ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
391 place. The CCLL would use this channel. Several flight spare ERM release/restraint devices were on  hand as ESDM spares. These would be used as the schedule precluded obtaining other release devices, and the available control electronics were designed to actuate this particular device. Spare ESDM potentiometers were also available, and one was used as as a sensor to verify the the state (locked, unlocked, or somewhere in between) of the launch lock. The driving requirements for the CCLL are listed in Table 3. 

Table 3. CCLL Driving Requirements 

Requirement Value  Keel Dimensions 17.24  x  19.48 x 8.50 inch (438  x 495  x 216 mm)  Keel Mass 90 lbm (40.9 kg)  Allowable Keel Deflection < 0.015 inch (0.4 mm)  Stowed Frequency ≥ 60 Hz  Cold Operational Temperature -40°C  Uncompensated Momentum ≤ 25 Nms  Uncompensated Torque ≤ 158 Nm 

The cryo-cooler is mounted to an aluminum Keel plate which houses the compressors, expander, and cold heads. This Keel is suspended beneath the instrument -X panel, near the focal plane array (FPA). The first stage is coupled to the telescope 170K cold shields, and the second stage cold head is coupled to the 43K FPA. Both stages utilize flexible conductive link heat straps to provide a thermal conduction path that would not transmit jitter loads from the cryo-cooler to the FPA or shields. Circleflex™ damping flexures were used to provide passive vibration isolation between the keel and the instrument structure. Use of these flexures reduced the cryo-cooler frequency to levels which would damage the cryo-cooler at launch. The Keel Assembly is shown in Figure 16. 

Figure 16. The Keel Assembly, as modifed with CCLL interfaces is shown. The 3 posts were added  to provide accessible interface s for for a launch lock mech anism mounted on the opposite  (interior) side of the -X instrument panel. The triangular frame was added later to increase overall  stiffness.
392 The only available volume available for the launch lock was within the instrument, opposite the externally  mounted keel. Posts would have to extend from the keel through clearance holes (that would have to be cut) in the -X instrument panel. The lock hardware would restrain these legs.  To properly launch lock the cryocooler, it must be constrained in all 6 degrees of freedom. Only a single  ERM-type release device could be used; hence the system could only be constrained at a single point. It was also desirable for the ERM to be mounted on the instrument exterior to facilitate easy access for manual resetting. Additionally, the cryo-cooler could never displace by more than 0.015" (0.38 mm) due to close clearances between the flexible conductive links and the FPA cold shield. Calculations showed that the flexures would displace the keel by no more than 0.010” (0.25 mm) when gravity was removed, or as the gravity vector changed when the instrument was rotated. The engaged launch lock carried the launch loads of the cryo-cooler keel, and when disengaged, allowed the keel to float on the passive isolator flexures. 

Figure 17. The CCLL Assembly, as shown with the instrument -Y and -Z panels removed. Note that  the Rotor Arm protrudes through the -Y panel, and the ERM and Lingage Guide Housing are  mounted on the exterior side of the -Y panel. A spare ESDM potentiometer is used to verify rotor  arm position and indicate proper deployment. It is mounted to the -X panel underneath the  rotation axis of the rotor arm, and is tied to the rotor arm using a compliant coupling similar to the  one used on the ESDM.
393   Figure 18. The overall architecture of the CCLL is shown above. The solid colored components  show the launch lock in the unlocked position, with the ERM Cap released from the ERM body.  The wireframe shows the position of the launch lock in the locked position, with the ERM Release  Cap locked to the ERM body. The Rotor Arm rotates ~9.5 degrees between the locked and  unlocked positions. Commercial grade tie rod end hardware was used in all of the linkages.
394 Pin Pusher Actuators  Several concepts were traded, with a 3-point constraint architecture selected. This used 3 drive pins on  the instrument going into three holes on the cryo-cooler keel structure posts. The dimensional clearance between the pins and the holes is only 0.002 inch (0.05 mm). Each point constrained 2 translations and 2 rotations, providing a 2-2-2 degree of freedom kinematic lock. To release the launch lock, the pins were pushed through the holes by spring pistons. To facilitate resetting, the pins were not completely removed from the holes, and the holes were given a toroidal profile. The drive pins were stepped down so that a thicker diameter matched the toroidal hole diameter, and a thinner diameter provided adequate free clearance. The pins were pushed the distance required to remove the larger diameter pin length from the toroidal bushing hole engagement diameter to a seating hole in the anchor fitting. Each restraint point had a pin pushing, compression-type spring, as shown in Figure 19. Each drive pin was tied to a central rotor disk via a drag link. As the pins are pushed out of the engagement bushings, the rotor disk rotated approximately 9.5 degrees, as shown in Figure 18. If the rotor disk is not allowed to rotate, the pins are held in place and the springs are compressed. If the disk is free to rotate the springs are free to expand and push out the drive pins. The disk also couples the motion of all of the pin pullers together; if any single spring fails, the rotor disk rotation, powered by the surviving springs, will pull the pin with the failed spring.  

Figure 19. The spring-powered, pin pusher actuato r shown locked (stowed) in the upper figure,  and unlocked in the lower figure. The ERM restraint device prevents the compressed drive spring  from pushing the the green Drive Pin out of the Toroidal Bushing. When the ERM releases the  central rotor, the drive spring is free to expand and pushes the Drive Pin out of the Toroidal  Bushing and into a guide hole in the anchor fitting. To reset the actuator, the the Drive Pin is  manually pulled back through the Toroidal Bushing (v ia the Reset Linkage a nd the Rotor Arm) and  connected to the reset ERM.  The titanium Anchor FIttings and Toroidal Bushings had a Tiodize® type II (Teflon® impregnated) finish. The Drive Pins are made of aluminum bronze, CDA 63020 per AMS 4590B. A light film of Braycote
395 Micronic 602EF grease was applied to the Drive Pin, and to the sliding surfaces in the anchor housings  and the toroidal bushings.   Release/Restraint Device  The ERM restraint device prevents the rotor disk from rotating. When fired, the ERM releases the disk, the pins retract, and the cryo-cooler unlocks. The ERM is not coupled directly to the rotor disk. An extension arm runs from the rotor disk through a clearance slot in the instrument -Y panel. A tie rod runs from the extension arm to the ERM mount. This allows the ERM to be mounted on the instrument exterior, facilitating reset access. A potentiometer is used to indicate rotor position, providing a positive indication of the locked or unlocked state. The ESDM potentiometer was used, as spares were immediately available. Telemetry lines to support this sensor were also available in the control electronics.   Cryo-Cooler Launch Lock Mechanism Development Issues   To reduce risk and expedite schedule, a functional breadboard unit was fabricated quickly to validate the overall architecture, and the flight unit and an engineering test unit were developed in parallel. The developmental intention was to qualify the design in parallel with the flight fabrication. There was no schedule available to environmentally qualify a unit and then fabricate the flight unit serially. It had to be done in parallel. The ETU unit was built onto a flight-like composite panel for environmental testing. The flight unit was built up directly onto the flight structure and was qualified at the instrument level. The ETU and flight units were identical. 

Figure 20. In the left figure, the flight Cryo-Cooler is shown in development at the vendor,  mounted within the flight Keel structure. Vibration testing of the ETU Keel Assembly (with CryoCooler mass simulators) to verify post stiffness is shown in the center figure. The CCLL functional  breadboard unit is shown in the right figure.  It became evident during the initial development and analysis of the mechanism that 60-Hz stiffness was going to be a very diffcult requirement to meet, with initial concepts predicted to be ~42 Hz. Several design changes were implemented, including optimizing the Drive Pin/Toroidal Bushing clearance and the addition of the Triangular Frame, which rasied the system stiffness to meet and finally exceed the requirement. The CCLL first mode frequency was measured at 65 Hz.  Upon actuation, the CCLL produces a peak torque of 99 N-m (877 lb-in), and an angular momentum of 12.4 Nms (110 in-lb-s), within the limits specified in Table 3.  The CCLL ETU successfully passed qualification testing in September 2011 and the flight unit has been installed into the TIRS instrument. Many issues were encountered and ultimately solved in this unique mechanism.
396   Figure 21. The CCLL is shown during installation into the TIRS instrument, with the -Y panel  removed. The left figure shows the rotor arm, anchor fittings, and prominent triangular frame. The  Cryo-Cooler legs were integrated into the CCLL before the keel, and attached to it later. The drive  springs are visible in the anchor fittings just beneath the triangular frame. The right figure shows  the instrument with the -Y panel replaced, with the ERM mounted on the right and the rotor arm  coming through the panel on left side, with the two joined by a linkage. The gold colored frames  that partially obscure the ERM hardware are box enclosures that will keep thermal blanketing  clear of the moving parts, while allowing access for manual ERM resetting.  Use of commercial grade mechanical components for space flight  Due to the compressed schedule, procuring flight quality tie rod ends and/or spherical bearings was  impossible. The only option was to fly commercial industrial components. Traceability paperwork to verify material type and quality was generally not available, or was not sufficiently trustworthy for mission assurance. To reduce risk, many extra components were procured, and used not as spares, but for testing. Components were proof tested to measure yield and ultimate tensile strength, and chemically analyzed to verify material type, finish, and surface hardness. Parts that were selected from the lot for flight were static-load tested to qualification load levels before use. The commercial components used in the CCLL successfully passed mechanism component level and instrument level testing. 

Lesson Learned: Generally available, commercial industrial hardware may be qualified for spaceflight use if a comprehensive set of tests to verify base material type, coatings,finish, yield and ultimate strength are carried out. The hardware should be purchased in lots and many indentical units tested to determine unit-to-unit statistical differences and aid in flight part selection from the lot.  Unpredicted behavior during vibration testing  During vibration testing it was found that the frequency of the system would increase as full level input  was approached. This occurred in the lateral directions only. The shifts in frequency for the Z lateral axes are shown in the Figure 22. The shift that was seen during testing in the Y-Axis was not as significant as the one seen for the Z-Axis. The Y-Axis shift was 10 Hz from -18 dB to Full Level. The shift for the Z-Axis was 31 Hz from -18 dB to Full Level. The predicted frequency was higher than the 65 Hz measured during testing, but that was for an idealized system, and did not take into account a number of factors that could be attributed to the difference.
397   Figure 22. The frequency of the stowed CCLL was shown to increase as the test level was  increased to full level. Note how the first mode frequency peaks shift the right (increase) as the  test level increases from -18 dB to full level. This behavior was not predicted, and occured in the 2  lateral (ZY mechanism plane) axis. It did not occur in the axial (X, thrust) axis.  The shift was a result of the gap clearance around the Drive Pin to Toroidal Bushing interface. The gap  was necessary to ensure that the mechanism would operate reliably at cold temperatures and was misalignment tolerant. As more energy was input into the system the gap would close, and the system would become stiffer. This characteristic made predicting the behavior of the system difficult, but established that smaller clearances made the sytem stiffer.  Lesson Learned: Mechanisms which utilize sliding interfaces are difficult to model accurately, and will  generally be less stiff than predicted . The measured system frequency will increase as test  loads increase  to full levels. This is due to sliding clearances being reduced in proportion to input level during vibration tests where the lines of action of the sliding device are close to the test axis. This was not seen when the test axis was perpendicular to the mechanism lines of action.  

Conclusions   The SSM, ESDM, and CCLL span a broad range of mechanisms types; from microradian precision mirror positioning to coarse deployment of the large earth shield. Due to the compressed schedule, it was not possible to fully develop engineering test unit mechanisms before the flight mechanisms. A protoflight approach had to be utilized to save time, and even then some prudent development processes, such as the encoder electronics signal integrity analysis, and the SSM bake out, were sacrificed. Higher levels of risk had to be accepted to meet the delivery schedule. We were able to mitigate the consequences resulting from the analytical and contamination control omissions, and the heritage justification to forego critical analysis was proven false. The building of a simple breadboard SSM mechanism, which verified 1.00E-081.00E-071.00E-061.00E-051.00E-041.00E-031.00E-021.00E-011.00E+00 10.00 100.00 1000.00ASD (g2/Hz) Frequency (Hz)Z Axis Vibration Response Overlay for Cryocooler Mount 19Z from -18dB of Full 19Z from -12dB of Full 19Z from -6dB of Full 19Z from Full Level -18dB-12dB-6dB Full Level
400                                          Figure 1.  The locations of two filter wheel assemblies and the one pupil imaging lens assembly  are shown on the module A (top) beryllium optical bench. Stray light shrouds appear transparent  in the figure to show the PIL located within. 

The FWA is designed to insert specific optical elements into the NIRCam optical beam. Each FWA unit  consists of two independently driven wheels that each hold 12 optical elements and is required to insert each optic element into a target position with a repeatability of ±75 microns. Each FWA unit is configured with primary mirror wave-front sensing elements, optical calibration sources, and numerous optical filters dependent upon which longwave or shortwave path the FWA resides.  An illustration of the FWA is shown in Figure 2. The primary function of the PIL assembly is to deploy a set of optics into the beam path. Once inserted, the NIRCam instrument focus is changed to image the 18-segment, 6.5-meter diameter, primary mirror instead of deep space. To achieve this function, the PIL must deploy its optics into the specified  point in the beam path with a repeatability of 0.016 degree. The PIL assembly is shown in ).  Figure 3 . 

         Figure 2. The filter wheel assembly (FWA).  Fi gure 3. The pupil imaging lens assembly (PIL).  PIL located within  optic shrouds on shortwave optic Shortwave filter wheel assembly Longwave filter wheel assembly
402 Finding the Problem 

Testing of the FWA prototype, and the FWA and PIL qualification units successfully demonstrated all performance requirements, but a planned disassembly of the PIL revealed an unforeseen problem in the bearing mount. The PIL qualification unit was fully assembled and performance tested at ambient temperatures and then subsequently subjected to proto-flight random and sinusoidal vibration testing. Because of the unique optic configuration of the PIL, the first vibration test of the system was performed with a surrogate PIL optic installed. After successfully completing this initial vibration test, the PIL was partially disassembled to swap out the surrogate optic with the actual optic. At this point, contamination was found at the base of the PIL shaft as shown in Figure 5.        Figure 5.  Contamination found in the PIL qualifica tion unit after vibration testing, and a diagram  of the PIL bearing mount illustrating the suspected contamination path. 

Investigation into the powdery contamination found in the PIL qualification unit initially focused on the  bearing sleeve which showed evidence that the some of the anodic coating was worn away, but the investigation was quickly expanded to look at the entire bearing assembly. Magnified optical inspection and scanning electron microscope (SEM) analysis of the sleeve showed evidence of wear with debris in the form of conglomerate particles. This type of conglomerate particle debris which is made up of many  sub-micron sized particles is characteristic of fretting wear 1. The wear debris found on the sleeve  contained mostly titanium from the shaft and sleeve base material as well as silicon and oxygen from the  anodic coating. However, it seemed unlikely that all the debris could come from the wear areas between the sleeve and shaft. Analysis of debris from other areas at the base of the PIL shaft showed the same conglomerate particles, but these particles consisted of titanium and iron, but did not contain silicon and oxygen. The iron was most likely from wear of the bearing. Titanium is a fretting-wear sensitive material, and it became clear that the microscopic motion of the flexure bearing mount during vibration testing resulted in the wear-generated particulate contamination seen on the PIL qualification unit. Since a similar bearing mount was used in the FWA design, the FWA qualification unit was disassembled to inspect for possible fretting wear.  After disassembling the FWA qualification unit, it became evident that the assembly suffered from the same fretting wear issue. Much of the bearing-to-shaft interface surfaces were covered with similar wear                                                    1 Fretting wear is a type of adhesive wear that will occur wh en contacting surfaces are undergoing small, oscillatory,  tangential displacements.  The relative sliding motion causes localized adhesion and disruption of the surface  generating fine particulates which then oxide and become imbedded back into the surface causing further abrasion  damage.  Contamination particles  Contamination path
403 particulates as seen on the PIL design. However, nodular buildup of fretting wear particles had collected  in relief areas of the design as shown in Figure 6 . These 0.127-mm (0.005-inch) diameter nodules had  the consistency of dry powder but would easily smear when touched or dissolve into a paste with  isopropyl alcohol. The material composition, as determined from a scanning electron microscope analysis, was the same as what was seen in the PIL. Due to the stringent cleanliness requirement for these mechanisms and the optics that surrounded them on the NIRCam bench, this contamination from fretting wear of the flexure bearing mount was unacceptable. The next step was to resolve the issue.   

Figure 6.  Similar fretting w ear found in FWA qualification uni t bearing shaft. The 0.127-mm   (0.005-inch) nodules collected from the bearing relief c onsisted of nanometer-sized particles of  titanium from the shaft and iron from the bearing and had the consistency of dry powder that  smeared when touched. 

Fretting Wear Resolution Trade Studies   The fretting wear contamination was clearly unacceptable, but what course of action could be taken that  would ensure no contamination of the nearby optics, no performance degradation of the different functionalities of the FWA and PIL mechanisms, and not result in significant delays to the program? The team considered a number of possible solutions including using a conventional hard-clamp bearing mount, trying to contain the contamination, fabricating the shaft from a less fret-wear sensitive material, adding different wear resistant coatings to the shaft, as well as combinations of these ideas. The team was uncomfortable with a hard-clamped bearing design for two main reasons. First, the team was already considering changing the material of the bearing flexure clamp. Secondly, the test history of the current design had met all other performance requirements, and a hard-clamped bearing mount design would likely require a new set of qualification testing. Therefore, the design team focused on changes to materials and coatings and possible containment.  There is a wealth of knowledge related to wear-resistance coatings and materials for many industrial applications as well as many aerospace applications. However, the knowledge base dwindles significantly for wear-resistant materials and coatings for applicatio ns at cryogenic temperatures that must operate in  a vacuum and also survive launch vibrations in a dry air environment. By investigating known heritage materials and coatings used for near-ambient temperature space applications and more recent coating technology innovations that show promise, a selection of four candidate solutions where chosen. It was clear that test data at cryogenic temperatures for the selection was all but non-existent, and even the ambient test data did not provide wear performance related to the FWA and PIL fretting wear issue. The team quickly determined that a test was the best method to discern the best solution for the problem.