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HYDRAULIC SYSTEM...

LANDING GEAR SYSTEM

LANDING GEAR SYSTEM

The landing gear system on the orbiter is a conventional aircraft tricycle configuration consisting of a nose landing gear and a left and right main landing gear. Each landing gear includes a shock strut with two wheel and tire assemblies. Each main landing gear wheel is equipped with a brake assembly with anti-skid protection. The nose landing gear is steerable. The nose landing gear is located in the lower forward fuselage, and the main landing gear are located in the lower left and right wing area adjacent to the midfuselage.

The nose landing gear is retracted forward and up into the lower forward fuselage and is enclosed by two doors. The main landing gear are also retracted forward and up into the left and right lower wing area, and each is enclosed with a single door. The nose and main landing gear can be retracted only during ground operations.

For retraction, each gear is hydraulically rotated forward and up during ground operations until it engages an uplock hook for each gear in its respective wheel well. The uplock hook locks onto a roller on each strut. Mechanical linkage driven by each landing gear mechanically closes the respective landing gear doors. All three landing gear doors have high-temperature reusable surface insulation thermal protection system tiles bonded to their outer surface with thermal barriers to protect and prevent the landing gear and wheel well from the high-temperature thermal loads encountered during the shuttle's entry into the atmosphere.

For deployment of the landing gear, the uplock hook for each gear is activated by the flight crew initiating a gear-down command. The uplock hook is hydraulically unlocked by hydraulic system 1 pressure applied to release it from the roller on the strut to allow the gear, assisted by springs and hydraulic actuators, to rotate down and aft. Mechanical linkage released by each gear actuates the respective doors to the open position. The landing gear reach the full-down and extended position within 10 seconds and are locked in the down position by spring-loaded downlock bungees. If hydraulic system 1 pressure is not available to release the uplock hook, a pyrotechnic initiator at each landing gear uplock hook automatically releases the uplock hook on each gear one second after the flight crew has commanded gear down.

The landing gears are deployed only after the spacecraft has an indicated airspeed of less than 300 knots (345 mph) and an altitude of approximately 250 feet.

The shock strut of each landing gear is the primary source of shock attenuation at landing. The struts have air/oil shock absorbers to control the rate of compression extension and prevent damage to the vehicle by controlling load application rates and peak values.

Each main landing gear wheel contains an electrohydraulic disc brake assembly with anti-skid control. The main landing gear brakes are controlled by the commander or pilot applying toe pressure to the rudder pedals; electrical signals produced by rudder pedal toe pressure control hydraulic servovalves at each wheel and allow hydraulic system pressure to perform braking. Main landing gear brakes cannot be applied until weight on the main gear has been sensed. The anti-skid system monitors wheel velocity and controls brake torque to prevent wheel lock and tire skidding. The braking/anti-skid system is redundant in that it utilizes system 1 and 2 hydraulic pressure as the active system with system 3 as standby and also utilizes all three main dc electrical systems.

The nose landing gear contains a hydraulic steering actuator that is electrohydraulically steerable through the use of the onboard general-purpose computers, the commander's or pilot's rudder pedals in conjunction with the orbiter flight control system in the control stick steering mode, or through the use of the commander's or pilot's rudder pedals in the direct mode. If hydraulic system 1 is inoperative, nose wheel steering changes to caster mode, and the commander or pilot would then apply toe pressure to the brake pedals to apply hydraulic pressure to the left and right main gear brakes as required for directional control using differential braking.

Each landing gear shock strut assembly is constructed of high-strength, stress- and corrosion-resistant steel alloys, aluminum alloys, stainless steel and aluminum bronze. Cadmium and chromium plating and urethane paint are applied to the strut surfaces for space flight protection. The shock strut is a pneudraulic shock absorber containing gaseous nitrogen and hydraulic fluid. Because the shock strut is subjected to zero-g conditions during space flight, a floating piston separates the gaseous nitrogen from the hydraulic fluid to maintain absorption integrity.

Landing gear wheels are made in two halves from forged aluminum and are primed and painted with two coats of urethane paint.

The LG hyd isol vlv 1, 2 and 3 switches on panel R4 control the corresponding landing gear isolation valve in hydraulic systems 1, 2 and 3. When the LG hyd isol vlv 1 switch on panel R4 is positioned to close , hydraulic system 1 is isolated from the nose and main landing gear deployment uplock hook actuators and strut actuators, nose wheel steering actuator and main landing gear brake control valves. A talkback indicator next to the switch would indicate cl. The landing gear isolation valves will not close or open unless the pressure in that system is at least 100 psi. When the LG hyd isol vlv 1 switch is positioned to open , it allows hydraulic system 1 source pressure to the main landing brake control valves and to the normally closed extend valve. The normally closed extend valve is not energized until a gear-down command is initiated by the commander or pilot on panel F6 or panel F8. The talkback indicator would indicate op . The LG hyd isol vlv 1 switch is left in the close position during the mission to prevent inadvertent gear deployment.

The LG hyd isol vlv 2 and 3 switches on panel R4 positioned to close isolate the corresponding hydraulic system from only the main landing gear brake control valves. The adjacent talkback indicator would indicate cl . When switches 2 and 3 are positioned to open, the corresponding hydraulic system source pressure is available to the main landing gear brake control valves. The corresponding talkback indicator would indicate op .

Thus, only hydraulic system 1 is used to deploy the nose and main landing gear and for nose wheel steering. When the nose- and main-landing-gear-down command is initiated by the commander or pilot on panel F6 or F8, hydraulic system 1 pressure is directed to the nose and main landing gear uplock hook actuators and strut actuators (provided that the LG hyd isol vlv 1 switch is in the open position) to actuate the mechanical uplock hook for each landing gear and allow the gear to be deployed and also provide hydraulic system 1 pressure to the nose wheel steering actuator. The main landing gear brake control valves receive hydraulic system 1 source pressure when the LG hyd isol vlv 1 switch is positioned to open. If hydraulic system 1 is unavailable, a pyrotechnic actuator attached to the nose and main landing gear uplock actuator would deploy the landing gear automatically one second after the gear-down command, actuate the mechanical uplock hook for each landing gear and allow the gear to be deployed. Because powered nose wheel steering would not be functional, directional control for steering would be accomplished by differential braking to caster the nose wheel.

The GPC position of the LG hyd isol vlv 1, 2 and 3 switches on panel R2 permits the onboard computer to automatically control the valves in conjunction with computer control of the corre sponding hydraulic system circulation pump. The LG hyd isol vlv 2 and 3 switches provide fluid circulation to only the main landing gear brake system, which dead-ends at the brake control valves. The LG hyd isol vlv 1 switch is left closed to prevent inadvertent gear deployment.

The normally open hydraulic system 1 redundant shutoff valve is a backup to the retract/circulation valve to prevent hydraulic pressure from being directed to the retract side of the nose and main landing gear uplock hook actuators and strut actuators if the retract/circulation valve fails to open during nose and main landing gear deployment.

The normally closed hydraulic system 1 dump valve is energized open to allow hydraulic system 1 fluid to return from the nose and main landing gear areas when deployment of the landing gear is commanded by the flight crew.

The activation/deactivation limits of the hydraulic fluid circulation systems can be changed during the mission by the flight crew or the Mission Control Center-Houston. The program also includes a timer to limit the maximum time a circulation pump will run and a priority system that automatically monitors hydraulic bootstrap pressure to allow all three circulation pumps to be on at the same time. The software timers allow this software to be used in contingency situations for ''time-controlled'' circulation pump operations in order to periodically boost an accumulator that is losing hydraulic fluid through a leaking priority valve or unloader valve.

During entry, if required, LG hyd isol vlv 1, 2 and 3 are positioned to GPC . At 19,000 feet per second, the landing gear isolation valve automatic opening sequence begins under GN&C software control. If the landing gear isolation valve is not opened automatically, the flight crew will be requested by the Mission Control Center to open the valve by positioning the applicable LG hyd isol vlv to open. Landing gear isolation valve 2 is automatically opened six minutes and 37 seconds later, and this is followed by the automatic opening of landing gear isolation valve 1 when orbiter velocity is at 800 feet per second or less. Landing gear isolation valve 3 is automatically opened at ground speed enable. Landing gear isolation valve 1 is next to last to ensure that an inadvertent gear deployment would occur as late (low airspeed) as possible.

Note that the hydraulic system 1 retract/circulation valve would be automatically closed when the landing gear system is armed for deployment.

The commander and pilot have a landing gear deployment arm and dn (down) guarded push button switch/light indicators and landing left, nose and right indicators. The commander's controls and indicators are on panel F6, and the pilot's controls and indicators are on panel F8. The dn push button, when depressed, energizes the hydraulic system 1 normally closed extend valve, permitting hydraulic system 1 source pressure for gear deployment and nose wheel steering.

The proximity switches on the nose and main landing gear doors and struts provide electrical signals to control the landing gear nose, left and right indicators on panels F6 and F8. The output signals of the landing gear and door uplock switches drive the landing gear up position indicators and the backup pyrotechnic release system. The output signals of the landing gear downlock switches drive the landing gear dn position indicators. The landing gear indicators are barberpole when the gear is deploying (or retracting).

The left and right main landing gear weight-on-wheels switches produce output signals to the guidance, navigation and control software to reconfigure the flight control system for landing.

The two weight-on-nose-gear signals run to the main landing gear brake/skid control boxes to prevent the main landing gear brakes from being applied until the nose gear is in contact with the runway and also to the GN&C software, which computes a nose wheel steering enable signal. This enable signal is then sent to the NWS control box to prevent NWS until the nose gear is in contact with the runway.

The six group 1 switches are signal conditioned by the landing gear proximity sensor electronics box 1, located in avionics bay 1. The six group 2 switches are signal conditioned by the landing gear proximity sensor electronics box 2, located in avionics bay 2.

Landing gear deployment is initiated when the commander or pilot depresses the guarded arm push button switch/light indicator and then the guarded dn push button switch/light indicator at least 15 seconds before predicted touchdown and at a speed no greater than 300 knots (345 mph).

Depressing the arm push button switch/light indicator energizes latching relays that close the hydraulic system 1 landing gear retract/circulation valve and the normally open redundant shutoff valve to the retract/circulation valve. It also arms the nose and main landing gear pyrotechnic initiator controllers and illuminates the yellow light in the arm push button switch/light indicator.

The dn push button switch/light indicator is then depressed. This energizes latching relays that open the hydraulic system 1 landing gear normally closed extend control valve, permitting the fluid in hydraulic system 1 to flow to the landing gear uplock and strut actuators and nose wheel steering. The relays also open the normally closed dump valve, allowing the landing gear retract line fluid to flow in to the hydraulic system 1 return line. The green light in the dn push button switch/light indicator is illuminated.

Hydraulic system 1 source pressure is routed to the nose and main landing gear uplock actuators, which releases the nose and main landing gear and door uplock hooks. As the uplock hooks are released, the gear begins its deployment and mechanical linkage attached to the doors and fuselage is powered by landing gear strut camming action, during gear extension, which opens the landing gear doors. There are two landing gear doors for the nose gear and one for each main gear. The landing gear free falls into the extended position, assisted by the strut actuators and airstream in the deployment. The hydraulic strut actuator incorporates a hydraulic fluid flow through orifice (snubber) to control the rate of landing gear extension and thereby prevent damage to the gear's downlock linkages.

If hydraulic system 1 fails to release the landing gear within one second after the dn push button is depressed, the nose and left and right main landing gear uplock sensors (proximity switches) will provide inputs to the pyro initiator controllers for initiation of the redundant NASA standard detonators (nose, left and right main landing gear pyrotechnic backup release system). They release the same uplock hooks as the hydraulic system. The nose landing gear, in addition, has a PIC and redundant NSDs that initiate a pyrotechnic power thruster two seconds after the dn push button is depressed to assist gear deployment.

The landing gear drag brace overcenter lock and spring-loaded bungee lock the nose and main landing gear in the down position.

The ldg gr/arm/dn reset switch positioned to reset on panel A12 unlatches the relays that were latched during landing gear deployment by the landing gear arm and dn push button light/switch indicators. This is primarily a ground function, which will be performed only during landing gear deactivation.

The reset position also will extinguish the yellow light in the arm push button switch/light indicator and the green light in the dn push button switch/light indicator. In addition, the hydraulic system 1 landing gear dump valve is closed, the extend control valve is closed, the retract/circulation valve is opened only if the switch is in the open position, and its redundant shutoff valve is opened (de-energized) and de-energizes the landing gear PIC circuits.

The nose landing gear tires are 32 by 8.8 inches and will withstand a burst pressure of not less than 3.2 times the normal inflation pressure of 300 psi. The inflation agent is gaseous nitrogen. The maximum allowable load per nose landing gear tire is approximately 45,000 pounds and rated at 224 knots (258 mph) landing speed.

The nose landing gear shock strut has a 22-inch stroke. The maximum allowable derotation rate is approximately 9.4 degrees per second or 11 feet per second, vertical sink rate.

The main landing gear tires are 44.5 by 16 and 21 inches. The normal inflation pressure is 315 psi, and the inflation agent is gaseous nitrogen. The maximum allowable load per main landing gear tire is 123,000 pounds. If the orbiter touches down with a 60/40 percent load distribution on a strut's two tires, with one tire supporting the maximum load, then the other tire can support a load of only 82,410 pounds. Therefore, the maximum tire load on a strut is 205,410 pounds with a 60/40 percent tire load distribution. The tires are rated at 225 knots (258 mph).

The main landing gear shock strut stroke is 16 inches. The allowable main gear sink rate for a 212,000-pound orbiter is 9.6 feet per second; for a 240,000-pound orbiter, it is 6 feet per second. With a 20-knot (23-mph) crosswind, the maximum allowable gear sink rate for a 212,000-pound orbiter is 6 feet per second; for a 240,000-pound orbiter, it is approximately 5 feet per second.

The landing gear tires have a life of one landing.

MAIN LANDING GEAR BRAKES

    Each of the orbiter's four main landing gear wheels has electrohydraulic disc brakes and an anti-skid system.

    Each main landing gear wheel has a disc brake assembly consisting of nine discs, four rotors, three stators, a backplate and a pressure plate. The carbon-lined beryllium rotors are splined to the inside of the wheel and rotate with the wheel. The carbon-lined beryllium stators are splined to the outside of the axle assembly and do not rotate with the wheel.

    Each of the four main landing gear wheel brake assemblies is supplied with pressure from two different hydraulic systems. Each brake hydraulic piston housing has two separate brake supply chambers. One chamber receives hydraulic source pressure from hydraulic system 1 and the other from hydraulic system 2. There are eight hydraulic pistons in each brake assembly. Four are manifolded together from hydraulic system 1 in a brake chamber. The remaining four pistons are manifolded together from hydraulic system 2. When the brakes are applied, the eight hydraulic pistons press the discs together, providing brake torque.

    In the event of the loss of hydraulic system 1 or 2 source pressure, switching valves provide automatic switching to the standby hydraulic system 3 when the active hydraulic system source pressure drops below approximately 1,000 psi. If hydraulic system 1 is unavailable, it has no effect on the braking system because standby system 3 would automatically replace system 1. Loss of hydraulic system 2 or both 1 and 2 would also have no effect on the braking system because system 3 would automatically switch to replace system 2 or 1 and 2. Loss of hydraulic systems 1 and 3 would cause the loss of half of the braking power on each wheel and additional braking distance would be required. Loss of hydraulic systems 2 and 3 would also cause the loss of half of the braking power on each wheel, requiring additional braking distance.

    As in the landing gear deployment, the landing gear isolation valve in hydraulic systems 1, 2 and 3 must be open to allow the applicable hydraulic source pressure to the main landing gear brakes.

    The brakes MN A, MN B and MN C switches are located on the flight deck display and control panels O14, O15 and O16 and allow electrical power to brake/anti-skid control boxes A and B. The antiskid switch located on panel L2 provides electrical power for enabling the anti-skid portion of the braking system boxes A and B. The brakes MN A, MN B and MN C switches are positioned to on to supply electrical power to brake boxes A and B and to off to remove electrical power. The antiskid switch is positioned to on to enable the anti-skid system and to off to disable the system.

    When weight is sensed on the main landing gear, the brake/anti-skid boxes A and B are enabled, permitting the main landing gear brakes to become operational.

    The main landing gear brakes controlled by the commander's or pilot's brake pedals are located on the rudder pedal assemblies at the commander's and pilot's stations. The pedals' positions are adjustable by a handle. The braking commands are accomplished by the commander or pilot initiating toe pressure on the top of the rudder pedal assembly.

    Each brake pedal (left and right) has four linear variable differential transducers. The left pedal transducer unit will output four separate braking signals through the brake/skid control boxes for braking control of the two left main wheels. The right pedal transducer unit does likewise for the two right main wheels. When toe pressure is applied to the brake pedal, the transducers transmit electrical signals of zero to 5 volts dc to the brake/anti-skid control boxes. If both right pedals are moved, the pedal with the greatest toe pressure becomes the controlling pedal through electronic OR circuits. The electrical signal is proportional to the toe pressure. The electrical output energizes the main landing gear brake coils proportionally to brake pedal deflection, allowing the desired hydraulic pressure to be directed to the main landing gear brakes for braking action. The brake system bungee at each brake pedal provides the braking artificial feel to the crew member.

    Each of the three hydraulic systems' source pressure of 3,000 psi is reduced by a regulator in each of the brake hydraulic systems to 1,500 psi.

    The anti-skid portion of the brake system provides optimum braking by preventing tire skid or wheel lock and subsequent tire damage.

    Each main landing gear wheel has two speed sensors that supply wheel rotational velocity information to the skid control circuits in the brake/skid control boxes. The velocity of each wheel is continuously compared to the average wheel velocity of all four wheels. Whenever the wheel velocity of one wheel is below 30 percent of the average velocity of the four wheels, skid control removes brake pressure from the slow wheel until the velocity of that wheel increases to an acceptable range. The brake system contains eight brake/skid control valves that receive signals from the brake/skid control boxes. Each valve controls the hydraulic brake pressure to one of the brake chambers. The brake/skid control valves contain a brake coil and a skid coil. The brake coil allows hydraulic pressure to enter the brake chambers. The skid coil, when energized by the skid control circuit, provides reverse polarity to the brake coil, preventing the brake coil from allowing brake pressure to the brake chamber.

    Anti-skid control is automatically disabled below 9 to 14 knots (11 to 17 mph) to prevent loss of braking for maneuvering and/or coming to a complete stop.

    The anti-skid system control circuits contain fault detection logic. The antiskid yellow caution and warning light located on the flight deck display and control panel F3 will be illuminated if the anti-skid fault detection circuit detects an open or short in a wheel speed sensor, open or short in a anti-skid control valve servocoil or a failure in an anti-skid control circuit. A failure of these items will only deactivate the failed circuit, not total anti-skid control. If the brake power switches are on and the antiskid switch is off , the antiskid caution and warning light will be illuminated.

    Insulation and electrical heaters are installed on the portions of the hydraulic systems that are not adequately thermally conditioned by the individual hydraulic circulation pump system because of stagnant hydraulic fluid areas.

    Redundant electrical heaters are installed on the main landing hydraulic flexible lines located on the back side of each main landing gear strut between the brake module and brakes. These heaters are required because the hydraulic fluid systems are dead-ended and fluid cannot be circulated with the circulation pumps. In addition, on OV-103 and OV-104, the hydraulic system 1 lines to the nose landing gear are located in a tunnel between the crew compartment and forward fuselage. The passive thermal control systems on OV-103 and OV-104 are attached to the crew compartment, which leaves the hydraulic system 1 lines to the nose landing gear exposed to environmental temperatures, thus requiring electrical heaters on the lines in the tunnel. Since the passive thermal control system on OV-102 is attached to the inner portion of the forward fuselage rather than the crew compartment, no heaters are required on the hydraulic system 1 lines to the nose landing gear on OV-102.

    The hydraulics brake heater A, B, and C switches on panel R4 enable the heater circuits. On OV-103 and OV-104, hydraulics brake heater switches A, B and C provide electrical power from the corresponding main buses A, B and C to the redundant heaters on the main landing gear flexible lines and the hydraulic system 1 lines in the tunnel between the crew compartment and forward fuselage leading to the nose landing gear. Thermostats on each electrical A, B and C system cycle the heaters automatically off or on.

    The hydraulics brake heater A, B and C switches on panel R4 enable the heater circuits on only the main landing gear hydraulic flexible lines on OV-102.

    Because problems were encountered with the main landing gear braking system in the majority of the first 24 landings, an improvement program has been implemented for the main landing gear and braking system in addition to a long-term improvement program for the main landing gear brakes.

    Main landing gear axle stiffness has been increased to reduce brake-to-axle deflections to preclude brake damage, which occurred in previous landings. This should also minimize tire wear. With the increased axle thickness, existing axle/bearing and axle/sensor interfaces are maintained. All main landing gear axles will be changed before the three orbiters return to flight.

    Six orifices were added to the hydraulic passages in the brake hydraulic piston housing to restrict circular fluid flow within the chambers in order to stop the whirl phenomenon, which has been identified as the cause of brake damage.

    The electronic brake control boxes were modified to provide hydraulic pressure balancing between adjacent brakes in order to equalize energy applications. This results in higher efficiency and allows full capability of adjacent brakes. The anti-skid circuitry that reduced brake pressure to the opposite wheel if a flat tire was detected was removed.

    The previous, thinner, carbon-lined beryllium stator discs are being replaced in two positions with thicker discs to provide a significant increase in braking energy capability. The additional material added to the stators improves heat capacity, with resulting lower temperatures, and provides the stators with greater strength. Note that the main landing gear brakes, which were exposed to two 14-million- foot-pound wear-in cycles added before installation on the orbiter, reduced damage to the brakes during landing. The thicker stator discs will provide approximately 65 million foot pounds of energy absorption, which is a significant increase over the thinner stator discs.

    A long-term structural carbon brake program is in progress to provide higher braking capability by increasing maximum energy absorption capability to 82 million foot pounds and to reduce refurbishment costs. These new brakes will consist of a five-rotor, disc-type carbon configuration for each main landing gear wheel brake. The goal is to demonstrate that the carbon heat sink brake design will have the capability of providing a one-time stop of 100 million foot pounds. The go-ahead for the carbon brake design was given in January 1986, with delivery scheduled in late April 1988.

    Upon the return to flight of the space shuttle, end-of-mission landings are planned for Edwards Air Force Base in California until the performance of the landing gear system is fully understood and a higher confidence in the weather prediction capability is established at the Kennedy Space Center shuttle landing facility runway area in Florida.

    Strain gauges have been added to each nose and main landing gear wheel in order to monitor tire pressure and provide the status of tire pressures during launch pad stay, launch, orbit, deorbit and landing to the flight crew and Mission Control Center in Houston.

    A landing gear tire improvement and runway-surface study is in progress at NASA's Langley Research Center to determine how best to decrease tire wear experienced during previous Kennedy Space Center landings and improve crosswind landing capability. Six new tire designs will be evaluated at the Langley Research Center with various yaw and tilt angles at speeds up to 253 mph. Additional tests at Langley Research Center are to provide the ability to mathematically model tire side-force characteristics.

    Modifications were made to the Kennedy Space Center shuttle landing facility runway. The full 300-foot width of 3,500-foot sections at both ends of the runway were ground to smooth the runway surface and remove cross grooves. The corduroy ridges are smaller than those they replace and run the length of the runway rather than across its width. The existing landing zone light fixtures were also modified, and the markings of the entire runway and overruns were repainted. The primary purpose of the modifications is to enhance safety by reducing tire wear during landing.

    The current strength of the orbiter landing system is a maximum of 240,000 pounds. Evaluations for landing loads of 256,000 pounds, associated with abort landings, are to be completed by the spring of 1988.

    Other studies in progress are arrest barriers at landing sites (except lake bed runways) to provide safe stops in the event of main landing gear brake failure or unforeseen wet runway conditions. The barrier net study will determine whether the barrier can safely stop a 256,000-pound orbiter traveling at 100 knots (115 mph) at the end of runway. Also under study are (1) the installation of a skid on the landing gear that could preclude the potential for a second blown tire on a gear on which one tire has blown, (2) a rim that would provide a predictable roll in the event of the loss of both tires on a single or multiple gear and (3) the addition of a drag chute.

NOSE WHEEL STEERING

    The orbiter nose wheel is steerable after nose wheel touchdown at landing. The nose wheel is electrohydraulically steerable through the use of the general-purpose computers and the commander's or pilot's rudder pedals, in conjunction with the orbiter flight control system in the control stick steering mode. In addition, the nose wheel may also be steerable through the use of the commander's or pilot's rudder pedals in the direct mode.

    Nose wheel steering is advantageous to allow positive lateral directional control of the orbiter during the rollout phase of a mission in the presence of high crosswinds and blown tires. Recent modifications of the nose wheel steering system have been incorporated to allow a safe high-speed engage of the nose wheel steering system.

    The first in the series of changes was to redefine the NWS switch on panel L2. The forward position of the switch now activates the direct mode of nose wheel steering, the center position activates the GPC mode of nose wheel steering, and the aft switch off position deactivates all nose wheel steering. A flexible handle extension was also added to the switch handle.

    The commander selects and activates the GPC mode while performing the pre-entry checklist of cockpit switches while his eyes and attention are inside the cockpit. No other flight crew action is required in order to engage active nose wheel steering at the time of nose wheel touchdown. Other added features provide assurance that the nose wheel will be positioned straight ahead at the moment of nose wheel touchdown, and the final enable signals in the GPC mode are sequenced at nose wheel touchdown and begin active nose wheel steering at that time.

    During derotation and nose wheel touchdown, the flight crew is looking out the windshield and at the head-up display. If the commander chooses to select the direct mode of nose wheel steering or deactivate all nose wheel steering by selecting off without visual verification of the actual switch position, he can sweep his left hand over the switch (forward motion for direct or aft motion for off ) and be certain of the position of the NWS switch without looking at the switch.

    Only hydraulic system 1 supplies hydraulic system supply pressure to the nose wheel hydraulic actuator for steering in either the GPC or direct mode. If hydraulic system 1 supply pressure is unavailable, the commander or pilot can apply the left and right main landing brakes by applying toe pressure on the rudder pedals differentially, which allows directional control by differential braking. The nose wheel steering actuator limits the motion of the nose wheel to plus or minus 10 degrees and also prevents nose wheel shimmey.

    When the NWS switch is set to either GPC or direct , the nose wheel steering system fault detection logic detects (1) the interruption of hydraulic system 1 supply pressure; (2) an open or short in the nose wheel steering servovalve circuitry; (3) an open or short in the position feedback, rate position error; (4) an open or short in the command transducer; and (5) broken linkage or loss of electrical power. At this time, the NWS fail C/W light on panel F3 is illuminated, and the nose wheel steering reverts automatically to the caster mode.

    General-Purpose Computer Mode. The NWS switch on panel L2, positioned to GPC, enables the nose wheel steering solenoid control valves, which supplies hydraulic system 1 pressure to the nose wheel steering servovalves. In addition to the GPC mode selection of the NWS switch, the flight control system roll/yaw CSS push button switch/light indicator on panel F2 or F4 must be depressed to enable GPC mode steering. When either push button switch/light indicator on panel F2 or F4 is depressed, a white light is illuminated within the push button.

    When the commander or pilot positions the rudder pedals in the GPC roll/yaw CSS mode, the rudder pedals command position is appropriately scaled within the GPC's software and transmitted to a summing network, along with accelerometer inputs from within the flight control system. The accelerometer inputs are utilized to prevent any sudden lateral deviation of the orbiter's velocity vector direction. From this summing network, a nose wheel steering command is sent to a comparison network as well as to the steering servo system.

    The three new steering position transducers in the unit added on the nose wheel strut receive redundant electrical excitation from the new steering position amplifier (added to the middeck ceiling), which receives redundant electrical power from data display unit 2 on the flight deck.

    Each of the three new transducers transmits nose wheel position feedback to a redundancy management mid-value-select software, which then transmits a nose wheel position signal to the comparison network. The orbiter nose wheel commanded and actual positions are compared for position error and for rates to reduce any error. Absence of an error condition will enable nose wheel steering after weight on the nose gear is sensed in the software. Weight on the nose gear requires weight on all three landing gear and a nose-down orbiter attitude. The enable signal permits hydraulic system 1 pressure to be applied to the nose wheel steering actuator. If hydraulic pressure is below 1,350 psi, the actuator remains in the shimmy damp mode, and a failure is annunciated to the NWS fail C/W yellow light on panel F3. If hydraulic system 1 pressure is above 1,350 psi, the actuator is configured to the GPC CSS mode to position the nose wheel utilizing the commander's or pilot's rudder pedals.

    If more than one of the three new position transducer feedback signals are lost, the comparison logic switches hydraulic system 1 supply pressure off to the nose wheel steering actuator, and pressure sensing immediately inhibits nose wheel steering in the GPC CSS mode automatically. Nose wheel steering will then revert to a nose wheel caster mode and the NWS fail yellow C/W light will be illuminated on panel F3. This redundancy in the GPC CSS mode provides immediate protection against any undesired lateral nose wheel response during nose wheel steering.

    If nose wheel steering has failed, the NWS switch on panel L2 positioned to off will extinguish the NWS fail C/W yellow light on panel F3, unless the GPC CSS mode comparison circuits caused the light to be illuminated. When the NWS switch is positioned to off, hydraulic system 1 supply pressure to the nose wheel steering system is removed, allowing the nose wheel to caster.

    Direct Mode. When the NWS switch on panel L2 is positioned to direct, the nose wheel steering solenoid control is enabled after weight is sensed on the nose gear. This supplies hydraulic system 1 supply pressure to the nose wheel steering servo system.

    When the commander or pilot positions the rudder pedals in the direct mode, the rudder pedals command position is shaped to provide a less sensitive rudder pedal command in the midrange and provides steering commands directly to the nose wheel steering servoactuator, bypassing the GPC software altogether.

    The two additional steering position transducer/amplifier boxes are from Honeywell, the NWS box was modified by Sterer, and the DDU 2 modification was made by Collins.

    A new additional nose wheel steering system design is expected to be available in late 1988 or early 1989.

    The contractors for the landing gear are B.F. Goodrich, Troy, Ohio (main and nose landing gear wheel and main landing gear brake assembly and the nose/main gear tires); Bertea Corp., Irvine, Calif. (main landing gear hydraulic uplock actuator, main landing gear strut actuator and nose landing gear uplock actuator); Menasco Manufacturing Co., Burbank, Calif. (main and nose landing gear shock struts and drag brace assembly); Sterer Engineering and Manufacturing, Los Angeles, Calif. (nose gear steering/damping and solenoid-operated landing gear uplock control valves); Crane Co., Hydro Aire, Burbank, Calif. (main landing gear brake anti-skid system); Eldec Corp., Lynwood, Wash. (landing gear proximity switch); OEA, Denver, Colo. (nose gear uplock release pyro thruster); Scott Inc., Downers Grove, Ill. (main landing gear uplock release thruster actuator).

    Click Here for CAUTION AND WARNING SYSTEM

Table of Contents


Information content from the NSTS Shuttle Reference Manual (1988)
Last Hypertexed Wednesday October 11 17:45:37 EDT 1995
Jim Dumoulin (dumoulin@titan.ksc.nasa.gov)



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