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 Lateral View of unflown (top) and flown (bottom) SE-6 Gemini Reentry Control Engines. Bottom engine has been fired during flight on Gemini 3 (external nozzle charing is the result of reentry heating).
 Front
 Lateral Obverse View
 Interior Thrust Chamber Assembly
 Inspection Stamps
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SE-6 pressure fed bipropellant engines 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. | ||||
 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  | ||||
 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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