Manned spacecraft fabrication techniques developed at McDonnell for the Mercury Spacecraft are being extensively applied in the Gemini program. Weight limitations combined with heat resistance, air load, and acoustic requirements have necessitated improved fabrication practices and have placed even greater demands on assembly techniques.
New welding techniques have been developed at McDonnell to meet many of these demands.
ADAPTER WELDING
The adapter skin is a magnesium-thorium alloy of 0.032-inch thick sheet stock. Welded on the inside is a continuous magnesium-thorium tee-bulb extrusion that serves both as a stiffening stringer and a closed passage for the coolant that converts the adapter to a radiator. Intimate contact between skin and extrusion to facilitate heat transfer is accomplished by using a weld-through sealer for conductivity and as a moisture barrier. Seam welding then mates the parts.
These skins are fabricated in quarter panels of two sizes. Radiator extrusions on each panel are 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 is already seam welded to the adapter skin, a mirror is necessary to make this difficult weld in the confined space.
Penetrant and X-ray inspection is then made and a sustained pressure check on the system to assure no leaks exist. A pressure drop test assures that no weld burn-through has occurred to restrict coolant flow.
CABIN STRUCTURE
Most welding problems center around the pressurized cabin. The cabin is 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 includes equipment bay doors and hatches, is made of welded titanium assemblies.
A bonded honeycomb structure was considered but discarded in favor of the welded assembly. An air-tight cabin is required to hold the life-sustaining atmosphere for the astronauts and a welded assembly is 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 make up the completed cabin section.
The flat walls of the cabin consist 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 are attached by spot welding and the wall panels are 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 appear on the outside of' the large pressure bulkhead (the backs of the astronauts' seats lie next to this bulkhead) while the skin on the inside is smooth. Both skins are 0.010-inch titanium and similar in construction to the cabin walls except the contour is spherical.
A special "birdcage" weld fixture holds the contour while skins are 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 accommodates the hatch door seals.
Part of the cone-shaped cabin is "scooped out" to allow the astronaut to look out the window in the hatch. This presents reverse contours in the "eyebrow" area.
A complete "double" hatch sill is made up of eighty-eight welded details and requires over seven hundred linear inches of automatic and hand-fusion welding.
CREW HATCHES
Hatches, for crew access to and from the pressurized cabin area, mate with close tolerance to the hatch sill. Fabrication problems are similar to the hatch sill but in addition to the conical shape and reverse contour of the "scooped out" area, an observation window installation is provided. The window frames are welded assemblies and in turn are welded into the hatch opening.
Two hundred eighty-five inches of hand-fusion welding are required to mate the thirteen titanium pieces of each hatch.
WELD CHAMBERS
Several of the titanium cabin subassemblies were fabricated by fusion welding which consists mainly of hand or automatic applications of tungsten-inert gas with or without filler wire.
Application of McDonnell-designed welding chambers has speeded production. The weld fixture containing the production assembly is put into a weld chamber, chamber air is purged by inert argon gas, and a skilled welder outside the chamber makes the fusion weld while peering through windows and holding the welder unit in rubber pressure gloves.
Several sizes of welding chambers have been designed by McDonnell to accommodate the various sizes of weld fixtures and production assemblies. Features of these chambers include 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 holds the assembly at any desired angle.
Some production assemblies presented unusual problems and required special weld chambers. One of these is 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 is welded to them. This tool is affectionately called the "green-house" because of its many plastic windows.
The amount of automatic and hand welding on the cabin assembly alone is about two hundred fifty feet. This does not include spot, stitch or seam welding.
AUTOMATIC FUSION WELDING
Two automatic welding units with the boom-mounted torch extending over the fixtures move on rails behind the line of automatic weld fixtures. This gives the T.I.G. welding head three-axis movement. Each weld fixture is 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 makes burn-through "T" welds of two pieces and also produces "T" legs at various angles. This fixture is the only one requiring cooling water in addition to copper chill bars for temperature control.
Strain relief fixtures are necessary after most welding operations to prevent warpage of materials involved.
AIR PADS
The innovation of air pads on weld fixtures has been of great importance in Gemini Spacecraft production. Air pads supporrt a heavy weld fixture making it possible to move a heavy assembly with the touch of a finger, while an assembly is spot or seam welded in a Sciaky welder. Six Sciaky machines are installed in a line. A smooth aluminum jig plate floor is located in front of each two welders. The air pads replace casters on the weld fixture and. with tandard shop air. support up to 400 pounds per pad. A ten-inch diameter pad has been developed at McDonnell to ride 0.003-inch above the jig plate floor.
FUSION WELDING INSPECTION
All welds get a visual (size and shape) inspection. A penetrant inspection is done on non-magnetic materials. One hundred percent radiographic inspection is done with few exceptions. Inspection fixtures were designed to check tolerances at various stages of assembly.
RESISTANCE WELDING INSPECTION
The welding machines are certified for specific material-thickness combinations. Test specimens made prior to production runs simulate the production spotweld and are used for shear and microscopic evaluations. Inspection for nugget penetration is made on spot, stitch and seam welds. When shear test samples are not made on stitch and seam weld, examination is made for minimum nugget diameter.
Production welds are penetrant inspected before final acceptance.
Many of the manufacturing aspects of spacecraft are typical of the aircraft industry. However, many are peculiar requirements which require significant refinement of old techniqes or the development of new techniques.
Some of the conspicuous special procedures have received considerable publicity. These are usually the ones involving assembly and test operations in the "clean rooms". There are other operations which have necessitated the improvement of old or the development of new techniques, however.
CORROSION PREVENTION
Materials used in the spacecraft are either inherently corrosion-resistant or are processed to resist corrosion in the various environmental conditions which may be encountered by the spacecraft. Also, materials which are not encapsulated or contained in hermetically sealed enclosures are either inherently fungus-resistant or are processed to resist fungus attack.
The prevention of corrosion in the magnesium extruded stringers used as a coolant Huid loop in the space radiator posed unique problems. These stringers are procured as extrusions of varying lengths up to 60 feet. Protection starts 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 is 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 is purged to remove all traces of the corrosion preventative compound.
Subsequent corrosion protection is provided by pressurizing the closed system with dry nitrogen gas until the spacecraft is serviced with coolant.
HEAT SHIELD NON-DESTRUCTIVE TESTING
An extremely important requirement is internal structural integrity of the spacecraft heat shield. The basic configuration consists of a honeycomb core of approximately one-quarter million cells which are filled with a poured- plastic compound. The completed shield is radiographically evaluated for internal structural soundness. The complete structure is 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 mate with index markings on the heat shield. Horizontal and vertical grid lines were marked on the cover to establish sections. Each section carries a location identity. This identification records on the X-ray film during radiographic exposure and become permanent location information. Discrepancies affecting structural soundness are readily visible on exposed X-ray films, but a Polaroid film pack is used to more accurately locate a discrepancy in a given grid section. A series of lead arrows are placed in the proximity of the affected area, the Polaroid film pack is laid up on heat shield and exposed with the radiographic unit. The affected cell can 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 is then marked with a map pin for re-work.
Cl.EAN ROOM PRODUCTION
Preparation of functional equipment, installation into the spacecraft, and subsequent testing takes place in clean rooms. The structural assemblies of the spacecraft sections ate completed in an assembly area outside the clean room. Equipment to be installed is tested in the clean room. When the structural assembly of a section is complete. the section is thoroughly cleaned and brought into the clean room. where the installation of equipment begins.
Clean room requirements for testing and installation of equipment in the spacecraft stems 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 have been defined for manufacturing such equipment, and the suppliers have clean rooms with controlled temperature, humidity, and air filtration. Equipment is protected during shipment by capping of openings and sealing in plastic enclosures with necessary dessicants included. Suppliers also provide information tags in the enclosures to define the degree of cleanliness of the part so that no one will inadvertently break the seal in an uncontrolled area.
Depending on the exact level of cleanliness deemed necessary these items are 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 are clad in white. Those in coveralls and sneakers are full-time clean room personnel. Those in smocks and disposable booties are various personnel who enter the clean room for only short periods of time.
In the Class 10 Clean Room, personnel wear hoods and gloves which provide even more complete coverage than the clothing used in the Class 6 room.
A third controlled-cleanliness area is maintained for fabrication of electrical subassemblies such as wire harnesses and relay panels.
QUALITY CONTROL
Spacecraft manufacturing quality control procedures differ from aircraft quality control procedures in degree rather than nature. McDonnell examines the condition of parts more thoroughly and maintains more exacting criteria for acceptance. For instance, no sampling inspection procedures are used for anything except standard nuts, bolts, and fittings. McDonnetl tests 100% by X-ray, in 3 views, all transistors procured for the Gemini and requires major Gemini electronics equipment suppliers to do the same. These X-ray photographs are examined tor foreign inclusions or other detectable internal abnormalities.
In maintaining cleanliness of components and tubing used in the gas or fluid systems, the flushing fluids are inspected until microscopic examination verifies satisfactory cleanliness, that is, contaminant particles in the 2-10 micron size range. As a graphic illustration of the degree of refinement this means, a typewritten period is approximately 500 microns in diameter.
A necessary facet of quality assurance on a program such as Gemini is the amount of record keeping required. Most equipment items down to the level of switches, relays, pyrotechnic cartridges, etc., are controlled by serial number. All serialized items require logs which record significant events in their life. Likewise. higher level equipments necessitate sets of data sheets recording exact results of all tests that they undergo. These data provide 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 manutacturing and performance history can be examined.
SEALING OF THE GEMINI SPACECRAFT
While much of the pressure vessel of the Gemini Spacecraft is welded titanium which eliminates leakage, there are a number of instances where construction involves 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 is manually operated by a mechanical latching mechanism from either the inside or the outside of the vessel. The latch forces dictate the use of a soft silicon rubber seal. For proper alignment for the hatch striker and seal in the manufacturing process, the channel is filled with soft putty to determine the striker contact. When the desired alignment is obtained, the putty is removed and the rubber seal is installed in the seal channel. The seal channel frame-to-structure joint is sealed with a room temperature vulcanizing General Electric silicon sealant.
Each of the ingress/egress hatches incorporates a visual observation window consisting of inner and outer glass assemblies. The outer assembly is a single flat pane gasketed on each side with a 0.04-inch Fiberfrax paper. A hollow metal O-ring provides the seal between the periphery of the glass and the frame. The inner window assembly consists of two flat panes and is sealed with silicon rubber flat gaskets and silicon rubber O-rings.
Other doors which provide access to the equipment compartments within the pressure vessel are sealed by means of a gasket design.
In cases where it is necessary to add a shelf, a piece of equipment, or provide an attachment after major assembly is complete, the area is reinforced by spot welding a stiffening member to the structural wall. A seal problem is created if a hole is necessary under the angle or doubler. Normally the hole in the double wall can be sealed by overlapping spot welds. The hole for a late attachment is sealed by placing a piece of 0.005 silicon coated Fiberglas tape in the hole area under the attaching member. The bolt pressure then seals off the air entrance to the hole in the wall and prevents pressure loss. Stat-O-Seal washers are 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 is accomplished by terminating electrical wire bundles at the walls of bulkheads with sealed connectors. The connectors are sealed at the structure surface by an O-ring grooved flange. All connectors are 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 requires a number of thermal radiation control coatings to reflect and/or re-emit external and internally generated energy when the spacecraft is subjected to heat loads during ascent, orbit or reentry. The Gemini Spacecraft utilizes 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 serves as a space radiator or heat exchanger for the dissipation of internally generated heat; therefore, the exterior surface of that part of the spacecraft has 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 require 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 prevents escape of heat from the interior of the adapter and keeps solar radiation off the internal equipment.
The Rene 41 shingles on the side of the reentry module are heat oxidized to provide a stable adherent, high temperature resistant high emittance surface finish. The beryllium shingles are chemically oxidized to provide a high emittance surface. An air curing silicate bonded black ceramic coating that has a significant degree of toughness at room temperature, good thermal shock resistance, and will withstand temperatures as high as 2300 degrees is used to repair any scratched or damaged areas that may affect the surface of the shingles prior to launch. The inside of the beryllium shingle is coated with a very thing 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 is not complete with only heat rejection features. Proper temperature of the equipment in the adapter section is 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 is on the daylight side of the Earth, and heat emmision from the base of the adapter when the spacecraft is on the dark side of the Earth. The effect of these two phenomena is controlled by coating the interior of the adapter structure with a coating which is highly reflective at relatively low temperature and covering the open end of the adapter with a surface which is relatively absorptive at the temperatures of solar heat. The interior of the adapter module of GT-3 is coated with aluminum foil tape which has a silcon pressure resistant adhesive.
A gold coating deposited at room temperature with a water base mixture is undergoing evaluation and may be utilized in later spacecraft due to it ease of application and its reduction in weight.
An intermediate coating of white epoxy is applied to the treated magnesium prior to the. application of the gold spray. This coating provide coating protection for the sub-structure and also provides a smooth surface so that the lowest emittance can be obtained with the gold coating. Very few silicons and acrylic bonded coatings are available that will 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 are applied are kept clean during assembly by a drillable low adhesive protective paper which is 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 is removed and the adapter is cleaned to remove any adhesive. Frayed surfaces are sealed to prevent any intrapped 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 is completed, it is cured at 310 degrees and then scrubbed with a special cleaner and steam cleaned until a water break free surface is obtained. The adapter is then sprayed with alcohol to remove surface water. Masking is accomplished with a low adhesive tape and the Fiberglas fairings are coated with a silicon bonded zinc oxide coating and allowed to dry. The adapter is then coated with a silicon bonded zinc oxide coating in several coats until a total thickness of 4 to 5 mils is achieved. This requires several sucessive coatings with each coat applied imediately after the carrier flashes off. The coating is cured at 310 degrees after several hours of air drying.
The coatings were selected after being sujected to tests in simulated launch temperatures using furnaces and vacuum chambers and ultraviolet radiation for a period equivalent to two weeks.
Copyright 1997-2005 by John
Duncan |