CURRENT COLLECTION OF APOLLO LUNAR MODULE AND SATURN V SPACEFLIGHT ARTIFACTS
Saturn S-IVB APS Attitude Control Engine (TR-204) Quad Redundant Solenoid Valves
Saturn V S-IVB APS - Diagram Reflects origin of 28 Volt DC Command Firing Impulse which opened the propellant solenoids and initiated flow of hypergolic propellants into the injectors of the TR-204 and SE 7-1 APS engines.
Saturn S-IVB APS TR-204 Attitude Control Engine Exit Nozzle with its Ablative Exit Cone Liner and Moly Throat Insert.
Saturn V S-IVB APS Attitude Control Engine - Looking through the Moly Throat Insert to the Injector with its 12 fuel and oxidizer ports arrayed in an "Unlike on Unlike" Doublet configuration
Diagram of S-IVB Auxiliary Propulsion System - location of TR-204 Attitude Control Engines highlighted.
||Flight Spare (Unflown)
SATURN V THIRD STAGE AUXILLARY PROPULSION ENGINE / TR-204
Thompson Ramo Wooldridge (TRW) Inc. Rocket Engine developed for application as the Apollo Saturn V Launch Vehicle S-IVB (Third Stage) Auxiliary Propulsion System (APS) Attitude Control Engine.
Three 150 pound thrust (670 Newton) TRW, Inc. TR-204 engines were employed in each of the two attitude control modules affixed to the S-IVB stage (employed as both the second stage on the Saturn IB and the third stage of the Saturn V launch systems). Hypergolic bi-propellants (Nitrogen Tetroxide combined with Monomethylhydrazine) fed at an oxidizer to fuel ratio of 1.60 was supplied to the engine under positive pressurization via independant helium and propellant storage tanks housed within each APS module. The APS provided roll control during J-2 engine powered phases of flight, pitch, yaw, and roll (attitude) control during coast periods; maneuvering impulse during undocking and extraction of the Lunar Module, and also propellant settling (ullage). Firing control for the TR-204 was initiated by the Instrument Unit Flight Control Computer. As a final act of glory, the APS was continuously fired until depletion to propel the spent S-IVB stage to lunar impact to produce seismic energy for detection by sensors emplaced on the moon's surface during preceding Apollo missions.
The TR-204 combustion chamber is lined with ablatively refrasil material. The engines have quadruple propellant injector valves for redundancy. The thrust chamber assembly (TCA) is integrally fabricated and composed of three major elements: the combustion chamber, the nozzle throat section and the nozzle expansion cone with an expansion ration of 33.9 to 1. The engine was qualified for a total pulse operation of 300 seconds. During the Saturn program, this engine achieved an overall reliability design goal of 0.992 at a 90-percent confidence level while operating.
The engine propellant valve assembly consists of eight normally closed, quad redundant propellant valves (four oxidizer and four fuel), arranged in two series parallel arrangements. Dual failure within the manifold fuel or oxidizer arrangement was required to cause failure of the rocket engine assembly. An assembly closed failure would have prevented any operation while an assembly open failure would have resulted in continuous flow and loss of all propellant. Assembly valve failure cannot occur unless two valves fail open in series or two valved failed closed in parallel. This high degree of reliability was nessessary to prevent the possibility of a mission abort do to loss of spacecraft/launch vehicle attitude control prior to separation of the Saturn third stage and the Apollo Command Service Module (CSM). The inlet ot the valve package contains a 100 mesh screen (150 microns) to protec the engine from large particulate conamination.
The injector valves provided positive on-off control of propellant flow upon command from an external power source (28 volts DC). Four valves, integral in an assembly were capable of simultaneous operation and were synchronized to allow or terminate propellant flow within 3 milliseconds of each other. The opening time for each valve assembly, defined as the time from initiation of open signal to fully open valve package did not exceed 23 milliseconds.
The injector consists of twelve pairs of unlike-on-unlike doublets arranged to minimize hot spots in the combustion chamber. The valve side of the injector is filled with a silver braze heat sink which reduces injector operating temperature. During the 300-second life requirement, the external wall temperature was designed not to exceed 600 degrees F and the maximum valve body external temperature did not exceed 165 degrees F.
Saturn V S-IVB (Third Stage) Auxiliary Propulsion System (APS) TR-204 Attitude Control Engine Schematic.
Saturn V S-IVB Auxiliary Propulsion System (APS) Attitude Control (TR-204) and Ullage (SE 7-1) imaged together for context (relative size, orientation within the APS and application). The SE 7-1 is addressed more comprehensively below following discussion of the TR-204.
Rocketdyne SE 7-1 (View 2)
SE 7-1 Propellant Ports and Solenoid Electrical Connectors
Thrust Chamber Data Plate
SE 7-1 Ullage Engine Component Diagram
Diagram of S-IVB Auxiliary Propulsion System - location of Ullage Engine highlighted.
||Flight Spare (Unflown)
SATURN V THIRD STAGE ULLAGE ENGINE - ROCKETDYNE SE 7-1
Rocketdyne SE 7-1 Rocket Engine developed for application as the Saturn-IVB (Third Stage) Auxiliary Propulsion System (APS) Ullage Engine. (Ullage is the free space within a tank above the liquid propellant. Derived from the term 'ullage' in winemaking, where it refers to the space above the liquid in a container such as a barrel or wine bottle - within the weightless environment of space during intervals when the Saturn Third stage J-2 main engine is not active, liquid propellant hydrogen and oxygen (LH2/LOX) may float around in the empty tank volume and inadvertently get sucked into the main engine resulting in the undesired effect of cavitation. The ignition of the Ullage engine prior to main engine restart induced propellant settling within the tank to inhibit this cavitation). Two of these engines, rated for a specific impulse of 274 Seconds (Vacuum) / 72 pounds thrust each, were installed on the S-IVB to support stage restart capability. The ablatively cooled, pressure fed thrust chamber ran hypergolic propellants Monomethylhydrazine (MMH) as fuel/Nitrogen Tetroxide (NTO) as oxidizer and has a total burn life of 425 seconds.
The APS Ullage engines were used to facilitate propellant settling after completion of the first J-2 burn on the Saturn third stage and during restart chill-down immediately prior to the second J-2 burn which placed the Apollo CSM/LM in Translunar Injection. They were also used on several Apollo missions for ground commanded, guided lunar impact trajectory burns of the S-IVB/IU spent stage after separation from the Command Service Module and extraction of the Lunar Module (S-IVB lunar surface impact was desired to generate seismic data in conjuction with an experiment supporting compositional analysis of the Moon's interior. Data was collected via seismometers left by earlier Apollo crews).
This SE 7-1 is a direct derivative of the (SE 7) 100 pound thrust engine used on the Gemini Orbital Attitude and Maneuver System (OAMS). Differences between the Gemini engine and the APS SE 7-1 include a reduction in operating chamber pressure (from 150 to 100 psia); rated thrust (from 95 to 72 lbs); and the propellant inlet fittings (from tube stub to dual redundant right angle fittings for adapataion of the engine to the Saturn APS).
The thrust chamber body is made in two segments; the combustion zone segment and the nozzle segement. The combustion zone segment is fabricated from a 6-degree oriented (referenced to engine centerline), resin-impregnated, high-silica fiber cloth. In addition, the thrust chamber body is wrapped with a layer of phenolic-bonded asbestos fiber to provide increased heat resistance and sealing capabilities. The bond line between the combustion chamber segment and the nozzle segment is located in a low-pressure, low-stress area aft of the throat insert. Structural support for the thrust chamber body assembly is provided by alternate layers of high-temperature high-strength glass cloth and filament-wound glass roving, bonded by phenolic resin. Additional layers of glass roving provide added strength in the injector attached and throat areas. The thrust chamber body is encased in a stainless steel shell to provide a positive seal between the thrust chamber and the launch vehicle. The engine combustion chamber contains a one-piece JTA graphite liner. A throat insert of solid silica carbide is used to resist the erosive effects of the combustion gases.
The thrust chamber injector is fabricated from stainless steel. It consists of 16 pieces of unlike doublets which impinge on a splash plate providing propellant mixing for high combustion efficiency.
Engine operation is controlled by two fast-acting electrically-operated solenoid propellant valves. These are attached to a mounting bracket which in turn is attached to the injector plate. The basic propellant valve design embodies a hermetically sealed solenoid. Valve sealing is accomplished through the use of a precision ground ball, attached to the armature, which rests on a Teflon seat in the closed position. A metal stop below the Teflon seat is incorporated to limit the armature stroke. Closing is accomplished through the use of a spring, and sealing force is obtained from the spring and pressure of propellant acting on the ball.
SE 7-1 Thrust Chamber Ablative Liner; Silicon Carbide Insert also visible
LVDC Memory Modules and Page Assembly Cards (View 2)
Memory Module Assembly Chip Set (Base)
Memory Module Assembly Chip Set (Side)
LVDC Page Assembly Side "A"
LVDC Page Assembly Side "B" (opposite side of Page Assembly card shown above)
Page Assembly top and Frame Assembly interface connector/pins
Launch Vehicle Digital Computer Component Diagram
Depiction of Memory Module Ferrite Core Matrix Subsection
||SATURN V and IB
||Flight Spare (Unflown)
SATURN V LAUNCH VEHICLE DIGITAL COMPUTER - MEMORY MODULES AND PAGE ASSEMBLY CARDS
Launch Vehicle Digital Computer (LVDC) Memory Module Assembly and Page Assembly Cards produced by International Business Machines (IBM) Corporation, Federal Systems Division, Rockville Maryland under NASA contract number NAS 8-11561. The LVDC was installed within the Saturn IB and Saturn V Instrument Unit (IU) to support prelaunch checkout; navigation, guidance and attitude control; flight sequence control and orbital checkout of vehicle systems. It represented the brains of the SATURN flight control system and for its time, the �state-of-the-art� in computational technology. Refer to a diagram of the LVDC to the left for a depiction of where the Page Assembly and Memory Modules were installed within the computer�s Magnesium-Lithium Chassis.
Logic Page assembly (cards), shown above and to the left, were fabricated from two Multilayer Interconnection Boards bonded back-to-back ( each side labeled "A" and "B" , and comprised the Logic section of the LVDC. Semiconductor chips are mounted on square ceramic wafers (side length 7.5mm) on which interconnecting wiring and film resistors have been depositied by silk screen printing and firing. The devices, called Unit Logic Devices, are soldered to multi-layer interconnection boards. Each multilayer interconnection board has a capacity of 35 Unit Logic Devices. Two multilayer interconnection boards are bonded back-to-back to a supporting metal frame to form a logic page assembly. Multilayer interconnection boards and pages are joined by connectors to a central multilayer printed circuit board.
The Memory Modules are self-contained assemblies with memory timing, drive, inhibit and sensing circuits arranged around the core array. They provided 4,096 word locations (28 bits each) of primary storage in each of up to eight memory modules for 32,789 words (4 memory modules are shown in this collection) and could be operated in either a simplex or duplex mode, as determined by the Memory Control Elements. Simplex operations took full advantage of the available capacity by storing different information in each mode. Duplex operation halved the capacity of the memory but increased its reliability through redundant storage. Each Memory Module operated independently on command from the Memory Control Element. The modules are divided into sixteen sectors, one of which is designated the �Residual� sector. Each sector contained 256 locations, or addresses. Sector selection was identical for each module and address selection was the same for every sector, except the Residual sector. Storage external to the memory is located predominantly in glass delay lines.
The Memory Module assemblies employed Toroidal (donut shaped) ferrite cores as the storage medium, utilizing coincident current addressing and destructive readout techniques. The hand-woven ferrite cores are arrayed in fourteen 128 x 64 magnetic core planes (refer to diagram to the left of this narrative) and can be magnetized in either of two directions. By establishing that a core �contained� a �1� when magnetized in one direction and a �0� when magnetized in the opposite direction, the core can be used to store a single �bit� of a binary number. Core magnetization was achieved by passing a direct current through the X and Y drive lines (copper wire conductors running through each of the cores comprising the magnetic core planes).
The LVDC employed the first computer application and architecture in which all critical circuits were triplicated (triple modulear redundancy), giving near-ultimate operative reliabiity. Designers selected seven functional sections where catastrophic failure might occur but, for reasons of reliability, could not be permitted to occur in flight. Each selected section was then placed in three identical but independant logic channels. Problems were presented to each module simultaniously, and the results of each, independently derived, went to a majority-rule voter circuit. Any dissenting "vote" was discarded as an error, and the only signal passed along by the voter circuit consisted of the identical signals from two of the modules.
Exposed 128 x 64 Bit Core Memory Plane (14 stacked plane's comprise the data storage resources for a single memory module) displaying 2,048 individual tiny donut magnets and the interleaved "X" and "Y" drive and sense lines - it was onto these magnets (via the drive lines) that the actual Saturn V Launch Vehicle program was recorded and stored for subsequent access and "destructive" readout (via the sense lines) by the computer.
Launch Vehicle Digital Computer With Top Cover Removed and Exposed Page Assembly Cards/Logic Subsystem (Courtesy U.S. Space and Rocket Center, Huntsville Alabama)
|Full Frontal View (1)
Panel Right Front
Panel Center Front
Panel Left Front
Lunar Module Panel Layout (LM 6 Typical) Panel 3 Highlighted In Red
||APOLLO LUNAR MODULE (LM)COMMANDERS & S.E. LOWER MAIN FLIGHT PANEL
||Probable Flight Ready Spare (Not Flown)
LUNAR MODULE CONTROL PANEL 3 - APOLLO 9 COMPLETE LUNAR MODULE COMMANDERS & S.E. LOWER MAIN FLIGHT PANEL ASSEMBLY
Below image of the panel after being reenergized with 400hz/115V AC replicated LM power source to commemorate 40th Anniversery of Apollo 11 Landing on the Moon. (click on picture for a higher resolution image
Energized Panel (Note Illuminated Blue Lunar Contact Light)
A Lunar Module Lower Main Panel Assembly which would provided the crew with radar control and indication, desent engine override, interior/exterior lighting, touch down indication (via the blue CONTACT LIGHT) and a number of other operator interfaces with the spacecraft. Based on a review of Grumman Drawings and chassis serial numbers, this panel has been confirmed to be an operational spare (OS) constructed for LM-3 Spider (Apollo 9) .
Panel 5 Forward
Panel 5 Backplane
Panel 8 Forward
Panel 8 Backplane
Panel 12 Forward
Panel 12 Backplane
||Test Articles (Unflown)
LUNAR MODULE CONTROL PANELS 5, 8 and 12
Apollo Lunar Module Panels 5, 8 and 12 manufactured by Grumman Corporation for use on the lunar lander under NASA Contract NAS9-1100. These panels, which were test articles, display post Apollo 1 accident modifications to reduce the risk of fire, including new fire-retardant potting (reddish/brown material on the back of the panels where the cables intersect the backplane), Dow Corning Beta Cloth wrapping and "booties" covering the backplane wiring and connectors, and wire bundle ties, and clamps.
Panel 5 contains lighting and mission timer controls, engine start and stop pushbuttons, and an X-translation pushbutton. The Desecent Rate Switch (DES RATE) permited establishing the LM rate of descent under the Primary Guidance Navigation System (PGNS) control in fixed increments. Each switch actuation provided a discrete pulse, changing the rate of descent by 1 foot per second. The Engine STOP pushbutton provided descrete stop signals to the Descent and Ascent engines, independent of the Engine Arm switch, except when in abort sequence. The Engine START pushbutton (installed immediately below the Engine Stop) permitted immediate manual firing of the Descent or Ascent engine, depending upon setting of the Engine Arm switch. The TIMER CONTROL Switch provided descrete signals to the Mission Timer indicator and the Slew Control provided slewing functions for setting the time in Minutes and Seconds respectively (Flight panels also included an Hour setting toggle). The OVERRIDE ANUN switch provided full voltage bypass of Caution and Warning array and the component caution light portion of the ANUM/NUM control.
At the left of the Commander's station is panel 8, which is canted up 15 degrees from the horizontal. This panel contains controls and displays for explosive devices and descent propulsion, and audio controls.
At the right of the LM Pilot's station is panel 12, which is canted up 150 from the horizontal. This panel contains audio, communications, and communications antenna controls and displays.
Lunar Module Panel Layout (LM 6 Typical) with Panels 5, 8 and 12 Highlighted in Red
LM Valve Package Assembly outboard view.
Lunar Module Valve Package Assembly top view showing propellant ducting which leads to the RS-18 Injector (not present on this artifact).
Lunar Module VPA propellant ducts. The fuel duct inspection and acceptance stamps are visible.
Lunar Module Propellant Valve Solenoid (Actuator) closeup (one of 4 installed on the VPA).
Bell Aerospace Label Plate
Ascent Engine Propellant System Schematic
||Lunar Module Ascent Engine Component
LUNAR MODULE ASCENT ENGINE VALVE PACKAGE ASSEMBLY
An Apollo Lunar Module Ascent Engine Valve Package Assembly (VPA) produced by Bell Aerospace Corporation under NASA contract NAS9-1100 for Grumman (Prime for the LM). This assembly regulated propellant flow into the engine injectors and combustion chamber of the engine; it had to function perfectly the first and only time it was employed or the crew would have perished on the surface of the moon.
The VPA accepted hypergolic fuel (hydrazine (N2H4) mixed with unsymmetrical dimethylhydrazine (UDMH), commercially known as Aerozine 50, and oxidizer (nitrogen tetroxide - N2O4) from the Ascent stage propellant storage tanks; supplying it to the propellant feed section/engine assembly interface, the oxidizer and fuel lines lead into the valve package assembly. The individual valves that make up the valve package assembly are in a series-parallel arrangement to provide redundant propellant flow paths and shutoff capability. The valve package assembly consists of eight propellant shutoff valves and four solenoid-operated pilot valve and actuator assemblies. valve assembly consists of one fuel shutoff valve and one oxidizer shutoff valve These are ball valves that are operated by a common shaft. which is connected to its respective pilot valve and actuator assembly. Shaft seals and vented cavities prevent the propellants from coming into contact with each other. The VPA was also connected to separate vent manifold assemblies which drained the fuel and oxidizer that leaks past the valve seals, and the actuation fluid (fuel in the actuators when the pilot valves close), overboard. The eight shutoff valves open simultaneously to permit propellant flow to the engine while it is operating; they close simultaneously to terminate propellant flow at engine shutdown. The four non-latching, solenoid-operated pilot valves control the actuation fluid (fuel).
Notice the fuel ducting appears larger, bulboelous then its companion oxidizer line. During testing of the Ascent engine it was noteced that pre-ignition pressure spikes were occuring, this was resolved by lengthing the path the fuel had to take so that timing of its arrival at the injector was delayed to occur until slightly after the oxidizer (the ducting wraps back on itself to extend the path length).
Ascent Engine Diagram depicting physical location of Valve Package Assembly - (extracted from the GRUMMAN APOLLO NEWS REFERENCE)
A Lunar Module RS-18 Ascent Engine with its installed Valve Package Assembly (VPA). The VPA is nearly identical to the artifact in this collection; the grey mylar sheath (absent from the artifact) was installed durinng finally assembly of the engine.
Engine Interface Assembly
||Lunar Module Ascent Engine Components
LUNAR MODULE ASCENT ENGINE PROPELLANT SYSTEM
Apollo Lunar Module Ascent Engine Propellant System tubing, interfaces and filters produced under subcontract to GAEC (Grumman Aircraft and Engineering Corporation)
for NASA Contract NAS9-1100. Components are as follows
1 - ORIFICE SUB ASSEMBLY, HYPERGOLIC FUEL; Operating Pressure 250 PSIG, Manufacturer STAINLESS STEEL PRODUCTS (Burbank Ca.)
2 - ENGINE INTERFACE SUB ASSEMBLY, OXIDIZER; Operating Pressure 250 PSIG, Manufacturer STAINLESS STEEL PRODUCTS (Burbank Ca.)
3 - FUEL FILTER ELEMENT; Manufacturer WINTEC Corp
4 - VENT LINE ASSEMBLY; Manufacturer BELL AEROSYSTEMS
5 - IN LINE MICRO-POROUS FILTER, Helium: Operating Pressure 4000 PSIG
6 - MICRO-PORUS FILTER
Saturn I S-IV Gimbal Actuator Deflection Gage
Saturn I S-IV Gimbal Actuator Obverse View
||SATURN I BLOCK II FLIGHT CONTROL SYSTEM
SATURN I FLIGHT CONTROL S-IV SERVO ACTUATOR
****DETAILED DISCUSSION OF THE SERVO ACTUATOR ASSEMBLY WILL BE UPDATED ON THIS SITE SHORTLY****
SATURN I SECOND STAGE (S-IV) RL-10 ENGINE GIMBAL HYDRAULIC ACTUATOR
Accumulator Reservoir View 2
Accumulator Reservoir View 3
Accumulator Reservoir View 4
Marshall Space Flight Center Tag
Functional Diagram of S-IVB Flight Control System
||SATURN V THIRD STAGE FLIGHT CONTROL SYSTEM
SATURN V (and IB) S-IVB STAGE ACCUMULATOR RESERVOIR ASSEMBLY
A Saturn V Third Stage (S-IVB) Flight Control System Accumulator Reservoir Assembly which provided hydraulic fluids for gimbaling the Rocketdyne J-2 engine. The Accumulator Reservoir was used to enable thrust vector steering, accomplished by gimbaling the J-2 engine for pitch and yaw control during the boost and separation phase. Two servoactuators were used to translate the steering signals received from the Instrument Unit Flight Control Computer into vector forces to position the engine.
Hydraulic fluid from the Accumulator Reservoir is supplied to an engine driven hydraulic pump for deflection of the J2 servoactuators at a rate proportional to the pitch and yaw steering signals issued from the onboard Flight Control Computer. An example of a flight control computer and gimbal servoactuator artifacts from an earlier variant of this stage (the S-IV and its affiliated RL-10�s) can be seen in the listing preceding this entry.
This Accumulator Reservoir, a prototype manufactured by Douglas Aircraft for NASA�s Marshall Space Flight Center under contact NAS7-101 (Prime Saturn S-IVB ) , was mounted as an integral unit on the thrust structure at the bottom of the stage adjacent to the J2. The reservoir section is a storage area for hydraulic fluid and has a maximum volume of 167 cubic inches. During system operation, between 60 and 170 psig is maintained in the reservoir by two pressure operated pistons contained in the accumulator section. In addition to maintaining pressure in the reservoir, the system accumulator supplies peak system demands and dampens high pressure surging.
SATURN V S-IVB (THIRD STAGE) WITH INSTALLED ACCUMULATOR RESERVOIR INDICATED.
ACCUMULATOR RESERVOIR INSTALLED U.S. SPACE AND ROCKET CENTER S-IVB (Photograph courtesy HEROICRELICS.org / Mike Jetzer).
Cover Removed, Dielectric Exposed(Top View)
Radiating Element Exposed
IU Side COAX Connector Interface Panel
Diagram Showing Antenna Location on Instrument Unit
||SATURN V INSTRUMENT UNIT COMMAND AND TELEMETRY
SATURN V INSTRUMENT UNIT (IU) DIRECTIONAL CCS PCM ANTENNA
SATURN V Instrument Unit Directional CCS/PCM Antenna array. The basic structure was fabricated by Metal Research and Mr. Chris Argus of Calumet Fiberglass under subcontract to IBM (NASA Contract NAS 8-1400 )
and measures approximately 8 X 5 X 23 inches . The antenna cover, made of RF transparent epoxy impregnated fiberglass, has been removed in the adjacent images to reveal dielectric foam inserts,
and 5 helical elements, 4 in quadrapole arrangement for high gain UHF S-BAND (2282.5 MHZ @ 20 watts) and 1 monopole element The antenna was recovered by a NASA engineer from a Marshall
Space Flight Center dumpster after disposal. Because it is likely a test article, it lacks the final white titanium dioxide paint coating which would have been applied after installation on the launch vehicle.
This transmit only high gain directional antenna was vehicle fixed on the IU and the radiation pattern was directed toward the earth by controlling the attitude of the spacecraft. It assumed responsibility for the
Omnidirectional antenna pair once the launch vehicle exited their range (approximately 6700 nm above the earth's surface) and provided the Command and Communications System (CCS) downlink and
Pulse Code Modulation (PCM)/Frequency Modulated (FM) telemetry signals to ground stations while also acting as a backup tracking transponder. Two antennas were installed for redundancy onboard the Saturn
V Instrument Unit in the +Z / Position "I" quadrant (see location diagram to lower left).
IDENTICAL CCS/PCM ANTENNA INSTALLED ON SATURN V IU
REDUNDANT CCS/PCM TELEMETRY ANTENNAS ON IU SA-504 (APOLLO 9 SATURN V LAUNCH VEHICLE)
||SATURN V CONTROL VALVE
SATURN V PNEUMATIC PROPELLANT CONTROL VALVE
A uni-directional Pneumatic Control Valve used to regulate the flow of Liquid Hydrogen Propellant (LH2) onboard the Saturn V (S-IVB) third stage, manufactured by Subcontractor Clary Corporation (San Gabriel California) Dec 1966 on behalf of Douglas Aircraft Company (Prime S-IVB) for NASA contact NAS7-101 . This component was part of the pneumatic control system which provided gaseous Helium (Ghe) pressure to actuate all S-IVB stage pneumatically operated valves with the exception of those provided as components of the J-2 engine.
||SATURN V HELIUM FILL MODULE
SATURN V COLD HELIUM FILL MODULE
An onboard Cold Helium Fill module used to facilitate loading of gaseous Helium (Ghe) into the onboard storage tanks; manufactured by subcontractor Fairchild Stratos (Manhattan Beach, California), March 1965 on behalf of Douglas Aircraft Company (Prime S-IVB) for NASA contract NAS7-101 . This component was part of the pneumatic control system which provided Ghe pressure to actuate all S-IVB stage pneumatically operated valves with the exception of those provided as components of the J-2 engine.
Overhead View showing Vent, Inlet and Outlet Ports
Redundant Actuation Cylinder Heads
Obverse View - Valve Outlet Ports A/B
||SATURN V VENT/RELIEF SYSTEM
SATURN V ACTUATION CONTROL MODULE
Apollo Saturn V Third Stage (S-IVB) Actuation Control Module affilated with the Liquid Hydrogen (LH2) propellant cryogenic storage tank Vent&Relief system; produced by CLARY CORPORATION under subcontract to DOUGLAS AIRCRAFT CORP (prime S-IVB) for NASA CONTRACT NAS 7-101.
The module, when command from an external ground signal (during fill operations) and from the Instrument Unit Flight Sequencer (during liftoff and flight) pneumatically triggered actuation of the Vent valve, off-loading excess hydrogen gas from via the non-propulsive
vent ducts on the stage. See S-IVB Vent and Vent/Relief Valves (listed elsewhere on this page) for related discussion.
||SATURN V VENT/RELIEF SYSTEM
SATURN V THIRD STAGE DIRECTIONAL CONTROL VALVE
Saturn V Third Stage Directional Control Fuel Vent Valve manufactured by CALMEC Corporation for Douglas Aircraft Corporation (primary integrator for the S-IVB under
NASA Marshall Space Flight Center Contract NAS7-101). During ground fueling operations, the valve routed Gaseous Hydrogen (GH2) overboard to the burn pond via the LH2 Ground Vent Disconnect Coupling (an example of which
can be seen below in this collection). During flight, the control valve diverted GH2 though the stage non-propulsive vents for explusion into space, or via the propulsive vents to support stage Ullage aumentation (propellant settling) in concert with the Auxiliary Propulsion System (APS). The valve was pneumatically operated utilizing 475 psig gaseous helium from onboard storage tanks.
S-IVB Directional Control Valve in its installed location at the top of the stage Liquid Hydrogen (LH2) Fuel Tank. Outlet "C" (left) leads to overboard discharge of hydrogen gas to the "Burn Pond" during ground based tanking operations, Inlet "A" to the right routes from the the Vent and Relief valve (See Vent and Relief Valve artifact photo-documented below on this page).
||SATURN V PROPELLANT SYSTEM
SATURN V FUEL TANK RELIEF VALVE
A Saturn V Third Stage (S-IVB) Fuel Tank Relief Valve utilized in conjunction with the Liquid Hydrogen (LH2) Cryogenic storage tank and propellant distribution system; manufactured by subcontractor W.O. Leonard Inc (Pasadena, California) un behalf of Douglas Aircraft Company (Prime S-IVB) for NASA contract NAS7-101. This component was part of the vent and relief system capable of relieving all excess pressure accumulated from over-pressurization or fuel boil-off during fill and flight operation.
S-IVB Relief Valve in its installed location at the top of the stage Liquid Hydrogen (LH2) Fuel Tank
Male Coupling(S-IVB Interface)
Exhaust Port (Towards Burn Pond)
||SATURN V PROPELLANT VENT AND RELIEF SYSTEM
SATURN V GASEOUS HYDROGEN GROUND VENT DISCONNECT
Gaseous Hydrogen Ground Vent Line Disconnect coupling manufactured by Fairchild Stratos (Western Division) for Douglas Aircraft Corporation on behalf of the NASA Marshall Spaceflight Center under contract NAS7-101 ( Saturn V S-IVB Contract ). The coupling was intefaced to the Saturn Third Stage via Launch Utility Tower (LUT) Service Arm 7 and provided for safe routing of excess Gaseous Hydrogen (GH2) exported from the stage Vent and Relief Valve to the Hydrogen burn pond.
Vent (Outlet) Ports
Diagram of Saturn V S-IVB Fuel Tank Pressurization and Vent System (location of Vent and Relief Valve highlighted).
||SATURN V THIRD STAGE
SATURN S-IVB THIRD STAGE FUEL TANK VENT AND RELIEF VALVE
Apollo Saturn V S-IVB (Third Stage) Gaseous Hydrogen (GH2) dual function Tank Vent and Relief Valve manufactured by CALMEC under subcontract to Douglas Corporation (prime for the S-IVB NASA CONTRACT NAS 7-101).
The Valve provided relief and overboard venting of excess Gaseous Hydrogen from the fuel tank which serviced the 6 RL-10 engines; routing via the
non-propulsive vent ducts into space (during flight) or the GH2 Burn Pond (during ground fueling operations).
Vent valve actuation
was commanded from an external ground signal during fill operations, and from the flight sequencer during liftoff and flight. The vent valve
was designed to open in a maxiumum of 0.1 second upon command. The relief valve, which provided a backup capablity in case of vent valve failure,
opened at 42 psia and reseated at 39 psia, and had a flow/relief capability of 2 pounds/second at sea level.
S-IVB Vent and Relief Valve in its installed location at the top of the stage Liquid Hydrogen (LH2) Fuel Tank. The Liquid Oxygen (LOX) tank also had a similar valve, and example of which is included inthe collection and is the next described artifact on this page (also see image below for visual comparison of LH2/LOX tank Vent and Relief Valves)
Comparision of S-IVB Vent and Relief Valve for LH2 (left) and LOX (right) tanks
S-IVB Oxidizer Vent and Relief Valve - Pilot Valve can be seen at the top
Vent and Relief Valve Outlet - LOX vapor S-IVB tank was routed out this port to the Propulsive vent for overboard discharge
Vent Check Valve
||SATURN V THIRD STAGE
SATURN S-IVB THIRD STAGE OXIDIZER (LOX) TANK VENT AND RELIEF VALVE
An Apollo Saturn V/IB S-IVB (Third Stage) Liquid Oxygen (LOX) tank propulsive venting Vent and Relief Valve manufactured by CALMEC under subcontract to Douglas Aircraft Corporation (primary design/builder for the S-IVB). The Vent and Relief valve is pneumatically operated using helium gas from the stage pneumatic control system. It is opened upon receipt of a ground command prior to propellant (LOX) fill operations, and the boiloff of LOX during loading is directed through the propulsive vent in the aft skirt of the stage. When LOX tank pressurization commences, the valve is closed and placed in the Relief position. The valve is calibrated to open at a maximum of 45 psia and reseat at 42 psia. During flight, relief venting is normally accomplished via a separate non-propulsive vent valve. Propulsive venting in flight is used to support ullage (positioning of propellants towards the bottom of the stage in preparation for J-2 engine start/restart).
A similar Vent and Relief valve was installed in the Liquid Hydrogen (LH2) tank; an example is included in this collection and is discussed in a listing immediately preceding (above) this artifact.
Vent and Relief Valve Inlet - LOX vapor from the S-IVB tank was routed to this port via a separate Non-Propulsive Vent Valve
Outlet Side of the valve. To the extreme left interior (at the 9 oclock position in this image) is the dual channel relief valve
Pneumatic Piston/Valve Actuator
Disk exibiting fluted channels for mass reduction.Recirculation Supply Line inlet port can be seen at the top.
Installed Prevalve (LH2 Propellant Feed System) on S-IVB-211 (located U.S. Space and Rocket Center)
Saturn S-IVB J2 Prevalve Diagram
Saturn S-IVB J2 Prevalve Functional Schematic
||SATURN V THIRD STAGE/SATURN IB SECOND STAGE
SATURN V S-IVB THIRD STAGE, SATURN IB SECOND STAGE PROPELLANT FEED SYSTEM PREVALVE
Saturn V Third Stage (S-IVB) Oxidizer (LOX) shutoff Prevalve manufactured by SNAP-TITE Inc (AERVALCO division) under subcontract to Douglas Aircraft Corp, prime contractor assigned to design/build the S-IVB. This 10 inch diameter inlet/8 inch outlet butterfly valve is one of two installed on the stage. The fuel prevalve is mounted to the tank integral elbow on the fuel duct (actually mounted on the OML of the vehicle). The LOX prevalve is mounted directly to the LOX tank sump and is located 'under' the thrust structure. The valve disk is actuated by a pneumatic piston (using helium supplied by the stage He control bottle) with a response time of 315 milliseconds. When fully opened, the valve facilitated a LOX flow rate of 386 pounds per second at 60 PSI. An integral relief valve supported propellant bleed off and back flow to the propellant tank when the Prevalve was closed to inhibit rupturing of the main propellant duct section.
The Prevalve is normally open during LOX (and LH2) tank filling to allow all feedducts and engine plumbing (upstream of the main engine propellant valves) to fill with cryogenic propellants resulting in feedsystem chilldown so that the Oxidizer (and Fuel) will remain in liquid form. The valve is also opened in conjunction with powered flight operations. The Prevalves are closed approximately five minutes before liftoff of the Saturn V (10 minutes preceeding liftoff of the Saturn IB) up to the point in the flight until just before J-2 engine start to isolate the propellant tanks and permit the propellant to flow via an alternate path in support of the recirculation chilldown system (this system uses the cryogenic propellant to condition the ducting and engine to the proper temperature level and to eliminate bubbles prior to pressurization and Third Stage start). The Prevalve was actuated shut and Recirculation system run prior to engine restarts to include the period preceding Trans Lunar Injection (TLI) burn.
The Prevalve also provided a redundant propellant shutoff capability, supplemental to the main Oxidizer and Fuel valves on the J-2 engine. A similar Prevalve is used on the Saturn V Second Stage (S-II); an example is also featured within the collection and immediately follows a description of this artifact (also see below image concurrently displaying the S-II and S-IVB prevalves )
S-IVB Prevalve installed on the LOX feed of S-IVB-211 located U.S. Space and Rocket Center (USSRC). Note Recirculation supply line leading to the Prevalve from the right.
Comparison image of S-IVB Prevalve and S-II Prevalve (next listed artifact)
Feedline Input (From Tank)
Output Port (To J2)
Valve Tag Data
||SATURN V SECOND STAGE PROPELLANT SYSTEM
SATURN V S-II STAGE J2 PROPELLANT LINE PRE-VALVE
A Saturn V Second Stage (S-II) Liquid Oxygen (LOX) and Liquid Hydrogen (LH2) Pre-Valve manufactured by North American Aviation (NAA) Space and Informations Systems Division in 1968 under NASA Contract NAS 7-200 (primary contract awarded to North American Rockwell for design/build of the Saturn V Second Stage) . The Pre-Valve was a critical component of the S-II Propellant Delivery System and regulated the flow of propellants through separate feedlines to each of the Rocketdyne J2 engines.
The 8 inch prevalves were normally open, pneumatically actuated, electrically controlled, butterfly-gate type valves. Supported propellant flow rates were 386 pounds of LOX per second at 132 psi; and 78 pounds of LH2 at 132 psi. A built-in four-way pneumatic control solenoid permited 750 +/- 50 psig helium pressure to actuate the butterfly-gate (response time was 1000 milliseconds). Should a loss of pneumatic or electrical power occur, the prevalves were designed to be spring actuated to return to the open position. During ground operations, the Pre-Valve was opened to permit propellant fueling and in conjunction with the recirculation subsystem, allow the propellant to cycle through the feed lines and valves servicing the J2 engines, maintaining uniform cryogenic density and temperature, and precluding the formation of gas in propellant plumbing. Following propellant tank loading and all the way up to the point of S-IC (Saturn V first stage) separation, the Pre-Valves were commanded to remain shut and were opened in conjunction with second stage ignition. They would normally remain open during S-II powered flight unless a signal was received from the engine shutdown system. The valves also provided a redundant shutoff of propellant concurrent with main valve closure.
Inlet Port (Top View)
Outlet Port (Bottom View)
Label Plate(Top View)
||SATURN V HYDROGEN VENT VALVE
SATURN V S-II SECOND STAGE HYDROGEN VENT VALVE ASSEMBLY
An Apollo Saturn V S-II (Second Stage) Hydrogen Vent valve produced by CALMEC Manufacturing Company under subcontract to North American-Rockwell Corporation (prime for the S-II). This assemblage is a component of the Venting Subsystem for the S-II (Saturn V Second Stage which employed five Rocketdyne J-2 LH2/LOX bipropellant fueled engines).
The venting subsystem is used during loading and flight operations of the Saturn V S-II. The valve was physically located atop the S-II Liquid Hydrogen (LH2) tank and provided overboard discharge of excess propellant. While the propellant tanks were being loaded, the vent valve was
opend by electrical signals from ground equipment to allow the gas created by propellant boil-off to leave the tanks. The valve is spring-loaded to be normally closed, but a relief valve would open if pressure in the tanks reached an excessive level.
The valve is capable of venting enough gas to relieve the pressure in its tank; two are provided in the LH2 propellant tank for redundancy.
S-II Vent and Relief Valve in its installed location at the top of the stage Liquid Hydrogen (LH2) Fuel Tank