Item	Nomenclature	Function
1	Separation Assembly, FLSC	Severed the spacecraft from the launch vehicle and the adapter equipment section from the adapter retrograde section.
2	Guillotine	Cut cables and wire buddies.
3	Cutter-Sealer	Cut and crimped seal the OAMS fuel and oxidizer lines.
4	Horizon Sensor Fairing	Jettisoned the horizon sensor fairing.
5	Horizon Sensor Release	Jettisoned the horizon sensors.
6	Ventilation Door Release	Jettisoned the ventilation door.
7	Cartridge, Pyrotechnic Valve	Opened or closed pressurant, propellant, or oxidizer lines.
8	Pyrotechnic Switch	Opened circuits prior to guillotine operation.
9	Separation Assembly, MDF	Separated rendezvous and recovery section from the reentry module.
10	Retrograde Rocket igniters	Ignited the retrograde rocket motors.
11	Retrograde Rocket	Provided thrust for velocity decrement necessary for reentry.
12	Parachute Mortar and Reefing Cutter	Parachute mortar deployed parachute in reefed condition, reefing cutters allowed complete parachute deployment.
13	Parachute Disconnect	Changed reentry module suspension and jettison parachute.
21	Separation Assembly, FLSC	Severed the reentry module from the retrograde section.

Escape Mode

14	MDF initiation System	Initiated the hatch actuators.
15	Hatch Actuator	Opened hatch before ejection seat operation.
16	Rocket Catapult	Provided ejection seat propulsion during ejection seat operation.
17	Harness Release Actuator	Unlatched seat restraint harness after seat ejection.
18	Seat-Man Separator	Separated astronaut from seat.
19	Ballute Deploy and Release Assembly	Deployed and release ballute according to altitude.
20	Drogue Mortar-Backboard Jettison Assembly	Deployed astronaut drogue parachute.

Electrically-initiated devices incorporated cartridges and detonators. They were used in the flexible linear shaped charge assemblies, guillotines, cutter sealers, release assemblies, valves, pyrotechnic switches, mild detonating fuse separation assembly, retrograde rocket initiators, parachute mortars, and parachute disconnects. The cartridges and detonators were provided with one or two independent electrical bridge wire circuits as reliability dictates.

Lanyard-initiated devices were the parachute reefing cutters, mild detonating fuse initiation system, harness release actuators, ballute deployment and release systems, and drogue mortar-backboard jettison assemblies.

Pressure-initiated devices were utilized in the hatch actuators, rocket catapults, and seat-man separators. These devices fired when exposed to pressures between 500 and 3000 pounds per square inch except for the rocket catapult which fired when exposed to 300 to 9000 pounds per square inch.

Flexible linear-shaped charge assemblies were sections of electrically initiated charges capable of separating sheet metal, wire bundles, straps, and tubes. Three such assemblies were provided to separate the launch vehicle, the adapter equipment section, and the adapter retrograde section from the reentry module. The assemblies completely severed inter-connecting tubes, wires, bundles, titanium straps and structural skin of the sections.

Guillotines were provided for cutting wire bundles, cables, and bolts. These were knife-like devices explosively driven through the item to be severed. Redundancy was provided through use of two guillotines, one on either side of the separation plane. The guillotines used for severing cables or bolts contained two cartridges. Function of either cartridge was sufficient to sever the cable or the bolt.

Cutter-sealers provided for sealing and cutting of the orbit attitude and maneuvering system propellant lines prior to section separation. The lines were sealed to prevent leakage of fuel or oxidizer residue at the same instant that the tubes were severed. Redundant cutting was provided by using two cutter-sealers, one on either side of the separation plane. Each contained one cartridge.

The horizon sensor fairing release assembly was secured to the spacecraft by a cartridge-activated single-point hold- down device containing two cartridges The cartridge was sufficient to jettison the fairing. The horizon sensor fairing release assembly was enclosed to preclude the escape of any pyrotechnic by-products, thus protecting the delicate horizon sensors against damage during fairing jettison.

The horizon sensor was secured to the spacecraft by a two-cartridge, single-point, hold-down device. The function of a single cartridge was sufficient to jettison the sensors and the cartridges from the spacecraft. This was accomplished following retrograde and prior to reentry. The electrical connectors, through which the cartridges were initiated, were disconnected by the force of ejection.

Normally open or normally closed pyrotechnic valves were provided for isolation and control of pressurants and propellants. Normally closed pyrotechnic valves isolated the pressurants and propellants in the tanks of the orbit attitude and maneuvering system and the reentry control system from the remainder of the system during prelaunch. Orbit attitude and maneuvering system valves were pyrotechnically opened shortly before launch. The normally closed reentry control system valves were opened prior to reentry. The orbit attitude and maneuvering system was provided with both normally closed and normally open pyrotechnic valves actuated in the event of pressure regulator malfunction.

Pyrotechnic switches were provided for opening electrical circuits in the wire bundles prior to severing. Each pyrotechnic switch contained a single cartridge with dual bridge wires.

A mild detonating fuse separator was provided to break all bolts that attached the rendezvous and recovery section to the reentry module. The separation assembly contained two detonators and an explosive ring assembly. Booster charges initiated by the detonators caused simultaneous detonation in both mild detonating fuse strands of the explosive ring assembly. Either strand had sufficient energy to break all the attachment bolts.

Each retrograde rocket contained two independent cartridge-actuated igniters mounted adjacent to the rocket nozzle. Each igniter contained an electrically actuated cartridge, 20 boron/potassium nitrate pellets, and a 0.04-pound solid propellant charge of polysulfide/ammonium prechlorate. When fired, the cartridge ignited the boron/potassium nitrate pellets which in turn ignited the solid propellant charge. The propellant would burn for 0.35 second, discharging its exhaust gases into the retrograde rocket cavity and igniting the basic retrograde rocket. Either of the two igniters provided for each rocket was sufficient to ignite the retrograde rocket.

The parachute mortar which deployed the pilot parachute, contained two electrically-initiated cartridges. Function of a single cartridge was sufficient to eject the pilot parachute from the mortar tube. The pilot parachute was deployed in a reefed condition.

Two lanyard-initiated reefing cutters, incorporating a 6-second pyrotechnic time delay, were provided to complete pilot parachute deployment. Three lanyard-initiated reefing cutters, incorporating a 10-second pyrotechnic time delay, were provided to complete main parachute deployment.

After the main parachute had deployed and the spacecraft was suspended from a single point, a parachute disconnect was actuated to allow the spacecraft to rotate to a two-point suspension. Two parachute disconnects jettisoned the main parachute upon impact. Each of these disconnects contained two electrically initiated charges.

Spacecraft Fabrication

Manned spacecraft fabrication techniques developed at McDonnell for the Mercury spacecraft were extensively applied in the Gemini program. Weight limitations combined with heat resistance, air load, and acoustic requirements necessitated improved fabrication practices and placed even greater demands on assembly techniques. New welding techniques were developed at McDonnell to meet many of these demands.

Adapter Welding

The adapter skin was a magnesium-thorium alloy of 0.032-inch thick sheet stock. Welded on the inside was a continuous magnesium-thorium tee-bulb extrusion that served both as a stiffening stringer and a closed passage for the coolant that converted the adapter to a radiator. Intimate contact between skin and extrusion to facilitate heat transfer was accomplished by using a weld-through sealer for conductivity and as a moisture barrier. Seam welding then mated the parts.

These skins were fabricated in quarter panels of two sizes. Radiator extrusions on each panel were joined on assembly by using filler wire and a small hand torch to make fillet welds at each end of the sleeve joint. Since the radiator extrusion was already seam welded to the adapter skin, a mirror was necessary to make this difficult weld in the confined space.

Penetrant and X-ray inspection was then made and a sustained pressure check on the system to assure no leaks exist. A pressure drop test assured that no weld burn-through had occurred to restrict coolant flow.

Cabin Structure

Most welding problems centered around the pressurized cabin. The cabin was made up of a spherical bulkhead on the large end and a flat circular bulkhead on the small end. Eighty-five percent of the cabin section, which included equipment bay doors and hatches, was made of welded titanium assemblies.

A bonded honeycomb structure was considered but discarded in favor of the welded assembly. An airtight cabin was required to hold the life-sustaining atmosphere for the astronauts and a welded assembly was ideal from this standpoint. It also provided a relatively smooth surface, even in welded areas, upon which to mount many of the over eight hundred stiffeners, brackets, clips and equipment holders that made up the completed cabin section.

The flat walls of the cabin consisted of two titanium sheets resistance-welded together - 0.010 inch beaded skin on the inside, 0.010 smooth skin on the outside. Brackets and doublers were attached by spot welding and the wall panels were attached to the mating structure by seam welding.

Advanced techniques, improved equipment and higher weld reliability from experience gained on the Mercury project were important design factors.

Large Pressure Bulkhead

Strengthening beads appeared on the outside of' the large pressure bulkhead (the backs of the astronauts' seats were next to this bulkhead), while the skin on the inside was smooth. Both skins were 0.010-inch titanium and similar in construction to the cabin walls except the contour was spherical. A special "birdcage" weld fixture held the contour while skins were being spot welded.

Hatch Sill

Unique problems were presented by the hatch sill. A groove 3/4 inch wide and 1/2 inch deep around each of the two hatch sill openings accommodated the hatch door seals. Part of the cone-shaped cabin was "scooped out" to allow the astronaut to look out the window in the hatch. This presented reverse contours in the "eyebrow" area. A complete "double" hatch sill was made up of eighty-eight welded details and required over seven hundred linear inches of automatic and hand-fusion welding.

Crew Hatches

Hatches for crew access to and from the pressurized cabin area mated with close tolerance to the hatch sill. Fabrication problems were similar to the hatch sill but, in addition to the conical shape and reverse contour of the "scooped out" area, an observation window installation was provided. The window frames were welded assemblies and in turn were welded into the hatch opening. Two hundred eighty-five inches of hand-fusion welding were required to mate the thirteen titanium pieces of each hatch.

Weld Chambers

Several of the titanium cabin subassemblies were fabricated by fusion welding, which consisted mainly of hand or automatic applications of tungsten-inert gas with or without filler wire.

Application of McDonnell-designed welding chambers speeded production. The weld fixture containing the production assembly was put into a weld chamber, chamber air was purged by inert argon gas, and a skilled welder outside the chamber made the fusion weld while peering through windows and holding the welder unit in rubber pressure gloves.

Several sizes of welding chambers were designed by McDonnell to accommodate the various sizes of weld fixtures and production assemblies. Features of these chambers included replaceable flat plastic windows to minimize visual distortion, elimination of the argon cylinder by piping the gas from a central source, and a small positioning fixture that held the assembly at any desired angle.

Some production assemblies presented unusual problems and required special weld chambers. One of these was a special weld chamber made to hold the hatch sills in the proper relationship to the environmental control system (life support and equipment cooling) box while the structural skeleton of the Gemini pressure vessel was welded to them. This tool was affectionately called the "green-house" because of its many plastic windows.

The amount of automatic and hand welding on the cabin assembly alone was about two hundred fifty feet. This did not include spot, stitch or seam welding.

Automatic Fusion Welding

Two automatic welding units with the boom-mounted torch extending over the fixtures moved on rails behind the line of automatic weld fixtures. This gave the TIG welding head three-axis movement. Each weld fixture was supplied with air for clamping and argon gas for backup. These units produce burr welds, angle welds, "T" welds, and corner welds in straight or contoured configuration. One unit made burn-through "T" welds of two pieces and also produced "T" legs at various angles. This fixture was the only one requiring cooling water in addition to copper chill bars for temperature control.

Strain relief fixtures were necessary after most welding operations to prevent warpage of materials involved.

Air Pads

The innovation of air pads on weld fixtures was of great importance in Gemini spacecraft production. Air pads supported a heavy weld fixture, making it possible to move a heavy assembly with the touch of a finger, while an assembly was spot or seam welded in a Sciaky welder. Six Sciaky machines were installed in a line. A smooth aluminum jig plate floor was located in front of each two welders. The air pads replace casters on the weld fixture and. with standard shop air. supported up to 400 pounds per pad. A ten-inch diameter pad was developed at McDonnell to ride 0.003-inch above the jig plate floor.

Fusion Welding Inspection

All welds got a visual (size and shape) inspection. A penetrant inspection was done on non-magnetic materials. One hundred percent radiographic inspection was done with few exceptions. Inspection fixtures were designed to check tolerances at various stages of assembly.

Resistance Welding Inspection

The welding machines were certified for specific material-thickness combinations. Test specimens made prior to production runs simulated the production spotweld and were used for shear and microscopic evaluations. Inspection for nugget penetration was made on spot, stitch and seam welds. When shear test samples were not made on stitch and seam weld, examination was made for minimum nugget diameter. Production welds were penetrant inspected before final acceptance.

Many of the manufacturing aspects of spacecraft were typical of the aircraft industry. However, many were peculiar requirements that require significant refinement of old techniques or the development of new techniques.

Some of the conspicuous special procedures received considerable publicity. These were usually the ones involving assembly and test operations in the "clean rooms". There were other operations that necessitated the improvement of old or the development of new techniques, however.

Corrosion Prevention

Materials used in the spacecraft were either inherently corrosion-resistant or were processed to resist corrosion in the various environmental conditions which could be encountered by the spacecraft. Also, materials which were not encapsulated or contained in hermetically sealed enclosures were either inherently fungus-resistant or were processed to resist fungus attack.

The prevention of corrosion in the magnesium extruded stringers used as a coolant loop in the space radiator posed unique problems. These stringers were procured as extrusions of varying lengths up to 60 feet. Protection started with the handling of the stringers at the supplier through the receipt of the material at McDonnell, its subsequent processing in fabrication both at McDonnell and at a West Coast chem-mill subcontractor, and up to and including its completed installation as a spacecraft plumbing system ready for servicing and functional application. A corrosion preventative compound was applied to provide protection of the material until completion of the assembly of the coolant loop within the adapter. Prior to the baking of the spacecraft adapter exterior paint finish and coupling spacecraft operational components and modules to the coolant loop, the system was purged to remove all traces of the corrosion preventative compound.

Subsequent corrosion protection was provided by pressurizing the closed system with dry nitrogen gas until the spacecraft was serviced with coolant.

Heat Shield Non-Destructive Testing

An extremely important requirement was internal structural integrity of the spacecraft heat shield. The basic configuration consisted of a honeycomb core of approximately one-quarter million cells which were filled with a poured- plastic compound. The completed shield was radiographically evaluated for internal structural soundness. The complete structure was recorded on X-ray film documenting out-of-tolerance conditions, i.e., lack of bond, inclusions, or voids.

The size of the assembly and number of cells to be evaluated dictated the need for a coordination technique that could achieve accurate and consistent correlation between exposed film and the heat shield structure. A 0.020 clear Mylar cover was tailored to heat shield dimensions and contour and provided with locators, which mated with index markings on the heat shield. Horizontal and vertical grid lines were marked on the cover to establish sections. Each section carried a location identity. This identification recorded on the X-ray film during radiographic exposure and become permanent location information. Discrepancies affecting structural soundness were readily visible on exposed X-ray films, but a Polaroid film pack was used to more accurately locate a discrepancy in a given grid section. A series of lead arrows were placed in the proximity of the affected area, the Polaroid film pack was laid up on heat shield and exposed with the radiographic unit. The affected cell could be pinpointed in a series of three to four exposures by reviewing the previous Polaroid "shot", repositioning the lead arrows closer to the defect and reshooting. The defective cell was then marked with a map pin for re-work.

Clean Room Production

Preparation of functional equipment, installation into the spacecraft, and subsequent testing took place in clean rooms. The structural assemblies of the spacecraft sections were completed in an assembly area outside the clean room. Equipment to be installed was tested in the clean room. When the structural assembly of a section was complete, the section was thoroughly cleaned and brought into the clean room, where the installation of equipment began.

Clean room requirements for testing and installation of equipment in the spacecraft stemmed basically from the fluid and gas systems, where small foreign particles or small amounts of corrosion could prevent or degrade the functioning of valves, regulators, pumps, etc., or could cause loss of fuel, oxygen, coolant fluid, pressurizing gas, or other expendables, by preventing the complete sealing of valves.

Specific cleanliness requirements were defined for manufacturing such equipment, and the suppliers had clean rooms with controlled temperature, humidity, and air filtration. Equipment was protected during shipment by capping of openings and sealing in plastic enclosures with necessary dessicants included. Suppliers also provided information tags in the enclosures to define the degree of cleanliness of the part so that no one would inadvertently break the seal in an uncontrolled area.

Depending on the exact level of cleanliness deemed necessary these items were uncovered, tested, and installed in either the McDonnell Class 10 (the most extreme requirement) or the Class 6 Clean Rooms.

The occupants of the Class 6 room were clad in white. Those in coveralls and sneakers were full-time clean room personnel. Those in smocks and disposable booties were various personnel who entered the clean room for only short periods of time.

In the Class 10 Clean Room, personnel wore hoods and gloves which provided even more complete coverage than the clothing used in the Class 6 room.

A third controlled-cleanliness area was maintained for fabrication of electrical subassemblies such as wire harnesses and relay panels.

Quality Control

Spacecraft manufacturing quality control procedures differed from aircraft quality control procedures in degree rather than nature. McDonnell examined the condition of parts more thoroughly and maintained more exacting criteria for acceptance. For instance, no sampling inspection procedures were used for anything except standard nuts, bolts, and fittings. McDonnell tested 100% by X-ray, in 3 views, all transistors procured for the Gemini and required major Gemini electronics equipment suppliers to do the same. These X-ray photographs were examined for foreign inclusions or other detectable internal abnormalities.

In maintaining cleanliness of components and tubing used in the gas or fluid systems, the flushing fluids were inspected until microscopic examination verified satisfactory cleanliness, that is, contaminant particles in the 2-10 micron size range. As a graphic illustration of the degree of refinement this meant, a typewritten period is approximately 500 microns in diameter.

A necessary facet of quality assurance on a program such as Gemini was the amount of record keeping required. Most equipment items down to the level of switches, relays, pyrotechnic cartridges, etc., were controlled by serial number. All serialized items required logs which recorded significant events in their life. Likewise, higher-level equipment necessitated sets of data sheets recording exact results of all tests that they underwent. These data provided a permanent record of the behavior of each item of equipment in the spacecraft. Should last minute concern arise as to the suitability of a particular equipment for an imminent mission, its total manufacturing and performance history could be examined.

Sealing Of The Gemini Spacecraft

While much of the pressure vessel of the Gemini spacecraft was welded titanium, which eliminated leakage, there were a number of instances where construction involved bolts and rivets with their attendant possibilities for loss of cabin pressure. In addition, the egress hatches require sealing as they open directly into the internal pressure vessel.

Each hatch was manually operated by a mechanical latching mechanism from either the inside or the outside of the vessel. The latch forces dictated the use of a soft silicon rubber seal. For proper alignment for the hatch striker and seal in the manufacturing process, the channel was filled with soft putty to determine the striker contact. When the desired alignment was obtained, the putty was removed and the rubber seal was installed in the seal channel. The seal channel frame-to-structure joint was sealed with a room temperature vulcanizing General Electric silicon sealant.

Each of the ingress/egress hatches incorporated a visual observation window consisting of inner and outer glass assemblies. The outer assembly was a single flat pane gasketed on each side with a 0.04-inch Fiberfrax paper. A hollow metal O-ring provided the seal between the periphery of the glass and the frame. The inner window assembly consisted of two flat panes and was sealed with silicon rubber flat gaskets and silicon rubber O-rings.

Other doors which provided access to the equipment compartments within the pressure vessel were sealed by means of a gasket design.

In cases where it was necessary to add a shelf, a piece of equipment, or provided an attachment after major assembly was complete, the area was reinforced by spot welding a stiffening member to the structural wall. A seal problem was created if a hole was necessary under the angle or doubler. Normally the hole in the double wall could be sealed by overlapping spot welds. The hole for a late attachment was sealed by placing a piece of 0.005 silicon coated Fiberglas tape in the hole area under the attaching member. The bolt pressure then sealed off the air entrance to the hole in the wall and prevented pressure loss. Stat-O-Seal washers were used on both sides of a joint of this type to prevent leakage around the bolt.

The transmission of electrical power and signals in and out of the pressure vessel was accomplished by terminating electrical wire bundles at the walls of bulkheads with sealed connectors. The connectors were sealed at the structure surface by an O-ring grooved flange. All connectors were potted with a room temperature vulcanizing silicon sealant compound to provide a moisture and pressure seal and to provide support for the wires at the soldered terminals.

Gemini Thermal Radiation Control Coatings

The Gemini spacecraft required a number of thermal radiation control coatings to reflect and/or re-emit external and internally generated energy when the spacecraft was subjected to heat loads during ascent, orbit or reentry. The Gemini spacecraft utilized super alloys, beryllium, high temperature thermal insulation, ablative heat shields, in addition to the thermal radiation control coatings, to alleviate these high and low temperature extremes. During orbit, the adapter section also served as a space radiator or heat exchanger for the dissipation of internally generated heat; therefore, the exterior surface of that part of the spacecraft had a very low solar absorption and a very high infrared thermal emittance to maintain its desired temperature characteristics.

The internal surface of the adapter module walls required a very low thermal emittance to reduce the heat transfer by radiation between the skin of the adapter and the interior equipment. Also a flexible gold-plated fabric thermocontrol cover prevented escape of heat from the interior of the adapter and kept solar radiation off the internal equipment.

The Rene 41 shingles on the side of the reentry module were heat oxidized to provide a stable adherent, high temperature resistant high emittance surface finish. The beryllium shingles were chemically oxidized to provide a high emittance surface. An air curing silicate bonded black ceramic coating that had a significant degree of toughness at room temperature, good thermal shock resistance, and would withstand temperatures as high as 2300 degrees was used to repair any scratched or damaged areas that could affect the surface of the shingles prior to launch. The inside of the beryllium shingle was coated with a very thin layer of gold to reduce the heat emittance radiated from the shingles to the interior of the rendezvous and recovery section and the reentry control system section.

Proper thermal balance of the spacecraft and its equipment was not complete with only heat rejection features. Proper temperature of the equipment in the adapter section was the result of balancing several conflicting phenomena: namely, radiation from their surfaces to the adapter structure, radiation of heat from direct exposure to solar absorption when the spacecraft was on the daylight side of the Earth, and heat emission from the base of the adapter when the spacecraft was on the dark side of the Earth. The effect of these two phenomena was controlled by coating the interior of the adapter structure with a coating, which was highly reflective at relatively low temperature and covering the open end of the adapter with a surface which was relatively absorptive at the temperatures of solar heat. The interior of the adapter module of GT-3 was coated with aluminum foil tape, which had a silicon pressure resistant adhesive. A gold coating deposited at room temperature with a water base mixture was utilized in later spacecraft due to it ease of application and its reduction in weight.

An intermediate coating of white epoxy was applied to the treated magnesium prior to the application of the gold spray. This coating provided coating protection for the sub-structure and also provided a smooth surface so that the lowest emittance could be obtained with the gold coating. Very few silicon and acrylic bonded coatings were available that would meet the 600 degree requirement for the primary radiator coating. A potassium silicate bonded zinc oxide was selected as the primary space radiator coating on the outside. A silicon-bonded zinc oxide was selected for minor touch-up of areas that would not require the maximum 600-degree ascent temperature resistance and as a primer for the Fiberglas fairings.

The porous Dow 17 Type 1 treated magnesium skins of the adapter to which the coatings were applied were kept clean during assembly by a drillable low adhesive protective paper, which was applied immediately after the treatment. This paper allows rivet patterns to be laid out and drilled with minimum damage or contamination of the clean surface. After assembly, the protective paper was removed and the adapter was cleaned to remove any adhesive. Frayed surfaces were sealed to prevent any entrapped chromate solution resulting from the processing of the assembled parts or resulting from the weld-through sealer used during the welding of the magnesium coolant tubes to the magnesium skins.

After the adapter coating was completed, it was cured at 310 degrees and then scrubbed with a special cleaner and steam cleaned until a water break free surface was obtained. The adapter was then sprayed with alcohol to remove surface water. Masking was accomplished with a low adhesive tape and the Fiberglas fairings were coated with a silicon bonded zinc oxide coating and allowed to dry. The adapter was then coated with a silicon bonded zinc oxide coating in several coats until a total thickness of 4 to 5 mils was achieved. This required several successive coatings with each coat applied immediately after the carrier flashes off. The coating was cured at 310 degrees after several hours of air-drying.

The coatings were selected after being subjected to tests in simulated launch temperatures using furnaces and vacuum chambers and ultraviolet radiation for a period equivalent to two weeks.