Limited Lifetime Warranty free from defects in material, workmanship and assembly, under normal use, in accordance with the specifications and warnings, for the life of the product.
Las especificaciones pueden cambiar y mejorar sin previo aviso. This manual is also suitable for: Ph Print page 1 Print document 22 pages. Rename the bookmark. Delete bookmark? Cancel Delete. Delete from my manuals? Sign In OR. Don't have an account? Browse our daily deals for even more savings!
Wiring holden gemini repair i just bought an 82 tf holden gemini with a 1. User manuals, philips tv operating guides and service manuals. It is not necessary to look directly into the light. Consult with your selling dealership or contact thor motor coach, inc. Philips values and respects your privacy. Read this user manual carefully before you use the product. User manuals, guides and specifications for your lg 42lby-tf led tv. The indicator is extinguishes if the attitude rates remain within acceptable limits but illuminates red if the rates exceed these 1imits.
The secondary guidance indicator on the command pilot's panel indicates which guidance system is in operation. The indicator is extinguished to indicate that primary guidance is being used. The indicator illuminates amber to indicate that secondary guidance has been selected. Both indicators illuminate red when the abort command is transmitted.
The indicator signals the crew to initiate immediately the abort mode appropriate for the altitude and velocity of the spacecraft. These modes are described under Sequence System Operation.
During the boost phases the crew has been reminded via the uhf communications link of the abort mode in effect. The stage 1 fuel end oxidizer meters on the command pilot's panel enable the crew to monitor the current status and progress of the boost phase, and to anticipate an abort condition if one should develop. These meters indicate the gas pressures in psia of the stage 1 fuel and oxidizer tanks.
Dual indicator needles are provided for redundancy. The range of the stage 1 meters is 35 to 5 psia. A time-versus- pressure scale near the bottom of the meter shows the minimum required pressure at 20, 40, and 60 seconds after lift-off.
Critical fuel tank pressure is indicated by a shaded column at the low end of the scale. After staging with no signals applied, the meters indicate maximum psia.
The stage 2 fuel and oxidizer meters on the command pilot's panel indicate stage 2 fuel and oxidizer tank pressure over a 70 to 10 psia range. Redundant pointers are used. Critical fuel tank pressures are indicated by a shaded column at the low end of the scale.
The S-flag at the psia mark indicates the minimum acceptable stored pressure in the tank before pressurization. After spacecraft separation, the meters indicate maximum psia. The accelerometer on the command pilot's panel indicates the rate in g's at which the launch vehicle engines are changing the velocity of the spacecraft. The range of the accelerometer is minus 6g's to 16 g's.
The meter has positive and negative memory pointers. The accelerometer enables the crew to monitor the effectiveness of the engines. It is a secondary indicator of staging. The guidance switch above the abort control handle permits the command pilot to manually change from primary guidance to secondary backup guidance. When back-up guidance has been selected either manually or automatically during stage I boost and the ground station determines that primary guidance is feasible during stage 2 boost, primary guidance can be selected again by momentarily placing the guidance switch to the RGS position.
A D-ring is provided on the ejection seat of each pilot. These rings are pulled to initiate mode I abort at altitude below 70, feet. Refer to Section III of this volume for the location and operation of these devices. The abort control handle is located on the command pilot's side of the cabin. These modes are effective above 25, feet. When the abort handle is moved to ABORT, an immediate spacecraft separation and retrograde sequence is performed. These sequences differs from the normal sequences in that they are performed without cues from the indicators on the main instrument panel.
The switches, indicators, and switches-indicators on the main instrument panel center console have the following nomenclature, place in the mission sequence, and functions. The jettison fairing switch is used at the end of second stage engine thrust decay, by the command pilot to jettison the nose fairing, and the horizon scanner head cover.
The separate spacecraft switch-indicator is used in the separation-insertion phase of the sequence. The command pilot presses the switch-indicator approximately 20 seconds after second stage engine cutoff when the IVI displays the delta-V required for insertion.
Pressing the switch-indicator causes several things to happen. Primarily, it detonates pyrotechnic devices which separate the spacecraft from the launch vehicle. Secondarily, it extends the uhf and diplexer antennas and readies the acquisition aid beacon for use.
As the spacecraft moves away from the launch vehicle, separation sensors close and energize the spacecraft separation relays. The relays illuminate the Indicator green. The Indicate retrograde attitude switch-indicator is illuminated amber when the electronics timer energizes the Tr second relay. The amber light cues the crew to press the switch-indicator at this time.
When pressured, a bias voltage is placed on the pitch needle of the FDI, and the inertial platform is electrically placed in the BEF mode.
When released the ember light is extinguished and a green light is illuminated. The battery power indicator illuminated amber by the Tr second relay. This change must be made because the adapter section will be jettisoned at retrograde. When all of the main battery switches are on, the indicator changes from amber to green.
The amber light cues the command pilot to activate the RCS firing the fuel, oxidizer, and pressurant isolation squibs. Pressing the switch-indicator energizes relays which fire the squibs.
The indicator changes from ember to green, indicating that the RCS has been activated. The separate OAMS lines indicator is illuminated amber by the Tr second relay is the prepare-to-go to retrograde phase. The amber light cues the crew to seal and sever the OAMS lines before jettisoning the adapter. Pressing the switch-indicator energizes relays which ignite the pyrotechnics used to seal and sever the lines.
The relays also fire pyrotechnic switches and wire guillotines severing some of the adapter-retrograde mating line wiring. The indicator changes from amber to green. The separate electrical indicator is also illuminated amber by the Tr second relay. Pressing the switch-indicator energizes the wire guillotine relay. The pyrotechnics are detonated and the wiring is cut. The indicator changes from amber to green to indicate that electrical separation has been accomplished.
The separate adapter indicator is illuminated amber by the Tr second relay. The amber light cues the crew to Jettison the adapter equipment section. Pressing the switch-indicator causes the adapter shaped charge and the Z7O tubing cutter pyrotechnic to be detonated, and the adapter section severed.
Separation of the adapter section is sensed by the equipment adapter separation sensors. Two closed sensors energize the sensor relay and change the indicator from amber to green. The arm automatic retrofire indicator is illuminated amber by the Tr second relay. The amber light cues the crew to arm the automatic retrofire circuits so that when the electronic timer closes the TR contacts at TR time, the retrorockets will fire automatically. Pressing the switch-indicator completes the patch from the retrograde common control bus to the timer TR contact, and also energizes the TR arm relay.
The relay changes the light from amber to green. Contact closure at Tr time energizes the Tr signal relay. The signal relay energizes the second time delay relay, fires the retro rockets at 5. The manual fire retrorockets switch connects the retrograde common control bus to the manual retrograde latch relay.
Contacts of this relay energizes the second time delay relay, fire the retrorockets at 5. The Jettison retrograde adapter indicator is illuminated amber by the second time delay relay 45 seconds after retrofire begins.
The amber light cues the crew to jettison the retrograde adapter. It fires the shaped charges which sever the retrograde adapter section from the re-entry vehicle. It energizes the Horizon Sensor System scanner head jettison relays which fire the jettison squibs and Jettison the scanner head.
It removes the retrograde attitude signals applied to the flight director needles at Tr seconds. It switches the FDI roll channel to the mix mode for re-entry. Ten Sequence System relay panels are installed in Gemini Spacecraft 5, 6, and 8 through Four relay panels are located in the re-entry vehicle, three in the retrograde section, one in the equipment section, and two in the rendezvous and recovery section. The following Sequence System relay panels are in the re-entry module.
The necessary functions required for adapter retrograde section separation are performed by the fourteen relays of the retrograde separation relay panel. The relays perform such functions as pyrotechnic switch and shaped charge ignition, Tr-3O second indication, automatic IGS free mode selection, and arming of the contacts of the Time Reference System.
Re-entry Control System squib firing, scanner cover and scanner heads jettison, abort interlock RCS amber light actuation, and RCS ring B squib firing test prior to launch are provided by the sixteen relays of the attitude control system scanner and RCS squib fire relay panel. The umbilical pyrotechnic switch relay panel contains two relays which apply landing squib bus I and 2 power to re-entry umbilical wiring pyrotechnic switch. The retrograde section contains the following three relay panels which control spacecraft separation, retrofire, and equipment section separation.
The equipment section contains the Orbit Attitude Maneuver System squib fire relay panel. The retrofire relay panel has twenty relays. These relays control the automatic, manual and salvo firing of the retro rockets, and time the 5. The retrograde sequence adapter separate relay panel contains twelve relays. The relays are used for equipment adapter shaped charge ignition, propellant line guillotine, electrical wire guillotine, and retrograde abort. The Rendezvous and Recovery section contains two Sequence System relay panels: the nose fairing Jettison relay panel, and the docking relay panel.
The nose fairing Jettison relay panel contains two relays which control the jettisoning of the nose fairing. The docking relay panel has eleven relays which extend the docking index bar, illuminate the MSG ACPT light, effect emergency release of the docking latches, release and jettison the locking latches at retrograde, jettison the index bar, and cover the docking latch ports.
The Sequence System contains two sets of separation sensors. Separation sensors are toggle switches which are normally open before separation is initiated.
The separating structure will close the sensors as it moves away from the spacecraft re-entry module. The closure of any two of a set of three sensors is sufficient to sense and indicate separation. No primary ac electrical power system Is provided for the spacecraft. Devices requiring ac power obtain this power from self-contained inverters within the individual systems. The Electrical Power System includes switches, circuit breakers, relay panels, ammeters, a voltmeter and telelights which provide control, distribution and monitoring for the system.
Also included as an Electrical Power System subsystem is the Reactant Supply System RSS which provides storage and control of the reactants hydrogen and oxygen used for fuel cell battery operation not applicable to spacecraft 6.
Provisions are made for utilizing external power and remote monitoring of the spacecraft power buses during ground tests and pre-launch operations.
The two fuel cell battery sections and four main batteries provide dc power to the spacecraft main power bus on spacecraft 6, the three adapter module batteries and the four main batteries provide dc power to the main bus.
On spacecraft 5 and 6, a dual-vertical-readout ammeter is located on the right instrument panel. On spacecraft 5, two FCAP indicator lamps are located on the right instrument panel. On spacecraft 6, a conventional voltmeter and ammeter with associated selector switches are located on the right instrument panel.
The FC O2 and H2 regulator and heater circuit breakers are inoperative on spacecraft 6. The O2 and H2 heater switches, quantity indicator and selector switch are identified as CRYO switches and indicator on spacecraft 1O, 11, and The power system relay panel, power distribution relay panel and adapter power supply relay panel are located in the left equipment area of the cabin. In order to conserve spacecraft battery power, external electrical power is utilized during the pre-launch phase of the mission.
External power is supplied to the spacecraft common control, main and squib power buses through umbilical cables connected to the re-entry module and equipment adapter section umbilical receptacles. Remote control of the spacecraft squib bus arming relays, and remote monitoring of the spacecraft power buses is also accomplished through the re-entry and adapter umbilical cables. On spacecraft 5 and 8 through 12, the fuel cell batteries are activated in sufficient time prior to launch, to insure launch readiness of the fuel cell batteries and RSS.
The common control bus and OAMS squib buses are switched from external power to the squib batteries in sufficient time prior to launch, to verify the squib battery circuits. The re-entry and equipment adapter section umbilicals are disconnected from the spacecraft Just prior to lift-off.
Normally, umbilical separation is accomplished by an electrical solenoid device. A backup method of separation is also provided by a lanyard initiated mechanism which is actuated by movement of the launch vehicle.
From launch time until booster separation and insertion into orbit, both the fuel cell battery sections module batteries on spacecraft 6 and the four main batteries are connected in parallel to the main power bus. In the event of an abort, all squib buses required for the abort function, which are not armed prior to launch, are armed via the abort relays controlled by the Sequential System.
These relays effectively bypass the applicable squib bus arming switches which normally arm these buses. All three squib batteries are connected to the common control bus through diodes for individual fault protection. Squib batteries 1 and 2 are connected to the two OAMS squib buses via the de-energized squib bus arming relays. This switch controls the O2 crossfeed valve, which provides the capability of connecting the ECS O2 supply to the fuel cell battery sections.
This switch, when in VENT position, initiates a pyrotechnic cutter device to perform this function. On spacecraft 5 and 8 through 12, a small percentage of the reactant gases must be purged from the fuel cell batteries periodically to insure that the impurities contained in the feed gases do not restrict reactant flow to the cells and to remove any accumulation of product water in the gas lines.
The pilots may increase the flow of gases to the fuel cell sections for more effective purging by setting the X-OVER switch to ON position. The retrograde rockets are used according to the mission requirements. There is no automatic switching provided for this function. This switch provides a method of arming the JETT RETRO switch center instrument panel , and is effectively an interlock to prevent inadvertent Jettisoning of the retrograde section prior to firing of the retrograde rockets.
All unnecessary electrical equipment will be deactivated to conserve the remaining spacecraft batteries for recovery equipment operation. Throughout the mission, visual displays of bus voltage and current are provided by the system voltmeter and ammeters. The ammeters monitor individual fuel cell stack current IA through 2C. The delta pressure indicator, in conjunction with a selector switch, provides a visual display of O2 versus H2 and O2 versus H2O differential pressure in the fuel cell battery sections.
In the event that the differential pressure exceeds the prescribed limits, the pilots must evaluate the fuel cell battery performance, and if a malfunction exists, shut down the malfunctioning fuel cell battery section by setting the applicable fuel cell power and stack control switches to OFF position.
The delta pressure indicator is inoperative on spacecraft 5. An out of tolerance delta pressure indication is also provided by the fuel cell delta pressure FCAP telelights on the center instrument panel. The lights are illuminated red when a malfunction exists. On spacecraft 5 only, two FCAP indicator lamps on the right instrument panel are illuminated red when a malfunction exists.
The BTRY PWR sequence light, located on the center instrument panel, is illuminated amber at Tr seconds during the mission by action of the Tr-5 relay in the power system relay panel. This informs the pilots that they must return the MAIN BATTERIES switches to ON position to insure continuity of main bus power due to the impending separation of the equipment adapter section containing the adapter power supply fuel cell battery sections on spacecraft 5 and 8 through 12 and silver-zinc batteries on spacecraft 6.
On spacecraft 5 and 6 the dual-vertical-readout section ammeter provides a display of section 1 and 2 main bus current. Section 1 includes 50 percent of the adapter power supply current fuel cell batteries or silver-zinc batteries as applicable plus main batteries 1 and 2 current. Section 2 includes 50 percent of the adapter power supply current plus main batteries 3 and 4 current.
The stack ammeter used for battery test ammeter on spacecraft 6 with selector switch in 1A, IB, 1C or 2A, 2B, 2C positions, displays 50 percent of adapter module battery current. Displays of common control bus voltage, main bus voltage, OAMS squib bus voltage, and adapter module battery voltage is provided by the system voltmeter and selector switch.
The FCAP telelights and reactant quantity indications are not operative on spacecraft 6. The squib batteries are special high-discharge-rate batteries which will maintain a terminal voltage of 18 volts for one second under a 75 ampere load. These batteries are used in lieu of fuel cell batteries. All of the silver-zinc batteries have an open circuit terminal voltage of The main and squib battery cases are made of titanium. The approximate activated wet weight for each squib battery is 8 lb.
The adapter module battery cases spacecraft 6 are constructed of magnesium and the approximate wet weight of each battery is lbs. The battery electrolyte consists of a 40 percent solution of reagent grade potassium hydroxide and distilled water. The main and squib batteries have a vent valve in each cell designed to prevent electrolyte loss and will vent the cell to atmospheric pressure in the event a pressure in excess of 40 psig builds up within the cell.
All of the silver-zinc batteries are equipped with relief valves which maintain a tolerable interior to exterior differential pressure in the battery cases. The batteries are capable of operating in any attitude in a weightless state. Prior to installation into the spacecraft, the batteries are activated and sealed at sea level pressure.
All of the batteries are coldplate mounted to control battery temperature. The power system relay panel contains relays necessary for controlling and sequencing power system functions. The panel contains the control relays for the fuel cell battery system and RSS, main battery power sequence light relay, Tr-5 relay and the squib bus arming relays.
The power distribution relay panel contains the relays required for arming the retrograde squib buses in the event of an abort. These relays are controlled by the Sequential System. The adapter power supply relay panel contains relays necessary for controlling adapter module power to the main power bus. The relay panel contains the stack power relays which connect the individual fuel cell stacks to the main bus. On spacecraft 6 the stack power relays connect the adapter module batteries to the main bus.
The panel also contains diodes used for reverse current protection between the adapter power supply and the spacecraft main power bus. The main bus section ammeter spacecraft 5 and 6 is a dual-edge-readout vertical reading meter having a ampere range with a total accuracy of two percent. The NO. The ammeter is shunt connected between the main power bus and spacecraft ground.
With the selector switch in 1A, 1B, 1C or 2A, 2B, 2C positions, the ammeter displays 50 percent of the applicable adapter module battery current. The meter has a ampere scale. On spacecraft 6 the voltmeter, used in conjunction with a selector switch, displays main bus, common control bus and squib bus voltage. The voltmeter has a vdc range.
The delta pressure indicator not operative on spacecraft 5 and voltmeter are used in conjunction with selector switches located just below on the instrument panel. The ammeters provide a display of individual fuel cell stack 1A through 2C current reading must be multiplied by 0.
The AC position on the selector switch Is inoperative on spacecraft 5. The voltmeter has a ac volt range and an dc volt range. The delta pressure indicator has a This indicator displays O2 versus H2 differential pressure and O2 versus H2O differential pressure for each fuel cell battery section. The fuel cell batteries used in the Gemini Spacecraft are of the solid ion exchange membrane type using hydrogen H2 for fuel and oxygen O2 for an oxidizer. The fuel cell battery system is comprised of two separate sections which are sealed in air tight pressure containers.
Each section is made up of three interconnected fuel cell stacks with plumbing for transferring hydrogen, oxygen and product water. Each fuel cell stack consists of 32 individual fuel cells. Each basic fuel cell is made up of two catalytic electrodes separated by a solid type electrolyte in laminated form. The electrolyte is composed of a sulfonated styrene polymer plastic approximately 0. Thin films of platinum catalyst, applied to both sides of the electrolyte, act as electrodes and support ionization of hydrogen on the anode side of the cell and oxidation on the cathode side of the cell.
A thin titanium screen, imbedded into the platinum catalytic electrode, reduces the internal resistance along the current flow path from the electrode to the current collector and adds strength to the solid electrolyte. On the hydrogen side of the fuel cell, a current collector is attached by means of a glass-cloth-reinforced epoxy frame which assures a tight seal around the edges of the cell, forming a closed chamber. Ribs in the collector are in contact with the catalytic electrode on the fuel cell, providing a path for current flow.
The hydrogen fuel is admitted through an inlet tube in the frame of the current collector and enters each gas channel between the collector ribs by way of a series of slots in the tube.
Another tube provides a purge outlet, making it possible to flush accumulated inert gases from the cell. The collector plate is made of approximately 0. On the oxygen side of the cell, a current collector of the same configuration and material as the hydrogen side collector is attached.
Its ribs, located at right angles to those of the other collector, provide structural support to the electrolyte-electrode structure.
A Dacron cloth wick, attached between the ribs, carries away the product water through capillary action, by way of a termination bar on one side of the assembly. Oxygen is admitted freely to this side of the fuel cell from the oxygen filled area of the section container. The cell cooling system consists of two separate tubes bonded in the cavity formed by the construction of the oxygen side current collector and the back side of the hydrogen current collector. Each tube passes through six of the collector ribs and has the cooling capacity to maintain operating temperature.
The cooling of the oxygen current collector, which holds the product water transport wicks provides the coldplate for water condensation from the warmer oxygen electrode. The individual fuel cell assemblies are arranged in series to form a stack.
When assembling the cells into a stack, the ribs of the oxygen side current collector contact the solid electrolyte of the fuel cell assembly. Titanium terminal plates are installed on the ends of the two outside cells to which connections are made for the external circuit.
End plates, which are honey-comb structures of epoxy-glass laminate 0. Stainless steel insulated tie rods hold the stack together and maintain a compression load across the area of each cell assembly. This assures proper contact of the solid electrolyte with the ribs of each current collector.
The fuel cell stacks are packaged in a pressure tight container, together with the necessary reactant and coolant ducts and manifolds, water separator for each stack, and required electrical power and instrumentation wiring.
The hydrogen inlet line, hydrogen purge line, and the two coolant lines for each cell lead from their respective common manifolds running the length of the stack. The manifolds are made of an insulating plastic material and the individual cell connections are potted in place after assembly to provide a leak-tight seal.
The oxygen sides of the cells are open to the oxygen environment surrounding the fuel cell assemblies within the container. An accessory pad is mounted on the outside of the fuel cell section container.
It includes the gas inlet and outlet fittings, purge and shutoff valves, water valve and electrical connectors. Structurally, the container is a titanium pressure vessel consisting of a central cylinder with two end covers and two mounting brackets. Within the container, the fuel cell stacks are mounted on fiberglass impregnated epoxy rails by bolts which pass through the stack plates.
These rails are in turn bolted to the mounting rings sandwiched between the two flanges on the section container. The hydrogen manifolds on each stack within a section are parallel fed with a hydrogen shutoff valve and check valve in the feed line to each stack. Oxygen is fed into the section container so that the entire free volume of the container contains oxygen at approximately The coolant reaches the fuel cell battery sections by two separate isolated lines.
Any malfunction in the coolant line in one section will not affect the cooling function of the coolant line in the other section. Each stack in the section has its own water-oxygen separators which are manifolded into a single line coming out of the section container.
All hydrogen, oxygen coolant, electrical and water storage pressure line connections at the section container are fastened to standard bulkhead fittings on the accessory pad. After the stacks are completely assembled within the container, all void spaces are filled with unicellular foam.
The purpose of this foaming is for vibration dampening, acoustical noise deadening and minimizing free gas volume to prevent possible fire propagation. Thin plastic covers are placed over the top and bottom of each stack to manifold oxygen to the stack and to keep the foam material from entering areas around the coolant manifolds and oxygen water separator. The basic principle by which the fuel cell operates to produce electrical energy and water, is the controlled oxidation of hydrogen.
This is accomplished through the use of the solid electrolyte ion-exchange membrane. On the hydrogen side of the fuel cell, hydrogen gas disassociates on the catalytic electrode to provide hydrogen ions and electrons. The electrons are provided a conducting path of low resistance by the current collector, either to an external load or to the next series-connected fuel cell. When a flow of electrons is allowed to do work and move to the oxygen side of the fuel cell, the reaction will proceed.
By use and replacement, hydrogen ions flow through the solid electrolyte to the catalytic electrode on the oxygen side of the fuel cell. When electrons are available on this surface, oxygen disassociates and combines with the available hydrogen ions to form water. The oxygen current collector provides the means of distributing electrons and condensing the product water on a surface to be transported away by the wick system through capillary action. The individual cell wicks are integrated into one large wick which routes the water to an absorbent material that separates the water from the gas.
By using the oxygen outlet pressure as a reference, a small pressure differential is obtained over the length of the water removal system. This pressure is sufficient to push the gas-free water toward the storage reservoir. Waste heat, generated during the fuel cell battery operation, is dissipated by means of the recirculating coolant provided by the spacecraft cooling system.
In addition, the total coolant flow provides the function of preheating the incoming reactant gases. In the spacecraft, the reactant gases are supplied to the fuel cell battery sections by the RSS. This system contains the reactant supply tanks, control valves, heat exchangers, temperature sensors and heaters required for management of the fuel cell reactants. The RSS is essentially a subsystem for the fuel cell battery sections.
The system provides storage for the cryogenic hydrogen and oxygen, converts the reactants to gaseous form and controls the flow of the gases to the fuel cell battery sections. This vessel also supplies O2 to the ECS. Two tanks are utilized to separately contain the cryogenic hydrogen and oxygen required for the operation of the fuel cell battery sections.
The tanks are thermally insulated to minimize heat conduction to the stored elements which would cause the homogeneous solution to revert to a mixture of gas and liquid. The tanks are capable of maintaining the stored liquids at supercritical pressures and cryogenic temperatures. The approximate total amount of liquid stored in the hydrogen vessel is The approximate total amount of liquid stored in the oxygen vessel is lb.
The hydrogen vessel is composed of titanium alloy and the oxygen vessel Is made of a high strength nickel base alloy. Both vessels are spherical in shape and double welded. A vacuum space between the inner and outer wall approximately one inch provides thermal insulation from ambient heat conduction.
The inner wall is supported in relation to the outer wall by an insulating material supplemented by compression loading devices. Each storage tank contains a fluid quantity sensor, a pressure sensor, a temperature sensor and an electrical heater the O2 tank on spacecraft 10, 11, and 12 has two heaters installed in the inner vessel in intimate contact with the stored reactants.
The fluid quantity sensor is an integral capacitance unit which operates in conjunction with an indicator control unit containing a null bridge amplifier. The sensor varies the capacitance in proportion to fluid level in a circuit connected to the null bridge amplifier. The amplified signal is then used to drive a servo motor, which in turn operates a visual indicator for quantity indication.
Power inverters supply cycle, 26 vac power to the fluid quantity circuits. The temperature sensor is a platinum resistance device capable of transmitting a source signal to a balanced bridge circuit. The sensor provides cryogenic fluid temperature monitoring for telemetry and AGE.
The pressure sensor is a dual resistive element, diaphragm type transducer. The sensor provides signals for cryogenic fluid pressure monitoring on a spacecraft meter.
The electrical heaters provide a method of accelerating pressure build-up in the reactant supply tanks. The heaters may be operated either in a manual or automatic mode. In the automatic mode a pressure switch removes power from the heater element when the tank pressure builds up to a nominal psig in the oxygen tank and a nominal psig in the hydrogen tank.
In the manual mode a spacecraft pressure meter indicates proper switch operation. The fill and vent valves provide a dual function in permitting simultaneous fill and vent operations. Quick disconnect fittings are provided for rapid ground service connection to both the storage tank fill check valve and the vent check valve.
When fill connections are made, the pressure of the ground service connection against the fill and vent valve poppet shaft simultaneously opens both the fill and vent ports. When ground service equipment is removed, the valve poppet automatically returns to its normally spring-loaded-closed position. The vent check valve is a single poppet type, spring-loaded-closed valve which opens when system pressure exceeds 20 psig to relieve through the fill and vent valve vent ports.
The supply temperature control heat exchangers are finned heat exchangers in which the supply fluid temperature is automatically controlled by absorbing heat from the recirculating coolant fluid of the spacecraft cooling system. The dual pressure regulator and relief valves are normally open poppet type regulators which control downstream pressure to the fuel cell battery sections. The regulators maintain the hydrogen pressure at approximately The oxygen side of the regulators is referenced to hydrogen pressure.
The hydrogen side of the regulators is referenced to produce H2O pressure. The relief valves provide overpressurization protection for the regulated pressure to the fuel cell battery sections. This valve is precalibrated to operate at a pressure of approximately 10 psia above the normal supply level. The high pressure relief valves are single poppet type, spring-loaded-closed valves which provide system overpressurization protection.
The valves vent system gas to ambient when pressure exceeds the system limits. The solenoid shutoff valves are solenoid operated latching type valves which eliminate fluid loss during the nonoperating standby periods. The valves are normally open and are closed only during fill and standby periods by applying power to the solenoids. The crossover valve is a solenoid operated latching type valve which provides the capability of selecting both dual pressure regulators to supply hydrogen and oxygen to a fuel cell battery section for the purpose of increasing flow rate for more effective purging.
The O2 crossfeed reactant valve is a solenoid operated, latching type valve which provides the capability of pressurizing the RSS with O2 from the ECS oxygen supply. This provides a redundant method of supplying the proper reactant O2 pressure to the fuel cell battery sections in the event of a malfunction in the RSS oxygen supply.
During pre-launch, the two separate reactant supply tanks are serviced using AGE equipment with liquid hydrogen and oxygen. After the tanks are filled, in order to accelerate pressure buildup within the tanks, the internal tank heaters are operated, utilizing external electrical power.
In approximately one hour the liquid is converted into a high density, homogeneous fluid at a constant pressure. During the fill operation, the solenoid shutoff valves between the storage tanks and the dual pressure regulators are closed. Once operating pressure is obtained, the solenoid shutoff valves may be opened by applying power to the coil of the valves.
The high density, homogeneous fluid will then flow upon demand. The fluid flows from the supply tanks to the heat exchangers. The heat exchangers absorb heat from the recirculating coolant fluid of the spacecraft cooling system. The reactants, now in gaseous form, flow through the heat exchangers, past the high pressure relief valves and AGE temperature sensors, to the supply solenoid shutoff valves.
During fuel cell battery operation, if the demand on the fluid flow is inadequate to keep tank pressures within limits, the high pressure relief valves will vent, externally, the excess fluid.
The AGE temperature sensors on the heat exchangers are used for pre-launch checkout only. The reactants flow through the supply solenoid shutoff valves to the dual pressure regulator and relief valves. The dual pressure regulators reduce the pressure of the reactants to approximately The gas now flows through the manual shutoff valves and is then directed to the fuel cell battery sections at a flow rate that is determined by both the electrical load applied and the frequency of purging.
0コメント