ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
8.     Herrera-Fierro, P.; Jones, W.R.; and Pepper, S.V.: “Interfacial Chemistry of a   Perfluoropolyether  Lubricant Studied by X-ray Photoelectro n Spectroscopy and Temperature Desorption  Spectroscopy.” J. Vac. Sci. and Technol. A, vol. 11, 1993, p. 354  9.    Kasai, P.H.: “Perfluoropolyethers: Intramolecular Disproportionation,” Macromolecules, 25,    pp. 6791–6799, (1992).  10.   Zabinski, J S; Bultman, JE; Sanders, JH; Hu, JJ: “Multi-environment lubrication    performance and  lubrication mechanism of MoS 2/Sb 2O3/C composite films”: Tribology Letters; August 2006; v.23,  no.2, p.155-163.  11.   Prasad, S; Zabinski, J: “Super Slippery Solids”: Nature;vol 387; June 19th, 1997;p 761-763;  12.   Zabinski, JS; Donley, MS; Walck, SD; Schneider, TR; Mcdevitt, NT: “Effects of dopants on the  chemistry and tribology of sputter-deposited MoS 2 films”: Tribology Transactions; Oct 1995; v.38,  no.4, p. 894-904.  13.    Stamatis, DH, “Failure Mode and Effect Analysis: FMEA from Theory to Execution,” American  Society for Quality, Quality Press, Milwaukee 53203, 2003  14. Fault Tree Handbook, NUREG-0492, U. S. Nuclear Regulatory Commission, January 1981 15. Platit website http://www.platit.com/   16. J Fontaine, 'Ultra-high Vacuum Friction Experiments to Simulate Material Transfer and Wear of the  Platinum Proof Mass Used in the GRADIO Accelerometer'; Proc 10 th ESMATS, ESA SP-524,  September 2003.  144
Lessons Learned Designing a Spherical Satellite Release Mechanism Ryan Hevner* Abstract  A low-cost mechanism, part of the CAPE ICU payload, was designed to contain and deploy two spherical  satellites from the Shuttle in December 2006. Overall the system successfully placed the satellites into  orbit but encountered an anomaly. This flight anomaly and subsequent investigation revealed several key  design issues. This paper discusses the lessons learned from the inaugural mission.  /g3 Figure 1. Ejection of ICU from CAPE aboard STS-116  /g3 /g3 Figure 2. Separation of ICU, Releasing the Satellites  /g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3/g3 /g3 * Planetary Systems Corporation, Silver Spring, MD  Proceedings of the 39th Aerospace Mechanisms Symposium, NASA Marshall Space Flight Center, May 7-9, 2008  145
The Designed Separation Event  The ICU was designed to separate from the CAPE by means of actuating the NEA operated CSS  Lightband. 32 compression springs on the CSS eject the ICU at a velocity of 0.52 m/s from the Shuttle.  After 40 seconds, the timer in the Avionics Deck sends  a signal to simultaneo usly separate ISS #1 and  ISS #2, releasing the satellites from their respective cylinder (Figure 7).  /g3 Figure 7. Anticipated ICU Separation Event  Flight Anomaly  By all accounts the ICU separated nominally from the CAPE. Video taken from the Shuttle shows the ICU  ejecting smoothly with no observable rotation. Velocity appeared normal. See Figure 1.  The ISS separation event had an anomaly. The satellites did not simultaneously eject from their  respective cylinders as expected. FCal separated from its cylinder. Its ejection rate was as expected  except when the equator reached the end of the cylinder it temporarily “stopped” for a few seconds and  then slowly continued out of the cylinder. The MAA did not come out of its cylinder. Instead of all bodies ejecting simultaneously away from the central Avionics Deck, the Avionics Deck “stuck” to the cylinder for  a few seconds before slowly tumbling away. The Avionics Deck showed a noticeable rotation rate of 1015 deg/sec while the MAA Cylinder had no ob servable rotation.  MAA was not seen to eject for the entire  20 minutes that Shuttle video tracked the MAA Cylinder. Ground operations confirmed that MAA did  finally eject approximately 1 hour after ISS separation (Figure 8).  /g3 Figure 8. Actual On-Orbit ICU Separation Event  149
were fine tuned to match the expected flight masses. The test fixture itself worked well, allowing for  verification and measurement of the separation velocities and satellite rotation rates. The spin-up snubber  was present in the fixture; however the Avionics Deck snubbers were not present due to an oversight in  the fixture design. Also, a simple means of implementing compressed o-rings could not be determined  and thus they were also left out of the test.  /g3 Figure 9. ICU 2D Separation Test Fixture The 2-D separation test revealed that the separation velocities were too high. An imposed design  requirement was that all components must separate from the Shuttle with a net velocity of 0.3 m/s.  Energy had to be removed from the separation event to reduce the velocity of the MAA Cylinder. The  required amount of energy removal was further impacted by the earlier decision to hastily increase the mass of the ICU Lid, effectively reducing the ICU separation velocity from the shuttle. It was deemed  essential that the initial separating force from the spring s remain the same to ensure proper operation of  the Lightbands and overcome any possible stiction in the o-rings. It was thus decided to reduce the travel  of all the compression springs from 20 mm to 5 mm. The final configuration consisted of 4 roller springs  per satellite and 4 regular springs per ISS. The fully compressed force of each spring was 88 N.  The springs were modified and the test was re-run. Separation rates were within limits for all scenarios:  simultaneous ISS separation, ISS #1 first and ISS #2 first. However, the test revealed a risk of the  Avionics Deck re-contacting a satellite during staggered separations. It was deemed that tip-off of the Avionics Deck could not be controlled well enough to guarantee clearance. Thus it was decided that a  simultaneous ISS separation was the best scenario for achieving proper separation of all 5 components.  151
/g3 Figure 10. Ratio of Imparted Kinetic Energy to Stored Compression Energy in a Viton O-Ring  Several trials at various compressions were performed at room temperature, while only a few trials were  performed hot and cold. This explains the scatter at  room temperature and very little scatter elsewhere.  However, from the data it is apparent that the energy efficiency of the o-ring increases with temperature.  It was actually negative for some tests. The assumption is stiction between the Viton and aluminum rings decreased velocity. The test set-up and fixture limited the minimum achievable o-ring temperature to 0°C.  It is still unknown how the o-ring would react at -25°C. If it follows the same trend however, it is likely that  the preloaded o-rings had little effect on the separation event.  The other question that remained was the exact preload on the satellites at separation. Strain gages were  bonded to the inside  cylinder walls, transforming them into load cells. To keep them away from the  satellites the strain gages were placed approximately 13 cm from the ends of the cylinders. These were  calibrated by loading mock-up plates that simulated the interface of the satellites to the pedestals.  Schedule and budget restraints prevented testing to correlate preload with respect to temperature. It also  prevented the monitoring of preload over time to correlate creep in the o-rings and create a model to  predict preload vs. time. The only data available was during a trial satellite installation where the satellites  remained in the ICU for a few hours (Figure 11).  153
/g3 Figure 11. Decrease in Satellite Preload as a Result of O-Ring Creep  This short timeframe makes it difficult to predict whether the preload would continue dropping or  eventually approach some asymptote.    The indicated preload readings were also subject to some  skepticism due to the extremely low strain in  the cylinders. Given the cylinders were 490 mm in diameter and 2.9-mm thick, the strain was on the order  of 2E-5 to 4E-5. The indicator could measure this strain. The problem arose in bolting components  together. Small variations in flatness often created localized stress increases that significantly changed  the indicated preload. For instance, simply bolting the ICU Lid to the FCal Cylinder changed the indicated  reading on FCal by 17,000 N. Because of this, care was taken to calibrate the load cells with everything  fastened as in flight. However, the addition of preload shims could easily change the local flatness. With it being that sensitive, the preload on the satellites has a large tolerance. An improvement would be to put  the strain gages on the outside of the cylinders so that they can be placed in the middle, as far away as  possible from any localized interface stresses. Also, the flatness of all adjoining planes should be held as  tight as reasonably possible.  Consideration was given to using something stiffer than Viton to contact and preload the satellites that  would be non-marring. A Delrin strip was considered however this  was too stiff. Since the shims can only  be manufactured to 0.25-mm thick, replacing one shim would result in a preload change of several  thousand newtons. The cylinder walls cannot be made much thinner without affecting the overall stiffness  of the ICU. An o-ring appears to be the best option. A consideration should be made to investigate the  performance of either low-temperature Viton or Butyl rubber (IIR) for this application given.  154
The 2D separation test proved to be of limited ben efit in predicting the flight event. The limitations of the  fixture produced results that cannot be assumed realistic. Developing a detailed simulation model in  software is a better allocation of resources.  A result of this mission and subsequent investigation is that it is crucial to test a mechanism throughout  the extremes of the flight environment. Assuming the system would perform at flight extremes as it did  under room temperature conditions was a significant ov ersight. The ability to predict the satellite preload  and understand the effects of a compressed elastomer on the separation event at various temperatures is essential to create accurate simulation models.  Lessons Learned  The following is a summary of the lessons learned during the design, test, integration and flight of the  ICU. These lessons will be applied to the follow-on ANDE mission. /g120 Stagger the satellite ejections. This reduces the degrees of freedom and interaction possibilities  by decreasing the number of sepa rating bodies from 5 to 3 for each event. It is also impossible to  get a truly simultaneous separation of both Lightbands. /g120 Tighten the flatness and alignment tolerances on all mating surfaces. This will improve the  accuracy of satellite preload determination. Flatness of 0.025mm may be necessary. /g120 Remove the penetrating conical snubbers from the satellites. They needlessly complicate the  separation event. /g120 Perform tests to confidently predict the preload creep on the satellites over several months. /g120 Ensure the material properties, specifically stored energy, of an elastomeric o-ring are fully  understood across the entire temperature range. /g120 Increase the velocity of the ICU from the CAPE. This increases the energy available to impart on  the satellite ejection event. /g120 Make both satellites out of aluminum to minimize any thermal induced preload changes. /g120 Reduce the risk of snagging by eliminating lock wire near the path of satellites. /g120 Painting scheme and alignment marks are valuable when viewing and analyzing on-orbit video. References  1. Eby, Matthew A. / The Aerospace Corporation. “ICU Dynamic Simulation.” 22 April 2007. 2. Ritterhouse, Scott R. and Johnnie P. Engelha rdt. Payload Deployment System with an Internal Cargo  Unit. U.S. Patent 6,776,375. August 17 th 2004.  3. Holemans, Walter.  Reusable, Separable, Structural Connector Assembly. U.S. Patent 6,227,493.  May 8th 2001. U.S. Patent 6,343,770. February 5th 2002. U.S. Patent 6,390,416. May 21st 2002.  156
Wear Life Testing of a Mission Critical Separation Interface  Jonathan P. Wood *and Joseph de la Fuente* Abstract This paper covers the wear life validation testing performed for a mission critical separation interface. The  test was performed for the purpose of risk reduction, to demonstrate wear life margin prior to performing official qualification testing. The following is an overview of the mechanism, the design rationale of the test equipment and method, the test results, and the lessons learned. In addition to the specifics of the  test, the implications of the test data on the design of launch lock mechanisms are discussed.  Introduction The LLA (Launch Lock Assembly) is the mechanism used to restrain and deploy a large Payload Wing  during its launch and mission sequence. The two halves of the LLA are mated by a preloading element in order to stow the payload wing for launch, as shown in Figure 1. These two halves separate by releasing  the preload. This necessitates a separable joint within the LLA, capable of withstanding axial and shear  loading. This joint, referred to as a shear tie, is subjected to significant wear under loads during launch. After launch, the joint must separate, in order for the payload wing to function. As such, the LLA shear ties represent a mission critical interface. Due to the large mass and relatively small surface area of the payload wing, the bulk of wear that occurs  during launch is due to the coupled loads between the launch vehicle and spacecraft. This creates a  relatively large displacement rocking motion between the male and female shear ties at low frequencies  (~20 Hz). This shear tie geometry and wear mode do not fit comfortably within heritage shear tie designs, so a unique test was devised to simulate the expected wear that would occur during ground test and  launch. Figure 1. LLA cross-section showing shear tie interface  * Lockheed Martin Space Systems Company, Sunnyvale CA Proceedings of the 39th Aerospace Mechanisms Symposium, NASA Marshall Space Flight Center, May 7-9, 2008 Separation Plane X YZ Male and Female Shear Tie Pair 157
acoustic energy), a random-vibration test was performed that re-created the primary rocking wear mode.  Results from this test, (with respect to the cumulated fretting distance) are shown in Table 2. Coating B consistently outperformed the Tiodize coating, which showed much higher wear levels, and localized freckles of galling. Figure 3 shows a photograph of a galled shear tie test specimen.  The discrepancy between test results was surprising because the material selection test had applied the  same contact stress level to the parts as the random-vibration test, and the maximum fretting distance  applied during each cycle was similar between the two tests (0.6 mm for the material selection test,  0.7 mm for the random-vibration test). The fretting distance at which galling occurred was also different between the two tests (7000 mm for the Tiodize coating in the material selection test, ~5000 mm for the  random-vibration test.) The coatings from both tests were cross-sectioned, and parameters such as  coating thickness and coating consistency were in-family with both batches of parts. Figure 2. Schematic showing material selection test configuration  Loading Male Test Specimen  Female Test Specimen  159
Table 2. Random-vibration wear test results  Coating Relative Wear Coating B Low - Moderate Coating B Low Coating B Low - Moderate Coating B Low Tiodize/Tiolube 460 Moderate-High Tiodize/Tiolube 460 Moderate Tiodize/Tiolube 460 Moderate Tiodize/Tiolube 460 High - Galling Tiodize/Tiolube 460 Moderate-High Tiodize/Tiolube 460 Moderate Tiodize/Tiolube 460 High - Galling Tiodize/Tiolube 460 Moderate By looking at the differences between the two tests, the following conclusions were drawn. First, it is  important to re-create the same wear mode during a wear test – simply applying the same bulk fretting distance under the same contact stress may not provide the same result. Second, the actual geometry of the mating parts (not just the contact stress), seems to influence test results. The shear tie specimens from the material selection test had a larger overall diameter, larger spherical radius, and shallower contact angle than the random-vibration test specimens. (The geometry for the material selection test coupons was based on a new shear tie design). Figure 3. Photograph showing galled test specimen (unacceptable wear pattern)  160
tests were performed at a drive frequency of 0.1 Hz. Two cycles were performed for each test run. During  cycling, the preload in the preload rod was monitored, and the peak preload value was recorded. The autocollimator measurements were also monitored. For the autocollimator reading, some drift was acceptable, provided it was not cyclic in nature.  Clevis Load Cell Turnbuckle ClevisShear Tie Center of  Rotation Figure 4. Schematic of test setup  Phase 2 Test Phase 2 tests were performed using the same basic procedure as phase 1, but at higher frequencies,  and with cycle counts corresponding to Table 3. The 1 /g86 test run was performed first, followed by the 2 /g86 run and the 3 /g86 run. During pre-test trials, it had been found that the displacement limits were tripped if the  drive frequency was increased too suddenly. For this reason, each test run was started at 0.5 Hz and  ramped up to the final drive frequency. During this test, the autocollimator was monitored to verify that  uncontrolled displacements were not inputting a significant level of additional wear. UUT Instron Ram  DirectionExtensometer  Measurement Points Mirror Flat for  Autocollimator Y/Z X 164
Figure 5. Photograph of test setup  Phase 3 Test After each specimen pair was subjected to all 1 /g86, 2/g86, and 3/g86 test runs, a separation test was performed.  The separation test fixture is shown in Figure 6. A cradle, which rides along a linear bearing, was used to support the bolt catcher bracket and clevis mount during separation. The function of this cradle is to limit rotation about the axis during the release of the preload. Unrestrained motion during this step may break any cold-welds that may have occurred, which could prevent the detection of shear tie failure. During each separation test, the cradle was first installed on the base plate. Then, the ball-tipped screws  were adjusted to contact the LLA. Next, the preload nut was unscrewed slowly, while monitoring preload.  If a popping noise was heard, (which indicates the breaking of a cold-weld), the load when this occurred  was recorded. Once the preload was removed, the shear ties were inspected and photographed.  Test Results  Anomalies and Failures Some anomalies did occur during testing. The first anomaly was caused by excessive slop in the linkage  between the Instron machine load cell and the bolt catcher bracket. The excessive slop was due to the clearance in the pinned joint between the test fixture turnbuckle and the Instron load cell. This affected the first set of specimens during the 1-sigma wear test by reducing the overall amplitude of wear cycles  throughout the test. This problem was fixed before proceeding to the 2-sigma wear test by replacing the “standard” pin block on the load cell with a pin block that could be preloaded with a jam large jam nut.  While testing the first set of specimens, it was also observed that the extensometers have a tendency to  slip on their mounting posts during the 5 Hz tests. Grooves were filed into the extensometer mounting  posts to mitigate this problem, but it continued to occur during testing of the second and third pairs of  specimens during 5 Hz tests. This anomaly is considered to be of minimal impact, because the low frequency characterization tests provide accurate data to determine the actual rotation amplitude that  165
occurred during each test. Issues with the Instron controller caused two to three extra cycles to be  applied to the second and third specimen sets. Restrain LLA by engaging with ball tipped screws  Unscrew Preload Nut to Remove Preload  Figure 6. Release of preload  Results Summary No galling or failure to self-separate occurred during testing. Photographs of a representative shear tie  pair are shown in Figure 7 & Figure 8. The level of shear tie wear was acceptable. The wear on the shear tie specimens is very directional, as expected, and shows a layer of burnished-in MoS 2 (Tiolube 460) that  is just beginning to flake away from the anodize layer. Inspection under a microscope shows that although the Tiolube layer is very thin, it is mostly intact. Comparison of this level of wear to the sequential wear patterns observed during life testing of this coating indicate that the coating has most of  its wear life remaining. 166
A representative set of load/deflection curves is shown in Figures 9 -11. It is interesting to note that each  curve shows a distinct Dahl slope, which represents a portion of rotation that is characterized by elastic deformation of the contact surface (due to shear stress) rather than sliding. The first 0.3 to 0.5 degree of displacement in each cycle was due to elastic shear. Once the torque to overcome friction is reached, the joint begins to slip, and the slope of the curve levels out. 1-Sigma Test SN 008 -1200-700-200300800 -1.5 -1 -0.5 0 0.5 1 1.5 Displacement (mm)Load (N) Outboard Extensometer Inboard Extensometer Figure 9.  1-sigma load/deflection curve  2-Sigma Test SN 008 -1500-1000-500050010001500 -3 -2 -1 0 1 2 3 Displacement (mm)Load (N) Outboad Extensometer Inboard Extensometer Figure 10. 2-sigma load/deflection curve  168
3-Sigma Test SN 008 -1500-1000-500050010001500 -5 -4 -3 -2 -1 0 1 2 3 4 5 Displacement (in)Load (lb) Outboard Extensometer Inboard Extensometer Figure 11. 3-sigma load/deflection curve The rocking angle that occurred during each test was calculated, and is compared to the desired value in  Table 4. Note that the 1-sigma test level for serial number 007 was about half the desired test level, as discussed in the anomalies and failures section. The friction coefficient was calculated from the load at which slip occurred during each test. The friction coefficient ranged from 0.03 to 0.08.  Table 4. Actual displacements during testing  Measured Desired Difference Measured Desired Difference 006 0.049 0.059 17% 0.48 0.59 19%007 0.029 0.059 51% 0.29 0.59 51% 008 0.049 0.059 17% 0.47 0.59 20%006 0.104 0.110 5% 1.05 1.18 11%007 0.104 0.110 5% 1.08 1.18 8%008 0.105 0.110 5% 1.03 1.18 13% 006 0.157 0.165 5% 1.60 1.77 10% 007 0.160 0.165 3% 1.83 1.77 -3% 008 0.160 0.165 3% 1.60 1.77 10%2-Sigma 3-SigmaTest Serial NoAngle Displacement 1-Sigma 169