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ON MAY 14, 1973

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JULY 13, 1973

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 Chapter Table of Contents
Board Membership
Figure List
Table List
I. The Skylab Program
II. The Flight of Skylab 1
III. Detailed Analysis of Flight Data
IV. The Meteoroid Shield Design
V. Highlights of the Meteoroid Shield History
VI. Analysis of Possible Failure Modes of the Meteoroid Shield
VII. Postulated Sequence of the Most Probable Failure Mode
VIII. Possible Impact of Costs and Schedules on the Meteoroid Shield
IX. The Skylab Management System
X. Significant Findings and Corrective Actions
 A Letter Establishing Board
 B Acronyms and Symbols




    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

Counsel to the Board

    S. Neil Hosenball Deputy General Counsel, NASA Headquarters

Executive Secretary

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

Project Liaison

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

Aerospace Safety Advisory Panel Observer

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

Technical Consultants

    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.

1-1     Skylab Cluster
1-2     SL-1 Vehicle
1-3     Skylab Mission Profile
2-1     Scalar Wind Speed at Launch Time of SL-1
2-2     Altitude Comparisons for Early Portion of S-V Launch Vehicle Boost Trajectory
2-3     SL-1 Alpha-Beta Limit at Maximum Bending Moment
2-4 A    Meteoroid Shield and Solar Array System (Stowed)
2-4 B    Meteoroid Shield and Solar Array System (Deployed)
2-5     Meteoroid Shield and Instrumentation Layout
3-1     Roll Rate Versus Range Time
3-2     Time Sequence of 63-Second Anomaly Instrumentation
3-3     Condition of Meteoroid Shield Instrumentation at R+60.90 Seconds
3-4     Condition of Meteoroid Shield Instrumentation at R+62.78 Seconds
3-5     Condition of Meteoroid Shield Instrumentation at R+62. 89 Seconds
3-6     Condition of Meteoroid Shield Instrumentation at R+62. 90 Seconds
3-7     Condition of Meteoroid Shield Instrumentation at R+62. 97 Seconds
3-8     Condition of Meteoroid Shield Instrumentation at R+64. 88 Seconds
3-9     SL-1 Retro-Rocket Impingement Force Schematic for S-II/SWS Separation
3-10     593 Second Anomaly Time Sequence
3-11     Explanation of 593 Second Anomaly
3-12     Explanation of 593 Second Anomaly
3-13     Plume Impingement Force on SAS-2
3-14     SAS-2 Wing Hinge
3-15     Engine Compartment Gas Temperature
3-16     Base Region Pressures - Assumed Failure Mode: Interstage Did Not Separate
3-17     Separation EBW Firing Unit Monitor Indications
3-18     Second Plane Separation System, S-11 (block diagram and location)
3-19     EBW Detonator and Detonator Blocks, Second Plane Separation System, S-II (installation)
3-20     S- 11- 13 Interstage Station 196 Tension Strap Analysis
3-21     Forward Interstage Internal Pressure
4-1     Meteoroid Shield
4-2     Butterfly Hinges Which Connect Meteoroid Shield to straps Running Under Main Tunnel
4-3     Photograph of Titanium Frame Springs in Auxiliary Tunnel
4-4     Trunnion Strap Assembly As Used In Rigging
4-5     Meteoroid Shield Deployment Ordnance and Foldout Panels
4-6     Meteoroid Shield In Its Stowed or Rigged Condition for Launch
4-7     Meteoroid Shield Partially Deployed
4-8     Meteoroid Shield Deployed for Orbit
4-9     Ordnance Schematic and Cross Section View for Meteoroid Shield Release
4-10     Photograph Showing Typical Swing Link and Latch Detail
4-11     Drawing of Typical Swing Link and Torsion Rod Assembly
4-12     Assembly View of Auxiliary Tunnel
4-13     Wiring Tunnel for TACS Running Inside Auxiliary Tunnel
4-14     Views Showing Vent Area Provision for Auxiliary Tunnel
4-15 A  Photographs of Auxiliary Tunnel Boot (Stowed)
4-15 B  Photographs of Auxiliary Tunnel Boot (Deployed)
4-16     Typical Cross Section Through Members of the Orbital Workshop Wall
4-17     Longitudinal Joint Detail of AIS
4-18  &>

Transfer interrupted!

 Rain Seal at Typical Top End of MS Flange
4-19     Thrust Block Detail (one of twelve)
4-20     Meteoroid Shield Laid Flat
4-21     Meteoroid Shield Laid Flat
6-1     View of Kapton Surface of the OWS Showing Forward Torsion Rod Swing Link
6-2     View of Kapton Surface of the OWS Showing Aft Torsion Rod Swing Link and Thrust Blocks
6-3     Auxiliary Tunnel Frame Spring Stiffness
6-4     Venting Locations in Meteoroid Shield
6-5     Ordnance Foldout Panel
6-6     Longitudinal Section Through Meteoroid Shield at Foldout Panel
6-7     Skylab, (SL-1, SA-513) Dynamic Pressure Profile for Boost Phase
6-8     Meteoroid Shield Area Design Differential Pressures for Smooth Configuration
6-9     SL-1 Auxiliary Tunnel Design Differential Pressures
6-10     Auxiliary Tunnel Forward Vent
6-11     Meteoroid Shield Response - Aft Auxiliary Tunnel Boot Sealed
6-12     Auxiliary Tunnel Leaks
6-13     Meteoroid Shield Response - Aft Boot Leakage
6-14     Compressibility Waves from the Forward Auxiliary Tunnel Fairing
6-15     Mathematical Model for Meteoroid Shield Divergence Analysis
6-16     Air Bladder Test Rig for Tunnel Deflection Tests
7-1     Definition of Axes and Positive Rotations
7-2     Possible Meteoroid Shield Motion from 60.12 Seconds to 62. 74 Seconds
7-3     Sketches of Possible Shield Dynamics During the 63 Second Anomaly
7-4     Photograph from Orbit Showing Longitudinal Alurninum Angle Bent Over the SAS-1 Wing



Tables appear following the figures at the end of each Chapter

I-1  Major Skylab Contractors
II-1  Maximum wind speed in high dynamic pressure region for Apollo/Saturn 501 through Saturn 513 vehicles
II-2  Extreme wind shear values in the high dynamic pressure region for Apollo/Saturn 501 through Saturn 513 vehicles
II-3  Orbit parameters
II-4  Normal major events
IV-1  OWS meteoroid shield swing link settings and measurements
IX-1  Orbital Workshop Program Meteoroid Shield Design Reviews
IX-2   Orbital Workshop Program Solar Array System Design Reviews



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.




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:

1. Marshall Space Flight Center (MSFC)

a. Performing overall systems engineering and integration to assure the compatibility and integration of the total mission hardware for each flight and for the orbital assembly.

b. Developing elements of the flight, hardware and related software, including: S-IB and S-V launch vehicles, OWS, AM, MDA, AM and payload shroud.

c. Developing assigned experiments and supporting hardware and integrating them into the flight hardware.

d. Supporting Kennedy Space Center (KSC) and Johnson Space Center (JSC) flight operations and performing mission evaluation.

2. Johnson Space Center (JSC)

a. Implementing all flight and recovery operations, including: mission analyses and associated systems engineering, related ground equipment and facilities, preflight preparations, and conducting the flight and recovery.

b. Providing and training flight crews and developing crew and medical requirements.

c. Developing elements of the flight hardware and related software. including: modified command and service modules, spacecraft launch adapter for manned launches, trainers and simulators. crew systems, medical equipment and food.

d. Developing assigned experiments. integrating those to be carried in the CSM, and providing for stowage of experiment data and hardware designated for return from orbit.

d. Performing mission evaluation.

3. Kennedy Space Center (KSC)

a. Providing launch facilities for the four Skylab 1 launches.

b. Preparing checkout procedures and accomplishing the pre-launch checkout of flight hardware and ground support equipment

c. Planning and executing launch operations.

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

fig_1_1.jpg (244302 bytes)    Figure 1-1   

fig_1_2.gif (31226 bytes)    Figure 1-2   

fig_1_3.gif (62084 bytes)    Figure 1-3 - Skylab mission profile   



Contractor Responsibility

Contract amount $ millions

Rockwell International Command and service module 354.3
General Electric Automatic checkout equipment reliability and quality assurance system engineering. 29.7
Martin Marietta Corp Payload and experiments integration and spacecraft support. 105.4
The Garrett Corp Portable astronaut life support assembly 11.9
International Latex Corp Space suits 16.9
ITEK Corp S190 - Multispectral photo facility 2.7
Black Engineering, Inc S191 - Infrared spectrometer 2.0
Cutler Hammer Airborne Instrument Lab S194 - L-band radiometer 1.5
General Electric S193 - Microwave radiorneter / scatterometer 11.3
Honeywell Corp S192 - 10-band multispectral scanner 10.8
Martin Marietta Corp Program support 11. 1
General Electric Electrical support equipment and logistics support 25.0
McDonnell Douglas S-IVB stage 25.7
Martin Marietta Corp Payload integration and multiple docking adapter assembly 215.5
Rockwell International (Rocketdyne Division) Saturn engine support-Saturn V and Saturn 1B 10.3
IBM Apollo telescope mount digital computer and associated items 29.2
Chrysler S-I B stage 30.0
  S-18 systems and integration 7.0
McDonnell Douglas-West Orbital workshop 383.3
McDonnell Douglas-East Airlock 267.7
General Electric Launch vehicle ground support equipment 12.6
IBM Instrument unit 30.7
Boeing S-IC stage 0.9
  System Engineering and integration 7.4
American Science & Engineering X-Ray spectrographic telescope - S054 8.3
High Altitude Observatory White light coronagraph - S052 14.7
Harvard UV spectrometer - S055 34.6
Naval Research Laboratory UV spectrograph / heliograph 40.9
Goddard Space Flight Canter Dual X-ray telescope 2.5
Chrysler Corp S-IB launch operations support 23.2
Boeing Co Saturn V launch vehicle and launch, complex 39, launch operations 14.4
Rockwell International Command and service module support 17.5
McDonnell Douglas S-IVB launch services 58.9
IBM Instrument unit, launch services 12.3
Delco Electronics Navigation and guidance launch operations 0.9
Martin Marietta Corp Multiple docking adapter support 7.2
Aerojet General Corp CSM service propulsion system (SPS) rocket engines 3.1
AiResearch Manufacturing Co CSM environmental control systems (ECS) 5.6
Aeronca Inc CSM honeycomb panels 1.5
AVCO Corp Command module heat shields 2.5
Beech Aircraft Corp CSM cryogenic gas storage system 4.0
Collins Radio CSM communications and data systems 4.7
Honeywell Inc CSM stabilization and control systems 3.1
Marguardt Co Service module reaction control system (RCS) engines 1.1
Northrop Corp Command module Earth landing system 0.8
Pratt & Whitney Aircraft CSM fuel cell powerplants 3.2
Bell Aerospace Co RCS propellant storage tanks 3.4
Simmonds Precision Products, Inc Propellant utilization gauging system 1.3
TRW Solar array system 23.7
Fairchild Miller Habitability support system 19.0
Hamilton Standard Division of United Aircraft Corp Centrifugal urine separators 9.6
Hycom Manufacturing Co Orbital workshop viewing window 0.9
AiResearch Manufacturing Co Molecular sieve 4.7




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.

fig_2_1.gif (36871 bytes)    Figure 2-1. - Scalar wind speed at launch time of SL-1.   

fig_2_2.gif (40936 bytes)    Figure 2-2. - Altitude comparisons for early portion of S-V launch vehicle boost trajectory.   

fig_2_3.gif (49285 bytes)    Figure 2-3. - SL-1 alpha-beta limit at maximum bending moment.   

Fig_2_4_a.gif (9571 bytes)   Figure 2-4. - Meteoroid shield and solar array system.  (a) STOWED.   

Fig_2_4_b.gif (15370 bytes)   Figure 2-4. - Meteoroid shield and solar array system.  (b) DEPLOYED.   

fig_2_5.gif (115381 bytes)    Figure 2-5. - Meteoroid shield and instrumentation layout.   



Vehicle No.




m/s (knots)




Km (ft)

Pitch, Wx,

M/s (knots)


Km (ft)

Yaw, Wz

M/s (knots)


Km (ft)






(37 700)




(37 700)




(29 500)






(42 600)




(42 650)




(51 700)






(49 900)




(49 500)




(51 800)






(38 480)




(38 390)




(37 500)






(46 520)




(45 280)




(48 720)






(37 400)




(36 680)




(39 530)






(46 670)




(46 670)




(44 780)





13. 58

(44 540)




(44 540)




(42 570)


52. 8




(43 720)




(43 720)




(33 460)






(45 110)




(45 030)




(44 040)






(38 880)




(38 880)




(50 850)






(39 945)




(39 945)




(37 237)










(42 732)




(41 584)



(Delta h = 1000 M)



Pitch Plane

Yaw Plane

Shear Sec-1


Km (ft)

Shear Sec-1


Km (ft)




(32 800)



(32 800)




(48 900)



(43 500)




(52 500)



(51 800)




(49 700)



(48 160)




(50 200)



(50 950)




(48 490)



(33 790)




(46 750)



(47 820)




(50 610)



(45 850)




(43 720)



(38 880)




(36 830)



(47 330)




(44 780)



(50 850)




(26 164)



(34 940)




(46 095)



(30 347)


Table II-3 -- Orbit Parameters *

Parmeter Actual Predicted Difference Between Actual and Predicted
Apogee, nautical miles 234.5 233.8 0.7
Perigee, nautical miles 233.8 233.8 0
Inclination, degrees 50.06 50.00 0.06
Ascending Node, west longitude 129.90 129.90 0

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


Table II-4 -- Normal Major Events

Major Event

Actual Time From R=0 Seconds

Predicted Time From R=0 Seconds

Difference Between Actual and Predicted Seconds

Guidance Reference Release (GRR) - 17.0 - 17.0 0
S-IC Engine Start Sequence Command - 8.9 - 8.9 0
Range TimeZero (1730:00) 0 0 0
All Holddown Arms Released 0.2 0.2 0
Liftoff, Begin Time Base 1 0.586 0.520 0
Begin Tower Avoidance Pitch and Yaw Maneuver 1.6 1.5 0.1
End Tower Avoidance Pitch Maneuver 5.8 5.7 0.1
Begin Pitch and Roll Program 12.2 11.2 1.0
S-IC Outboard Engine Cant 20,5 20.5 0
Mach 1 61.1 61.5 - 0.4
Maximum Dynamic Pressure (Max Q) 73.5 75.0 - 1.5
S-IC Center Engine Cutoff (CECO) 140.7 140.6 0.1
Begin Time Base 2 140.8 140.7 0.1
S-IC Outboard Engine Cutoff Enable 152.4 152.4 0
Begin Tilt Arrest (Stop Pitch) 158.1 157.1 1.0
S-IC Engine 1 and 3 Cutoff 158.2 158.2 0
S-IC Engine 2 and 4 Cutoff 158.2 158.2 0
Begin Time Base 3 158.2 158.2 0
S-IC/S-II Separation 159.9 159.9 0
S-II Engine Start Sequence Command 160.6 160.6 0
Arm-1, S-II Aft Interstage Separation 183.2 183.2 0
Arm-2, S-II Aft Interstage Separation 183.3 183.3 0
S-II Aft Interstage Separation Commiand-1 (Second Plane Separation Command 1) 189.9 189.9 0
S-II Aft Interstage Separation Command-2 (Second Plane Separation Command-2 [Backup]) 190.0 190.0 0
Start Iterative Guidance Mode (IGM) Phase 1 197.1 196.2 0.9
Start Steering Misalignment Calculation 216.4 216.8 - 0.4
S-II Center Engine Cutoff 314.0 314.2 - 0.2
Start IGM Phase 1:1 314.5 314.3 0.2
S-II Engine Mixture Ratio (EMR) Shift 403.7 402.6 1.1
Start IGM Phase III 404.0 402.5 1.5
Begin Terminal Steering 568.8 563.7 5.1
S-II Outboard Engine Cutoff (OECO) 589.0 588.3 0.7
Begin Time Base 4 589.2 588.5 0.7
S-II/Saturn Workshop (SWS) Separation Command / Fire Retro Motors-1 591.1 590.5 0.6
S-II/(SWS) Separation Command / Fire Retro Motors-2 (Backup) 591,2 590.6 0.6
Initiate S-II Timer 591.2 590.6 0.6
Orbit Insertion 599.0 598.3 0.7
Start Local Reference Maneuver (Local Vertical Attitude) 599.6 598.5 1.1
Initiate S-II Safing vent 805 1 800.6 4.5
Start Payload Shroud Jettison / Begin Time Base 4A 919.2 932.3 - 13.1
Payload Shroud Jettison 920.4 934.0 - 13.6
Start Solar Inertial Maneuver 958.8 972.3 - 13.5
Initiate ATM Deployment 999.1 998.5 0.6
Initiate ATM Solar Arrays Deployment 1492.3 1491.7 0.6
ATM Telemetry On 2209.1 2208.5 0.6
Initiate OWS Solar Array System Deployment 2465.7 2465.1 0.6
Initiate MS Deployment 5764.1 5763.5 0.6
Thruster Attitude Control System (TACS) Command Transfer to ATM 17400.7 17400.1 0.6
Begin Time Base 5 29399.5 29398.6 0.9




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.

fig_3_1.gif (48914 bytes)    Figure 3-1 - Roll rate versus range time.   

fig_3_2.gif (101478 bytes)    Figure 3-2.  Time Sequence of 63-sec Anomaly Instrumentation   

fig_3_3.gif (101049 bytes)    Figure 3-3. - Condition of meteoroid shield instrumentation at R + 60.90 sec   

fig_3_4.gif (99370 bytes)    Figure 3-4. - Condition of meteoroid shield instrumentation at R + 62.78 sec   

fig_3_5.gif (97535 bytes)    Figure 3-5. - Condition of meteoroid shield instrumentation at R + 62.89 sec.   

fig_3_6.gif (99184 bytes)    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.   




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.   

    Figure 4-15. - Photographs of auxiliary tunnel boot. (a) STOWED POSITION.   

    Figure 4-15. - Photographs of auxiliary tunnel boot. (b) DEPLOYED POSITION.   

    Figure 4-16. - Typical cross section through members of the orbital workshop wall.   

    Figure 4-17. - Longitudinal joint detail of MS.   

    Figure 4-18. - Rain seal at typical top end of MS flange.   

    Figure 4-19. - Thrust block detail (one of twelve).   

    Figure 4-20. - Meteoroid shield laid flat.   

    Figure 4-21. - Meteoroid shield laid flat.   



Swing Link No.

Swinglink Effective Length


Flight Angles

Notes on Flight Values

Torque Angle

Swing Angle

Residual Angle

SAS-1 Released But Jammed

SAS-1 Deployed

1 6.8" F

6.8" A

1800 F

1800 A

1390 F

1390 A

410 F

410 A

65.570 F

85.730 A

136.710 F

130.840 A





Exists of:






2 8.7" F

8.7" A

1650 F

1650 A

150.50 F

150.50 A

14.50 F

14.50 A

86.030 F

161.110 A

84.380 F
3 10.5" F

10.5" A

1730 F

1730 A

157.80 F

157.80 A

15.20 F

15.20 A

172.420 F

167.230 A

4 12.5" F

12.5" A

1800 F

1800 A

163.30 F

163.30 A

16.70 F

16.70 A

173.340 F

179.510 A

5 12.55" F

12.55" A

1800 F

1800 A

162.90 F

162.90 A

17.10 F

17.10 A

186.490 F

173.450 A

6 10.1" F

10.1" A

1730 F

1730 A

152.40 F

152.40 A

20.60 F

20.60 A

165.450 F

163.250 A

173.450 F
7 8.2" F

8.2" A

1650 F

1650 A

148.30 F

148.30 A

No stop

16.70 F

16.70 A

171.010 F

165.260 A

8 6.9" F

6.9" A

1800 F

1800 A

1370 F

1370 A

430 F

430 A

144.300 F 143.420 F






The meteoroid shield of the OWS had its origin in a letter request of November 1, 1966 from MSM to the Douglas Aircraft Company to submit an expedited ECP for a description of "systems feasible as protection against probable meteoroids." A brief proposal was submitted to MSFC in response to this request on December 7, 1966. There followed a submittal of design criteria for the MS by MSFC stating, among other things, that it "shall be designed as a structurally integrated part of stage 209 capable of withstanding the dynamic forces imposed during the orbital workshop mission" and that "the weight of the bumper system shall be a primary design consideration." Protection from meteoroid penetration with a probability of 0.9950 of no penetration for a 12-month mission was also specified. The ECP of December 7, 1966 was approved on March 16, 1967 and work on the design and development of the shield was initiated, leading to the following project milestones and events.

May 2-10, 1967 Preliminary Design Review on the orbital workshop.
December 12-15, 1967 A "Delta" Preliminary Design Review on the orbital work shop.
July 22, 1969 NASA decision to change from a "wet" Saturn S-IVB launched workshop to a "dry" Saturn V launched workshop. At the same time, "Skylab" as the program was to be later designated, became a major line-item program in its own right, independent of Apollo funding and schedules.
August 20, 1969 Supplemental agreement No. 2 to contract NAS9-6555 authorizing the change from the "wet" to the "dry" workshop. In accordance with Skylab program policy that no unnecessary changes be made, and based on the full confidence of MSFC and MDAC-W management in the current design, the basic concept and general design of the meteoroid shield was retained.
August 1969 Provisions made in the meteoroid shield for the ground access door, the wardroom window and the scientific airlocks.
November 1969 The introduction, via a MDAC-W internal memo, of vents and internal baffles in the auxiliary tunnel to reduce aero-dynamic loads on the tunnel frames.
December 1969 Cluster System Review
September 14-18, 1970 Critical Design Review of orbital workshop.
November 1970 - April 1971 The "Mathews" Subsystem Review
January 1971 Rework of the butterfly hinges and adjacent panels of the MS to accommodate a misalignment in the bonding of the tension straps to the OWS.
March - April 1971 First reference to possibility of meteoroid shield 'burst" pressures in an internal MDAC-W memo; venting of shield by drilling holes in the panel joints proposed as solution in a responding memo.
March 17-18, 1971 Critical Mechanisms Review
April 29, 1971 Completion of re-qualification of expandable tube explosive assembly.
May 10, 1971 First deployment test at MDAC W. Shield failed to fully deploy because of gravity loading on the linkage mechanisms.
May 14, 1971 Test failure of expandable ordnance tube requiring re-design and retest.
May 18, 1971 Completion of qualification testing of MS ordnance system.
August 1971 An internal MDAC-W memo suggesting a change to an unbaffled auxiliary tunnel with the aft end sealed and a vent at the forward end.
September 14, 1971 ECP for design changes arising, in part, from the failure of deployment test. Changes included improvements to ordnance, in creased torque on swing links, increased auxiliary tunnel frame thickness and redesigned hinge pins. Decision to conduct all future deployment tests at MSFC on the static test article (STA).
February 10, 1972 Completion of re-qualification of expandable tube and strap assembly after tube wall rupture.
March 1972 Completion of series 1 deployment tests. Distance between fold-over panel and tank insufficient to provide meteoroid protection. Rigidity of shield also insufficient to cause panel to chord. Scroll springs accordingly added to four corners of the fold-over panel.
April 1972 Completion of series 2 deployment test. Latches failed to engage. Latch modified and spring relocated to provide greater angular motion.
April 15, 1972 Qualification of full-scale (22 feet) redesigned expandable tube and strap assembly on back-up structure.
May 31, 1972 Doublers added to butterfly hinges on SAS-2 side of main tunnel because of failure of hinges during tank pressurization test at MSM.
May 1972 Special development test on deployment latch mechanism.
June 19-23, 1972 Design Certification Review of orbital workshop. August 1972 The completion of four mechanical and three ordnance deployment tests at MSFC on STA using OWS- 1 MS flight hardware.
September 8, 1972 OWS- 1 shipped to KSC. Arrives on September 22, 1972.
October 28, 1972 Discrepancy Report DR 0136 developed at KSC due to difficulty encountered in rigging the shield tight to the tank (see next section entitled "Rigging the Meteoroid Shield for Flight").
November 10, 1972 Discrepancy Report DR 180 written at KSC on gaps between the shield and tank. Contains detailed mapping of the areas of such gaps and MRB disposition to "use as is."
April 3, 1973 Hardware Integrity Review
April 18, 1973 Flight Readiness Review
May 14, 1973 Launch

As may be noted from the foregoing, the meteoroid shield experienced a remarkable stability of design, both in basic concept and in most of its design details, throughout its six year history. Principal development difficulties, and hence engineering effect, concerned the achievement of an ordnance system that would release the shield without a rupture of the expandable tube and in several minor details of deployment such as locating latches or obtaining the desired deployment distances. The major structural failure experienced in testing was of the butterfly hinge, which was more of a ground test than a flight problem, and was readily solved by the addition of hinge doublers.

No deployment tests were conducted under vacuum conditions, which is quite acceptable in view of the low rate of motion of the deployment. Vibration, acoustic and flutter tests were specifically omitted in the test specifications because of the design requirement that the shield be "tight to the tank." This design requirement and pervading philosophy of design and development also served to omit all aerodynamic tests of the meteoroid shield. The major difficulty experienced with the meteoroid shield was in getting it stowed and rigged on the OWS. Handling such a large, lightweight structure proved difficult, requiring the coordinated action of a large group of technicians, and considerable adjustments to the assembly of the various panels were necessary in an effort to obtain a snug fit between the shield and the OWS wall. The specific procedure used for rigging the shield for flight is discussed in the following section.

Rigging the Meteoroid Shield for Flight

The condition of the MS at launch is sensitive to the rigging procedures used to secure the shield around the OWS. For this reason, the rigging procedures used at KSC to prepare the US for launch are summarized in the following.

The flight MS was shipped to KSC from MDAC-W fully installed, but not flight rigged, on the OWS except for the installation of the ordnance panel. This panel was later installed at KSC after the mechanical deployment tests were completed. The deployment tests which were conducted earlier at MSFC utilized a static test article (STA) version of the shield.

The rigging procedure that was to be used at KSC was developed jointly by MSFC and MDAC using the STA at MSFC. The STA shield was, however, different from the flight MS in four significant aspects. On the flight MS: (1) the double butterfly hinges on the SAS 1 side of the main tunnel were bonded to the tension straps while on the STA they were present but unbonded; (2) the butterfly hinges on each side of the main tunnel were cut in the middle of a longitudinal joint and refitted to the adjacent panels at a slight angle as mentioned earlier. The longitudinal edges of the panels were also modified to suit the altered hinge line. This change to the flight MS at MDAC was necessary to accommodate the misalignment which occurred in the location of the tension straps on the OWS; (3) a longitudinal misplacement of the tension straps of 0.15 inch too high also resulted in some binding of the forward weather seal and torsion rods that had to be refitted at KSC; and (4) the trunnion bolts. nuts and washers were initially not lubricated on either the flight MS or the STA. This lack of lubrication caused difficulties in the final rigging of the shield at KSC, which was subsequently corrected by applying a solid film lubricant.

The MS delivered for flight was therefore not identical to that used to develop the rigging procedures. In this sense, the flight shield was also the development and qualification unit.

An abbreviated version of the rigging procedure developed at MSFC on the STA is presented below. This procedure was used initially at KSC in rigging the flight MS.

1. Install zero-g kit.
2. Rotate shield against tank wall. (This process required the coordinated action of a group of technicians, one each at the forward and aft end of each of the sixteen panels.)
3. Loosely install trunnion bolts.
4. Remove auxiliary tunnel covers.
5. Spread each of the 28 auxiliary tunnel frames 1-3/8 inches + /- 1/8 inch with a spreading fixture.
6. Adjust trunnion bolt nuts finger tight, with distance between trunnion straps uniform within + 0.10 inch.
7. By sequential end-to-end passes, torque the center 12 trunnion bolts to a value of 100 inchpounds. This condition to be reached while maintaining a uniform spacing between trunnion straps of less than + 0.10 inch. Sequential adjustments of all bolts to be such that at steps of 50 and 75 inch-pounds, an equal torque on all 12 bolts is obtained.
8. Remove the auxiliary tunnel spreading fixture.
9. Back off the 12 trunnion bolt nuts 3 revolutions.
10. Retorque the 12 trunnion nuts up to 1 revolution, not to exceed 45.0 inch-pounds. (Revolution was the controlling factor.)
11. Torque the top and bottom trunnion bolt nuts to 20 inch-pounds.
12. Torque the trunnion bolt jam nuts to 100 inchpounds and lock wire.

This rigging procedure did not produce a satisfactory fit of the field to the OWS tank wall. Several "bulges" were evident where the shield was not snug to the tank and significant gaps, of up to an inch in extent at one point, existed between the shield and the tank along the upper and lower edges of the shield assembly. The rigging procedure furnished to KSC therefore had to be modified considerably in an effort to produce a tightly fitting shield. These modifications are noted in a discrepancy report (D.R. No. 0136), a summary of which is presented below.

1. Loosen seal retainers on bottom of two pairs of panels.
2. Loosen panel joint bolts along bottom 36 inches of a pair of panels and push panels against tank. Retighten bolts.

(A gap still existed.)

3. Loosen 6 bolts in the other pair of panels, install a "puller" to one of these bolts, and pull panel out in mid-region and push bottom in against tank. Repeat this pushing and pulling 10 times.
4. While pushing the shield against the tank, loosen 4 other bolts and tighten the others. Push and pull again in this region 10 times.
5. Repeat step 4 with another pair of bolts loose. Re-tighten 2 bolts.
6. Repeat step 5 with another pair of bolts loosened.
7. Continue this procedure until one has worked his way up along the entire length of two longitudinal joints between two pairs of panels. Remove the "puller" and replace flight bolts.
8. Tighten bottom two trunnion bolts 2 turns, with torque not to exceed 45 inch-pounds.
9. Tighten third from bottom trunnion bolt 1 turn, torque not to exceed 45 inch-pounds.
10. Remove shims on bearing blocks, loosen seal retainers on two more panels, and bolts from a strap on the ordnance panel.
11. Loosen three bottom trunnion bolts (i.e., undo steps 8 and 9).
12. Measure tangential length of ordnance panel, the gap between trunnion straps, and the spread of tunnel frames.
13. Torque bolts to 90 inch-pounds.
14. Repeat measurements of step 12.
15. Install flight panel over access door.
16. Remove top and bottom trunnion bolts.
17. Loosen up and re-torque the center 12 bolts to 45 inchpounds.
18. Remove these 12 bolts, apply lubricant, and torque to 45 inch-pounds.
19. Torque up the trunnion bolts in sequential endto-end passes to 96 inch-pounds, maintaining a uniform gap between straps within + 0.06 inch.
20. Back off trunnion bolt nuts and torque to 45 inch-pounds.
21. Install top and bottom trunnion bolts, torque to 20 inch-pounds.
22. Install jam nuts and lock wire.
23. Measure and record torsion bar / strut fitting relationship (all okay).
24. Put the bulb seal, back in place.

(A gap still existed along edge of panel.)

25. Remove splice plate between two panels (one of the pair previously re-worked) and slot holes in shield flange.
26. Reinstall splice plate loosely and push in on the panel joint. While pushing, tighten up the splice plate.
27. Set the bulb seal against the tank.
28. Verify torque on aft bolt at 20 inch-pounds, re-install jam nuts and lock wire.

At the conclusion of the rigging described above, the contact areas, that is, the areas of the shield which were snug to the tank, were mapped, and it was determined that only 62% of the shield was in contact with the unpressurized OWS. The OWS was then pressurized to 8 psi above ambient and the contact areas again mapped. In this condition, about 95% of contact was achieved. Much of the remaining gapping occurred along the forward edge of the MS. Since the flight differential pressure was substantially higher than 8 psi, it was felt that the contact area during the flight would be substantially higher than 95%. The condition of the MS was therefore formally accepted as satisfactory for flight on January 10, 1973. No further adjustments were made to the MS prior to the flight.

From the foregoing, it is apparent that the MS was very difficult to rig to the tank. Many hands were required to push and pull on various joints in the shield while groups of panel bolts were sequentially loosened and tightened in an effort to obtain a snug fit. As a result, the final rigging used prior to flight differed markedly from that used any previous time. More importantly, the actual condition of the shield in terms of final tension in the tunnel springs or in the trunnion bolts were uncertain at best. Some gaps undoubtedly existed between the forward and aft ends of the shield and the tank walls at the time of launch, which could well have increased as the flight progressed due to the non-uniform growth of the tank.




During the course of this investigation, a total of ten possible failure modes of the meteoroid shield were postulated and examined by the Board. These failure modes are presented and discussed, in increasing order of probability, in the followi>

Transfer interrupted!



Premature Firing of the Meteoroid Shield Separation Ordnance  

The flight data shows that the three ordnance break wire event sensors on the ordnance strap were still intact after the first indication of the MS anomaly. The in-flight firing command was therefore not issued prematurely. The ground firing command to fire the ordnance was not given. It was therefore concluded that the failure of the MS was not a result of a premature firing of the separation ordnance.    


Failure of Butterfly Hinges  

The flight load on the butterfly hinges at 63 seconds after lift-off was calculated to be 44 lbs/inch. This load is created by pretensioning of the auxiliary tunnel frames prior to lift-off and by the circumferential growth of the OWS due to internal pressurization. Assuming uniform circumferential loading, the factor of safety of the butterfly hinge on the auxiliary tunnel side was approximately 5 and on the SAS-1 side was 11. Because of the friction between the MS and OWS that must be overcome in rigging the shield for flight, the ground loads on these hinges was undoubtedly higher than the flight loads. This fact, and the high factor of safety of the hinges, leads to the conclusion that they did not initiate the failure.

3. Failure of the Trunnion Bolts or Straps
Because of the friction between the MS and OWS, loads greater than the flight load were probably imposed m the trunnion system in torquing up the bolts for ground rigging. The calculated flight load, without friction, on the trunnion bolts and straps at the 63 second event was 44 lbs/inch (see fig. 4-4). This load was arrived at by assuming that during the boost phase there was sufficient acoustic and vibration energies to distribute the load circumferentially even in the MS. With a factor of safety of 6 in the trunnion bolts, 4 in the trunnion straps and 3. 5 in the trunnion strap rivets, the Board concluded that the trunnion system did not initiate the failure.
4. Failure of MS due to Thrust Block Slippage
There are twelve thrust block supports approximately equally spaced circumferentially between the aft end of the MS and the forward end of the aft OWS skirt to take the MS boost loads. The fourteen tension straps under the main tunnel are bonded to the OWS wall and attach to the MS through the butterfly hinges. They also take a portion of MS boost loads. The mating support blocks are machined at a 10 degree angle so that during a normal deployment there will be minimum separation interference. A simple test showed that this 10 degree slope could produce a radial displacement of the MS away from the OWS wall when subjected to vibratory loads. This radial movement is, however, resisted by the pretensioning of the shield by the auxiliary tunnel frames, along with the additional tension buildup due to the thrust blocks moving radially outward under tank pressurization. The bonded tension straps that also support the MS introduces a binding-friction cockedload between the OWS wall and shield when the shield is subjected to boost loads. From photographs taken of the swing links in orbit by the astronauts (see figs. 6-1 and 6-2) it is evident that none of them were bent downward to an extent which would indicate that the shield slid down over the aft skirt. Consideration of the above facts leads to the conclusion that a slippage over the thrust blocks did not initiate the failure.
5. Buckling Failure at Support Blocks
Two other potential failure modes exist in relation to the thrust blocks. Conceivably, a failure might result from buckling and cripling of the longitudinal panel joints due to high concentrated loads at the block support points. This could take two forms:

a. A barrell-like mode of buckling failure where the shield deflects into a circumferentially oriented corrugated or bellow like shape and slides down the OWS wall.

b. A crippling mode of failure where the flanges of the longitudinal panel joints deflect laterally (circumiferentially) and collapse.

Each of these modes is hindered by friction between the meteoroid shield and the workshop wall. In addition, the butterfly hinge would force asymmetric docking of the MS, resulting in increased friction and diagonal forces which would tend to prevent buckling.
Elementary analyses have been made by both MDAC and LRC, including conservative estimates of radial support from hoop tension and disregarding friction, the butterfly hinge restraint, and asymmetric cocking. These elementary computations reinforced the intuitive analyses that buckling at the thrust block was not the initiating failure mode.
6. Crushing Loads on Auxiliary Tunnel Frames
Aerodynamic crush pressures on the auxiliary tunnel tend to flatten the tunnel frame springs and thus reduce the tension of the MS around the OWS. If the crush pressure is very high, the possibility exists that the preloading of the shield could be lost and the spring frame could collapse (buckle).
Data from analyses made at both MDAC and LRC are shown in figure 6-3 and indicate that the required loss in preload would not occur until the crush pressures are more than about 6 times the expected maximum crush pressure occurring at the aft end of the tunnel. The expected maximum crush pressure will reduce the preload at 63 seconds by approximately 10 percent.
Although crush pressures do slightly relax the MS preload, these results indicate that buckling or complete loss of preload is not a probable failure mode.
7. Meteoroid Shield Flutter
The possibility of panel flutter was considered in detail in two separate analyses by MSFC and MDAC. Both concluded that brief periods-of low amplitude. flutter might have occurred, but that high stresses and structural fatigue were highly unlikely. No reason to question this finding has been found and all remainig aerodynamic discussion will center on steady or slowly varying pressures.
8. Small Volumes of Entrapped Air Under the MS
Although the meteoroid shield was intended to conform tightly to the OWS (except under the auxiliary tunnel) there were many small volumes of enclosed air between the MS and the OWS which had to be vented in order to prevent development of burst pressures during ascent through the atmosphere. Chief among these are the spaces outside the wardroom window and scientific airlocks. Smaller volumes were the channels formed along the longitudinal joints between shield panels, unavoidable slight wrinkles, gaps under hinges and between the panels at the pyrotechnic foldout assembly, and at the butterfly hinges. Venting was provided by many holes drilled along the panel joints as shown in figure 6-4. In addition, the inherent construction of many details of the shield provided additional venting such as through the butterfly hinges, etc. Capacity for outflow from these enclosed volumes was considered adequate.
9. Lifting of the Forward Edge of the Meteoroid Shield
The possibility that the forward edge of. the MS projected far enough into the slip stream to experience high pressures was examined at length. Several design features of the ordnance fold-out panel make it by far the most likely candidate among all portions of the forward edge as indicated in figures 6-5 and 6-6. In this region, the total height of the MS edge above the OWS surface is greater than elsewhere because: (1) there are three layers of 0.025 inch skin instead of only one; (2) these layers are separated by stiffeners as shown; (3) the two hinges add to the bulk; and (4) a torsion link is exerting an outward force of 18 pounds on me side of this panel. In addition to these features of the fold-out panel, it must be expected that an additional standoff of about 0.12 inch resulted during flight from the swelling of the OWS due to internal pressure. When pressurized, the tank grows about 0.12 inch less in radius at the flanges than over the middle sections, and the lightly loaded shield does not conform to these new contours. The sum of all the above features results in the forward edge of the MS extending to at least within 0.11 inch of the outer surface of the forward skirt fairing. Postflight wind tunnel measurements conducted at MSFC and MDAC indicated that the design was such that high bursting pressures could have been produced during transonic flight in this region under the MS. This situation must therefore be classified as a possible failure mode, although the actual flight data indicate that another failure mode was more probable.
10. Auxiliary Tunnel Venting
When the 63 second anomaly occurred, the flight dynamic pressure was near its maximum value as shown in figure 6-7 and the velocity was near Mach 1. A diagram of the general distribution of pressures predicted to exist over the Skylab spacecraft at Mach 1 is shown in figure 6-8. These pressures are deduced from wind tunnel data and represent the pressure which would occur if the vehicle were to fly steadily at Mach 1 and there were no flow fluctuations.
The local changes in pressure produced by the auxiliary tunnel are shown schematically in figure 6-9. Dips in pressure result from expansion of the flow around the blunt base of the auxiliary tunnel forward fairing and around the base of the auxiliary tunnel.
The design intent for the meteoroid shield and its auxiliary tunnel was that the aft end of the tunnel be sealed for "no leakage" and that the forward end be vented into the base region of the forward fairing so as to discharge air into the forward low-pressure region. The 9-to-10 square inch vent area shown in figure 6-10 was intended to provide a crushing pressure (an external pressure exceeding the internal pressure) over the entire tunnel, as shown schematically on figure 6-11.
Post-flight investigation revealed that the aft end of the tunnel was, however, not completely sealed because of:

a. An unexplained omission to seal or cap two hollow structural stringers on the aft skirt which extended into the aft fairing of the auxiliary tunnel (see fig. 6-12). These stringers yielded about 2.2 square inches of leakage area.

b. An inadequate metal-to-metal fit between the aft fairing of the auxiliary tunnel and the two stringers to which it was secured resulted in approximately 2 square inches of additional leakage area. This vent area is shown in figure 6-12, and occurs at the flare of the aft fairing where the flare mates and seals with the aft end of the auidliary tunnel.

c. An unplanned venting resulting from leakage past a molded rubber boot used to seal the movable joint and rearward facing end of the auxiliary tunnel. This boot and adjacent details are shown in figure 6-12. A metal yoke provided a positive clamp to a molded flange on the bottom of the boot over the rigid aft fairing. A bonded seal was achieved between the upper molded flange on the boot and the auxiliary tunnel.

Because the auxiliary tunnel was required to lift freely away from the OWS and move circumferentially upon deployment in orbit into the position shown in figure 4-15, only a wiping butt seal could be achieved along the bottom edges of this boot "seal". When a differential pressure is applied to the boot, the butt seal deflects away from the OWS surface and creates two orifices of "semi-oval" cross section whose size depends upon the applied pressure differential.

A full-scale test was performed at MDAC to determine the leak rate at the bottom edge of the rubber boot. These data indicated that a pressure dependent leakage area of 1.8 square inches would occur under the flight environment existing at 63 seconds.

The above three sources of unplanned leakage resulted in a differential pressure distribution over the auxiliary tunnel significantly different from the design distribution shown in figure 6-11. Post-flight calculations performed at MDAC, MSFC, and LRC indicated that, for the total 6 square inches of leakage area into the aft end of the tunnel, the pressure distribution along the tunnel at Mach 1.0 would be approximately as shown in figure 6-13. These deduced pressures produce large lifting forces on much of the tunnel and part of the adjacent shield areas near the forward edge of the MS.
With this leakage into the aft end of the tunnel, the effect is to lift the forward end of the auxiliary tunnel and the adjacent shield until a critical position is reached where high velocity ram air rushes under the shield and tears it outward from its mountings on the OWS.
Another means of producing bursting pressures under the forward edge of the MS in the region of the auxiliary tunnel is illustrated in figure 6-14. This is an illustrative view of the wave pattern produced by the flared portion of the auxiliary tunnel forward fairing. At low supersonic flight speeds, the high and low pressure regions (the compression and expansion from the flare) extend to considerable distances away from the tunnel itself. - High pressure over the MS forward edge and lower pressures aft tend to lift the overall structure. Lifting due to this mechanism would be indistinguishable from that due to auxiliary tunnel leakage described above.
Because lifting of the auxiliary tunnel as a result of internal pressure is a prime failure mode, both analytical and experimental post-flight studies have been performed to determine the elastic behavior of the MS in the region of the tunnel. Analytical studies were performed by MDAC using a finite element model of an area of the meteoroid shield as shown on figure 6-15. These analyses indicate that the deflection of the auxiliary tunnel and shield away from the OWS tends to become divergent. That is, as an area of the shield becomes exposed to a differential burst pressure, the shield lifts up exposing additional area, which results in further lifting.
Experimental studies were done by MSFC using the air bladder test rig shown on figure 6-16 to generate burst pressures over the forward edge of the auxiliary tunnel area and crush pressures over the rear. The conclusion of both of these efforts is that the MS and auxiliary tunnel are quite limp and easily lifted from the OWS tank into the slip stream. A burst pressure of about 0.5 psi was found to be sufficient to effect this failure mode.
The above postulated onset of failure is judged to be the most probable means by which the MS was lifted into the airstream.

Summary and Recommendations

The preceding analysis and discussion of possible failure modes of the meteoroid shield have identified at least two ways that it could fail in flight. Although the most probable cause of the present failure was the lifting of the shield from the OWS tank by excessive pressures in the auxiliary tunnel, other failure modes could have occurred in other regions of flight or under more severe flight environments than were encountered by Skylab 1. Among these other modes of potential failure, which could combine in various ways under varying conditions of flight, are excessive pressures under the forward edge of the shield, or inadequate venting of the folded ordnance panel. The inherently light spring force of the auxiliary tunnel frames, the crushing loads on these frames in flight, the inherent longitudinal flexibility of the shield assembly, the forces applied by the swing links to deploy the shield, the possible "breathing" of the shield panels as cavities are vented, the non-cylindrical nature of the underlying pressurized tank, and the uncertain tension loads applied to the shield in rigging for flight all contribute to a lack of rigidity of the shield and a weakness of its structural integrity with the underlying tank structure.

A simple and straightforward solution to these inherent problems of the present shield design is therefore not likely. A fundamentally different design concept seems in order. One solution is, of course, to simply omit the meteoroid shield, suitably coat the OWS for thermal control and accept the meteoroid protection afforded by the OWS tank walls. Although the Board has not conducted an analysis, meteoroid flux levels are now know to be considerably lower than those used in the original calculations. A new analysis, based on these flux levels, may show acceptable protection.

Should some additional meteoroid protection be required, the Board is attracted to the concept of a fixed, non-deployable shield. Although the inherent weight advantages of a separable bumper are not available in this approach, the mission of Skylab could probably be satisfied in this manner. One concept would be to bond an additional layer of metal skin to the surface of the tank with a layer of non-venting foam between the OWS tank and the external skin. The problem being statistical in nature, the entire shell of the OWS would not have to be covered.

    Figure 6-1. - View of Kapton surface of the OWS showing forward torsion rod swing link.   

    Figure 6-2. - View of Kapton surface of the OWS showing aft torsion rod swing link and thrust blocks.   

    Figure 6-3. - Auxiliary tunnel frame spring stiffness.   

    Figure 6-4. - Venting locations in meteoroid shield.   

    Figure 6-5. - Ordnance foldout panel.   

    Figure 6-6. - Longitudinal section through meteoroid shield at foldout panel.   

    Figure 6-7. - Skylab (SL-1, SA-513) dynamic pressure profile for boost phase.   

    Figure 6-8. - Meteoroid shield area design differential pressures for smooth configuration (M = 1.0).   

    Figure 6-9. - SL-1 auxiliary tunnel design differential pressures (M = 1.0).   

    Figure 6-10. - Auxiliary tunnel forward vent.   

    Figure 6-11. - Meteoroid shield response - aft auxiliary tunnel boot sealed.   

    Figure 6-12. - Auxiliary tunnel leaks.   

    Figure 6-13. - Meteoroid shield response - aft boot leakage.   

    Figure 6-14. - Compressibility waves from the forward auxiliary tunnel fairing.   

    Figure 6-15. - Mathematical model for meteoroid shield divergence analysis.   

    Figure 6-16. - Air bladder test rig for tunnel deflection test.   




The availability of flight data from the instrumentation on the MS and the vehicle disturbances, the design features of the MS, the SAS, photographs taken in orbit, descriptions by the astronauts. and other information permit the following postulation of the probable sequence of events associated with the MS failure.

Figure 7-1 is furnished to indicate the pitch. roll, and yaw axes and directions of positive rotations. These axes are related to the position designations I, II, III, and IV as indicated on the sketch. A view of the shield and tunnel separating from the butterfly hinge, rotating outwardly, and producing roll torques is shown to provide initial orientation.

In figure 7-2 and 7-3, sketches and details of salient events are correlated to the roll rate data around the 63 second anomaly period. The events are designated on the figures by times which are consistent with the available data.

60. 12 Seconds - MS liftoff and local inflation in the vicinity of the auxiliary tunnel was indicated by a small shift in position of the torsion rod on the forward edge just to the left of the tunnel (see fig. 7-2).

61. 78 Seconds - Air entered the forward facing opening, raised the pressure under the shield and high mass flows escaped through the adjacent holes in the butterfly hinge. This flow produced reactive forces causing a gradual decrease in roll rate between 61. 78 seconds and 62.74 seconds.

62. 74 to 62.79 Seconds - Burst pressure under the auxiliary tunnel and adjacent MS caused a large tangential load on the forward section of the butterfly hinge, causing the whole hinge to unzip.

Fly around inspection indicated that the failure of the butterfly hinge occurred at the hinge line adjacent to the main tunnel.

The butterfly hinge was now completely broken. Aerodynamic drag on the M including the bulky auxiliary tunnel produced tension in the shield and pulled on the vehicle so as to roll it in the direction shown, that is, opposite to that noted earlier. The large area and mass of this metal "flag" induced a more rapid change in roll rate than the earlier jetting through the butterfly hinge. This process terminated as the MS started to wrap around and lift the SAS-2 wing.

62.79 to 62.90 Seconds - During this interval the shield was wrapping around SAS-2 wing producing a negative roll torque in the vehicle. At about 62.85 seconds the SAS-2 tie-downs were broken.

62.90 Seconds - Upon release of SAS-2, the tension in the shield was transferred to the trunnions, causing failure of the trunnion straps. Upon separation of this section of the shield, the negative roll torgue ended.

62.90 to 62.95 Seconds - In this interval, the remaining section of the MS began unwinding, introducing a large positive roll torque.

63.17 Seconds - A large shock was detected by the Instrument Unit (IU) upper mounting ring vibration sensor due to the impact of the separated section of the MS upon the conical adapter between the OWS and the S-II stage.

63.7 Seconds - The MS continued to unwind and whip until 63.7 seconds when it reached SAS-1 wing. As the MS began to wrap around the SAS-1 wing, a negative roll torque resulted. The MS then ripped apart from top to bottom at the longitudinal joint adjacent to SAS-1, pulling a portion of the joint assembly over the SAS-1 wing as the MS section departed. From this point on the vehicle showed normal response to its roll control system. Figure 7-4 is a photograph taken by the astronauts which shows a portion of the MS that remained and how it was attached to the SAS - 1 wing.

    Figure 7-1. - Definition of axes and positive rotations.   

    Figure 7-2. - Possible meteoroid shield motion from 60.12 to 62.74 sec.   

    Figure 7-3. - Sketches of possible shield dynamics during the 63 second anomaly.   

    Figure 7-4. - Photograph from orbit showing longitudinal aluminum angle bent over the SAS-1 wing. (This angle was later cut by astronaut and released the SAS-1 for full deployment. )   




The origin of Skylab in late 1966 -- as an extension of the use of Apollo hardware for experiments in earth orbit imposed an initial environment of limited funding and strong schedule pressures on the program. Skylab, then designated the Apollo Applications Program (AAP), was to "fit in" among the Apollo flights under schedules imposed by the main-line Apollo program. Funding was provided out of the Apollo program and thus the needs of Skylab competed with those of the higher priority Apollo program.

The situation changed in mid 1969 when Skylab became a major line item in its own right and was to use a Saturn-V launch vehicle with a dedicated, dry, OWS. From that point on, increased funding and new flight schedules were established for Skylab. Nonetheless, the original concept of the meteoroid shield was retained when the OWS changed from a Saturn-IB propulsion stage to a dry workshop launched by a Saturn-V. The Board was therefore interested in determining the extent, if any, that either the initial limitation of funds and time, or any subsequent limitations, determined the design or thoroughness of development of the meteoroid shield. This inquiry was limited to the possible effect of funding and schedule of the meteoroid shield as designed and flown on Skylab 1 and did not consider whether meteoroid protection could have or should have been provided in some other way had the program not evolved as it did.

In the Board's review of the evolution of the meteoroid shield from initial design concept. through testing and development. to final assembly for flight, particular attention was devoted to any impacts arising from limitation of funds or time. Extensive discussions were also held with management personnel of MDAC-W, MSFC, JSC and NASA Headquarters on this matter. In no instance could the Board find any evidence that the design or testing of the meteoroid shield was compromised by lack of funds or time. Program personnel, both government and contractor, had full confidence in the basic concept of the MS and thus saw no need to alter the design when the change to a dry, Saturn-V launched OWS occurred. Given the concept that the shield was to be maintained tight to the OWS tank, and thus structurally integrated with the well-established S-IVB structure. the emphasis of testing given to ordnance reliability and shield deployment was considered proper. Neither the records of Skylab nor the memories of key personnel revealed any tests or analyses of the meteoroid shield that were considered desirable at the time and which were precluded by lack of funds or time.




The management system utilized for the Skylab program was derived directly from that which had been developed and used in the Apollo program. As such, it included a series of formal reviews and certifications at progressive points in the program life cycle that are intended to provide visibility to contractor and NASA management on program status, problems and their resolution.. The selected review points and their primary purpose are set forth in Skylab Program Directive No. 11A, which is summarized as follows:

PRR    Preliminary Requirements Review. "To verify by formal review the suitability of the conceptual configuration and to establish the requirements and action necessary to achieve a design baseline."
PDR    Preliminary Design Review. "To verify by formal review the suitability of the baseline design of the Contract End Item."
CDR    Critical Design Review. "To verify by formal review the suitability of the design of a Contract End Item when the design is essentially complete."
CI    Configuration Inspection. "To certify that the configuration for the Contract End Item as being offered for delivery is in conformance with the baseline established at the CDR."
COFW    Certification of Flight Worthiness. "To certify that each flight stage module and experiment is a complete and qualified item of hardware prior to shipment."
DCR    Design Certification Review. "To examine the design of the total mission complex for proof of design and development maturity."
FRR    Flight Readiness Review. "A consolidated re view of the hardware, operational and support elements to assess their readiness to begin the mission."

The primary thrust of these key program milestones is thus a formal review and certification of equipment design or program status; the primary purpose being served is to provide visibility into these matters to senior NASA and contractor program management. As noted in the Skylab Program Directive, the organization and conduct of the review is a major responsibility of a senior program or management official. For each review, specific objectives are to be satisfied, in conformance with preestablished criteria and supported by specified documentation. The reviews are, thus highly structured and formal in nature, with a major emphasis on design details, status of various items and thoroughness of documentation. Several hundred specialists, subsystem engineers and schedule managers are generally in attendance.

The material presented in these reviews is, of course, developed over a period of time in many lower-level reviews and in monthly progress reports dealing with various systems and subsystems. In addition, several other major reviews peculiar to Skylab were conducted, including the following:

Cluster System Review of December 1967

Mathews' Subsystem Review Team of August 1970 - July 1971

Critical Mechanisms Review of March 1971

Systems Operations Compatibility Assessment Review of October 1971 - June 1972

Structural/Mechanical Subsystem Reviews of July 1971 - May 1972

Hardware Integrity Review of March 1973

MSM Center Directors' Program Reviews

There was thus no shortage of reviews. In order to determine the consideration given to the meteoroid shield throughout the program, the Board examined the minutes, presentation material, action items, and close-out of data of each of these reviews and progress reports. In every case, complete records and documentation were available for inspection. In no case did the Board uncover any conflict or inconsistency in the record. All reviews appeared to be in complete conformance to Program Directive 11A and were attended by personnel appropriate to the subject matter under consideration. The system was fully operational.

And yet, a major omission occurred throughout this process - consideration of aerodynamic loads on the meteoroid shield during the launch phase of the mission. Throughout this six.year period of progressive reviews and certifications the principal attention devoted to the meteoroid shield was that of achieving a satisfactory deployment in orbit and containment of the ordnance used to initiate the deployment. As noted in the preceding section on possible failure modes, design attention was also given to the strength of the hinges, trunnion straps and bolts, to the crushing pressures on the frames of the auxiliary tunnel, to flutter and to the venting of both the auxiliary tunnel and the several panels of the shield. But never did the matter of aerodynamic loads on the shield or aeroelastic interactions between the shield and its external pressure environment during launch receive the attention and understanding during the design and review process which in retrospect it deserved.

This omission, serious as it was, is not surprising. From the beginning, a basic design concept and requirement was that the shield be tight to the tank. As clearly stated in much of the early documentation, the meteoroid shield was to be structurally integral with the S-IVB tank - a piece of structure that was well proven in many previous flights. The auxiliary tunnel frames, the controlled torque on the trunnion bolts and the rigging procedure itself were all specifically intended to keep the shield tight against the tank. The question of whether the shield would stay there under the dynamics of flight through the atmosphere was simply not considered in any coordinated manner - at least insofar as the Board could determine by this concentrated investigation.

Possibly contributing to this oversight was the basic view of the meteoroid shield as a piece of structure. Organizationally, responsibility for the meteoroid shield at MDAC-W was established to develop it as one of the several structural subsystems, along with such items as spacecraft structure and penetrations, pressure vessels, scientific airlocks, protective covers and finishes. Neither the government, (MSFC), or the contractor, (MDAC-W), had a full-time subsystem engineer assigned to the meteoroid shield. While it is recognized that one cannot have a full-time engineer on every piece of equipment, it is nonetheless possible that the complex interactions and integration of aero-dynamics, structure, rigging procedures, ordnance, deployment mechanisms, and thermal requirements of the meteoroid shield would have been enhanced by such an arrangement. Clearly, a serious failure of communications among aerodynamics, structures, manufacturing and assembly personnel, and a breakdown of a systems engineering approach to the shield, existed over a considerable period of time. Further, the extensive management review and certification process itself, in its primary purpose of providing visibility of program status to management, did not identify these faults.

Further insight into this treatment of the meteoroid shield as one of several structural subsystems is obtained by a comparison of a listing of the design reviews conducted on both the MS and the SAS. At MDAC-W, the SAS was considered a major subsystem and was placed under the direction of a full-time project engineer. A comparison of the design reviews held jointly between MDAC-W and MSFC on the MS and on the SAS is presented in tables IX-1 and IX-2, respectively. The more concentrated and dedicated treatment received by the SAS is evident in this comparison.

The Board is impressed with the thoroughness, rigor and formalism of the management review system developed by Apollo and used by Skylab. Great discipline is imposed upon everyone by this system and it has served very well. In a large program as geographically dispersed and intrinsically complex as Skylab, such visibility of program status and problems, is a management necessity. We therefore have no wish to alter this management system in any basic manner. But all systems created by man have their potential flaws and inherent hazards. Such inherent flaws and weaknesses must be understood by those who operate the system if it is not to become their master. We therefore wish to identify some of those potential flaws as they have occurred to us in this investigation, not to find fault or to identify a specific cause of this particular flight failure but to use this experience to further strengthen the management processes of large and complex endeavors.

As previously noted the management system developed by NASA for manned space flight places large emphasis on rigor, detail and thoroughness. In hand with this emphasis comes formalism, extensive documentation, and visibility in detail to senior management. While nearly perfect, such a system can submerge the concerned individual and depress the role of the intuitive engineer or analyst. It may not allow full play for the intuitive judgment or past experience of the individual. An emphasis on management systems, can, in itself, serve to separate the people engaged in the program from the real world of hardware. To counteract these potential hazards and flaws, we offer the following suggestions.



Date     Title
May 1967     OWS Preliminary Design Review (PDR)
Feb 1970     MS Internal Design Review QDR)
Feb 1970     MS Workshop Design Review (WDR)
Sep 1970     OWS Critical Design Review (CDR)
Mar 1971     OWS Ordnance & Deployment System Review
Jun 1971     MS Test Review
Jul 1971     Structures, Propulsion & Thermal Control Subsystem Review
Sep 1971     Environmental Control, Pneumatic, Structures & Ordnance Subsystem Review
Oct 1971     Environmental Control, Pneumatic, Structures & Ordnance Subsystem Review
Nov 1971     Pneumatic, Structures, Environmental Control Subsystem Review
Jan 1972     Structures, Ordnance & Pneumatic Subsystem Review
Mar 1972     Structures, Ordnance & Pneumatic S ubsystem Review
May 1972    TACS, Structures, Pneumatic Subsystem Review
Jun 1972    OWS Structural Subsystem DCR
Oct 1972     OWS Design Certification Review (DCR)
Mar 1973     OWS Flight Readiness Review (HIR)
Apr 1973     OWS Flight Readiness. Review (FRR)



Date     Title
Feb, 1970     SAS Internal Design Review (IDR)
Jun 1970     SAS Workshop Design Review (WDR)
Jun 1970     SAS Preliminary Design Review (PDR)
Jan 1971     SAS Critical Design Review (CDR)
Mar 1971     OWS Ordnance & Deployment System Review
Jun 1971     SAS Subsystem Review
Sep 1971     SAS Subsystem Review
Oct 1971     SAS Subsystem Review
Nov 1971     SAS Subsystem Review
Dec 1971     SAS Subsystem Review
Jan 1972     SAS Subsystem Review
Mar 1972     SAS Subsystem Review
Apr 1972     SAS Subsystem Review
May 1972     SAS Subsystem Review
Jun 1972     OWS Structural Subsystem DCR
Oct 1972     OWS Design Certification Review (DCR)
Mar 1973     OWS Hardware Integrity Review (HIR)
Apr 1973     OWS Flight Readiness Review (FRR)




Significant Findings

  1. The launch anomaly that occurred at approximately 63 seconds after lift-off was a failure of the meteoroid shield of the OWS.
  2. The SAS-2 wing tie downs were broken by the action of the meteoroid shield at 63 seconds. Subsequent loss of the SAS-2 wing was caused by retro-rocket plume impingement on the partially deployed wing at 593 seconds.
  3. The failure of the S-II interstage adapter to separate in flight was probably due to damage to the ordnance separation device by falling debris from the meteoroid shield.
  4. The most probable cause of the failure of the meteoroid shield was internal pressurization of its auxiliary tunnel. This internal pressurization acted to force the forward end of the tunnel and meteoroid shield away from the OWS and into the supersonic air stream. The resulting forces tore the meteoroid shield from the OWS.
  5. The pressurization of the auxiliary tunnel resulted from the admission of high pressure air into the tunnel through several openings in the aft end. These openings were: (1) an Imperfect fit of the tunnel with the aft fairing; (2) an open boot seal between the tunnel and the tank surface; and (3) open stringers on the aft skirt under the tunnel.
  6. The venting analysis for the tunnel was predicated on a completely sealed aft end. The openings in the aft end of the tunnel thus resulted from a failure to communicate this critical design feature among aerodynamics, structural design, and manufacturing personnel.
  7. Other marginal aspects of the design of the meteoroid shield which, when taken together, could also result in failure during launch are:
  8. a. The proximity of the MS forward reinforcing angle to the air stream

    b. The existence of gaps between the OWS and the forward ends of the MS

    c. The light spring force of the auxiliary tunnel frames

    d. The aerodynamic crushing loads on the auxiliary tunnel frames in flight

    e. The action of the torsion-bar actuated swing links applying an outward radial force to the MS

    f The inherent longitudinal flexibility of the shield assembly

    g. The non-uniform expansion of the OWS tank when pressurized

    h. The inherent difficulty in rigging for flight and associated uncertain tension loads in the shield.

  9. The failure to recognize many of these marginal design features through six years of analysis, design and test was due, in part, to a presumption that the meteoroid shield would be "tight to the tank" and "structurally integral with the S-IVB tank" as set forth in the design criteria.
  10. Organizationally, the meteoroid shield was treated as a structural subsystem. The absence of a designated "project engineer" for the shield contributed to the lack of effective integration of the various structural, aerodynamic, aeroelastic, test, fabrication, and assembly aspects of the MS system.
  11. 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 MS during launch was not revealed by the extensive review process.
  12. No evidence was found to indicate that the design, development and testing of the meteoroid shield were compromised by limitations of funds or time. The quality of workmanship applied to the MS was adequate for its intended purpose.
  13. Given the basic view. that the meteoroid shield was to be completely in contact with and perform as structurally integral with the S-IVB tank, the testing emphasis m ordnance performance and shield deployment was appropriate.
  14. Engineering and management personnel on Skylab, on the part of both contractor and government, were available from the prior Saturn development and were highly experienced and adequate in number.
  15. The failure to recognize these design deficiencies of the meteoroid shield, as well as 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.

Corrective Actions

  1. If the back-up OWS or a similar spacecraft is to be flown in the future, a possible course of action is to omit the meteorold shield, suitably coat the OWS for thermal control, and accept the meteoroid protection afforded by the OWS tank walls. if, on the other hand, additional protection should be necessary, the Board is attracted to the concept of a, fixed, nondeployable shield.
  2. To reduce the probability of separation failures such as occurred at the S-II interstage Second Separation Plane, both linear shaped charges should be detonated simultaneously from both ends. In addition, all other similar ordnance applications should be reviewed for a similar failure mode.
  3. "Structural" systems that have to move or deploy, or that involve other mechanisms, equipment or components for their operation, should not be considered solely as a piece of structure nor be the exclusive responsibility of a structures organization.
  4. Complex, multi-disciplinary 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.

Observations on the Management System

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.




A - 1



May 22, 1973


Mr. Bruce T. Lundin


Lewis Research Center

21000 Brookpark Road

Cleveland, OH 44135

Subject: Skylab 1 Investigation Board

Dear Bruce:

As you know certain anomalies have occurred during the launch and earth-orbit of Skylab 1, and they have jeopardized the full attainment of the Skylab mission. It is important to NASA that the cause of the anomalies be established and that appropriate preventative measures be taken for future NASA launches.

I am, therefore, establishing a Board to investigate the anomalies and request that you act as Chairman of this Board. The Board is established in accordance with the policy stated in NMI 8621.1A. I have enclosed a NASA Notice which describes the Board's purpose, authority, and responsibilities among other matters. I ask that you recommend to me a slate of candidates for membership on the Board on or before May 29, 1973. I also request that you meet with the Associate Administrator for Manned Space Flight and the Director of the Marshall Space Flight Center within a week after the launch of Skylab 2 to establish a schedule for the Board's activities.









This Notice establishes the Skylab 1 Investigation Board and sets forth its responsibilities and membership.


a. The Skylab 1 Investigation Board is hereby established because it is in the public interest to determine the actual or probable cause(s) of the anomalies which occurred during the launch and initial earth orbits of Skylab 1 and to recommend appropriate preventative measures for future NASA launches.

b. The Chairman of the Board will report to the Administrator and Deputy Administrator.


a. The Board will:

(1) obtain and analyze whatever evidence, facts, and opinions it considers relevant by relying upon reports of studies, findings, recommendations, and other actions by NASA officials and contractors or by conducting inquiries, hearings, tests, and other actions de novo. In so doing, it may take testimony and receive statements from witnesses.

(2) Impound property, equipment and records to the extent that it considers necessary.

(3) Determine the actual or probable cause(s) of the Skylab 1 anomalies.

(4). Develop recommendations for preventive and other appropriate actions.

(5) Provide a final written report to the Administrator.

(6) Carry out any other responsibilities that may be requested by the Administrator or Deputy Administrator.

b. The Chairman will:

(1) Conduct Board activities in accordance with the provisions of this Notice and any other instructions that the Administrator or Deputy Administrator may issue.

(2) Establish and document, to the extent considered necessary, rules and procedures for the organisation and operation of the Board, including any subgroups, and for the format and content of oral or written reports to the Board and by it.

(3) Designate any representatives, consultants, experts, liiaison officers, or other individuals who may be

required to support the activities of the Board and define their duties and responsibilities.

(4) Establish and announce a target date for submitting a final report and keep all NASA officials concerned, informed of the Board's plans, progress, and findings.

(5) Designate another member of the Board to act as Vice-Chairman.


The Chairman, members of the Board, observers and supporting staff are designated in Attachment A.


The Chairman will arrange for the conduct of all meetings and for such records or minutes of meetings an he considers necessary.


a. The Directors of the John F. Kennedy Space Center, the George C. Marshall Space Flight Center, and the Johnson Space Center will arrange for providing office space and other facilities and services that may be requested by the Chairman or his designee.

b. All elements of NASA will cooperate fully with the Board and provide any records, data, and other administrative or technical support and services that may be requested.


The Administrator will terminate the Board when it has fulfilled his requirements.


This Notice is automatically canceled one year from its effective date.



SDL, 1


A - Members and affiliates of the Skylab 1 Investigation Board  (listed on page ii of this report)




AAP    Apollo Applications Program
AM    Airlock Module
ATM    Apollo Telescope Mount
CECO    Center Engine Cutoff
CSM    Command and Service Module
CW    Clockwise
EBW    Exploding Bridge Wire
ECP    Engineering Change Proposal
EMR    Engine Mixture Ratio
g    Gravity
IU    Instrument Unit
IGM    Iterative Guidance Mode
JSC    Johnson Space Center
KSC    Kennedy Space Center
LSC    Linear Shaped Charged
MAX-Q    Maximum Dynamic Pressure
MDA    Multiple Docking Adapter
MDAC    McDonnell Douglas Astronautics Company
MDAC-W    McDonnell Douglas Astronautics Company-West
MS    Meteoroid Shield
ms    millisecond
MSFC    Marshall Space Flight Center
OECO    Outboard Engine Cutoff
OWS    Orbital Workshop
R    Range Time
S-IB    Saturn IB, First Stage
S-IC    Saturn V, First Stage
S-II    Saturn V, Second Stage
S-IVB    Saturn IB, Second Stage
S-V    Saturn V
SA-513    Launch Vehicle for Skylab 1
SAS    Solar Array System
SAS-1    Solar Array System 1
SAS-2    Solar Array System 2
SL-1    Skylab 1
SL-2    Skylab 2
STA    Static Test Article
SWS    Saturn Workshop
TACS    Thruster Attitude Control Subsystem
Vdc    Volts Direct Current
w    watts