Index

    Bruce T. Lundin Director, Lewis Research Center

    Vincent L. Johnson, Vice-Chairman Deputy Associate Administrator for Space Science, NASA Headquarters

    Thomas N. Canning Systems Development Branch, NASA Ames Research Center

    William R. Dunbar Deputy Director of Launch Vehicles, NASA Lewis Research Center

    E. Barton Geer Director for Systems Engineering and Operations, NASA Langley Research Center

    Lt. Col. Perry W. Harker Manager, Titan III Launch Services, Space -and Missile Test Center, Vandenberg Air Force Base

    Capt. Robert E. McKean Chief, Titan Ill Launch Controller, Space and Missile Test Center, VandenbergAir Force Base

    Merland L. Moseson Deputy Director of Systems Reliability, NASA Goddard Space Flight Center

    S. Neil Hosenball Deputy General Counsel, NASA Headquarters

    Edward A. Richley Chief, Operations Analysis and Planning, NASA Lewis Research Center

    Haggai Cohen Director, Reliability, Quality and Safety, Office of Manned Space Flight, NASA Headquarters

    Gilbert L. Roth Special Assistant, Aerospace Safety Panel, NASA Headquarters

    Vernon L. Alley, Jr Chief, Engineering and Analysis for Systems Engineering and Operations, NASA Langley Research Center

    Robert T. Wingate Head, Engineering Analysis Branch, Systems Engineering Division, NASA Langley Research Center

Figures appear at the end of each chapter.

Index

Tables appear following the figures at the end of each Chapter

Index

At approximately 63 seconds into the flight of Skylab 1 on May 14, 1973, an anomaly occurred which resulted in the complete loss of the meteoroid shield around the orbital workshop. This was followed by the loss of one of the two solar array systems on the workshop and a failure of the interstage adapter to separate from the S-II stage of the Saturn V launch vehicle. The investigation reported herein identified the most probable cause of this flight anomaly to be the breakup and loss of the meteoroid shield due to aerodynamic loads that were not accounted for in its design. The breakup of the meteoroid shield, in turn, broke the tie downs that secured one of the solar array systems to the workshop. Complete loss of this solar array system occurred at 593 seconds when the exhaust plume of the S-II stage retro-rockets impacted the partially deployed solar array system. Falling debris from the meteoroid shield also damaged the S-II interstage adapter ordnance system in such a manner as to preclude separation.

Of several possible failure modes of the meteoroid shield that were identified, the most probable in this particular flight was internal pressurization of its auxiliary tunnel which acted to force the forward end of the meteoroid shield away from the shell of the workshop and into the supersonic air stream. The pressurization of the auxiliary tunnel was due to the existence of several openings in the aft region of the tunnel. Another possible failure mode was the separation of the leading edge of the meteoroid shield from the shell of the workshop (particularly in the region of the folded ordnance panel) of sufficient extent to admit ram air pressures under the shield.

The venting analysis for the auxiliary tunnel was predicated on a completely sealed aft end; the openings in the tunnel thus resulted from a failure of communications among aerodynamics, structural design, and manufacturing personnel. The failure to recognize the design deficiencies of the meteoroid shield through six years of analysis, design and test was due, in part, to a presumption that the shield would be "tight to the tank" and "structurally integral with the S-IVB tank" as set forth in the design criteria. In practice, the meteoroid shield was a, large, flexible, limp system that proved difficult to rig to the tank and to obtain the close fit that was presumed by the design. These design deficiencies of the meteoroid shield, as well as the failure to communicate within the project the critical nature of its proper venting, must therefore be attributed to an absence of sound engineering judgment and alert engineering leadership concerning this particular system over a considerable period of time.

The overall management system used for Skylab was essentially the same as that developed in the Apollo program. This system was fully operational for Skylab; no conflicts or inconsistencies were found in the records of the management reviews. Nonetheless, the significance of the aerodynamic loads on the meteoroid shield during launch were not revealed by the extensive review process. Possibly contributing to this oversight was the basic view of the meteoroid shield as a piece of structure, rather than as a complex system involving several different technical disciplines. Complex, multidisciplinary systems such as the meteoroid shield should have a designated project engineer who is responsible for all aspects of analysis, design, fabrication, test and assembly.

The Board found no evidence that the design deficiencies of the meteoroid shield were the result of, or were masked by, the content and processes of the management system that were used for Skylab. On the contrary. the rigor, detail, and thoroughness of the system are doubtless necessary for a program of this magnitude. At the same time, as a cautionary note for the future, it is emphasized that management must always be alert to the potential hazards of its systems and take care that an attention to rigor, detail and thoroughness does not inject an undue emphasis on formalism, documentation, and visibility in detail. Such an emphasis can submerge the concerned individual and depress the role of the intuitive engineer or analyst. It will always be of importance to achieve a cross-fertilization and broadened experience of engineers in analysis, design, test or operations. Positive steps must always be taken to assure that engineers become familiar with actual hardware, develop an intuitive understanding of computer-developed results, and make productive use of flight data in this learning process. The experienced "chief engineer," who can spend most of his time in the subtle integration of all elements of the system under his purview, free of administrative and managerial duties, can also be a major asset to an engineering organization.

Index

Program Objectives

Skylab missions have several distinct goals: conduct of earth resources observations. advance scientific knowledge of the sun and stars; study the effects of weightlessness on living organisms, particularly man; study and understand methods for the processing of materials in the absence of gravity. The Skylab mission utilizes man as an engineer and as a research scientist, and provides an opportunity for assessing his potential capabilities for future space missions.

Skylab Hardware

Skylab utilizes the knowledge, experience and technical systems developed during, the Apollo program along with specialized equipment necessary to meet the program objectives.

Figure 1-1 shows the Skylab in orbit. Its largest element is the Orbital Workshop (OWS), a cylindrical container 48 feet long and 22 feet in diameter weighing some 78. 000 pounds. The basic structure of the OWS is the upper stage, or S-IVB stage, of the S-IB and S-V rockets which served as the Apollo program launch vehicle. The OWS has no engines, except attitude control-thrusters, and has been modified internally to provide a large orbiting space laboratory and living quarters for the crew. The Skylab 1 (SL-1) space vehicle included a payload consisting of four major units (OWS, Airlock Module (AM)., Multiple Docking Adapter (MDA), Apollo Telescope Mount (ATM)) and a two-stage Saturn-V (S-IC and S-II) launch vehicle as depicted in figure 1-2. To provide meteoroid protection and thermal control, an external meteoroid shield (MS) was added to cover the OWS habitable volume. A solar array system (SAS) was attached to the OWS to provide electrical power.

The original concept called for a "Wet Workshop". In this concept, a specially constructed S-IVB stage was to be launched "Wet" as a propulsive stage on the S-IB Launch System filled with propellants., The empty hydrogen tank would then be purged and filled with a life-supporting atmosphere. A major redirection of Skylab was made m July 22, 1969, six days after the Apollo 11 lunar landing. As a result of the successful lunar landing, S-V launch vehicles became available to the Skylab program. As a result, it became feasible to completely equip the S-IVB on the ground for immediate occupancy and use by a crew after it was in orbit. Thus it would not carry fuel and earned the name of "Dry Workshop".

Skylab Mission Plan

The nominal Skylab-mission (fig. 1-3) called for the launch of the unmanned S-V vehicle and workshop payload SL-1 into a near circular (235 nautical miles) orbit inclined 50 degrees to the equator. Then about 24 hours after the first launch., the manned Skylab 2 (SL-2) launch would take place using a Command Service Module (CSM) payload atop the S-IB vehicle. After CSM rendezvous and docking with the orbiting cluster, the crew enters and activates the workshop; Skylab is then ready for its first operational period of 28 days. At the end of this period, the crew returns to earth with the CSM, and the Skylab continues in an unmanned quiescent mode for some 60 days. The second three man crew is launched with a second S-IB, this time for a 56-day period of manned operation. After return of the second crew to earth, the Skylab again operates in an unmanned mode for approximately one month. The third three-man crew is then launched with the third S-IB for a second 56-day period In orbit after which they will return to earth. The total Skylab mission activities cover a period of roughly eight months, with 140 days of manned operation.

Skylab Program Environment

The Skylab Program Office in the Office of Manned Space Flight in NASA Headquarters is responsible for overall management of the program. The NASA Center responsibilities are as follows:

The major Skylab prime and first tier subcontractors and their. responsibilities are shown in table I-1.

    Figure 1-1   

    Figure 1-2   

    Figure 1-3 - Skylab mission profile   

Index

Index

Launch and Environment

Skylab 1 was launched at 1730:00 (Range time, R=0) on May 14, 1973, from Complex 39 A, Kennedy Space Center. At this time. the Cape Kennedy launch area was experiencing cloudy conditions .with warm temperatures and gentle surface winds. Total sky cover consisted of scattered cumulus at 2,400 feet. scattered stratocumulus at 5,000 feet, broken altocumulus at 12,000 feet, and cirrus at 23,000 feet. During ascent, the vehicle passed through the cloud layers but no lightning was observed in the area. As shown in tables II-1 and II-2, upper area wind conditions were benign compared to most other Saturn-V flights. Figure 2-1 shows a comparison between wind speed. altitude, and time during the launch. Figure 2-2 shows altitude vs. range time. Figure 2-3 is a plot showing SL-1 history in the region of maximum bending moment. As can be seen. the flight environment was quite favorable.

Major Events

The automatic countdown proceeded normally with Guidance Reference Release occurring at R-17.0 seconds and orbit insertion, occurring at R+599.0 seconds. Table II-3 lists the pertinent orbit parameters and table II-4 is a summary of the normal major events through orbit insertion. All times are referenced from Range time, R-0, which is defined as the last integral second prior to liftoff. As can be seen from table II-4, the OWS solar array deployment was commanded on time; however, real time data indicated that the system did not deploy fully.

Description of Solar Array System and Meteoroid Shield

The Solar Array System (SAS) on the OWS consists of two large beams enclosing three major sections of solar cell assemblies within each. During ascent, the sections are folded like an accordion inside the beams which in turn are stowed against the workshop as shown in figure 2-4. The MS is a lightweight structure wrapped around the converted S-IVB stage orbital workshop and is exposed to the flight environment. The MS, and its attachment to the OWS, is described in detail in Chapter IV of this report. The two hinged SAS wings are secured to the OWS by tie downs above and below the MS. Seals attached to the SAS perimeter actually press against the shield to form an airtight cavity prior to launch. Once in orbit, the SAS beams are first deployed out 90 degrees. The MS is deployed later to a distance of about five inches from the OWS wall (see fig. 2-4). After the ordnance release is fired, MS deployment is effected by torsion rods and swing links spaced around the structure fore and aft. The rods are torqued prior to launch and simply "unwind" in orbit to move the MS away from the tank. Detection of pertinent conditions associated with the MS and SAS is afforded by measuring various parameters by telemetered instrumentation. Figure 2-5 shows a plan view of the MS and SAS configuration and identifies the location of instrumentation sensors.

Early Indication of Anomalies

When the OWS Solar Array System was commanded to deploy, telemetered data indicated that events did not occur as planned. The flight data was analyzed by flight operations personnel to reveal the possible source of the problem. At about R+60 seconds, the S-II telemetry reflected power increased slightly. At about 63 seconds, numerous measurements indicated the apparent early deployment and loss of the MS. At this time, the vehicle was at about 28,600 feet altitude and at a velocity of about Mach 1.

At this time, vehicle dynamic measurements such as vibration, acceleration, attitude error and acoustics indicated strong disturbances. Measurements which are normally relatively static at this time, such as torsion rod strain gages, tension strap breakwires, temperatures, and SAS position indicators, indicated a loss of the MS and unlatch of the SAS-2 wing. Further preliminary evaluation revealed abnormal vehicle accelerations, vibrations, and SAS temperature and voltage anomalies at about R+593 seconds. Temperature data loss and sudden voltage drops indicated that the SAS-2 wing was separated from the OWS at this time. Other data later in the flight indicated the SAS-1 wing did not fully deploy when commanded to do so. Although not apparently associated with the 63-second and 593-second anomalies, the S-II stage Range Safety Receiver signal strengths showed several drops throughout the flight beginning at about R+260 seconds.

    Figure 2-1. - Scalar wind speed at launch time of SL-1.   

    Figure 2-2. - Altitude comparisons for early portion of S-V launch vehicle boost trajectory.   

    Figure 2-3. - SL-1 alpha-beta limit at maximum bending moment.   

   Figure 2-4. - Meteoroid shield and solar array system.  (a) STOWED.   

   Figure 2-4. - Meteoroid shield and solar array system.  (b) DEPLOYED.   

    Figure 2-5. - Meteoroid shield and instrumentation layout.   

Index

TABLE II-1. - MAXIMUM WIND SPEED IN HIGH DYNAMIC PRESSURE REGION FOR APOLLO / SATURN 501 THROUGH SATURN 513 VEHICLES

Index

TABLE II-2. - EXTREME WIND SHEAR VALUES IN THE HIGH DYNAMIC PRESSURE REGION FOR APOLLO / SATURN 501 THROUGH SATURN 513 VEHICLES

Index

Table II-3 -- Orbit Parameters *

* Data source is from radar data processed by the Mission Operations Computer at JSC.

Index

Table II-4 -- Normal Major Events

Index

63 Second Anomaly - Loss of MS

The Investigation Board, evaluated the telemetry data in order to explain the various anomalies that occurred on Skylab 1. The first anomalous indication was an increase in S-II telemetry reflected power from a steady 1.5w beginning at R+ 59. 80 seconds. At this time the telemetry forward power remained steady at 58.13w. By 61.04 seconds, the reflected power had reached 1.75w, and by 80.38 seconds, the reflected power had stabilized at about 2.0w. This abnormal increase in power might be indicative of a vehicle physical configuration change which altered the antenna ground plane characteristic.

Shortly after the telemetry reflected power increase, the MS torsion rod 7 forward (measurement G7036) indicated a slight change toward the deployed condition (see fig. 2-5 for instrumentation layout). This occurred at R+60.12 seconds, and at 61.78 seconds the vehicle roll rate decreased slightly from a normal value of 1.1 degrees per second clockwise (CW) looking forward. Figure 3-1 is a graph of the roll rate versus range time during the time of interest. The next torsion rod 7 forward sample at about 62.52 seconds revealed a further relaxation. The increase in telemetry reflected power and the movement of torsion rod 7 forward tend to indicate meteoroid shield lifting between positions I and II (see fig. 2-5).

Between R+62.75 and 63.31 seconds, several vehicle dynamic measurements indicated a significant disturbance. A sensor on the OWS film vault showed an abnormal vibration at 62.75 seconds followed by disturbances sensed by X and Y accelerometer pickups in the Instrument Unit (IU), the pitch, yaw, and longitudinal accelerometers, and the pitch, yaw, and roll rate gyros. At 62.78 seconds, the roll rate gyro sensed a sudden CW roll rate resulting in a peak amplitude of 3.0 degrees per second CW at 62.94 seconds. A sensor at the X upper mounting showed a maximum peak-to-peak shock of 17.2 g's at 63.17 seconds. In addition, the S-II engine actuators experienced pressure fluctuations caused by vehicle movement against the inertia of the non-thrusting engine nozzles.

During the time the vehicle was sensing a disturbance, several slower-rate MS and SAS measurements experienced drastic changes. Because these measurements are sampled only once every 0.1 to 2.4 seconds, there is that period of uncertainty as to when the measurement has actually changed. Figure 3.2 is a graphic representation of the applicable measurements associated with the 63-second anomaly. Where only a single point is shown, the sampling is continuous or has no significant bearing on the hypotesization of the MS failure mode. For the MS and SAS data sampled at 0.1, 0.8, and 2.4 seconds per sample, the last normal and first abnormal times are shown. Figures 3-3 , 3-4 , 3-5 , 3-6 , 3-7 , 3-8 are pictorial representations of the status of the MS and SAS measurements at the indicated period of time. Figure 3-3 is a time slice at R+60.90 seconds where all measurements are known to be normal for the last time (except for the slight movement of torsion rod 7 forward beginning at 60.12 seconds).

Figure 3-4 is a time slice at the first indication of a measurement failure (R+62.78 seconds). The measurements K7211, C70132, K7010, K7011, and K7012 can be considered normal here because they were normal during the previous sample and were sampled later than 62.78 seconds and found to still be normal. At this time period, C7011 (a temperature measurement) was lost. The cause of this measurement failure could have been due to the sensor or its cabling (shown in fig. 3-4 by dashed lines) being damaged. This was most likely a result of the NS failure in the area between the SAS-2 wing and the main tunnel (between positions I and H). Furthermore, both SAS wing secure indications and the ordnance tension strap indications are known to be good. This evidence leads to two conclusions at this point: the meteoroid shield failure began prior to the SAS-2 wing becoming unlatched, and the ordnance did not fire prematurely.

Figures 3-5 and 3-6 are time slices at R+62.89 and R+62.90, respectively, that show the failure of measurements C7012, K7010, K7011, and K7211, while K7212 (SAS-1 secured) and C7013 (MS temperature) were known to be normal by a later sample, The abnormal telemetry indications C7012, K7010 and K7011, like C7011 at R+62.78 seconds, could have been due to sensor or wiring damage. Measurements K7010, K7011, and K7012 are, in fact, only breakwires placed across the ordnance tension strap. Measurement K7211, however, reveals that the SAS-2 wing was no longer secure to the OWS. This is an indication that the SAS-2 wing had moved out at least between 0.651 and 2.821 degrees, or between 4.66 and 20.2 inches as measured at the aft end of the wing perpendicular to the OWS.

Figure 3-7 represents a later time, 62.97 seconds. At this time, K7012 (tension strap) was detected as failed. Slightly later, at R+63.04 seconds, the first indication of increased SAS voltage appeared. Measurement M0103 showed a slight increase in voltage which is attributed to sunlight illuminating exposed sections of the partially deployed (unlatched) SAS-2 wing. Other SAS voltages fluctuated throughout the remainder of the launch phase for the same reason. Between 62.97 and 64.92 seconds, an of the MS failure-related measurements became abnormal. Figure 3-8 shows that the SAS-1 wing secure measurement (K7212) was still normal.

The data indicate that the most probable sequence of Meteoroid Shield failure was initial structural failure of the MS between the SAS-2 wing and the main tunnel (between positions I and II). The initial failure propagation from this area appears likely since the wardroom window thermocouple indication (C7013) remained normal at 62. 94 seconds after SAS- 2 indicated unlatched at 62.90 seconds and after the K7010 and K7011 tension strap measurements failed.

593 Second Anomaly

As a consequence of the MS failure at approximately 63 seconds, the SAS-2 wing was unlatched and partially deployed as evidenced by minor variations in the main SAS electrical voltages and SAS-2 temperatures. Full deployment was prevented due to the aerodynamic forces and accelerations during the remainder of powered flight.

At the completion of the S-II phase of flight the four 35, 000 pound thrust retro-rockets fired for approximately two seconds commencing at R+591.10 seconds followed by spacecraft separation at 591.2 seconds. The effect of retro-rocket plume impingement (refer to fig. 3-9 for location and orientation of the retro-rockets relative to the SAS-2 wing) was observed almost immediately on the SAS-2 temperature and on vehicle body rates.

The time sequence of observed changes in the affected measurements is demonstrated in figure 3-10. The response of the vehicle and the corrective action of the attitude control system may be seen in figures 3-11 and 3-12.

An analysis of the impingement forces on the wing was made and compared to the force required to produce the observed vehicle motion. This comparison provides a reasonable fit for the first 50 to 60 degrees of wing rotation as shown in figure 3-13.

At 593.4 seconds the wing imparted momentum to the vehicle, probably by hitting and breaking the 90 degree fully deployed stops and at 593.9 imparted a final kick as it tore completely free at the hinge link. In-orbit photographs show clearly the hinge separation plane and the various wires which were torn loose at the interface (see fig. 3-14).

Interstage Second Plane Separation Anomaly

Post-flight analysis revealed unexpectedly high temperatures and pressures in the S-II engine compartment following ignition and continued high after interstage, separation command as shown in figures 3-15 and 3-16. The unusually high temperatures from S-II ignition and until the S-II interstage separation signal are considered by MSFC to be caused by a change in the engine heat shield skirts introduced on this flight, and therefore do not indicate a problem. However, the increasing temperatures after the time of normal S-II interstage separation are indicative of an abnormal condition. More detailed investigation based on performance evaluation and axial acceleration time history revealed that the interstage had not been jettisoned; however, due to the vehicle performance characteristics and performance margin, the desired orbit was achieved.

Data analysis confirms that the primary ordnance command was properly issued at R+189.9 seconds. The back-up command was issued 100 milliseconds (ms) later but the exploding bridge wire circuit discharge was characteristic of an open circuit consistent with separation of the interstage disconnect by a minimum of 0.2 5 inch as shown in figure 3 -17.

The linear shaped charge (LSC) is mounted circumferentially around the S-II interstage as shown in figures 3-18 and 3-19. When fired by the primary command, the charge cuts the tension straps (in the direction of position II to position I) allowing the skirt to drop away. Normal propagation time of the LSC is approximately 4ms. Assuming a failure to propagate completely around the structure, analyses were made by appropriate contractor and the government personnel to determine what area must remain intact in order to retain the skirt and what area must have been cut to allow rotation of the skirt sufficient to disconnect the connector panel. An example of the results of one analysis is shown in figure 3-20. The various analyses isolate the region of failure to an are extending from approximately 0 = 100 degrees to as much as 0 = 200 degrees.

This ordnance installation was different from prior Saturn flights. Previously, a single fire command from the instrumentation unit was issued which simultaneously detonated the LSC from both ends allowing the charge to propagate from both directions. On this flight, in an attempt to provide redundant firing commands, the detonators at each end of the LSC were separately connected to two command channels spaced 100 milliseconds apart due to the characteristics of the airborne equipment. As a result of the partial cutting of the interstage, it rotated sufficiently to separate the electrical connector prior to issuing the back-up command.

A review of the history of manufacturing, acceptance, checkout, qualification and flight environment revealed no basic cause for failure. The most probable cause is secondary damage as a result of the MS failure, attributed to falling debris as evidenced by the various shock and acoustic disturbances occurring in the 63-second time period.

The redundant mode of ordnance operation of all prior Saturn flights in which both ends of the LSC are fired at once from a single command would probably have prevented the failure, depending on the extent of damage experienced by the LSC.

Forward Interstage Internal Pressure Anomaly

Flight data indicated a deviation of the S-II forward interstage pressure from analytical values commencing at approximately 63 seconds. Inasmuch as the deviation from the analytical curve of internal pressure versus time appeared to be coincident with the MS failure (see fig. 3-21) it was postulated that a portion of the shield had punctured the forward interstage. On this basis, it was possible to correlate the flight data with either an assumed 2.0 square foot hole in the conical section or an assumed 0.75 square foot hole in the cylindrical section.

Range Safety Receiver Anomaly

During the S-II portion of the flight, the signal strength indications from both range safety receivers showed drops in level. From liftoff through R+259 seconds, both receivers maintained relatively stable values above range requirements. At R+259.57 seconds, receiver 2 signal strength began to drop and between this time and 522.1 seconds, both receivers indicated various degrees of signal strength shift. These signal strength shifts dropped below the 12 db safety margins required by Air Force Eastern Test Range (AFETR) Manual 127-1. At R+327.81 seconds, the receiver 2 signal strength dropped briefly below its threshhold sensitivity. At this instant this receiver probably would not have responded to any range safety commands. Receiver 1 was, however, capable of receiving commands. At R+521.16, receiver 2 strength again dropped briefly to its threshhold sensitivity. None of these drops could be correlated to ground system performance.

Analysis indicates that the most probable cause of the S-II receiver signal strength dropout was a variable phase shift within the vehicle's hybrid coupler due to the changing aspect angle produced by the moving vehicle and the fixed transmitting site. Because the decrease in receiver signal strength occurred with only one receiver at a time, range safety commands could have been received continuously throughout power flight. During two of these drops, however, the planned redundancy of range safety receivers was not available.

During this investigation, it was revealed that the Wallops Island and Bermuda ground stations did not continuously record ground transmitter power levels. The Board considers that such continuous recordings would be of value.

    Figure 3-1 - Roll rate versus range time.   

    Figure 3-2.  Time Sequence of 63-sec Anomaly Instrumentation   

    Figure 3-3. - Condition of meteoroid shield instrumentation at R + 60.90 sec   

    Figure 3-4. - Condition of meteoroid shield instrumentation at R + 62.78 sec   

    Figure 3-5. - Condition of meteoroid shield instrumentation at R + 62.89 sec.   

    Figure 3-6. - Condition of meteoroid shield instrumentation at R + 62.90 sec.   

    Figure 3-7. - Condition of meteoroid shield instrumentation at R + 62.97 sec.   

    Figure 3-8. - Condition of meteoroid shield instrumentation at R + 64.88 sec.   

    Figure 3-9. - SL-1 retro-rocket impingement force schematic for S-III / SWS separation.   

    Figure 3-10. - 593 sec anomaly time sequence.   

    Figure 3-11. - Explanation of 593 second anomaly.   

    Figure 3-12. - Explanation of 593 second anomaly.   

    Figure 3-13. - Plume impingement force on SAS-2.   

    Figure 3-14. - SAS-2 wing hinge.   

    Figure 3-15. - Engine compartment gas temperature.   

    Figure 3-16 . - Base region pressures - assumed failure mode: interstage did not separate.   

    Figure 3-17. - Separation EBW firing unit monitor indications   

    Figure 3-18. - Second plane separation system, S-II (block diagram and location).   

    Figure 3-19. - EBW detonator and detonator blocks, second plane separation system, S-II (installation).   

    Figure 3-20 . - S-II-13 interstage station 196 tension strap analysis   

    Figure 3-21 . - Forward interstage internal pressure.   

Index

Overall Description

Although fairly simple in concept, the meteoroid shield had to provide such a variety of functions that it was, in fact, a quite complicated device. It was, foremost, a very lightly built cylindrical structure 270 inches in diameter (in the deployed condition) by 265 inches long.

The general layout of the MS is illustrated in figure 4-1. The OWS, which it surrounds, is deleted in this figure for clarity. In brief, the MS is formed of a set of sixteen curved sheets of 2014 T6 aluminum panels, 0.025 inches thick, assembled at flanges and other fittings to form the cylinder shown. The forward and aft ends were reinforced with curved 7075 T6 angles.

Various special details were included in the assembly in order to hold it in place, deploy it in orbit, and provide access to the OWS interior during prelaunch activities. The principal means of holding the shield in place in orbit (and to a lesser extent during powered flight) was a set of tension straps under the main tunnel illustrated at position II in figure 4-1. These straps were bonded to the OWS wall and fitted with a hinge on each end to take the butterfly hinge that attaches to the adjacent MS panel as indicated in figure 4-2. These butterfly hinges were designed to rotate so as to lie against the sides of the main tunnel which enclosed the tension straps and various cable runs on the OWS.

Proceeding clockwise from the tension straps and butterfly hinges in figure 4-1, the next special feature is the auxiliary tunnel. This tunnel extends in an arch between panels of the thin meteoroid shield. The 28 titanium frames of this tunnel (fig. 4-3) provide a very springy section in the relatively rigid hoop provided by the rest of the shield. The auxiliary tunnel also encloses a smaller tunnel covering the wiring for the thruster attitude control system. Farther around, in position I, there are two curved rectangular smaller panels, included to provide access to the OWS.

Between positions I and IV, the two halves of the MS overlap and are joined by a series of 14 trunnion bolts and straps (shown in detail in 4-4). These trunnion bolts were used to adjust the tension with which the MS was held against the OWS. Adjusting the bolts in the trunnion assemblies was a major aspect in positioning and tightening the MS against the OWS (rigging).

In order to provide the extra 30 inches of perimeter required when the MS was deployed, a foldout panel assembly (released by ordnance) is included in the panel adjacent to the trunnions. The mode of operation of this foldout panel is indicated in figure 4-5. Detailed descriptions of the ordnance and its function are given later. The only remaining distinctive features of the MS are the panels located over the scientific airlock and wardroom window at position III. The MS is completed at the butterfly hinges and tension straps at position I.

Deployment Provisions

The deployment of the 265 inch long MS was accomplished by providing two folding panel sections on each side of a contained explosive pyrotechnic chain which extended axially for the full length of the MS except for short end reinforcements. These folding panels and a schematic version of the redundant pyrotechnics were previously presented in figure 4-5. Illustrations of the shield in the stowed, partially deployed, and deployed configurations are shown in figures 4-6, 4-7, and 4-8, respectively. A cross section through the ordnance and its schematic are shown in figure 4-9. When the ordnance strip is fired and separates the "fold-over" panel, the segments are released and the shield is deployed. After release of this folded panel, a number of swing arms are used to displace the shield away from the OWS wall and hold it there. A rotational force is applied to these swing arms by a total of sixteen torsion rods suitably spaced around the ends of the MS as shown in figures 4-10 and 4-11. When the MS is stowed for launch, there is a larger twist in the torsion rods than after deployment. Both stowed and deployed torque settings are tabulated in table IV-1. It can be noted in figure 4-7 how the links on one side of the ordnance chain swing in a direction opposite to those on the other side. The butterfly hinges on each side of the main tunnel permit the radial displacement of the shield at the location of the tension straps.

The MS should therefore be regarded as a very limp system, which depends on being stretched tight around the OWS to withstand the aerodynamic, vibratory, flutter and thrust loads at launch. After deployment, it needs very little strength to serve its primary objective as a meteoroid shield.

The Auxiliary Tunnel

The auxiliary tunnel, an assembly of which is shown in figures 4-12 and 4-13, extends from the forward skirt, down the full length of the MS shield, and below the MS by about 57 inches. Venting of this tunnel was provided through an outlet of 10 square inches under the corrugations of the tunnel cover at the aft end of the forward fairing as detailed in figure 4-14. The tunnel was intended to be sealed at the aft end by a rubber boot assembly shown in the photographs of figure 4-15 in both the stowed (A) and deployed (B) position, Note that the tunnel is displaced some 5 or 6 inches circumferentially upon deployment of the shield.

The main structural members of the auxiliary tunnel are titanium, arch shaped, frame springs. These frames provide the structural tie between two MS panels and provide both regulation of the preloading of the MS to the OWS and act as a flexible relief for diametrical changes resulting from thermal and pressure changes of the OWS.

The tunnel also serves to protect the thrust attitude control system cables located in a small channel shaped cover permanently attached to the OWS and shown in figure 4-13. A segmented and corrugated outer skin form an aerodynamic fairing for the complete system and seals between forward and aft fairings.

Thermal Control

Although the primary purpose of the meteoroid shield is that of providing protection of the OWS from meteoroids, it also plays a significant role in the thermal control system. Much of the overall thermal design was accomplished passively by. painting the outer surfaces of the MS black except for a large white cross-shaped pattern on the earth side during flight. The entire surface of the OWS wall was covered with gold foil. The overall choice of finishes biased the thermal design toward the cold side, it being easier to vernier control by heating rather than cooling.

Friction Between MS and OWS Wall

To provide a uniform tension throughout the MS upon assembly and rigging for flight, and to permit transfer of the trunnion bolt tension into the frames of the auxiliary tunnel, it was necessary to minimize friction between the MS and the extemal surface of the OWS. This was accomplished by applying a teflon coating to the entire inner surface of the MS assembly. Special care was also taken to assure that all fastening rivets be either flush with or below the teflon surface of the MS. In addition to considerations of friction, the elimination of rivet head protrusions was important in not damaging the rather delicate gold surface used to provide the proper emissivity of the outer OWS wall surfaces as mentioned above. This was a vapor deposited gold surface applied to a kapton backing and bonded to the outer workshop wall with an adhesive. A typical cross section through the entire workshop wall members is shown in figure 4-16.

Panel Details

The 16 panels comprising the meteoroid shield were formed of 0.025 inch thick aluminum stock fitted with doublers and angles to permit their assembly. A typical detail of the longitudinal joints between the sixteen panels is shown in figure 4-17. In each of these panel joints, 96 holes of 1/8-inch diameter were drilled to vent any air trapped under the MS skin. In detail B of figure 4-17 is shown the special panel joint required next to the SAS-1 wing because of the unavailability of sufficiently wide panel stock for the panel under SAS-1. It was a strap" of metal of this special joint that became embedded in the SAS-1 cover and prevented automatic deployment of SAS-1 in orbit. It is, perhaps, of passing interest to note the longer length of exposed bolts in this particular joint.

Around the top of the panels is located an angle and a neoprene rubber rain or weather seal as shown in figure 4-18. This seal was not intended to be an aerodynamic seal and could not be expected to accommodate significant relative deflections between the OWS and MS surfaces. To provide meteoroid protection at the two ends of the MS, small strips of thin stainless steel "fingers" were squeezed down between the OWS and the MS when stowed. These fingers, deployed, are visible in the photograph of figure 4-10. The thrust load of the

MS, which weighs some 1200 pounds, is transferred to the forward flange of the aft skirt through a group of twelve thrust blocks as shown in figure 4-19. Figures 4-20 and 4-21 depict the.MS as laid out flat to identify the relative locations of the various panels, openings, joints and other features of the complete assembly.

    Figure 4-1. - Meteoroid shield.   

    Figure 4-2. - Butterfly hinges which connect meteoroid shield to straps running under main tunnel.   

    Figure 4-3. - Photograph of titanium frame springs in auxiliary tunnel.   

    Figure 4-4. - Trunnion strap assembly as used in rigging   

    Figure 4-5. - Meteoroid shield deployment ordnance and foldout panels.   

    Figure 4-6. - Meteoroid shield in its stowed or rigged condition for launch.   

    Figure 4-7. - Meteoroid shield partially deployed.   

    Figure 4-8. - Meteoroid shield deployed for orbit.   

    Figure 4-9. - . Ordnance schematic and cross section view for meteoroid shield release.   

    Figure 4-10. - Photograph showing typical swing link and latch detail.   

    Figure 4-11. - Drawing of typical swing link and torsion rod assembly.   

    Figure 4-12. - Assembly view of auxiliary tunnel.   

    Figure 4-13. - Wiring tunnel for TACS running inside auxiliary tunnel.   

    Figure 4-14. - Views showing vent area provision for auxiliary tunnel.