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 TR-201 Propellant Inlet Ports
 TR-201 Obverse View
 TR-201 Overhead View
 TR-201 Interior Nozzle View - Pintle Injector. To the extreme left of the image can be seen the annular ports which drizzled propellant down the upper thrust chamber walls, supporting film cooling.
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TRW TR-201 Bipropellant Rocket Engine. The thrust chamber was initially developed for the Apollo Lunar Module and was subsequently adopted for the Delta Expendable Launch Vehicle 2nd stage. The engine made 10 flights during the Apollo program and 77 during its Delta career between 1974-1988. This TR-201 has been configured as a fixed thrust version of the Lunar Module Descent Engine (LMDE) for Delta's stage 2. Multi start operation is adjustable up to 55.6 kN and propellant throughput up to 7,711 kg; and the engine can be adapted to optional expansion ratio nozzles. Development of the innovate thrust chamber and pintle design is
credited to TRW Aerospace Engineer Dr. Peter Staudhammer.
The combustion chamber consists of an ablative-lined titanium alloy case to the 16:1 area ratio. Fabrication of the 6A1-4V alloy titanium case was accomplished by machining the chamber portion and the exit cone portion from forgings and welding them into one unit at the throat centerline. The ablative liner is fabricated in two segments and installed from either end. The shape of the nozzle extension (not installed on the example in this collection) is such that the ablative liner is retained in the exit cone during transportation, launch and boost. During engine firing, thrust loads force the exit cone liner against the case. The titanium head end assembly which contains the Pintle Injector and propellant valve subcomponents is attached with thirty-six A-286 steel ¼ inch bolts.
In order to keep the maximum operating temperatures of the titanium case in the vicinity of 800 degrees (F), the ablative liner was designed as a composite material providing the maximum heat sink and minimum weight. The selected configuration consisted of a high density, erosion-resistant silica cloth/phenolic material surrounded by a lightweight needle-felted silica mat/phenolic insulation.
The installed Pintle Injector, unique to TRW designed liquid propulsion systems, provides improved reliability and less costly method of fuel oxidizer impingement in the thrust chamber then conventional coaxial distributed-element injectors typically used on liquid bipropellant rocket engines.
Number flown: 77 (Delta 2000 configuration)
Dry mass: 300 pounds (with Columbian Nozzle Extension Installed)
Length: 51 inches - Gimbal attachment to nozzle tip (minus nozzle extension)
Maximum diameter: 34 inches (minus nozzle extension)
Mounting: gimbal attachment above injector
Engine cycle: pressure fed (15.5 atm reservoir)
Fuel: 50/50 N_2O_4/UDMH at 8.92 kg/s
Oxidizer: monomethyl hydrazine at 5.62 kg/s
O/F ratio: 1.60
Thrust: 42.923 kN vac
Specific impulse: 303 s vacuum
Expansion ratio: 16:1, 43:1 (with Expansion Nozzle)
Cooling method: Film cooled (upper thrust chamber); quartz phenolic chamber ablation (lower thrust chamber) and columbium (niobium) nozzle radiation (Nozzle extension)
Chamber pressure: 7.1 atm
Ignition: hypergolic, started by 28 V electrical signal to on/off solenoid valves
Burn time: 500 s for total of 5 starts; 10 350 s single burn
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 Viking Lander MR-50F Full View
 Obverse View
 40:1 Expansion Nozzle
 Viking Reaction Control System (MR-50F) Thrust Chamber Component Diagram
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A Rocket Research Corporation (RRC) MR-50F monopropellant hydrazine thruster developed for application on the Viking Mars Spacecraft. This 8 pound thrust engine had a rated impulse of 38.69-N and a maximum duty cycle of 20,000 pulses (the thruster is capable of steady state firing for a total of 3,504 seconds). 12 of these engines were installed in groups of 3 on four quadrants of the Viking Lander’s protective Aeroshell Reaction Control System (to provide three-axis attitude control thrust from shortly after the Lander separated from the Orbiter until the aeroshell was jettisoned; as well as the de-orbit impulse required to alter the Lander’s trajectory for the Mars encounter). An additional 4 of these thrusters were installed on the Lander itself as part of the Terminal Descent System (the four MR-50F thrusters enabled roll control on the Lander and augmented the 3 Terminal Descent System main engines).
A description of the engine follows; refer to the diagram at the lower left for graphic presentation of the component locations discussed. Monopropellant hydrazine flows from a trim orifice through the thrust chamber vale and on through an all-welded single feed tube to a rigimesh dispersion element. Hydrazine is distributed from an integrated rigimesh spud through the upper 0.25-inch, 25-30 mesh Shell (Corporation) 405 catalyst bed. An intermediate bed plate with screens welded on each side separates the upper bed from the lower bed, which is 0.65 inches of 14-18 mesh Shell 405 catalyst. A lower bed plate with screen retains the catalyst. Decomposed hydrazine exists through a 40:1 expansion ratio Rao nozzle. A 1/16-inch OD x 0.010-inch wall tube is provided to measure thrust chamber pressure. The propellant valve is solenoid-operated and was manufactured by Parker Aircraft Company and was actuated by 19 watts @ 33 VDC.
The chamber and nozzle assembly consists of a catalyst chamber, nozzle and chamber pressure tap fitting. The chamber-nozzle body is fabricated from Haynes Alloy No. 25 one-piece machining with a cylindrical lower catalyst bed section and a straight 40:1 contoured nozzle (the choice of Haynes alloy was driven by its thermal properties at elevated temperature coupled with its easy machine-ability). The chamber section is designed with a minimum wall thickness of 0.020 inches. The lower end of the cylindrical portion of the chamber contains a step to which a diaphragm assembly is heliarc welded. The nozzle convergent zone is a 50-degree angle conical section blending into the 0.197 inch-diameter nozzle throat. The nozzle divergent section employs a Rao contoured nozzle.
The engine is enclosed in a shield for thermal control. At the aft end of the chamber, a diaphragm is welded between the chamber and heat shield to provide for lateral structural support, differential thermal expansion and for high thermal resistance to limit heat transfer from the chamber to the heat shield. The rocket engine assembly (REA) is mounted from the upper heat-shield flange. Total weight is 1.2 pounds.
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 Viking Orbiter Engine Removed from Transport Case
 Overhead View - Propellant Intake Ports
 Nozzle Interior
 Thrust Chamber/Nozzle Coupling and View of Injector
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Viking Orbiter Main Propulsion/Orbit Insertion engine produced by North American Rockwell /Rocketdyne (manufacturer designation RS-21) under contract to the Jet Propulsion Laboratory (JPL) for the tandem 1975-1980 Viking Orbiter/Landers Robotic exploration missions to Mars. The RS-21 Viking Propulsion system received the further Rocketdyne designation RS-2101c to distinguish it from the RS-2101a (a similar engine developed for the Mariner Mars spacecraft).This bipropellant, REGEN cooled engine is fed by a hypergolic mixture of Nitrogen Tetroxide and Monomethylhydrazine (N204/MMH); use of hypergolics increased engine reliability as the constituent propellants ignited on contact when applied through the injector plate into the thrust chamber, eliminating the requirement for a separate ignition source. The RS-21 is a derivative of the RS-14 (the RS-14 was used on the Minuteman ICBM post-boost phase warhead delivery vehicle). It is capable of producing thrust (in a vacuum) of 136 kgf with a specific impulse (ISP) of 294 sec (approximately 1323 Newtons) translating to a delta-V of 1480 m/s (3310 mph). The engine was a flight ready spare and exhibits evidence of having been test fired. It measures 20 inches in length by 10.5 inches (maximum diameter of Nozzle) independent of its Rocketdyne transport case and includes the Gimbal Ring assembly which supported up to 9 degrees off-axis nozzle positioning for thrust vectoring (spacecraft steering).
The engine was utilized to provide midcourse trajectory corrections while the Viking was enroute to Mars and executed the orbital insertion and orbit trim maneuvers of the Orbiter/Lander spacecraft upon arrival at the red planet. Orbital insertion of Viking 1 required a long engine burn-38 minutes of thrust, which consumed 1063 kilograms of propellant, slowing the spacecraft from its initial approach speed of 14400 kilometers per hour (8948 MPH) to 10400 kilometers per hour (6462 MPH). To bring the spacecraft to the proper point at periapsis (1511 Kilometers/939 miles above the planets surface), the spacecraft was placed in a long, looping 42.6-hour revolution of the planet, reaching first periapsis; orbital apoapsis was ultimately trimmed to 32800 Kilometers (20,381 Miles above the Martian surface).
The primary objectives of the Viking orbiters were to transport the landers to Mars, perform reconnaissance to locate and certify landing sites, act as a communications relay for the companion landers, and to perform their own scientific investigations. The orbiter, loosely based on the earlier Mariner 9  | ||||
 Thrust Chamber Throat
 Propellant Manifold
 Throttle Valve
 Regenerative Feed Line/Fuel Inlet
 Transport Cover
 TD-339 Vernier Engine installed on Surveyor S-10 Engineering Model (National Air & Space Museum)
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A Thiokol Chemical Corporation - Reaction Motors Division (RMD) TD-339 regenerative fuel-cooled Vernier thrust chamber developed for the Surveyor Lunar Lander program. Three of these gold-plated, engines were attached to the “knee” of each of the Surveyor’s three legs; 2 in a “fixed” configuration and the third (on Leg 1) gimbalable +/- 6 degrees about an axis allowing its use for roll control of the spacecraft. The purpose of the Vernier was to provide propulsive power for mid-course correction maneuvers, attitude control before and during the main terminal descent phase from 66 nautical miles altitude above the lunar surface to touchdown, and prime retro power after the main solid-fuel Thiokol retro was jettisoned from the Surveyor spacecraft.
The 100 pound engine utilized hypergolic Monomethyl Hydrazine Hydrate (MMH-H2O) as the fuel and mixed oxides of nitrogen - Nitrogen Tetroxide with 10 percent Nitric Oxide (MON-10) as the oxidizer at a maximum chamber pressure of 250 psia. Assembly consists of a linked upstream throttle valve (each engine was individually throttleable between 30 and 104 pounds), a regeneratively fuel-cooled cylindrical thrust chamber of concentric stainless steel with vortex injector and crematic liner of Zirconium Oxide ( Rokide-Z ), Silicon Carbide insert throat block and a molybdenum nozzle extension (the nozzle extension is absent from this artifact); collectively these engines subassemblies implemented a “Voramic Chamber” concept which was introduced by RMD to mitigate heat transfer issues associated with potential boiling and decomposition of the propellant applied to regeneratively cool the Vernier. A fuel regulator also maintained constant high pressure to inhibit fuel boiling within the chamber cooling jacket. The fuel itself was chosen for its thermal stability so it could operate through the full range of Surveyor’s throttling modes. The gold plating seen on the exterior of the engine is 0.0001 inch (.00025 cm) thick and was applied for rejection of thermal/cosmic radiation from the sun.
This artifact is displayed in its transport case mounting, exactly as it would have been received by the Jet Propulsion Laboratory (JPL) from the manufacture - RMD. It exhibits evidence of having been test fired and has installed desicant plugs for shipment.
The Surveyor project was conceived by JPL scientists primarily as a series of robotic precursor missions to prepare the way for human landings and exploration on the lunar surfaces during the Apollo project. The Vernier Propulsion System (VPS) was one of the most difficult developments undertaken by the Surveyor project . Seven Surveyor missions were conducted between May 1966 and January 1968, two of which (Surveyors 2 and 4) were not successful. A significant achievement occurred in 1967 with the TD-339 enabling the first flight from the Lunar Surface with the lift off of Surveyor 6 as it performed a 2.5 meter “hop” maneuver (the maneuver was executed to observe surface disturbances produced by the initial landing and the effects of firing rocket engines close to the lunar surface – important information required to properly engineer the Apollo Lunar Module Descent Engine).
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 Obverse View
 RP1 and LOX Propellant Manifold
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This modified ROCKETDYNE LR-101 was developed by legandary rocket pioneer Capt Bob Truax and engineer Bill M. Sprague. The new design harvests the originate 4130 steel thrust chamber and propellant valves, transitioning them from a manifold gimbal assembly to a new gimbal design which incorporates actuators and flexible propellant coupling's.
The enhancements permit dual axis gimbaling without reliance on the complex pneumatic systems used in the legacy LR-101 manifold design as deployed on the Atlas and Delta launch vehicle vernier (steering rockets); an example of which may be seen in the next artifact following this entry (Altas LR-101 Vernier).
The TRUAX-SPRAGUE engine, which uses RP-1 (highly refined kerosene) and LOX is a derivative of Capt Truax's efforts to design a commercial man-rated launch vehicle (the X-3 or " VOLKS ROCKET ") capable of parabolic flights to an altitude of 50 nautical miles. Video of engine test firing is viewable on the
TRUAX ENGINEERING MULTIMEDIA ARCHIVES .
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 Anterior view
 Lateral View
 Rocketdyne Label Plate
 Diagram Depicting Significant Assemblage Components
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An LR101-NA Vernier rocket engine assemblage manufactured by ROCKETDYNE Corporation for installation onboard an Atlas SM-65 ICBM (Atlas “A”/XLR 89 NA-1). These engines were employed in various configurations to provide attitude (roll, pitch and yaw) control onboard the Mercury-Atlas, Atlas, Thor ICBM, Delta propulsion systems. A fixed-thrust, single-start, liquid bipropellant engine producing of maximum of 1000 pounds of thrust (nominal seal level), the engine design allows postoperative purging, regenerative cooling, thrust chamber gimbaling, and full-thrust runs of 325 seconds duration. It has a dry weight of 54 pounds and measures approx 28 x 27 ¼ x 20 inches (normal gimbaling arcs included). Designed propellant mixture is combination RP1 (highly refined liquid Kerosene) and LOX (liquid oxygen).
The engine consists of a thrust chamber assembly (a steel double-walled structure with a copper spiral regenerative cooling coils between the inner and outer walls), a pneumatically operated propellant valve with a valve position-indicating switch, an electrically fired igniter assembly, a pneumatically controlled oxidizer bleed valve, a fuel manifold pressure switch, a manifold gimbal assembly, propellant orifices, and pneumatic purge check valves. These components along with interconnecting electrical cabling and tubing assemblies are fixed in position on a welded tubular engine mount.
Gimbaling is facilitated via a pitch gimbal shaft, which provides for movement of the thrust chamber through a pitch-roll correct arc of 70 degrees on either side of the neutral position; and a yaw gimbal shaft which permits movement of the vernier thrust chamber through a yaw correction arc of 30 degrees (outboard) and 20 degrees (inboard) of the neutral position. In addition to performing the thrust direction gimbal function, the yaw shaft serves as a manifold for passage of fuel and oxidizer to the thrust chamber.
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 Front
 Lateral Obverse View
 Interior Thrust Chamber Assembly
 Inspection Stamps
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An SE-6 pressure fed bipropellant engine manufactured by ROCKETDYNE Corporation for installation onboard the Gemini spacecraft Re-Entry Control System (RCS).
These thrusters utilized a hypergolic propellant combination of Nitrogen Tetroxide (NNH) as the Oxidizer and Monomethyl Hydrazine (MMH) as the fuel.
The ablatively cooled pulse-operated engine was rated for a total life span of 96 seconds while providing a vacuum thrust of 23.5 pounds/107 Newtons.
The Thrust Chamber Assembly (TCA) consists of two propellant valves (fuel and oxidizer), injection system, calibrated orifices, combustion chamber and expansion nozzle. The fuel and oxidizer solenoid valves are the normally-closed quick acting type which open simultaneously upon command. The action permits fuel and oxidizer flow into the injector system. The injectors use precise jets to impinge fuel and oxidizer streams on one another for controlled mixing and combustion.
The calibrated orifices are fixed devices used to control propellant flow. Hypergolic ignition occurred in the combustion chamber.
The combustion chamber is non-regenerative; it is lined with ablative materials and insulation to absorb and dissipate heat and control external TCA wall temperature.
When integrated to the Gemini spacecraft, the engine was installed within the RCS module mold line, with the nozzle terminating flush with the outer mold line. A total of (2) sets of eight SE-6 thrusters were paired at approximate radially symmetrical points on the RCS module in a location suitable to execute yaw/roll/pitch maneuvering and attitude control after loss (vehicle separation) of the Equipment Section/Orbit Attitude and Maneuver System (OAMS). The OAMS employed the more potent SE-7 (25, 85 and 100 pound) engines and provided primary attitude control during flight. Firing of the RCS TCA's were
commanded via the Gemini spacecraft's Attitude Control and Maneuver Electronics System (ACME).
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 Nozzle Closure and Throat Insert
 Expansion Cone
 Thiokol Data Tag
 TE-M-385 Motor Schematic
 Location of Retrograde Motors
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A TE-M-385 Star-13E Gemini Retro Rocket nozzle assembly manufactured by Thiokol Elkton Chemical Corp. in the first quarter of 1964 (the company was later known as Morton Thiokol and is currently a subdivision of ATK. It was built as a spare for the Gemini manned space program. This polysulfide propellant engine is a derivative of the Thiokol STAR-13 TE-M-485 orbit insertion motor. Four TE-M-385’s were housed in the Gemini spacecraft’s Retrograde Adapter Module and were provisioned to retard spacecraft orbital velocity for re-entry and to provide distance and velocity to clear the launch vehicle in the event of an abort during ascent from launch. The motors were symmetrically located about the longitudinal axis of the spacecraft and individually, optically aligned prior to mating the adapter to the re-entry module. Firing was initiated using the spacecraft cabin squib arm switches, which applied voltages that ignited the four retrofire rockets via open contacts of either the retro rocket automatic or manual fire relays (located on the Retro Fire Relay panel). The motor was individually rated for an ISP of 211.0 Seconds with a total impulse of 63.16 kNs.In flight configuration, this nozzle would have been mounted to a 13” diameter propellant motor case comprised of titanium alloy. The nozzle assembly, a partially submerged, type, consists of the expansion cone, throat insert and nozzle bulkhead. The nozzle bulkhead is a machined titanium alloy, with 24 bolt holes to facilitate mounting to the motor case flange. The expansion cone is compression molded of vitreous silica phenolic resin and is threaded to the nozzle bulkhead. The throat insert is machined from high density graphite and is pressed into the nozzle bulkhead. The throat insert is insulated from the bulkhead by a plastic material to reduce heat transfer during motor operation and is designed to be recessed into the motor case to reduce nozzle assembly length (when mounted to the titanium propellant case). A rubber nozzle closure is sandwiched between the throat insert and nozzle bulkhead which incorporates a shear grove that permits ejection at a predetermined internal pressure level, or basically at motor ignition. A test adapter fitting is designed into the closure to permit pressure checking of the rocket motor. The Gemini capsule retrograde rocket motors were designed to function in two modes: Normal and Abort. In the normal mode of operation, the rocket motors were used to initiate Gemini spacecraft re-entry. They were fired sequentially at 5.5 second intervals, by hot gases from pyrotechnic igniter assemblies. The Retro rockets could also be salvo fired in conjunction with a launch abort scenario above 15,000 feet to separate the spacecraft from the Titan Launch vehicle. | ||||
 Electrical Receptical and Vent Valve
 Lateral View
 Tag Data
 Diagram of Skylab M518 Canister
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This Skylab flight back-up Gallium Arsenide Crystal Growth Experiment with M518 Furnance, designated experiment M555 , is a flight ready spare for the unit utilized onboard SKYLAB. The canister consists of the Multipurpose Electric Furnace (M518) and a M555 cartridge containing the experiment sealed within the oven. Onboard Skylab, the assembly was inserted into the M512 Materials Processing Facility (located within the Multiple Docking Adapter section of the Spacecraft) .
M555 experimental objectives included the growth of single gallium arsenide crystals in the weightless environment of space from solution in order to produce material of exceptionally high chemical and crystalline perfection, have better doping homogeneity, have more homogeneous starting melts, and achieve uniform growth. M555 was conducted during SKYLAB II (first crew).
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 ISS HATCH - IVA (Inboard) Side
 Capture Mechinism
 Latch Mechinism
 Hatch Installed on Quest Airlock
 Hatch Installed on Unity Node
 HATCH DIAGRAM
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This hatch, manufactured by Boeing Corporation was designed for employment on the International Space Station and supports application in the Quest Joint Airlock, Unity (Node 1), Harmony (Node 2), Node 3, Node X (Node 4), Destiny, Columbus, Kibo (Japanese Experiment Module). The hatch can be opened or closed from either side and includes a pressure interlock feature, which prevents the hatch from being opened if there is a negative differential pressure between opposite sides of the structure.
The hatch measures 53 x 53 inches and mates to a 51 inch square port; and its 4-foot-plus width gives the US subsections of the International Space Station capabilities that are unprecedented in 30 years of orbital operations, going back to NASA's Skylab and Russia's MIR. New capabilities include the ability to transfer large hardware which otherwise would have had to been permanently installed prior to launch (and thus non-removable subsequent to arrival on-orbit). The 4-foot-wide hatch makes it possible to bring in equipment as big as a standard-size refrigerator; without it the ISS would be limited to flying smaller experiments. The use of a smaller hatch design on the MIR precipitated its decommissioning as inoperable/obsolete equipment could not be trasfered via that space stations 31 inch diameter tunnels and hatch.
The large surface area of the hatch requires a designed rating capable of withstanding 20 tons of force from the interior station pressure. An additional design challenge arose from the square shape which results in non-uniform distribution of forces around the edge. To mitigate this, the hatch design includes milled inboard truss bars which provide structural reinforcement.
The Hatch Plate Assembly consists of the following main components: Hatch Plate, Hatch Window Assembly, Manual Pressure Equalization Valve (MPEV), Drive Mechanism Assembly, Hatch Crank Assembly, Roller Assembly, Stowage Assembly, Alignment Guides, Hard & Soft Handles, and Decals / Labels. The Common Hatch assembly in this collection lacks the MPEV and Window Assembly; otherwise it is complete. Reference the hatch diagrams to the lower left of this narrative which illustrates the different Hatch Plate Assembly components and their installation; a detailed description of the individual components is given hereafter.
The Hatch Plate is a rectangular plate with rounded corners made of aluminum alloy. It provides for the installation of all hatch components with the exception of the Hatch Tracks The interior (IVA) side of the Hatch Plate is also referred to as the dome side, due to the smooth, clean, convex curved appearance. The exterior (EVA) side of the hatch plate, referred to as the ribbed side (since providing 24 radial ribs), is distinguished by the latches, latch drive mechanism and tension rods. The Hatch Window Assembly (also referred to as the View Port or Viewing Port) enables the Crew to look inside the space station node with the Hatch closed and consists of two pressure panes, each comprised of two pieces of 0.16 in. thick glass separated by a Dow-Corning 93-500 Silicon layer, leak check ports around each pane, and window rings and seals. One pressure pane can be accessed from the IVA side and the other from the EVA side. The window includes redundant pressure seals. The Manual Pressure Equalization Valve (MPEV) is used to reduce the pressure differential across the Hatch to approximately zero, prior to hatch opening and to allow manual air sampling.
The Drive Mechanism Assembly operates the Latch Mechanisms which secure the Hatch to the Space Station
bulkhead when closed and provide a pre–loading of the pressure seals. The Drive Mechanism Assembly, consists of eight sliders on a rotating ring and the stationary ring. The eight sliders connect to eight-tension rod assemblies, which then connect to the eight
latch assemblies, configured around the exterior perimeter of the Hatch Plate Assembly. The Drive Mechanism Assembly is provided on the ribbed side of the Hatch Plate. Actuation is performed manually by the Crew using the Hatch Crank. The Hatch Crank Assembly actuates the Latch Drive Mechanism, which seats and unseats the hatch on the
ISS bulkhead and pre-loads the Hatch seals. Two are provided, one on the dome side and one on the
ribbed side of the Hatch Plate. The crank can be disengaged from the latch drive mechanism by pulling straight out on the handle. The Hatch is opened by turning the crank handle counter-clockwise on the IVA side and clockwise on the EVA side. A pawl
located on the crank handle allows Crew members to select latch or unlatch. The crank handle engages a ratchet
mechanism to overcome the maximum force required to turn the handle. The crank handle must be stowed prior
to opening the Hatch, because the hatch crank handle has a high profile in the deployed position. Roller Assemblies are bolted to the IVA side of the Hatch Plate. The purpose of the rollers is to guide the Hatch from the closed position, to the full open position, and back to the closed position. Each Roller Assembly contains a roller bearing with a bumper mounted on a shaft. The shaft is clamped in a bracket that is bolted to the Hatch Plate. The bumpers are coated with lubricative, plastic, non-conducting material (called Vespel) which allows the rollers to react to side loads on the tracks and to prevent binding between the roller and the tracks. The two lower rollers have a complete disk of Vespel material on the tip of the roller. In the case of the two upper rollers (or the bonding rollers), the bumper is in the form of a ring with a metallic electrical contact inside the ring that provides contact for the S-class bond. The bearing runs in the hatch tracks to guide the Hatch during opening and closing.
The Stowage Assembly secures the Hatch Plate in the stowed (open) position. It consists of the stowage handle
with a push-button, the cable leading to the latching assembly, the latch housing assembly, the latch pin, and the
manual stowage knob. The stowage latch is operated manually by means of the stowage handle; the latch pin secures the hatch in its stowed position. The Hatch Plate is guided into position for latching by the use of Alignment Guides located at six places around
the perimeter of the Hatch Plate. The Alignment Guides are adjustable to allow smooth operation of the Hatch Plate while maintaining alignment with the Hatch Plate opening and ensuring that the pressure seals are correctly placed with respect to the Hatch Plate. Handles are placed on the Hatch to aid the Crew in operating the Hatch, and in handling the Hatch after it has
been removed during a maintenance operation. There are two types of handles used on the Hatch, hard and soft handles.
The ribbed side of the Hatch is equipped with two rigid, aluminum hard handles. The Crew handle on the dome side of the Hatch is a soft handle. The soft handle has been likened to a suitcase handle in that it is made from NOMEX webbing fastened to the Hatch. This makes the soft handle flexible and less likely to cause interference during Crew operations. Labels are provided on both the interior and exterior side of the Hatch Plate supporting the Crew in all nominal Hatch operations by providing operating and safety instructions. The Hatch Assembly is guided from its stowed position to its closed position by a pair of tracks, right-hand and
left-hand. The axial hatch track installation includes left and right-hand strut assemblies, mounting brackets, and
roller stops. Air leakage at the Hatch sealing surface is minimized by a single–fault tolerant Gask–O–Seal. The seal
comprises four (4) identical quarter sections that are fastened to the Space Station bulkhead. The identical quarter
sections are each made up of two beaded seals which provide, single-fault tolerance.
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 WF/PC 1 WideField (f/12.9) Camera Assembly
 WF/PC 1 Planetary (f/30) Camera Assembly
 Interior Image Showing Cassegrain Mirror Assembly
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Four Optical Camera Channel and Baffle assemblies from the Wide Field and Planetary Camera (WF/PC) 1 recovered from the Hubble Space Telescope during HST Service Mission 1. This collection of artifacts comprises 4 of 8 such assemblies originally onboard WFPC 1, two other assemblies are located on display at the National Air & Space Museum (NASM) in the Explore the Universe exhibit - a link to an online NASM page dedicated to this exhibit may be found HERE; additional photographs of that exhibit are below. The location of the remaining 2 camera’s is undetermined but may currently remain with JPL (original integrator of the WFPC payload).
These 4 camera assemblies, which lack only the affiliated 800 x 800 CCD detectors include two of the original four f/12.9 WideField camera’s (the shorter 13 inch in length tubes) and two of the original four f/30 Planetary camera’s (longer 19.5 inch tubes). Each assembly houses a Cassegrain (folded optics) design including a primary and secondary mirror which focused light on the CCD’s to facilitate the imaging process. The optics remain intact within all four assemblies. It was these cameras which were used for HST’s on-orbit diagnosis of the faulty Hubble primary mirror.
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 Full Frontal View (1)
 Frontal View (2)
 Arm Deployed
 Rear View |
This is the life-test unit/prototype of the Faint Object Camera (FOC) M1 Field Mirror Mechanism that was ultimately installed as part of the COSTAR (Corrective Optics Space Telescope Axial Replacement) payload during Space Shuttle Mission STS-61 (Hubble Service Mission 1) to correct errors in the primary mirror onboard the Hubble Space Telescope. The error was the result of a residual aberation polished into the primary due to a mis-assembled nulling apparatus; the error resulted in the Hubble's primary mirror being ground about 2 micrometers to flat (1/40 the thickness of a human hair). Scientists and engineers devised COSTAR with four small mirrors, about the size of dimes and quarters. The small mirrors were intentionally produced with a flaw identical to and opposite the flaw on the primary Hubble mirror.After installation of the COSTAR, a set of mechanical arms, no longer then a human hand, deployed from the optical bench, placing one of a pair of mirrors (the M1) in front of the opening that admits light into the scientific instruments supporting the FOC and redirected the light to a second mirror (the M2) which properly refocused the light entering the FOC. In the mirror pair, the first corrective mirror passed an image of the Hubble's 94.5 inch-diameter primary mirror onto a second mirror. The second glass was formed like the Hubble's mirror, only in exact reverse, thus canceling the flaw. In a sense, COSTAR put "eyeglasses" in front of FOC (as well as two other scientific payloads - the FOS and GHRS) to correct the telescope's vision, although the eyeglasses were mirrors, not lenses. Conceived by NASA Engineer James H. Crocker and Ball Aerospace Optical expert Murk Bottema, the COSTAR FOC M1 Mirror Mechanism was considered to be one of the most technically challenging aspects of the COSTAR payload to develop because of the complexity affiliated with mirror deployment, tip and tilt. The corrective mirrors were manufactured/hand figured by Tinsley Laboratories. The COSTAR was assembled by Ball Aerospace in Boulder Colorado; the arm is made of Beryllium, the body of Titanium, and the balance of the remaining components constructed of Titanium, Aluminum and other materials. The addition of the COSTAR M1 assembly seen here in concert with the COSTAR M2 corrector dramatically improved FOC's resolution and allowed Hubble to single out individual stars in distant star clusters. COSTAR enabled the FOC instruments to distinguish between objects that were 0.05 arcseconds apart - which is roughly the width of a human hair viewed from a distance of 1 kilometer. If human eyes had this ability, we would be able to distinguish between a pair of automobile headlights 5,000 miles away! The Faint Object Camera was replaced by the Advanced Camera for Surveys (ACS) in March 2002. The photographs to the left depict the arm both in the stowed position (arm not extended) and the deployed position (arm extended). This is the only known working prototype of its kind in the public domain. Its mission complete, COSTAR was anticipated to be retrieved and replaced by an upgraded optics package during the final Hubble Servicing Mission (orginally scheduled for 2004 prior to the loss of STS-107). As part of President Bush's Jan 04 space initiative however, the 5th service mission to HST has been canceled. The cancellation of the plans to retrieve COSTAR for subsequent return to earth make it likely that this artifact is the only remaining M1 FOC Assembly remaining. View an animation of COSTAR ARM DEPLOYMENT (courtesy of Lockheed and NASA). REAL PLAYER Media software is required. |
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Two flown Hubble Space Telescope Solar Cells extracted from a returned array after removal by the crew of STS-61 during Hubble Service Mission One. The lucite encapsulated cells were presented by the European Space Agency (ESA) as a gift to
a select few individuals and VIP's who contributed to the design, production and funding of the ESA manufactured solar
array's. The array's were produced in 1995 and flew onboard HST between 1990-93. Replacement was initiated to correct for thermal degradation of the original arrays and upgrade power production capability (the new array's provided 30% increase in total output).Only one of the replaced array's was returned by STS-61, the other was detached and jettisioned for destruction during re-entry. STS-61 and Hubble Service Mission 1 is most noted for the installation of COSTAR optics which compensated for the sperical aberation on Hubble's primary mirror. See preceding artifacts for a detailed discussion of COSTAR. | ||||
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