ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
156 Once all 111 fasteners have been disengaged, the Front P anel is disconnected from the STIS instrument  but still directly attached to the FCP tool via the four  studs. Since the FCP is still “soft docked” to the  handrail stanchions, it will not float away. The operator can use both hands to move the tool / Front Panel  assembly away from the worksite.  Tool Design Considerations

Design of the fastener capture technology involved many  tests, prototype iterations, and trade-offs. Many  design considerations involved deciding how to deal with possible anomalies and common fastener problems, such as stripped head recesses. A proble m encountered often on Earth is when, after a  fastener is fully disengaged, its last thread incident ally re-engages with its mating female thread. One  solution implemented was having a so ft, compliant material within the fastener cavity that holds the  fastener in place after it is disenga ged. This is illustrated in Figure 16 in a cross-section view of the  Handrail Removal Tool. This concept can also be utilized if the design needs to allow the fasteners to be re-engaged, because the fastener is held fairly stationary for bit re-engagement. Design details that must be taken into consideration include particle contamination, optimal geometries, and adequate compliant material durometer. 

Figure 16 – Illustration of Fastener C avity Compliant Material Concept 

Accessibility and operator visibility of the worksite are major design considerations. The fastener capture technology works best when the operator can see the fastener heads clearly and closely without  obstruction. This makes it easier to determine if t he screwdriver bit is properly engaged, recognize if a  fastener is turning properly, and realizing the fasten er is all the way out. The geometry necessary for  implementing this does not always align with other essential design features, leading to important tradeoff decisions.  Whether or not the screwdriver bits already exist or still are in design stage must factor into the tool  design. Reach and access evaluations may deem the bit must engage a fastener at a slight angle. Also  material strength analysis may influence the bit’s diameter and other geometries. The Access Holes in the tools must accommodate these types of variables.  In order to ensure the fasteners and washers are proper ly contained, the tool’s geometry is important but  also its position relative to the fasteners and mating su rfaces. For the STIS Fastener  Capture Plate, this is  dependent upon the four anchor points and proper installation. The operator must be given some sort of feedback that the tool is installed properly , meaning sufficient enough to guarantee successful  functionality. This feedback is provided by Depth Indicators on the FCP that tell the operator that the tool is seated flush enough to its mating surface such that  there are no gaps to allow fasteners to escape. 

Clear Window Compliant  Material NASA/CP-2010-216272
157   Figure 17 – Illustration of  Depth Indicators 

Due to tolerance stack-ups and anomalies, it would be impossible to guarantee that the FCP is installed  exactly flush with its mating surface. Therefore, some compliance must be built into the system. This is accomplished with a gasket barrier around the peri meter of the underside of the tool. The compliant  gasket creates a seal at the mating interface, prov iding containment to prevent disengaged fasteners or  washers from getting free.  A potential problem of significant co ncern early on in the design phase was the presence of a Center of  Gravity label on the Front Panel that covered up so me of the fasteners that needed to be removed. The  label is a sticker made of sheet metal and is not simply removable by hand. A lot of testing was employed to come up with a solution. Drilling through the label was evaluated, as well as a mechanism for pealing it off. In the end, included in the STIS FCP is a mechanism for cutting t he label in a manne r that removed  just enough label material to reveal the fasteners it covered up.  Another design requirement of the STIS FCP involved providing the operator a means of prying the Front Panel off with enough force to overcome a sticky gasket that is holding it in place underneath. This was accomplished by including a power tool driven Drive- Off Mechanism that could be actuated if the force  was too much to overcome by hand.  

Tool Development and Evaluation

The fastener capture technology was developed throu gh many tool design reviews as well as evaluation  and testing of prototypes. Many of  these evaluations oc curred at the NASA Neut ral Buoyancy Lab, where  astronauts train and practice  for EVA’s underwater to help simulate ze ro-gravity. Early prototypes of the  tools allowed the operators to make the following deductions: - Removal of the Front Panel and fasteners should be done by an EVA crew member and not an  automated tool.  - Each fastener at the worksite is ac cessible to be removed individually.   - Use of the fastener capture technology of the FCP was preferable to removing and stowing each  fastener individually.  - Disengaging the fasteners would require a clear line-of-sight to each fastener.  - An astronaut in a spacesuit could accomplish remo val of all fasteners wit hout being over-fatigued.   - Operators demonstrated the ability to diagnose and remove a stuck fastener.  - Demonstrated reliable operation of the Center of Gravity Label Cutter.  Depth Indicator Locations  Locations (4 places) Uninstalled Fully InstalledNo visible marking Black line exposed  NASA/CP-2010-216272
158  

Figure 18 – STIS FCP Being Evaluated Underwater at the Neutral Buoyancy Lab  

Figure 19 – ACS FCP Being Evaluated on a Mockup 

On-Orbit Tool Performance   HST SM4 occurred in May of 2009 and was a complete success. The failed electronics boards inside  both STIS and ACS instruments were replaced on-orbit. The fastener capture tools contained all removed  fasteners as designed. Images of the tools on -orbit can be seen in Figures 20, 21, and 22. 

NASA/CP-2010-216272
160   Figure 22 – Handrail Removal Tool Installed on Handrail On-Orbit  

Summary 

The Fastener Capture Plate technology  proved that it could solve the problem of capturing fasteners that  were not originally intended to be disengaged on-orbit in mi crogravity, specifically for NASA’s Hubble  Space Telescope Servicing Mission #4. The technology demonstrated a very time-efficient method for  disengaging fasteners and removing hardware while minimizing the chances of losing parts or generating debris. The four astronaut tools that implement this concept were used successfully on-orbit. There were  many design variables determined for overcoming anomalies, maximizing operator efficiency, and  minimizing space debris. This technology and the res earch put into developing the tools could be utilized  in any future need for on-orbit repairs where space debris and contamination is a concern.  

Acknowledgements  The development of the Fastener Capture Plate technology, related EVA tools, and STIS Repair  engineering would not have been possible without the thoughtful work of Jon Penn, Tim Horner, Matt Ashmore, Torrie Kelly, Hans Raven, Kristoff Paulson, Mike Cortina, Jill McGuire, and Justin Cassidy. Thank you for your help and insight. 

References 

1. NASA Human Spaceflight: Gallery. < http://spaceflight.nasa.gov/gallery >  2. NASA: Hubble Space Telescope. <http://www.nasa.gov/mission_p ages/hubble/main/index.html > 

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162  

Figure 1. CMGs 1-4 Mounted on Z1 Truss 

Space Station Z1/CMG Configuration2 

The four CMGs operate as momentum storage devices that exchange momentum with the ISS through induced gyroscopic torques. These torques are created by a motor-driven constant-speed momentum  flywheel mounted inside two orthogonal gimbals. Both gimbals have torquer motors and position resolvers mounted on the rotational axis and move by use of a gear train system. The momentum  flywheel is mounted inside the inner gimbal and is supported by bearings mounted on each side of the flywheel. This configuration is shown in Figure 2.  Double Gimbal CMGs were selected for the ISS  because of the spherical momentum envelope. The co ntrol law is heuristic with no singularities and  should the active number of CMGs change, the entire momentum envelope varies in proportion to the  initial momentum sphere. 3 

Figure 2.  CMG Gimbal and Flywheel Assemb ly Figure 3.  CMG Spin Bearing Assembly  The flywheel is mounted on two spin bearings, one on the Spin Motor (SM) side, and one on the Hall Resolver (HR) side. The Spin Motor is a low torque motor mounted on the spin axis with a control loop to maintain the commanded wheel-speed. The Hall Resolv er provides digital wheel-speeds for downlinking  via telemetry. The Spin Motor Command Current (S MCC) ranges from 0 to 1.6 Amps and nominally  averages 0.37 Amps. The SMCC varies due to viscosity changes in the spin bearing lubricant caused by thermal variations and high gimbal rates that induce radial loads on the spin bearings. The SMCC is used as an indicator of spin bearing health and status.    NASA/CP-2010-216272
163 An accelerometer is mounted to the inner gimbal of each CMG to monitor flywheel unbalance. The  accelerometer signal is filtered to a bandwidth of 20 to 120 Hz to eliminate frequencies/disturbances not  associated with the flywheel unbalance frequency.  The Spin Bearings, shown in Figure 3, are single angular-contact ball bearings with modified inner and outer races. There are 15 steel balls  mounted in a retainer to maintain consistent ball motion. The spin  bearings are lubricated with KG80, a super refined mineral  oil, in an active oiler system. Pairs of cartridge  heaters mounted near the outer race of each bearing are provided to maintian the lubricant above the  pour point temperature and in the range of 15.6 oC to 26.7oC (60oF to 80oF) during cold orbital conditions. 

Space Station Operational History 

ISS Assembly   Assembly of the ISS began in November, 1998, with  the first Russian component, the Functional Cargo  Block. Since then, the ISS has expanded with the addi tion of many sections, several of which are of  international origin. Major components include Node 1 (a connector module), the Russian Service Module, the Z1 Truss, the US Labor atory Module, a Canadian robotic arm,  the Integrated Truss Structure  with eleven segments, four sets of solar arrays, Node 2, the European Space Agency Columbus module,  and the Japanese Experiment Module. Node 3 will  be added in early 2010. Russian components were  transported via their own launch v ehicles. All other components were carried into orbit using the US  Shuttle Transportation System (STS). The ISS ha s been permanently inhabited by international crews  since October, 2000.  

Figure 4. International Space Station, Present Configuration   To date, the ISS has flown in th ree different orbital attitudes:  • X Body Axis on the Velocity Vector  ( ±XVV),  also referred to as Local Vertical, Local Horizontal  (LVLH). The ISS design is optimized for this attitude for microgravity, altit ude re-boosts, and vehicle  dockings, hence this is the primary flight attit ude. Normally the +XVV attitude is flown, but during  docked Space Shuttle missions, the attitude is chang ed to –XVV to allow for the Space Shuttle to not  be in the forward velocity direction to av oid the potential for micro-meteorite debris.  • X-Principle Axis Perpendicular to the Orbit Plane ( ±X-POP)  was utilized during early Space Station  construction since full solar arra y capability was not achieved until F light 12A. During certain solar  beta angles, ISS would be maneuvered to this attitude to  allow the existing solar arrays to point more  directly to the sun.  • Y Body Axis on the Velocity Vector ( ± YVV)  is a similar but alternate attitude from the ±XVV attitude.  During YVV, the Space Station is ro tated 90 degrees from the typical ±XVV attitude. This attitude was  NASA/CP-2010-216272
167  

Figure 7. Hard Failure and Shutdown GMT 159/2002/12:00-18:00 

CMG3 “Soft” Failure   CMG3 operated continuously at 691 rad/s (6600 rpm)  from its activation on February 12, 2001 until the  unbalance sensor level reached the criteria (0.2 g) for removal from the active steering law on Day 272, September 29, 2006. This occurred as the CMGs we re being pre-positioned for Momentum Management  startup in the XVV flight attitude.  CMG3  was kept out of the steering law pending further analysis.  Several times the unbalance sensor levels reached 0.2 g while the CMG was operational but not used to actively control the ISS. On Day 283, October 10, 2006, the unit was shut down because the unbalance  sensor reached the shut down criteria of 0.4 g. Figure 8 shows the performance signatures and events that preceeded the shut down of CMG3. CMG3 was returned to L-3 Communications Corporation Space  & Navigation (L-3 S&N) on March 11, 2008 for disassembly and a failure investigation.   NASA/CP-2010-216272
168  

Figure 8. CMG3 Time Line to Shutdown 

Forensic Findings   CMG1 Forensic Findings

The failure investigation confirmed that the HR spin bearing of CMG1 suffered a failure after 1.3 years of  operation. This spin bearing was severely damaged and the rings were seperated. Exposure to high  temperature was clearly evident. Verification testing of  the active oiler system flowrate could not be  performed due to the extensive damage from high te mperature. The inner raceway contained a significant  amount of smeared metal. A secondary raceway was generated when some of the balls were driven from  the manufactured raceway and became wedged between the outer race land and inner race land on the load side of the inner race. Two balls were fused together with severe skidding damage at both the inner and outer race contact locations. The remaining 13 balls were of various size and shape. The inner race was seized to the flywheel shaft and required the machining of two axial slots for removal from the flywheel shaft. The outer raceway also contained sm eared metal and was stuck in the bearing housing  steel sleeve. The outer race was seized in the bear ing cartridge steel sleeve and was also machined out  using a similar technique. No damage was done to bore of  the steel sleeve during the outer race removal  process. The retainer was almost completely destro yed, with only two significant pieces remaining,  showing that the cage appeared to split circumferentially down the center line. The two major fragments remaining each had portions of several ball pockets. A smaller segment containing several pads on which the cage is guided by the inner race land was also  recovered. There was no obvious evidence that the  cage had experienced unusual wear either within the ball pockets or at the pad-land interface.    The SM bearing survived the 72-minute deceleration period. Radial and axia l control of the spinning mass  was unconstrained by the failed bearing and the spinni ng flywheel coned about the SM bearing. However,  radial constaint is provided by sequential gaps between the rotating mass and the stationary parts, i.e.,  NASA/CP-2010-216272
169 the stator support labyrinth seal/flywheel shaft air g ap. These features allow for a controlled fail safe  deceleration.     The SM side bearing (non-failed side) was in good c ondition. Adequate lubrication was noted in the ball  race contact zone and at the retainer pad and inner race land interface. The balls were bright with some wear tracks that are considered consistent with op eration time. Except for one uneven wear track on the  outer race, the wear tracks on both races are cons idered nominal for the operational time. A ring of  degraded lubricant was noted on the inner race land, which is abnormal for the accumulated operating time. The retainer was not damaged and did not contain any evidence of unstable operation or distress.  Inspection verified the location of the SM bearing wear tracks were consistent with CMG1’s lifetime load conditions. Additionally, testing of the active oiler syst em verified that the lubricant flow was consistent  with the presence of fresh oil in the bearing.  As the investigation continued, the evidence support ed that a compromised bearing preload led to the  failure. The following findings are evidence that t he spin bearing preload system was compromised:   1. Some “fretting” type marks were observed on the steel sleeve of the SM bearing cartridge steel  sleeve. The fretting marks were 180 degrees apart from the area on inboard section to the area on the outboard section, an indication of bore misali gnemnt. If these conditions occurred before the  failure, either during on earth acceptance testing or during operation, this misalignment would have affected the preload system. In addition, a particula r mark coined a “galled fret” was observed on the  CMG1 HR steel sleeve. This galled fret may have sp ecifically been the one that prevented the outer  race to slide properly which would compromise preload and initiate ball skidding. 

Figure 9. HR Side Bearing Housing Steel Sleeve   2. Some scoring (record grooving) on both the Spin Motor and Hall Resolver side preload nuts has  been observed. This condition inhibited the spring action to react properly during the load and release cycle during axial motion of the flywheel. Preload was reduced to cause ball skid. 

NASA/CP-2010-216272
171 just in its last days of op eration. Additionally, testing of the active  oiler systems verified that the lubricant  flow was consistent with the presence of fresh oil in the both bearings.   Evidence to further support the CMG1 failure conclu sions and provide support that the bearing preload  was reduced to cause the unbalance monitor to reach the “take action criteria” for the CMG3 soft failure  are:  1. Similar “fretting” type marks were observed on the steel sleeve of the bearing cartridge steel  sleeves (Figure 11). The fretting marks were near 180 degrees apart from the area on inboard section to the area on the outboard section, an indication of bearing bore misalignment. If these  conditions occurred before the failure, either during on earth acceptance testing or during  operation, this misalignment would have affected the preload system.   

Figure 11. Typical Cartridge Fretting   NASA/CP-2010-216272
172 2. Some scoring (record grooving) on both the Spin Motor and Hall Resolver side preload nuts has  been observed (Figure 12). This condition inhibited the spring action to react properly during the  load and release cycle during axial motion of the flywheel. Preload was reduced to cause low preload and ball skid. 

Figure 12. Typical Preload Nut Record Grooving 

3. Assembly evaluations also identified a built in  misalignment caused by the inner gimbal covers  distorting the inner gimbal structure and altering the precision bearing bore alignment. This bore misalignment exacerbated the wear observed on the steel sleeves of the Bearing Housings.  Similar distortation characteristics to the CMG1 covers were confirmed on CMG3. 

CMG1 and CMG3 Most Probable Causes 

After a thorough review of the forensic findings and t he on-orbit telemetry the failure investigation teams  concluded the most probable causes for the CMG1 bearing failure and CMG3 excessive unbalance were:   • Fretting/galling at the bearing Outside Diameter (OD)/Steel Sleeve interface inhibited free motion  of bearing preload mechanism, this reduced bearing preload to a threshold level to initiate ball  skid which ultimately caused failure.  • Scoring (record grooving) on the Spin Moto r and Hall Resolver side preload nuts adversely  affected the bearing preload and bore alignment al so contributing to a compromised preload that  initiated ball skid.  • Bearing misalignment introduced by the inner gim bal covers exacerbates the wear conditions at  the preload nut and sleeve interfaces.   • Additionally, research conducted during the CM G Root Cause Investigations showed that the  adverse mechanical conditions described above we re exacerbated by spin bearing radial loads  induced by highest allowable gimbal rates. Oper ational limits were put in place to control both  gimbal angular velocity and acceleration to restri ct the amount of radial loading imparted to the  bearings. 

Design Improvements 

The failure investigation, consisting of test, disa ssembly, evaluation and analysis of CMG1 and CMG3,  identified the wear characteristics at the bearing outer race/bearing housing steel sleeve interface, scoring of the preload and inner gimbal cover induce d gimbal distortion. These conditions adversely lower  the  bearing preload to the point where ball skid can occur. This can cause a bearing failure as in CMG1,  NASA/CP-2010-216272
174 Acknowledgements 

The authors and co-authors of this paper express their gratitude to the failure investigation teams. The  collaborative efforts of NASA, Boeing, NESC, Battelle, The Aerospace Corporation, L-3 S&N, and independent consultants resulted in comprehensive inve stigations that resulted in identifying the most  probable cause for both anomalie s. Accolades we re expressed by Boeing and NASA upper management  many times for the team work, leadership and efficiency of the teams identified below:  

   CMG1 Failure Investigation Team     NASA/CP-2010-216272
175  

CMG3 Failure Investigation Team 

References                                                         1 L-3 S&N CMG Sell Sheet L3SN_24_0603.qxd. “Cleared by Department of Defense/DFOISR for Public  Release under 02-S-0862 on February 27, 2002”   2 Burt, Richard “AAS 03-072 “Failure analysis of International Space Stati on Control Moment Gyro” 26th  Annual AAS Guidance and Control Co nference, Breckenridge, Colorado.    3 Kennel H.F. “Steering Law for Parallel Mounted Doubl e-Gimbaled Control Moment Gyros – Revision A”  NASA TM-82390  dated January 1981 

    NASA/CP-2010-216272
177 International Space Station Powered Bolt Nut Anomaly and   Failure Analysis Summary 

Daniel E. Sievers* and Harry K. Warden* 

Abstract   A key mechanism used in the on-orbit assembly of the International Space Station (ISS) pressurized  elements is the Common Berthing Mechanism. The mechani sm that effects the structural connection of  the Common Berthing Mechanism halves is the Powered Bolt Assembly. There are sixteen Powered Bolt  Assemblies per Common Berthing Mechanism. Th e Common Berthing Mechanism has a bolt which  engages a self aligning Powered Bolt Nut (PBN) on the mating interface (Figure 1). The Powered Bolt Assemblies are preloaded to approximately 84.5 kN ( 19000 lb) prior to pressurization of the CBM. The  PBNs mentioned below, manufactured in 2009, will be used on ISS future missions. An on orbit functional failure of this hardware would be unacceptable and in  some instances catastrophic due to the failure of  modules to mate and seal the atmosphere, risking loss of crew and ISS functions. The manufacturing processes that create the PBNs need to be strictly cont rolled. Functional (torque vs. tension) acceptance  test failures will be the result of processes not being strictly followed. Without the proper knowledge of  thread tolerances, fabricat ion techniques, and dry film lubricant a pplication processes,  PBNs will be, and  have been manufactured improperly.  The knowledge gained from acceptance test failures and the  resolution of those failures, thread fabrication tec hniques and thread dry film lubrication processes can be  applied to many aerospace mechanisms to enhance t heir performance. Test data and manufactured PBN  thread geometry will be discussed for both failed and successfully accepted PBNs. 

Figure 1: ACBM Powered Bolt as mated to CBM Nut. 

                                                     * The Boeing Company, Huntsville, AL 

Proceedings of the 40th Aerospace Mechanisms Symposium, NASA Kennedy Space Center, May 12-14, 2010  Bolt  Encapsulated nut  Powered Bolt Assembly Powered Bolt Nut Assembly  NASA/CP-2010-216272