The upper and lower forward fuselage, crew compartment, forward reaction control system and aft fuselage were manufactured at Rockwell's Space Transportation Systems Division facility in Downey and were transported overland from Downey to Rockwell's Palmdale, Calif., assembly facility.
The midfuselage was manufactured by General Dynamics, San Diego, Calif., and transported overland to Rockwell's Palmdale assembly facility.
The wings (including elevons) were manufactured by Grumman, Bethpage, Long Island, N.Y., and transported by ship from New York via the Panama Canal to Long Beach, Calif., and then transported overland to Rockwell's Palmdale assembly facility.
The vertical tail (including rudder/speed brake) were manufactured by Fairchild Republic, Farmingdale, Long Island, N.Y., and transported overland to Rockwell's Palmdale assembly facility.
The payload bay doors were manufactured at Rockwell International's Tulsa, Okla., facility and transported overland to Rockwell's Palmdale assembly facility.
The body flap was manufactured at Rockwell International's Columbus, Ohio, facility and transported overland to Rockwell's Palmdale assembly facility.
The aft orbital maneuvering system/reaction control system pods were manufactured by McDonnell Douglas, St. Louis, Mo., and transported by aircraft to Rockwell's Palmdale assembly facility. They were also transported by aircraft from Rockwell's Palmdale assembly facility to the Kennedy Space Center.
Approximately 250 major subcontractors supplied various systems and components to Rockwell's Palmdale assembly facility.
Rockwell's Palmdale assembly facility is where all the individual parts, pieces and systems came together and were assembled and tested. Upon completion, the spacecraft was turned over to NASA for transport overland from Palmdale to Edwards Air Force Base, California. NASA's Dryden Flight Research Facility at Edwards Air Force Base is the site of the mate-demate facility for mating or demating the spacecraft and the shuttle carrier aircraft.
For mating atop the shuttle carrier aircraft, the orbiter is raised horizontally in the mate facility until the shuttle carrier aircraft can be towed under the orbiter. The orbiter is then lowered and attached to two aft and one forward attach points on the shuttle carrier aircraft. These attach points on the orbiter are the same attach points where the external tank is attached to the orbiter.
For ferry flights of the orbiters after delivery from Palmdale, dummy orbital maneuvering system/reaction control system pods were used along with a tail cone installed over the aft section of the orbiter to streamline airflow.
The space shuttle main engines were manufactured by Rockwell International's Rocketdyne Division in Canoga Park, Calif. They are shipped separately from Rocketdyne to the National Space Technology Laboratories, then to the Kennedy Space Center.
The starting and sustaining heater system for each fuel cell power plant was modified to prevent overheating and loss of heater elements. A stack inlet temperature measurement was added to each fuel cell power plant for full visibility of thermal conditions.
The product water from all three fuel cell power plants flows to a single water relief control panel. The water can be directed from the single panel to the environmental control and life support system's potable water tank A or to the fuel cell power plant water relief nozzle. Normally, the water is directed to water tank A. In the event of a line rupture in the vicinity of the single water relief panel, water could spray on all three water relief panel lines, causing them to freeze and preventing water discharge.
The product water lines from all three fuel cell power plants were modified to incorporate a parallel (redundant) path of product water to ECLSS potable water tank B in the event of a freeze-up in the single water relief panel. If the single water relief panel freezes up, pressure would build up and discharge through the redundant paths to water tank B.
A water purity sensor (pH) was added at the common product water outlet of the water relief panel to provide a redundant measurement of water purity (a single measurement of water purity in each fuel cell power plant was provided previously). If the fuel cell power plant pH sensor failed in the past, the flight crew had to sample the potable water.
Improved auxiliary power units are scheduled for delivery in late 1988. A new turbine housing increases the life of the housing to 75 hours of operation (50 missions); a new gas generator increases its life to 75 hours; a new standoff design of the gas generator valve module and fuel pump deletes the requirement for a water spray system that was required previously for each APU upon shutdown after the first OMS thrusting period or orbital checkout; and the addition of a third seal in the middle of the two existing seals for the shaft of the fuel pump/lube oil system (previously only two seals were located on the shaft, one on the fuel pump side and one on the gearbox lube oil side) reduces the probability of hydrazine leaking into the lube oil system.
The deletion of the water spray system for the gas generator valve module and fuel pump for each APU results in a weight reduction of approximately 150 pounds for each orbiter. Upon the delivery of the improved units, the life-limited APUs will be refurbished to the upgraded design.
In the event that a fuel tank valve switch in an auxiliary power unit is inadvertently left on or an electrical short occurs within the valve electrical coil, additional protection is provided to prevent overheating of the fuel isolation valves.
1. The thickness of the main landing gear axle was increased to provide a stiffer configuration that reduces brake-to-axle deflections and precludes brake damage experienced in previous landings. The thicker axle should also minimize tire wear.
2. Orifices were added to hydraulic passages in the brake's piston housing to prevent pressure surges and brake damage caused by a wobble/pump effect.
3. The electronic brake control boxes were modified to balance hydraulic pressure between adjacent brakes and equalize energy applications. The anti-skid circuitry previously used to reduce brake pressure to the opposite wheel if a flat tire was detected has now been removed.
4. The carbon-lined beryllium stator discs in each main landing gear brake were replaced with thicker discs to increase braking energy significantly.
5. A long-term structural carbon brake program is in progress to replace the carbon-lined beryllium stator discs with a carbon configuration that provides higher braking capacity by increasing maximum energy absorption.
Other studies involve arresting barriers at the end of landing site runways (except lake bed runways), installing a skid on the landing gear that could preclude the potential for a second blown tire on the same gear after the first tire has blown, providing ''roll on rim'' for a predictable roll if both tires are lost on a single or multiple gear and adding a drag chute.
Studies of landing gear tire improvements are being conducted to determine how best to decrease tire wear observed after previous Kennedy Space Center.landings and how to improve crosswind landing capability.
Modifications were made to the Kennedy Space Center.shuttle landing facility runway. The full 300-foot width of 3,500-foot sections at both ends of the runway were ground to smooth the runway surface texture and remove cross grooves. The modified corduroy ridges are smaller than those they replaced and run the length of the runway rather than across its width. The existing landing zone light fixtures were also modified, and the markings of the entire runway and overruns were repainted. The primary purpose of the modifications is to enhance safety by reducing tire wear during landing.
The low-temperature thermal protection system tiles on Columbia's midbody, payload bay doors and vertical tail were replaced with advanced flexible reusable surface insulation blankets.
Because of evidence of plasma flow on the lower wing trailing edge and elevon landing edge tiles (wing/elevon cove) at the outboard elevon tip and inboard elevon, the low-temperature tiles are being replaced with fibrous refractory composite insulation (FRCI-12) and high-temperature (HRSI-22) tiles along with gap fillers on Discovery and Atlantis. On Columbia only gap fillers are installed in this area.
Also, because of the detailed analysis of actual descent flight data, room-temperature vulcanizing silicone rubber material was bonded to the lower midfuselage from bays 4 through 11 to act as a heat sink, distributing temperatures evenly across the bottom of the midfuselage, reducing thermal gradients and ensuring positive margins of safety.
The hardware changes required to the orbiters would enable the flight crew to equalize the pressurized crew compartment with the outside pressure via a depressurization valve opened by pyrotechnics in the crew compartment aft bulkhead that would be manually activated by a flight crew member in the middeck of the crew compartment; pyrotechnically jettison the crew ingress/egress side hatch in the middeck of the crew compartment; and bail out from the middeck of the orbiter through the ingress/egress side hatch opening after manually deploying the escape pole through, outside and down from the side hatch opening. One by one, each crew member attaches a lanyard hook assembly, which surrounds the deployed escape pole, to his parachute harness and egresses through the side hatch opening. Attached to the escape pole, the crew member slides down the pole and off the end. The escape pole provides a trajectory that takes the crew members below the orbiter's left wing.
Changes were also made in the software of the orbiter's general-purpose computers. The software changes were required for the primary avionics software system and the backup flight system for transatlantic-landing and glide-return-to-launch-site aborts. The changes provide the orbiter with an automatic-mode input by the flight crew through keyboards on the commander's and/or pilot's panel C3, which provides the orbiter with an automatic stable flight for crew bailout.
Note that the side hatch jettison feature could be used in a landing emergency.
The emergency egress slide replaces the emergency egress side hatch bar, which required the flight crew members to drop approximately 10.5 feet to the ground. The previous arrangement could have injured crew members or prevented an already-injured crew member from evacuating and moving a safe distance from the orbiter.
Inadvertent closure of either valve in a 17-inch disconnect during main engine thrusting would stop propellant flow from the external tank to all three main engines. Catastrophic failure of the main engines and external tank feed lines would result.
To prevent inadvertent closure of the 17-inch disconnect valves during the space shuttle main engine thrusting period, a latch mechanism was added in each orbiter half of the disconnect. The latch mechanism provides a mechanical backup to the normal fluid-induced-open forces. The latch is mounted on a shaft in the flowstream so that it overlaps both flappers and obstructs closure for any reason.
In preparation for external tank separation, both valves in each 17-inch disconnect are commanded closed. Pneumatic pressure from the main propulsion system causes the latch actuator to rotate the shaft in each orbiter 17-inch disconnect 90 degrees, thus freeing the flapper valves to close as required for external tank separation.
A backup mechanical separation capability is provided in case a latch pneumatic actuator malfunctions. When the orbiter umbilical initially moves away from the ET umbilical, the mechanical latch disengages from the ET flapper valve and permits the orbiter disconnect flapper to toggle the latch. This action permits both flappers to close.
In addition to the Phase II improvements, additional changes in the SSME have been incorporated to further extend the engines' margin and durability. The main changes were to the high-pressure turbomachinery, main combustion chamber, hydraulic actuators and high-pressure turbine discharge temperature sensors. Changes were also made in the controller software to improve engine control.
Minor high-pressure turbomachinery design changes resulted in margin improvements to the turbine blades, thereby extending the operating life of the turbopumps. These changes included applying surface texture to important parts of the fuel turbine blades to improve the material properties in the pressure of hydrogen and incorporating a damper into the high-pressure oxidizer turbine blades to reduce vibration.
Main combustion chamber life has been increased by plating a welded outlet manifold with nickel. Margin improvements have also been made to five hydraulic actuators to preclude a loss in redundancy on the launch pad. Improvements in quality have been incorporated into the servo-component coil design along with modifications to increase margin. To address a temperature sensor in-flight anomaly, the sensor has been redesigned and extensively tested without problems.
To certify the improvements to the SSMEs and demonstrate their reliability through margin (or limit testing), an aggressive ground test program was initiated in December 1986. From December 1986 to December 1987, 151 tests and 52,363 seconds of operation (equivalent to 100 shuttle missions) were performed. The SSMEs have exceeded 300,000 seconds total test time, the equivalent of 615 space shuttle missions. These hot-fire ground tests are performed at the single-engine test stands at the NASA National Space Technology Laboratories in Mississippi and at Rockwell International's Rocketdyne Division's Santa Susana Field Laboratory in California.
An "SRM Redesign Project Plan" was developed to formalize the methodology for SRM redesign and requalification. The plan provided an overview of the organizational responsibilities and relationships, the design objectives, criteria and process; the verification approach and process; and a master schedule. The companion "Development and Verification Plan" defined the test program and analyses required to verify the redesign and the unchanged components of the SRM.
All aspects of the existing SRM were assessed, and design changes were required in the field joint, case-to-nozzle joint, nozzle, factory joint, propellant grain shape, ignition system and ground support equipment. No changes were made in the propellant, liner or castable inhibitor formulations. Design criteria were established for each component to ensure a safe design with an adequate margin of safety. These criteria focused on loads, environments, performance, redundancy, margins of safety and verification philosophy.
The criteria were converted into specific design requirements during the Preliminary Requirements Reviews held in July and August 1986. The design developed from these requirements was assessed at the Preliminary Design Review held in September 1986 and baselined in October 1986. The final design was approved at the Critical Design Review held in October 1987. Manufacture of the RSRM test hardware and the first flight hardware began prior to the Preliminary Design Review (PDR) and continued in parallel with the hardware certification program. The Design Certification Review will review the analyses and test results versus the program and design requirements to certify the redesigned SRM is ready to fly.
The SRM field-joint metal parts, internal case insulation and seals were redesigned and a weather protection system was added.
In the STS 51-L design, the application of actuating pressure to the upstream face of the O-ring was essential for proper joint sealing performance because large sealing gaps were created by pressure-induced deflections, compounded by significantly reduced O-ring sealing performance at low temperature. The major change in the motor case is the new tang capture feature to provide a positive metal-to-metal interference fit around the circumference of the tang and clevis ends of the mating segments. The interference fit limits the deflection between the tang and clevis O-ring sealing surfaces caused by motor pressure and structural loads. The joints are designed so that the seals will not leak under twice the expected structural deflection and rate.
The new design, with the tang capture feature, the interference fit and the use of custom shims between the outer surface of the tang and inner surface of the outer clevis leg, controls the O-ring sealing gap dimension. The sealing gap and the O-ring seals are designed so that a positive compression (squeeze) is always on the O-rings. The minimum and maximum squeeze requirements include the effects of temperature, O-ring resiliency and compression set, and pressure. The clevis O-ring groove dimension has been increased so that the O-ring never fills more than 90 percent of the O-ring groove and pressure actuation is enhanced.
The new field joint design also includes a new O-ring in the capture feature and an additional leak check port to ensure that the primary O-ring is positioned in the proper sealing direction at ignition. This new or third O-ring also serves as a thermal barrier in case the sealed insulation is breached.
The field joint internal case insulation was modified to be sealed with a pressure-actuated flap called a J-seal, rather than with putty as in the STS 51-L configuration.
Longer field-joint-case mating pins, with a reconfigured retainer band, were added to improve the shear strength of the pins and increase the metal parts' joint margin of safety. The joint safety margins, both thermal and structural, are being demonstrated over the full ranges of ambient temperature, storage compression, grease effect, assembly stresses and other environments. External heaters with integral weather seals were incorporated to maintain the joint and O-ring temperature at a minimum of 75 F. The weather seal also prevents water intrusion into the joint.
Redesigned SRM certification is based on formally documented results of development motor tests; qualification motor tests and other tests and analyses. The certification tests are conducted under strict control of environments, including thermal and structural loads; assembly, inspection and test procedures; and safety, reliability, maintainability and quality assurance surveillance to verify that flight hardware meets the specified performance and design requirements. The "Development and Verification Plan" stipulates the test program, which follows a rigorous sequence wherein successive tests build on the results of previous tests leading to formal certification.
The test activities include laboratory and component tests, subscale tests, full-scale simulation and full-scale motor static test firings. Laboratory and component tests are used to determine component properties and characteristics. Subscale motor firings are used to simulate gas dynamics and thermal conditions for components and subsystem design. Full-scale hardware simulators are used to verify analytical models; determine hardware assembly characteristics; determine joint deflection characteristics; determine joint performance under short-duration hot-gas tests, including joint flaws and flight loads; and determine redesigned hardware structural characteristics.
Fourteen full-scale joint assembly demonstration vertical mate/demate tests, with eight interspersed hydro tests to simulate flight hardware refurbishment procedures, were completed early for the redesigned capture-feature hardware. Assembly loads were as expected, and the case growth was as predicted with no measurable increase after three hydro-proof tests.
Flight-configuration aft and center segments were fabricated, loaded with live propellant, and used for assembly test article stacking demonstration tests at Kennedy Space Center. These tests were pathfinder demonstrations for the assembly of flight hardware using newly developed ground support equipment.
In a long-term stack test, a full-scale casting segment, with live propellant, has been mated vertically with a J-seal insulation segment and is undergoing temperature cycling. This will determine the compression set of the J-seal, aging effects and long-term propellant slumping effects.
The Structural Test Article (STA-3), consisting of flight-type forward and aft motor segments and forward and aft skirts, was subjected to extensive static and dynamic structural testing, including maximum prelaunch, liftoff and flight (maximum dynamic pressure) structural loads.
Redesigned SRM certification includes testing the actual flight configuration over the full range of operating environments and conditions. The joint environment simulator, transient pressure test article, and the nozzle joint environment simulator test programs all utilize full-scale flight design hardware and subject the RSRM design features to the maximum expected operating pressure, maximum pressure rise rate and temperature extremes during ignition tests. Additionally, the Transient Pressure Test Article (TPTA) is subjected to ignition and liftoff loads as well as maximum dynamic pressure structural loads.
Four TPTA tests have been completed to subject the redesigned case field and case-to-nozzle joints to the above-described conditions. The field and case-to-nozzle joints were temperature-conditioned to 75 F. and contained various types of flaws in the joints so that the primary and secondary O-rings could be pressure-actuated, joint rotation and O-ring performance could be evaluated and the redesigned joints could be demonstrated as fail safe.
Six of the seven Joint Environment Simulators (JES) tests have been completed. The JES test program initially used the STS 51-L configuration hardware to evaluate the joint performance with prefabricated blowholes through the putty. The JES-1 test series, which consisted of two tests, established a structural and performance data base for the STS 51-L configuration with and without a replicated joint failure. The JES-2 series, two tests, also used the STS 51-L case metal-part joint but with a bonded labyrinth and U-seal insulation that was an early design variation of the J-seal. Tests were conducted with and without flaws built into the U-seal joint insulation; neither joint showed O-ring erosion or blow-by. The JES-3 series, three tests, uses almost exact flight configuration hardware, case field-joint capture feature with interference fit and J-seal insulation.
Four of five nozzle JES tests have been successfully conducted. The STS 51-L hardware configuration hydro test confirmed predicted case-to-nozzle-joint deflection. The other three tests used the radially bolted RSRM configuration.
Seven full-scale, full-duration motor static tests are being conducted to verify the integrated RSRM performance. These include one engineering test motor used to (1) provide a data base for STS 51-L-type field joints; (2) evaluate new seal material; (3) evaluate the ply-angle change in the nozzle parts,; (4) evaluate the effectiveness of graphite composite stiffener rings to reduce joint rotation; and (5) evaluate field-joint heaters. There were two development motor tests and three qualification motor tests for final flight configuration and performance certification. There will be one flight Production Verification Motor that contains intentionally induced defects in the joints to demonstrate joint performance under extreme worse case conditions. The QM-7 and QM-8 motors were subjected to liftoff and maximum dynamic pressure structural loads, QM-7 was temperature-conditioned to 90 F., and QM-8 was temperature-conditioned to 40 F.
An assessment was conducted to determine the full-duration static firing test attitude necessary to certify the design changes completely. The assessment included establishing test objectives, defining and quantifying attitude-sensitive parameters, and evaluating attitude options. Both horizontal and vertical (nozzle up and down) test attitudes were assessed. In all three options, consideration was given to testing with and without externally applied loads. This assessment determined that the conditions influencing the joint and insulation behavior could best be tested to design extremes in the horizontal attitude. In conjunction with the horizontal attitude for the RSRM full-scale testing, it was decided to incorporate externally applied loads. A second horizontal test stand for certification of the RSRM was constructed at Morton Thiokol. This new stand, designated as the T-97 Large Motor Static Test Facility, is being used to simulate environmental stresses, loads and temperatures experienced during an actual Shuttle launch and ascent. The new test stand also provides redundancy for the existing stand.
Alternative designs with long-lead-time implications were also developed. These designs focus on the field joint and cast-to-nozzle joint. Since fabrication of the large steel components dictates the schedule, long-lead procurement of maximum-size steel ingots was initiated. This allowed machining of case joints to either the new baseline or to an alternative design configuration. Ingot processing continued through forging and heat treating. At that time, the final design was selected. A principal consideration in this configuration decision was the result of verification testing on the baseline configuration.
In addition to the NRC, the redesign team has a design review group of 12 expert senior engineers from NASA and the aerospace industry. They have advised on major program decisions and serve as a "sounding board" for the program.
Additionally, NASA requested the four other major SRM companies -- Aerojet Strategic Propulsion Co., Atlantic Research Corp., Hercules Inc. and United Technologies Corp.'s Chemical Systems Division -- to participate in the redesign efforts by critiquing the design approach and providing experience on alternative design approaches.