutral position before he could initiate another pulse to increase his angular rate or effect a braking action in the other direction.

Reentry rate command mode was utilized for manual attitude control during reentry. This mode provided similar operating characteristics to that of the rate command mode, except that damping dead bands were wider and roll rate crossfeed was included in the yaw damping loop, providing for conservation of fuel because the fine control provided by rate command was not necessary for manual performance of this phase of the mission.

The horizon scan mode provided for automatic control of the spacecraft about the pitch and roll axes during the orbit phase of the mission. ACME received attitude information from the horizon sensor and generated an output to the proper thrusters to maintain the attitude within the damping dead band. When in this mode, the ACME supplied a nose-down pitch bias which enabled the flight crew to view the horizon out of the window.

In the reentry mode spacecraft pitch and yaw angular rates were automatically- maintained within a damping dead band by the ACME. A roll attitude was determined by inputs from the computer to the ACME. The computer either provided a bank angle command or a fixed roll rate depending on the relationship between the predicted and the desired touchdown points. ACME would not accept rate commands from the attitude control handle when in the reentry mode.

Two horizon sensors, one primary and one secondary or standby unit, provided reference signals for alignment of the inertial platform and error signals to the ACME for controlling the spacecraft attitude about its pitch and roll axes. Horizon sensors operated by tracking the Earth's infrared horizon.

Two attitude displays, each incorporating a three-axis attitude reference ball with 360 degrees of rotation about each axis, were provided on the right and left instrument panels. These displays, built by the Lear-Siegler Corporation, were slaved to the positions of the normal inertial platform gimbals and provided a continuous all-attitude reference of roll, pitch, and yaw.

Integral rate and command flight director needles displayed control movements required to position the spacecraft in a commanded attitude or rate. When the commanded attitude or rate was achieved, the needles were centered.

An attitude hand controller was mounted on the console between the astronauts and provided a means of manually controlling the spacecraft attitude and rate in three axes. The controller could be operated by either astronaut, while either was in the restrained position, through simple wrist articulation and palm pivot motion. The controller was spring loaded to provide an increasing resistance as the handle was moved away from neutral. The total travel of the hand controller was +/-10 degrees from neutral in all three axes. Displacement or rotation of the controller caused the spacecraft to turn in the direction of displacement or rotation.

A maneuver hand controller initiated translation of the spacecraft. The controller contained centering springs and six switches, one for initiation of spacecraft displacements in each of six directions along with three major axes. Movement of the handle in any of these six directions initiated corresponding spacecraft translation. The handle could be removed and stored when not in use, providing clearance in the event of seat ejection.

Guidance System Functions

Mission Phase	Applicable Guidance Mode	Function
Launch	Ascent	Provided backup guidance in event of a radio guidance system (RGS) malfunction. Provided attitude error data for astronaut evaluation of mission status. Provided navigation initial conditions for reentry guidance in event of an abort during boost.
Post-SSECO	Ascent	Provided capability to correct apogee and perigee altitude deviations resulting from insertion velocity, flight path angle, and altitude errors.
Orbit	Catch-up	Provided the capability for displaying and applying ground determined velocity changes.
Orbit	Prelaunch	Provided a diagnostic sum check of computer syllable two instructions. Provided standby guidance mode wherein digital acquisition was provided and MDIU/DCS locations were telemetered and provided different telemetry quantities.
Retro	Reentry	Provided a real-time navigation during retrofire sequence. Display of retrograde incremental velocity to astronauts.
Reentry	Reentry	Performed real-time navigation. Performed continuous prediction of the spacecraft zero lift impact point. Provided control logic to deliver steering commands to control crossrange and downrange travel. Provided an automatically controlled reentry in conjunction with reentry attitude control mode.

Attitude Control System Characteristics

Mission Phase	Applicable Attitude Control Mode	Function
Launch	Direct	Rate information to flight crew.
Post-SSECO	Direct	ON-OFF commands to attitude thrusters.
	Rate Command	Vehicle angular rates were proportional to attitude control handle displacements.
Orbit	Horizon Scan	Automatic pitch and roll attitude control to a horizon sensor reference.
		Pulse mode capabilities retained for manual over-ride about all axes.
		Five-degree small-end down pitch bias included to maintain astronaut view of Earth's horizon.
	Pulse	Angular rate could be changed in incremental steps by commanding thrust for a calibrated period of time.
		Effective about all three axes and could be used for fine attitude control.
	Rate Command	Same as described for Post-SSECO.
	Direct	Same as described for Post SSECO.
Retro	Rate Command	Same as described for Post SSECO.
	Direct	Same as described for Post-SSECO.
Reentry	Automatic rate damping about pitch and yaw axes.
		Roll attitude and rate controlled by computer and ACME.
		Roll rate cross-feed was included in yaw damping loop.
Reentry	Reentry-Rate Command	Astronaut control of roll orientation.
		Rate damping provided about all axes.
		Roll rate cross-feed was included in yaw damping loop.
	Direct	Astronaut control of roll orientation.
		Manual rate damping required in all axes.

ACME Inputs and Outputs

Inputs

• Pitch and roll attitude error signals from the horizon sensors.
• Roll rate or roll attitude signals from the digital computer.
• Pitch, yaw, and roll on-off attitude acceleration commands from the attitude control handle.
• Pitch, yaw, and roll proportional rate commands from the attitude control handle.
• Pulse initiation signal from attitude control handle.
• Translation commands from the maneuver control handle.
• Mode of operation change commands as selected by the crew.
• Test inputs as necessary for satisfactory checkout during ground testing of the ACME in conjunction with associated equipment.
• Redundancy selection commands initiated by the crew.

Outputs

• Attitude maneuver control "on-off" command signals to the affected OAMS and RCS thrust chamber solenoids.
• Pitch, roll, and yaw rate signals.
• Signals to telemetry systems as required for monitoring ACME operation.

Time Reference System

The time reference system consisted of an electronic timer. a time correlation buffer, an event timer, and a clock.

An interface existed between the time reference system and the digital command system, the digital computer, and the data transmission system.

The electronic timer recorded elapsed time in 1/8th-second increments from liftoff through impact; it counted time-to-retrograde from liftoff to zero in 1/8th-second increments; and it counts time to equipment reset on command in 1/8th second increments. The electronic timer exchanged signals with the digital command system, the digital computer, and the data transmission system.

The electronic timer had a crystal controlled time reference accurate to 35 parts in 1 million for 24-hour period. Stability over a 3-hour period was 10 parts in 1 million at 25 degrees C 10 degrees C. The timer was mounted behind the center instrument panel.

Updated or revised time-to-go was forwarded to he electronic timer by the digital command system.

To prevent inadvertent or premature countdown to retrofire, the electronic timer was provided with a lockout set at 512 seconds. It would not accept any time to go quantity of less than 512 seconds.

The time correlation buffer accepted elapsed time and clock information from the time reference system pulse code modulation input. The time correlation buffer provided outputs to the voice tape recorder and the two biomedical tape recorders. Information to the recorders from the time correlation buffer was updated every 2.4 seconds. It provided serial data and clock data outputs to a buffer register every 2.4 seconds.

The event timer provided a cockpit decimal time displayed in minutes and seconds to a maximum 99 minutes and 59 seconds. This display permitted countup and countdown timing by the astronauts. The display could be manually positioned or it could be started by face-mounted switching or independent electrical remote signal. The unit operated completely independent of the electronic timer.

The spacecraft clock displayed Greenwich Mean Time (GMT) in hours and minutes. Launches were conducted at Cape Kennedy on GMT. The clock included an additional minute hand and a second hand, which could be stopped and reset to zero mechanically at any time. A calendar day display was also provided.

Gemini instrumentation and recording systems monitored specific spacecraft systems, conditions, and events, and sensed, conditioned, encoded, reproduced, and transmitted data to ground stations.

The instrumentation and recording system performed the following:

• Accepted physiological signals and signals from the sequential, environmental, electrical, adapter module propulsion, reentry module propulsion, communication, and guidance and control systems through sensing devices that were included as components of these various systems.
• Sensed structural signals, temperature, accelerations, aerodynamic pressures, gas partial pressures, and cabin pressures and temperatures.
• Sensed or accepted signals from experimental devices.
• Provided signal conditioning equipment to transmit signals for the crew station display from the instrumentation system sensors. It also accepted signals from sensing devices that were part of the crew station instrumentation display system.
• Included signal conditioning equipment as necessary to prepare signals for transmission to instrument panel displays.
• Included signal conditioning to adapt signals to the input requirements of multiplexer-encoding equipment.
• Encoded data for telemetry and recording.
• Recorded and reproduced data for delayed telemetry transmission.
• Recorded general purpose photographic data.
• Recorded astronaut conversations.

The instrumentation system had both an operational and a diagnostic function. The system provided real-time telemetry transmission of data generated in the spacecraft and required at ground stations for monitoring the progress of the spacecraft mission, for assessing the spacecraft status, and for making decisions concerning flight safety. In addition, Gemini instrumentation included the necessary hardware for generating signals, not available through individual spacecraft systems, for operation of the spacecraft cabin displays. In its diagnostic function, the spacecraft instrumentation system provided a means for documenting significant events and data, throughout the entire mission, by three methods: Real-time transmission, delayed transmission, and onboard recording. Electro-Mechanical Research supplied the data acquisition system for Project Gemini.

Data Transmlssion System Programmer

The data transmission system programmer provided for data multiplexing, analog-to-digital data conversion, and digital-data multiplexing.

The programmer included a high-level multiplexer, a low-level multiplexer and a delayed transmission recorder/reproducer.

The high-level multiplexer functioned as a high-level analog commutator and on-off digital data multiplexer. It provided for the sampling of 32 high-level data channels, 24 bi-level signals and 16 inverted bi-level pulse signals.

The low-level multiplexer functioned as a differential-input analog input commutator and provided a sequential sampling of 32 low-level signals. The delayed transmission recorder/reproducer recorded data during the time the spacecraft was in orbit and out of range of worldwide tracking stations. When the spacecraft was within range of a tracking station, the recorder/reproducer was triggered by either a ground signal or by the astronauts, reversed its tape direction, and plays back the recorded data at an accelerated rate.

Photographic recording was accomplished by two cameras, one a 70-millimeter camera for still photography and the second a 16-millimeter movie camera. These cameras were hand held and operated by the astronauts and were stored in the cabin when not in use.

Voice recording was accomplished on a tape recorder installed in the center pedestal. Recording was limited to astronaut conversations and would occur at any time that the mode selection switched on the voice control center was placed in a record position.

Biomedical recorders were installed in the spacecraft and accept and recorded data from the biomedical instrumentation centers attached to the astronauts.

Propulsion Systems

Secondary propulsion systems (as contrasted to the primary propulsion systems of the launch vehicle) were installed in the reentry module and the adapter module to provide capability for separation from the launch vehicle either under normal conditions or in emergency conditions; for translational maneuvering in six basic directions: up, down, left, right, forward and rearward and for attitude control about the pitch, roll and yaw axes. One system, the retrograde rocket system, provided the necessary velocity decrement to initiate reentry.

Propulsive thrust to perform these functions was generated by three individual systems: the orbit attitude and maneuver system (OAMS), the reentry control systems (RCS), and the retrograde rocket system (RRS). The orbit attitude and maneuver system was installed in the equipment and retrograde sections of the adapter module. The reentry control systems were located in the reentry module. The retrograde rocket system was clustered in the center of the retrograde section, just aft of the reentry module heat shield.

Orbit Attitude And Maneuver System (OAMS)

The OAMS was a liquid bipropellant rocket engine propulsion system constructed on a modular basis consisting of thrust chamber assemblies; pressure storage tanks; pressure regulators; propellant tanks; propellant shut-off valves; propellant and pressurant lines; propellant line cutter-sealer assemblies; and five component packages which provided means of ground testing, pressurant and propellant tank filling, burst diaphragms, relief valves and instrumentation. The components were divided according to function into a thrust chamber assembly group, an oxidizer/fuel group and a pressurizing group.

The OAMS thrust chamber assembly group consisted of 16 engines, each mounted in a fixed position and operated at a fixed thrust level. (The amount of force desired from each engine for maneuvering was obtained by varying the time it was operated at the fixed thrust level.) Eight of the engines develop about 25 pounds; two produce about 85 pounds and six develop approximately 100 pounds of thrust.

The engines provided attitude and maneuver control from spacecraft separation from the launch vehicle until the adapter equipment section was jettisoned. The OAMS engines were operated by signals from the orbit attitude and maneuver electronics system. Pitch, roll and yaw torques were obtained by firing the 25 pounds thrust engines in pairs. Maneuvering was accomplished by firing the 100 pounds thrust engines for lateral, vertical and forward movement. Two of these fired to the aft to provide thrust for spacecraft separation from the launch vehicle. The two 85 pounds thrust engines fired forward to provide rearward motion.

Each engine assembly consisted of two propellant valves with calibrated orifices and filters, a fuel and oxidizer injection system, a combustion chamber, and an expansion nozzle. The propellant valves were quick acting and were normally closed and operated by solenoid action; that is, the valves open upon application of an appropriate electrical signal to permit the flow of fuel and/or oxidizer to the injector to which they were fitted. In the event of electrical malfunction of any kind, no signal was transmitted and therefore the solenoid valves remain closed. Valve construction was such that the parts in contact with the fuel and oxidizer were not adversely affected by the corrosive nature of the propellants.

The combustion chamber and the expansion nozzle were lined with an ablative material, which maintained internal geometry and protected the external wall from temperature damage.

The engines operated on storable hypergolic propellants. The oxidizer was nitrogen tetroxide; the fuel was monomethylhydrazine. Fuel quantities varied according to mission requirements.

The oxidizer and fuel propellant tanks were all-welded titanium spherical structures. Propellant tank volume and arrangement varied in accordance with mission requirements. Two propellant shut-off valves permitted isolation of the propellants from the engine chamber assemblies in the event of engine malfunction. One valve was in the oxidizer feed system and the other valve was in the fuel feed system. The oxidizer/fuel group valves were electric motor operated.

Since the OAMS functioned under a weightless condition during much of its operational lifetime, it was necessary to use a pressurization system to expel propellants on demand.

To achieve positive expulsion of propellants, both the oxidizer and the fuel were in bladder containers inside storage tanks. Gas was induced between the tank wall and the bladder to provide a "squeezing" pressure that forced propellant to the engines.

The pressurant, which forced the oxidizer and fuel to the engine, was helium. The pressurization system included a storage tank, component packages and pressure regulation group.

The pressurant tanks also were all-welded titanium spheres each with a volume of 1700 cubic inches. The pressure in the propellant tanks was regulated at 295 pounds per square inch. The number of tanks installed varied with mission requirements.

Reentry Control Systems (RCS)

The reentry control systems were completely independent of the orbit attitude and maneuver system and were installed in the reentry module just forward of the pressurized cabin. There were two completely independent reentry control systems. Each was a liquid bipropellant rocket engine propulsion system constructed on a modular basis. Both were identical and provided redundancy in the event of a malfunction of one system.

There were eight fixed-thrust-level, fixed mounted engines in each of the two systems. As in the OAMS, they operated on storable hyperbolic propellants supplied by a cold gas, pressurized, positive expulsion feed system. Electrically operated valves controlled the oxidizer and fuel flow in the thrust chamber assembly. The basic operation of each RCS was identical to that of the orbit attitude and maneuver system. The reentry control systems responded to electrical signals from the orbit attitude and maneuver electronics system, the same unit which operated to provide input to the orbit attitude and maneuver engines.

In the event of a malfunction of one system, the remaining RCS had sufficient total impulse capacity and thrust to assure attitude control during retrograde, stabilization for a safe reentry, and thrust as necessary to steer the reentering spacecraft to the desired landing point.

Attitude was controlled through the attitude control and maneuvering electronics from inputs of the astronaut's hand controller. Pitch, roll and yaw torques were obtained by selective firing of pairs of engines. Each had a nominal rocket engine thrust output rated at about 25 pounds.

The fluids involved in the operation of the RCS were identical to those in the OAMS except for the pressurant, which was nitrogen in the case of the RCS. The pressurizing group and the fuel packages were contained in the non-pressurized section of the reentry module. The pressurant tank had a fluid volume of 185 cubic inches. The pressure regulator maintained a 295 pounds per square inch to the propellant tanks.

The propellant tanks, one oxidizer and one fuel, were all-welded titanium cylindrical structures. The oxidizer tank had a fluid volume capacity of 439 cubic inches; the fuel tank had a fluid volume capacity of 546 inches. The engines of the reentry control systems had characteristics similar to those of the orbit attitude and maneuvering system 25 pounds thrust engines.

Retrograde Rocket System (RRS)

The Thiokol RRS, an independent system, consisted of four solid propellant rocket motors mounted in the retrograde section of the adapter and located symmetrically about the longitudinal axis of the spacecraft.

The retrograde rockets, each with approximately 2500 pounds of thrust, provided an impulse to the reentry module resulting in a sufficient velocity decrement to initiate reentry into the Earth's atmosphere. The retrograde rockets were fired automatically by an electrical signal from an onboard electronic timer. There was also a manual backup initiation capability. The rockets fired sequentially at a nominal 5.5-second interval.

The spent retrograde rockets were jettisoned with the retrograde section of the adapter approximately 45 seconds after rocket firing initiation. In the event of an abort before orbital altitude and velocities were achieved, the retrograde rockets could be salvo fired by the flight crew to aid in separation of the spacecraft from the launch vehicle.

Each rocket consisted of a motor case, a partially submerged contour nozzle, and dual pyrogenic igniters with removable pressure cartridge initiators. The motor case was an all-welded titanium alloy sphere slightly greater than one foot in diameter. The nozzles include an expansion cone a throat insert and a nozzle bulkhead. The pyrogenic igniter was a small rocket of short burn duration, which fired into the charge of the retrorocket thus igniting it.

Landing System

A parachute landing system provided for final descent of the spacecraft to the Earth's surface. Deployment of the parachutes reduced the reentry module trim angle and decelerated the reentry module to a rate well within acceptable limits for water impact.

The parachute landing system included a high altitude drogue parachute, cable guillotines, a pilot parachute, a main parachute, a bridle assembly, attachment and disconnect assemblies, mortar assemblies, reefing cutters, displays and controls.

Landing system operation began at approximately 50,000 feet, with the astronaut deployment of the high altitude drogue parachute. It was deployed by firing mortar cartridges, which ejected the drogue parachute from its container.

When the spacecraft reached approximately 10,600 feet the pilot parachute was deployed by the flight crew. Approximately 2 1/2 seconds later the rendezvous and recovery section automatically separated from the reentry module by action of a mild detonating fuse. The pilot parachute then pulled the R and R section free of the reentry module, and as it was pulled free, it drew the main parachute out of its container. The main parachute deployed in a reefed condition.

After landing, the parachute was jettisoned by the astronauts. Transfer from the single point suspension to a two-point suspension was effected by the astronaut depressing the landing attitude switch.

High Altitude Drogue Parachute

The high altitude drogue parachute was an 8.3-foot diameter conical ribbon chute of nylon material. It was packaged in a deployment bag stored in the drogue parachute mortar tube at the top of the R and R section. Approximately 16 seconds after the high altitude drogue parachute was deployed two redundant, lanyard-actuated reefing cutters unreefed the drogue parachute. The riser from the drogue parachute was attached to the R and R section by three steel cables. The unreeled drogue parachute stabilized the reentry module within 23 degrees of the vertical axis as it descended from 50,000 feet to 10,700 feet.

Cartridge-actuated guillotines severed the high altitude drogue parachute attach cables. Actuation of these guillotines was switch-controlled by the crew. With the attach cables severed, the high altitude drogue parachute pulled the pilot parachute from its mortar, resulting in pilot parachute deployment.

Pilot Parachute

The pilot parachute was an 18.3-foot diameter ring-sail parachute constructed of nylon material. The parachute, its reefing line, reefing cutters and risers were packaged in a deployment bag stored in the pilot parachute mortar tube in the rendezvous and recovery section. The pilot parachute was deployed in a reefed condition and within 2.5 seconds 0.125 seconds after deployment, the pilot parachute drew the R and R section (released by action of a mild detonating fuse) away from the reentry module. As it did so, it deployed the main parachute in a reefed condition.

Within approximately 6 seconds after pilot parachute deployment, two redundant lanyard-actuating reefing cutters severed the pilot parachute reefing line. The unreefed pilot parachute lowered the R and R section to the sea.

An auxiliary landing sequence was available in event of malfunction of the high altitude parachute. Use of this switch-controlled sequence enabled the astronauts to actuate drogue parachute attachment guillotines and the apex cable guillotine, to fire the pilot parachute mortar and to initiate R and R section separation, which resulted in main pilot parachute deployment.

Main Parachute

The main parachute was an 84.2-foot diameter ring-sail parachute constructed of nylon material with alternating white and international orange gores. The main parachute was attached to the reentry module by a riser and a bridle arrangement. The parachute reefing lines, reefing cutters and risers were packaged in a nylon web cotton sateen bag and stored in the main parachute, a Fiberglas cylinder inside the R and R section. The main parachute was deployed in a reefed condition at approximately 9,700 feet. Within 10 seconds 2.2 seconds after main parachute deployment, redundant lanyard-actuating reefing cutters severed the reefing line. Operation of any one of three reefing cutters resulted in the unreeling of the main parachute. The unreefed main parachute provided a rate of descent that resulted in allowable loads upon water impact.

Only one main parachute was provided in the Gemini landing system. Crew escape was possible through use of the ejection system in event of spacecraft main parachute malfunction.

The spacecraft was suspended from the parachute by means of a "bridle" assembly. The bridle included a forward and an aft leg. The forward leg was stored in a Fiberglas tray in the reentry control system section of the reentry module. It incorporated a loop at one end for connection to a disconnect assembly. The aft bridle leg was stored in a trough in the reentry module skin and incorporated a loop at one end for connection to an aft disconnect assembly. The trough, located between the hatches, was covered with a frangible insulating material, which allowed the aft bridle leg to tear free when two-point suspension was initiated. Following main parachute deployment, the initial mode of descent was by single-point suspension from a single riser. Transfer from the single point suspension to a two-point suspension occurred when an astronaut depressed the landing attitude switch. After landing, the parachute was jettisoned by the astronauts.

After the water landing, the astronaut depressed the parachute jettison switch, which released both the forward and aft bridle leg attachment/disconnect assemblies. This also energized a cartridge-actuated guillotine, which severed door restraints permitting a spring loaded hoist loop and attached flashing recovery light to extend. The Gemini reentry module, unlike the Mercury spacecraft, had no attached landing bag to absorb landing shock. The Gemini spacecraft was brought to a "pilot heads-up" position causing it to impact the water at the edge of the heat shield. Water landing forces resulting were less than the touch-down forces of Mercury, even though Mercury had a landing bag installed.

Postlanding and Survival Equipment

Postlanding and survival equipment provided recovery personnel with visual and radio reference aid in locating the reentry module after landing. It provided the crew with water, food and survival equipment and included a mooring lanyard assembly and hoist loop for recovery of the spacecraft after sea landing.

The equipment included a flashing recovery light, dye marker, survival equipment, UHF and HF rescue communications and beacons, splash curtains, a hoist loop, flotation material, and electrical power supplies.

The recovery light was a high intensity flasher deployed upon parachute jettisoning and located on top of the reentry module as it floats in the water. The light had a minimum flashing rate of 15 flashes per minute and was visible on a clear night at a distance of 50 nautical miles.

Fluorescent dye marker was installed in the forward end of the reentry control system section just below the flotation line. The dye was green-yellow and stored in a container having openings covered with a water-soluble film for automatic deployment upon exposure to the water.

The radio recovery aids included a UHF recovery beacon, UHF voice transceiver, HF voice transceiver and the rescue beacon in the survival pack.

The mooring lanyard assembly consisted of a lanyard with a fitting at one end for attachment to the "D" ring on the astronaut's parachute harness. A life raft was attached to the lanyard at a point about 8 to 12 feet from the astronaut.

During an ejection sequence the lanyard assembly extracted the life raft and a survival pack from containers in the seat as the backboard of the seat falls away from the astronaut. A cylinder charged with carbon dioxide was packed with the life raft for inflation. The life raft could be manually extracted following a normal landing in the spacecraft.

Ejection System

The Gemini escape system for safe escape of the astronauts through use of the ejection seats while the launch vehicle was still on the launch pad during boost and after spacecraft reentry. While ejection capability was available throughout the mission, actual usage was determined by the altitude, the type of emergency, the system condition, the mission phase, and the astronaut evaluation of the problem. Subcontractor for the ejection system was Weber Aircraft.

Two ejection seats were provided in the Gemini spacecraft for emergency escape. These seats included the seat structural assembly, a backboard assembly and an egress kit assembly. The seats were arranged side-by-side facing the small end of the crew compartment. They were constructed to be compatible with crewmembers wearing fully inflated pressure suits. The escape system incorporated a survival kit containing water, food, life raft, fishing gear, a radio transmitter and a machete, all packed into the seat.

The ejection seats functioned as one complete system. Should the need to abort arise, the decision to eject was made by the astronauts themselves. Once it was determined to eject, either man could pull the escape ring located between his knees, ejecting both astronauts. Only when manually initiated by an astronaut would the ejection sequence occur.

When the system was actuated, the remainder of the operation was fully automatic. First, both hatches of the Gemini spacecraft were opened simultaneously, then the rocket powered escape seats were propelled out of the vehicle.

Then, 1.1 seconds after ejection, the seats and men separated; 2.3 seconds later, a drogue gun fired, extracting a pilot chute from the astronaut's back pack. The deployed pilot chute then pulled a 28-foot diameter main parachute out, allowing for full canopy inflation.

In a pad abort, just 10 seconds after leaving the Gemini spacecraft, both astronauts descended to a safe landing in a nearby cleared landing area. After reentry the astronauts would have had the option of riding the spacecraft to water impact beneath a Northrop-Ventura ringsail parachute or ejecting themselves and landing much like a paratrooper.

In most instances the latter method would not be used unless the spacecraft entered the Earth's atmosphere at a point where a dry landing had to be performed.

The Gemini rocket catapult, furnished by Rocket Power, Inc., had a total impulse of 2650 pound-seconds. Burning time for the rocket was approximately 1/4 second. Rocket ignition occurred 0.2 seconds after catapult ignition. The astronauts would be subjected to maximum ejection acceleration of 24 g.

Trajectories that could be expected for "off-the-pad'' ejections would land the astronauts at least 500 feet from the launch vehicle. Tests indicate that the landing point would actually be closer to 800 feet from the launch site. The height of trajectories on 'off-the-pad' tests was approximately 350 feet above the terrain when launched from 150-foot height.

Escape systems testing was begun at the US Naval Ordnance Test Center, China Lake, California, in 1962, using a 150-foot-high tower. The 150-foot tower was used to simulate escape from the spacecraft while still on the pad. During the field testing operation, environmental studies were conducted on the system's related components, including harnesses, back pack, straps and pyrotechnics to determine component reaction under extreme heat, cold, humidity, shock, and vacuum.

High speed track tests were conducted with a full-scale Gemini boilerplate spacecraft mounted on a rocket powered sled. During sled tests, every possible escape condition was simulated by firing each seat at different attitudes to determine how the system would perform under adverse conditions. The sled was accelerated to 550 mph and the escape system actuated, causing the seats to be ejected out and away from the vehicle to qualify the system for high speed ejections. High altitude tests were conducted at the Naval Parachute Facility, El Centro, California, utilizing an F-106 supersonic fighter flying at Mach 1.75 and 20,000 feet. More than 100 studies and tests were conducted in the laboratory and field before the system was man-rated as operational Gemini equipment.

Ballute

If the returning astronauts used their ejection seats at high altitude, the ejection sequence would function as in a pad abort situation with one exception. At altitudes above 7,500 feet a device called a ballute were employed. Designed and built by Goodyear, the ballute would stabilize and decelerate the ejected astronauts during the free fall before parachute deployment.

The ballute, a contraction for balloon-parachute, was a stabilizing device included as part of the ejection system. The ballute was packaged in the ejection seat and was utilized during high altitude aborts to achieve desired stabilization of the astronaut.

The ballute was a balloon-shaped device looking much like a child's spinning top. It was constructed of inflatable rubberized fabric and was packaged in a deflated condition in the ejection seat during flight. The inflated balloon was approximately 48 inches in diameter and 54 inches long. The astronauts could use the ballute at heights up to 14 miles. When an altitude of approximately 5700 feet was reached, the personal parachute would open automatically and bring the astronaut safely down to Earth.

Pyrotechnic Devices

Pyrotechnic (explosive) devices were used throughout the Gemini mission for switching operations and major configuration changes and in recovery and escape systems. All pyrotechnic systems were designed with inherent redundancies. Pyrotechnic devices were initiated by either pressure, electrically, or by means of a lanyard.

The pyrotechnic devices were installed in such a manner that explosive effects were confined within the device or directed outward and shielded in such a manner as to preclude damage to nearby equipment or inward release of shrapnel