ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
applied in both Z- and X-axes. A borescope was used after testing in each axis to qualitatively assess the  magnitude of internal wear and provide a comparison from axis to axis. For the Z- and X-axis testing, the  hardware was exposed to the target levels with no significant (i.e., operation-inhibiting) damage observed  and thus deemed successful.             Figure 5.  Typical Pre- and Post-Qualification MWS Differences (Z-axis [left]; X-axis [right])  The Y-axis was anticipated to be the worst environment for the hardware. MWS and acceptance levels  were completed in the Y-axis. However, during the qualification level test, a distinct change in the sound  and response produced by the hardware was observed 27 seconds into the full level qualification run. A lack of clear visual failure indicators coupled with ill-defined criteria for test stoppage resulted in  continuation of the testing to the full duration (3 minutes). Upon completion of the test, the NESC team found the secondary piston had failed at the aft tooling hole and the lockwire restraining the forward lug from rotating had failed (Figure 6). Due to the failure of the spring strut hardware, no further random  vibration testing could be completed and the hardware was taken back for investigation and performance testing. Note that even though the test hardware incurred a failure, the testing was ultimately considered successful because it flagged design deficiencies before they were encountered late in the test and verification cycle. Figure 6. Structural Failure of Secondary Piston at Aft Tooling Hole [left];   Lockwire Failure at Forward Lug [right] In the performance test, the development spring strut was exercised at a slow rate from the flight  configuration (stowed) to the extended position (Figure 7) to measure the load versus displacement characteristics and observe the function of the anti-back-travel latches. In the wear-in test, the strut was  Pre-Qualification  Post-Qualification  395
exercised 15 times at nominal deployment speeds (394 mm/s (15.5 in/s)). During the test, the strut was  stroked from the flight length to the extended position to reach steady state performance and to observe potential wear or performance degradation. After wear-in, the strut was stored at the flight configuration with the retention tool applied and shipped to the random vibration testing facility.  After failure during the random vibration testing, the strut was returned to the performance testing facility with a truncated secondary piston to remove the failed portion of the strut. Using a cupping interface,  post-vibe performance deployment tests were performed with the strut assembled in the load frame. The  post-vibration performance test served primarily to assess whether there were any large excursions from the nominal force-deflection curve. These excursions could be indicative of significant damage to the  spring strut sliding surface. The post-vibration performance test satisfied the nominal expected end of  travel force of 165 N ± 27 N (37 ± 6 lb f) and resulted in nominal engagement of the anti-back travel  latches. Force versus displacement within the expected tolerance yielded acceptable force margin based  on the next high assembly assumptions and calculations. A comparison of the pre- and post-random  vibration performance test is shown in Figures 8.                  Figure 7. Test Setup for Performance Testing (Pre-Vibe [left] and Post-Vibe [right])

396
Figure 8. Force versus Displacement Plot for Slow Speed Extension Root Cause Investigation and Findings Immediately following the strut failure, the joint NESC and LM team generated a fishbone diagram. Individual elements were classified as aspects of design, analysis, test, and/or workmanship. Related and supporting evidence was gathered to disposition those elements as either most probable cause, contributor, credible, credible but unlikely, or non-credible. The NESC/LM joint investigation occurred over  a month whereupon, due to MPCV Program schedule pressure, LM chose to implement corrective actions addressing identified proximate causes and assumed ownership of the CM-SM umbilical spring strut development test program. The NESC team continued the investigation in an effort to identify root  cause and the most probable failure scenario. Supporting analyses were completed, photographic  evidence was reviewed, and inspections were conducted to provide rationale for fishbone element  classifications shown in Figure 9. Detailed information regarding the root cause investigation, including  classification definition, can be found in the full NESC assessment report [1].  397
Figure 9.  Fishbone Diagram Failure Scenario Supporting Elements The following key findings related to the root cause investigation elements were utilized to devise a most  probable failure scenario. 1.Test Video and Strain Gauge Data : Immediately after application of the qualification levels, the SM  side assembly (aft cap, actuator housing, and forward cap) rotated clockwise (looking SM to umbilical  side). Lagging this rotation, the secondary piston was observed to rotate counterclockwise. Inspection of data from strain gauges mounted on the secondary piston and visual indicators of mounted accelerometers indicated a rotation of approximately 90 degrees. Strain gauge data with indicators of tooling hole orientation relative to the Y-axis applied load can be found in Figure 10. Approximately 27  seconds after application of qualification levels (242 seconds) there was a noticeable decrease in noise and strut dynamic response. No noticeable damage was observed, so the test was continued to the full duration at 395 seconds. 2.Forward Lug Locking Patch Design : Load requirements for the locking patch were not defined prior  to developmental testing. Designers did not anticipate any applied loosening torque to aid in the patch  sizing in part due to the fact that the strut would be in compression during vibration testing.Additionally, it is not standard practice to perform supporting analyses of locking patch capability.Locking patches are not intended to serve as torque reaction features; rather, they should be used to reduce the rate of preload loss in a joint. The prevailing torque requirement was less than the minimum recommended running torques specified for fine thread series threads 38.1 mm (1.500 in) in diameter or less. The locking patch vendor indicated the patch size was small relative to the thread  size and pitch to which they were applied.   3.Joint Characteristics : The threads utilized for testing had an as-designed preload limited to 25  percent of the tensile yield strength (Fty), which was constrained by Material Usage Agreement  (MUA) dictating sustained stresses to be less than 50 percent Fty due to stress corrosion cracking (SCC) susceptibility of the joint material, aluminum 7050-T7451. Preload is the primary means of joint  loosening when subjected to cyclic, transverse loads. Typical torques sufficient to prevent loosening  398
overcome the aft cap-to-actuator housing secondary locking feature with increased capability  (augmented after Z-axis testing). 6.Spring Effects : In addition to bias torques induced by compression of the in-line spring, hand  calculations confirmed the potential for an offset spring mass. Given the free length, compressed  length, spring diameter, and pitch, the titanium spring would have experienced buckling in its compressed state. As a significant mass contributor, this buckling would have further offset the CG from the strut centerline. Evidence of this offset can be found in wear observed internal to the actuator housing from an apparent contact with the spring. 7.End Fitting Rotational Offset : Indications of wear are evident on the end fittings and adjoining  clevises (Figure 11). Contact due to misalignment/rotational offset would have imparted a force offaxis from the strut and induced a torque applied to the end fittings. If the force was sufficient to  overcome resistive elements of the joint, then the end fittings could have tightened or loosened.  Loosening at the forward strut interface would have permitted an applied torque to more easily rotate the secondary piston relative to the forward lug.  

Figure 10. Strain Gauge Da ta and Strut Rotation    Figure 11.  Wear Post-Vibration Testing (Forward Lug [left]; Test Fixture [right]) Failure Scenario The following failure scenario was identified to best correlate with the evidence collected during the root  cause investigation. Due to the varied classifications of elements identified, this failure scenario should  not be considered definitive. Technical limitations for failure scenario supporting elements are provided  and discussed in reference 1. 400
LM performed a residual life assessment on all parts to identify any other components at risk. Those items with insufficient life were replaced. Numerous design changes as shown in Figure 12 were also  incorporated: the secondary piston was redesigned to eliminate the joint between the rod end and the  secondary piston by replacing the assembly with a rod end integral to the secondary piston, the tooling hole interface to the retention tool was replaced with a clamp-style tool, and the aft cap-to-actuator housing interface was adhesively bonded after application of the required preload. The new design was placed through a new random vibration test, this time using an approach that included qualification for AVT per reference 5, and the unit passed the vibration test and post-vibration performance testing. Figure 12. Updated Spring Strut Design Lessons Learned The assessment and the subsequent failure investigation resulted in a number of lessons learned. Those  described above are reflected below in addition to others more fully described in the NESC final report [1]. 1.Care must be taken in adapting heritage designs to new applications.  An actuator design from a  mars lander support leg design was adapted for the spring strut application. As the design evolved  and was scaled to the new application, some of the features became obsolete, but were retained; this  was the case with the forward lug/secondary piston interface. A focused evaluation of the interface would have led to its elimination earlier in the design cycle. 2.Threaded aluminum parts should only be used be u sed in lightly loaded ap plications (e.g., an  application that does not see severe, cyclic bending loads). The threaded interface in the  bending load path of the piston was able to be eliminated in the subsequent design. With regard to  the aft cap, an alternative to the aluminum threaded interface could be designed to include a flange  mounted interface with a pattern of standard fasteners providing the structural connection.   3.Avoid designs that have the potential to utiliz e fastener thread lock ing features to react  applied or induced torque in the higher level assembly. Thread locking features are intended only  to resist self-loosening. However, in this strut design they were used in a way that depended on them  to react any torque that was applied to the entire strut by the environments. Thread locking featuresare not intended or designed to react applied torques that are part of the structural load path and should not be used to do so. 4.Ensure sufficient preloads are obtained (e.g., th rough joint material selection) to reduce the  potential for joint loosening. Due to the choice of aluminum 7050-T7451, the permissible preload of  the joints was reduced to avoid the potential for stress corrosion cracking. Given preload is the  primary means to combat joint loosening, the design was highly susceptible to the failure observed when subjected to a significant random vibration environment. 402
Conclusions The strut, with corrective actions employed, has since successfully completed qualification and  acceptance vibration testing with no failures or anomalies and is slated to fly on the Exploration Flight Test-1 (EFT-1) in September, 2014.  While no single root cause was identified, numerous contributors and credible elements were found that  may have attributed to the failure. Based on these elements, a most probable failure scenario was identified that points to a number of design and process deficiencies that led to a contact condition with  the interfacing test hardware, that, when combined with a center of gravity offset of the mechanism  assembly and significant applied environments, produced a loosening torque at the forward lug-tosecondary piston interface overcoming the resistive torque of the remaining locking elements and a subsequent rotation of the secondary piston. The tooling holes in the secondary piston, nominally located in the neutral bending axis, were then subjected to maximum bending stresses. When the fatigue life at the hole stress concentration was exceeded, the secondary piston failed. Results from this assessment identified key lessons learned for the spring strut design and mechanisms  in general. These key takeaways focus on design as well as process concerns in areas of design, analysis, test, and workmanship. The valued added by the development test program will help satisfy the future needs of the MPCV through increased confidence in the mechanism’s reliability and ability to meet future programmatic constraints.  References 1. Dervan, Jared; Staab, Lucas; Robertson, Brandan; Song, Kyong; Irvine, Tom; Knight, Joseph. /g48/g88/g79/g87/g76/g16 /g51/g88/g85/g83/g82/g86/g72/g3/g38/g85/g72/g90/g3/g57/g72/g75/g76/g70/g79/g72/g3/g11/g48/g51/g38/g57/g12/g3/g38/g85/g72/g90/g3/g48/g82/g71/g88/g79/g72/g3/g11/g38/g48/g12/g3/g177/g3/g54/g72/g85/g89/g76/g70/g72/g3/g48/g82/g71/g88/g79/g72/g3/g11/g54/g48/g12/g3/g56/g80/g69/g76/g79/g76/g70/g68/g79/g3/g54/g83/g85/g76/g81/g74/g3/g54/g87/g85/g88/g87/g3/g55/g72/g86/g87/g76/g81/g74 . Doc. no. 11-00747. November 14, 2013 2./g39/g72/g86/g76/g74/g81/g3/g68/g81/g71/g3/g39/g72/g89/g72/g79/g82/g83/g80/g72/g81/g87/g3/g53/g72/g84/g88/g76/g85/g72/g80/g72/g81/g87/g86/g3/g73/g82/g85/g3/g48/g72/g70/g75/g68/g81/g76/g86/g80/g86 . NASA-STD-5017. June 13, 2006 3./g49/g36/g54/g36/g3/g54/g83/g68/g70/g72/g3/g55/g85/g68/g81/g86/g83/g82/g85/g87/g68/g87/g76/g82/g81/g3/g54/g92/g86/g87/g72/g80/g29/g3/g54/g83/g68/g70/g72/g3/g54/g75/g88/g87/g87/g79/g72/g3/g38/g85/g76/g87/g72/g85/g76/g68/g3/g73/g82/g85/g3/g51/g85/g72/g79/g82/g68/g71/g72/g71/g3/g37/g82/g79/g87 s. NSTS 08307A. July  6, 1998 4./g49/g36/g54/g36/g3/g55/g72/g70/g75/g81/g76/g70/g68/g79/g3/g48/g72/g80/g82/g85/g68/g81/g71/g88/g80/g29/g3/g51/g85/g72/g79/g82/g68/g71/g72/g71/g3/g45/g82/g76/g81/g87/g3/g36/g81/g68/g79/g92/g86/g76/g86/g3/g48/g72/g87/g75/g82/g71/g82/g79/g82/g74/g92/g3/g73/g82/g85/g3/g54/g83/g68/g70/g72/g3/g41/g79/g76/g74/g75/g87/g3/g54/g92/g86/g87/g72/g80/g86 . NASA-TM-106943. December 1995 5. Mechanical Engineering Design, 5th Edition. Joseph Shigley. November 1, 1988 404
design tool. Additional optimization could be included at a later date to allow: tapered members, different  spreader sizes from root to tip, sizing based on specific member loads rather than maximum loads, andsizing based on nominal, as opposed to worst case loads (with safety factors). For instance, a fully  extended manipulator pose is unlikely to be used for most applications; typical operations would have the  links at some angle. In addition, the worst-case torsional load results from a highly unlikely manipulator pose. With a well-defined and actively controlled work envelope, the severity of the load cases could be  reduced. The current sizing for tip stiffness assumes purely passive operation, i.e. the members are sized  so that when a tip load is applied the structure deforms to a set tip displacement. However, with full  antagonistic control, the tip displacement can be driven back to zero by actuating the tendons. This operation would result in a mid-manipulator displacement rather than a tip displacement, but could  significantly reduce the mass of the system if tip stiffness is a primary objective, but at the cost of greater  structural flexibility. Depending on the application and control system this might be an attractive option for  future analysis. Since the tip stiffness requirement also determines the tendon sizing; a reduction in  required passive tip stiffness would reduce the tendon diameters allowing smaller/faster gear trains and motors. This could allow greater actuation speeds that might be necessary to actively control the dynamicresponse of a flexible system. Any of these optimization features could be added in the future if desired.  The TALISMAN sizing tool is primarily used to explore the parameter design space and currently only  calculates the structural mass. Future work will include adding mass estimating tools for joints, motors, gear trains, electronics and any additional hardware needed. A simple, pinned truss analysis tool, based  on matrix structural analysis, was written in MATLAB and used successfully to verify and validate the  results of the spreadsheet output and the analytical equations. To enable more refined analysis of a specific TALISMAN design, an ABAQUS Finite Element Model script was also written to recreate the  geometry from the spreadsheet design tool. The Finite Element Model can then be used for dynamic, frequency and global buckling analyses. The script can create different operational poses, and can be integrated with structural models of candidate spacecraft. Loads can be applied at other points along the  structure if needed and all internal loads and displacements can be interrogated. With these tools the  TALISMAN architecture can be adapted and optimized for a wide range of applications and designs created that have accurate estimates of their structural performance and system mass. 412
Table 1. Comparison of TALISMAN to Shuttle Remote Manipulator System. (Refs. 1, 2) Design Parameter SRMS TALISMAN Total manipulator length 15.3 m (50 ft) 15.3 m (50 ft) Number of joints in manipulator 6 (2 shoulder, 1 elbow, 3 wrist) 5 (2 base, 4 links) Number of links in manipulator 2 4 Tube/Link System Mass [kg] 46 kg (101.4 lbf) 7.03 kg (15.5 lbf) Manipulator Mass 410 kg (904 lbf) 36.1 kg (79.6 lbf) Packaged Volume 1.74 m3(61.4 ft3) 0.23 m3(8 ft3) Table 2. Results from tendon-routing option evaluation 414
Table 3. Specifications for two TALISMAN applications Design Parameter, Specification LR: ISS Payload Application CS: Asteroid Capture  Application Total manipulator length, m (ft) 300.0 (984.3) 20.0 (65.6) Number of links in manipulator 6 4 Link length, m (ft) 50.0 (164) 5.0 (16.4) Number/types of degrees-of-freedom (dof)7 total: 1 base rotation; 6  single dof hinges5 total: 1 base rotation; 4  single dof hinges Number of motor/gear-boxes per hinge joint22 Control method Fully antagonistic Fully antagonistic Hinge joint articulation range, degrees From 0: +125; -180 From 0: +180; -180 Location of motor/gear-boxes Link mid-point Link end-point Link structure 1m by 1m polymeric  composite truss with titanium  joints7.62cm by 15.24cm (3in by  6in) polymeric composite  integral truss Structural assembly concept Assembled robotically on orbit  using Electron-Beam weldingFully assembled & packaged  on ground, deployed in space Spreader height (tip to joint axis), m(ft) 4.5 (14.8) 1.0 (3.3) Passive tension stiffening? Yes: tower height 4.5 m (same  as spreaders)No Tower base location Link mid-point NA Design load: manipulator fully  extended, transverse load at tip, N (lb)100.0 (22.5) 66.7 (15.0) Tip deflection, m (ft) 5.38 (17.7) [1.8% of length] 0.20 (0.66) [1.0% of length] Mass estimate, kg (lb) 2200 (4840) 48.8 (107) Mass per unit length, kg/m (lb/in) 7.33 (0.41) 2.44 (0.14) 415
Figure 1. Planar embodiment of the Invention with Multi ple JointsJoint Tension Elements  to Actuate Joints Tension Elements to Stiffen LinkJoint  SpreaderLink (1 of 3)Pulle y Spring CapstanMotor Joint Axis  Link TowerA C BD I II III IV Figure 2. Variations of Tension Actuated JointMultiple SpreadersLifted off Joint Articulation Terminated at Link End Constrained Lift offSingle Spreader Mounted Motor 1o f  4L i n k  Mounted MotorsJoint Articulation  Terminated Mid Link Single Link  Mounted Motor 416
Deployed Figure 3. Packaging of TALISMAN4.23 m (166.51 inch) 0.353 x.1524 (out of plane) m (13.93 x 6 inch)Packaged Shuttle Remote Manipulator Envelope TALISMAN Packaged  Shuttle Remote Manipulator Composite Tube Diameter Figure 4. Novel Joint Provides 360 Degree Rotation 0 90 180 270 3603 Hinge LinesCB A Spreader  Support F 0.35 SMAN Packaged  417
The Use, Evolution and Lessons Learnt of Deployable Static  Solar Array Mechanisms Mark Ferris*and Andrew Haslehurst* Abstract This paper focuses on the mechanisms incorporated into SSTL's static deployable arrays; namely the  sprung-hinges and hold down and release mechanism (HDRM). Combined, the HDRM and hinges form  the hold down release system (HDRS). The deployable static solar array HDRS has been successfully  used on several missions, first launched upon the DMC-CFESAT spacecraft in 2007 for a U.S. customer(Figure 1), and later used on DMC-UK2 and EXACTVIEW-1 launched in 2009 and 2012, respectively. The simple, robust and low-cost solution HDRS has been evident in allowing missions to satisfy an everincreasing power demand, allowing the solar arrays to increase in size and have a preferable sun angle for increased cell efficiency. The system is now being employed on the first mission out of SSTL’s U.S.  office (SST-US) on the Orbital Test Bed platform. This paper shall cover details of the original design and development program, problems incurred on latter missions, and evolution of the HDRS for the present Orbital Test Bed mission. Both the original development and recent evolutions have taken place in rapid  timescales, to satisfy the high-turnaround of SSTL missions. Figure 1. DMC-CFESAT platform with depl oyable solar panels partially deployed (Ph. Credit LeRoy N. Sanchez94) Introduction The SSTL approach to Spacecraft Engineering Surrey Satellite Technology Ltd (SSTL) is a unique supplier of low-cost satellites to international  customers from its headquarters based near London, UK. SSTL’s success within the industry and to meeting the increasing market need for more cost-effective satellite solutions has meant SSTL expanding  into the U.S. with a dedicated office (SST-US) with design and assembly integration and test (AIT)  facilities. SST-US' first mission, Orbital Test Bed, is now live, with AIT of the spacecraft commencing in  summer 2014. SSTL-UK shall be supplying the modules and platform bus, while SST-US provides the  payloads and conducts the remaining AIT campaign on the entire spacecraft. SSTL’s traditional simple, low-cost platforms used fixed body-mounted solar arrays on all 4 side-panels (non-Earth and Space facing facets) as the satellite’s power source. As SSTL’s mission capabilities grew and payloads became more power hungry, SSTL had to achieve higher-power capacity platforms. Options to do this included higher-performance solar cells; static deployable solar arrays; and tracking deployable arrays. Static  deployable arrays were selected over the other two options for the most popular SSTL-100 class of  *Surrey Satellite Technology Ltd, Surrey, United Kingdom  /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g23/g21/g81/g71/g3/g36/g72/g85/g82/g86/g83/g68/g70/g72/g3/g48/g72/g70/g75/g68/g81/g76/g86/g80/g86/g3/g54/g92/g80/g83/g82/g86/g76/g88/g80/g15/g3/g49/g36/g54/g36/g3/g42/g82/g71/g71/g68/g85 /g71/g3/g54/g83/g68/g70/g72/g3/g41/g79/g76/g74/g75/g87/g3/g38/g72/g81/g87/g72/g85/g15/g3/g48/g68/g92/g3/g20/g23/g16/g20/g25/g15/g3/g21/g19/g20/g23/g3 421