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CURRENT COLLECTION OF APOLLO COMMAND SERVICE MODULE LUNAR PROGRAM SPACEFLIGHT ARTIFACTS
VIEW	ITEM TYPE	FLOWN STATUS	DESCRIPTION
Backlit Image of Apollo Command Module Parachute Main Canopy; originating from the top (apex) is the Pilot Parachute riser, routed via the inverted Deployment pack. Pilot Parachute in foreground. Close-up of Apollo Parachute Main canopy vent ports (canopy collapsed in this image) Deployment Pack (after reinversion). This pack housed the complete Apollo Main Parachute assembly under compression in the forward compartment of the Command Module (1 of 3 packs per spacecraft). Deployment Pack - tag data showing Apollo Main Parachute Assembly production data, packing date, NASA Contract Number, Assembly Part Number, Inspection stamps. Drawing showing key components of the Apollo Main Parachute Deployment Pack and Retention Assembly. First and Second Stage Reefing Line holders and pockets which housed the pyrotechnically actuated cutters. The Apollo Command Module Main Parachute was disreefed in two stages, at 6 and 10 seconds after extraction to reduce canopy shock loads. Apollo Main Parachute fabric riser terimation which interconnected with the steel riser leading to the Command Module "Flowerpot". Apollo Main Parachute Deployment bag bridle attachment point to the top of the main canopy. Apollo Command Module Main Parachute Assembly stored in the orignal recovery bag used by the U.S. Navy Divers (Frogmen) for retrieval at the Splashdown location. Apollo 16 Command Module (Casper) approaching splashdown with all 3 Main parachutes fully deployed. White pilot parachute can be seen streaming from the center of the Main canopy; inverted (dark green) Deployment Packs are also visable. Apollo Command Module Earth Landing System parachute deployment sequence.	Apollo Command Module Earth Landing System	Flown Command Module 117 (Skylab 3)	APOLLO COMMAND MODULE MAIN PARACHUTE ASSEMBLY An Apollo Command Module Main Parachute assembly manufactured by Northup Corporation’s Ventura Division under subcontract to North American Rockwell, primary contractor for development and integration of the Apollo Command Service Module (CSM) under NASA contract NAS9-150. This parachute flew onboard CSM 117 for 59 days, 11 hours in support of Skylab 3 (Second Manned Mission - SLM-2 which included crew members - Astronauts Alan Bean, Jack Lousma and Owen Garriott) and as one of three Main parachutes provided deceleration of the Command Module during reentry to ensure safe descent velocity suitable for water landing. The design flew on all Apollo CSM spacecraft as part of the CM Earth Landing System, which assured safe landing for two primary landing modes: landing after a completed mission; and landings by means of the launch escape system (LES) from the time the Apollo crew is in the spacecraft prior to take-off to approximately 300,000 feet after second stage booster ignition. The flown artifact in this collection includes a fully intact, uncut Main 83.5 foot main canopy, 7.2 foot Pilot parachute, Deployment / Retention pack and all riser, suspension lines. The parachute was originally transfered from the Johnson Space Flight Center in 1975 at the conclusion of the Apollo program to the National Air and Space Museum (NASM). NASM determined the chute was in excess of their requirements and designated it for destruction (along with several others). The owner of a private institution, Mr. Cole Palen of Rhinebeck Aerodrome in upstate New York was offered this parachute (and one other unflown example) in lew of destruction. Just prior to his death in December 1993, the parachute was bequeathed to an employee of the Aerodrome, Mr. Chris Rogine. After a subsequent legal challenge by the Aerodrome, Mr Rogine was determined to be the lawful owner and placed this Apollo Main Parachute up for sale at which point it was acquired into my collection. Apollo Main Parachute stowed in its recovery bag. To the top of the image is the Pilot parachute canopy laying on the Main Parachute deployment pack. The Pilot parachute suspension lines, fabric and steel riser (which interconnected with the Deployment pack and extracted the Main parachute), and bridle are also visible. The steel riser was required because of possible contact with the Command Module during deployment. The Main Parachute Deployment pack which housed the hard packed parachute under highly compressed form to a density of .0245 lb/cubic inch (packed under vacuum using specially developed Parachute Packing presses) were installed beneath the Apex Cover in the forward compartment of the Command Module. It was restrained during flight with daisy chain retainers on three sides. The retention system connected directly to the deployment bag without intermittent flaps. The Deployment pack incorporates several layers of Dacron felt for heat protection. Main Parachute Assembly in Deployment Pack installed in forward comparment of Apollo Command Module. (Archival Image) The Pilot parachute is a 7.2 foot diameter , 12 gore Ringslot design permanently attached to the Main parachute deployment bag through a steel riser cable (both the Pilot and Main parachute deployment bag remain permanently attached to the apex of the main parachute). Suspension line is 600 pound nylon cord, 130.4 feet in length and canopy construction is of 1.1 oz per square yard Ripstop nylon. Constructed porosity is 22 percent with a drag area of 24.4 feet squared. A Dacron cover protects a 346.1 inch length fabric riser interconnecting the Pilot parachute suspension line with the 111 inch steel riser. The chute is retained within a mortar tube assembly until deployment and is installed with a sabot placed between the packed chute and the motar charge to support safe ejection. The Main parachute is a 83.5 foot diameter conical Ringsail design actively reefed in two stages. Calculated total porosity is 12 percent. The first stage is actively reefed to 8.4 percent diameter (yielding 285 foot squared drag area) and utilizes redundant reefing lines with two 6.0 second pyrotechnically activated reefing line cutters on each line. The second stage is actively reefed to 24.8 percent diameter (yielding 1080 foot squared drag area) and utilizes two 10.0 second reefing line cutters on a single reefing line. When fully disreefed, the Main parachute Drag area is 4200 foot squared). The Main parachute canopy consists of twelve rings of sails (hence its nomenclature Ring-Sail parachute) with each ring divided into 68 gores and is manufactured with 1.6 oz per yard ripstop nylon (at the crown) and 1.1 oz per yard ripstop nylon (lower gores) The canopy terminates in 68 suspension lines, 1443.9 inches in length which are attached by six 42 inch long steel connector links to six individual legs of a 15.5 inch fabric riser. The six legs of the fabric riser converge into a single leg which connects to the end of a steel cable riser (the steel cable riser is not present on this artifact). The steel cable riser is attached to the Command Module through the parachute attachment and disconnect assembly (informally referred to as the “Flowerpot” because of its appearance). The Main canopy incorporated a two stage reefing system retarded canopy inflation during deployment; this was essential to prevent the destruction of the parachutes and reduce the canopy weight. The reefing system incorporated two reefing lines on the first stage, which each line having its individual set of two pyrotechnically actuated cutters (first stage disfreefing occurred 6 seconds after Main canopy deployment); the second stage had only one reefing line with two cutters pyrotechnically actuated 10 seconds after Main canopy deployment. A steel riser connected the Main parachute to the spacecraft and transmitted parachute loads following entry from terrestrial orbit, lunar flight, or mission abort conditions. The steel riser was interfaced to the Main parachute fabric riser at one end and to the Main Parachute disconnect mechanism (referred to informally as the “Flowerpot” ) on the Command Module; the steel riser is absent from the artifact in this collection. Links are fabricated from Inconel 718, Dacron protective booties cover the fabric riser’s on both the Pilot and Main chutes. Deployment of the Main parachutes commenced at 11,000 after two 16.5 foot diameter ribbon-type mortar deployed drogues stabilized Command Module attitude and provided initial deceleration. Following detachment of the drogues, the Pilot parachute (one affiliated with each Main chute) was mortar deployed simultaneously at 90 degrees to the CM vertical axis, providing the force necessary to release the Main parachute retention system and extract the Main parachute pack assemblies from the Command Module’s upper deck (the parachutes were protected during reentry by a forward heatshield Apex Cover which was jettisoned at 24,000 feet). The deployment sequence was controlled by the fully automatic redundant sequencing system with a manual override mode available as back-up system at the astronauts discretion. As the main parachute packs were pulled away from the command module, the parachutes were extracted from their deployment bags. The Main parachute inflated through the two reefing states to the fully open configuration. Collectively, the three main parachute assemblies decelerated the command module to the final descent velocity of 32 feet per second (two parachute descent rate, i.e. in the event one of the three failed to properly deploy resulted in an off nominal but still very survivable 36 fps descent rate – as demonstrated with the flight failure of one Main chute during Apollo 15’s splashdown). Photograph taken Command Module 117 returning from Skylab 3 during reentry at approximately 10,000 feet ASL (photo taken immediately after extraction of Main parachutes); note that the canopies have not yet been fully disreefed. Parachute in this collection is one of the 3 imaged in the above photograph.
Apollo Command Module FDAI - Lateral View. Apollo Command Module FDAI - power and IMU electrical interface connector cables Apollo Command Module FDAI - Hermetically sealed backplane and support pins for attachment to coldplates. Apollo Command Module FDAI - Johnson Space Center (Manned Space Flight Center Houston) asset and production tags. Apollo FDAI Presentation	Apollo Command Module Stablization and Control System	Unflown	APOLLO COMMAND MODULE FLIGHT DIRECTION ATTITUDE INDICATOR An Apollo Command Module Flight Direction Attitude Indicator (FDAI) manufactured by Sperry Corporation under subcontract to North American Aviation (Primary Contractor for design/build of the Apollo Command Service Module). Two FDAI (for redundancy) were integrated as part of the CSM Stablization and Control System (SCS) to support the crew with display of angular velocity (rate), attitude error, and total attitude. The body rate (roll, yaw, or pitch) displayed on either or both FDAI's was drived from the Command Module Inertial Measurement Unit (IMU) Body Mounted Attitude Gyros (BMAGs). Positive angular rates were indicated by a downward displacement of the pitch rate needle and by leftward displacement of the yaw and roll rate needles. The angular rate displacements were "fly-to" indications as related to rotation control direction of motion requirement to reduce the indicated rates to zero. Servometric meter movements are used for the three rate indicator needles. Specific functions performed by the Apollo Command Module FDAI: (1) Provided an inside-out display of the Command Module attitude with respect to a select inertial frame of reference by means of graduated three-axis ball and appropriate reference indexes. (2) Provided a "fly-to" display of the spacecraft angular position relative to an inertial reference in all three axis by means of attitude-error needles. (3) Provided a fly-to display of the Command Module angular rate about each of the three mutually perpendicular spacecraft body axis before the .05g level during entry (After 0.05g was reached, the displays provided the angular rate about the spacecraft pitch axis and about the roll and yaw axes.) (4) Provided a centralized display of attitudes, vehicular rates, and attitude errors. (5) Provided coarse attitude orientation (obtainable with the ball) and fine orientation using attitude-error needles. (6) Indicated the IMU maneuver limits and the approach of these limits during maneuvers. (7) Provided a monitor and a cross-check of reference-equipment conditions by comparing body-axis rates adjacent to body-axis errors. (8) Provided a monitor of the SCS execution of reference-system commands in three axes by three command needles. (9) Provided variable scale factors in rate and error indications for varying precision of the maneuvers required during the different mission phases. Installed Flight Direction Attitude Indicators onboard the Apollo 16 Command Module "Casper".
Overhead View of Apollo Fuel Cell Group Transport Cover Installed Fuel Cell Simulator Obverse View (1) View of Fuel Cell Stack - Up through Base Overhead View Label Plate Label Plate Electrical and LOX/LOH Connectors	APOLLO COMMAND SERVICE MODULE FUEL CELLS	UNFLOWN	APOLLO COMMAND SERVICE MODULE FUEL CELL POWERPLANT GROUP A group of Apollo Command Service Module Fuel Cell Assemblies manufactured by Pratt & Whitney Aircraft Corp under subcontract for North American Aviation (NAA) arrayed as they would have been installed onboard an Apollo Spacecraft. 3 of these cells were employed to generate primary power and and potable water for the Command Module. This sub-collection of Fuel Cell, is comprised of operational (production representaive units) delivered for the Apollo Command Service module and one Simulator (identified as Serial Number 1 on its label plate), which has been modified to support ground simulation testing; outside the addition of the simulator interface, it is identical to production models employed on the CSM. Each of the assemblies measures 44 inches high, 22 inches in diameter, and weighs 245 pounds; and were designed for installation in Sector (Bay) 4 of the SM. Primarily constructed of titanium, stainless steel, and nickel, the Fuel Cells are rated at 27 to 31 volts under normal loads. There are 31 separate cells in a stack, each producing 1 volt, with potassium hydroxide and water as electrolyte. Each cell consists of a hydrogen and an oxygen electrode, a hydrogen and an oxygen gas compartment and the electrolyte. Each gas reacts independently to produce a flow of electrons. The fuel cells are nonregenerative. They are normally operated at 400 degrees F with limits of 385 and 500 degrees. Water-glycol is used for temperature control. The fuel cells use hydrogen, oxygen, and nitrogen under regulated pressure to produce power and, as a by-product, water. Detailed discussion of functionality is addressed in the following paragraphs. The Bacon-type fuel cell powerplant, was configured in a cluster of 3 systems to comprise the CSM power plant; each cell individually coupled to a heat rejection (radiator) system, the hydrogen and oxygen cryogenic storage systems, a water storage system, and a power distribution system. The powerplants generate dc power on demand through an exothermic chemical reaction. A byproduct of this chemical reaction is water, which is fed to a potable water storage tank in the Command Module (CM) where it is used for astronaut consumption and for cooling purposes in the environmental control subsystem. The amount of water produced is proportional to the ampere-hours. . Apollo Fuel Cell Power Plants installed in Command Service Module Bay Sector 4 The water separation, reactant control, and heat transfer components are mounted in a compact accessory section attached directly above the pressure jacket. Powerplant temperature is controlled by the primary (hydrogen) and secondary (glycol) loops. The hydrogen pump, providing continuous circulation of hydrogen in the primary loop, withdraws water vapor and heat from the stack of cells. The primary bypass valve regulates flow through the hydrogen regenerator to impart exhaust heat to the incoming hydrogen gas as required to maintain the proper cell temperature. The exhaust gas flows to the condenser where waste heat is transferred to the glycol, the resultant temperature decrease liquifying some of the water vapor. The motor-driven centrifugal water separator extracts the liquid and feeds it to the potable water tank in the CM. The temperature of the hydrogen-water vapor exiting from the condenser is controlled by a bypass valve which regulates flow through a secondary regenerator to a control condenser exhaust within desired limits. The cool gas is then pumped back to the fuel cell through the primary regenerator by a motor-driven vane pump, which also compensates for pressure losses due to water extraction and cooling. Waste heat, transferred to the glycol in the condenser, is transported to the radiators located on the fairing between, the CM and SM, where it is radiated into space. Radiator area is sized to reject the waste heat resulting from operation in the normal power range. If an emergency arises in which an extremely low power level is required, individual controls can bypass three of the eight radiator panels for each powerplant. This area reduction improves the margin for radiator freezing which could result from the lack of sufficient waste heat to maintain adequate glycol temperature. This is not a normal procedure and is considered irreversible due to freezing of the bypassed panels. Reactant valves provide the connection between the powerplants and the cryogenic system. They are opened during pre-launch fuel cell startup and closed only after a powerplant malfunction necessitating its isolation from the cryogenic system. Before launch, a valve switch is operated to apply a holding voltage to the open solenoid of the hydrogen and oxygen reactant valves of the three powerplants. This voltage is required only during boost to prevent inadvertent closure due to the effects of high vibration. The reactant valves cannot be closed with this holding voltage applied. After earth orbit insertion, the holding voltage is removed and three circuit breakers are opened to prevent valve closure through inadvertent activation of the reactant valve switches. Nitrogen is stored in each powerplant at 1500 psia and regulated to a pressure of 53 psia. Output of the regulator pressurizes the electrolyte in each cell through a diaphragm arrangement, the coolant loop through an accumulator, and is coupled to the oxygen and hydrogen regulators as a reference pressure. Cryogenic oxygen, supplied to the powerplants at 900 +/- 35 psia, absorbs heat in the lines, absorbs additional heat in the fuel cell powerplant reactant preheater, and reaches the oxygen regulator in a gaseous form at temperatures above 0 degrees F. The differential oxygen regulator reduces pressure to 9.5 psia above the nitrogen reference, thus supplying it to the fuel cell stack at 62.5 psia. Within the porous oxygen electrodes, the oxygen reacts with the water in the electrolyte and the electrons provided by the external circuit to produce hydroxyl ions. Cryogenic hydrogen, supplied to the powerplants at 245 (+15, -20) psia, is heated in the same manner as the oxygen. The differential hydrogen regulator deduces the pressure to 8.5 psia above the reference nitrogen, thus supplying it in a gaseous form to the fuel cells at 61.5 psia. The hydrogen reacts in the porous hydrogen electrodes with the hydroxyl ions in the electrolyte to produce electrons, water vapor, and heat. The nickel electrodes act as a catalyst in the reaction. The water vapor and heat are withdrawn by the circulation of hydrogen gas in the primary loop and the electrons are supplied to the load. Each of the 31 cells contains electrolyte which on initial fill consists of approximately 83 percent potassium hydroxide (KOH) and 17 percent water by weight. The powerplant is initially conditioned to increase the water ratio, and during normal operation, water content will vary between 23 and 28 percent. At this ratio, the electrolyte has a critical temperature of 360 degrees F. Powerplant electrochemical reaction becomes effective at the critical temperature. The powerplants are heated above the critical temperature by ground support equipment. A load on the powerplant of approximately 563 watts is required to maintain it above the normal minimum operating temperature of 385§F. The automatic in-line heater circuit will maintain powerplant temperature in this range with smaller loads applied.
Rear Assembly Close View (1) Lateral View Lateral View	APOLLO COMMAND MODULE OXYGEN CONTROL PANEL	UNFLOWN	APOLLO COMMAND MODULE (BLOCK I) OXYGEN CONTROL PANEL ASSEMBLY Apollo Command Module (CM) Oxygen Control Panel (identified as panel 314 in the APOLLO OPERATIONS HANDBOOK for spacecraft 012), manufactured circa 1967/1968 (derived from date stamps on artifact) which comprise part of the Environmental Control system. This panel, as a component of the oxygen subsystem, controls the flow of oxygen within the CM, provides access to a reserve supply for use during entry and emergencies, regulates the pressure of oxygen supplied to subsystem and pressure suit circuit components, controls cabin pressure, controls pressure in water tanks and the glycol reservoir, and provides for purging the pressure suit circuit. Significant differences between the BLOCK I O2 Control panel in this collection and BLOCK 2 variant include relocation of the Regulator Valve assembly from the primary panel (Block one O2 regulator was onboard the primary panel, and shifted to a separate subpanel for the Block II spacecraft) and the provisioning of a Portable Life Support System (PLSS) port (Block II 02 panel had no PLSS port). Inspection Stamps
Filter Obverse Lateral View (1) Data Plate	APOLLO COMMAND MODULE ENVIRONMENTAL CONTROL	UNFLOWN	APOLLO COMMAND MODULE CO2 ABSORBER ELEMENT An expendable Command Module CO2 absorber element produced by AIRESEARCH; used in conjunction with the CM Environmental Control Unit to regulate particulates and carbon dioxide within the cabin. The unit weighs 8 pounds and measures 7x7x5 inches. CO2 absorber elements of Lithium Hydroxide (LiOH) and activated Carbon were used to remove trace contaminants and metabolic carbon dioxide; each element was sized for 1.5 man-days of operation at the designed metabolic loads, and the elements were changed by the crew alternately every 12 hours. 20 elements were carried for 8-10 mission days initially; this was increased to 30 for the extend "J" missions (Apollo’s 15-17). A similar CO2 absorber element was employed onboard the Lunar Module but was sized/shaped differently (the CM used this "Mailbox" scrubber; the LM employed a cylindrically shaped canister). During the Apollo 13 Mission, as a result of an explosion in the Command Service Module Liquid Oxygen tank which provided the crew and Fuel Cells with oxygen, the Lunar Module was needed as a "life boat" but was only designed to accomidate 2 crew members; the requirement to support the third crewmember and affiliated increased metabolic loading rapidly expended the LM CO2 absorber elements. The Command Module CO2 absorber elements were needed in the LM. Working with Mission Control in Houston, the astronauts fabricated a novel adapter from existing material onboard the spacecraft (plastic bags, platic-coated cue cards, spacesuit hoses and lots of duct tape) allowing the Command Module's square canisters to be interfaced to the Lunar Module Environmental Control System (ECS), preventing a dangerous buildup of carbon dioxide. Command Module CO2 Absorber Element Component Diagram
Operator Interface Lateral View/NAA Label Plate Rear View CSM (Block I) Panel 201	APOLLO COMMAND MODULE WASTE DISPOSAL SELECTOR VALVE	UNFLOWN	APOLLO COMMAND MODULE (BLOCK I) WASTE DISPOSAL SELECTOR VALVE An Apollo Command Service Module (Block I) Waste Disposal Selector Valve assembly manufactured circa 1965 under Nasa Apollo Contract NAS9-150 by Accessory Products Company (APCO). Part of the Environmental Control System (ECS) Waste Management System / Waste Management-Selector, this valve was used to route human (liquid and solid) bio-waste products originating from a Fecal Canister, Urine Receptacle, Urine Volume Sampling Measuring Systems Unit (UVSMS) and Vacumn Cleaner assembly to overboard discharge. The Waste Management System (including this Valve) was redesigned for the Block II CSM. Diagram at left shows location of valve assembly handle on Panel 201 of the Block I CSM.
Lateral View Inlet/Outlet Ports Tag Data	APOLLO COMMAND MODULE ENVIRONMENTAL CONTROL	UNDETERMINED	APOLLO COMMAND MODULE (BLOCK II) PRIMARY WATER-GLYCOL COOLANT PUMP Apollo Command Module primary water-glycol coolant pump manufactured in 1969 by Airesearch under subcontract to North American Aviation / Rockwell (CSM integrator under NASA contract NAS9-150 ). Two of these centrifugal pumps (primary/backup) were installed as part of the Environmental Control System to circulate 200 lb/hr of water/ethylene glycol coolant through the heat absorption and rejection equipment in the CSM. Typically throughout the course of a mission, only one pump would be utilized (in continous mode) with the secondary pump being reserved in the event of primary unit/cooling loop failure.
Valve Assembly Lateral View 1 Lateral View 2 Tag Data CSM BLOCK II PANEL 303 Command Module Cabin Primary Temperature Control Valve Flow Diagram	APOLLO COMMAND MODULE PRIMARY TEMP CONTROL ASSEMBLY	UNDETERMINED	APOLLO COMMAND MODULE (BLOCK II) PRIMARY TEMPERATURE CONTROL A Block II Apollo Command Service Module Environmental Control System (ECS) Primary Cabin Temperature Flow Control Valve and Manual Crew Selector Knob assembly manufactured circa 1968 by AIRESEARCH under CSM contract NAS 9-150. This assembly regulated the primary glycol loop in the Apollo Command Module and was normally positioned automatically by the cabin temperature control or manually by means of an override control on the face of the valve. The valve was located on the CSM panel 303 (a diagram of the panel is viewable to the left of this narrative) The motor-operated valve is manually controlled by the back-up mode control knob. Rotational movement from H to C is approximately ½ turn with the dual valve on a single shaft permitting water-glycol flow to route to the heat exchanger. Rotation toward H (heat) position results in proportional increase in cabin temperature by directing warm water-glycol to the cabin heat-exchanger. Rotation towards C (cool) position results in proportional decrease in cabin temperature by directing cool water-glycol to the cabin heat exchanger.
Matched Ground and Service Module Oxygen Vent Line Couplers Mated Ground and Service Module Oxygen Vent Couplers Service Module Oxygen Fill Coupler Data Tag Ground Hydrogen Vent Coupler Stamped Data	APOLLO SERVICE MODULE FUEL CELL CRYOGENIC SUB-SYSTEM	UNFLOWN	APOLLO SERVICE MODULE CRYOGENIC HYDROGEN AND OXYGEN COUPLERS This sub-collection is comprised of Cryogenic couplers installed onboard the Service Module and affilated ground umbilicles which supplied liquid Oxygen and Hydrogen to the Apollo Fuel Cell storage tanks. The Service Module coupling devices were manufactured by Beech Aircraft; the ground couplers were subcontracted by North American Aviation to Stratos-Western (Fairchild Hiller Corporation) under NASA Contract NAS9-150. Cryogenic Hydrogen and Oxygen comprised the constituant reactants used by the (3) onboard Bacon Fuel cells to provide power, drinking water and heating (an example of a fuel cell is also documented within this collection in a preceeding entry). The Oxygen tanks also supplied metabolic breathing oxygen to the crew Cryogenic Fill and Vent Port Locations
O2 Regulator Assembly Top View O2 Regulator Assembly Bottom View	APOLLO COMMAND MODULE OXYGEN REGULATOR ASSEMBLY	Unflown	APOLLO COMMAND MODULE OXYGEN REGULATOR An Apollo Command Module Oxygen (O2) Redundant Main Regulator; date of production assessed as 1973 which likely affiliates this component as a flight ready spare for the Apollo-Skylab Command Modules - Spacecrafts AS116 through AS118 or the ATSP/Skylab Rescue vehicle (AS119). The main regulator reduces the O2 supply pressure to 100 + 10 psig for use by subsystem components. The regulator assembly is a dual unit which is normally operated in parallel. Selector valves at the inlet to the assembly provide a means of isolating either of the units in case of failure, or for shutting them both off. Integral relief valves limit the downstream pressure to 140 psig maximum. APOLLO 14 Oxygen Control Panel - switch tabs (toggled to the up or "OPEN" position) for Oxygen Regulator Assembly can be seen immediately to left of "Oxygen Control Panel" label* (image courtesy of Ray Holt)
AEROJET Tag Assembly Lateral View Assembly Detail View Apollo SPS Main Engine Functional Flow Diagram	APOLLO SERVICE PROPULSION SYSTEM BIPROPELLANT VALVE SUBSYSTEM	Likely Flight Ready Spare (Unflown)	APOLLO SERVICE PROPULSION SYSTEM PNEUMATIC CONTROL ASSEMBLY An Apollo Service Propulsion System (SPS) Primary Bipropellant Valve Pneumatic Control Assembly used to regulate the introduction of fuel and oxidizer into the Command Service Modules main engine. Manufactured by General Tire's AEROJET Liquid Rocket Operations Division (NASA Prime Contractor for the Apollo program Service Module propulsion system). Both a primary and secondary assembly were installed ontop the Apollo SPS thrust chamber; each assembly consisting of one each nitrogen pressure vessel, injector prevalve, nitrogen regulator, nitrogen relief valve, two solenoid control valve, and two actuators. Determination that this is the primary assembly is based on Regulator labeling (tag carries the "A" designation) and the 2 Solenoid Control valves which carry the "-1" and the "-2" designation respectively stamped on the exterior of each valve. A functional diagram of the SPS MAIN ENGINE depicting the relationship of the Bipropellant valve assembly may be viewed by clicking here or on the image to the lower left; a detailed description of this assembly follows below. Nitrogen tanks are mounted on the bipropellant valve assembly to supply pressure to the injector prevalves. One tank is in the primary pneumatic control system (A) and the other tank is in the secondary pneumatic control system (B). The tanks each contain 5.8 cubic inches of nitrogenÄ enough to operate the valves 43 times with an initial nominal pressure of 2500 psi. The injector prevalves are two-position, solenoid operated valves, one for each pneumatic control system, and identified as A and B. The valve is energized open and spring-loaded closed. The prevalves, when energized open, allow nitrogen supply tank pressure to be directed through the regulator into a relief valve and to a pair of solenoid control valves. The single-stage regulator is installed in each pneumatic control system between the injector prevalves and the solenoid control valves. The regulator reduces the nitrogen pressure to 190-230 psi. The pressure relief valve is located downstream from the regulator to limit the pressure applied to the solenoid control valves in case a regulator malfunctions. The orifice between the injector prevalve and regulator restricts the flow of nitrogen and allows the relief valve to relieve the pressure overboard in the event the regulator malfunctions, preventing damage to the solenoid control valves and actuators. Four solenoid-operated, three-way, two-position control valves (from both the primary and secondary systems) are used for actuator control. Two solenoid control valves are located in each pneumatic control system. The solenoid control valves in the primary system are identified as l and 2 and the two in the secondary system are identified as 3 and 4. The solenoid control valves in the primary system control actuator and ball valves 1 and 2. The two solenoid control valves in the secondary system control actuator and ball valves 3 and 4. Four piston-type, pneumatically operated actuators control the eight propellant ball valves. Each actuator piston is mechanically connected to a pair of propellant ball valves, one fuel and one oxidizer. When the solenoid control valves are opened, pneumatic pressure is applied to the opening side of the actuators. The spring pressure on the closing side is overcome and the actuator piston moves. Utilizing a rack and pinion gear, linear motion of the actuator connecting arm is converted into rotary motion, which opens the propellant ball valves. When the engine firing signal is removed from the solenoid control valves, the solenoid control valves close, removing the pneumatic pressure source from the opening side of the actuators. The actuator spring pressure then forces the actuator piston to move in the opposite direction, causing the propellant ball valves to close. The piston movement forces the remaining nitrogen on the opening side of the actuator back through the solenoid control valves where it is vented overboard. Each actuator contains a pair of linear position transducers. One supplies information on the position of the ball valve to the main display console and the other information to telemetry.The eight propellant ball valves are used to distribute fuel and oxidizer to the engine injector assembly.
Top View Label Plate Lateral View - O2 Relief Assembly Base View System Tank Valve Module Component Diagram	APOLLO OXYGEN/HYDROGEN SERVICE MODULE	Flight Ready Spare (Unflown)	APOLLO OXYGEN/HYDROGEN SERVICE MODULE A System (Tank) Valve Module (Apollo Oxygen Hydrogen Service Module) manufactured by PARKER-HANNIFIN circa 1967-1971 used to provide overboard relief of elevated pressures within the stored LH2 (Liquid Hydrogen) and LO2 (Liquid Oxygen) tanks in the Apollo Command Service Module (CSM). This system also sensed LH2/LO2 tank pressures, regulated the tank heaters and controlled the tank destratification motors (motors used to execute the CRYO-STIR in the LOX/LH2 tanks for "slush" mitigation). The modules were in-line between the LH2/LO2 cryogenic tanks and the CSM Fuel Cells. The Systems (Tank) Valve Modules for the hydrogen and oxygen systems are functionally identical. Each module contrains two relief valves, two pressure transducers, two pressure switches and one check valve. The relief valves are a differential type designed to be unaffected by back pressure in the downstream plumbing, The valve has temperature compensation and a self-aligning valve seat. Relief crack pressure is 273 psig minimum for hydrogen tanks and 983 psig minimum for oxygen tanks. Full flow pressures (LO2) is 1010 psig (maximum) and for LH2 285 psig (maximum). Module pressure switches are a double pole single throw absolute device. A positive reference pressure (less then atmospheric) is used to trim the mechanical trip mechanism to obtain the absolute switch actuation settings. Reference pressure is typically between 4 to 10 psia. The motor driven switch controls power to both the tank heaters and destratification motors. The pressure transducers, are an absolute (vacuum reference) device. Each transducer consists of a silicon pickup comprised of four sensors mounted on a damped edge diaphragm and an integral signal conditioner. The unit senses tank pressure through the discharge line from the tank; signal conditioner output is 0-5 VDC analog output which is linearly proportioned to tank pressure. The check valve, is designed to open at a differential pressure of approximately 1 psia. The single poppet is spring loaded and has a large area to prevent chattering during flow in the normal direction. This large area also helps in obtaining a positive seal if pressurized in the reverse direction.
RCS Helium Check Valve (Obverse)	APOLLO CM/SM HELIUM CHECK VALVE	Unflown	APOLLO CM/SM RCS CHECK VALVE A Reaction Control System (RCS) Helium Check Valve assembly manufactured by Accessory Products Corporation (APCO) for the Apollo Command Module and Service Module (Service Propulsion System) Reaction Control System (RCS) Propellant Pressurization Section under NASA contract NAS9-150. Produced circa 1964, two check valve assemblies, one for oxidizer and one for fuel, permit helium flow to the tanks and prevent propellant or propellant vapor flow into the presurization system if seepage or failure occurs in the propellant tank bladders.
He Regulator(Obverse) He Regulator(Top) Valve Schematic Service Module RCS System Diagram	APOLLO SPS HELIUM REGULATOR	Unflown	APOLLO SERVICE MODULE RCS HELIUM REGULATOR Reaction Control System (RCS) Helium Pressure Regulator assembly manufactured by Fairchild Stratos Corp (Western Division) for the Apollo Service Module (Service Propulsion System), Reaction Control System pressurization system under NASA contract NAS9-150 for North American Aviation (NAA). The dual, series configured regulators (produced 1965) received gas from the SPS helium tank isolation valve and reduced ambient tank pressure of 4150 PSIA to approx 220 PSIA (nominal) before forwarding to the downstream RCS check valve (see RCS Check Valve description which precedes this entry). Each assembly incorporates two (primary and secondary) regulators connected in series. If the primary regulator fails open, the secondary regulator will maintain slightly higher but acceptable pressures. Diagram to the left depicts functional relationship/location of regulator in the RCS system
He Relief Valve Label Data	APOLLO SERVICE PROPULSION SYSTEM HE RELIEF VALVE	Unflown	APOLLO SERVICE PROPULSION SYSTEM HELIUM RELIEF VALVE Helium Pressure Relief Valve manufactured by CALMEC Manufacturing Corp for the Apollo Service Propulsion System (SPS) pressurization subsystem under subcontract to NAA (NASA contract NAS9-150). The pressure relief valve (one of two which would have been integrated into the SPS) consists of a relief valve, a diaphragm, and a filter. In the event that excessive helium or propellant vapor ruptured the diaphragm, the relief valve opened and vented the system. The relief valve would close and reseal after the excessive pressure had returned to the operating level. The diaphragm provided a more positive seal of helium than a relief valve. The filter prevents any fragments from the diaphragm from entering onto the relief valve seat. The relief valve opens at a pressure of 212 psi after the diaphragm ruptures at about 213 psi. The valve will close when pressure drops to 208 psi. Regulated helium was utilized by the SPS to provide positive pressure to the fuel and oxidizer storage tanks and drive the propellant fluids into the Apollo Command Service Module's main engine.
Complete view Side view Detailed view of QA Stamps Post Flight sample of Apollo 4 Heat Shield AVCOAT 5026-39HCG recovered by North American Aviation	Apollo Thermal Protection System	UnFlown (Avcoat 5026 Block)/ Flown SC017 Specimen	APOLLO COMMAND SERVICE MODULE AVCOAT 5026 HEAT SHIELD BLOCK AND FLOWN SPECIMEN An AVCOAT 5026-39HCG HeatShield block (1 square foot/12x12inch; 2 inch thick) manufactured in 1966 as a component of the Command Service Module (CSM) Thermal Protection System. Also included for comparison is a flown example recovered from SpaceCraft 017 (Apollo 4 Unmanned Test Flight). AVCOAT was produced by AVCO corporation, with integration into the Command Module's Thermal Protection System (TPS) by NAA (North American Aviation). The block is comprised of an extremely lightweight fiberglass honeycomb filled with an epoxy-novalac resin reinforced with quartz fibers and phenolic microballons. Command Module installation required application of the material into a honeycomb matrix (comprised of 3/8 inch diameter individual cells) that has been previously vacuum bonded to a stainless-steel substructure which in turn was braze welded to the Heat Shield subassemblies. The AVCOAT 5026-39HCG compound was applied to the individual honeycomb cells with a dielectrically heated (160 degree F.) hypodermic device similar to a caulking gun (the operation was termed "gunning"). After completion of filling, the subassemblies were vacuum bagged and the ablator oven cured for 16 hours at 200 degrees F, machined on a turret lathe to the design-thickness and x-rayed to detect any defects. Thickness of the TPS ablator ranged from .7 to 2.7 inches and varied with the projected local environment on the spacecraft's Aft, Crew Compartment and Forward headshields (a function of the re-entering Command Module's attitude with respect to the atmosphere). The unflown example in this collection has 1,024 individual handfilled cells; the completed CSM heatshield had 370,000 cells. The ablative charactoristic (i.e. the material essentially melts and chars away during re-entry) mitigates the 20,000 degree heat experienced by the CSM as it transits through the extreme thermo-dynamic phase of the re-entry cooridor. Images at the left show the cellular construct of the heat block. This technology was also selected in April 2009 to protect NASA's Orion Crew Module, the follow-on spacecraft for the U.S. Space Shuttle. Orion, being developed under Project Constellation is eventually planned to facilitate U.S. return to the Lunar Surface (in combination with a new lunar lander (Altar) and launch vehicle family (Ares 1 and Ares V). NASA's formal announcement selecting TEXTRON (who acquired AVCO in 1986) to provide Orion's Thermal Protection System incorporating AVCOAT may be accessed HERE and discussion of the test/evaluation process HERE. Bisected specimen of Command Module 017 (Apollo 4) Thermal Protection System (Avcoat 5026-39HCG) displaying significant ablation post-reentry and recovery.

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