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CURRENT COLLECTION OF APOLLO COMMAND SERVICE MODULE LUNAR PROGRAM SPACEFLIGHT ARTIFACTS
VIEW	ITEM TYPE	FLOWN STATUS	DESCRIPTION
Apollo Command Module Master Events Sequential Controller (MESC) interface connectors to control panel, DC power and subordinate controllers which comprise the Sequential Events Control System (SECS). Apollo Command Module MESC DC fuse boxes. Internal view of the Master Events Sequential Controller showing extensive potting with Dow Corning RTV to inhibit component vibration. Installed location of MESC within Apollo Block II Command Module in Right Hand Equipment Bay	APOLLO COMMAND MODULE SEQUENTIAL EVENTS CONTROL SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE MASTER EVENTS SEQUENTIAL CONTROLLER (MESC) An Apollo Command Module (Block II) Master Events Sequencer Controller (MESC) which was a component of the spacecraft's Sequential Events Control System (SECS) subsystem. The SECS regulated automatically sequenced functions during mission ascent, entry, flight and in the event of an abort. 2 MESC's were installed as part of redundant SECS "A" and "B" systems. The MESC governed activities included Launch Escape Tower (LET) jettison, Service Launch Adapter (SLA) separation, Command Module/Service Module (CM/SM) separation and Earth Landing System (ELS) deployment; and directed Launch Escape System (LES) actuation and Service Propulsion System abort modes. As the Master Controller for the SECS, MECS interfaced with and commanded the following controllers: Emergency Detection System (EDS), SM Jettison Controller (SMJC), Lunar Module Separation Sequence Controller (LMSSC), Command Module Reaction Control System Controller (RCSC), Earth Landing Sequence Controller (ELSC) and the Translation Controller. Relationship of Apollo Command Module Master Events Sequence Controller (MESC) to the Sequential Events Control System subordinate controllers. The MESC and the other SECS controllers received power from the spacecraft Entry and Post Landing batteries (these intern were charged from power provided by the Service Module Fuel Cells). The SMJC, RCSC and ELSC are further detailed as artifacts in this collection in subsequent entries following discussion of the MESC. Part of the CM/SM separation process include the destruction of the CM to SM wiring umbilical. A prerequisite to this action was to "open" certain circuits known to be hot (i.e. active with electrical signals/current flow). To accomplish this, an additional responsibility of the MESC was to trigger pyrotechnically actuated interupters to dead-face these hot circuits in conjunction with CM separation from the SM.
Apollo Command Module Earth Landing System Sequence Controller Label Plate Information Apollo Command Module ELSC installed location (Right Hand Equipment Bay depicted)	APOLLO COMMAND MODULE SEQUENTIAL EVENTS CONTROL SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE EARTH LANDING SEQUENCE CONTROLLER (ELSC) An Apollo Command Module (Block II) Earth Landing Sequence Controller ELSC, manufactured by Northrup Ventura (the same company also developed the rest of the CM Earth Landing System including the 83.5 foot diameter parachutes, an example of which is include the collection and detailed in a subsequent entry on this site). 2 of these units were installed for redundancy in the spacecrafts Forward Equipment Bay. The sequencer, was a subsystem of and employed by the Sequential Events Control System (SECS) to initiate deployment of the Command Modules Apex Cover (forward heat shield), the two 16.5 foot diameter conical-ribbon stabilization Drogues, three 7.2 foot ring-slot Pilot which extracted the three 83.5 foot white and orange stripped Main parachutes (the Drogue and Pilot were mortar deployed). The front interface panel of the ELSC includes a Static Air Port on the lower right of the panel which served as the input for the 24,000 and 10,000 foot baroswitches; their closure armed and subsequently triggered initiation of the 3 mortar deployed pilot parachutes to extract the main canopies.
Apollo Command Module Pyro Continuity Verification Box internal components exposed, heavily potted with RTV. Redundant architecture right/left sides. Apollo Command Module PCVB installed location (Right Hand Equipment Bay depicted)	APOLLO COMMAND MODULE SEQUENTIAL EVENTS CONTROL SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE PYRO CONTINUITY VERIFICATION BOX (PCVB) An Apollo Command Module Earth Landing Pyro Continuity Verification Box (PCVB), operated in series with the Earth Landing Sequence Controller (ELSC), provisioned to monitor the pyrotechnics responsible for jettison of the Apex Cover (forward heatshield), and deployment mortars associated with the Drogue and Pilot parachutes. The box includes a Ground Support Equipment (GSE) access connector for ground checkout of ELS pyrotechnics. This unit was a subsystem of the SECS (Sequential Events Control System). System redundancy was achieved in a single unit (unlike the MESC/ELSC which had two boxes per subsystem).
Apollo Command Module Service Module Jettison Controller (SMJC) Internal View Apollo Command Module SMCJ installed location (Forward bulkhead, Service Module depicted)	APOLLO COMMAND MODULE SEQUENTIAL EVENTS CONTROL SUBSYSTEM	UNFLOWN	APOLLO COMMAND MODULE SERVICE MODULE JETTISON CONTROLLER (SMJC) An Apollo Command Service Module (CSM) Service Module Jettison Controller (SMJC), produced by North American Aviation's Autonetics Division installed as part of the spacecraft's Sequential Events Control System (SECS). The SMJC was responsible for commanding activities associated with detachment of the Apollo Command Module from the Service Module just prior to the CM's reentry into the Earth's atmosphere. Two (redundant) SMJC assemblies were installed in the forward bulkhead of the Service Module and supported by battery power so that they could autonomously perform their function after Command/Service Module separation. Physical separation required severing of all the connections between the modules, transfer of electrical control, and firing of the Service Module Reaction Control System to increase the distance between the CM and SM.
Lateral view of the Apollo Command Module Inverter Apollo Command Module Inverter connectors showing input interface from the spacecraft's Main DC Bus and 3 phase AC outputs to the Command Module load. Inverter Control Box, Motor Switches which enabled flow of power from the DC Main Bus "A" and "B" to the selected Inverter. One each motor switch was apportioned to align Main Bus "A" with Inverter 1, Main Bus "B" with Inverter 2, with the two remaining motor switches interconnected to Inverter 3 so that either bus "A" or "B" could be selected for redundancy. Liberal application of green colored Ladicote, a fire-proofing material introduced into the Block II Command Module design as a result of lessons learned from the Apollo 1 Fire can be seen throughout the artifact. North American Aviation (NAA) I.D. Tag with contract information, part numbers Functional diagram of Apollo Command Module DC Power Distribution with Inverter and Inverter Control Box highlighted	APOLLO COMMAND MODULE ELECTRICAL POWER SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE DC-AC INVERTER AND INVERTER DC CONTROL BOX An Apollo Command Module Electrical Power Subsystem Inverter (manufactured by Westinghouse Electric's Aerospace Electrical Division, Lima, Ohio) and an affiliated Inverter DC Control Box (produced by Kinetics Corporation, Solana Beach California). Three of these 53 pound solid-state inverters and one cotrol box were installed in the lower equipment bay; each inverter converted Fuel Cell originated 28 volt Direct Current (DC) to 115-volt Alternating Current (AC), 3 phase 400 Hertz power which was supplied to spacecraft electronics, environmental control system pumps and space suit compressors. The DC Control Box regulated activation of the inverters via motor actuated switching (on/off) of inputted DC voltage from the Command Module's two DC Main Buses "A" and B". At any give time, two of three inverters were online to ensure uninterupted flow of AC power to the spacecraft. Installed location of the Apollo Command Module Inverters (3) and Inverter DC Control Box within the lower equipmet bay. The inverter is composed of a oscillator, an eight-stage digital countdown section, a dc line filter, two silicon-controlled rectifiers, a magnetic amplifier, a buck-boost amplifier, a demodulator, two dc filters, an eight-stage power inversion section, harmonic neutralization transformer, an ac output filter, current sensing transformers, Zener diode reference bridge, low-voltage control and an overcurrent trip circuit. The inverter normally sourced a 6.4 khz square wave synchronizing signal from the Central Timing Equipment (CTE). DC voltage from the Fuel Cell powerplants (located in bay IV of the Service Module) was connected to the main DC buses "A" or "B"; Inverter No. 1 powered from DC Main Bus "A", No. 2 by bus "B" and Inverter 3 through either DC Main Bus "A" or "B" by switch selection. Each of the inverters had a separate circuit breaker and power control motor switch (contained within the Inverter DC Control Box). Inverter output was routed (via a separate series of control motor switches) to the Command Module AC buses.
Functional relationship of EA within the Apollo Command Module Telecommunications Subsystem. Blocks highlighted in green are representative of artifacts in the collection. The location of artifacts discussed in this and the following 11 entries when installed in their native environment (within the Block II Apollo Command Module). This profile depicts the Lower Equipment bay where the majority of the spacecraft telecommunications subsystem electronics were housed). For reference the Astronauts feet (when laying in the crew couch) are oriented towards the Equipment bay).	Apollo Command Module Telecommunications Subsystem	N/A	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM COMPONENTS The following 11 entries detail a subcollection of Apollo Command Service Module (CSM) Telecommunications Subsystem electronics assemblies (EA) comprising the Radio Frequency (RF) and associated baseband hardware employed on the spacecraft The artifacts are discussed below as individual components; for a depiction of their physical placement and functional relationship (as a system) in the Apollo spacecraft Telecommuncations Subsystem architecture, refer to the diagrams to the left. Where feasible the enclosures have been opened to display internal design. The Apollo Command Module telecommunications subsystem provided voice, television, telemetry and tracking and ranging between the spacecraft and earth, between the CM and Lunar Module (LM), and between the spacecraft and astronauts wearing the portable life support system durig Extra-Vehicular Activity (EVA). It also provided communications amoung the astronauts in the spacecraft and included the central timing equipment for synchronization of other equipment and correlation of telemetry equipment. Most of the components were produced under the direction of Collins Radio Corporation of Cedar Rapids Iowa, with major supporting subcontractor rolls provided by Motorola, RCA, Leach, Radiation Inc and others.
Lateral view of the Apollo Command Module Unified S-Band Transponder Lateral view of the Apollo Command Module Unified S-Band Transponder Top-down view of the Apollo Command Module Unified S-Band Transponder Comparision of Apollo Command Module Unified S-Band Transponder and Lunar Module S-Band Tranceiver (also in this collection and documented on the Lunar Module page) Motorola Corporation News Bureau release photograph of Apollo Command Module Unified S-Band Equipment (USBE) Transponder - the release reads: "The two-way radio on the Apollo Command Module requires less power to communicate with Earth from the vicinity of the Moon then the power used by the lightbulbs in your refrigerator. The small unit, produced by Motorola Government Electronics Division, is the only communications link with the Apollo Command Module crew has with Earth from beyond 30,000 miles away providing all voice contat, TV pictures and mission data. Lovely Motorola technician Mandy Biondi shows the sophisticated unit which has functioned perfectly on every mission." (Image courtesy Motorola/GDAIS).	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE UNIFIED S-BAND TRANSPONDER Apollo Command Module Unified S-Band Transponder (manufactured by Motorola, Inc., Military Electronics Division, Scottsdale, Ariz.). The Unified S-Band Transponder was the only method of exchanging voice communications, tracking, biomedical, and ranging, transmission of pulse code modulated (PCM) data and television, and reception of uplinked data from Mission Control once the Apollo Command Module was outside a range of 1500 nautical miles and line of sight from Manned Space Flight Network (MSFN) ground stations strung around the Earth (within that range, VHF was available). The term "Unified" is applicable because the communications system combined the functions of (signal) acquisition, telemetry, command, voice, television and tracking on one radio link. This design resulted in fewer antennas/electronics assemblies (and thus decreased complexity and weight) on both the spacecraft and the ground station segments of the MSFN. The Unified S-Band Equipment (USBE) onboard the Apollo Command Module, Lunar Module, Lunar Rover were absolutely critical to the successful execution of the Apollo program; and reliabilty was assured through the implementation of full redudant, heavily tested design. The Electronic assembly hosts a redundant architecture consisting of two phase-locked transponders and one frequency modulated transmitter housed in single, gasket-sealed, machined aluminum case, 9.5 by 6 by 21 inches. The unit weighs 32 pounds, operated from 400 Hertz power, with RF output of 300 milliwatts, with a fixed transmit frequency of 2287.5 Megahertz (MHZ) / receive frequency 2106.4 MHZ. Apollo Unified S-Band Transponder electronic assembly with cover removed showing redundant subcomponents. The S-band transponder is a double-superheterodyne phase-lock loop receiver that accepted a phase-modulated radio frequency signal containing the updata and up-voice subcarriers, and a pseudo-random noise code when ranging was desired. This signal is supplied to the receiver via the triplexer integral to the S-band power amplifier equipment and presented to three separate detectors: the narrow- band loop phase detector, the narrow-band coherent amplitude detector, and the wide-band phase detector. In the wide-band phase detector, the intermediate frequency is detected, and the 70-kiloHertz up-data and kilohertz up-voice subcarriers are extracted, amplified, and routed to the up-data and up-voice discriminators in the premodulation processor. When operating in a ranging mode, the pseudo-random noise ranging signal is detected, filtered, and routed to the S-band transmitter as a signal input to the phase modulator. In the loop- phase detector, the intermediate frequency signal is filtered and detected by comparing it with the loop reference frequency. The resulting dc output is used to control the frequency of the voltage-controlled oscillator. The output of the voltage controlled oscillator is used as the reference frequency for receiver circuits as well as for the transmitter. The coherent amplitude detector provided the automatic gain control for receiver sensitivity control. In addition, it detected the amplitude modulation of the carrier introduced by the high-gain antenna system. This detected output was returned to the antenna control system to point the high- gain antenna to the ground station. When the antenna pointed at the ground station, the amplitude modulation was minimized. An additional function of the detector was to select the auxiliary oscillator to provide a stable carrier for the transmitter, whenever the receiver lost lock. The S-band transponders could transmit a phase- modulated signal with the initial transmitter frequency obtained from one of two sources: the voltage controlled oscillator in the phase-locked disband receiver or the auxiliary oscillator in the transmitter. Selection of the excitation was controlled by a coherent amplitude detector. The S-band equipment also contains a separate FM transmitter which permited scientific, television, or playback data to be sent simultaneously to the ground while voice, real-time data, and ranging were being sent via the transponder.
Apollo Command Module Unified S-Band Amplifier electronics assembly with cover removed, displaying redundant architecture including Traveling Wave Tube (TWT) amplifiers and integral triplexer	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE UNIFIED S-BAND AMPLIFIER Apollo Command Module S-Band Power Amplifier (Manufactured by Collins Radio Corp.) This traveling wave- tube (TWT) power amplifier is housed in a sealed and pressurized case 5.75 by 5.56 by 22.26 inches and weighs 32 pounds. It has two independent amplifiers, either of which can amplify outputs of either the phase- modulated or frequency modulated unified S-band transmitters within the S-Band Transponder Assembly (see previous artifact entry for discussion of the S-Band Transponder). It operatee on 3-phase, 1 15/200 volt, 400 Hertz power and supported two power outputs of 2.8 and 11.2 watts. The amplifier increased the low-power output of the unified S-band equipment to high power. Apollo Unified S-Band Amplifier electronic assembly with cover removed showing redundant subcomponents and Triplexer. The S-band power amplifier equipment was used to amplify the radio frequency output from the S-band transmitters when additional signal strength was required for adequate reception by the ground. The amplifier equipment consists of a triplexer, 2 traveling-wave tubes for amplification, power supplies, and the necessary switching relays and control circuitry housed within a single electronics package located in the Apollo Command Module's lower equipment bay. . All received and transmitted S-band signals passed through the triplexer. For the receive leg (Mission Control or the Lunar Module to the Command Module), the S-band carrier uplinked to the spacecraft entered the triplexer from either the s-band Omni-Directional or High Gain antenna equipment. The triplexer then relayed the signal straight through to the S-band receiver. The downlinked signal (CSM to the LM or Mission Control/JSC) from the S-band transponder entered the S-band power amplifier where it was either bypassed directly to the triplexer and out to the S-band antenna equipment, or amplified first and then fed to the triplexer. There ere two power amplifier modes of operation: low power and high power. The high-power mode is automatically chosen for the power amplifier connected to the FM transmitter
Apollo Command Module Telemetry VHF Transmitter Receiver electronics assembly lateral view Apollo Command Module Telemetry VHF Transmitter Receiver electronics assembly production/inspection stamps Apollo Command Module Telemetry VHF Transmitter Receiver electronics assembly block diagram	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE VHF TRANSMITTER / RECEIVER Apollo Command Module VHF/AM Transmitter-Receiver (manufactured by RCA Defense Electronic Products Communications System Division, Camden, N.J.) . Apollo Very High Frequency (VHF) Transmitter Receiver electronic assembly with cover removed showing redundant subcomponents Housed in a single case are dual transmitters and receivers for simplex or duplex operation. The enclosure contains 11 subassemblies, two coaxial relays, and two band pass filters mounted within a three-piece hermetically sealed case installed in the Command Module lower equipment bay. Powered from 28 volts dc, the 5-watt (RF output) unit weighs 13-1/2 pounds and is 6 by 4.7 by 12 inches. The VHF/AM transmitter-receiver equipment provided two-way voice communications among the CM, the ground, the LM, astronauts outside the CM, and recovery forces, relay of two-way voice from either the LM or extravehicular astronauts to the ground (via the S-band); reception of pulse-code modulated data from the LM; and reception of biomedical data from extravehicular astronauts. The equipment includes two independent VHF/ AM transmitters and two independent VHF/AM receivers. Various modes of operation were possible in both the simplex and duplex configurations. These included Simplex "A" (TX/RX on 296.8 MHZ for voice only); Simplex "B" (TX/RX on 259.7 MHZ for Voice only); Duplex "A" (Transmit 296.8 MHZ/Receive 259.7 MHZ voice and biomed data); Duplex "B" (Transmit 259.7/Receive 296.8 voice and ranging); Receive "A" (Receive 296.8 MHZ only); Receive "B" (Receive Lunar Module data 259.7 MHZ); Relay (Interfaces with the Unified S-Band system for relay to the Manned Space Flight Network). The transmitters and receivers operate on different frequencies and one receiver accepts data as well as voice. The receiver circuits are isolated up to the final common output. The VHF/AM transmitter-receiver equipment is controlled by switches on the main display console and push-to-talk buttons. A squelch control varies the level of squelch sensitivity The transmitters and receivers connect with the main display console, the audio center, and the triplexer. The equipment is connected through the triplexer and antenna control switch to either of the VHF Scimitar omni-directional antennas in the service module or the VHF recovery antenna No. 2 in the command module.
Apollo Command Module Telemetry VHF Recovery Beacont electronics assembly internal architecture Apollo Command Module Telemetry VHF Recovery Beacon block diagram	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE VHF RECOVERY BEACON Very High Frequency (VHF) Recovery Beacon manufactured by Collins Radio (Cedar Rapids Iowa). The solid stage 1000 hz square wave tone-modulated AM transmitter was activated automatically upon deployment of the Command Module main parachutes, broadcasting a 243 Mhz carrier frequency transmitted for two second intervals interupted by 3 seconds. The unit weights 2.7 pounds and is 4 x 4 x 6.75 inches, requires a maximum of 10 watts/28 volts DC. The emitted signal provided line-of-site direction finding for recover forces.
Apollo Command Module Telemetry PCM Multiplexer electronics assembly lateral view - test ports Apollo Command Module PCM Telemetry PCM Multiplexer NASA inspection stamps and production information. Apollo Command Module PCM Telemetry PCM Multiplexer assembly serial/part number. Apollo Command Module PCM Telemetry PCM Multiplexer assembly block diagram.	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE PCM TELEMETRY EQUIPMENT Apollo Command Module PuIse-Code Modulation (PCM) Telemetry Equipment (manufactured by Radiation, Inc., Melbourne, Fla.) This unit weighs 42.1 pounds and is 13 by 7 by 14 inches. It received and sampled analog, parallel digital, and serial digital information, which consists of astronaut biomedical, spacecraft operation, and scientific data; and converted it to a single serial output for transmission to earth. This single-output signal was routed to the Premodulation Processor (See following artifact for an example of an the Premod Processor) for transmission to the ground or to Data Storage Equipment (DSE). The pulse-code modulated telemetry equipment was located in the lower equipment bay. Incoming signals were of three general types: high-level analog, parallel digital and serial digital. Two modes of operationwere possible: the high (normal) bit-rate mode of 51.2 kilobits per second and the low- (reduced) bit-rate mode of 1.6 kilobits per second. The analog multiplexer accommodated up to 365 high-level analog inputs in the high-bit rate mode. These analog signals were gated through the multiplexer, the high-speed gates, and then fed into the coder. In the coder, the 0-to-5 volt analog signal was converted to an 8-bit binary digital representation of the sample value. This S-bit word was parallel-transferred into the digital multiplexer where it was combined with 38 external 8-bit digital parallel inputs to form the output in non-return-to zero (NRZ) format. The digital parallel information was transferred into the output register and combined with the digital serial input, and then transferred serially into the data transfer buffer. From here the information was passed on to the premodulation processor for preparation for transmission. The pulse-code modulation telemetry equipment received 512-kiloHertz and 1-Hertz timing signals from the central timing equipment. If this source failed, its programmer used an internal timing reference. The timing source being used was telemetered. Two calibration voltages were also telemetered as a confidence check of the operation of the telemetry equipment.
Apollo Command Module Up-Data Link Equipment electronics assembly lateral view Apollo Command Module Up-Data Link Equipment electronics assembly input/output connectors	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE UP-DATA LINK EQUIPMENT Apollo Command Module Up-Data Link Equipment (Manufactured by Motorola) This 21-pound device is 6 by 18.3 by 9.6 inches. It received, verifies, and distributed digital updating information sent to the spacecraft from the Manned Space Flight Network at various times throughout the mission to update or change the modes of the telecommunications systems. Data was received by the S-band receiver and routed to the up-data link. The up-data link equipment received, verified and distributed digital information sent to the spacecraft by the ground to update or change the status of operational systems. The up-data link consists of detecting and decoding circuitry, a buffer storage unit, output relay drivers, and a power supply and provides the means for the ground to update the computer and the central timing equipment, and to select certain vehicle function. Up-data information was transmitted to the spacecraft as part of the 2-gigaHertz S-band signal. When this signal was received by the unified S-band receiver, the 70-kiloHertz subcarrier containing the up-data information was extracted and sent to the up-data discriminator in the premodulation processor. The resulting composite-audio frequency signal was routed to the sub-bit detector in the up-data link which converted it to a serial digital signal. The digital output from the sub-bit detector was fed to the remaining up-data link circuitry, which checked and stored the digital data, determined its proper destination, and transfered it to the appropriate system or equipment.
Apollo Command Module Pre-Modulation Processor electronics assembly lateral view Apollo Command Module Pre-Modulation Processor electronics assembly production info/inspection stamps. Unfortuately the tag was scraped by North American Aviation employees at the conclusion of the Apollo program but some information remains discernable Apollo Command Module Pre-Modulation Processor electronics assembly functional relationship to Communications Subsystem	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE PREMODULATION PROCESSOR Apollo Command Module Premodulation Processor (manufactured by Collins Radio Corporation) is of solid state design and modular construction and has redundant circuitry. It weighs 14.5 pounds and is 4.7 by 6 by 10.5 inches and required 12.5 watts of power at 28 volts dc. The Premod Processor provided the interface connection between the spacecraft data-gathering equipment and the S-band RF electronics and was essentially the brains of the telecommunications system. It accomplished signal modulation and demodulation, signal mixing, and the proper switching of signals so that the correct intelligence corresponding to a given mode of operation was transmitted.
Apollo Command Module DSE obverse view Apollo Command Module Data Storage Equipment (DSE) with primary cover removed, protective cover installed over the Record and Reproduce head assembly and drive mechanism Apollo Command Module DSE bottom cover removed exposing electronics Apollo Command Module DSE inspection stamps Apollo Command Module DSE residual tag imprint. The tag was removed by Rockwell at the conclusion of the Apollo program in conjunction with scrapping of this artifact. However the information on the tag is gleenable do to the imprint and reads: NAA S&ID PN ME 435-0035-0037 Date of MFR 4/8/69 Collins (or Leach) P/N 514-0051-015 Serial #053338060018 Apollo DSE signal flow showing tape track apportionment. Comparison of Apollo Command Module Block II (top) and Block I Data Storage Equipment.	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE DATA STORAGE EQUIPMENT (DSE) Handling the recording of voice and data aboard the Command Module was a very sophisticated unit - the Data Storage Equipment (DSE). This self-contained device includes two eight-inch reels that spooled through read-write heads approximately 2,250 feet of one-inch Mylar magnetic tape. The 14-track tape had a storage capacity of over four hours of voice and data. Subsystem information, normally sent directly from the spacecraft, was recorded by the DSE along with voice at a high or low bit rate and could then be transmitted to the ground by Mission Control. The DSE was used during the critical Lunar Orbit Insertion (LOI) burn performed by the CSM while on the far side of the Moon when the spacecraft was out of communication with the Earth. During this period, the DSE recorded crew voices, along with important engine and system parameters, that were then dumped to the ground for engineering analysis as soon as the vehicle flew into Earthrise and regained radio communication with Mission Control. The equipment, manufactured by Leach Corporation was installed in the lower equipment bay and had tape speeds of 3.75, 15, and 120 inches per second. The tape speed was selected automatically based on the data rate, The tape has fourteen parallel tracks, four Command Module Pulse Code Modulated (PCM) digital data, one digital clock, one allocated to Lunar Module (LM) PCM data, one CM-LM voice, three scientific data and four spare. Apollo Command Module Block II DSE with drive mechanism and Record/Reproduce head exposed The DSE had several modes of operation, including Single Directional with rewind mode, Automatic Selection (in which the tape speed is determined by the data rate); Remote Control (complete remote operation of the DSE was achieved from Mission Control via the Up-Data Link/S-Band). A total of 2250 feet of tape on the real yielded recording times of 2 hours at 3.75 inches per second and 30 minutes at 15 inches per second.
Apollo Command Module Central Timing Unit (CTE) electronics assembly lateral view Apollo Command Module Central Timing Unit (CTE) block diagram	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE CENTRAL TIMING EQUIPMENT (CTE) Central Timing Equipment maufactured by General Time Corporation. The 10 pound Central Timing Unit Equipment provided precision square wave timing pulses of several frequencies to correlate all time-sensitive functions. It also generated and stored real time day, hour, minute and second of mission elapse and time in binary-coded decimal format for transmission to Mission Control in Houston Texas via the Manned Space Flight Network (MSFN). In primary or normal operation, the Apollo Command Module computer provded a 1,024-kiloHertz synch pulse to the CTE for automatic synchronization with the computer. If the pulse failed, the CTE instantaniously defaulted to the secondary mode of operation using its internally provisioned crystal oscillator. For redundancy, the CTE also contained two power supplies, each fed from different spacecraft power bus/breaker sources. Timing signals generated by the Central Timing Equipment were supplied to the following components on the spacecraft: Pulse Code Modulation (PCM) Equipment (512 Khz square wave for synchronization of the internal clock, 1 hz frame synchronization, and 25-bit parallel time code input for time-coorelation of PCM data); Premodulation Processor (512 Khz square wave for S-Band emergency key transmission); Electrical Power Subsystem Inverters (64 khz square wave for synchronization of 400-cycle AC power); Digital Event Timer (10 hz Pulse for digital clock); Environmental Control System (to discharge water from astronaut suits at the rate of 1 pulse every 10 minutes); Scientific Data Equipment (serial time code output for time-correlation of data)
Apollo Command Module Signal Conditioner electronics assembly test connector, purge valve in view, main cover removed Apollo Command Module Signal Conditioner Equipment (SCE), cover removed Apollo Command Module Signal Conditioner Equipment (SCE), Block Diagram	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE SIGNAL CONDITIONER EQUIPMENT (SCE) Signal Conditioner Equipment (SCE) electronics assembly manufactured by North American Rockwell's (NAR) Autonetics Division. The 45 pound hermetically sealed package resided in the Apollo Command Module's lower equipment bay and drew 28 volts dc/35 watts. The SCE transformed signals from sensors and transducers to basic 0 to 5 volt DC instrumentation analog (coded measurements) voltage level. Signals were then distrubuted to the Pulse Code Modulated (PCM) telemetry and Command Module displays; and provided excitatio voltages to some of the instrumentation sensors and transducers. The SCE contains a DC differential amplifier assembly, DC differential bridge amplifier, AC to DC converter, DC active attenuator, and redundant 20, 10 and 5 volt DC power supplies. Also included is an error detction circuit which automatically switched to the redundant power supply if the primary power supply voltages fell out of tolerance. A switch on spacecraft's main console also permitted manual toggling between either power supply (it was this switch that was toggled by astronaut Alan Bean during the launch of Apollo 12 which prevented an abort and restored telemetry after a lightening strike impacted the launch vehicle disabling spacecraft electronics).
Apollo Command Module Audio Center Equipment (ACE) electronics assembly internal architecture Apollo Command Module Audio Center Equipment (ACE) functional location in spacecraft telecommunications subsystem.	APOLLO COMMAND MODULE TELECOMMUNICATIONS SUBSYSTEM	UNDETERMINED	APOLLO COMMAND MODULE AUDIO CENTER EQUIPMENT (ACE) Audio Center manufactured by Collins Radio Co, Cedar Rapids Iowa. The 7.9 pound unit, measuring 4.7 by 4.56 by 8.65 inches is a 28-volt, 20 watt gasket-sealed box enclosing three identical headset amplifiers, one for each Astronaut audio station on the Apollo Command Module. It enabled communications amoung astronauts, between astronauts and launch personnel, and post-landing recovery frogmen; via the external communications links (Unified S-Band and VHF) to JSC Houston and the Lunar Module, as well as routing of audio to the Data Storage Equipment (DSE) for recording of audio signals. The ACE consists of three electically identical sets of circuitry which provide parallel selection, isolation, gain control and amplification of all voice communications. Each set of circuitry contains isolated pad, diode switch and gain control for each receiver inupt, and intercom channel along with isolation pad and diode switch for each transmitter modulation output and intercom channel; an earphone and microphone amplifier and voice-operated relay (VOX) circuitry with externally controlled sensitivity. The equipment operates with three remote control panels to form three audio statios, each providing an astronaut with independent control of all common functions.
Apollo Command Module RCS SE-8 Obverse View. Apollo Command Module RCS SE-8 Nozzle, ablation/charring evident from firing Apollo Command Module RCS SE-8 Propellant Inlets Apollo Command Module RCS SE 8-2 Data Plate indicating production date, part and contract number. Rocketdyne SE-8 Component Diagram Location of Rocketdyne SE-8 RCS engines on Apollo Command Module Rocketdyne SE 8-2 undergoing static testing during engine development (Image courtesy Rocketdyne)	Apollo Command Module Reaction Control System	Unflown	ROCKETDYNE SE-8 (SE 8-2) REACTION CONTROL SYSTEM THRUST CHAMBER Rocketdyne SE 8-2 Apollo Command Module Reaction Control System (RCS) bi-propellant thrust chamber assembly manufactured second quarter 1964. The Command Module (CM) RCS provided the impulses required for controlling spacecraft rates and attitude during the terminal phases of the mission after separation of the CM from the Service Module (SM) through atmospheric reentry. Attitude control was critical to maintaining correct heatshield orientation while transiting the reentry corridor to splashdown. 12 of these pressure fed engines were installed on the CM and utilized the propellants Nitrogen Tetroxide (N2O4) as oxidizer and Monomethylydrazine (MMH) as fuel at a 2:1 ratio (by weight); pressurized helium gas served as the the propellant transferring agent. This propellant combination is hypergolic so no independent ignition source is necessary. SE-8's were integrated on the CM in 2 (redundant) systems of 6 engines, with individual thrust chambers provisioned to enable clockwise and counterclockwise roll, positive and negative yaw, and positive and negative pitch. The engine has the option of being pulse fired, producing short thrust impulses or continuously fired for steady state thrust level of 93 lbs/414 N with a specific impulse of 274 seconds. Because it is ablatively cooled, the SE 8-2 has a limited operational life (maximum aggregate firing time) of 130 seconds (Pulsing) or 273 seconds (including 200 seconds of steady state operations). The fuel and oxidizer engine injector valves contain orifices which meter the propellant flow at the 2.1:1 ratio. Injector valves were automatically controlled by the CM's controller reaction jet ON-OFF assembly (commanded via the Apollo Guidance Computer) or manually enabled in Direct Manual Mode for rotational maneuvers with the Rotation controller. Direct Manual mode also served as a backup mode to control RCS engine operation in the event ON-OFF mode fails. The SE 8-2 injector valves are spring-loaded closed, energized open and utilize two coaxially wound coils, one each respectively for the manual and direct manual mode. Ignition is approximately 14 millisecond from receipt of electrical command "ON" signal. Cutaway of SE 8-2 Apollo CM RCS Engine - shown with installed variable length nozzle extension not included as part of artifact in this collection. (Image courtesy Rocketdyne). The Injector consists of a 16 unlike doublet splash plate design, containing 16 fuel and 16 oxidizer passages that impinge upon a splash plate within the combustion chamber. The thrust chamber is fabricated in four segments, the combustion chamber ablative sleeve, throat insert, ablative material, asbestos and a fiberglass wrap. The inner section of the engine incorporates a 6 degree wrap ablative sleeve (pre-charred to eliminate the requirement for a crematic liner) with a graphite, zirconium diboride/silicon composite refractory throat insert. A 45 degree nozzle body comprised of high-silica, fabric-reinforced laminated phenolic wrapped ablator is externally covered by a fiber glass overwrap. The combustion chamber casing is a welded stainless steel can. Expansion ratio is 9:1. A conformal nozzle extension (not included on the artifact in this collection) is provided to duct the engine exhaust gases through the spacecraft heatshield (the extension does not significantly alter the engines expansion ratio and is scarfed after installation to match the precise contour of the Apollo Command Module's thermal protection system). A temperature control system incorporating transducers mounted on the engine injector was used to provide crew situational awareness (via a control panel indicator) of the engine temperature status. Prior to firing, if engine temperature was not within specs, the crew had the option of energizing the fuel and oxidizer solenoid control valve direct coils to heat up the engines to the requisite minimum temperature of -10F. Comparison of SE 8-2 with other Rocketdyne SE (Space Engine) series ablative thrusters in this collection (Gemini SE-6 and Saturn APS SE 7-1); artifacts are also documented in independent entries on this site.
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) Imprinted information on Apollo Command Module Main Parachute Canopy, Gore Number 1 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 termination 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. Parachute production at Northrup Ventura. More then 3.5 miles of thread and over two million individual stitches were required by Northrup workers to complete each Apollo main parachute.	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) Diagram of Apollo Command Module Forward Compartment (normally covered by Apex Heat Shield) showing location of Earth Landing System components with installed Main Parachutes. 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 6 strand 9/32 cres steel cable riser (the steel cable riser, rated for a design load of 31,000 pounds, 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 SubSystem	Unflown	APOLLO COMMAND MODULE FLIGHT DIRECTION ATTITUDE INDICATOR An Apollo Command Module Flight Direction Attitude Indicator (FDAI) manufactured by Honeywell 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 FDAI or "8-ball" was one of the most important instruments in the Apollo Command Module (and Lunar Module). Designers had originally intended to give the crew three separate displays to show their attitude; one each for roll, pitch and yaw. Being pilots, the crews quickly threw out the three displays for a development of the artificial horizon familiar from aircraft instrument panels. In an aircraft, the Earth's horizon provides an obvious reference against which the artificial horizon displays attitude. In a spacecraft, such an obvious reference may not exist so the inertial platform at the centre of the IMU (Inertial Measurement Unit) provides one - a reference that is constant relative to the stars, known as an 'inertial' reference. The basic function of the FDAI was to display the spacecraft's attitude with respect to the orientation of this platform. Pitch and yaw can be read off the ball directly; roll is shown by a pointer around the edge of the 8-ball. Three meters around the display (black needles) show the rate of rotation around the three axes. Three additional yellow needles in front of the ball graphically display the difference between the current and desired attitude of the spacecraft. 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".
Apollo Command Module Reaction Jet On-Off Engine Controller (RJ/EC) Internal View. Apollo Command Module RJ/EC label plates displaying contract, serial numbers, modification dates and inspection stamps Apollo Command Module RJ/EC installed location within the CM lower equipment bay	Apollo Command Module Stablization and Control SubSystem	Likely Flown(unconfirmed)	APOLLO COMMAND SERVICE MODULE REACTION JET ON-OFF/ENGINE CONTROLLER (RJ/EC) An Apollo Command Module Reaction Jet and Engine On-Off Control Assembly (RJ/EC) manufactured by Honeywell - part of the CSM Stabilization and Control System, responsible for translating Service Propulsion System (SPS) Main Engine and CM/SM Reaction Control System (RCS) firing commands originated from either the vehicle flight computer or manually from the astronaut (using the Translation Controller) to firing impulses. The RJ/EC accepted the firing commands and outputted 28 volt dc which opened the affiliated SPS/RCS solenoids on each of the engines, flowing propellant and thus firing the main engine and/or R-4Ds (Service Module)/SE-8's (Command Module) as applicable. In effect this box functioned as an electronic "gas pedal". Per the diagram at the left, the unit was installed in the Command Modules lower equipment bay. Functional location of the Apollo Command Service Module Reaction Jet On-Off/Engine Controller (RJ/EC) within the Stabilization and Control Subsystem.
Apollo Command Service Module Thrust Vector Position Servo Amplifier (TVSA) Internal View. Apollo Command Service Module TVSA label plates displaying contract, serial number, and inspection stamps Apollo Command Service Module TVSA installed location within the CM lower equipment bay	Apollo Command Module Stablization and Control SubSystem	Likely Flown(unconfirmed)	APOLLO COMMAND SERVICE MODULE THRUST VECTOR POSITION SERVO AMPLIFIER (TVSA) An Apollo Command Service Service Module Thrust Vector Position Servo Amplifier (TVSA), a component of the spacecraft's Stabilization and Control Subsystem. The TVSA, manufactured by Honeywell Corporation, provided the electrical interface between the command electronics and gimbal actuator for positioning the Service Propulsion System (SPS) main engine. The TVSA received main engine actuator position and deflection rates (velocity) and forwarded that information to the Gimbal Position/Fuel Pressure Indicator on the main display console. In addition, it interpreted Automatic Thrust Vector commands issued by the Command Module Computer as well as manual commands originated by the crew inputs to the Rotation Controller, converting them to a clutch current which resulted in servo-actuator deflection and as a result, gimbaling of the SPS main engine. Functional location of the Apollo Command Service Thrust Vector Position Servo Amplifier (TVSA) within the Stabilization and Control Subsystem.
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 Environmental Control System 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 Environmental Control System	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 Service Module Environmental Control System	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 Service Module Environmental Control System	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 Service Module Environmental Control System	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 Service Module Environmental Control System	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 � 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 Command Module Environmental Control 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 Environmental Control System	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 Command Service Module Service Propulsion System	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 nitrogen� 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 Command Service Module Environmental Control System	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 Command Service Module Service Propulsion System	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 Command Service Module Service Propulsion System 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 Command Module 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. Phill Parker offers an excellent discussion of the thermal environment encounted by the Apollo Command Module and AVCOAT's role in ensuring the safe return of the spacecraft and its crew during the reentry phase of flight. 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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