ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
explained by a forced rotation of an outer bearing ring of a bearing pair as a consequence of a transient  jamming force produced by the cage. After rotation of the bearing ring to its new position, the cage stability  might have been improved. This theory would also comply to the fact that for a stiffer cage the noise issue got worse (higher forces produced by the cage).  The key question was now to find the origin and influencing parameters for the tr ansient forces obviously  produced by the cage with the consequence of generating the observed bearing noise. The only possible  explanation had to be searched in the cage configuration itself.   A special tool was manufactured to allow inspection of the cage circularity and the EQM cage dimensions  were found to be within the “as designed” tolerances. Also the pocket dimensions and clearance to the  inner race was in tolerance (chosen such that at cold operation the shrinking cage could not jam on the inner race due to different thermal expansion between race and cage).  The conclusion drawn from all the previous investigations was that the bearing cages might be deflected  or pushed out of their nominal position from time to time thus generating high local friction forces and  subsequent bearing noi se. After relaxation of the individual force peak, and nominal stab le cage rotation,  the noise was not observed for a certain time. This theory was supported by the fact that the noise issue  could be deleted for a certain time after bearing re-lubrication leading to lower friction and consequently  also to lower built up cage forces.   As a consequence, a method had to be identified allo wing to avoid the genera tion of stresses between  bearing cage and balls or to damp the dynamic motion of the cage in a way suitable to avoid noise generation. As a result a quite un-conventional idea was realized. Three bearing balls of each bearing  were dismounted from the positions 12 0 deg apart, i.e. three of the 39 cage pockets were left without a  ball. This measure was performed after finding a positive Hertzian stress margin for a 36-ball configuration.  The effect was significant. The noise issue was not observed any longer and even during temperature test  at qualification temperatures the bearing run smoothly and completely nominal. Also during a subsequent  life test lasting for 4 months at nominal speed (corresponding to 100 million bearing revs), no further noise  issue was observed. Though the effect was not completely understood, it could be concluded that the bearing issue could be  solved by changing the dynamic behavior of the bearing cage. This was achieved by taking out dedicated balls from three positions of the cage pockets. Figure 10 illustrates schematically the measures taken. The baseline Cage Configuration using 39 Ball pockets in an un-symmetric configuration with uneven pocket  spacing was maintained, however 3 balls were taken out of this configuration (every 13th ball) to come to the final 36-ball configuration.  3 Ca ge Pockets left free Figure 10: Bearing cage configuration and final bearing configuration using 36 balls only  127
bearings are definitively required for the application. The lead times for standard off the shelf bearings are  often very short compared to specifically designed space bearings.  Even if a bearing supplier offers dedicated space bearings, the main production process will normally  focus towards manufacturing for industrial or military application. Therefore the support in case of bearing  issues observed after bearing delivery is often limited and the implementation of any necessary changes might cause significant implications within the supplier standard manufacturing process.   In addition, the mechanism developer should be able to dismantle and re-integrate the procured bearings  on site for later inspection, lubric ation, pre-load or cage modification ac cording to in-house procedures in  order to mitigate the remaining technical and schedule risk.  By following the process as described below, extreme advantages were already gained in follow on  projects w.r.t. schedule, technical risk and costs: The procured standard bearings are dismantled after  delivery by the mechanism developer. The cages are designed according to the needs and produced from the material best suited for the application. The bearings are then adjusted to the requested bearing  preload by grinding the race shoulders to the required dimension and lubrication is performed according to  the needs. Then the bearings are integrated on site and inspection as well as early life test under representative worst case conditions is performed.  High speed bearings should be tested in any case in a representative bearing test set up at worst case  operational conditions before implementing it into the space  mechanism design. This becomes possible  within the normal (tight) development time frame, if the procurement approach as described above is realized.  130
The resulting surface coating evaluations produced a robust solution but do not serve well as a strict  scientific study. Due to schedule constraints, multiple variables were not controlled and recorded as  thoroughly as desired. As a result, coatings which failed the testing shown here should not be discarded  for other applications because conformity to important factors like surface preparation before coating  application were not sufficiently controlled. 

Previous Design Background  The investigation presented here started with the final qualification of a prior, virtually complete design.  Upon receipt of three flight units, several problems immediately came up that were due to mechanism  interference and jamming. After these problems were resolved and parts were prepared for testing, a  question arose regarding performance under required preloading. A light pre-load was applied to the  latching mechanism and it would not release. The mechanism was required to operate from a high output  paraffin actuator (HOP) producing less than 222N of  activation force. The de sign was altered to  accommodate slim-line piezo-electric load cells to determine how much activation force was being applied by the HOP. It was found that  even with a slight mechanism pre-l oad of much less than the  required 4448N, over 1779N of activation force was required to actuate it. Upon actuation, the unit  exerted an audible shock. It became clear that the problem was that the pre-load was transmitted  perpendicularly into the mechanism such that a roller had to be rammed up a ramp (Figure 1). A solution  to balance the undesired moment exerted by the pre-load and a sloped ramp to make pre-load  application tend to open the mechanism were implemented and successful releases were achieved. This design change, however, now had a tendency of making the mechanism pop open spontaneously under  static loading. This new problem resulted in the discardin g of the prior design due the possible critical  mission failure that could result in an accidental LL release.   Figure 1. Original LL design problem  Design and development of the original unit did not focus on understanding failure modes. Requirements  were considered but a positive method of meeting the requirement was fo rmulated and was not critically  analyzed from a mechanism functionality perspective. For example, the accepted demonstration of the  device being 0.999 reliable was that the HOP’s had a manufacturer’s reliability of 0.97. Since the release  was redundant with separate actuators, the reliability was therefore 1 – (1-0.97)*(1-0.97) = 0.9991. No analysis into the failure modes of the actual mechanism was done. Even the use of the manufacturer’s  0.97 reliability number was applied incorrectly since the HOP was allowed to misalig n by more than the  operational requirement of 3 degrees due to the mechanism’s clearances. With the original design’s failure, the second design was conducted with a large amount of oversight developed to make sure the  next design would work.    132
derived requirements can serve to guide designers back to how they coul d cause failure. This centralized  use of a fault tree was utilized in the LL re-design. The idea was originally driven by oversight requesting  a strong analytical evaluation of the reliability of the device but the resulting fault tree analysis was found  to be very beneficial to assist in orienting one’s mind to eliminate critical failure modes. This process easily incorporates standard  activities such as stress m argin assessments a nd test plan creation. The  specifics of general fault tree analysis are not discussed in this paper. Some good resources on this topic  are contained in the references [13, 14].  Launch Lock Design and Qualification Process

The design functionality of the re-designed LL device is illustrated in Figure 2. It constrains a gimbal at a  single location. The design has a spherical interface that allows the main housing to slide in a plane in order to meet an unknown tolerance stack-up without having to do any match machining. This requires  lapping of the spherical parts to get a smooth fit. Release is accomplished via a low shock non-explosive  actuator (NEA) 9131 [2].       The fault tree was created early in the design process based on ‘brainstorming’ sessions within the  design team. Fifty-one mechanical events were identified, which were sources of either complete or  partial mission failure (Figure 3). Fault tree events were then noted in drawings for critical dimensions and testing activities so that each part can be referenced to the fault tree (Figure 4). The fault tree work also  lead to the development of mechanism-specific testing that went beyond a standard qualification process.    Figure 2. LL operation 

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Figure 3. LL fault tree with 51 mechanical events  Stiffness and Pre-load   Since the LL constrains the gimbaled system at a single location, the overall stiffness of the system has a  strong dependence on mechanism stiffness. Assessment of stiffness and pre-load are the refore critical to  device performance. In the previous design, this was recognized but compliance with requirements was  not correctly verified. The original design stiffness was quantified through a finite element analysis (FEA)  that assumed that entire interfaces were rigidly joined. This was discovered to be a poor assumption for  the pre-load and joint to length ratios present for the design. The stiffness was placed into the system  level model as six springs with rigid connections between multiple interface points (Figure 5).   

In the re-design effort, careful consideration of failure modes lead this incorrect model to be replaced with  a mandate to experimentally measure the stiffne ss for the new design and to then use this data to tune a  geometrically accurate FEA model via three springs at the spherical interface. This first order assessment  of dynamic performance was considered sufficient since the 1st mode of the LL is much higher than the  system level modes of interest. A test setup that allowe d general force applic ation to the separation  element was designed (Figure 6). Monitoring of displacements was accomplished via LVDT’s.    Figure 4. Drawing package connection to failure events   Stiffness was evaluated experimentally as a 3X3 matrix (translations only with rotation s free) about the  center of rotation of the LL spherical joint. This experimental assessment was compared to the same  135
entity calculated from modeling results. The resulting stiffness correlation achieved accuracy of 4 and 2%  in the two directions that were critical to system level dynamics. The system level dynamics were  considerably different after this process had been completed. The first mode had dropped by 6% and  considerable changes in mode shapes and position had occurred. Load levels changed by up to 20% as a result of the changes.    Figure 5. Original modeling (L) New modeling with experimental verification (R)  It is clear from these results that good modeling practice for single point constraint mechanisms should  include detailed modeling as opposed to the six spring approach that was originally used. Use of a  geometrically correct finite element analysis model, which has been verified experimentally, minimizes the  possibility of oversimplified or incorrect modeling. In cases where the coupling is not accurately modeled  as first order, even more advanced techniques such as frequency response function techniques could be  implemented to evaluate joint characteristics [3-5]. These techniques could clearly benefit future system level modeling efforts but would require extensive development and validation to be used.   Figure 6. Stiffness setup and required improvements   The fault tree analysis reminded the design team of the importance of pre-load. The pre-load of the LL  affects stiffness considerably. Use of a strain gauge rather than relying on torque was therefore required. Prior satellite hardware programs had used strain gauges and found that on occasions, the gauges  sometimes did not work correctly when pre-load was applied. This lead to the need to be able to verify  136
designed such that root mean square force of a mass attached to the top is equivalent to that of the  acceleration spectral density (ASD) required. The similarity of the test setup to the flight unit with respect  to joint slipping was limited due to scaling effects. Later evaluations of the fixture indicated that RV levels  at the top of the fixture were actually much higher tha n predicted for the flight interface and that, from this  perspective, the test set up provides a conservative test case. In addition, testing was conducted at  qualification levels +6dB above the required ASD on three orthogonal axes for three minutes each (Figure  9). Thermal vacuum testing was conducted with the unit oriented so that the weight of the separation  element was free to fall. This 2-N weight was considered a lower bound for the force the gimbal would be  able to exert if any adhesion occurred.  Figure 7. Gimbal configuration   The proto-type unit had 2 microns of Boron Carbide (B 4C) coating applied to both the separation element  and main housing. B 4C is an extremely hard (micro hardness 4700 kg/mm square) ceramic film with a  nano-crystalline structure applied via physical vapor deposition (PVD) [7]. It was initially chosen based on  the requirement for a “good hard coat ing”. Given the limited initial requ irements of good adhesion to  Titanium and applicable temperature range, no specific control of the application process was required of  the vendor and no adherence measurements were made before testing. No lubrication was placed  between the opposing surfaces. When the elements were separated after vibration testing, a powder had  formed between the two parts and extensive delaminating of the coating was visible (  Figure 10). No adhesion force significant to the application was observed between the parts. Energy  dispersive spectrometry (EDS) analysis and scanning electronic microscope (SEM) images of the coating indicated that pockets of argon gas had been trapped in the coating layers during PVD application leading  to poor wear properties (Figure 11). The coating was immediately rejected as unacceptable due to  contamination potential and the fear that the delaminating could leave exposed clean, bare Ti to Ti in a launch sequence leading to cold welding. This also resulted in a search for a new, more appropriate  coating for the hardware.   Due to scheduling constraints, a quick solution was desired. The separation element material was  changed to passivated stainless steel AM-355 that was the least reactive alternative per MIL-STD-889B  that also met strength requirements. A flight-like proto-type unit was re-lapped and DLN was applied on both elements and Braycote 600EF applied between them. This test failed because significant pre-load  was not applied (Figure 12). A bare AM-355 to Ti6Al4V test was also run in hope a high friction interface  would remain static and not create particulate (Figure 12). Separation occurre d after the thermal vacuum  cycles but significant fretting corrosion had taken place (Figure 12).   

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Figure 11. EDS analysis and SEM images of B 4C coating with argon gas bubbles  A case with only Braycote 600EF between the surfaces was also run but too much was applied without  any dilution. It was learned afterward that Braycote must be applied in diluted form using the standard  removal solvent. After doing application studies it was found that Braycote is very difficult to uniformly  apply in diluted form. After RV and T-vac testing significant chemical reaction had occurred between the  annealed Ti6Al4V and the Braycote (Figure 14). Reactions of these types are described in [8, 9]. This is  the only case that did not separate at low temperature. The Braycote became very viscous and acted as a sealant, creating a suction cup-like  phenomena. When room temperature had been reached, the  elements had separated at an unobserved time. Another run with Tiodize and MoS 2 doped with SbO 3 and  Au revealed Tiodize did not have good enough wear properties to survive the required environment (Figure 13).  Figure 12. Right: DLN coating failure due to lack of pre-load.  Left: Bare metal case  Figure 13 L: Braycote to Ti6Al4V interface R: Tiodize and MoS 2 after RV and T-vac Testing  140
Figure 14. Separation element Braycote EDS analysis  A new coating dubbed “chameleon coating” utilizing MoS 2 and graphite was investigated but not tested  due to its lack of flight history [10]. The friction data collected for this coating was promising and indicated  excellent friction characteristics in wet and dry conditions. This may be an ideal coating for similar future  applications where a complete testing pro cess can be applied to guarantee flight performance.  Figure 15. Chameleon coating friction characteristics  After further investigation, a final working interface wa s found. The initial attempt at this interface was to  deposit one and a half microns of TiCN via PVD on the Ti6Al4V housing and one micron of MoS 2 via PVD  doped with SbO 3 and Au to make the MoS 2 environment tolerant [11, 12] on the separation element. This  first attempt initially had two failed regions (Figure 16). A second attempt at this approach included  development of a specification for cleaning and preparation before coatings were applied. This process comprises an automated aqueous detergent ultrasonic cleaning, an in-situ cleaning and coating cycle  within a chamber having capabilities of controlled atmospheres and sub atmospheric pressures and a  post deposition tribo-finishing. This combination survived a total of 45 minutes of flight, proto-flight, and qualification levels. The toughness of TiCN with proper surface preparation on both elements coupled  with the high lubricity of MoS 2 allowed the sliding interface to survive four times longer th an required for  testing and flight. A similar combination called “Starvic” by  Platit [15] is claimed to have some flight history  on space applications. Work concerning TiCN’s low adhesion properties is contained in reference [16]. All  of the cases are listed in Table 1 .

It was found that coating thickness is very important. Four micron MoS 2 has considerably higher residual  stresses and significant delamination for only 3 minutes of flight levels occurs. Scratch testing revealed this weakness (Figure 17). The 1.0 to 1.5 microns of MoS 2 is the confirmed working thickness. 

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Table 1. Coating cases evaluated 

Figure 16 L: First TiCN-MoS 2 try.  R:  Successful TiCN-MoS 2 qual testing  The use of single point spherical interfaces with a moment arm ratio of the system described is  successfully protected in this application by 1.5 to 2 microns of TiCN and 1.0 to 1.5 microns of MoS 2  doped with SbO 3 and Au with proper surface preparation of both elements. This combination represents a  highly robust solution that survives significant wear due to slipping under long duration RV environments.  It was learned that recalibration of coating equipment should be required using sample specimens.  Samples should always be included in every coating application run with flight hardware, should always have a region masked and coating thic kness measured, and EDS analysis, scratch testing, and SEM  should be applied immediately after application to verify compliance.   

Conclusion  The fault tree method proposed is a valuable tool for carrying out the entire design process as opposed to  an uncontrolled original design effort. Applying this method to the LL re-design produ ced a very robust  solution to the failed LL design wi th very thorough qualification proc ess. Methods for me asuring stiffness  of the unit were successfully implemented and controlled. It was found that verified FEA methods are  important for single point restraint devices and that six spring representations are not adequate for accurate modeling. The proposed fault tree process, if used correctly, is highly beneficial. However, it can  142
become over-burdensome if difficult assessments, such as the margin assessment mentioned earlier, are  required. It is recommended that this process become a standard practice for critical mechanism  applications. Many other methods for careful thinking coul d be proposed but it is extremely important that  the mechanism community standardize methods for doing good design and clearly communicate this information to new personnel. 

Through this fault tree driven design process, the major design issue that arose concerned finding a coating/lubricant combination that would not create debris under RV conditions. It was found that it is  better under the single point spherical socket conditions de scribed to allow slipping to occur freely over a  smooth hard coating with very low friction rather than trying to constrain the interface completely. A robust solution was produced using TiCN and MoS 2 doped with SbO 3 and Au. This solution is strongly  dependent on applying the correct coating thickness. It is recommended that future applications carefully  weigh the no-match-machining benefit of a single point spherical interface constraining a significant  inertia against use of a high friction flat or conical interface which will have fewer wear and coating issues  [6].    Figure 17. SEM Images of 4 micron thick (L) and 1.5 micron thick MoS 2 after scratch testing 

Acknowledgments   This work was an extensive process executed by several people who played key roles and deserve  acknowledgement. The design effort was aided by Roger Sirois, Prasad Somuri, Scott Nicolaysen, Chris Sorensen, Elmer Collins, Steven Moya, and Mark V aughn. Testing and qua lification were aided by  Nathaniel Roberts, Paul Larkin, Brendan Rogillio, and Rand Garfield. Stiffness validation was led by Vit  Babuska who was aided by Troy Skousen, and James Freymiller. Detailing support was lead by Wilbur Martin. Many others also helped with this work to a lesser degree than those mentioned.  References 

1.   Wilbur Shapiro, Frank Murray, Roy Howarth, and Robert Fusaro, “Space Me chanisms Lessons  Learned Study,” NASA, 1995 http://www.grc.nasa.gov/WWW/spacemech/lessonscover.html 2.    NEA electronics corporate home page:  http://www.neaelectronics.com/ 3.    Ren, Y; Beards, C. F., “Identification of joint properties of a structure using FRF data,” Journal of  Sound and Vibration; 7 Sept. 1995; vol 185, no. 5 p.845-66  4.   Yang, TC; Fan, SH; Lin, CS; “Joint stiffness identification using FRF measurements” Computers and  Structures; Nov. 2003; v.81, no.28-29, p.2549-2556.  5.    Lee, U; Cho, JY ; Cho, CD, “Spectral FEM-based dynamic characterization of structural joints: Theory  and experiment,” Key Engineering Materials; 2006; v.326-328, pt.1-2, p. 1629-1632.  6.    Santos, LV; Trava-Airoldi, VJ; Corat, EJ; Nogueira, J; Leite, NF; “DLC cold welding prevention films  on a Ti6Al4V alloy for space applications.”  Surface & Coatings Technology; 24 Jan. 2006; vol.200,  no.8, p2587-93  7.     Boron Carbide “Diamond Black” Bodycote information flyer:          http://mcna.bodycote.com/index.asp?sid=process&content=process.asp?ID=253 143