LANDING/DECELERATION SYSTEM

Description

Landing Gear

Landing Gear Doors

Landing/Deceleration Interfaces

Gear Retraction

Nose Gear and Door Uplock Mechanism

Gear Deployment

Nose Landing Gear Deployed

Main Landing Gear Deployed

Shock Struts

Wheels and Tires

Nose Landing Gear Stowed

Main Landing Gear Stowed

Drag Chute

Drag Chute Configuration

Nominal Sequence of Drag Chute Deployment, Inflation, and Jettison

Pilot and Main Chutes

Drag Chute Controls

Main Landing Gear Brakes

Panels F2, F3

Landing Gear Hydraulics

Brake System Hydraulic Power

Anti-Skid

Brake/Skid Control System Overview

ANTI SKID FAIL Caution and Warning Light on Panel F3

Anti-Skid Fail Light Status

Temperature Control

Nose Wheel Steering

OI-21 NWS Functional Drawing

Operations

Landing Gear

Commander’s LANDING GEAR Controls on Panel F6

Pilot’s LANDING GEAR Controls on Panel F8

BRAKE ISOL VLV Switches and Talkbacks on Panel R4

Ground Reset

Drag Chute

LG ARM/DN RESET Switch on Panel A12

Brake Controls

BRAKES Switch on Panel O14

BRAKES Switch on Panel O16

BRAKES Switch on Panel O15

ANTISKID Switch on Panel L2

NOSE WHEEL STEERING Switch on Panel L2

ROLL/YAW CSS Pushbutton Panel F2

NWS FAIL Light on Panel F3

Landing/Deceleration System Summary Data

Landing/Deceleration System Rules of Thumb

Description

The orbiter, unlike previous space vehicles, has the capability of landing on a runway using a conventional type of landing system. Once the orbiter touches down, the crew deploys the drag chute, begins braking, and starts nose wheel steering operations.

The orbiter drag chute, first used on the maiden flight of OV-105, improves the orbiter's deceleration and eases the loads on the landing gear and brakes.

Braking is accomplished by a sophisticated system that uses electro-hydraulic disk brakes with an anti-skid system. Only the two main gear sets have braking capability, and each can be operated separately.

Two primary steering options are available. By applying variable pressure to the brakes, the crew can steer the vehicle by a method called differential braking. Also, by selecting nose wheel steering, the crew can use the rudder pedal assembly to operate a hydraulic steering actuator incorporated in the nose landing gear. The crew can also use the rudder to assist steering while at higher ground speeds.

Landing Gear

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. The nose landing gear is located in the lower forward fuselage, and the main landing gear is located in the lower left and right wing area adjacent to the mid-fuselage.

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.

Landing Gear Doors

The nose landing gear has two doors, and each main gear has one door. When the crew commands gear deployment, the doors open automatically as the gear is dropped. This is accomplished by the door extend/retract mechanism, which is actuated by the dropping gear. The nose landing gear doors have two door hooks that hold the doors closed, and the main gear doors have four door hooks.

In addition, the doors have door-assist bungee assemblies. These assemblies exert additional force on the inside of the doors to assist in door deployment to overcome the aerodynamic forces acting against the doors and/or in case the pressure inside the wheel wells is less than the outside pressure. The nose landing gear bungee assist assemblies exert 2,000 pounds of force on the doors; the main landing gear bungee assist assemblies exert approximately 5,000 pounds of force on the doors over the first 2 inches of travel.

The nose landing gear also contains a pyro boost system to further assure nose gear door and gear extension in case high aerodynamic forces on the nose gear door are present. This pyro system is fired each time the landing gear is deployed.

Landing/Deceleration Interfaces

Landing/Deceleration Interfaces

Each of the landing gear doors has high temperature reusable surface insulation tiles on the outer surface and a thermal barrier or door seal to protect the landing gear from the high temperatures encountered during reentry.

Gear Retraction

During retraction, each gear is hydraulically rotated forward and up by ground support equipment until it engages an up-lock hook for each gear in its respective wheel well. The up-lock hook locks onto a roller on each strut. A mechanical linkage driven by each landing gear mechanically closes each landing gear door.

The nose landing gear is retracted forward and up into the lower forward fuselage and is enclosed by two doors. The main landing gear is 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.

Nose Gear and Door Uplock Mechanism

Nose Gear and Door Uplock Mechanism

Gear Deployment

When deployment of the landing gear is commanded by the crew, the up-lock hook for each gear is unlocked by hydraulic system 1 pressure. (See Section 2.1 for more information on orbiter hydraulic systems.) Once the hook is released from the roller on the strut, the gear is driven down and aft by springs, hydraulic actuators, aerodynamic forces, and gravity. A mechanical linkage released by each gear actuates the 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 down-lock bungees. If hydraulic system 1 pressure is not available to release the up-lock hook, a pyrotechnic initiator at each landing gear up-lock hook automatically releases the up-lock hook on each gear 1 second after the flight crew has commanded gear down.

The landing gear are deployed at 300 ± 100 feet and a maximum of 312 knots equivalent airspeed (KEAS).

Nose Landing Gear Deployed

Nose Landing Gear Deployed 

Main Landing Gear Deployed

Main Landing Gear Deployed

Shock Struts

The shock strut of each landing gear is the primary source of impact 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 landing gear shock strut assembly is constructed of high-strength, stress- and corrosion-resistant steel alloys, aluminum alloys, stainless steel, and aluminum bronze. The shock strut is a pneumatic-hydraulic 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.

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

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 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. (Current maximum operational crosswind is limited to 15 knots.)

Wheels and Tires

Landing gear wheels are made in two halves from forged aluminum. The nose landing gear tires are 32 by 8.8 inches and have a normal nitrogen inflation pressure of 350 psi prior to launch. The maximum allowable load per nose landing gear tire is approximately 45,000 pounds. Nose landing gear tires are rated for 217 knots maximum landing speed. They may be reused once.

The main landing gear tires are 44.5 by 21 inches and have 16 cord layers in a bias-ply design. They are normally inflated with nitrogen to a pressure of 370 psi. The maximum allowable load per main landing gear tire is 132,000 pounds. With a 60/40 percent tire load distribution, the maximum tire load on a strut is 220,000 pounds. The main gear tires are rated at 225 knots maximum ground speed and have a life of one landing.

Nose Landing Gear Stowed

Nose Landing Gear Stowed 

Main Landing Gear Stowed

Main Landing Gear Stowed

 

Drag Chute

The orbiter drag chute was designed to assist the deceleration system in safely stopping the vehicle on the runway at end of mission (EOM) and abort weights. Design requirements included the ability to stop a 248,000 lb TAL abort orbiter in 8,000 feet with a 10 knot tailwind on a hot (103° F) day and maximum braking at 140 knots ground speed or one half runway remaining. The drag chute, housed at the base of the vertical stabilizer, is manually deployed by redundant commands from the CDR or PLT at de-rotation. Drag chute deployment may be done between main gear touchdown and de-rotation only for vehicle mass moments 1.53 million foot pounds. The drag chute is jettisoned at 60 (±20) knots ground speed to prevent damage to the main engine bells. The drag chute will be used on lake bed and concrete runways except with crosswinds greater than 15 knots or in the presence of main engine bell repositioning problems. The drag chute may be deployed without engine bell repositioning if landing/rollout control problems exist.

Drag Chute Configuration

Drag Chute Configuration 

Nominal Sequence of Drag Chute Deployment, Inflation, and Jettison

Nominal Sequence of Drag Chute Deployment, Inflation, and Jettison 

Pilot and Main Chutes

When drag chute deployment is initiated, the door is blown off of the chute compartment by pyros and a mortar fires deploying a nine foot pilot chute. The pilot chute in turn extracts the 40 foot, partially reefed conical main chute. The main chute is reefed to 40 percent of its total diameter for about 3.5 seconds to lessen the initial loads on the vehicle. The main chute trails the vehicle by 89.5 feet on a 41.5 foot riser.

Drag Chute Controls

The drag chute deployment and jettison pushbuttons, ARM 1(2), DPY 1(2), and JETT 1(2) are installed on either side of both the CDR's and PLT's HUDs. Activation of each lighted pushbutton initiates a signal through the primary and redundant paths simultaneously. The deployment sequence requires that both the ARM and DPY pushbuttons be activated together. The JETT pushbutton signal will only be effective if the ARM command has previously been initiated. (ARM and JETT may be initiated simultaneously.) Circuit breakers for the drag chute controls are located on panels O15 and O16.

Main Landing Gear Brakes

Each of the orbiter's four main landing gear wheels has electro-hydraulic disc brakes and an associated anti-skid system. The disc brake assembly consists of nine discs, five rotors, four stators, a back-plate, and a pressure plate. The carbon-lined rotors are splined to the inside of the wheel and rotate with the wheel. The carbon-lined stators are splined to the outside of the axle assembly and do not rotate with the wheel.

The brakes are controlled by the commander or pilot applying toe pressure to the upper portion of the rudder pedals; electrical signals produced by rudder pedal toe pressure control hydraulic servo-valves at each wheel and allow hydraulic system pressure to actuate braking. Brakes cannot be applied until about 1.9 seconds after weight on the main gear has been sensed. The anti-skid system monitors wheel velocity and controls brake pressure 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 it also utilizes all three main dc electrical systems.

Panels F2, F3

Panels F2, F3

Landing Gear Hydraulics

Landing Gear Hydraulics 

Brake System Hydraulic Power

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.

When hydraulic system 1 or 2 source pressure drops below approximately 1,000 psi, switching valves provide automatic switching to hydraulic system 3. Loss of hydraulic system 1, 2, or both would have no effect on braking capability, because standby system 3 would automatically replace either system. Loss of hydraulic system 3 and either 1 or 2 would cause the loss of half of the braking power on each wheel, and additional braking distance would be required.

The brake valves in hydraulic systems 1, 2, and 3 must be open to allow hydraulic pressure to the brakes. All three valves are automatically commanded open after weight on the main landing gear is sensed. The 3,000 psi hydraulic pressure is reduced by a regulator in each of the brake hydraulic systems to 2,000 psig.

Anti-Skid

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 60 percent below 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. 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 approximately 10 to 15 knots 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 yellow ANTI SKID FAIL caution and warning light on panel F3 will be illuminated if the anti-skid fault detection circuit detects an open circuit or short in a wheel speed sensor or control valve servo-coil, 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 BRAKES switches on panels O14, O15, and O16 are ON, and the ANTISKID switch on panel L2 is OFF, the ANTISKID FAIL caution and warning light will also be illuminated.

Brake/Skid Control System Overview

Brake/Skid Control System Overview

ANTI SKID FAIL Caution and Warning Light on Panel F3

ANTI SKID FAIL Caution and Warning Light on Panel F3

Anti-Skid Fail Light Status

Anti-Skid Fail Light Status

Temperature Control

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, OV-104, and -105, 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, OV-104, and OV-105 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, OV-104 and -105, 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.

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.

HYDRAULICS BRAKE HEATER Switches on Panel R4

Nose Wheel Steering

The nose landing gear contains a hydraulic steering actuator that responds to electronic commands from the commander's or pilot's rudder pedals. Two types of operation are available, GPC and caster. The GPC mode is supported by hydraulic systems 1 and 2 through the selection of nose wheel steering (NWS) system 1 or 2 (NWS 1(2)). This provides redundant avionics modes regardless of hydraulic system support. NWS 1 or 2 will work with either hydraulic system 1 or 2.

In the GPC modes (NWS 1(2)), the flight control software uses accelerometer assembly feedback to modify commands from the rudder pedal transducer assemblies (RPTAs) to automatically counter hard overs or large lateral accelerations due to gear or tire malfunctions. If GPC 1 or 2, or FF1 or 2 is inoperative, steering down modes to caster if NWS 1 is selected. Similarly, the loss of GPC 3 or 4, or FF3 or 4 will prevent the use of NWS 2 and cause a down mode. The loss of either NWS system will only cause a down mode to castor. The other NWS system must be selected if required. In caster, no positive control over the nose wheel position is available, and differential braking and rudder are used for directional control.

A hydraulic servo-actuator mounted on the nose strut permits orbiter nose wheel steering up to 9° left or right after system activation. Hydraulic systems 1 and 2 provide redundant hydraulic pressure to either NWS 1 or 2. If the pressure in one system is more than twice that in the other, the higher pressure system provides hydraulic power for NWS. If NWS is not activated, or if hydraulic systems 1 and 2 fail, the NWS actuator acts as a nose wheel shimmy damper in the caster mode.

NWS can only be enabled after certain preconditions are met. Among these preconditions are two major milestones: weight-on-wheels (WOW) and weight-on-nose-gear (WONG). There are three sensors on each main gear designed to sense when main gear touchdown (MGTD) occurs so that WOW can be set. One sensor is a proximity sensor and the other two are wheel speed sensors (one per tire). Once WOW is set, the speed brake is commanded full open, flat turn discrete is set, half gain RHC is enabled, and the HUD format down modes. After WOW is set on one strut, brakes are also enabled.

OI-21 NWS Functional Drawing

OI-21 NWS Functional Drawing

WONG can be set by either of two proximity sensors located on the nose gear. Once WONG is set (presupposing WOW is already set and the vehicle attitude (theta) is less than zero), the ground speed enable "flag" is set. This enables NWS and the I-loaded downward deflection of the elevons for tire load relief. As a backup to the WOW and WONG discretes, the crew nominally selects MANUAL ET or SRB SEP and presses the associated pushbutton. This will manually bypass the WOW/WONG discretes and set the ground speed enable "flag."

Operations

Landing Gear

Landing gear deployment is initiated when the commander (on panel F6) or pilot (on panel F8) depresses the guarded LANDING GEAR ARM pushbutton and then the guarded DN pushbutton at least 15 seconds before predicted touchdown at a speed no greater than 312 KEAS at 300 ±100 feet above ground level (AGL).

NOTE

Deploying the landing gear at equivalent airspeeds greater than 312 knots may result in high aerodynamic loads on the doors and interference with the normal opening sequence.

Commander’s LANDING GEAR Controls on Panel F6

Commander’s LANDING GEAR Controls on Panel F6

 

Pilot’s LANDING GEAR Controls on Panel F8

Pilot’s LANDING GEAR Controls on Panel F8

Depressing the ARM pushbutton energizes latching relays for the landing gear extend valves 1 and 2 in preparation for gear deploy. It also arms the nose and main landing gear pyrotechnic initiator controllers and illuminates the yellow light in the ARM pushbutton. This is normally performed by the pilot at approximately 2,000 feet AGL.

The DN pushbutton is then depressed. This energizes latching relays that open the hydraulic system 1 extend valve 1 and hydraulic system 2 landing gear extend valve 2. Fluid in hydraulic system 1 flows to the landing gear uplock and strut actuators and the nose wheel steering switching valve. The green light in the DN pushbutton indicator is illuminated.

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 in transit.

The left and right main landing gear WOW switches produce output signals to the guidance, navigation, and control software to reconfigure the flight control system for landing and rollout gains.

The two WONG signals, along with WOW and theta (pitch angle) less than 0°, allow the GNC software to issue a nose wheel steering enable signal. This signal is then sent to the steering control box to enable nose wheel steering.

Six gear proximity switches are signal conditioned by the landing gear proximity sensor electronics box 1, located in avionics bay 1. Six additional gear proximity switches are signal conditioned by the landing gear proximity sensor electronics box 2, located in avionics bay 2. All WOW proximity switches are redundant through two signal conditioners.

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. During gear extension, a camming action opens the landing gear doors. 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.

The BRAKE 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 switch is positioned to CLOSE, hydraulic system 1 is isolated from the main landing gear brakes. The talkback indicator above the switch would indicate CL. The landing gear isolation valves cannot be opened or closed with hydraulic pressures less than approximately 100 psi. When the valve is open, hydraulic system 1 pressure is available to the main landing gear brake control valves. The normally closed landing gear extend valve 1, located downstream of the landing gear isolation valve, is not energized until a LANDING GEAR DN command is initiated by the commander or pilot on panel F6 or panel F8.

The BRAKE ISOL VLV 2 and 3 switches on panel R4 positioned to CLOSE isolate the corresponding hydraulic system from 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. Landing gear extend valve 2 is located downstream of brake isolation valve 2. This valve further isolates hydraulic system 2 supply pressure from the nose wheel steering and nose landing gear deploy actuators and is opened by a LANDING GEAR DN command.

When the nose and main LANDING GEAR DN command is initiated, hydraulic system 1 pressure is directed to the nose and main landing gear uplock hook actuators and strut actuators (provided that the LG/NWS HYD SYS switch is in the AUTO 1/2 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 landing gear/nose wheel steering hydraulic system switching valve will automatically select hydraulic system 2 supply pressure if system 1 should fail, thereby providing redundant hydraulics for NWS actuation and nose gear deploy.

NOTE

Hydraulic system 1 is the only hydraulic system for deploy of the main landing gear.

The GPC position of the BRAKE ISOL VLV 1, 2, and 3 switches on panel R4 permits the onboard computer to automatically control the valves in conjunction with computer control of the corresponding hydraulic system circulation pump.

BRAKE ISOL VLV Switches and Talkbacks on Panel R4

BRAKE ISOL VLV Switches and Talkbacks on Panel R4 

Two series valves, landing gear retract control valve 1 and 2, 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.

Prior to entry, the BRAKE ISOL VLV 1, 2, and 3 switches are positioned to GPC. This allows automatic opening of the valves after weight on main gear is sensed with GPC command via MDM FA1, FA2, and FA3, respectively. At 19,000 feet per second, the landing gear isolation valve automatic opening sequence begins under GNC software control. If the landing gear isolation valve is not opened automatically, the flight crew will be requested to manually open the valve by positioning the applicable BRAKE ISOL VLV switch to OPEN.

If the hydraulic system fails to release the landing gear within 1 second after the DN pushbutton is depressed, the nose and left and right main landing gear uplock sensors (proximity switches) will provide inputs to the pyro initiator controllers (PICs) for initiation of the redundant NASA standard initiators (pyro system 1 and 2). They release the same uplock hooks as the hydraulic system. As mentioned earlier, the nose landing gear, in addition, has a PIC and redundant NASA standard initiators that initiate a pyrotechnic power thruster 2 seconds after the DN pushbutton is depressed to assist gear deployment. This "nose gear pyro assist" pyro fires every time the gear are deployed.

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

Ground Reset

Landing gear reset is primarily a post landing function, which will be performed by the crew.

The LG ARM/DN RESET switch on panel A12 positioned to RESET unlatches the relays that were latched during landing gear deployment by the LANDING GEAR ARM and DN pushbutton indicators. The primary function of this procedure is to remove power to the PIC circuits that are still charged as a backup landing gear deploy method. The RESET position also will extinguish the yellow light in the ARM pushbutton indicator and the green light in the DN pushbutton.

Drag Chute

During entry, as the vehicle decelerates from 8000 to 3500 fps, the main engine bells are repositioned 10° below the nominal to preclude damage during drag chute deployment. The crew cannot monitor the bell repositioning but can determine that the system is enabled at item 19 on the SPEC 51 OVERRIDE display. The crew can inhibit the repositioning while in MM 301-304 by toggling item 19 in SPEC 51. Nominal EOM drag chute deployment will be initiated only if main engine bell repositioning is enabled.

LG ARM/DN RESET Switch on Panel A12

LG ARM/DN RESET Switch on Panel A12

Although the drag chute may be deployed at speeds up to 230 KEAS, current EOM procedures call for its deployment at 190 KEAS (195 KEAS for heavyweight vehicles) with a crosswind component no greater than 15 knots. If the drag chute is deployed above 230 KEAS, the drag chute pivot pin is designed to fail, resulting in the chute being jettisoned. Approximately one second after the CDR or PLT presses the ARM 1(2) and DPY 1(2) pushbuttons simultaneously, the pilot chute deploys. Within one second, the pilot chute extracts the main chute which deploys to its 40percent reefed diameter. After about 3.5 seconds of reefed deployment, two cutters sever the reefing ribbon allowing the main chute to inflate to its full 40 foot diameter.

WARNING

Deployment of the drag chute between 135 and 40 ft AGL can cause loss of control of the vehicle. Drag chute jettison must be initiated immediately to prevent loss of the vehicle and crew.

For pre-derotation deployment, an unreefed chute will produce a large nose pitch-up moment for a vehicle mass moment 1.53 million foot pounds. The large pitch-up may produce handling difficulties for the crew that could lead to loss of the vehicle and crew.

At 60 (±20) KGS, the drag chute will be jettisoned. Below 40 KGS, if drag chute jettison has not been initiated, the chute will be retained until the orbiter has stopped to minimize damage to the main engine bells.

Brake Controls

The BRAKES MN A, MN B, and MN C switches are located on panels O14, O15, and O16. These switches 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 OFF to disable the system.

When weight is sensed on the main landing gear, the brake/anti-skid boxes A and B are enabled and brake isolation valves 1, 2, and 3 are opened permitting the main landing gear brakes to become operational.

BRAKES Switch on Panel O14

BRAKES Switch on Panel O14

BRAKES Switch on Panel O16

BRAKES Switch on Panel O16

BRAKES Switch on Panel O15

BRAKES Switch on Panel O15

ANTISKID Switch on Panel L2

ANTISKID Switch on Panel L2

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. Pressure on the toe of the adjustable brake/rudder pedals results in a command to the wheel braking system.

Each brake pedal has four linear variable differential transducers. The left pedal transducer unit outputs 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 the brake pedal is deflected, the transducers transmit electrical signals of 0 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 proportionately 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 artificial braking feel to the crewmember.

Nose Wheel Steering

GPC Mode

The NOSE WHEEL STEERING switch on panel L2 positioned to NWS 1(2) enables the corresponding NWS (avionics) system. In addition to the NWS mode selections of the switch, the FLIGHT CNTLR POWER switch on panel F8 must be positioned to ON, and the flight control system ROLL/YAW CSS pushbutton on panel F2 or F4 must be depressed to enable the GPC for nose wheel steering. When either pushbutton is depressed, a white light illuminates the pushbutton.

NOSE WHEEL STEERING Switch on Panel L2

NOSE WHEEL STEERING Switch on Panel L2

When the commander or pilot makes an input to the rudder pedals in the NWS 1(2) mode, the rudder pedal command position is appropriately scaled within the GPC's software and transmitted to a summing network, along with lateral accelerometer inputs from within the flight control system. The accelerometer inputs are utilized to prevent any sudden orbiter lateral deviation. From this summing network, a nose wheel steering command is sent to a comparison network, as well as to the steering servo system.

ROLL/YAW CSS Pushbutton Panel F2

ROLL/YAW CSS Pushbutton on Panel F2 

NWS FAIL Light on Panel F3

NWS FAIL Light on Panel F3

Steering position transducers on the nose wheel strut receive redundant electrical excitation from the steering position amplifier, which receives redundant electrical power from data display unit 2.

Each of the three transducers transmits nose wheel position feedback to a redundancy management mid-value-select software. It 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 allow nose wheel steering to be enabled after WOW, WONG, and theta less than 0° are sensed in the software. The enable signal permits hydraulic system 1(2) pressure to be applied to the nose wheel steering actuator via the NWS switching valve. If hydraulic system 1 is lost, hydraulic system 2 provides the pressure for nose wheel steering. If both systems' pressures drop below approximately 1,325 psi, the actuator remains in the caster mode and a failure is annunciated to the NWS FAIL C/W yellow light on panel F3.

Landing/Deceleration System Summary Data

Landing/Deceleration System Rules of Thumb