MAIN PROPULSION SYSTEM

The main propulsion system, assisted by the two solid rocket boosters during the initial phases of the ascent trajectory, provides the velocity increment from lift-off to a predetermined velocity increment before orbit insertion. The two SRBs are jettisoned after their fuel has been expended, but the MPS continues to thrust until the predetermined velocity is achieved. At that time, main engine cutoff is initiated. The external tank is jettisoned, and the orbital maneuvering system is ignited to provide the final velocity increment for orbital insertion. The magnitude of the velocity increment supplied by the OMS depends on payload weight, mission trajectory and system limitations.

Coincident with the start of the OMS thrusting maneuver (which settles the MPS propellants), the remaining liquid oxygen propellant in the orbiter feed system and space shuttle main engines is dumped through the nozzles of the three SSMEs. At the same time, the remaining liquid hydrogen propellant in the orbiter feed system and SSMEs is dumped overboard through the hydrogen fill and drain valves for six seconds. Then the hydrogen inboard fill and drain valve is closed, and the hydrogen recirculation valve is opened, continuing the dump. The hydrogen flows through the engine hydrogen bleed valves to the orbiter hydrogen MPS line between the inboard and outboard hydrogen fill and drain valves, and the remaining hydrogen is dumped through the outboard fill and drain valve for approximately 120 seconds.

During on-orbit operations, the flight crew vacuum inerts the MPS by opening the liquid oxygen and liquid hydrogen fill and drain valves, which allows the remaining propellants to be vented to space.

Before entry, the flight crew repressurizes the MPS propellant lines with helium to prevent contaminants from being drawn into the lines during entry and to maintain internal positive pressure. MPS helium is also used to purge the spacecraft's aft fuselage. The last activity involving the MPS occurs at the end of the landing rollout. At that time, the helium remaining in onboard helium storage tanks is released into the MPS to provide an inert atmosphere for safety.

The MPS consists of the following major subsystems: three SSMEs, three SSME controllers, the external tank, the orbiter MPS propellant management subsystem and helium subsystem, four ascent thrust vector control units, and six SSME hydraulic servoactuators.

The main engines are reusable, high-performance, liquid-propellant rocket engines with variable thrust. The propellant fuel is liquid hydrogen and the oxidizer is liquid oxygen. The propellant is carried in separate tanks in the external tank and supplied to the main engines under pressure. Each engine can be gimbaled plus or minus 10.5 degrees in the yaw axis and plus or minus 10.5 degrees in the pitch axis for thrust vector control by hydraulically powered gimbal actuators.

The main engines can be throttled over a range of 65 to 109 percent of their rated power level in 1-percent increments. A value of 100 percent corresponds to a thrust level of 375,000 pounds at sea level and 470,000 pounds in a vacuum. A value of 104 percent corresponds to 393,800 pounds at sea level and 488,800 pounds in a vacuum; 109 percent corresponds to 417,300 pounds at sea level and 513,250 pounds in a vacuum.

At sea level, the engine throttling range is reduced due to flow separation in the nozzle, prohibiting operation of the engine at its 65-percent throttle setting, referred to as minimum power level. All three main engines receive the same throttle command at the same time. Normally, these come automatically from the orbiter general-purpose computers through the engine controllers. During certain contingency situations, manual control of engine throttling is possible through the speed brake/thrust controller handle. The throttling ability reduces vehicle loads during maximum aerodynamic pressure and limits vehicle acceleration to 3 g's maximum during boost.

Each engine is designed for 7.5 hours of operation over a life span of 55 starts. Throughout the throttling range, the ratio of the liquid oxygen-liquid hydrogen mixture is 6-to-1. Each nozzle area ratio is 77.5-to-1. The engines are 14 feet long and 7.5 feet in diameter at the nozzle exit.

The SSME controllers are digital, computer system, electronic packages mounted on the SSMEs. They operate in conjunction with engine sensors, valve actuators and spark igniters to provide a self-contained system for monitoring engine control, checkout and status. Each controller is attached to the forward end of the SSME.

Engine data and status collected by each controller are transmitted to the engine interface unit, which is mounted in the orbiter. There is one EIU for each main engine. The EIU transmits commands from the orbiter GPCs to the main engine controller. When engine data and status are received by the EIU, the data are held in a buffer until the EIU receives a request for data from the computers.

Three orbiter hydraulic systems provide hydraulic pressure to position the SSME servoactuators for thrust vector control during the ascent phase of the mission in addition to performing other functions in the main propulsion system. The three orbiter auxiliary power units provide mechanical shaft power through a gear train to drive the hydraulic pumps that provide hydraulic pressure to their respective hydraulic systems.

The ascent thrust vector control units receive commands from the orbiter GPCs and send commands to the engine gimbal actuators. The units are electronics packages (four in all) mounted in the orbiter's aft fuselage avionics bays. Hydraulic isolation commands are directed to engine gimbal actuators that indicate faulty servovalve position. In conjunction with this, a servovalve isolation signal is transmitted to the computers.

The SSME hydraulic servoactuators are used to gimbal the main engine. There are two actuators per engine, one for pitch motion and one for yaw motion. They convert electrical commands received from the orbiter GPCs and position servovalves, which direct hydraulic pressure to a piston that converts the pressure into a mechanical force that is used to gimbal the SSMEs. The hydraulic pressure status of each servovalve is transmitted to the ATVC units.

The orbiter MPS propellant management subsystem consists of the manifolds, distribution lines and valves by which the liquid propellants pass from the external tank to the main engines and the gaseous propellants pass from the main engines to the external tank. The SSMEs' gaseous propellants are used to pressurize the external tank. All the valves in the propellant management subsystem are under direct control of the orbiter GPCs and are either electrically or pneumatically actuated.

The orbiter MPS helium subsystem consists of a series of helium supply tanks and regulators, check valves, distribution lines and control valves. The subsystem supplies the helium used within the engine to purge the high-pressure oxidizer turbopump intermediate seal and preburner oxidizer domes and to actuate valves during emergency pneumatic shutdown. The balance of the helium is used to actuate all the pneumatically operated valves within the propellant management subsystem and to pressurize the propellant lines before re-entry.

ORBITER MAIN PROPULSION SYSTEM HELIUM SUBSYSTEM

The MPS helium subsystem consists of seven 4.7-cubic-foot helium supply tanks; three 17.3-cubic-foot helium supply tanks; and associated regulators, check valves, distribution lines and control valves. Four of the 4.7-cubic-foot helium supply tanks are located in the aft fuselage, and the other three are located below the payload bay liner in the midfuselage in the area originally reserved for the cryogenic storage tanks of the power reactant storage and distribution system. The three 17.3-cubic-foot helium supply tanks are also located below the payload bay liner in the midfuselage.

The tanks are composite structures consisting of a titanium liner with a fiberglass structural overwrap. The large tanks are 40.3 inches in diameter and have a dry weight of 272 pounds. The smaller tanks are 26 inches in diameter and have a dry weight of 73 pounds. The tanks are serviced before lift-off to a pressure of 4,500 psi.

Each of the larger supply tanks is plumbed to two of the smaller supply tanks (one in the midbody, the other in the aft body), forming three sets of three tanks for the engine helium pneumatic supply system. Each set of tanks normally provides helium to only one engine and is commonly referred to as left, center, or right engine helium, depending on the engine serviced. Each set normally provides helium to its designated engine for in-flight purges and provides pressure for actuating engine valves during emergency pneumatic shutdown.

The remaining 4.7-cubic-foot helium tank is referred to as the pneumatic helium supply tank. It normally provides pressure to actuate all of the pneumatically operated valves in the propellant management subsystem.

There are eight helium supply tank isolation valves grouped in pairs. One pair of valves is connected to each engine helium supply tank cluster, and one pair is connected to the pneumatic supply tank. In the engine helium supply tank system, each pair of isolation valves is connected in parallel, with each valve in the pair controlling helium flow through one leg of a dual-redundant helium supply circuit. Each helium supply circuit contains two check valves, a filter, an isolation valve, a regulator and a relief valve. The two isolation valves connected to the pneumatic supply tanks are also connected in parallel; however, the rest of the pneumatic supply system consists of a filter, the two isolation valves, a regulator, a relief valve and a single check valve. Each engine helium supply isolation valve can be individually controlled by its He isolation A left , ctr , right open , GPC , close and He isolation B left , ctr , right , open , GPC, close switches on panel R2. The two pneumatic helium supply isolation valves are controlled by a single pneumatic He isol , open, GPC, close switch on panel R2.

All of the valves in the helium subsystem (with the exception of the supply tank isolation valves) are spring loaded to one position and electrically actuated to the other position. The supply tank isolation valves are spring loaded to the closed position and pneumatically actuated to the open position. Valve position is controlled via electrical signals from either the onboard GPCs or manually by the flight crew. All of the valves can be controlled automatically by the GPCs, and the flight crew can control some of the valves.

The helium source pressure of the pneumatic, left, center and right supply systems can be monitored on the helium , pneu , l (left), c (center), r (right) meters on panel F7 by positioning the tank, reg (regulator) switch below the meters to tank . In addition, the regulated pressure of the pneumatic, left, center and right systems can be monitored on the same meters by placing the switch to reg .

Each of the four helium supply systems operates independently until after main engine cutoff. Each engine helium supply has two interconnect (crossover) valves associated with it, and each valve in the pair of interconnect valves is connected in series with a check valve. The check valves allow helium to flow through the interconnect valves in one direction only. One check valve associated with one interconnect valve controls helium flow in one direction, and the other interconnect valve and its associated check valve permit helium flow in the opposite direction. The in interconnect valve controls helium flow into the associated engine helium distribution system from the pneumatic helium supply tank. The out interconnect valve controls helium flow out of the associated engine helium supply system to the pneumatic distribution system.

Each pair of interconnect valves is controlled by a single switch on panel R2. Each He interconnect , left , ctr , right switch has three positions- in open/out close , GPC , and in close/out open. With the switch in the in open/out close position, the in interconnect valve is open and the out interconnect valve is closed. The in close/out open position does the reverse. With the switch in GPC, the out interconnect valve opens automatically at the beginning of the liquid oxygen dump and closes automatically at the end of the liquid hydrogen dump.

In a return-to-launch-site abort, the GPC position will cause the in interconnect valve to open automatically at MECO and close automatically 20 seconds later. If an engine was shut down before MECO, its in interconnect valve will remain closed at MECO. At any other time, placing the switch in GPC results in both interconnect valves being closed.

An additional interconnect valve between the left engine helium supply and pneumatic helium supply would be used if the pneumatic helium supply regulator failed. This crossover valve would be opened and the pneumatic helium supply tank isolation valves would be closed, allowing the left engine helium supply system to supply helium to the pneumatic helium supply. The crossover helium valve is controlled by its own three-position switch on panel R2. The pneumatics l (left) eng He xovr (crossover) switch positions are open, GPC and close. The GPC position allows the valve to be controlled by the flight software loaded in the GPCs.

Manifold pressurization valves located downstream of the pneumatic helium pressure regulator are used to control the flow of helium to propellant manifolds during a nominal propellant dump and manifold repressurization. There are four of these valves grouped in pairs. One pair controls helium pressure to the liquid oxygen propellant manifolds, and the other pair controls helium pressure to the liquid hydrogen propellant manifold.

The liquid hydrogen RTLS dump pressurization valves located downstream of the pneumatic helium pressure regulator are used to control the pressurization of the liquid hydrogen propellant manifolds during an RTLS liquid hydrogen dump. There are two of these valves connected in series. Unlike the liquid hydrogen manifold pressurization valves, the liquid hydrogen RTLS dump pressurization valves cannot be controlled by flight deck switches. During an RTLS abort, these valves are opened and closed automatically by GPC commands. An additional difference between the nominal and the RTLS liquid hydrogen dumps is in the routing of the helium and the place where it enters the liquid hydrogen feed line manifold. For the nominal liquid hydrogen dump, helium passes through the liquid hydrogen manifold pressurization valves and enters the feed line manifold in the vicinity of the liquid hydrogen feed line disconnect valve. For the liquid hydrogen RTLS dump, helium passes through the RTLS liquid hydrogen dump pressurization valves and enters the feed line manifold in the vicinity of the liquid hydrogen inboard fill and drain valve on the inboard side. There is no RTLS liquid oxygen dump pressurization valve since the liquid oxygen manifold is not pressurized during the RTLS liquid oxygen dump.

Each engine helium supply tank has two pressure regulators operating in parallel. Each regulator controls pressure in one leg of a dual-redundant helium supply circuit and is capable of providing all of the helium needed by the main engines.

The pressure regulator for the pneumatic helium supply system is not redundant and is set to provide outlet pressure between 715 to 770 psig. Downstream of the regulator are two more regulators: the liquid hydrogen manifold pressure regulator and the liquid oxygen manifold pressure regulator. These regulators are used only during MPS propellant dumps and manifold pressurization. Both regulators are set to provide outlet pressure between 20 to 25 psig. Flow through the regulators is controlled by the appropriate set of two normally closed manifold pressurization valves.

Downstream of each pressure regulator, with the exception of the two manifold repressurization regulators, is a relief valve. The valve protects the downstream helium distribution lines from overpressurization if the associated regulator fails fully open. The two relief valves in each engine helium supply are set to relieve at 785 to 850 psig and reseat at 785 psig. The relief valve in the pneumatic helium supply circuit also relieves at 785 to 850 psig and reseats at 785 psig.

There is one pneumatic control assembly on each of the three space shuttle main engines. The PCA is essentially a manifold pressurized by one of the engine helium supply systems and contains solenoid valves to control and direct pressure to perform various essential functions. The valves are energized by discrete on/off commands from the output electronics of the associated SSME controller. Functions controlled by the PCA include the high-pressure oxidizer turbopump intermediate seal cavity and preburner oxidizer dome purge, pogo system postcharge and pneumatic shutdown.

MAIN PROPULSION SYSTEM PROPELLANT MANAGEMENT SUBSYSTEM

Within the orbiter aft fuselage, liquid hydrogen and liquid oxygen pass through the manifolds, distribution lines and valves of the propellant management subsystem.

During prelaunch activities, this subsystem is used to control the loading of liquid oxygen and liquid hydrogen in the external tank. During SSME thrusting periods, propellants from the external tank flow into this subsystem and to the three SSMEs. The subsystem also provides a path that allows gases tapped from the three SSMEs to flow back to the external tank through two gas umbilicals to maintain pressure in the external tank's liquid oxygen and liquid hydrogen tanks. After MECO, this subsystem controls MPS dumps, vacuum inerting and MPS repressurization for entry.

All the valves in the MPS are either electrically or pneumatically operated. Pneumatic valves are used where large loads are encountered, such as in the control of liquid propellant flows. Electrical valves are used for lighter loads, such as in the control of gaseous propellant flows.

The pneumatically actuated valves are divided into two types: those that require pneumatic pressure to open and close the valve (type 1) and those that are spring loaded to one position and require pneumatic pressure to move to the other position (type 2).

Each type 1 valve actuator is equipped with two electrically actuated solenoid valves. Each solenoid valve controls helium pressure to an ''open'' or ''close'' port on the actuator. Energizing the solenoid valve on the open port allows helium pressure to open the pneumatic valve. Energizing the solenoid on the close port allows helium pressure to close the pneumatic valve. Removing power from a solenoid valve removes helium pressure from the corresponding port of the pneumatic actuator and allows the helium pressure trapped in that side of the actuator to vent overboard. Removing power from both solenoids allows the pneumatic valve to remain in the last commanded position. This type of valve is used for the liquid oxygen and liquid hydrogen feed line 17-inch umbilical disconnect valves (two), the liquid oxygen and liquid hydrogen prevalves (six), the three liquid hydrogen and liquid oxygen inboard and outboard fill and drain valves (four), and the liquid hydrogen 4-inch recirculation disconnect valves.

Each type 2 valve is a single electrically actuated solenoid valve that controls helium pressure to either an open or a close port on the actuator. Removing power from the solenoid valve removes helium pressure from the corresponding port of the pneumatic actuator and allows helium pressure trapped in that side of the actuator to vent overboard. Spring force takes over and drives the valve to the opposite position. If the spring force drives the valve to the open position, the valve is referred to as a normally open valve. If the spring force drives the valve to a closed position, the valve is referred to as a normally closed valve. This type of valve is used for the liquid hydrogen RTLS inboard dump valve (NC), the liquid hydrogen RTLS outboard dump valve (NC), the liquid hydrogen feed line relief shutoff valve (NO), the liquid oxygen feed line relief shutoff valve (NO), the three liquid hydrogen engine recirculation valves (NC), the two liquid oxygen pogo recirculation valves (NO), the liquid hydrogen topping valve (NC), the liquid hydrogen high-point bleed valve (NC), and the liquid oxygen overboard bleed valve (NO).

The electrically actuated solenoid valves are spring loaded to one position and move to the other position when electrical power is applied. These valves also are referred to as either normally open or normally closed, based on their position in the de-energized state. Electrically actuated solenoid valves are the gaseous hydrogen pressurization line vent valve (NC), the three gaseous hydrogen pressurization flow control valves (NO) and the three gaseous oxygen pressurization flow control valves (NO).

There are two 17-inch-diameter MPS propellant feed line manifolds in the orbiter aft fuselage, one for liquid oxygen and one for liquid hydrogen. Each manifold has an outboard and inboard fill and drain valve in series that interface with the respective port (left) and starboard (right) T-0 umbilical. The port T-0 umbilical is for liquid hydrogen; the starboard, for liquid oxygen. In addition, each manifold connects the orbiter to the external tank in the lower aft fuselage through a port 17-inch liquid hydrogen disconnect valve umbilical and a starboard 17-inch liquid oxygen disconnect valve umbilical.

There are three outlets in both the liquid oxygen and liquid hydrogen 17-inch manifolds between the orbiter-external tank 17-inch umbilical disconnect valves and the inboard fill and drain valve. The outlets in the manifolds provide liquid oxygen and liquid hydrogen to each SSME in 12-inch-diameter feed lines.

The prevalve in each of the three liquid oxygen and liquid hydrogen 12-inch feed lines to each engine isolates liquid oxygen and liquid hydrogen from each engine or permits liquid oxygen and liquid hydrogen to flow to each engine. Each prevalve is controlled by an LH 2 or LO 2 prevalve , left , ctr , right switch on panel R4. Each switch has open, GPC and close positions.

The 8-inch-diameter liquid hydrogen outboard and inboard fill and drain valves are also controlled by their own switches on panel R4. Each propellant fill/drain LH 2 , outbd , inbd switch has open, gnd and close positions, as does each LO2, outbd, inbd switch.

Each 17-inch liquid hydrogen and liquid oxygen manifold has a 1-inch-diameter line that is routed to a feed line relief isolation valve and feed line relief valve in the respective liquid hydrogen and liquid oxygen system. The LO 2 and LH 2 feed line rlf (relief) isol (isolation) switches on panel R4 have open , GPC and close positions. When a feed line relief isolation valve is opened, the corresponding manifold can relieve excessive pressure overboard through its relief valve.

The liquid hydrogen feed line manifold has another outlet directed to the two liquid hydrogen RTLS dump valves in series. Both valves are controlled by the MPS prplt dump LH 2 valve switch on panel R2, which has backup LH 2 vlv open , GPC , close positions. When opened, these valves enable the liquid hydrogen dump during RTLS aborts or provide a backup to the normal liquid hydrogen dump after a nominal main engine cutoff. In an RTLS abort dump, liquid hydrogen is dumped overboard through a port at the outer aft fuselage's left side between the orbital maneuvering system/reaction control system pod and the upper surface of the wing.

The MPS propellant management subsystem also contains two 2-inch-diameter manifolds, one for gaseous oxygen and one for gaseous hydrogen. Each manifold individually permits ground support equipment servicing with helium through the respective T-0 umbilical and provides initial pressurization of the external tank's liquid oxygen and liquid hydrogen orbiter/external tank disconnect umbilicals. Self-sealing quick disconnects are provided at the T-0 umbilical and the orbiter/external tank umbilical.

Six 0.63-inch-diameter pressurization lines, three for gaseous oxygen and three for gaseous hydrogen, are used after SSME start to pressurize the external tank's liquid oxygen and liquid hydrogen tanks.

In each SSME, a small portion of liquid oxygen is diverted into the engine's oxidizer heat exchanger, and the heat generated by the engine's high-pressure oxidizer turbopump converts the liquid oxygen into gaseous oxygen and directs it through a check valve to two orifices and a flow control valve for each engine. During SSME thrusting periods, liquid oxygen tank pressure is maintained between 20 and 22 psig by the orifices in the two lines and the action of the flow control valve from each SSME. The flow control valve is controlled by one of three liquid oxygen pressure transducers. When tank pressure decreases below 20 psig, the valve opens. If the tank pressure is greater than 24 psig, it is relieved through the liquid oxygen tank's vent and relief valve.

In each SSME, gaseous hydrogen from the low-pressure fuel turbopump is directed through two check valves to two orifices and a flow control valve for each engine. During the main engine thrusting period, the liquid hydrogen tank's pressure is maintained between 32 and 34 psia by the orifices and the action of the flow control valve from each SSME. The flow control valve is controlled by one of three liquid hydrogen pressure transducers. When tank pressure decreases below 32 psia, the valve opens; and when tank pressure increases to 33 psia, the valve closes. If the tank pressure is greater than 35 psia, the pressure is relieved through the liquid hydrogen tank's vent and relief valve. If the pressure falls below 32 psia, the LH 2 ullage press switch on panel R2 is positioned from auto to open , which will cause all three flow control valves to go to full open and remain in the full-open position.

The single gaseous hydrogen manifold repressurization line connects to the hydrogen line vent valve, which is controlled by the H 2 press line vent switch on panel R4. This valve is normally closed, and the switch is positioned to open when vacuum inerting the gaseous hydrogen pressurization lines after MECO and the liquid hydrogen dump. The gnd position allows the launch processing system to control the valve during ground operations.

MPS EXTERNAL TANK

The external tank is attached to the orbiter at one forward and two aft attach points. At the two aft attach points are the two external tank/orbiter umbilicals for the fluid, gas, signal and electrical power connections between the orbiter and the external tank. Each external tank umbilical plate mates with a corresponding umbilical plate on the orbiter. The umbilical plates help maintain alignment of the various connecting components. The corresponding umbilical plates are bolted together; and when external tank separation is commanded, the bolts are severed by pyrotechnics.

At the forward end of each external tank propellant tank is a vent and relief valve that can be opened by GSE-supplied helium before launch for venting or by excessive tank pressure for relief. The vent function is available only before launch; after lift-off only the relief function is operable. The liquid oxygen tank relieves at an ullage pressure of 25 psig, while the liquid hydrogen tank relieves at an ullage pressure of 38 psi. The flight crew has no control over the position of the vent and relief valves before launch or during ascent. Normal range of the tank ullage pressure of the liquid hydrogen tank during ascent is 32 to 39 psia. During prelaunch activities, the liquid hydrogen tank is pressurized to 44.1 psi to meet the start requirement of the main engine LPFT. The liquid oxygen and liquid hydrogen tanks' ullage pressures are monitored on the panel F7 eng manf LO2 and LH2 meters as well as on a cathode ray tube display.

In addition to the vent and relief valve, the liquid oxygen tank has a tumble vent valve that is opened during the external tank separation sequence. The thrust force provided by opening the valve imparts an angular velocity to the external tank to assist in the separation maneuver and provide more positive control of the external tank's re-entry aerodynamics.

There are eight propellant depletion sensors. Four of them sense fuel depletion and four sense oxidizer depletion. The oxidizer depletion sensors are mounted in the external tank's liquid oxygen feed line manifold downstream of the tank. The fuel depletion sensors are located in the liquid hydrogen tank. During prelaunch activities, the launch processing system tests each propellant depletion sensor. If any are found to be in a failed condition, the LPS sets a flag in the computer's SSME operational sequence, sequence logic that will instruct the computer to ignore the output of the failed sensor or sensors. During main engine thrusting, the computer constantly computes the instantaneous mass of the vehicle, which constantly decreases due to propellant usage from the external tank. When the computed vehicle mass matches a predetermined initialized-loaded value, the computer arms the propellant depletion sensors. After this time, if any two of the good fuel depletion sensors (those not flagged before launch) or any two of the good oxidizer depletion sensors indicate a dry condition, the computers command main engine cutoff. This type of MECO is a backup to the nominal MECO, which is based on vehicle velocity. The oxidizer sensors sense propellant depletion before the fuel sensors to ensure that all depletion cutoffs are fuel-rich since an oxidizer-rich cutoff can cause burning and severe erosion of engine components. To ensure that the oxidizer sensors sense depletion first, a plus 700-pound bias is included in the amount of liquid hydrogen loaded in the external tank. This amount is in excess of that dictated by the 6-1 ratio of oxidizer to fuel. The position of the oxidizer propellant depletion sensors allows the maximum amount of oxidizer to be consumed in the engines and allows sufficient time to cut off the engines before the oxidizer turbopumps cavitate (run dry).

Four ullage pressure transducers are located at the top end of each propellant tank (liquid oxygen and liquid hydrogen). One of the four is considered a spare and is normally off-line. Before launch, GSE normally checks out the four transducers; and if one of the three active transducers is determined to be bad, it can be taken off-line and the output of the spare transducer selected. The flight crew can also perform this operation after lift-off via the computer keyboard; however, because of the time involved from lift-off to MECO, this would probably be impractical. The three active ullage pressure sensors provide outputs for CRT display and control of ullage pressure within their particular propellant tanks. For CRT display, computer processing selects the middle value output of the three transducers and displays this single value. For ullage pressure control, all three outputs are used.

The external tank/orbiter aft umbilicals have five propellant disconnects: two for the liquid oxygen tank and three for the liquid hydrogen tank. One of the liquid oxygen propellant umbilicals carries liquid oxygen and the other carries gaseous oxygen. The liquid hydrogen tank has two disconnects that carry liquid hydrogen and one that carries gaseous hydrogen. The external tank liquid hydrogen recirculation disconnect is the smaller of the two disconnects that carry liquid hydrogen and is used only during the liquid hydrogen chill-down sequence before launch.

In addition, the external tank/orbiter umbilicals contain two electrical umbilicals, each made of many smaller electrical cables. These cables carry electrical power from the orbiter to the external tank and the two solid rocket boosters and bring telemetry back to the orbiter from the SRBs and external tank. The operational instrumentation telemetry that comes back from the SRBs is conditioned, digitized and multiplexed in the SRBs themselves. The external tank OI measurements that return to the orbiter are raw transducer outputs and must be processed within the orbiter telemetry system.

The external tank's liquid oxygen tank is serviced at the launch pad before prelaunch from ground support equipment through the starboard T-0 umbilical of the orbiter, the MPS outboard and inboard fill and drain valves, the MPS 17-inch liquid oxygen line, and the orbiter/external tank 17-inch umbilical disconnect valves. Once the liquid oxygen is loaded and ready for main engine ignition, the liquid oxygen tank's vent and relief valve is closed, and the tank is pressurized to 21 psig by GSE-supplied helium. During SSME thrusting, liquid oxygen flows out of the external tank through the orbiter/external tank umbilical into the orbiter MPS and to each SSME. Pressurization in the tank is maintained by gaseous oxygen tapped from the three main engines and supplied to the liquid oxygen tank through the orbiter/external tank gaseous oxygen umbilical.

The external tank's liquid hydrogen tank is serviced before launch from GSE at the launch pad similarly to the liquid oxygen tank but through the port T-0 umbilical and port orbiter/external tank umbilical. When the liquid hydrogen is loaded and ready for main engine ignition, the liquid hydrogen tank's vent and relief valve is closed, and the tank is pressurized to 42.5 psia by GSE-supplied helium.

Approximately 45 minutes after loading starts, three electrically powered liquid hydrogen pumps in the orbiter begin to circulate the liquid hydrogen in the external tank through the three SSMEs and back to the external tank through a special recirculation umbilical. This recirculation chills down the liquid hydrogen lines between the external tank and the high-pressure fuel turbopump in the SSMEs so that the path is free of any gaseous hydrogen bubbles and is at the proper temperature for engine start. Recirculation ends approximately six seconds before engine start. During engine thrusting, liquid hydrogen flows from the external tank and through the orbiter/external tank liquid hydrogen umbilical into the orbiter MPS and to the main engines. Tank pressurization is maintained by gaseous hydrogen tapped from the three SSMEs and supplied to the liquid hydrogen tank through the orbiter/external tank gaseous hydrogen umbilical.

SPACE SHUTTLE MAIN ENGINES

Oxidizer from the external tank enters the orbiter at the orbiter/external tank umbilical disconnect and then the orbiter's main propulsion system liquid oxygen feed line. There it branches out into three parallel paths, one to each engine. In each branch, a liquid oxygen prevalve must be opened to permit flow to the low-pressure oxidizer turbopump.

The LPOT is an axial-flow pump driven by a six-stage turbine powered by liquid oxygen. It boosts the liquid oxygen's pressure from 100 psia to 422 psia. The flow from the LPOT is supplied to the high-pressure oxidizer turbopump. During engine operation, the pressure boost permits the HPOT to operate at high speeds without cavitating. The LPOT operates at approximately 5,150 rpm. The LPOT, which is approximately 18 by 18 inches, is connected to the vehicle propellant ducting and supported in a fixed position by the orbiter structure.

The HPOT consists of two single-stage centrifugal pumps (a main pump and a preburner pump) mounted on a common shaft and driven by a two-stage, hot-gas turbine. The main pump boosts the liquid oxygen's pressure from 422 psia to 4,300 psia while operating at approximately 28,120 rpm. The HPOT discharge flow splits into several paths, one of which is routed to drive the LPOT turbine. Another path is routed to and through the main oxidizer valve and enters into the main combustion chamber. Another small flow path is tapped off and sent to the oxidizer heat exchanger. The liquid oxygen flows through an anti-flood valve that prevents it from entering the heat exchanger until sufficient heat is present to convert the liquid oxygen to gas. The heat exchanger utilizes the heat contained in the discharge gases from the HPOT turbine to convert the liquid oxygen to gas. The gas is sent to a manifold and is then routed to the external tank to pressurize the liquid oxygen tank. Another path enters the HPOT second-stage preburner pump to boost the liquid oxygen's pressure from 4,300 psia to 7,420 psia. It passes through the oxidizer preburner oxidizer valve into the oxidizer preburner and through the fuel preburner oxidizer valve into the fuel preburner. The HPOT is approximately 24 by 36 inches. It is attached by flanges to the hot-gas manifold.

Fuel enters the orbiter at the liquid hydrogen feed line disconnect valve, then flows into the orbiter gaseous hydrogen feed line manifold and branches out into three parallel paths to each engine. In each liquid hydrogen branch, a prevalve permits liquid hydrogen to flow to the low-pressure fuel turbopump when the prevalve is open.

The LPFT is an axial-flow pump driven by a two-stage turbine powered by gaseous hydrogen. It boosts the pressure of the liquid hydrogen from 30 psia to 276 psia and supplies it to the high-pressure fuel turbopump. During engine operation, the pressure boost provided by the LPFT permits the HPFT to operate at high speeds without cavitating. The LPFT operates at approximately 16,185 rpm. The LPFT is approximately 18 by 24 inches. It is connected to the vehicle propellant ducting and is supported in a fixed position by the orbiter structure 180 degrees from the LPOT.

The HPFT is a three-stage centrifugal pump driven by a two-stage, hot-gas turbine. It boosts the pressure of the liquid hydrogen from 276 psia to 6,515 psia. The HPFT operates at approximately 35,360 rpm. The discharge flow from the turbopump is routed to and through the main valve and then splits into three flow paths. One path is through the jacket of the main combustion chamber, where the hydrogen is used to cool the chamber walls. It is then routed from the main combustion chamber to the LPFT, where it is used to drive the LPFT turbine. A small portion of the flow from the LPFT is then directed to a common manifold from all three engines to form a single path to the external tank to maintain liquid hydrogen tank pressurization. The remaining hydrogen passes between the inner and outer walls to cool the hot-gas manifold and is discharged into the main combustion chamber. The second hydrogen flow path from the main fuel valve is through the engine nozzle (to cool the nozzle). It then joins the third flow path from the chamber coolant valve. The combined flow is then directed to the fuel and oxidizer preburners. The HPFT is approximately 22 by 44 inches. It is attached by flanges to the hot-gas manifold.

The oxidizer and fuel preburners are welded to the hot-gas manifold. The fuel and oxidizer enter the preburners and are mixed so that efficient combustion can occur. The augmented spark igniter is a small combination chamber located in the center of the injector of each preburner. The two dual-redundant spark igniters, which are activated by the engine controller, are used during the engine start sequence to initiate combustion in each preburner. They are turned off after approximately three seconds because the combustion process is then self-sustaining. The preburners produce the fuel-rich hot gas that passes through the turbines to generate the power to operate the high-pressure turbopumps. The oxidizer preburner's outflow drives a turbine that is connected to the HPOT and the oxidizer preburner pump. The fuel preburner's outflow drives a turbine that is connected to the HPFT.

The HPOT turbine and HPOT pumps are mounted on a common shaft. Mixing of the fuel-rich hot gas in the turbine section and the liquid oxygen in the main pump could create a hazard. To prevent this, the two sections are separated by a cavity that is continuously purged by the MPS engine helium supply during engine operation. Two seals minimize leakage into the cavity. One seal is located between the turbine section and the cavity, and the other is between the pump section and cavity. Loss of helium pressure in this cavity results in an automatic engine shutdown.

The speed of the HPOT and HPFT turbines depends on the position of the corresponding oxidizer and fuel preburner oxidizer valves. These valves are positioned by the engine controller, which uses them to throttle the flow of liquid oxygen to the preburners and, thus, control engine thrust. The oxidizer and fuel preburner oxidizer valves increase or decrease the liquid oxygen flow, thus increasing or decreasing preburner chamber pressure, HPOT and HPFT turbine speed, and liquid oxygen and gaseous hydrogen flow into the main combustion chamber, which increases or decreases engine thrust, thus throttling the engine. The oxidizer and fuel preburner valves operate together to throttle the engine and maintain a constant 6-1 propellant mixture ratio.

The main oxidizer valve and the main fuel valve control the flow of liquid oxygen and liquid hydrogen into the engine and are controlled by each engine controller. When an engine is operating, the main valves are fully open.

A coolant control valve is mounted on the combustion chamber coolant bypass duct of each engine. The engine controller regulates the amount of gaseous hydrogen allowed to bypass the nozzle coolant loop, thus controlling its temperature. The chamber coolant valve is 100 percent open before engine start. During engine operation, it will be 100 percent open for throttle settings of 100 to 109 percent for minimum cooling. For throttle settings between 65 to 100 percent, its position will range from 66.4 to 100 percent open for maximum cooling.

Each engine main combustion chamber receives fuel-rich hot gas from a hot-gas manifold cooling circuit. The gaseous hydrogen and liquid oxygen enter the chamber at the injector, which mixes the propellants. A small augmented spark igniter chamber is located in the center of the injector. The dual-redundant igniter is used during the engine start sequence to initiate combustion. The igniters are turned off after approximately three seconds because the combustion process is self-sustaining. The main injector and dome assembly is welded to the hot-gas manifold. The main combustion chamber also is bolted to the hot-gas manifold.

The inner surface of each combustion chamber, as well as the inner surface of each nozzle, is cooled by gaseous hydrogen flowing through coolant passages. The nozzle assembly is a bell-shaped extension bolted to the main combustion chamber. The nozzle is 113 inches long, and the outside diameter of the exit is 94 inches. A support ring welded to the forward end of the nozzle is the engine attach point to the orbiter-supplied heat shield. Thermal protection for the nozzles is necessary because of the exposure that portions of the nozzles experience during the launch, ascent, on-orbit and entry phases of a mission. The insulation consists of four layers of metallic batting covered with a metallic foil and screening.

The five propellant valves on each engine (oxidizer preburner oxidizer, fuel preburner oxidizer, main oxidizer, main fuel, and chamber coolant) are hydraulically actuated and controlled by electrical signals from the engine controller. They can be fully closed by using the MPS engine helium supply system as a backup actuation system.

The low-pressure oxygen and low-pressure fuel turbopumps are mounted 180 degrees apart on the orbiter's aft fuselage thrust structure. The lines from the low-pressure turbopumps to the high-pressure turbopumps contain flexible bellows that enable the low-pressure turbopumps to remain stationary while the rest of the engine is gimbaled for thrust vector control. The liquid hydrogen line from the LPFT to the HPFT is insulated to prevent the formation of liquid air.

The main oxidizer valve and fuel bleed valve are used after shutdown. The main oxidizer valve is opened during a propellant dump to allow residual liquid oxygen to be dumped overboard through the engine, and the fuel bleed valve is opened to allow residual liquid hydrogen to be dumped through the liquid hydrogen fill and drain valves overboard. After the dump is completed, the valves close and remain closed for the remainder of the mission.

The gimbal bearing is bolted to the main injector and dome assembly and is the thrust interface between the engine and orbiter. The bearing assembly is approximately 11.3 by 14 inches.

Overall, a space shuttle main engine weighs approximately 7,000 pounds.

POGO SUPPRESSION SYSTEM

A pogo suppression system prevents the transmission of low-frequency flow oscillations into the high-pressure turbopump and, ultimately, prevents main combustion chamber pressure (engine thrust) oscillation. Flow oscillations transmitted from the space shuttle vehicle are suppressed by a partially filled gas accumulator, which is attached by flanges to the high-pressure oxidizer turbopump's inlet duct.

The system consists of a 0.6-cubic-foot accumulator with an internal standpipe, helium precharge valve package, gaseous oxygen supply valve package and two recirculation isolation valves (one located on the orbiter).

During engine start, the accumulator is charged with helium 2.4 seconds after the start command to provide pogo protection until the engine heat exchanger is operational and gaseous oxygen is available.

The accumulator is partially chilled by liquid oxygen during the engine chill-down operation. It fills to the overflow standpipe line inlet level, which is sufficient to preclude gas ingestion at engine start.

During engine operation, the accumulator is charged with a continuous gaseous oxygen flow maintained at a rate governed by the engine operation point.

The liquid level in the accumulator is controlled by the overflow standpipe line in the accumulator, which is orificed to regulate the gaseous oxygen overflow over the engine's operating power level. The system is sized to provide sufficient replenishment of gaseous oxygen at the minimum flow rate and to permit sufficient gaseous oxygen overflow at the maximum decreasing pressure transient in the low-pressure oxidizer turbopump discharge duct. At all other conditions, excess gaseous and liquid oxygen are recirculated to the the low-pressure oxidizer turbopump inlet through the engine oxidizer bleed duct. The pogo accumulator is charged (pressurized) at engine shutdown to provide a positive pressure at the HPOT inlet, which prevents HPOT overspeed in the zero-gravity environment.

SPACE SHUTTLE MAIN ENGINE CONTROLLERS

The controller is an electronics package mounted on each SSME. It contains two digital computers and the associated electronics to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings.

Each controller operates in conjunction with engine sensors, valves, actuators and spark igniters to provide a self-contained system for engine control, checkout and monitoring. The controller provides engine flight readiness verification; engine start and shutdown sequencing; closed-loop thrust and propellant mixture ratio control; sensor excitation; valve actuator and spark igniter control signals; engine performance limit monitoring; onboard engine checkout, response to vehicle commands and transmission of engine status; and performance and maintenance data.

Each engine controller receives engine commands transmitted by the orbiter's general-purpose computers through its own engine interface unit. The engine controller provides its own commands to the main engine components. Engine data are sent to the engine controller, where they are stored in a vehicle data table in the controller's computer memory. Data on the controller's status compiled by the engine controller's computer are also added to the vehicle data table. The vehicle data table is periodically output by the controller to the EIU for transmission to the orbiter's GPCs.

The engine interface unit is a specialized multiplexer/demultiplexer that interfaces with the GPCs and with the engine controller. When engine commands are received by the EIU, the data are held in a buffer until the EIU receives a request for data from the GPCs. The EIU then sends data to each GPC. Each EIU is dedicated to one space shuttle main engine and communicates only with the engine controller that controls its SSME. The EIUs have no interface with each other.

The controller provides responsive control of engine thrust and propellant mixture ratio throughout the digital computer in the controller, updating the instructions to the engine control elements 50 times per second (every 20 milliseconds). Engine reliability is enhanced by a dual-redundant system that allows normal operation after the first failure and a fail-safe shutdown after a second failure. High-reliability electronic parts are used throughout the controller.

The digital computer is programmable, allowing engine control equations and constants to be modified by changing the stored program (software). The controller is packaged in a sealed, pressurized chassis and is cooled by convection heat transfer through pin fins as part of the main chassis. The electronics are distributed on functional modules with special thermal and vibration protection.

The controller is divided into five subsystems: input electronics, output electronics, computer interface electronics, digital computer and power supply electronics. Each subsystem is duplicated to provide dual-redundant capability.

The input electronics receive data from all engine sensors, condition the signals and convert them to digital values for processing by the digital computer. Engine control sensors are dual-redundant, and maintenance data sensors are non-redundant.

The output electronics convert computer digital control commands into voltages suitable for powering the engine spark igniters, the off/on valves and the engine propellant valve actuators.

The computer interface electronics control the flow of data within the controller, data input to the computer and computer output commands to the output electronics. They also provide the controller interface with the vehicle engine electronics interface unit for receiving engine commands that are triple-redundant channels from the vehicle and for transmitting engine status and data through dual-redundant channels to the vehicle. The computer interface electronics include the watchdog timers that determine which channel of the dual-redundant mechanization is in control.

The digital computer is an internally stored, general-purpose computer that provides the computational capability necessary for all engine control functions. The memory has a program storage capacity of 16,384 data and instruction words (17-bit words; 16 bits for program use, one bit for parity).

The power supply electronics convert the 115-volt, three-phase, 400-hertz vehicle ac power to the individual power supply voltage levels required by the engine control system and monitor the level of power supply channel operation to ensure it is within satisfactory limits.

Each orbiter GPC, operating in a redundant set, issues engine commands to the engine interface units for transmission to their corresponding engine controllers. Each orbiter GPC has SSME subsystem operating program applications software residing in it. Engine commands are output over the engine's assigned flight-critical data bus (a total of four GPCs outputting over four FC data buses). Therefore, each EIU will receive four commands. The nominal ascent configuration has GPCs 1, 2, 3 and 4 outputting on FC data buses 5, 6, 7 and 8, respectively. Each FC data bus is connected to one multiplexer interface adapter in each EIU.

The EIU checks the received engine commands for transmission errors. If there are none, the EIU passes the validated engine commands on to the controller interface assemblies, which output the validated engine commands to the engine controller. An engine command that does not pass validation is not sent to the controller interface assembly. Instead, it is dead-ended in the EIU's multiplexer interface adapter. Commands that come through MIAs 1 and 2 are sent to CIAs 1 and 2, respectively. Commands that come to MIAs 3 and 4 pass through a CIA 3 data-select logic. This logic outputs the command that arrives at the logic first, from either MIA 3 or 4. The other command is dead-ended in the CIA 3 select logic. The selected command is output through CIA 3. In this manner, the EIU reduces the four commands sent to the EIU to three commands output by the EIU.

The engine controller vehicle interface electronics receive the three engine commands output by its EIU, check for transmission errors (hardware validation), and send controller hardware-validated engine commands to the controller A and B electronics. Normally, channel A electronics are in control, with channel B electronics active, but not in control. If channel A fails, channel B will assume control. If channel B subsequently fails, the engine controller will shut down the engine pneumatically. If two or three commands pass voting, the engine controller will issue its own commands to accomplish the function commanded by the orbiter GPCs. If command voting fails and two or all three commands fail, the engine controller will maintain the last command that passed voting.

The backup flight system computer, GPC 5, contains SSME hardware interface program applications software. When the four primary GPCs (1, 2, 3 and 4) are in control, the BFS GPC does no commanding. When GPC 5 is in control, the BFS sends commands to, and requests data from, the EIU; and in this configuration, the four primary GPCs neither command nor listen. The BFS, when engaged, allows GPC 5 to command FC buses 5, 6, 7 and 8 for main engine control through the SSME HIP. The SSME HIP performs the same main engine command functions as the SSME subsystem operating program. The command flow through the EIUs and engine controllers is the same when the BFS is engaged as for the four-GPC redundant set.

The engine controller provides all the main engine data to the GPCs. Sensors in the engine supply pressures, temperatures, flow rates, turbopump speeds, valve position and engine servovalve actuator positions to the engine controller. The engine controller assembles these data into a vehicle data table and adds status data of its own to the vehicle data table. The vehicle data tables output channels A and B to the vehicle interface electronics for transmission to the EIUs. The vehicle interface electronics output over both data paths. The data paths are called primary and secondary. The channel A vehicle data table is normally sent over both primary and secondary control (channel A has failed); then the vehicle interface electronics output the channel B vehicle data table over both the primary and secondary data paths.

The vehicle data table is sent by the controller to the EIU. There are only two data paths versus three command paths between the engine controller and the EIU. The data path that interfaces with CIA 1 is called primary data. The path that interfaces with CIA 2 is called secondary data. Primary and secondary data are held in buffers until the GPCs send a data request command to the EIUs. The GPCs request both primary and secondary data. Primary data is output only through MIA 1 on each EIU. Secondary data is output only through MIA 4 on each EIU.

During prelaunch, the orbiter's computers look at both primary and secondary data. Loss of either primary or secondary data will result in data path failure and either an engine ignition inhibit or a launch pad shutdown of all three main engines.

At T minus zero, the orbiter GPCs request both primary and secondary data from each EIU. For no failures, only primary data are looked at. If there is a loss of primary data (which can occur between the engine controller channel A electronics and the SSME SOP), the secondary data are looked at.

There are two primary written engine controller computer software programs: the flight operational program and the test operational program. The flight operational program is an on-line, real-time, process-control program that processes inputs from engine sensors; controls the operation of the engine servovalves, actuators, solenoids and spark igniters; accepts and processes vehicle commands; provides and transmits data to the vehicle; and provides checkout and monitoring capabilities. The test operational program supports engine testing. Functionally, it is similar to the flight operational program but differs with respect to implementation. The computer software programs are modular and are defined as computer program components, which consist of a data base organized into tables and 15 computer program components. During application of the computer program components, the programs perform data processing for failure detection and status to the vehicle. As system operation progresses through an operating phase, different combinations of control functions are operative at different times. These combinations within a phase are defined as operating modes.

The checkout phase initiates active control monitoring or checkout. The standby mode in this phase is a waiting mode of controller operation while active control sequence operations are in process. Monitoring functions that do not affect engine hardware status are continually active during the mode. Such functions include processing of vehicle commands, status update and controller self-test. During checkout, data and instructions can be loaded into the engine controller's computer memory. This permits updating of the software program and data as necessary to proceed with engine-firing operations or checkout operations. Also in this phase, component checkout, consisting of checkout or engine leak tests, is performed on an individual engine system component.

The start preparation phase consists of system purges and propellant conditioning, which are performed in preparation for engine start. The purge sequence 1 mode is the first purge sequence, including oxidizer system and intermediate seal purge operation. The purge sequence 2 mode is the second purge sequence, including fuel system purge operation and the continuation of purges initiated during purge sequence 1. The purge sequence 3 mode includes propellant recirculation (bleed valve operation). The purge sequence 4 mode includes fuel system purge and the indication engine is ready to enter the start phase. The engine-ready mode occurs when proper engine thermal conditions for start have been attained and other criteria for start have been satisfied, including a continuation of the purge sequence 4 mode.

The start phase covers operations involved with starting or firing the engines, beginning with scheduled open-loop operation of propellant valves. The start initiation mode includes all functions before ignition confirmed and the closing of the thrust control loop. The thrust buildup mode detects ignition by monitoring main combustion chamber pressure and verifying that closed-loop thrust buildup sequencing is in progress.

The main stage phase is automatically entered upon successful completion of the start phase. The normal control mode has initiated mixture ratio control, and thrust control is operating normally. In case of a malfunction, the electrical lock mode will be activated. In that mode, engine propellant valves are electrically held in a fixed configuration, and all control loop communications are suspended. There is also the hydraulic lockup mode, in which all fail-safe valves are deactivated to hydraulically hold the propellant valves in a fixed configuration and all control loop functions are suspended.

The shutdown phase covers operations to reduce main combustion chamber pressure and drive all valves closed to effect full engine shutdown. Throttling to minimum power level is the portion of the shutdown in progress at a programmed shutdown thrust reference level above the MPL. The valve schedule throttling mode is the stage in the shutdown sequence at which the programmed thrust reference has decreased below the MPL. Propellant valves closed is the stage in the shutdown sequence after all liquid propellant valves have been closed, the shutdown purge has been activated, and verification sequences are in progress. The fail-safe pneumatic mode is when the fail-safe pneumatic shutdown is used.

The post-shutdown phase represents the state of the SSME and engine controller at the completion of engine firing. The standby mode is a waiting mode of controller operations whose functions are identical to those of standby during checkout. It is the normal mode that is entered after completion of the shutdown phase. The terminate sequence mode terminates a purge sequence by a command from the vehicle. All propellant valves are closed, and all solenoid and torque motor valves are de-energized.

Each controller utilizes ac power provided by the MPS engine power left, ctr, right switches on panel R2.

Each controller has internal electrical heaters that provide environmental temperature control and are powered by main bus power through a remote power controller. The RPC is controlled by the main propulsion system engine cntrl htr left, ctr, right switches on panel R4. The heaters are not normally used until after main engine cutoff and are only turned on if environmental control is required during the mission.

MALFUNCTION DETECTION

There are three separate means of detecting malfunctions within the main propulsion system: the engine controllers, the caution and warning system and the GPCs.

The engine controller, through its network of sensors, has access to numerous engine operating parameters. A group of these parameters has been designated critical operating parameters, and special limits defined for these parameters are hard-wired and limit sensed within the caution and warning system. If a violation of any limit is detected, the caution and warning system will illuminate the red MPS caution and warning light on panel F7. The light will be illuminated by an MPS engine liquid oxygen manifold pressure above 249 psia; an MPS engine liquid hydrogen manifold pressure below 28 psia or above 60 psia; an MPS center, left or right helium pressure below 1,150 psia; an MPS center, left or right helium regulated pressure above 820 psia; or an MPS left, center or right helium delta pressure/delta time above 29 psia. Note that the flight crew can monitor the MPS press helium pneu, l, c, r meter on panel F7 when the switch is placed in the tank or reg position. The MPS press eng manf LO 2 , LH2 meter can also be monitored on panel F7. A number of the conditions will require crew action. For example, an MPS engine liquid hydrogen manifold pressure below the minimum setting will require the flight crew to pressurize the external liquid hydrogen tank by setting the LH2 ullage press switch on panel R2 to open , and a low helium pressure may require the flight crew to interconnect the pneumatic helium tank and the engine helium tanks using the MPS He interconnect valve switches on panel R2 for the engine helium system that is affected.

The engine controller also has a self-test feature that allows it to detect certain malfunctions involving its own sensors and control devices. For each of the three engines, a yellow main engine status left, ctr, right light (lower half) on panel F7 will be illuminated when the corresponding engine helium pressure is below 1,150 psia or regulated helium pressure is above 820 psia.

The lower half of the main engine status left, ctr, right light on panel F7 may also be illuminated by the SSME SOP (GPC- detected malfunctions). The yellow light may be illuminated due to an electronic hold, hydraulic lockup, loss of two or more command channels or command reject between the GPC and the SSME controller, or loss of both data channels from the SSME controller to the GPC of the corresponding engine. In an electronic hold for the affected SSME, loss of data from both pairs of the four fuel flow rate sensors and the four chamber pressure sensors will result in the propellant valve actuators being maintained electronically in the positions existing at the time the second sensor failed. (To fail both sensors in a pair, it is only necessary to fail one sensor.) In the case of either the hydraulic lockup or an electronic hold, all engine-throttling capability for the affected engine is lost; thus, subsequent throttling commands to that engine will not change the thrust level.

The red upper half of the main engine status left, ctr, right light on panel F7 will be illuminated if the corresponding engine's high-pressure oxidizer turbine's discharge temperature is above 1,760 degrees R, the main combustion chamber's pressure is below 1,000 psia, the high-pressure oxidizer turbopump's intermediate seal purge pressure is below 170 or above 650 psia, or the high-pressure oxidizer turbopump's secondary seal purge pressure is below 5 or above 85 psia. Because of the rapidity with which it is possible to exceed these limits, the engine controller has been programmed to sense the limits and automatically cut off the engine if the limits are exceeded. Although a shutdown as a result of violating operating limits is normally automatic, the flight crew can, if necessary, inhibit an automatic shutdown through the use of the main engine limit shut dn switch on panel C3. The switch has three positions: enable, auto and inhibit. The enable position allows only the first engine that violates operating limits to be shut down automatically. If either of the two remaining engines subsequently violates operating limits, it would be inhibited from automatically shutting down. The inhibit position inhibits all automatic shutdowns. The main engine shutdown left, ctr, right push buttons on panel C3 have spring-loaded covers (guards). When the guard is raised and the push button is depressed, the corresponding engine shuts down immediately.

The backup caution and warning processing of the orbiter GPCs can detect certain specified out-of-limit or fault conditions of the MPS. The backup C/W alarm light on panel F7 is illuminated, a fault message appears on all CRT displays, and an audio alarm sounds if the MPS engine liquid oxygen manifold pressure is zero or above 29 psia; the MPS engine liquid hydrogen manifold pressure is below 30 or above 46 psia; the MPS left, center or right helium pressure is below 1,150 psia; or the MPS regulated left, center or right helium pressure is below 680 or above 820 psia. This is identical to the parameter limit sensed by the caution and warning system; thus, the MPS red light on panel F7 will also be illuminated.

The SM alert indicator on panel F7 is illuminated, a fault message appears on all CRT displays, and an audio alarm is sounded when MPS malfunctions/conditions are detected by the SSME SOP or special systems-monitoring processing. The first four conditions are detected by the SSME SOP and are identical to those that illuminate the yellow lower light of the respective main engine status light on panel F7 due to electronic hold, hydraulic lockup, loss of two or more command channels or command reject between the GPC and the SSME controller, or loss of both data channels from the SSME controller to the orbiter GPC. The last four conditions are special systems-monitoring processing and illuminate the SM alert light on panel F7, sound an audio alarm and provide a fault message on all CRTs because of an external tank liquid hydrogen ullage pressure below 30 psia or above 46 psia or an external tank liquid oxygen ullage pressure of zero or above 29 psia. (Note that the main engine status lights on panel F7 will not be illuminated.)

ORBITER HYDRAULIC SYSTEMS

The three orbiter hydraulic systems supply hydraulic pressure to the main propulsion system for providing thrust vector control and actuating engine valves on each SSME.

The three hydraulic supply systems are distributed to the MPS TVC valves. These valves are controlled by hydraulics MPS/TVC 1, 2, 3 switches on panel R4. A valve is opened by positioning its respective switch to open. The talkback indicator above each switch indicates op or cl for open and close.

When the three MPS TVC hydraulic isolation valves are opened, hydraulic pressure actuates the engine main fuel valve, the main oxidizer valve, the fuel preburner oxidizer valve, the oxidizer preburner oxidizer valve and the chamber coolant valve. All hydraulically actuated engine valves on an engine receive hydraulic pressure from the same hydraulic system. The left engine valves are actuated by hydraulic system 2, the center engine valves are actuated by hydraulic system 1, and the right engine valves are actuated by hydraulic system 3. Each engine valve actuator is controlled by dual-redundant signals: channel A/engine servovalve 1 and channel B/engine servovalve 2 from that engine controller electronics. As a backup, all of the hydraulically actuated engine valves on an engine are supplied with helium pressure from the helium subsystem left, center and right engine helium tank supply system. In the event of a hydraulic lockup in an engine, helium pressure is used to actuate the engine's propellant valves to their fully closed position when the engine is shut down.

Hydraulic lockup is a condition in which all of the propellant valves on an engine are hydraulically locked in a fixed position. This is a built-in protective response of the MPS propellant valve actuator/control circuit. It takes effect any time low hydraulic pressure or loss of control of one or more propellant valve actuators renders closed-loop control of engine thrust or propellant mixture ratio impossible. Hydraulic lockup allows an engine to continue to thrust in a safe manner under conditions that normally would require that the engine be shut down; however, the affected engine will continue to operate at approximately the throttle level in effect at the time hydraulic lockup occurred. Once an engine is in a hydraulic lockup, any subsequent shutoff commands, whether nominal or premature, will cause a pneumatic helium shutdown. Hydraulic lockup does not affect the capability of the engine controller to monitor critical operating parameters or issue an automatic shutdown if an operating limit is out of tolerance; however, the engine shutdown would be accomplished pneumatically.

The three MPS thrust vector control valves must also be opened to supply hydraulic pressure to the six main engine TVC actuators. There are two servoactuators per SSME: one for yaw and one for pitch. Each actuator is fastened to the orbiter thrust structure and to the powerhead of one of the three SSMEs. The two actuators per engine provide attitude control and trajectory shaping by gimbaling the SSMEs in conjunction with the solid rocket boosters during first-stage ascent and without the SRBs during second-stage ascent. Each SSME servoactuator receives hydraulic pressure from two of the three orbiter hydraulic systems; one system is the primary system and the other is a standby system. Each servoactuator has its own hydraulic switching valve. The switching valve receives hydraulic pressure from two of the three orbiter hydraulic systems and provides a single source to the actuator. Normally, the primary hydraulic supply is directed to the actuator; however, if the primary system were to fail and lose hydraulic pressure, the switching valve would automatically switch over to the standby system, and the actuator would continue to function on the standby system. The left engine's pitch actuator utilizes hydraulic system 2 as the primary and hydraulic system 1 as the standby. The engine's yaw actuator utilizes hydraulic system 1 as the primary and hydraulic system 2 as the standby. The center engine's pitch actuator utilizes hydraulic system 1 as the primary and hydraulic system 3 as the standby, and the yaw actuator utilizes hydraulic system 3 as the primary and hydraulic system 1 as the standby. The right engine's pitch actuator utilizes hydraulic system 3 as the primary and hydraulic system 2 as the standby. Its yaw actuator utilizes hydraulic system 2 as the primary and hydraulic system 3 as the standby.

The hydraulic systems are distributed among the actuators and engine valves to equalize the hydraulic work load among the three systems.

The hydraulic MPS/TVC isol vlv sys1, sys2, sys3 switches on panel R4 are positioned to close during on-orbit operations to protect against hydraulic leaks downstream of the isolation valves. In addition, there is no requirement to gimbal the main engines from the stow position. During on-orbit operations when the MPS TVC valves are closed, the hydraulic pressure supply and return lines within each MPS TVC component are interconnected to enable hydraulic fluid to circulate for thermal conditioning.

MPS THRUST VECTOR CONTROL

The space shuttle ascent thrust vector control portion of the flight control system directs the thrust of the three main engines and two solid rocket boosters to control attitude and trajectory during lift-off and first-stage ascent and the main engines alone during second-stage ascent.

Ascent thrust vector control is provided by avionics hardware packages that supply gimbal commands and fault detection for each hydraulic gimbal actuator. The MPS ATVC packages are located in the three aft avionics bays in the orbiter aft fuselage and are cooled by cold plates and the Freon-21 system. The associated flight aft multiplexers/demultiplexers are also located in the aft avionics bays.

The MPS TVC command flow starts in the general-purpose computers, in which the flight control system generates the TVC position commands, and terminates at the SSME servoactuators, where the actuators gimbal the SSMEs in response to the commands. All the MPS TVC position commands generated by the flight control system are issued to the MPS TVC command subsystem operating program, which processes and disburses them to their corresponding flight aft MDMs. The flight aft MDMs separate these linear discrete commands and disburse them to ATVC channels, which generate equivalent command analog voltages for each command issued. These voltages are, in turn, sent to the servoactuators, commanding the SSME hydraulic actuators to extend or retract, thus gimbaling the main engines to which they are fastened.

Six MPS TVC actuators respond to the command voltages issued by four ATVC channels. Each ATVC channel has six MPS drivers and four SRB drivers. Each actuator receives four identical command voltages from four different MPS drivers, each located in different ATVC channels.

Each main engine servoactuator consists of four independent, two-stage servovalves, which receive signals from the drivers. Each servovalve controls one power spool in each actuator, which positions an actuator ram and the engine to control thrust direction.

The four servovalves in each actuator provide a force-summed majority voting arrangement to position the power spool. With four identical commands to the four servovalves, the actuator's force-sum action prevents a single erroneous command from affecting power ram motion. If the erroneous command persists for more than a predetermined time, differential pressure sensing activates an isolation driver, which energizes an isolation valve that isolates the defective servovalve and removes hydraulic pressure, permitting the remaining channels and servovalves to control the actuator ram spool provided the FCS channel 1, 2, 3, 4 switch on panel C3 is in the auto position. A second failure would isolate the defective servovalve and remove hydraulic pressure in the same manner as the first failure, leaving only two channels remaining.

Failure monitors are provided for each channel on the CRT and backup caution and warning light to indicate which channel has been bypassed for the MPS and/or SRB. If the FCS channel 1, 2, 3, or 4 switch on panel C3 is positioned to off, that ATVC channel is isolated from its servovalve on all MPS and SRB actuators. The override position of the FCS channel 1, 2, 3, 4 switch inhibits the isolation valve driver from energizing the isolation valve for its respective channel and provides the capability of resetting a failed or bypassed channel.

The ATVC 1, 2, 3, 4 power switch is located on panel O17. The on position enables the ATVC channel selected; off disables the channel.

Each actuator ram is equipped with transducers for position feedback to the TVC system.

The SSME servoactuators change each main engine's thrust vector direction as needed during the flight sequence. The three pitch actuators gimbal the engine up or down a maximum of 10 degrees 30 minutes from the installed null position. The three yaw actuators gimbal the engine left or right a maximum of 8 degrees 30 minutes from the installed position. The installed null position for the left and right main engines is 10 degrees up from the X axis in a negative Z direction and 3 degrees 30 minutes outboard from an engine centerline parallel to the X axis. The center engine's installed null position is 16 degrees above the X axis for pitch and on the X axis for yaw. When any engine is installed in the null position, the other engines cannot collide with it.

The minimum gimbal rate is 10 degrees per second; the maximum rate is 20 degrees per second.

There are three actuator sizes for the main engines. The piston area of the one upper pitch actuator is 24.8 square inches, its stroke is 10.8 inches, it has a peak flow of 50 gallons per minute, and it weighs 265 pounds. The piston area of the two lower pitch actuators is 20 square inches, their stroke is 10.8 inches, their peak flow is 45 gallons per minute, and they weigh 245 pounds. All three yaw actuators have a piston area of 20 square inches, a stroke of 8.8 inches and a peak flow of 45 gallons per minute and weigh 240 pounds.

HELIUM, OXIDIZER AND FUEL FLOW SEQUENCE

At T minus five hours 15 minutes, the fast-fill portion of the liquid oxygen and liquid hydrogen loading sequence begins under the control of the launch processing system.

At T minus five hours 50 minutes, the SSME liquid hydrogen chill-down sequence is initiated by the LPS. It opens the liquid hydrogen recirculation valves and starts the liquid hydrogen recirculation pumps. As part of the chill-down sequence, the liquid hydrogen prevalves are closed and remain closed until T minus 9.5 seconds.

At T minus three hours 45 minutes, the fast fill of the liquid hydrogen tank to 98 percent is complete, and a slow topping off process that stabilizes to 100 percent begins. At T minus three hours 30 minutes, the liquid oxygen fast fill is complete. At T minus three hours 15 minutes, liquid hydrogen replenishment begins and liquid oxygen replenishment begins at T minus three hours 10 minutes.

During prelaunch, the pneumatic helium supply provides pressure to operate the liquid oxygen and hydrogen prevalves and outboard and inboard fill and drain valves. The three engine helium supply systems are used to provide anti-icing purges.

When the flight crew enters the orbiter, all 10 helium supply tanks are fully pressurized to approximately 4,400 psi. The filling of the helium tanks from 2,000 psi to their full pressure begins at T minus three hours 20 minutes. This process is gradual to prevent excessive heat buildup in the supply tank. Regulated helium pressure is between 715 to 775 psi. The helium supply tank and regulated pressures are monitored on the MPS press, pneu, l, c, r meters on panel F7. The MPS press tank, reg switch positions on panel F7 select the supply or regulated pressures to be displayed on the meters. Engine helium and regulated pressures are also available on the CRT display.

When the flight crew enters the orbiter, the eight MPS He isolation A and B switches; the MPS pneumatics l eng to xovr and He isol switches; and the MPS He interconnect left, ctr, right switches on panel R2 are in the GPC position. With the switches in these positions, the eight helium isolation valves are open, and the left engine crossover and the six helium interconnect valves are closed.

At T minus 16 minutes, one of the first actions by the flight crew is to place the six MPS He isolation A and B switches and the MPS pneumatics He isol switch on panel R2 in the open position. This will not change the position of the helium isolation valves, but it inhibits LPS control of valve position.

During prelaunch, liquid oxygen from ground support equipment is loaded through the GSE liquid oxygen T-0 umbilical and passes through the liquid oxygen outboard fill and drain valve, the liquid oxygen inboard fill and drain valve and the orbiter liquid oxygen feed line manifold. The liquid oxygen exits the orbiter at the liquid oxygen feed line umbilical disconnect and enters the liquid oxygen tank in the external tank. During loading, the liquid oxygen tank's vent and relief valves are open to prevent pressure buildup in the tank due to liquid oxygen loading; and the main propulsion system propellant fill/drain LO 2 outbd and inbd switches on panel R4 are in the gnd (ground) position, which allows the LPS to control the positions of these valves as required. When liquid oxygen loading is complete, the LPS will first command the liquid oxygen inboard fill and drain valve to close. The liquid oxygen in the line between the inboard and outboard fill and drain valves is then allowed to drain back into the GSE, and the LPS commands the outboard fill and drain valve to close.

Also during prelaunch, liquid hydrogen supplied through the GSE liquid hydrogen T-0 umbilical passes through the liquid hydrogen outboard fill and drain valve, the liquid hydrogen inboard fill and drain valve and the liquid hydrogen feed line manifold. The liquid hydrogen then exits the orbiter at the liquid hydrogen feed line umbilical disconnect and enters the liquid hydrogen tank in the external tank. During loading, the liquid hydrogen tank's vent valve is left open to prevent pressure buildup in the tank due to boiloff. The main propulsion system propellant fill/drain LH 2 inbd and outbd switches on panel R4 are in the gnd position, which allows the LPS to control the position of these valves as required.

At T minus four minutes, the fuel system purge begins, followed at T minus three minutes 25 seconds by the beginning of the engine gimbal tests. During the tests, each gimbal actuator is operated through a canned profile of extensions and retractions. If all actuators function satisfactorily, the engines are gimbaled to predefined positions at T minus two minutes 15 seconds. The engines remain in these positions until engine ignition. In the predefined start positions, the engines are gimbaled in an outward direction (away from one another) so that the engine start transient will not cause the engine bells to contact one another during the start sequence.

At T minus two minutes 55 seconds, the LPS closes the liquid oxygen tank vent valve, and the tank is pressurized to 21 psi with GSE-supplied helium. The liquid oxygen tank's pressure can be monitored on the MPS press eng manf LO 2 meter on panel F7 as well as on the CRT. The 21-psi pressure corresponds to a liquid oxygen engine manifold pressure of 105 psia.

At T minus one minute 57 seconds, the LPS closes the liquid hydrogen tank's vent valve, and the tank is pressurized to 44 psia with GSE-supplied helium. The pressure is monitored on the MPS press eng manf LH 2 meter on panel F7 as well as on the CRT display. A liquid hydrogen tank pressure of 44 psia corresponds to a liquid hydrogen engine manifold pressure of 44.96 psia.

At T minus 31 seconds, the onboard redundant set launch sequence is enabled by the LPS. From this point on, all sequencing is performed by the orbiter GPCs in the redundant set, based on the onboard clock time. The GPCs still respond, however, to hold, resume count and recycle commands from the LPS.

At T minus 16 seconds, the GPCs begin to issue arming commands for the SRB ignition pyro initiator controllers, the hold-down release PICs and the T-0 umbilical release PICs.

At T minus 9.5 seconds, the engine chill-down sequence is complete, and the GPCs command the liquid hydrogen prevalves to open (the liquid oxygen prevalves are open during loading to permit engine chill-down). The main propulsion system LO2 and LH2 prevalve left, ctr, right switches on panel R4 are in the GPC position.

At T minus 16 seconds, helium flows out of the nine helium supply tanks through the helium isolation valves, regulators and check valves and enters the engine at the inlet of the pneumatic control assembly. The PCA is a manifold containing solenoid valves that control and direct helium pressure under the control of the engine controller to perform various essential functions. The valves are energized by discrete on/off commands from the output electronics of the engine controller. One essential function from T minus 6.6 seconds to main engine cutoff plus six seconds is the purging of the high-pressure oxidizer turbopump's intermediate seal cavity. This cavity is between two seals, one of which contains the hot, fuel-rich gas in the oxidizer turbine. The other seal contains the liquid oxygen in the oxidizer turbopump. Leakage through one or both of the seals and mixing of the propellants could result in a catastrophic explosion. Continuous overload purging of this area prevents the propellants from mixing as they are dumped overboard through drain lines. The PCA also functions as an emergency backup for closing the engine propellant valves with helium pressure. In a normal engine shutdown, the engine propellant valves are hydraulically actuated.

At T minus 6.6 seconds, the GPCs issue the engine start command, and the main fuel valve in each engine opens. Between the opening of the main fuel valve and MECO, liquid hydrogen flows out of the external tank/orbiter liquid hydrogen disconnect valves into the liquid hydrogen feed line manifold. From this manifold, liquid hydrogen is distributed to the engines through the three engine liquid hydrogen feed lines. In each line, liquid hydrogen passes through the prevalve and enters the main engine at the inlet to the low-pressure fuel turbopump. In the engine, the liquid hydrogen cools various engine components and in the process is converted to gaseous hydrogen. The majority of this gaseous hydrogen is burned in the engine; the smaller portion is directed back to the external tank to maintain liquid hydrogen tank pressure. The flow of gaseous hydrogen back to the external tank begins at the turbine outlet of the LPFT. Gaseous hydrogen tapped from this line first passes through two check valves and then splits into two paths, each containing a flow control orifice. One of these paths also contains a valve normally controlled by one of three pressure transducers located in the liquid hydrogen tank.

When the GPCs issue the engine start command, the main oxidizer valve in each engine also opens. Between the opening of the main engine oxidizer valve and MECO, liquid oxygen flows out of the external tank and through the external tank/orbiter liquid oxygen umbilical disconnect valves into the liquid oxygen feed line manifold. From this manifold, liquid oxygen is distributed to the engines through the three engine liquid oxygen feed lines. In each line, liquid oxygen passes through the prevalve and enters the main engine at the inlet to the low-pressure oxidizer turbopump. In the engine, a small portion of the liquid oxygen is diverted into the oxidizer heat exchanger. In the heat exchanger, heat generated by the high-pressure oxidizer turbopump is used to convert liquid oxygen into gaseous oxygen, which is directed back to the external tank to maintain oxidizer tank pressure. The flow of gaseous oxygen back to the external tank begins at the outlet of the heat exchanger. From this point, gaseous oxygen passes through a check valve and then splits into two paths, each containing a flow control orifice. One of these paths also contains a valve that normally is controlled by one of three pressure transducers located in the liquid oxygen tank. Downstream of the two flow control orifices and the pressure control valves, the gaseous oxygen lines empty into the orbiter gaseous oxygen pressurization manifold. This single line exits the orbiter at the gaseous oxygen pressurization disconnect and passes through the orbiter/external tank gaseous oxygen umbilical into the top of the liquid oxygen tank.

At T minus 6.6 seconds, if the PIC voltages are within limits and all three engine controllers are indicating engine ready, the GPCs issue the engine start commands to the three main engines. If the PIC conditions are not met in four seconds, the engine start commands are not issued, and the GPCs proceed to a countdown hold.

If all three SSMEs reach 90 percent of their rated thrust by T minus three seconds, then at T minus zero the GPCs will issue the commands to fire the SRB ignition PICs, the hold-down release PICs and the T-0 umbilical release PICs. Lift-off occurs almost immediately because of the extremely rapid thrust buildup of the SRBs. The three seconds to T minus zero allow the vehicle base bending loads to return to minimum by T minus zero.

If one or more of the three main engines do not reach 90 percent of their rated thrust at T minus three seconds, all SSMEs are shut down, the SRBs are not ignited, and a pad abort condition exists.

Beginning at T minus zero, the SSME gimbal actuators, which were locked in their special preignition positions, are first commanded to their null positions for SRB start and then allowed to operate as needed for thrust vector control.

Between lift-off and MECO, as long as the SSMEs perform nominally, all MPS sequencing and control functions are executed automatically by the GPCs. During this period, the flight crew monitors MPS performance; backs up automatic functions, if required; and provides manual inputs in the event of MPS malfunctions.

During ascent, the liquid hydrogen tank's pressure is maintained between 33 and 35 psig by the orifices in the two lines and the action of the flow control valve. There are three such systems, one for each SSME. When the pressure in the liquid hydrogen tank reaches 35 psig, the valve closes. It opens when the pressure drops below 33 psig. Tank pressure greater than 38 psia will cause the tank to relieve through the tank vent valve. If tank pressure falls below 33 psia, the flight crew positions the MPS LH 2 ullage press switch on panel R2 to open . This allows the three flow control valves to go to the full-open position. Normally, the MPS LH 2 ullage press switch is in the auto position. Downstream of the two flow control orifices and the flow control valves, the gaseous hydrogen line empties into the gaseous hydrogen pressurization manifold. This single line then exits the orbiter at the gaseous hydrogen umbilical and enters the top of the liquid hydrogen tank. During ascent, the liquid oxygen tank's pressure is maintained between 20 and 22 psig by the orifices in the two lines and the action of the flow control valve. When the pressure in the tank reaches 22 psig, the valve closes. It opens when pressure drops below 20 psig. A pressure greater than 25 psig will cause the tank to relieve through its vent and relief valve.

The SSME thrust level depends on the flight: it may be 100 percent or 104 percent for some missions involving heavy payloads or may require the maximum thrust setting of 109 percent for emergency situations. The initial thrust level normally is maintained until approximately 31 seconds into the mission, when the GPCs throttle the engines to a lower thrust to minimize structural loading while the orbiter is passing through the region of maximum aerodynamic pressure. This normally occurs around 63 seconds, mission elapsed time. At approximately 65 seconds, the engines are once again throttled to the appropriate higher percent and remain at that setting for a normal mission until 3-g throttling is initiated.

The solid rocket boosters burn out at approximately two minutes, mission elapsed time, and are separated from the orbiter by a GPC command sent via the mission events controller and by the SRB separation PICs. The flight crew can initiate SRB separation manually if the automatic sequence fails; however, the manual separation sequence does not bypass the separation sequence logic circuitry.

Beginning at approximately seven minutes 40 seconds, mission elapsed time, the engines are throttled back to maintain vehicle acceleration at 3 g's or less. Three g's is an operational limit devised to prevent physical stresses on the flight crew. Approximately eight seconds before main engine cutoff, the engines are throttled back to 65 percent.

Although MECO is based on the attainment of a specified velocity, the engines can also be shut down due to the depletion of liquid oxygen or liquid hydrogen before the specified velocity of MECO is reached. Liquid oxygen depletion is sensed by four sensors in the liquid oxygen feed line manifold. Liquid hydrogen depletion is sensed by four sensors in the bottom of the liquid hydrogen tank. If any two of the four sensors in either system indicate a dry condition, the GPCs will issue a MECO command to the engine controller.

Once MECO has been confirmed, the GPCs execute the external tank separation sequence. The sequence takes approximately 18 seconds and includes arming the external tank separation PICs, closing the liquid oxygen and liquid hydrogen prevalves, firing the external tank tumble system pyrotechnic, closing the liquid hydrogen and liquid oxygen feed line 17-inch disconnect valves, gimbaling the SSMEs to the MPS propellant dump position (full down), turning the external tank signal conditioners' power off (deadfacing), firing the umbilical unlatch pyrotechnics, and retracting the umbilical plates hydraulically.

At this point, the computers check for external tank separation inhibits. If the vehicle's pitch, roll and yaw rates are not less than 0.2 degree per second, automatic external tank separation is inhibited. If these conditions are met, the GPCs issue the commands to the external tank separation pyrotechnics. In crew-initiated external tank separation or return-to-launch-site aborts, the inhibits are overriden.

At separation, the orbiter begins a reaction control system minus Z translation separation maneuver to move it away from the external tank. This maneuver takes approximately 13 seconds and results in a negative Z-delta component of approximately 11 feet per second.

After MECO occurs (whether because the specified velocity is attained or the liquid oxygen or liquid hydrogen is depleted) and before external tank separation, the GPCs isolate the orbiter liquid hydrogen feed line from the external tank by closing the two liquid hydrogen 17-inch disconnect valves (one on each side of the separation interface) and the two liquid oxygen 17-inch disconnect valves (one on each side of the separation interface). At orbiter/external tank separation, the gaseous oxygen and gaseous hydrogen feed lines are sealed at the umbilicals by the self-sealing quick disconnects.

The MPS pneumatic control assembly on each main engine provides an emergency backup method of closing the engine propellant valves pneumatically using helium pressure. The normal engine shutdown of the engine propellant valves is by hydraulic actuation.

At MECO, the GPCs open the liquid oxygen feed line relief isolation valve, allowing any pressure buildup generated by oxidizer trapped in the orbiter liquid oxygen feed line manifold to be vented overboard through the relief valve provided the main propulsion system feedline rlf isol LH2 switch on panel R4 is in the GPC position. The GPCs also open the liquid hydrogen feed line relief isolation valve, and any pressure buildup from fuel trapped in the orbiter liquid hydrogen feed line manifold is vented overboard through the relief valve provided the main propulsion system feedline rlf isol LH 2 switch on panel R4 is in the GPC position.

At MECO, the pneumatic control assembly for each engine performs a 16-second purge of the engine preburner oxidizer domes and a two-second postcharge of the pogo accumulator. This purge ensures that no residual propellant remains in these areas to cause an unsafe condition and prevents a water hammer effect in the liquid oxygen manifolds of the main engines. This helium usage and the purge of the high-pressure oxidizer turbopump's intermediate seal cavity can be observed on the MPS helium l, c, r meters on panel F7 and are also available on the CRT.

Ten seconds after main engine cutoff, the RTLS liquid hydrogen dump valves are opened for 30 seconds to ensure that the liquid hydrogen manifold pressure does not result in operation of the liquid hydrogen feed line relief valve.

After the completion of the 16-second purge, the GPCs interconnect the pneumatic helium and engine helium supply system by opening the three out interconnect valves provided the MPS He interconnect left, center, right switches on panel R2 are in the GPC position. This connects all 10 helium supply tanks to the common manifold and ensures sufficient helium is available to perform the liquid oxygen and liquid hydrogen propellant dumps, which are required after external tank separation.

After external tank separation, approximately 1,700 pounds of propellant is still trapped in the SSMEs and an additional 3,700 pounds of propellant remains trapped in the orbiter's MPS feed lines. This 5,400 pounds of propellant represents an overall center-of-gravity shift for the orbiter of approximately 7 inches. Non-nominal center-of-gravity locations can create major guidance problems during re-entry. The residual liquid oxygen, by far the heavier of the two propellants, poses the greatest impact on center-of-gravity travel. The greatest hazard from the trapped liquid hydrogen occurs during re-entry, when any liquid or gaseous hydrogen remaining in the propellant lines may combine with atmospheric oxygen to form a potentially explosive mixture. In addition, if the trapped propellants are not dumped overboard, they will sporadically outgas through the orbiter liquid oxygen and liquid hydrogen feed line relief valves, causing vehicle accelerations of such a low level that they cannot be sensed by onboard guidance, yet represent a significant source of navigation error when applied over an entire mission. Outgassing propellants are also a potential source of contamination of scientific experiments contained in the payload bay.

Approximately 18 seconds after MECO occurs, the external tank separates from the orbiter. Approximately 102 seconds later, at MECO plus two minutes, the first thrusting period of the orbital maneuvering system begins. Coincident with the start of the OMS-1 thrusting, the GPCs automatically initiate the liquid oxygen dump provided the MPS prplt dump sequence LO2 switch on panel R2 is in the GPC position. The computers command the two liquid oxygen manifold repressurization valves to open (the main propulsion system manf press LO 2 switch on panel R4 must be in the GPC position), command each engine controller to open its SSME main oxidizer valve, and command the three liquid oxygen prevalves to open (the main propulsion system LO 2 prevalves left, ctr, right switch must be in the GPC position). The liquid oxygen trapped in the feed line manifolds is expelled under pressure from the helium subsystem through the nozzles of the SSMEs. If the main propulsion system manf press LO2 switch on panel R4 is left in the GPC position, the pressurized liquid oxygen dump continues for 90 seconds. At the end of this period, the GPCs automatically terminate the dump by closing the two liquid oxygen manifold repressurization valves, wait 30 seconds and then command the engine controllers to close their SSME main oxidizer valve. The three liquid oxygen prevalves remain open.

If necessary, the crew can perform the liquid oxygen dump manually utilizing the start and stop positions of the MPS prplt dump sequence LO 2 switch on panel R2. When the liquid oxygen dump is initiated manually, all valve opening and closing sequences are still automatic. Positioning the MPS prplt dump sequence LO 2 switch to start causes the GPCs to immediately begin commanding all of the required valves to open automatically and in the proper sequence. The liquid oxygen dump will continue as long as the switch is in the start position, but the pressurized portion with the two liquid oxygen manifold repressurization valves open is still limited to 90 seconds. Placing the switch in the stop position causes the GPCs to begin commanding all of the required valves to close automatically and in the proper sequence. The earliest a manual liquid oxygen dump can be performed is MECO plus 20 seconds since the SSMEs require a cool-down of at least 20 seconds after MECO.

The GPC software's MPS dump sequence automatically initiates the liquid oxygen dump at one time only-the beginning of the OMS-1 thrusting period. If the MPS prplt dump sequence LO 2 switch on panel R2 is not in the GPC position at that time, the liquid oxygen dump must be initiated manually. In addition, once the liquid oxygen dump has been initiated and the MPS prplt dump sequence LO 2 switch is placed in the stop position, the GPCs no longer monitor any of the positions of this switch. For this reason, the liquid oxygen dump cannot be reinitiated, manually or automatically.

Simultaneously with the liquid oxygen dump, the GPCs automatically initiate the MPS liquid hydrogen dump provided the MPS prplt dump sequence LH2 switch on panel R2 is in the GPC position. The GPCs command each engine controller to command a 10-second helium purge of its SSME's fuel lines downstream of the main engine fuel valves, command the liquid hydrogen manifold repressurization valve to open provided the main propulsion system manf press LH 2 switch on panel R4 is in the GPC position, and command the two liquid hydrogen fill and drain valves (inboard and outboard) to open.

The liquid hydrogen trapped in the orbiter feed line manifold is expelled overboard under pressure from the helium subsystem through the liquid hydrogen fill and drain valves for six seconds. Then the inboard fill and drain valve is closed; the three liquid hydrogen prevalves are opened; and liquid hydrogen flows through the engine bleed valves into the orbiter MPS, through the topping valve, between the inboard and outboard fill and drain valves, and overboard through the outboard fill and drain valve for approximately 88 seconds. The GPCs automatically terminate the dump by closing the two liquid hydrogen manifold repressurization valves and 30 seconds later closing the liquid hydrogen topping and outboard fill and drain valves.

If necessary, the flight crew can perform the liquid hydrogen dump manually utilizing the start and stop positions of the MPS prplt dump sequence LH 2 switch on panel R2. When the liquid hydrogen dump is initiated manually, all valve opening and closing sequences are still automatic. Placing the MPS prplt dump sequence switch in the start position causes the GPCs immediately to begin commanding all the required valves to open automatically and in the proper sequence. The liquid hydrogen dump continues as long as the switch is in the start position, but the pressurized portion of the dump with the two liquid hydrogen manifold repressurization valves open is still limited to 88 seconds. Placing the switch in the stop position causes the GPCs to begin commanding all of the required valves to close automatically and in the proper sequence.

At the end of the liquid oxygen and liquid hydrogen dumps, the GPCs close the helium out interconnect valves and all of the supply tank isolation valves provided the MPS He isolation left ctr, right A and B; pneumatic He isol; and He interconnect left, ctr, right switches on panel R2 are in the GPC position. After the dumps are complete, the space shuttle main engines are gimbaled to their entry positions with the engine nozzles moved inward (toward one another) to reduce aerodynamic heating.

Approximately 19 minutes into the mission and after the MPS liquid oxygen and liquid hydrogen dumps, the flight crew initiates the procedure for vacuum inerting the orbiter's liquid oxygen and liquid hydrogen lines. Vacuum inerting allows any traces of liquid oxygen or liquid hydrogen remaining after the propellant dumps to be vented into space.

The liquid oxygen vacuum inerting is accomplished by opening the liquid oxygen inboard and the outboard fill and drain valves. They are opened by placing the main propulsion system propellant fill/drain LO 2 outbd, inbd switch on panel R4 to the open position.

For liquid hydrogen vacuum inerting, the liquid hydrogen inboard and outboard fill and drain valves are opened by placing the main propulsion system propellant fill/drain LH 2 outbd, inbd switch on panel R4 to open. The external tank gaseous hydrogen pressurization manifold also is vacuum inerted by opening the hydrogen pressurization line vent valve by placing the main propulsion system H 2 line vent switch on panel R4 to open.

Helium for actuating the valves is provided by the two pneumatic helium isolation valves by placing the MPS pneumatic He isol switch on panel R2 to open. These isolation valves are closed by the GPCs at the end of the MPS liquid hydrogen dump. If additional helium is required to open and close the fill and drain valves, it can be obtained by opening the helium out interconnect valves by placing the MPS He interconnect left, ctr, right switches on panel R2 in the in close/out open position. These valves also are closed by the GPCs at the end of the MPS liquid hydrogen dump.

The liquid oxygen and liquid hydrogen lines are inerted simultaneously. Approximately 30 minutes is allowed for vacuum inerting. At the end of the 30 minutes, the flight crew closes the liquid oxygen outboard fill and drain valve by placing the main propulsion system propellant fill/drain LO2 switch on panel R4 to close . The inboard fill and drain valve is left open. To conserve electrical power after the completion of the liquid oxygen vacuum inerting sequence, the main propulsion system propellant fill/drain LO 2 outbd, inbd switch on panel R4 is placed in the gnd position. This position removes power from the opening and closing solenoids of the corresponding valves; and because they are pneumatically actuated, the valves remain in their last commanded position. At the end of the same 30-minute period, the liquid hydrogen outboard fill and drain valve and the hydrogen pressurization line vent valve are closed by positioning the main propulsion system propellant fill/drain LH2 outbd switch and the main propulsion system H 2 press line vent switch on panel R4 to close . The liquid hydrogen inboard fill and drain valve is left open. The main propulsion system propellant fill/drain LH 2 inbd, outbd and H 2 press line vent switches on panel R4 are positioned to gnd to conserve power. The hydrogen pressurization vent line valve is electrically activated; however, it is normally closed (spring loaded to the closed position), and removing power from the valve solenoid leaves the valve closed.

After vacuum inerting, the helium isolation valves and interconnect valves (if they were used) are closed by placing the MPS He isolation pneumatics He isol switch on panel R2 to close and the He interconnect left, ctr, right switches on panel R2 to GPC . This ensures that the helium supply tanks are isolated from any leakage in the downstream lines during orbital operations.

The electrical power to each engine controller and engine interface unit is turned off by positioning the MPS engine power left, ctr, right switches on panel R2 to off ; and the engine controller heaters are turned on by positioning the main propulsion system engine cntlr htr, left, ctr, right switches on panel R4 to auto .

During the early portion of the entry time line, the propellant feed line manifolds and the external tank pressurization lines are repressurized with helium from the helium subsystem. This prevents atmospheric contamination from being drawn into the manifolds and feed lines during entry. Removing contamination from the manifolds or feed lines can be a long and costly process since it involves disassembly of the affected part. Manifold repressurization is an automatic sequence performed by the GPCs.

After the orbital maneuvering system engines have been fired for deorbit and the orbiter begins to sense the presence of atmosphere, the GPCs start another vacuum inerting sequence. The liquid oxygen and liquid hydrogen prevalves that were left open at the end of the liquid oxygen and liquid hydrogen dump sequences remained open during the entire mission. Similarly, the liquid oxygen and liquid hydrogen inboard fill and drain valves that were left open at the end of the manual vacuum inerting sequence remained open during the entire mission. As re-entry begins, the left engine's helium isolation valve B and the pneumatic helium isolation valves are opened providing the MPS He isolation left B and the MPS pneumatic He isol switches on panel R2 are in the GPC position; the left engine's pneumatic crossover valve and in interconnect valve are opened; and the center and right engines' out interconnect valves are opened providing the MPS pneumatics l eng He xovr and MPS He interconnect left, ctr and right switches on panel R2 are in the GPC position. Also, the MPS liquid hydrogen topping valve, outboard fill and drain valves, and inboard and outboard RTLS drain valves are opened providing the propellant fill/drain LO 2 and LH2 outbd and inbd switches are in the gnd position. As orbiter re-entry continues, its velocity decreases. When the velocity drops below 20,000 feet per second, the liquid oxygen outboard fill and drain valve opens.

This vacuum inerting continues until the orbiter's velocity drops below 4,500 feet per second (between 110,000 and 130,000 feet altitude depending on the re-entry trajectory). Then the MPS liquid oxygen and liquid hydrogen outboard fill and drain valves, the liquid hydrogen inboard and outboard RTLS drain valves, and the liquid oxygen prevalves are closed; the MPS liquid oxygen and liquid hydrogen manifold repressurization valves and the MPS helium blowdown supply valves are opened; and a 650-second timer is started. This provides a positive pressure in the liquid oxygen and liquid hydrogen manifolds and in the aft fuselage and the OMS/RCS pods and prevents contamination. The 650-second timer runs out approximately one minute after touchdown. After the timer expires, the purge of the aft fuselage and OMS/RCS pods is terminated when the MPS helium supply blowdown valves are closed. The manifold repressurization continues until the ground crews install the throat plugs in the main engine nozzles.

If MECO is preceded by an RTLS abort, the subsequent MPS liquid oxygen dump will begin 10 seconds after the external tank separation command is issued, and the liquid hydrogen dump will begin simultaneously. The liquid oxygen and liquid hydrogen dumps are initiated and terminated automatically by the GPCs regardless of the positions of the MPS prplt dump sequence LO2 and LH2 switches on panel R2.

During an RTLS abort, liquid oxygen initially is dumped through the SSMEs and 30 seconds later via the liquid oxygen fill and drain valves. This dump is performed without helium pressurization and relies on the self-boiling of the trapped liquid.

In the RTLS liquid oxygen dump, the GPCs terminate the dump whenever the orbiter's velocity drops below 3,800 feet per second. The liquid oxygen is dumped through the nozzles of the main engines; however, each engine is gimbaled to the entry position rather than the normal dump position. The liquid oxygen feed line manifold is not pressurized in this mode, and the two liquid oxygen manifold repressurization valves remain closed throughout the entire dump. The liquid oxygen system is repres surized when the 3,800-feet-per-second velocity is attained, and repressurization continues as in a nominal entry. The main propulsion system prevalves LO2, left, ctr, right switches on panel R4 are in the GPC position, and the GPCs command the engine controllers to open each engine main oxidizer valve for the dump.

In the RTLS mode, the liquid hydrogen dump is initiated and terminated automatically by the GPCs simultaneously with the liquid oxygen dump regardless of the position of the MPS prplt dump sequence LH 2 switch on panel R2. The two RTLS dump valves and the two RTLS manifold repressurization valves are opened, and the liquid hydrogen trapped in the feed line manifold is expelled under pressure from the helium subsystem for 80 seconds through a special opening on the port side of the orbiter between the wing and the OMS/RCS pod. After 80 seconds, the liquid hydrogen fill and drain valves are opened, resulting in vacuum inerting of residual liquid hydrogen through bulk boiling. The GPCs terminate the liquid hydrogen dump and vacuum inerting automatically when the orbiter reaches the 3,800-feet-per-second velocity. At that time, the inboard and outboard RTLS dump valves, the inboard and outboard fill and drain valves, and the two RTLS manifold repressurization valves are closed. The liquid hydrogen system is repressurized after an RTLS liquid hydrogen dump, and repressurization continues as in a nominal entry.

CONTRACTORS. The Rocketdyne Division of Rockwell International, Canoga Park, Calif., is the prime contractor for the space shuttle main engines. Other contractors include Aeroflex Laboratories, Plainview, N.Y. (MPS vibration mounts); Airite Division, Sargent Industries, El Segundo, Calif. (MPS surge pressure receiver); Ametek Calmec, Pico Rivera, Calif. (1.5-inch and 2-inch liquid oxygen and liquid hydrogen shutoff valve, 4-inch liquid hydrogen disconnect and 2-inch gaseous hydrogen/gaseous oxygen disconnect); Ametek Straza, El Cajon, Calif. (8-inch liquid hydrogen/liquid oxygen fill and drain, 2- and 4-inch liquid hydrogen recirculation lines, high-point bleed line manifold and gimbal joint); Arrowhead Products, division of Federal Mogul, Los Alamitos, Calif. (12- to 17-inch-diameter liquid oxygen and liquid hydrogen feed lines and flexible purge gas connector); Astech, Santa Ana, Calif. (MPS heat shield); Brunswick, Lincoln, Neb. (17.3- and 4.7-cubic-foot capacity helium tanks); Brunswick-Circle Seal, Anaheim, Calif. (helium check valves, gaseous oxygen and gaseous hydrogen 1-inch helium pressurization line, 0.375-inch liquid hydrogen relief valve and engine isolation check valves); Brunswick-Wintec, El Segundo, Calif. (helium filter); Coast Metal Craft, Compton, Calif. (metal flex hose); Conrac Corp., West Caldwell, N.J. (engine interface unit); Consolidated Controls, El Segundo, Calif. (oxygen pressure primary flow control valve and hydraulic valve, hydrogen/oxygen pressurant flow control valves, 20-psi helium regulator, 850-psi helium relief valve and 750-psi helium regulator); Fairchild Stratos, Manhattan Beach, Calif. (12-inch prevalves, 1.5-inch liquid oxygen disconnect, 8-inch liquid oxygen and liquid hydrogen fill and drain valves, and gaseous nitrogen and gaseous hydrogen disconnects); Gulton Industries, Costa Mesa, Calif. (pogo pressure transducer); K-West, Westminister, Calif. (liquid oxygen and liquid hydrogen external tank ullage pressure signal conditioner, MPS differential pressure transducer and electronics propellant head pressure); Megatek, Van Nuys, Calif. (MPS line flange cryo seals); Moog Inc., East Aurora, N.Y. (main engine gimbal actuators); Parker Hannifin Corp., Irvine, Calif. (1-inch relief isolation valves, pogo check valves, 17-inch liquid hydrogen and liquid oxygen disconnects, 8-inch liquid oxygen and liquid hydrogen disconnects, and liquid oxygen and liquid hydrogen relief valves); Simmonds Precision Instruments, Vergennes, Vt. (liquid oxygen and liquid hydrogen point sensors and electronics); Sterer Engineering, Los Angeles, Calif. (main engine hydraulic solenoid shutoff valve); Whittaker Corp., North Hollywood, Calif. (750-/250-psi helium regulator); Wright Components Inc. Clifton Springs, N.J. (two-way pneumatic solenoid valve, three-way helium solenoid valve and hydraulic latching solenoid valve).

ORBITER/EXTERNAL TANK SEPARATION SYSTEM

The orbiter/external tank separation system consists of the oxygen and hydrogen umbilical disconnects located at the lower left and right aft fuselage, one forward structural attach point just aft of the nose landing gear doors and two structural attach points located in the orbiter/external tank umbilical disconnect cavities. An umbilical retraction system retracts the orbiter umbilicals within the orbiter aft fuselage, and umbilical doors close over each of the umbilical cavities after separation.

The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three space shuttle main engines. The respective 17-inch disconnects also provide the capability for external tank fill and drain of oxygen and hydrogen through the orbiter main propulsion system and the T-0 umbilicals. The liquid hydrogen interface between the orbiter and the ground storage tank is provided by a T-0 umbilical located on the left side of the aft fuselage. The liquid oxygen interface between the orbiter and the ground storage tank is provided by a T-0 umbilical on the right side of the aft fuselage.

17-INCH DISCONNECT

Each mated pair of 17-inch disconnects contains two flapper valves, one on the orbiter side of the interface and one on the external tank side of the interface. Both valves in each disconnect pair are opened to permit propellant flow between the orbiter and the external tank. Before the separation of the external tank, both valves in each mated pair of disconnects are commanded closed by pneumatic (helium) pressure from the main propulsion system. The closure of both valves in each disconnect pair prevents propellant discharge from the external tank or orbiter at separation. Valve closure on the orbiter side of each disconnect also prevents contamination of the orbiter main propulsion system during landing and ground operations.

Inadvertent closure of either valve in a 17-inch disconnect during space shuttle main engine thrusting would stop propellant flow from the external tank to all three main engines. Catastrophic failure of the main engines and external tank feed lines would result.

To prevent inadvertent closure of the 17-inch disconnect valves during the main engine thrusting, a latch mechanism was added in the orbiter half of the disconnects. The latch mechanism provides a mechanical backup to the normal fluid-induced-open forces. The latch is mounted on a shaft in the flowstream so it overlaps both flappers and obstructs closure for any reason.

In preparation for external tank separation, both valves in each 17-inch disconnect are commanded closed. Pneumatic (helium) pressure from the main propulsion system causes the latch actuator to rotate the latch shaft in each orbiter 17-inch disconnect 90 degrees, thus freeing the flapper valves to close as required for external tank separation.

If the latch pneumatic actuator malfunctions, a backup mechanical separation capability is provided. When the orbiter umbilical initially moves away from the external tank umbilical, the mechanical latch disengages from the external tank flapper valve and permits the orbiter disconnect flapper to toggle the latch. This action permits both flappers to close.

During ground mating of the external tank to the orbiter, the latch engagement mechanism in each 17-inch disconnect provides a go/no-go verification that flapper angle rigging is within stability limits. Misrigged flappers will prevent full engagement of latch. The angle of each flapper in each disconnect is still carefully rigged within specific tolerances to assure basic stability independently of the latch safety feature.

EXTERNAL TANK SEPARATION SYSTEM

The external tank is separated from the orbiter at three structural attach points. Separation from the orbiter occurs before orbit insertion and is automatically controlled by the orbiter's general-purpose computers. External tank separation can be manually initiated by the flight crew using the same jettison circuits as the automatic sequence. Separation is controlled by the ET separation auto, man switch on panel C3 and the sep push button on panel C3. In the auto position, the onboard GPCs initiate separation. To manually initiate separation, the ET separation switch is positioned to man and the sep push button is depressed.

The forward structural attachment consists of a shear bolt unit mounted in a spherical bearing. The bolt separates at a break area when two pressure cartridges are initiated. The pressure from one or both cartridges drives one of a pair of pistons to shear the bolt, with the second piston acting as a hole plugger to fill the cavity left by the sheared bolt. A centering mechanism rotates the unit from the displacement position to a centered position, aligning the bearing flush with the adjacent thermal protection system mold line.

The aft structural attachment consists of two special bolts and pyrotechnically actuated frangible nuts that attach the external tank strut hemisphere to the orbiter's left- and right-side cavities. At separation the frangible nuts are split by a booster cartridge initiated by a detonator cartridge. The attach bolts are driven by the separation forces and a spring into a cavity in the tank strut. The frangible nut, cartridge fragments and hot gases are contained within a cover assembly, and a hole plugger isolates the fragments in the container.

The aft separation involves right and left umbilical assemblies. Each assembly contains three dual-detonator frangible nut and bolt combinations that hold the orbiter and external tank umbilical plates together during mated flight. Each bolt has a retraction spring that, after release of the nut, retracts the bolt to the external tank side of the interface. On the orbiter side, each frangible nut and its detonators are enclosed in a debris container that captures nut fragments and hot gases generated by the operation of the detonators, either of which will fracture the nut.

The right aft umbilical assembly consists of an electrical disconnect, the gaseous oxygen 2-inch pressurization disconnect used for pressurization of the external tank's oxygen tank and the 17-inch liquid oxygen disconnect.

The left aft umbilical assembly consists of an electrical disconnect plate, the gaseous hydrogen 2-inch pressurization disconnect used for pressurization of the external tank's hydrogen tank, the 4-inch recirculation disconnect used during prelaunch to precondition the main engine and the 17-inch liquid hydrogen disconnect.

After release of the three frangible nuts and bolts at each aft umbilical, three lateral support arms at each orbiter umbilical plate hold the plates in the lateral position when the external tank separates from the umbilical plates. Each 17-inch disconnect has been commanded closed. The orbiter umbilical plates are retracted inside the orbiter aft fuselage approximately 2.5 inches by three hydraulic actuators and locked to permit closure of the umbilical doors in the bottom of the aft fuselage. Hydraulic system 1 source pressure is supplied to one actuator at each umbilical, hydraulic system 2 source pressure is supplied to the second actuator at each umbilical, and hydraulic system 3 source pressure is supplied to a third actuator at each umbilical.

The retraction of each umbilical disconnects the external tank and orbiter electrical umbilical in the first 0.5 of an inch of travel and releases any fluids trapped between the 17-inch disconnect flappers.

ORBITER UMBILICAL DOORS

An electromechanical actuation system on each umbilical door closes the left and right umbilical cavities after the external tank is jettisoned and the umbilical plates retracted inside the orbiter's aft fuselage. Each umbilical door is approximately 50 inches square.

The doors are held in the full-open position by two centerline latches, one forward and one aft. They are opened before the mating of the orbiter to the external tank in the Vehicle Assembly Building.

The orbiter umbilical doors normally are controlled by the flight crew with switches on panel R2. In return-to-launch-site aborts, the doors are controlled automatically. The ET umbilical door mode switch on panel R2 positioned to GPC enables automatic control of the doors. The GPC/man position enables manual flight crew control of the doors.

The ET umbilical door centerline latch switch on panel R2 positioned to gnd permits ground control of the door centerline latches during ground turnaround operations. The stow position, enables flight crew manual control of the door centerline latches. The talkback indicator above the switch indicates sto when the door centerline latches are stowed, which permits closure of the doors, and barberpole when the latches are latched or the doors are in transit.

The ET umbilical door left and right latch, off, release switches on panel R2 are used by the flight crew to unlatch the corresponding centerline latches during normal operations. Positioning the respective switch to release provides electrical power to redundant ac reversible motors which operate an electromechanical actuator for each centerline latch that causes the latch to rotate and retract the latch blade flush with the reusable thermal protection system mold line. It takes approximately six seconds for the latches to complete their motion. The talkback indicator above the respective switch indicates rel when the corresponding latches are released. The latch position of each switch is used during ground turnaround operations to latch the respective door open, and the talkback indicator indicates lat when the latches have latched the doors in the open position. The talkback indicators indicate barberpole when the latches are in transit. The off position of the switches removes power from the motors, which stops the latches.

The ET umbilical door left and right , open , off , latch switches on panel R2 normally are used by the flight crew to close the umbilical doors. Positioning the switches to close provides electrical power to redundant ac reversible motors, which position the doors closed through a system of bellcranks and push rods. It takes approximately 24 seconds for the doors to close; and when they are within 2 inches of the closed position, ready-to-latch indicators activate the door uplatch system. Three uplatch hooks for each door engage three corresponding rollers near the outboard edge of the door and lock the door in preparation for entry. The motors are automatically turned off. The talkback indicator above the respective switch indicates cl when two of the three ready-to-latch switches for that door have sensed door closure. The open position of the switches is used during ground turnaround operations to open the doors. The talkback indicator indicates op when the doors are open and barberpole when they are in transit. The off position removes power from the motors, which stops the doors' movement.

The ET umbilical door switch on panel R2 positioned to GPC provides a backup method of releasing the centerline latches and closing the umbilical doors through guidance, navigation and control software through cathode ray tube display item entry during an RTLS abort. The operation of the centerline latches and closing of the umbilical doors are completely automated after external tank separation when the ET umbilical door switch on panel R2 is positioned to GPC . Two seconds after external tank separation, the centerline latches release the doors and the latches are stowed. The ET umbilical door centerline latch talkback indicator indicates sto when the centerline latches complete their motion eight seconds after external tank separation. The left and right umbilical doors are closed, and the ET umbilical door left and right talkback indicates cl 32 seconds after separation. The left and right umbilical door latches latch the doors closed, and the ET umbilical door left and right talkback indicates lat 38 seconds after separation.

Each umbilical door is covered with reusable thermal protection system in addition to an aerothermal barrier that required approximately 6 psi to compress to seal the door with adjacent thermal protection system tiles.

A closeout curtain is installed at each of the orbiter/external tank umbilicals. After external tank separation, the residual liquid oxygen in the main propulsion system is dumped through the three space shuttle main engines and the residual liquid hydrogen is dumped overboard. The umbilical curtain prevents hazardous gases (gaseous oxygen and hydrogen) from entering the orbiter aft fuselage through the umbilical openings before the umbilical doors are closed. The curtain also acts as a seal during the ascent phase of the mission to permit the aft fuselage to vent through the orbiter purge and vent system, thereby protecting the orbiter aft bulkhead at station Xo 1307. The curtain is designed to operate in range of minus 200 F to plus 250 F. The umbilical doors are opened when the orbiter has stopped at the end of landing rollout.

Various parameters are monitored and displayed on the flight deck control panel and CRT and transmitted by telemetry.

Contractors for the separation system include Hoover Electric, Los Angeles, Calif. (external tank umbilical centerline latch and actuator; umbilical door actuator and umbilical door latch actuator); U.S. Bearing, Chatsworth, Calif. (external tank/orbiter spherical bearing); Bertea Corp., Irvine, Calif. (umbilical retractor actuator); Space Ordnance Systems Division, Trans Technology Corp., Saugus, Calif. (orbiter/external tank separation bolt/cartridge detonator assembly, 0.75-inch frangible nut orbiter/external tank umbilical separation and 2.5-inch frangible nut/pyro components in orbiter/external tank aft attach separation system).

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