The space shuttle coordinate reference system is a means of locating specific points on the shuttle. The system is measured in inches and decimal places; Xo designates the longitudinal (forward and aft) axis, Yo the lateral (inboard and outboard) axis and Z o the vertical (up and down) axis. The subscript ''o'' indicates orbiter; similar reference systems are used for the external tank (T), solid rocket booster (B), and overall space shuttle system (S).
In each coordinate system, the X-axis zero point is located forward of the nose tip; that is, the orbiter nose tip location is 236 inches aft of the zero point (at X o 236), the external tank nose cap tip location is at XT 322.5, and the solid rocket booster nose tip location is at XB 200. In the orbiter, the horizontal X o , Y o reference plane is located at Z o 400, which is 336.5 inches above the external tank horizontal XT , YT reference plane located at ZT 400. The solid rocket booster horizontal XB , YB reference plane is located at Z B 0 and coincident with the external tank horizontal plane at ZT 400. The solid rocket booster vertical XB , ZT planes are located at + Y S 250.5 and -YS 250.5. Also, the orbiter, external tank, and shuttle system center X, Z planes coincide.
From the X = 0 point, aft is positive and forward is negative for all coordinate systems. Looking forward, each shuttle element Y-axis point right of the center plane (starboard) is positive and each Y-axis point left of center (port) is negative. The Z axis of each point within all elements of the shuttle except the SRBs is positive, with Z = 0 located below the element. In the SRBs each Z-coordinate point below the XB , YB reference plane is negative and each point above that plane is positive.
The shuttle system and shuttle element coordinate systems are related as follows: the external tank XT 0 point coincides with XS 0, the SRB XB 0 point is located 543 inches aft, and the Y o , Zo reference plane is 741 inches aft of X S 0.
The orbiter structure is divided into nine major sections: the forward fuselage, which consists of upper and lower sections that fit clamlike around a pressurized crew compartment; wings; midfuselage; payload bay doors; aft fuselage; forward reaction control system; vertical tail; orbital maneuvering system/reaction control system pods; and body flap. The majority of the sections are constructed of conventional aluminum and protected by reusable surface insulation.
The forward fuselage structure is composed of 2024 aluminum alloy skin-stringer panels, frames and bulkheads.
The crew compartment is supported within the forward fuselage at four attachment points and is welded to create a pressure-tight vessel. The three-level compartment has a side hatch for normal passage and hatches in the airlock to permit extravehicular and intravehicular activities. The side hatch can be jettisoned.
The midfuselage is a 60-foot section of primary load-carrying structure between the forward and aft fuselages. It includes the wing carry-through structure and the payload bay doors. The skins consist of integral-machined aluminum panels and aluminum honeycomb sandwich panels. The frames are constructed from a combination of aluminum panels with riveted or machined integral stiffeners and a truss structure center section. The upper half of the midfuselage consists of structural payload bay doors hinged along the side and split at the top centerline. The doors are graphite epoxy frames and honeycomb panel construction.
The aft fuselage includes a truss-type internal structure of diffusion-bonded elements that transfer the main engine thrust loads to the midfuselage and external tank. (In OV-105 , the truss-type internal structure is of a forging construction.) The aft fuselage's external surface is of standard construction except for the removable OMS/RCS pods, which are constructed of graphite epoxy skins and frames. An aluminum bulkhead shield with reusable insulation at the rear of the orbiter protects the rear portion of the aft fuselage.
The wing is constructed of a conventional aluminum alloy, using a corrugated spar web, truss-type ribs and riveted skin-stringer and honeycomb covers. The elevons are constructed of aluminum honeycomb and are split into two segments to minimize hinge binding and interaction with the wing.
The vertical tail, a conventional aluminum alloy structure, is a two-spar, multirib, integrally machined skin assembly. The tail is attached to the aft fuselage by bolted fittings at the two main spars. The rudder/speed brake assembly is divided into upper and lower sections, which are split longitudinally and actuated individually to serve as both rudder and speed brake.
These major structural assemblies are mated and held together by rivets and bolts. The midfuselage is joined to the forward and aft fuselage primarily by shear ties, with the midfuselage overlapping the bulkhead caps at stations Xo 582 and Xo 1307. The wing is attached to the midfuselage and aft fuselage primarily by shear ties, except in the area of the wing carry-through, where the upper panels are attached with tension bolts. The vertical tail is attached to the aft fuselage with bolts that work in both shear and tension. The body flap, which has aluminum honeycomb covers, is attached to the lower aft fuselage by four rotary actuators.
The forward fuselage consists of the upper and lower fuselages. It houses the crew compartment and supports the forward reaction control system module, nose cap, nose gear wheel well, nose gear and nose gear doors.
The forward fuselage is constructed of conventional 2024 aluminum alloy skin-stringer panels, frames and bulkheads. The panels are single curvature and stretch-formed skins with riveted stringers spaced 3 to 5 inches apart. The frames are riveted to the skin-stringer panels. The major frames are spaced 30 to 36 inches apart. The Y o 378 upper forward bulkhead is constructed of flat aluminum and formed sections riveted and bolted together; the lower is a machined section. The bulkhead provides the interface fitting for the nose section.
The nose section contains large machined beams and struts. The structure for the nose landing gear wheel well consists of two support beams, two upper closeout webs, drag-link support struts, nose landing gear strut and actuator attachment fittings, and the nose landing gear door fittings. The left and right landing gear doors are attached by hinge fittings in the nose section. The doors are constructed of aluminum alloy honeycomb, and although the doors are the same length, the left door is wider than the right. Each door has an up-latch fitting at the forward and aft ends to lock the door closed when the gear is retracted, and each has a pressure seal in addition to a thermal barrier. Lead ballast in the nose wheel well and on the X o 378 bulkhead provides weight and center-of-gravity control. The nose wheel well will accommodate 1,350 pounds of ballast, and the X o 378 bulkhead will accommodate a maximum of 2,660 pounds.
The forward fuselage carries the basic body-bending loads (a tendency to change the radius of a curvature of the body) and reacts nose landing gear loads.
The forward fuselage is covered with reusable insulation, except for the six windshields, two overhead windows and side hatch window areas around the forward RCS engines. The nose cap is also a reusable thermal protection system. It is constructed of reinforced carbon-carbon and has thermal barriers at the nose cap-structure interface.
The forward fuselage skin has structural provisions for installing antennas, deployable air data probes and the door eyelet openings for the two star trackers. Two openings are required in the upper forward fuselage for star tracker viewing. Each opening has a door for environmental control.
The forward orbiter/external tank attach fitting is at the Xo 378 bulkhead and the skin panel structure aft of the nose gear wheel well. Purge and vent control is provided by flexible boots between the forward fuselage and crew compartment around the windshield windows, overhead observation window, crew hatch window and star tracker openings. The forward fuselage is isolated from the payload bay by a flexible membrane between the forward fuselage and crew compartment at Xo 582.
Six forward outer pane windshields are installed on the forward fuselage. They are described in the section on windows. The window structural frames in the forward fuselage are five-axis machined parts.
The forward RCS module is constructed of conventional 2024 aluminum alloy skin-stringer panels and frames. The panels are composed of single-curvature and stretch-formed skins with riveted stringers. The frames are riveted to the skin-stringer panels. The forward RCS module is secured to the forward fuselage nose section and forward bulkhead of the forward fuselage with 16 fasteners, which permit the installation and removal of the module. The components of the forward RCS are mounted and attached to the module, which will have a reusable thermal protection cover, in addition to thermal barriers installed around it and the RCS engine interfaces and the interface-attachment area to the forward fuselage.
The three-level crew compartment is constructed of 2219 aluminum alloy plate with integral stiffening stringers and internal framing welded together to create a pressure-tight vessel. The compartment has a side hatch for normal ingress and egress, a hatch into the airlock from the middeck, and a hatch from the airlock through the aft bulkhead into the payload bay for extravehicular activity and payload bay access.
Redundant pressure window panes are provided in the six forward windshields, the two overhead viewing windows, the two aft viewing windows and the side hatch windows; they are described in the window section. Approximately 300 penetrations in the pressure shell are sealed with plates and fittings. A large removable panel in the aft bulkhead provides access to the interior of the crew compartment during initial fabrication and assembly and provides for airlock installation and removal. The compartment supports the environmental control and life support system; avionics; guidance, navigation and control equipment; inertial measurement units; displays and controls; star trackers; and crew accommodations for sleeping, waste management, seats and an optional galley.
The crew compartment is supported within the forward fuselage at only four attach points to minimize the thermal conductivity between them. The two major attach points are located at the aft end of the crew compartment at the flight deck floor level. The vertical load reaction link is on the centerline of the forward bulkhead. The lateral load reaction is on the lower segment of the aft bulkhead.
The compartment is configured to accommodate a crew of four on the flight deck and three in the middeck. In OV-102, four can be accommodated in the middeck. The crew cabin arrangement consists of a flight deck, middeck and lower level equipment bay.
The crew compartment is pressurized to 14.7 psia, plus or minus 0.2 psia, and is maintained at an 80-percent nitrogen and 20-percent oxygen composition by the ECLSS, which provides a shirt-sleeve environment for the flight crew. The crew compartment is designed for 16 psia.
The flight deck is the uppermost compartment of the cabin. The commander's and pilot's work stations are positioned side by side in the forward portion of the flight deck. These stations have controls and displays for maintaining autonomous control of the vehicle throughout all mission phases. Directly behind and to the sides of the commander and pilot centerline are the mission specialist seats.
The commander's and pilot's seats have two shoulder harnesses and a lap belt for restraints. The shoulder harnesses have an inertia reel lock/unlock feature. The unlocked position allows the shoulder harness to move. The commander and pilot can move their seats along the orbiter's Z (vertical) and X (longitudinal) axes so they can reach and see controls better during the ascent and entry phases of flight. Seat movement for each axis is provided by a single ac motor. The total travel distance for the Z and X axes is 10 and 5 inches, respectively. Seat adjustment controls are located on the left side of the seat pan and consist of a three-position toggle switch for power bus selection and one spring-loaded, three-position toggle switch each to control horizontal and vertical seat movement. To operate the seat, the commander and pilot position the pwr buss sel switch to AC2 or AC3 for power; to move the seat along the horizontal axis, the commander and pilot position the horiz contr switch to fwd to move the seat forward and to aft to move the seat aft. Similarly, to move the seat along the vertical axis, the commander and pilot position the vert contr switch to up to move the seat upward and to down to move the seat down. The commander and pilot can position the pwr buss sel switch to off, removing power from the seat. If the seat motors fail, the seat can be adjusted manually. However, manual seat adjustment can only take place on orbit and is accomplished with a special seat adjustment tool provided in the in-flight maintenance tool kit. Manual horizontal and vertical seat adjustment controls are located under the seat pan cushion and on the aft side of the fixed seat structure. The seat adjustment tool is a ratchet-driven, 3/16-inch allen wrench, which is inserted into the vertical or horizontal manual adjustment to move the seat along the Z or X axis. The seats accommodate stowage of in-flight equipment and have removable seat cushions and mounting provisions for oxygen and communications connections to the crew altitude protection system.
Each mission and payload specialist's seat has two shoulder harnesses and a lap belt for restraints. The specialists' seats have controls to manually lock and unlock the tilt of the seat back. Each seat has removable seat cushions and mounting provisions for oxygen and communications connections to the CAPS. The specialists' seats are removed and stowed in the middeck on orbit. No tools are required since the legs of each seat have quick-disconnect fittings. Each seat is 25.5 inches long, 15.5 inches wide and 11 inches high when folded for stowage.
The aft flight deck has two overhead and aft viewing windows for viewing orbital operations. The aft flight deck station also contains displays and controls for executing attitude or translational maneuvers for rendezvous, stationkeeping, docking, payload deployment and retrieval, payload monitoring, remote manipulator system controls and displays, payload bay door operations and closed-circuit television operations.
The forward flight deck, which includes the center console and seats, is approximately 24 square feet. However, the side console controls and displays add approximately 3.5 square feet more. If the center console is subtracted from the 24 square feet, this would amount to approximately 5.2 square feet.
The aft flight deck is approximately 40 square feet.
Directly beneath the flight deck is the middeck. Access to the middeck is through two interdeck openings, which measure 26 by 28 inches. Normally, the right interdeck opening is closed and the left is open. A ladder attached to the left interdeck access allows easy passage in 1-g conditions. The middeck provides crew accommodations and contains three avionics equipment bays. The two forward avionics bays utilize the complete width of the cabin and extend into the middeck 39 inches from the forward bulkhead. The aft bay extends into the middeck 39 inches from the aft bulkhead on the right side of the airlock. Just forward of the waste management system is the side hatch. The completely stripped middeck is approximately 160 square feet; the gross mobility area is approximately 100 square feet.
The side hatch in the middeck is used for normal crew entrance/exit and may be operated from within the crew cabin middeck or externally. It can be jettisoned for emergencies, as discussed in the escape system section. It is attached to the crew cabin tunnel by hinges, a torque tube and support fittings. The hatch opens outwardly 90 degrees down with the orbiter horizontal or 90 degrees sideways with the orbiter vertical. It is 40 inches in diameter and has a 10-inch clear-view window in the center of the hatch. The window consists of three panes of glass. The side hatch has a pressure seal that is compressed by the side hatch latch mechanisms when the hatch is locked closed. A thermal barrier of Inconel wire mesh spring with a ceramic fiber braided sleeve is installed between the reusable surface insulation tiles on the forward fuselage and the side hatch. The total weight of the side hatch is 294 pounds.
Depending on the mission requirements, bunk sleep stations and a galley can be installed in the middeck. In addition, three or four seats of the same type as the mission specialists' seats on the flight deck can be installed in the middeck. Three seats over the normal three could be installed in the middeck for rescue missions if the bunk sleep stations were removed.
The waste management system, located in the middeck, can also accommodate payloads in the pressurized crew compartment environment.
The middeck also provides a stowage volume of 140 cubic feet. Accommodations are included for dining, sleeping, maintenance, exercising and data management. On the orbiter centerline, just aft of the forward avionics equipment bay, an opening in the ceiling provides access to the inertial measurement units.
The middeck floor contains removable panels that provide access to the ECLSS equipment. The middeck equipment bay below the middeck floor houses the major components of the waste management and air revitalization systems, such as pumps, fans, lithium hydroxide, absorbers, heat exchangers and ducting. This compartment has space for stowing lithium hydroxide canisters and five separate spaces for crew equipment stowage with a volume of 29.92 cubic feet.
Modular stowage lockers are used to store the flight crew's personal gear, mission-necessary equipment, personal hygiene equipment and experiments. The modular lockers are made of sandwich panels of Kevlar/epoxy and a non-metallic core. This reduced the lockers' weight by 83 percent compared to all-aluminum lockers, a reduction of approximately 150 pounds. There are 42 identical boxes, which are 11 by 18 by 21 inches.
An airlock, located in the middeck, is composed of machined aluminum sections welded together to form a cylinder with hatch mounting flanges. The upper cylindrical section and bulkheads are constructed of aluminum honeycomb. Two semicylindrical aluminum sections are welded to the airlock's primary structure to house the ECLSS and electrical support equipment. Each semicylindrical section has three feedthrough plates for plumbing and cable routings from the orbiter to the airlock.
Normally, two extravehicular mobility units are stowed in the airlock. The EMU is an integrated space suit assembly and life support system that enables flight crew members to leave the pressurized orbiter crew cabin and work outside the cabin in space.
The airlock is normally located inside the middeck of the spacecraft's pressurized crew cabin. It has an inside diameter of 63 inches, is 83 inches long and has two 40-inch- diameter D-shaped openings that are 36 inches across. It also has two pressure-sealing hatches and a complement of airlock support systems. The airlock's volume is 150 cubic feet.
The airlock is sized to accommodate two fully suited flight crew members simultaneously. Support functions include airlock depressurization and repressurization, extravehicular activity equipment recharge, liquid-cooled garment water cooling, EVA equipment checkout, donning and communications. The EVA gear, checkout panel and recharge stations are located on the internal walls of the airlock.
The airlock hatches are mounted on the airlock. The inner hatch is mounted on the exterior of the airlock (orbiter crew cabin middeck side) and opens into the middeck. The inner hatch isolates the airlock from the orbiter crew cabin. The outer hatch is mounted inside the airlock and opens into the airlock. The outer hatch isolates the airlock from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock to the payload bay when open.
Airlock repressurization is controllable from the orbiter crew cabin middeck and from inside the airlock. It is performed by equalizing the airlock's and cabin's pressure with equalization valves mounted on the inner hatch. The airlock is depressurized from inside the airlock by venting the airlock's pressure overboard. The two D-shaped airlock hatches open toward the primary pressure source, the orbiter crew cabin, to achieve pressure-assist sealing when closed.
Each hatch has six interconnected latches and a gearbox/actuator, a window, a hinge mechanism and hold-open device, a differential pressure gauge on each side and two equalization valves.
The 4-inch diameter window in each airlock hatch is used for crew observation from the cabin/airlock and the airlock/payload bay. The dual window panes are made of polycarbonate plastic and mounted directly to the hatch by means of bolts fastened through the panes. Each hatch window has dual pressure seals, with seal grooves located in the hatch.
Each airlock hatch has dual pressure seals to maintain pressure integrity. One seal is mounted on the airlock hatch and the other on the airlock structure. A leak check quick disconnect is installed between the hatch and the airlock pressure seals to verify hatch pressure integrity before flight.
The gearbox with latch mechanisms on each hatch allows the flight crew to open and close the hatch during transfers and EVA operations. The gearbox and the latches are mounted on the low-pressure side of each hatch; with a gearbox handle installed on both sides to permit operation from either side of the hatch.
Three of the six latches on each hatch are double-acting and have cam surfaces that force the sealing surfaces apart when the latches are opened, thereby acting as crew assist devices. The latches are interconnected with push-pull rods and an idler bell crank that is installed between the rods for pivoting the rods. Self-aligning dual rotating bearings are used on the rods for attachment to the bellcranks and the latches. The gearbox and hatch open support struts are also connected to the latching system by the same rod/bellcrank and bearing system. To latch or unlatch the hatch, the gearbox handle must be rotated 440 degrees.
The hatch actuator/gearbox is used to provide the mechanical advantage to open and close the latches. The hatch actuator lock lever requires a force of 8 to 10 pounds through an angle of 180 deg rees to unlatch the actuator. A minimum rotation of 440 deg rees with a maximum force of 30 pounds applied to the actuator handle is required to operate the latches to their fully unlatched positions.
The hinge mechanism for each hatch permits a minimum opening sweep into the airlock or the crew cabin middeck. The inner hatch (airlock to crew cabin) is pulled or pushed forward to the crew cabin approximately 6 inches. The hatch pivots up and to the right side. Positive locks are provided to hold the hatch in both an intermediate and a full-open position. A spring-loaded handle on the latch hold-open bracket releases the lock. Friction is also provided in the linkage to prevent the hatch from moving if released during any part of the swing.
The outer hatch (airlock to payload bay) opens and closes to the contour of the airlock wall. The hatch is hinged to be pulled first into the airlock and then forward at the bottom and rotated down until it rests with the low-pressure (outer) side facing the airlock ceiling (middeck floor). The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke. The hatch has a hold-open hook that snaps into place over a flange when the hatch is fully open. The hook is released by depressing the spring-loaded hook handle and pushing the hatch toward the closed position. To support and protect the hatch against the airlock ceiling, the hatch incorporates two deployable struts. The struts are connected to the hatch linkage mechanism and are deployed when the hatch linkage is rotated open. When the hatch latches are rotated closed, the struts are retracted against the hatch.
The airlock hatches can be removed in flight from the hinge mechanism using pip pins, if required.
The airlock air circulation system provides conditioned air to the airlock during non-EVA periods. The airlock revitalization system duct is attached to the outside airlock wall at launch. Upon airlock hatch opening in flight, the duct is rotated by the flight crew through the cabin/airlock hatch, installed in the airlock and held in place by a strap holder. The duct has a removable air diffuser cap, installed on the end of the flexible duct, which can adjust the air flow from 216 pounds per hour. The duct must be rotated out of the airlock before the cabin/airlock hatch is closed for airlock depressurization. During the EVA preparation period, the duct is rotated out of the airlock and can be used for supplemental air circulation in the middeck.
To assist the crew member before and after EVA operations, the airlock incorporates handrails and foot restraints. Handrails are located alongside the avionics and ECLSS panels. Aluminum alloy handholds mounted on each side of the hatches have oval configurations 0.75 by 1.32 inches and are painted yellow. They are bonded to the airlock walls with an epoxyphenolic adhesive. Each handrail has a clearance of 2.25 inches between the airlock wall and the handrail to allow the astronauts to grip it while wearing a pressurized glove. Foot restraints are installed on the airlock floor nearer the payload bay side. The ceiling handhold is installed nearer the cabin side of the airlock. The foot restraints can be rotated 360 degrees by releasing a spring-loaded latch and lock in every 90 degrees. A rotation release knob on the foot restraint is designed for shirt-sleeve operation and, therefore, must be positioned before the suit is donned. The foot restraint is bolted to the floor and cannot be removed in flight. It is sized for the EMU boot. The crew member first inserts his foot under the toe bar and then rotates his heel from inboard to outboard until the heel of the boot is captured.
There are four floodlights in the airlock.
If the airlock is relocated to the payload bay from the middeck, it will function in the same manner as in the middeck. Insulation is installed on the airlock's exterior for protection from the extreme temperatures of space.
For Spacelab pressurized module missions, the airlock remains in the crew compartment middeck, and a tunnel adapter that mates with the airlock and the Spacelab tunnel is installed in the payload bay.
The airlock tunnel adapter, hatches, tunnel extension and tunnel permit the flight crew members to transfer from the spacecraft's pressurized middeck crew compartment to Spacelab's pressurized shirt-sleeve environment.
In addition, the airlock, tunnel adapter and hatches permit the EVA flight crew members to transfer from the airlock/tunnel adapter in the space suit assembly into the payload bay without depressurizing the crew cabin and Spacelab.
The Spacelab tunnel and Spacelab are accessed via the tunnel adapter, which is located in the payload bay and is attached to the airlock at orbiter station Xo 576 and the tunnel extension at X o 660. The tunnel adapter has an inside diameter of 63 inches at its widest section and tapers in the cone area at each end to two 40-inch- diameter D-shaped openings 36 inches across. A 40-inch- diameter D-shaped opening 36 inches across is located at the top of the tunnel adapter. Two pressure-sealing hatches are located in the tunnel adapter, one in the upper area of the tunnel adapter and one in the aft end of the tunnel adapter. The tunnel adapter is a welded structure constructed of 2219 aluminum with 2.4- by 2.4-inch exposed structural ribs on the exterior surface and external waffle skin stiffening.
The hatch located on the middeck side of the airlock is mounted on the exterior of the airlock and opens into the middeck. The hatch isolates the airlock from the crew cabin. The hatch located in the tunnel adapter's aft end isolates the tunnel adapter/airlock from the tunnel extension, tunnel and Spacelab. This hatch opens into the tunnel adapter. The hatch located in the tunnel adapter at the upper D-shaped opening isolates the airlock/tunnel adapter from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock/tunnel adapter to the payload bay when open. This hatch opens into the tunnel adapter.
The hinge mechanism for each hatch permits a minimum opening sweep into the tunnel adapter or the spacecraft crew cabin middeck. The airlock crew cabin hatch in the middeck is pulled/pushed forward to the middeck approximately 6 inches. The hatch pivots up and right. Positive locks are provided to hold the latch in both an intermediate and a full-open position. A spring-loaded handle on the latch hold-open bracket releases the lock. Friction is provided in the linkage to prevent the hatch from moving if released during any part of the swing.
The aft hatch is hinged to be pulled first into the tunnel adapter and then forward at the bottom. The top of the hatch is rotated towards the tunnel and downward until the hatch rests with the Spacelab side facing the tunnel adapter floor. The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro.
The upper (EVA) hatch in the tunnel adapter opens and closes to the left wall of the tunnel adapter. The hatch is hinged to be pulled first into the tunnel adapter and then forward at the hinge area and rotated down until it rests against the port wall of the tunnel adapter. The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro.
The hatches can be removed in flight from the hinge mechanisms via pip pins, if required.
The crew compartment, bunk sleep stations (if installed), airlock and modular stowage lockers are built by Rockwell's Space Transportation Systems Division, Downey, Calif. The original crew seat contractor was AMI of Colorado Springs, Colo., but later Rockwell's Space Transportation Systems Division. The Spacelab pressurized module tunnel adapter and tunnel contractor is McDonnell Douglas Astronautics, Huntington Beach, Calif.
The orbiter windows provide visibility for entry, landing and on-orbit operations. For atmospheric flight, the flight crew needs forward, left and right viewing areas. On-orbit mission phases require visibility for rendezvous, docking and payload-handling operations.
The six windows located at the forward flight deck commander and pilot stations provide forward, left and right viewing. The two overhead windows and two payload-viewing windows at the aft station location on the flight deck provide rendezvous, docking and payload viewing. There is also a window in the middeck side hatch.
The six planform-shaped forward windows are the thickest pieces of glass ever produced in the optical quality for see-through viewing. Each consists of three individual panes. The innermost pane is constructed of tempered aluminosilicate glass to withstand the crew compartment pressure. It is 0.625 of an inch thick. Aluminosilicate glass is a low-expansion glass that can be tempered to provide maximum mechanical strength. The exterior of this pane, called a pressure pane, is coated with a red reflector coating to reflect the infrared (heat portion) rays while transmitting the visible spectrum.
The center pane is constructed of low-expansion, fused silica glass because of its high optical quality and excellent thermal shock resistance. This pane is 1.3 inches thick.
The inner and outer panes are coated with a high-efficiency, anti-reflection coating to improve visible light transmission. These windows withstand a proof pressure of 8,600 psi at 240 F and 0.017 relative humidity.
The outer pane is made of the same material as the center pane and is 0.625 of an inch thick. The exterior is uncoated, but the interior is coated with high-efficiency, anti-reflection coating. The outer surface withstands approximately 800 F.
Each of the forward six windows' outer panes measures 42 inches diagonally, and the center and inner panes each measure 35 inches diagonally. The outer panes of the forward six windows are mounted and attached to the forward fuselage. The center and inner panes are mounted and attached to the crew compartment. Redundant seals are employed on each window. No sealing/bonding compounds are used.
The two overhead windows at the flight deck aft station are identical in construction to the six forward windows except for thickness. The inner and center panes are 0.45 of an inch thick, and the outer pane is 0.68 of an inch thick. The outer pane is attached to the forward fuselage, and the center and inner panes are attached to the crew compartment. The two overhead windows' clear view area is 20 by 20 inches. The left-hand overhead window provides the crew members with a secondary emergency egress. The inner and center panes open into the crew cabin, and the outer pane is jettisoned up and over the top of the orbiter. This provides a secondary emergency exit area of 20 by 20 inches.
On the aft flight deck, each of the two windows for viewing the payload bay consists of only two panes of glass, which are identical to the forward windows' inner and center panes. The outer thermal panes are not installed. Each pane is 0.3 of an inch thick. The windows are 14.5 by 11 inches. Both panes are attached to the crew compartment.
The side hatch viewing window consists of three panes of glass identical to the six forward windows. The inner pane is 11.4 inches in diameter and 0.25 of an inch thick. The center pane is 11.4 inches in diameter and 0.5 of an inch thick. The outer pane is 15 inches in diameter and 0.3 of an inch thick.
During orbital operations, the large window areas of transparency expose the flight crew to sun glare; therefore, window shades and filters are provided to preclude or minimize exposure. Shades are provided for all windows, and filters are supplied for the aft and overhead viewing windows. The window shades and filters are stored in the middeck of the orbiter crew compartment. Attachment mechanisms and devices are provided for their installation at each window on the flight deck.
The forward station window shades (W-1 through W-6) are fabricated from Kevlar/epoxy glass fabric with silver and Inconel-coated Teflon tape on the outside surface and paint on the inside surface. When the shade is installed next to the inner window pane, a silicone rubber seal around the periphery deforms to prevent light leakage. The shade is held in place by the shade installation guide, the hinge plate and the Velcro keeper.
The overhead window shades (W-7 and W-8) are nearly the same as the forward shades; but the rubber seal is deleted, and the shade is sealed and held in place by a separate seal around the window opening, a hinge plate and secondary frame, and Velcro retainer. The overhead window filters are fabricated from Lexan and are used interchangeably with the shades.
The aft window shades (W-9 and W-10) are the same as the overhead window shades except that a 0.63-inch-wide strip of Nomex Velcro has been added around the perimeter of the shade. The shade is attached to the window by pressing the Velcro strip to the pile strip around the window opening. The aft window filters are the same as the overhead window filters except for the addition of the Velcro hook strip. The filters and shades are used interchangeably.
The side hatch window cover is permanently attached to the window frame and is hinged to allow opening and closing.
The contractor for the windows is Corning Glass Co., Corning, N.Y.
The wing is an aerodynamic lifting surface that provides conventional lift and control for the orbiter. The left and right wings consist of the wing glove; the intermediate section, which includes the main landing gear well; the torque box; the forward spar for mounting the reusable reinforced carbon-carbon leading edge structure thermal protection system; the wing/elevon interface; the elevon seal panels; and the elevons.
The wing is constructed of conventional aluminum alloy with a multirib and spar arrangement with skin-stringer-stiffened covers or honeycomb skin covers. Each wing is approximately 60 feet long at the fuselage intersection and has a maximum thickness of 5 feet.
The forward wing box is an extension of the basic wing that aerodynamically blends the wing leading edge into the midfuselage wing glove. The forward wing box is a conventional design of aluminum ribs, aluminum tubes and tubular struts. The upper and lower wing skin panels are stiffened aluminum. The leading edge spar is constructed of corrugated aluminum.
The intermediate wing section consists of the conventional aluminum multiribs and aluminum tubes. The upper and lower skin covers are constructed of aluminum honeycomb. A portion of the lower wing surface skin panel includes the main landing gear door. The intermediate section houses the main landing gear compartment and reacts a portion of the main landing gear loads. A structural rib supports the outboard main landing gear door hinges and the main landing gear trunnion and drag link. The support for the inboard main landing gear trunnion and drag link attachment is provided by the midfuselage. The main landing gear door is conventional aluminum honeycomb.
The four major spars are constructed of corrugated aluminum to minimize thermal loads. The forward spar provides the attachment for the thermal protection system reusable reinforced carbon-carbon leading edge structure. The rear spar provides the attachment interfaces for the elevons, hinged upper seal panels, and associated hydraulic and electrical system components. The upper and lower wing skin panels are stiffened aluminum.
The elevons provide orbiter flight control during atmospheric flight. The two-piece elevons are conventional aluminum multirib and beam construction with aluminum honeycomb skins for compatibility with the acoustic environment and thermal interaction. The elevons are divided into two segments for each wing, and each segment is supported by three hinges. The elevons are attached to the flight control system hydraulic actuators at points along their forward extremities, and all hinge moments are reacted at these points. Each elevon travels 40 degrees up and 25 degrees down.
The transition area on the upper surface between the torque box and the movable elevon consists of a series of hinged panels that provide a closeout of the wing-to-elevon cavity. These panels are of Inconel honeycomb sandwich construction outboard of wing station Y w 312.5 and of titanium honeycomb sandwich construction inboard of wing station Y w 312.5. The upper leading edge of each elevon incorporates titanium rub strips. The rub strips are of titanium honeycomb construction and are not covered with the thermal protection system reusable surface insulation. They provide the sealing surface area for the elevon seal panels.
The wing is attached to the fuselage with a tension bolt splice along the upper surface. A shear splice along the lower surface in the area of the fuselage carry-through completes attachment interface.
Prior to the manufacturing of the wings for Discovery (OV-103) and Atlantis (OV-104), a weight reduction program resulted in a redesign of certain areas of the wing structure. An assessment of wing air loads was made from actual flight data that indicated greater loads on the wing structure. As a result, to maintain positive margins of safety during ascent, structural modifications were incorporated into certain areas of the wings. The modifications consisted of the addition of doublers and stiffeners.
The midfuselage structure interfaces with the forward fuselage, aft fuselage and wings. It supports the payload bay doors, hinges, tie-down fittings, forward wing glove, and various orbiter system components and forms the payload bay area.
The forward and aft ends of the midfuselage are open, with reinforced skin and longerons interfacing with the bulkheads of the forward and aft fuselages. The midfuselage is primarily an aluminum structure 60 feet long, 17 feet wide and 13 feet high. It weighs approximately 13,502 pounds.
The midfuselage skins are integrally machined by numerical control. The panels above the wing glove and the wings for the forward eight bays have longitudinal T-stringers. The five aft bays have aluminum honeycomb panels. The side skins in the shadow of the wing are also numerically control machined but have vertical stiffeners.
Twelve main-frame assemblies stabilize the midfuselage structure. The assemblies consist of vertical side elements and horizontal elements. The side elements are machined; whereas the horizontal elements are boron/aluminum tubes with bonded titanium end fittings, which reduced the weight by 49 percent (approximately 305 pounds).
In the upper portion of the midfuselage are the sill and door longerons. The machined sill longerons not only make up the primary body-bending elements, but also take the longitudinal loads from payloads in the payload bay. The payload bay door longerons and associated structure are attached to the 13 payload bay door hinges. These hinges provide the vertical reaction from the payload bay doors. Five of the hinges react the payload bay door shears. The sill longeron also provides the base support for the payload bay manipulator arm (if installed) and its stowage provisions, the Ku-band rendezvous antenna, the antenna base support and its stowage provisions, and the payload bay door actuation system.
Plumbing and wiring in the lower portion of the midfuselage are supported by fiberglass milk stools.
Because of additional detailed analysis of actual flight data concerning descent stress thermal gradient loads, torsional straps were added to the lower midfuselage stringers in bays 1 through 11. The torsional straps tie all stringers together similarly to a box section, which eliminates rotational (torsional) capabilities to provide positive margins of safety.
Also, because of additional detailed analysis of actual flight data during descent, room-temperature vulcanizing silicone rubber material was bonded to the lower midfuselage from bay 4 through 12 to act as a heat sink and distribute temperatures evenly across the bottom of the midfuselage, which will reduce thermal gradients and ensure positive margins of safety.
The contractor for the midfuselage is General Dynamics Corp., Convair Aerospace Division, San Diego, Calif.
The payload bay doors are opened shortly after orbit is achieved to allow exposure of the environmental control and life support system radiators for heat rejection of the orbiter's systems. The payload bay doors consist of port and starboard doors hinged at each side of the midfuselage and latched mechanically at the forward and aft fuselage and at the split-top centerline. Thermal seals on the doors provide a relatively air-tight payload compartment when the doors are closed and latched. During prelaunch and postlanding, the purge, vent and drain system permits purging of undesirable gases and maintains a positive delta pressure for venting of payloads within the payload area when the doors are closed.
The port and starboard doors are 60 feet long with a combined area of approximately 1,600 square feet. Each door is made up of five segments that are interconnected by circumferential expansion joints. Each door hinges on 13 Inconel 718 external hinges (five shear and eight idlers). The lower half of each hinge attaches to the midfuselage sill longeron. The hinges rotate on bearings with dual rotational surfaces. There are five shear hinges and eight floating hinges. The floating hinges allow fore and aft movement of the door panels for thermal expansion.
Each door actuation system provides the mechanism to drive each door side to the open or closed position. Each mechanism consists of an electromechanical power drive unit and six rotary gear actuators, which are connected by torque tubes to each other and to the power drive unit. Linkages transmit torque from the rotary actuators to the doors.
The forward 30-foot sections of both doors incorporate radiators that can be deployed; they are hinged and latched to the door inner surface in order to reject the excess heat of the Freon-21 coolant loops from both sides of the radiator panels when the doors are open. An electromechanical actuation system on the door unlatches and deploys the radiators when open and latches and stows the radiators when closed. The radiators may be left in the stowed position for a given flight and will only radiate the excess heat from the one side. Fixed radiator panels are installed on the forward end of the aft payload bay doors and radiate from one side only. Kitted fixed radiator panels may be installed on the aft end of the aft payload bay doors when required by a specific mission; they also will radiate from only one side.
When the payload bay doors are closed, they are fixed at the aft fuselage bulkhead and allowed to move longitudinally at the forward fuselage. The doors also accommodate vehicle torsional loads (a force that causes a body, such as a shaft, to twist about its longitudinal axis), aerodynamic pressure loads and payload bay vent lag pressures. The payload bay is not a pressurized area.
Thermal and pressure seals are used to close the gaps at the forward and aft fuselage interface, door centerline and circumferential expansion joints.
The doors are 60 feet long. Each consists of five segments interconnected by expansion joints. The chord of each half of these curved doors is approximately 10 feet, and the doors are 15 feet in diameter.
The doors are constructed of graphite epoxy composite material, which reduces the weight by 23 percent over that of aluminum honeycomb sandwich. This is a reduction of approximately 900 pounds, which brings the weight of the doors down to approximately 3,264 pounds. The payload bay doors are the largest aerospace structure to be constructed from composite material.
The composite doors will withstand 163-decibel acoustic noise and a temperature range of minus 170 to plus 135 F.
The doors are made up of subassemblies consisting of graphite epoxy honeycomb sandwich panels, solid graphite epoxy laminate frames, expansion joint frames, torque box, seal depressor, centerline beam intercostals, gussets, end fittings and clips. There are also aluminum 2024 shear pins, titanium fittings, and Inconel 718 floating and shear hinges. The assembly is joined by mechanical fasteners. Lightning strike protection is provided by aluminum mesh wire bonded to the outer skin.
Extravehicular activity handholds are attached in the torque box areas.
The payload bay doors are covered with reusable surface insulation.
The left door with attached systems weighs approximately 2,375 pounds and the right weighs about 2,535 pounds. The right door contains the centerline latch active mechanisms, which accounts for the weight difference. These weights do not include the radiator panel system, which adds 833 pounds per door.
The PL bay door open/stop/close switch on panel R13 initiates the payload bay door power and control system through the aft flight deck data processing system, general-purpose computer and associated cathode ray tube display and keyboard. The normal operational mode for opening and closing the payload bay door bulkhead latches, centerline latches and payload bay doors is through keyboard entries in an automatic mode in which the latches are cycled and the doors controlled in a predetermined sequence. If a problem occurs in the predetermined automatic sequence, a manual keyboard capability permits selection of automatic sequence groupings that can be commanded individually. The open position of the switch on panel R13 provides the signals to a GPC to initiate and sustain the automatic or keyboard manual bulkhead latches and door opening sequence. The close position accomplishes the same as the open position except for the closing sequence. The stop position removes the open and close signals, stopping the sequence in progress.
The PL bay door talkback indicator on panel R13 is functional only in the automatic sequence and would remain in its initial state in the manual keyboard mode. The signal source for the talkback indicator is a combination of ready-to-latch and door-open limit switch inputs that are processed by software to establish the talkback indicator state. The talkback indicator indicates op when the bulkhead latches, centerline latches and doors are open; cl when the doors are closed and the centerline and bulkhead latches are closed; and barberpole when the bulkhead and centerline latches and doors are in transit or are stopped between open and closed.
When closed, the doors are latched to the forward and aft bulkheads and along the upper centerline of the doors. The latching system consists of 16 bulkhead latches (eight aft and eight forward) and 16 payload bay door centerline latches. The forward and aft bulkhead latches are in groups of four ganged latch hooks. The centerline latches are also in groups of four ganged latches. Each centerline latch gang incorporates four latches, bellcranks, push rods, levers, rollers and an electromechanical actuator.
The forward and aft bulkhead latches are arranged in groups of four ganged latches. Each group is opened or closed by an electromechanical actuator with two redundant, three-phase ac reversible motors that receive ac power from mid motor controller assemblies when commanded in the automatic predetermined sequence or by manual keyboard entries. In the automatic mode, the forward and aft bulkhead latches operate simultaneously.
The forward and aft bulkhead latch groups consist of two ac reversible motors. These groups also control an actuator output arm, which positions active latch mechanisms and disengages or engages four latch hooks on four corresponding passive rollers on the bulkhead. The two ac motors of a bulkhead latch group are commanded through limit switches to open or close that group of latches. When the ac motor is in operation, the brake associated with that motor is released and is applied when power is removed from the motor. The limit switches apply or remove ac power from the motor when that latch group reaches its open or closed position. When both motors are operating, the latch group operating time is 30 seconds; it is 60 seconds when only one motor is operating. In addition, each MMCA has its own timer set to twice the normal operating time to allow for single-motor operation of a bulkhead latch group without causing a sequence fail signal PLB doors CRT message and SM alert.
During latching operations for a bulkhead group, the payload bay door comes in contact with a bulkhead switch module striker when the door is nearly closed. A two-out-of-three voting logic of the ready-to-latch switches precludes premature start signals to the bulkhead latch motors. The ready-to-latch switch then activates the bulkhead latch ac motors, which latches the door closed. The door-closed limit switches turn the ac motors off. The limit switch contact closures are sent to the CRT display under micro-sw stat (switch status), which permits the flight crew to observe the change in the status of the microswitches. Telemetry can also monitor the microswitch status. Torque limiters in each bulkhead latch group permit slippage if a limit switch fails to turn off the ac motors or the mechanism jams during latching operations in order to prevent damage to the motors or mechanisms. Extravehicular activity disconnects are provided to permit an EVA flight crew member to close the door latch manually from inside the payload bay if the mechanism jams when the doors close.
The payload bay door centerline latch groups are controlled automatically in a predetermined sequence or manually by individual latch groups through keyboard entries in a manner similar to the bulkhead latch groups. The 16 centerline latches are arranged into groups of four, similar to the bulkhead latches.
Each centerline latch group consists of two ac reversible electric motors that drive a rotary shaft and bellcrank and four hooks to engage a corresponding passive roller to latch the door closed or disengage the passive roller to unlatch the door. All 16 centerline hook assemblies contain alignment rollers to eliminate payload bay door overlap due to thermal distortion. Passive shear fittings in each centerline latch group align door closure and cause the fore and aft shear loads to react once the doors are closed.
The centerline latch group ac reversible motors are automatically turned off by limit switches when the latches are opened or closed. Each motor has a brake that operates similarly to the brakes in the bulkhead motors. When both motors are operating, the nominal operating time is 20 seconds. If only one motor is operating, the time is 40 seconds. Each mid motor controller assembly has its own timer set to twice the normal operating time to allow single-motor operation of the centerline latch group without causing a sequence fail signal PLB doors CRT message and SM alert.
Torque limiters in the centerline latch groups allow slippage if limit switches fail to turn off an electrical drive motor or the mechanisms jam to prevent damage to the motors or mechanism.
EVA disconnects in a centerline latch group can be used to isolate a jammed latch from the group.
The payload bay doors are driven by a rotary actuator consisting of two electrical three-phase reversible ac motors per power drive unit. There is one power drive unit for right doors and one for the left doors.
The power drive unit drives a 55-foot-long torque shaft. The shaft turns the rotary actuators, which causes the push rod, bell crank and link to push the doors open. The same arrangement pulls the doors closed.
The payload bay door opening and closing sequence is controlled automatically through in a predetermined sequence or manually through keyboard entries. The starboard doors must be opened first and closed last due to the arrangement of the centerline latching mechanism and the structural and seal overlap. Limit switches on each power drive unit turn the ac motors off when the doors are open or closed. Each ac motor has an associated brake that operates similarly to the bulkhead and centerline latch motors. When both motors are operating, the nominal time for payload bay door opening or closing is 63 seconds. If only one motor is operating, the time is 126 seconds. Each MMCA has its own timer set to twice the normal operating time to allow single-motor operation of the payload bay doors without causing a sequence fail signal PLB doors CRT message and an SM alert .
Torque limiters are incorporated into the rotary actuators to avoid damaging the drive motors or mechanisms if limit switches fail to turn off an electrical drive motor or the mechanisms jam.
Two bolts on the bellcrank and the bolt connecting the link to the rotary actuator can be EVA disconnect points if the linkage fails when the doors close. The power drive unit can be disengaged manually on the ground or on orbit.
The payload bay doors open through an angle of 175.5 degrees.
Two radiator panels on each forward payload bay door can be deployed when the doors are opened on orbit and stowed when the doors are closed before entry, or they can be left in the stowed position for a given flight. Freon-21 coolant loop 1 flows through the left-hand radiator panels, and the No. 2 loop flows through the right-hand panels. On orbit, the panels radiate excess heat collected by the Freon-21 coolant loops from heat exchangers and cold plates throughout the orbiter. Coolant flows through the radiators from aft to forward. The radiator panels mounted on the forward end of the aft payload bay doors are fixed to the bay doors.
The radiator deploy and stow operation is controlled manually from the aft flight deck panel R13. The PL bay mech (payload bay mechanisms) pwr, radiator latch and radiator control sys switches control the panels. Four indicators show the radiator latch and deploy status.
When the payload bay doors are fully open, the PL bay mech sys 1 and sys 2 switches are positioned to on . The sys 1 and sys 2 switches positioned to on provide ac bus power to both right- and left-side radiator latch control actuators.
The radiator latch control sys A switch positioned to release applies ac power to one ac reversible drive motor on each starboard and port panel. When each motor is in operation, the brake is removed. Each ac drive motor rotates a torque shaft, which operates push rods that unlatch six latches on each of the two right and two left radiator panels. The linkages and latches are attached to the payload bay doors, and passive rollers are attached to the radiator panels. The operating time for releasing the latches with one motor is approximately 52 seconds. Limit switches remove power from the ac motors. The brake is applied for each motor. The radiator stdb (starboard) and port talkback indicators above the latch control sys A and B switches indicate rel when the corresponding latches are released and barberpole when in transit. When the radiator latch control sys B switch is positioned to release, ac power is applied to the remaining ac reversible drive motor on each right panel and each left panel. This remaining ac drive motor will operate the same rotating shaft and unlatch the same six latches on each of the two right and two left radiator panels. Separate limit switches remove power from these ac motors. The radiator stdb and port talkback indicators above the latch control sys A and B switches indicate rel when the corresponding latches are released and barberpole when in transit and have the same operating time as in system A. If both switches were positioned to rel simultaneously, the operating time would be approximately 26 seconds.
Positioning the radiator latch control sys A and/or B switch to latch reverses the action and latches the radiator panels to the payload bay doors. The talkback indicators indicate lat when the panels are latched and barberpole when in transit.
The off position of the radiator latch control sys A and/or B switch removes power from the corresponding control system, which stops the motors and latch system movement.
Torque limiters in the power drive system prevent damage to the system in the event of jamming or binding during operation.
The radiator control sys A switch positioned to deploy applies ac power to one ac reversible drive motor on the right panel and one ac reversible drive motor on the left panel. The motors are not operable until the MMCAs have received two signals from the radiator panel unlatch drives, which prevents inadvertent deployment of the radiators while still latched. When power is applied to the left and right motors, the brake is removed and the rotary actuator shaft rotates and pushes the respective radiator panels away from the payload bay doors to the deployed position. Separate limit switches turn the ac motors off and apply the brake for each motor. The operating time for deployment with one motor is 86 seconds. The radiator stdb and port talkback indicators above the radiator control sys A and B switches indicate dpy when the corresponding panels are deployed and barberpole when in transit. When the radiator control sys B switch is positioned to deploy, ac power is applied to the remaining ac reversible drive motor on the right radiator panel and the remaining ac reversible drive motor on the left panel. This remaining ac drive motor operates the same rotary actuator shaft and pushes the respective radiator panel away from the payload bay doors to the deployed position. Separate limit switches turn the ac motors off. The radiator stdb and port talkback indicators above the radiator control sys A and B switches indicate dpy when the corresponding panels are deployed and barberpole when in transit and have the same operating time as in system A. If both switches are positioned to deploy simultaneously, the operating time is 43 seconds.
Positioning the radiator control sys A and/or B switch to stow reverses the action, stowing the radiators to the payload bay doors. The talkback indicators indicate sto when the panels are stowed and barberpole in transit.
The off position of the radiator control sys A and/or B switch removes power from the corresponding control system, stopping the motors and radiator panel movement.
When the radiators are deployed, they are 35.5 degrees from the payload bay doors.
Torque limiters on the power drive system prevent damage to the system in the event of jamming or binding during operation.
Each rotary crank can be disengaged from the rotary actuator (via EVA operations) by retracting a shear pin. Retraction allows the crank to rotate around an alternate pivot and permits the crew to stow the panels if the system fails. If the power drive unit fails, all four shear pins must be removed to allow manual stowing of the radiators. The pins are accessible when the radiators are fully deployed. No disengagement is planned if the radiators fail to deploy.
The contractors are Rockwell's Tulsa Division, Tulsa, Okla. (payload bay doors); Curtiss Wright, Caldwell, N.J. (payload bay door power drive unit, rotary actuators, drive shafts, torque tubes and couplings, radiator deploy/latch actuator and latch mechanism); Hoover Electric, Los Angeles, Calif. (payload bay door electromechanical rotary actuators); Vought Corp., Dallas, Texas (radiators); Rockwell's Space Transportation Systems Division, Downey, Calif. (latches, linkages and actuators).
The aft fuselage consists of an outer shell, thrust structure and internal secondary structure. It is approximately 18 feet long, 22 feet wide and 20 feet high.
The aft fuselage supports and interfaces with the left-hand and right-hand aft orbital maneuvering system/reaction control system pods, the wing aft spar, midfuselage, orbiter/external tank rear attachments, space shuttle main engines, aft heat shield, body flap, vertical tail and two T-0 launch umbilical panels.
The aft fuselage provides the load path to the midfuselage main longerons, main wing spar continuity across the forward bulkhead of the aft fuselage, structural support for the body flap, and structural housing around all internal systems for protection from operational environments (pressure, thermal and acoustic) and controlled internal pressures during flight.
The forward bulkhead closes off the aft fuselage from the midfuselage and is composed of machined and beaded sheet metal aluminum segments. The upper portion of the bulkhead attaches to the front spar of the vertical tail.
The internal thrust structure supports the three SSMEs. The upper section of the thrust structure supports the upper SSME, and the lower section of the thrust structure supports the two lower SSMEs. The internal thrust structure includes the SSMEs, load reaction truss structures, engine interface fittings and the actuator support structure. It supports the SSMEs, the SSME low-pressure turbopumps and propellant lines. The two orbiter/external tank aft attach points interface at the longeron fittings.
The internal thrust structure is composed mainly of 28 machined, diffusion-bonded truss members. In diffusion bonding, titanium strips are bonded together under heat, pressure and time. This fuses the titanium strips into a single hollow, homogeneous mass that is lighter and stronger than a forged part. In looking at the cross section of a diffusion bond, one sees no weld line. It is a homogeneous parent metal, yet composed of pieces joined by diffusion bonding. (In OV-105, the internal thrust structure is a forging.) In selected areas, the titanium construction is reinforced with boron/epoxy tubular struts to minimize weight and add stiffness. This reduced the weight by 21 percent, approximately 900 pounds.
The upper thrust structure of the aft fuselage is of integral-machined aluminum construction with aluminum frames except for the vertical fin support frame, which is titanium. The skin panels are integrally machined aluminum and attach to each side of the vertical fin to react drag and torsion loading.
The outer shell of the aft fuselage is constructed of integral-machined aluminum. Various penetrations are provided in the shell for access to installed systems. The exposed outer areas of the aft fuselage are covered with reusable thermal protection system.
The secondary structure of the aft fuselage is of conventional aluminum construction except that titanium and fiberglass are used for thermal isolation of equipment. The aft fuselage secondary structures consist of brackets, buildup webs, truss members, and machined fittings, as required by system loading and support constraints. Certain system components, such as the avionics shelves, are shock-mounted to the secondary structure. The secondary structure includes support provisions for the auxiliary power units, hydraulics, ammonia boiler, flash evaporator and electrical wire runs.
The two external tank umbilical areas interface with the orbiter's two aft external tank attach points and the external tank's liquid oxygen and hydrogen feed lines and electrical wire runs. The umbilicals are retracted, and the umbilical areas are closed off after external tank separation by an electromechanically operated beryllium door at each umbilical. Thermal barriers are employed at each umbilical door. The exposed area of each closed door is covered with reusable surface insulation.
The aft fuselage heat shield and seal provide a closeout of the orbiter aft base area. The aft heat shield consists of a base heat shield of machined aluminum. Attached to the base heat shield are domes of honeycomb construction that support flexible and sliding seal assemblies. The engine-mounted heat shield is of Inconel honeycomb construction and is removable for access to the main engine power heads. The heat shield is covered with a reusable thermal protection system except for the Inconel segments.
The orbital maneuvering system/reaction control system left- and right-hand pods are attached to the upper aft fuselage left and right sides. Each pod is fabricated primarily of graphite epoxy composite and aluminum. Each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod is divided into two compartments: the OMS and the RCS housings. Each pod houses all the OMS and RCS propulsion components and is attached to the aft fuselage with 11 bolts. The pod skin panels are graphite epoxy honeycomb sandwich. The forward and aft bulkhead aft tank support bulkhead and floor truss beam are machined aluminum 2124. The centerline beam is 2024 aluminum sheet with titanium stiffeners and graphite epoxy frames. The OMS thrust structure is conventional 2124 aluminum construction. The cross braces are aluminum tubing, and the attach fittings at the forward and aft fittings are 2124 aluminum. The intermediate fittings are corrosion-resistant steel. The RCS housing, which attaches to the OMS pod structure, contains the RCS thrusters and associated propellant feed lines. The RCS housing is constructed of aluminum sheet metal, including flat outer skins. The curved outer skin panels are graphite epoxy honeycomb sandwich. Twenty-four doors in the skins provide access to the OMS and RCS and attach points.
The two graphite epoxy pods per spacecraft reduce the weight by 10 percent, approximately 450 pounds. The pods will withstand 162-decibel acoustic noise and a temperature range from minus 170 to plus 135 F.
The exposed areas of the OMS/RCS pods are covered with a reusable thermal protection system, and a pressure and thermal seal is installed at the OMS/RCS pod aft fuselage interface. Thermal barriers are installed, and they interface with the RCS thrusters and reusable thermal protection system.
The body flap thermally shields the three SSMEs during entry and provides the orbiter with pitch control trim during its atmospheric flight after entry.
The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. The main upper and lower forward honeycomb skin panels are joined to the ribs, spars and honeycomb trailing edge with structural fasteners. The removable upper forward honeycomb skin panels complete the body flap structure.
The upper skin panels aft of the forward spar and the entire lower skin panels are mechanically attached to the ribs. The forward upper skin consists of five removable access panels attached to the ribs with quick-release fasteners. The four integral-machined aluminum actuator ribs provide the aft fuselage interface through self-aligning bearings. Two bearings are located in each rib for attachment to the four rotary actuators located in the aft fuselage, which are controlled by the flight control system and the hydraulically actuated rotary actuators. The remaining ribs consist of eight stability ribs and two closeout ribs constructed of chemically milled aluminum webs bonded to aluminum honeycomb core. The forward spar web is of chemically milled sheets with flanged holes and stiffened beads. The spar web is riveted to the ribs. The trailing edge includes the rear spar, which is composed of piano-hinge half-cap angles, chemically milled skins, honeycomb aluminum core, closeouts and plates. The trailing edge attaches to the upper and lower forward panels by the piano-hinge halves and hinge pins. Two moisture drain lines and one hydraulic fluid drain line penetrate the trailing edge honeycomb core for horizontal and vertical drainage.
The body flap is covered with a reusable thermal protection system and an articulating pressure and thermal seal to its forward cover area on the lower surface of the body flap to block heat and air flow from the structures.
The aft fuselage is built by Rockwell's Space Transportation Systems Division, Downey, Calif. The OMS/RCS pods are built by McDonnell Douglas, St. Louis, Mo. The body flap is built by Rockwell's Columbus, Ohio, division.
The vertical tail consists of a structural fin surface, the rudder/speed brake surface, a tip and a lower trailing edge. The rudder splits into two halves to serve as a speed brake.
The vertical tail structure fin is made of aluminum. The main torque box is constructed of integral-machined skins and strings, ribs, and two machined spars. The fin is attached by two tension tie bolts at the root of the front spar of the vertical tail to the forward bulkhead of the aft fuselage and by eight shear bolts at the root of the vertical tail rear spar to the upper structural surface of the aft fuselage.
The rudder/speed brake control surface is made of conventional aluminum ribs and spars with aluminum honeycomb skin panels and is attached through rotating hinge parts to the vertical tail fin.
The lower trailing edge area of the fin, which houses the rudder/speed brake power drive unit, is made of aluminum honeycomb skin.
The hydraulic power drive unit/mechanical rotary actuation system drives left- and right-hand drive shafts in the same direction for rudder control of plus or minus 27 degrees. For speed brake control, the drive shafts turn in opposite directions for a maximum of 49.3 degrees each. The rotary drive actions are also combined for joint rudder/speed brake control. The hydraulic power drive unit is controlled by the orbiter flight control system.
The vertical tail structure is designed for a 163-decibel acoustic environment with a maximum temperature of 350 F.
All-Inconel honeycomb conical seals house the rotary actuators and provide a pressure and thermal seal that withstands a maximum of 1,200 F.
The split halves of the rudder panels and trailing edge contain a thermal barrier seal.
The contractor for the vertical tail and rudder/speed brake is Fairchild Republic, Farmingdale, N.Y.
A passive thermal control system helps maintain the temperature of the orbiter spacecraft, systems and components within their temperature limits. This system uses available orbiter heat sources and heat sinks supplemented by insulation blankets, thermal coatings and thermal isolation methods. Heaters are provided on components and systems in areas where passive thermal control techniques are not adequate. (The heaters are described under the various systems.)
The insulation blankets are of two basic types: fibrous bulk and multilayer. The bulk blankets are fibrous materials with a density of 2 pounds per cubic foot and a sewn cover of reinforced acrylic film Kapton. The cover material has 13,500 holes per square foot for venting. Acrylic film tape is used for cutouts, patching and reinforcements. Tufts throughout the blankets minimize billowing during venting.
The multilayer blankets are constructed of alternate layers of perforated acrylic film Kapton reflectors and Dacron net separators. There are 16 reflector layers in all, the two cover halves counting as two layers. Covers, tufting and acrylic film tape are similar to that used for the bulk blankets.
The contractors are Hi-Temp Insulation Inc., Camarillo, Calif. (fibrous insulation); Scheldahl, Northfield, Minn. (cover materials and inner layers); Apex Mills, Los Angeles, Calif. (separators).
The purge, vent and drain system on the orbiter is designed to perform the following functions: provide unpressurized compartments with gas purge for thermal conditioning and prevent accumulation of hazardous gases, vent the unpressurized compartments during ascent and entry, drain trapped fluids (water and hydraulic fluid) and condition window cavities to maintain visibility.
Three purge circuits are connected by the T-0 umbilical to ground equipment before launch during the preflight countdown and postlanding phases. Purge gas (cool, dry air and gaseous nitrogen) is provided to three sets of distribution plumbing: the forward fuselage, orbital maneuvering system/reaction control system pods, wings and vertical stabilizer; the midfuselage; and the aft fuselage. The purge gas makes all the unpressurized volumes inert, maintains constant humidity and temperature, forces out any hazardous gases and ensures that external contaminants cannot enter.
The vent and purge system is controlled exclusively through guidance, navigation and control software. The active ports are positioned by the software on the basis of mission time or mission events during ascent, entry and aborts and by crew inputs on the CRT and keyboard in the crew compartment flight deck.
There are 18 active vents in the orbiter fuselage, nine on each side. Each vent has a door that can be positioned for a specific purpose at various phases of flight. For identification, each door is numbered, starting at the nose of the orbiter. Each compartment has a dedicated vent on the left and right side of the orbiter for redundancy.
Internal vents are used to vent compartments that have no vent doors of their own, such as the nose wheel well, the two main wheel wells and the vertical tail section. Passive vents are used to back up vent 7 of the forward wing compartment, which responds to a delta pressure to open a check valve (passive vent) during ascent to vent the wing to the midbody if vent 7 fails; or, on descent, the midfuselage pressurizes the wing if vent 7 fails at a delta pressure of 0.72 to 1 psid. The aft bulkhead (X o 1307) has 14 one-way check valves that vent the payload bay into the aft fuselage at a delta pressure of 0.004 to 0.04 psid. Vent 8 vents the OMS/RCS pods, which are joined by a duct that enables the pod to vent through the opposite side of the vehicle if vent 8 fails to open.
All vent doors are driven by an electromechanical actuator. Vent doors located near each other share common actuators and controls. Vents 1 and 2, 4 and 7, and 8 and 9 share drive mechanisms on the left and right side. The 18 doors are divided into six groups of four ac motors each and are staggered so that all 24 motors do not run at the same time. All vent doors are driven inward, and each door has a pressure seal and thermal seal. The normal opening or closing time of a door with two motors operating is five seconds.
Vent doors 1, 2, 8 and 9 have purge positions that control flow from the forward and aft volumes, respectively. Vent 6 has two purge positions and a closed position that accommodates the different purge flow rates available to the payloads and payload bay. These doors are in the purge position before launch.
Two minutes 20 seconds before launch, the launch processing system reduces the purge flow in anticipation of closing vent 6. At T minus 35 seconds, vent 6 is closed. At T minus 25 seconds, the onboard general-purpose computers are enabled and take over the sequences.
At T minus 10 seconds, the vents are configured for launch. Vents 3, 4, 5, 6 and 7 are closed to limit sound pressure levels in the payload bay. Vents 1, 2, 8 and 9 are opened. If a launch abort occurs from T minus 10 seconds to T minus zero, the vent doors reposition to the prelaunch configuration. At T minus four seconds, any vent door out of configuration and not overridden causes the onboard GPCs to call a hold.
At T plus 10 seconds, all vent doors are commanded open. At T plus 80 seconds, vent doors 8 and 9 are commanded closed to prevent hazardous gases from entering the aft fuselage; at T plus 122 seconds, vents 8 and 9 are commanded open. All vent doors remain open during the remainder of ascent and on-orbit operations.
In preparation for entry, the onboard operational sequence software (OPS 3) closes all vent doors. The doors remain closed until the velocity of the orbiter reaches 2,400 feet per second, when all vents are opened by the onboard GPCs.
At the end of the mission, after the orbiter stops on the runway, vent doors 1, 2, 6, 8 and 9 are configured to their purge positions for ground cooling.
The purge and vent ducting is now made of Kevlar/epoxy (115 pieces up to 11 inches in diameter), which replaced the fiberglass or aluminum ducts and reduced the weight of the ducts 33 percent, or approximately 200 pounds.
The window cavity conditioning system prevents moisture from entering into the windshields and the cavities of the overhead and payload-viewing windows. It also depressurizes and repressurizes these cavities during flight and supplies the purge conditioning to dry them during ground operations. The side hatch window is self-contained.
A hazardous gas detection system detects hazardous levels of explosive or toxic gases. The onboard orbiter sample lines duct the compartment gases to the ground support equipment at the T-0 right-hand umbilical panel and to the ground-based mass spectrometer for analysis at the launch pad.
Information content from the NSTS Shuttle Reference Manual (1988)
Last Hypertexed Wednesday October 11 17:42:26 EDT 1995
Jim Dumoulin (email@example.com)
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