The Gemini spacecraft was a conical structure nearly 5.8 m high, 3.05 m in diameter at its base and weighing over 3175 kg.

The spacecraft was designed to endure the aerodynamic pressures, temperature loading, vibration and acoustical noise of launch; the temperature and vacuum of orbital flight; and the extreme heat of reentry, and the impact forces of water landing while providing life support for two astronauts and the necessary equipment for planned missions and experiments.

Gemini's design reflected the knowledge obtained from the development, manufacture and flight operation of the Mercury spacecraft, which, like Gemini, was produced for NASA by McDonnell Aircraft Corporation, St. Louis, Missouri.

Gemini was launched by an Air Force Titan II launch vehicle built by the Martin Company.

The spacecraft consisted of two major parts, a reentry module and an adapter module. The reentry module was designed to withstand the extreme heat of reentry. Its side walls were protected by heat resistant shingles and the large bulkhead by an ablative heat shield. The sections of the adapter module remained in decaying orbits and were burned up during reentry.

The spacecraft was primarily "skin-stringer" construction. Ring stabilized stringers carry near all axial loads. Structural materials and construction methods exhibited the influence of Gemini engineers' search for optimum strength-to-weight ratios. Titanium and magnesium were the principal metals used.

Reentry Module

The reentry module (the dark portion of the spacecraft) had three primary sections: rendezvous and recovery (R and R), reentry control system (RCS), and a cabin section. Integral to the reentry module was a heat shield attached to the large end of the module. An aerodynamic cover protected horizon sensors at the midpoint of the reentry module.

Materials used in the conical section of the reentry module varied greatly because of the effects of reentry heating. Super alloys such as Rene 41 and L-605 were used for the outer skin and skin attachments, which were thermally isolated from the inner structure by Johns-Manville MIN-K, Fiberglas and Thermoflex RF insulation. The basic load- carrying shell was titanium. Aluminum was used inside the cabin where heat was not a structural problem.

Rendezvous and Recovery (R and R) Section

The R and R section provided sufficient volume for a rendezvous radar system and a parachute landing system. Structural rings, stringers, and bulkheads were made of titanium. The external surface was covered with beryllium shingles. The nose fairing was reinforced plastic and Fiberglas-laminate. The R and R section was attached to the RCS section with 24 frangible bolts, which were fractured to jettison the R and R section upon deployment of the pilot chute following reentry.

Under the beryllium shingles were Thermoflex RF blankets held in place by a titanium mesh attached to the stringers. The outer surfaces of the rings and stringers were insulated with 0.038 mm Inconel-foil-encased Min-K held in Fiberglas channels.

Reentry Control System (RCS) Section

The RCS section housed reentry control system fuel and oxidizer tanks, and thrust chamber assemblies. This section was located between the rendezvous and recovery section and the cabin section. The cylindrical RCS section was an inner skinned ring-stringer titanium structure with an outer skin of beryllium shingles. External heat protection was essentially the same as for the rendezvous and recovery section.

Cabin Section

The cabin of the Gemini spacecraft was a truncated cone (a conical enclosure with approximately the top one-third cut away), which housed the Gemini crew, electrical, and life support equipment, and various experimental devices. The pressure vessel (crew compartment) provided adequate space for the two-man crew plus instrumentation and life support equipment.

The pressure vessel had a fusion-welded titanium frame attached to side panels and fore and aft bulkheads. The side panels and pressure bulkheads were double thickness, thin-sheet titanium (0.25 mm) with the outer sheet beaded for stiffness. A hatch was provided over each astronaut for entering and leaving the spacecraft.

Equipment bays, which contained a variety of electrical and electronic equipment, were located outside the pressure vessel. Unlike the Mercury spacecraft, which had nearly, all systems inside the pressure shell, the Gemini spacecraft had most system components located in unpressurized equipment bays. These components either required no pressurization or were internally pressurized. Since equipment was normally only one layer deep within the compartments, launch crews could remove a hatch, quickly pull out a malfunctioning unit and insert a new one, reinstall the hatch and proceed with the launch.

Experiments for NASA and the US Air Force were installed in bays on the "bottom" of the reentry module, in the adapter section, and in the pressurized crew compartment.

Personnel Hatches

Two hatches, contoured to the shape of the conical cabin, were located in what was the top of the spacecraft during orbital flight. A hatch was located over each astronaut and was manually operated by handles inside and outside the spacecraft. The latching mechanism was mechanical. The hinge was on the outboard side of the door.

Each hatch incorporated an observation window consisting of one outer and two inner panes of glass with an air space between each pane. The outer panes were high-temperature 96% silica glass. The innermost pane was of temper-toughened alumino-silicate glass for structural strength. The surface of each pane, with the exception of the outer one, was coated to reduce reflection and glare and to aid in attenuating ultraviolet radiation.

Skin and beam construction made up the structural design of the personnel access hatches. A silicon rubber seal around each hatch sill and around the two inner panes of glass in the window prevented loss of cabin pressure when the hatches were closed.

A hatch curtain was stowed alongside the hinge of each hatch, which, after a water landing, prevented water from entering the cabin when the hatches were opened. In emergency situations the hatches opened by a 3-sequence operation actuated by a pyrotechnic (explosive) device. When initiated, these actuators unlocked the mechanical latches, opened the hatches, and finally supplied a hot gas that ignited a seat-ejection catapult rocket.

Heat Shield

The heat shield was a dish-shaped structure that formed the large end of the reentry module.

The ablative substance of the Gemini heat shield was a paste-like material, which hardened in standard atmosphere after being poured into a honeycomb form.

Starting with a load-carrying Fiberglas sandwich structure consisting of two 5-ply faceplates of resin-impregnated glass cloth separated by a 0.65inch thick Fiberglas honeycomb core, an additional Fiberglas honeycomb was bonded to the convex side of the sandwich and filled with Dow-Corning DC-325 ablative material. The entire shield was encircled with a Fiberite ring. The basic ablative substance of the heat shield was developed by McDonnell and was marketed by Dow-Corning.

Shingles

The surface of the reentry module was covered with overlapping shingles, which provided aerodynamic and heat protection and held in place shaped pads of flexible insulation. The composition of the beaded (corrugated) Rene 41 shingle (0.41 mm thickness) on the sides of the cabin was 53% nickel, 19% chromium, 11% cobalt, 9.75% molybdenum, 3.15% titanium, 1.6% aluminum, .09% carbon, .005% boron, and less than 2.75% iron. The shingles were identical in composition and manufacturing technique to those used on Mercury. Extra large holes in the shingles at the attachment bolts allowed each to expand during aerodynamic and solar heating. Oversize washers covered these holes to minimize heat and air flow penetration.

The R and R and RCS section surfaces were unbeaded shingles of cross-rolled beryllium. The plate was supplied to McDonnell, by Brush Beryllium Company, in sheets ranging in thickness from 7.6 mm to 14.1 mm and was finished by McDonnell to a thickness of 2.28 mm to 7.11 mm. The shingles were attached to the spacecraft by beryllium retainers fabricated from similar plates.

Beryllium was previously used on Project Mercury using sections fabricated from hot-pressed beryllium blocks rather than cross-rolling procedures. The new manufacturing technique permitted much higher strength and shock resistance. Gemini rendezvous flights required almost twice the strength and impact resistance available with hot-pressed beryllium blocks.

Both Rene 41 and beryllium shingles were coated with ceramic paint on the outer surface to permit high thermal radiation from the spacecraft. The inner surface of the beryllium shingles had a very thin gold coating to attenuate thermal radiation into the spacecraft.

Adapter Module

The most obvious structural difference between the Gemini and Mercury spacecraft was the integral adapter module, which was part of the orbital configuration of Gemini.

The adapter module, 2.29 m high and 3.05 m across its largest diameter, provided volume for systems and equipments needed for long-duration orbital flights, as well as the mating structure between the reentry module and the launch vehicle.

The adapter was a ring stiffened skin-stringer structure consisting of circumferential aluminum rings, extruded magnesium alloy stringers, and magnesium skin. The T-shaped stringers had a hollow bulbous portion to provide a path for the flow of liquid coolant, which transferred heat to the adapter skin for subsequent radiation into space.

The outer surface of the adapter module was coated with white ceramic paint and the inner surface was covered with aluminum foil to reduce emissivity. The adapter was joined to the reentry module by three titanium tension straps external to the structure of both the reentry module and the adapter section.

Retrograde Section

At the small end of the adapter module was the retrograde section containing crossed aluminum "I" beams on which were mounted four retrograde rockets. This section made up the first 100 cm of the adapter module. One retrograde rocket was mounted in each quadrant of the section. In addition to retrograde rockets there were six orbit attitude and maneuvering system (OAMS) thrust chamber assemblies. Four of these assemblies permit orbital translation, up, down, left and right. Non-operating "dummy thrusters" were installed in place of these four assemblies on Spacecraft. No. 3. Two, whose nozzles face toward the reentry module, provided for the rearward or "backing away" maneuver of the spacecraft.

Equipment Section

At the large end of the adapter module the equipment section provided volume and attach points for several system modules, including orbit attitude and maneuvering system propellant tanks, the environmental control system primary oxygen supply, batteries, coolant, and electrical and electronic components. A honeycomb blast shield between the two sections protected the equipment section and the dome of the Titan launch vehicle from excessive (explosion-causing) heat should it be necessary to fire the retrorockets in an abort condition.

Ten OAMS thrust chamber assemblies were mounted in the equipment section providing for roll, pitch, yaw, and forward maneuvering of the spacecraft during orbital flight.

A Fiberglas cover over the open end of the adapter protected the equipment inside from solar radiation after separation from the launch vehicle. A forged aluminum alloy ring mated the spacecraft and the Titan II launch vehicle, 3.05 m in diameter with 20 lugs through which bolts were fastened to secure the mating. When the spacecraft was separated from the launch vehicle, a pyrotechnic charge was fired to sever the adapter section approximately 40 cm above the launch vehicle/spacecraft mating point. This charge cut through the metal skin of the adapter section instantly but in the same way a metal shear would cut through sheet metal.

The adapter structure was constructed as a single unit. The two sections were separated by a shaped-charge pyrotechnic device prior to reentry.

Communications And Tracking

The communication and tracking system provided two-way voice communication. ground-to-spacecraft command link. spacecraft-to-ground telemetry transmission, radar tracking signals and recovery aids. Subsystems consisted of telemetry, tracking. voice communications, digital command, antennas. and recovery aids.

Voice Communications

The voice communications subsystems included the voice control center, the UHF voice transmitters-receivers and HF voice transmitter-receiver, built by Collins Radio. The voice communication subsystem was operational from prelaunch through postlanding.

The voice subsystem provided communication between the astronauts, between the blockhouse and the spacecraft during launch, between ground stations and the spacecraft from launch through reentry, and between astronauts and frogmen during the water recovery. The voice subsystem also provided communication between the spacecraft and recovery forces during landing and postlanding.

During the reentry, voice communication would be lost on two occasions. Under worst conditions. the first extended from approximately 1310 seconds after retrofire to 1775 seconds after retrofire, nearly 8 minutes (under nominal conditions, this time was approximately 6 minutes). This loss of communication was caused by ion sheath formation around the spacecraft. The second period lasted about 30 seconds, starting at main parachute deployment, and was due to a delay between loss of the nose stub transmitting antenna and the erection of the descent antenna.

The voice control center provided for intercommunication between the astronauts, for control and distribution of audio to and from the transceivers, and supplied a tone for direction finding (HF-DF). The voice control center provided for individual selection and control of various functions in UHF, HF and intercommunication circuits.

Dual controls permitted mode switching and volume controls for HF, intercommunications, and UHF. A common section in the panel provided squelch control of UHF and HF, receiver selection of HF and UHF, and keying.

The HF voice transceiver provided for over-the-horizon spacecraft-to-ground communication and a direction finding signal when in the HF-DF mode.

The voice control center was located in the pressurized cabin of the Gemini spacecraft. Two UHF transceivers and one HF transceiver were located in the reentry module of the spacecraft outside the pressurized cabin.

Digital Command

The Motorola built digital command subsystem consisted of a receiver-decoder and associated relay units, which permitted spacecraft utilization of ground commands. Located in the equipment section, this system was operational from pre-launch until jettison of the equipment section. The digital command system received and decoded command transmissions from the global network of ground stations and transformed them into a digital format.

Digital commands were categorized as either "real-time commands" or "stored program commands." Real-time commands operated DCS relays that controlled equipment input power or energized relays in the spacecraft electrical system to control equipment usage. The stored program commands were used to provide such units as the time reference system and the computer with updated data.

Antenna Subsystem

The antenna subsystem consisted of antennas, coaxial switches, a diplexer and a quadriplexer. The antenna subsystem was operational from prelaunch through postlanding and provided radiation coverage for all communication and beacon tracking signals between ground stations and the spacecraft. The system included C-band tracking helical, C-band slot, S-band, HF whip, UHF whip, descent, recovery, and UHF nose stub antenna.

The antenna system provided radiation coverage for the communication system during all mission phases. Coverage varied with the mission phase depending upon spacecraft stabilization mode and ground coverage requirements. During the launch phase, when continuous C-band and UHF coverage were required for flight safety reasons, the antenna system provided roll symmetrical antenna patterns to optimize the ground coverage.

During the orbital phase, S-band and HF were added. The orbital antenna system provided yaw symmetrical, horizon- oriented, hemispherical patterns for optimum coverage in stabilized orbit attitude. For drifting flight with uncontrolled spacecraft attitude, the antenna/communication system provided complementary coverage.

Complementary coverage was obtained by use of yaw symmetrical and roll symmetrical antenna patterns.

The communications systems used these patterns simultaneously.

The astronauts select the antenna system to obtain the optimum pattern for voice and telemetry. During the reentry phase, UHF and C-band coverage was identical to launch phase coverage. During the recovery phase, antenna capability was provided for HF and UHF.

Telemetry Subsystem

The telemetry subsystem provided real-time, delayed-time and standby telemetry transmission.

The frequency modulated telemetry transmitters were employed during all phases of the Gemini mission when the spacecraft was in contact with the ground stations. These transmitters were energized either by the astronauts or automatically by the ground digital command system. The standby transmitter was used as a replacement for either the real-time or delayed-time transmitter in event of a failure. A delayed-time transmitter sent pulse code modulated information, which was stored in the tape recorder of the data transmission system. The real-time transmitter sends current information from the data transmission system programmer.

Recovery Aid Subsystem

The recovery aid subsystem included a UHF recovery beacon, a UHF rescue beacon transceiver and a flashing light. The HF voice transmitter in the reentry module was tone modulated in the HF-DF mode as part of recovery aid equipment.

A pulsed UHF output signal (energized upon landing impact) supplied continuous direction-finding information for recovery forces on the international distress frequency, and the flashing light provided the visible indication of the spacecraft location should recovery operations be conducted at night. The light was designed to be visible from an altitude of 12,000 feet at a distance of 50 nautical miles on a starlit, moonless night. The recovery aid subsystem was operational only during landing and postlanding phases of the Gemini mission.

Tracking Subsystem

The tracking subsystem included a C-band radar beacon, an S-band radar beacon and an acquisition aid beacon. The C- and S-band beacons provided tracking responses to interrogation signals from ground stations. Either or both of these beacons could be energized by ground command via the DCS. The astronauts could also energize either beacon, and select the antenna system for the C-band beacon to achieve a roll symmetrical or yaw symmetrical pattern. The acquisition aid beacon provided a radio frequency signal from the spacecraft to ground communication facilities for spacecraft acquisition.

The C-band radar beacon was operational from prelaunch through the landing phase, the S-band radar beacon was operational from prelaunch to immediately prior to retrograde and the acquisition aid beacon was operational from prelaunch to immediately prior to retrograde. S- and C-band beacons were built by ACF Electronic Division. The acquisition aid beacon was built by Vector Manufacturing Company.

Electrical System

The Gemini GT-3 spacecraft electrical system was a two-wire, grounded system using silver-zinc Eagle Picher batteries as sources of 25 VDC electrical power. There was no primary AC electrical power system on the spacecraft. Devices utilizing AC power obtain that power from self-contained inverters within the individual systems.

Prior to launch, external electrical power was provided to the spacecraft through the umbilical to prevent undue depletion of the spacecraft power supply. The battery subsystem was capable of supplying sufficient power to the electrically operated equipment for all phases of the planned GT-3 mission. In addition, sufficient power was available for a prelaunch period of two hours and a postlanding period of 36 hours for operation of necessary recovery equipment and for 12 hours of post landing suit compressor operation.

Ten silver-zinc batteries were provided, each activated and sealed at sea level pressure. The battery cases were vented to permit the escape of gases. Battery temperatures were controlled by mounting the battery cases in direct contact with spacecraft coldplates.

The reentry module main batteries supply a portion of the main bus electrical power during launch and all of the main bus electrical power during reentry, landing and postlanding. These four batteries were located in the right-hand equipment bay outside the pressurized area of the reentry module.

The squib batteries supply electrical power for squib-activated pyrotechnic devices throughout the entire mission. Squib batteries were isolated both electrically and mechanically from all other batteries. The three squib batteries were located ir the right-hand equipment bay outside the pressurized area of the reentry module.

It was possible to connect squid circuitry to the reentry main batteries in case of squib battery malfunction.

The three adapter batteries supply power to the main bus and were capable of supplying all of the electrical power necessary for spacecraft operation until separation of the adapter module.

To insure smooth system operation, complete electrical system management by the crew was provided. Extensive circuit protection was incorporated throughout the system and indicators were mounted on the instrument panel for use by the astronauts in systems monitoring.

Environmental Control System

The Gemini environmental control system was capable of providing life support for the biological systems of two astronauts. It provided for ingestion of appropriate gases and fluids, and the dispersal of by-products which were created, as well as for cooling of spacecraft equipment, and the cabin interior.

The system consisted of a water management subsystem, an oxygen supply subsystem, and a cooling subsystem. It provided gaseous oxygen for breathing, for suits and cabin pressurization, and for suit and cabin ventilation. It provided for the removal of small solids, carbon dioxide, odors, and moisture from the suit and cabin atmosphere. For the flight of GT-3, it provided a drinking water supply-. It provided for storage and disposal of water accumulated as a condensate, and for disposal of urine. It included a dual, recirculating coolant system for regulating the temperature of the suits, the cabin and items of electrical equipment.

Primary Oxygen Subsystem

The primary oxygen subsystem stored and dispensed oxygen for breathing and for suit and cabin pressurization. It supplied oxygen during the entire flight, commencing two hours prior to launch and terminating with the equipment section jettison at retrograde.

Oxygen pressure in the crew compartment was limited to 5 5 to 6.0 pounds per square inch above ambient by the cabin pressure relief valve. Primary oxygen supply capacity for a two-day mission was 15.3 pounds, located in a single spherical container in the equipment section, and was the primary source of oxygen during prelaunch, ascent and orbit. Oxygen was stored in this container in cryogenic form. It was heated to a gaseous state by a heat exchanger, passed to a pressure reducing regulator, then on to the cabin pressure regulator which automatically maintained cabin pressure as desired for the mission. Oxygen remaining in the primary supply was jettisoned with the tank when the equipment section was jettisoned prior to retrograde. The two-day absorber cartridge removed both odors and up to eleven pounds of carbon dioxide. The suit heat exchanger transfers heat from suit circuit oxygen to coolant flow. The heat transfer capacity was 1500 BTU per hour. The exchanger also removed moisture from the suit circuit oxygen and transfers it to the water management subsystem.

Pressure was maintained automatically- in the cabin through all phases of the Gemini mission. During the launch phase, cabin pressure relief valve permitted outflow of any overpressures which might exist. This valve seals the cabin when the ambient pressure falls to 5.5 pounds per square inch below cabin pressure. This would occur between 20 and 40 seconds after launch.

If the cabin should decompress for any reason, the supply of oxygen provided through the dual cabin pressure regulator automatically turns off when the pressure reached 4.0 pounds per square inch. When decompression occurs, either as programmed or as a result of a malfunction, the astronaut's pressure suit automatically took over the pressurization responsibilities previously provided as a cabin environment.

If one or both astronauts chose to work with the face-plate of their suits open, a manually operated valve permitted circulating cabin air through the suit circuit for carbon dioxide and water vapor removal. The same valve provided for relief of the vacuum created at the compressor inlet if the snorkel valve was momentarily closed by water after a water landing.

The crew compartment was maintained at a nominal 65 degrees F during orbital flight. It was expected to rise to a high of approximately 120 degrees F during reentry. The maximum acceptable temperature in the crew compartment during launch and reentry of the Gemini spacecraft was considered to be 200 degrees F.

Secondary Oxygen Subsystem

Secondary oxygen was contained in two tanks in the pressurized compartment of the reentry module. There was sufficient quantity in each of the secondary oxygen supply tanks to provide oxygen adequate for one orbit at a normal flow rate and reentry at a nominal oxygen high rate of 0.08 pounds per minute to each astronaut.

The secondary oxygen subsystem operated when the pressure in the primary oxygen line falls below an allowable 75 pounds per square inch. When the primary oxygen container was jettisoned. the secondary oxygen subsystem assumed the primary role.

The oxygen line from the suit-demand and cabin pressure regulators was common to both the primary and secondary supplies so that flow was continuous to the suit circuit if a malfunction occurs.

Egress Oxygen Subsystem

Oxygen for breathing and for suit pressurization in the event the astronaut's initiate ejection at 45,000 feet or below during launch or reentry was provided by the egress oxygen subsystem.

A tank containing approximately 1/3rd a pound of usable oxygen was located in the seat-mounted egress kit for each astronaut. The egress oxygen circuit was lanyard-opened with seat ejection.

Water Management Subsystem

The water management subsystem collected and stored water for drinking, dumps waste water overboard, and managed the water used for cooling. Components of the subsystem include a water tank, a urine receptacle, a drinking nozzle, controls and valves, an evaporator, a reservoir, and a water pressure regulator.

The water was stored in a tank located in the equipment section. For GT-3, it contained approximately 16 pounds of water stored at 7.5 psig. A second water tank in the reentry module also contains approximately 16 pounds of water. Another seven pounds of liquid was contained in the launch cooling heat exchanger reservoir. A tank in the equipment section was charged with oxygen at 1,000 psig and this pressure was available upon demand to pressurize the water storage tanks. Oxygen-pressurized diaphragms forced the water in the line to the cabin water tank and to the water dispenser.

Urine Disposal

Urine Disposal equipment was Government furnished, installed by McDonnell. It consisted of a urine line, bellows assembly, quick disconnect coupling, and a uriceptacle. On GT-3, urine was routed to the water evaporator for disposition upon actuation of the cabin water dump valve.

Temperature Control Subsystem

A distinctively new feature of the Gemini spacecraft was the fluid coolant system, which maintained cabin temperature, astronauts' suit temperature, and equipment temperature within acceptable limits.

Since the Gemini equipment and crew would generate heat at approximately three times the rate of the Mercury spacecraft, and would do so for almost ten times as long, it was necessary to provide a new method of heat rejection, namely, a space radiator. The entire outer skin of the adapter module served as the radiating surface, and the hollow stringers, through which the coolant passes, transferred the heat to the skin.

At the heart of this system was the space radiator in the adapter module, a coolant fluid reservoir, a low-level coolant-sensing device, two identical positive displacement pumps for each of two redundant loops of coolant lines, and two regenerative heat exchangers. The redundant systems provided protection against loss of a loop due to failure such as meteorite penetration.

A silicon ester coolant fluid (Monsanto's MCS 198) was pumped through each, then through heat exchangers, which heated the primary cryogenic oxygen supply, cooled the cabin and suits, cooled such equipment as electrical power supplies, and various electronic equipments. The cabin and suit heat exchangers were conventional heat exchangers, and the electronic equipment heat exchangers were coldplates on which the equipment was mounted.

In the cabin each coolant circuit was divided into two parallel paths, one for the cabin and one for the suit. Ahead of each heat exchanger was a manual valve, which permitted crew adjustment of the flow through the heat exchanger.

In the section involving coldplates (there were 24 of them in each spacecraft), as many of them as possible were arranged in parallel to minimize pressure loss. However, all of the flow was required through some high power density units, and these coldplates were arranged in series.

Two regenerative heat exchangers in the equipment section were provided for each of the loops. A temperature-sensitive valve modulated to maintain an outlet coolant temperature at between 36 degrees F and 42 degrees F. When the coolant temperature was 36 degrees F or lower, the valve caused all coolant to flow to the regenerative heat exchanger. When the coolant temperature was 42 degrees or above, the full coolant flow was directed to the space radiator. The adapter structural shell provided an area of 165 square feet to radiate the excess heat into space.

Each of the redundant coolant loops followed alternate stringers in the adapter, so that when operating on one loop, the tubes of the other loop were prevented from freezing by heat from the adjacent warm tubes. The coolant fluid, however, had a freezing point well below -100' F. and had a characteristic of low viscosity at extremely low temperatures.

During preflight operation a ground heat exchanger removed heat from the spacecraft by circulating a coolant fluid from a servicing cart through a secondary loop in the heat exchanger. During launch, and for a short period thereafter, the space radiator was too warm for effective cooling. Then, a water boiler cooled the coolant fluid by the simple process of absorbing the heat through vaporization.

In thermal balance tests run in the 30-foot space chamber at McDonnell, it was determined that the emissivity for the GT-3 spacecraft, which did not had fuel cells, was actually higher than desired. The GT-3 adapter was striped with .66 inch aluminum tape with black coating for deliberate and controlled reduction of the total heat emission of the radiating surface. There were 72 strips on the retrograde section, and 88 on the equipment section. A tape strip was aligned vertically between each stringer.

The temperature control subsystem circuits, which connected with the reentry module, were lost when the adapter section was jettisoned. Temperature control subsequent to initiation of retrograde was provided by cabin wall insulation and the heat shield.

Supercritical Storage of Cryogenic Fluids

Primary oxygen was stored in the equipment section of the Gemini spacecraft in a cryogenic state in especially designed cryogenic containers. Because cryogenic storage was at a lower pressure than tanks of equal size holding the same quantity of oxygen, danger of structural failure was reduced; associated components were lighter; and the heat absorption capacity of the fluid was available to help dispose of spacecraft waste heat.

The difference in weight between a vessel for storage of oxygen as a gas and in its cryogenic state was significant. Gaseous oxygen required a heavily constructed tank to withstand a pressure of 7,500 pounds per square inch. In the supercritical cryogenic system used in Gemini, oxygen was first stored as a liquid, then in a confined tank volume, it was heated to a point where it became a homogeneous compressed fluid. This process took place in the cryogenic storage container at a pressure of only 750 pounds per square inch. A cryogenic storage subsystem weighed less than four-tenths of a pound for each pound of oxygen stored. A high-pressure vessel of comparable capacity would weigh five times as much.

Gemini's supercritical storage of fluids insured a steady delivery of oxygen under all gravity conditions just as high-pressured gas does.

Guidance and Control

The guidance system included the inertial guidance system (IGS), information interfaces with the digital command system (DCS) and the data acquisition system (DAS), with the time reference system (TRS) functioning as an integral part of the guidance system. This system performed navigation and steering logic computation for trajectory management. The guidance system output provided guidance information directly to control systems and/or to instrument displays which the crew interpreted for manual control inputs. The Gemini guidance computer, guidance analysis, and guidance system integration subcontractor was IBM Federal Systems Division's Space Guidance Center.

The control system provided commands to the thrusters, acting in response to outputs from the horizon sensors, the computer, attitude control and maneuver handles, and the attitude control and maneuver electronics (ACME). The propulsion system was an integral part of the control system. Both automatic and manual modes of operation provided angular control of the spacecraft about its three major axes and for maneuvering. spacecraft translational maneuvers were manually controlled.

Inertial Guidance System

The inertial guidance system included an inertial measuring unit (IMU), digital computer, manual data insertion unit (MDIU), incremental velocity indicator (IVI), and an auxiliary computer power unit (ACPU).

The inertial measuring unit, built by Honeywell, was a stabilized inertial platform including an electronic unit and a power supply. The inertial measurement unit provided a stable attitude reference and incremental velocity data.

The inertial platform was a four-gimbal structure with an all-attitude capability. The gimbals, in sequence, starting with the innermost gimbal, were pitch, inner-roll, yaw, and outer-roll. The stable element contained three single-degree-of- freedom rate-integrating gyros and three accelerometers of the force rebalance type, mounted orthogonally (mutually perpendicular). Each gimbal contained two analog pickoffs for angle measurement between body and platform axes. One set of pickoffs operated in conjunction with the computer to generate digital representation of the angles. The second set operated with the attitude display group to provide attitude information to the flight crew.

The IMU subsystem electronics provided circuitry and equipment for alignment, stabilization, and gimbal torquing, and accelerometer rebalancing. IMU malfunction detection circuitry was also included.

The IGS power supply operating from 28-volt DC spacecraft power provided the alternating current and direct current power necessary for operations of the IMU, the computer, the MDIU, the IVI, the ACME, and the horizon sensors. The attitude control and maneuver electronics had a separate power supply, which could be operated when the inertial guidance system was not operating. The IGS power supply provided alternating current for components normally supplied by the attitude control and maneuver electronics unit, in the event of a ACME power supply failure.

Digital Computer

The onboard digital computer was a binary, fixed point, stored program, general purpose, digital computer that provided executive functions, computations, timing and signal processing for spacecraft guidance and control. The Gemini computer had a memory of 4,096 words (39 bits/ word). It was random access with a non-destructive readout. Division of the memory into three syllables provided flexibility in instruction and data storage location assignment. It could add, subtract, and conduct a transfer operation in 140 microseconds. Multiplication (full precision) was accomplished in 420 microseconds. Division (full precision) was achieved in 840 microseconds. Multiplication and division could be programmed concurrently with addition and subtraction or transfer operations. The clock rate of the Gemini digital computer was 500 KC arithmetic bit rate; 250 KC memory cycle rate.

Computer programs included an executor program, operational programs and standard computational sub-routines. The executor program performed diagnostic checks, determined elapsed time, selected the desired operational program, and performed all data input/output sub-routines.

The basic computer program was inserted at McDonnell prior to computer installation in the spacecraft. After installation, minor updating of constant words and time variable words was accomplished by the digital command system or the manual data insertion unit and was monitored by the data acquisition system. Operational program requirements were functions of the computer mode. These modes were varied in accordance with the missions to be accomplished by the individual spacecraft. For Spacecraft 3, four modes were available: Prelaunch, Ascent, Catch-Up, Reentry.

In the prelaunch mode, the computer was programmed to perform diagnostic checkout routines and sub-routines. In this mode, sum checks were performed in the section of the computer memory that contained constants and fixed programs that could not be varied during the mission.

Guidance in the ascent mode served a backup function in the event of malfunction of the radio guidance system (RGS).

In the ascent mode, information was displayed for flight crew evaluation of mission status. The computer also provided navigational data for reentry guidance in the event an abort should become necessary during the launch phase.

While the RGS had primary responsibility for launch vehicle guidance from liftoff through launch vehicle/spacecraft separation, the inertial guidance system provided backup guidance by effectively duplicating RGS functioned in an inertial frame, defined by the launch pad vertical and the desired azimuth at insertion. Stage 1 of standby guidance consisted of a time-sequenced roll and pitch program based on the design constraints of the RGS. Stage 2 guidance also satisfied the guidance constraints defined for the RGS. The ascent program could be updated by the digital command system with ground-determined velocity data to reduce the effects of IGS uncertainties. In the event of a guidance switchover, the IGS provided faded (smoothed) attitude error signals to the launch vehicle secondary autopilot. Switchover could occur at anytime. Should switchover occur before liftoff, the switchover signal simultaneously directed engine shut-down, preventing liftoff on backup guidance.

The IGS provided orbit insertion guidance so that the effect of insertion errors on apogee and perigee altitude was nullified. The velocity increment to be added for proper insertion was computed from IGS navigational data and desired insertion condition data stored in the onboard computer. The velocity increment to be added was displayed on the incremental velocity indicator. The attitude error command necessary for proper vehicle orientation was displayed on the flight director indicator. The inertial guidance system then performed navigation computations as the insertion sequence proceeds, driving the incremental velocity indicator display in accordance with the velocity increment added. Orbit insertion guidance began immediately after a normal second-stage engine cutoff (SSECO), or in the event of a premature engine shut-down, at least 300 seconds from liftoff.

After the incremental velocity was applied at perigee, the flight crew could read out of the computer, via the MDIU, an incremental velocity to be added at apogee to correct the perigee altitude. The time to apply this velocity could also be read by the MDIU.

In the event of a launch abort after a velocity greater than 21,000 feet per second was achieved, reentry navigation equations were initiated from the position and velocity information generated by the inertial guidance system. Thirty seconds after equipment section separation, an event initiated by the flight crew, the computer would generate transformation coefficients to relate the platform coordinate system to that required for reentry guidance equations. Following initiation of reentry navigation, the data flow was identical to that which occurred in the reentry mode.

Between orbit insertion and reentry, the inertial guidance system was operated in the catch-up mode. Various maneuvers were performed in this mode on the GT-3 flight to study the spacecraft's capabilities relative to future missions. The catchup mode provided a means of displaying and applying a ground-command velocity change. The time-to-go to apply thrust and the velocity change to be applied were computed at the integrated mission control center on the ground, based upon knowledge of the time of liftoff or upon ground tracking information or both. Time and velocity change components were transmitted to remote tracking sites for transmission to the spacecraft by voice and digital communication. Velocity change was normally transmitted by the digital communication system; however, it could be transmitted by voice and inserted into the onboard computer by the astronauts via the manual data insertion unit. In addition to the MDIU and DCS capability, the velocity change to be applied could also be displayed by the incremental velocity indicator. Spacecraft maneuver commands were manually applied using orbit attitude and maneuvering system thrusting.

Reentry Guidance Application

The Gemini spacecraft was designed for guided reentries from orbit to permit spacecraft landing at pre-selected, prepared sites. This capability was provided through the use of the inertial guidance system to develop steering commands for proper orientation of the aerodynamic lift vector arising from an offset center of gravity position.

Preparation for reentry began significantly ahead of retrograde to precisely establish the inertial platform reference and the initial conditions for navigation computation in the reentry mode. Following retrograde, the retrograde section was manually jettisoned and the spacecraft manually controlled to a flight crew "heads down" full-lift attitude. Reentry steering was initiated when the spacecraft entered the atmosphere and total acceleration achieved a prescribed predetermined value. Roll commands controlled horizontal and vertical components of lift as a function of downrange and cross-range errors. Roll errors were supplied to the attitude control and maneuver electronics for automatic control and also to the flight director indicator for manual control capability. When the density/altitude parameter reached a pre-selected point within the atmosphere, a maximum lift attitude was commanded and reentry guidance was terminated.

The reentry configuration for Gemini had a center of gravity location resulting in angle-of-attack conditions producing an appreciable lift vector during the atmospheric regions of the reentry. The roll attitude of the spacecraft established the direction of lift application; zero bank angle produced range extension, banking to the right a left turn, banking to the left the opposite.

For standard reentry, ground computers determined the future Earth tracks of the spacecraft, determined approximately when the desired landing site would become available, and then performed iterative trajectory solutions to establish a retrograde time that would require about one-half the downrange extension provided by the reentry lift. This time, along with the position and velocity associated with the retrograde time and the target coordinates, were transmitted to the spacecraft and inserted into the IGS.

The spacecraft computer, using the initial conditions transmitted from the ground, maintained a knowledge of the trajectory through conventional navigation equations. It then predicted a touchdown on the basis of "zero lift" and compared the predicted touchdown point with the desired target site. Downrange and crossrange errors were used to develop bank angle commands.

The system utilized the ground complex to develop a very accurate orbit position for use in subsequent reentry guidance. Ground stations locations were planned to maintain excellent coverage of the spacecraft during the majority of its flight.