ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
the end of the development test program. In order to improve the durability of the capture sensor flight  design, a grease lubricant was added to the wear surfaces of the capture sensor, with a plan to maintain  the lubricant between flights. When qualification 6DoF testing of the modified design was completed in  2016, the unit completed 216 docking cycles with one re-lubrication midway through the test matrix. The  lubricated sensors showed no visible signs of wear on the anodize coating or opposing ring contact area,  indicating that the grease lubricant was effective at mitigating damage to these surfaces.  Figure 14. Capture sensor and opposing ring wear pattern during development testing  A key design decision that was validated during 6DoF testing was actuator optimization. The selection of  actuator type and careful design of actuator components were important in optimizing the responsiveness  of the soft capture system. Because system requirements allowed enough time for the docking vehicles to  be slowly maneuvered into position, relatively little force output was needed from the linear actuators. As a  result, the need for aggressive gearing was avoided.   Due to the fidelity compromises in the engineering development unit (EDU) used for the first round of 6DoF  testing, the linear actuators were less optimized than the flight design. The capture sensors on the passive  Space Station side of the docking interface (similar to those on the active side) were also omitted. The  second round of 6DoF testing used a high-fidelity qualification test article for both the active and passive  halves of the docking system. In the qualification 6DoF test, the linear actuator‘s design fidelity was  increased, and the passive side capture sensors were included. The optimized actuator design reduced the  impact energy between the soft capture rings at the instant of capture, while the addition of the passiveside capture sensors helped further cushion the impact. The net effect of these changes is evident in the  axial docking loads measured by the 6DoF facility load cells that were mounted underneath the passive  test article. Figure 15 compares the axial docking loads for the EDU and Qualification 6DoF tests for the  case of an 18 t chaser vehicle docking to the ISS forward docking adapter, with a 5 deg Roll offset between  the two mechanisms at the instant of contact, and a 4.5 cm/sec approach velocity. The results indicate a  significant reduction in peak axial force at the moment of capture.  NASA/CP—2018-219887 483
Figure 15. Axial docking force from 6DoF testing of an 18 t chaser vehicle with ISS with a 5 deg  Roll misalignment  The results of the 6DoF testing validated the SCS design approach. The use of optimized linear actuators  allows the SCS to control its motion and limit interface loads without the need for force feedback. As  demonstrated by component-level thermal testing of the linear actuators, this capability extends over the  expected range of on-orbit thermal environments. Furthermore, the design is extensible to a variety of target  and chaser vehicle mass properties through adjustment of control law parameters.  Conclusion  The NDS is NASA’s first docking system since the Apollo-Soyuz Test Project of the 1970’s, and the first  implementation of the new International Docking System Standard. Its direct electrical architecture  leverages modern technology to greatly reduce mechanical complexity compared to the heritage APAS  design used for the space shuttle. Its innovative controls approach eliminates the need for post contact  thrusting to effect capture, reducing operational complexity for vehicle providers, flight controllers and crew.  Its flexible architecture allows for controller settings to be adjusted either on the ground or in flight in order  to capture target vehicles ranging from small 5 t landers to the 450 t International Space Station. Importantly,  it also provides a robust capability to recover safely from a wide variety of failure scenarios. As a key enabler  of modular mission architectures, NDS will be critical to making NASA’s vision for space exploration a  reality.  References  1. Fehse, W. Automated Rendezvous and Docking of Spacecraft. Cambridge University Press. 2008. p. 3 2. Report SD 72-CS-0048-3, Apollo/Salyut Test Mission Study Report, June 1972, Space Division North American Rockwell 3.Motaghedi, P. and Ghofranian, S. “Feasibility of the SIMAC for the NASA Docking System.” AIAA Space and Astronautics Forum and Exposition (SPACE 2014) ; 14 Jul 2014. NASA/CP—2018-219887 484
Capture Latch Assembly for the NASA Docking System  Brandon N. Dick*, Nathan Mauch* and Timothy Rupp** Abstract  This paper will present a summary of the Design, Development, and Qualification of the Capture Latch  Assembly (CLA) for the NASA Docking System (NDS) Block 1 (NDSB1). The CLA is an integral part of the  Soft Capture System (SCS) of the NDSB1, serving the purpose of connecting the mating SCS Rings of two  docking vehicles. The paper will present an overview of the function of the CLA and its basic concept of  operations, including a summary of the major components of the CLA. The development, qualification, and  production of the CLA will then be described. Particular focus will be provided on two major issues that  occurred during production and qualification of the CLA. The first issue was failures of the CLA Motor (CLM)  during acceptance testing. The failures of the CLM were ultimately determined to be due to design defects  and manufacturing errors in the motor commutation sensor assembly. The second issue was failure of the  secondary release mechanism, or Contingency Capture Latch Release (CCLR) mechanism during  development and qualification testing. The CCLR failures were found to be a result of excess free play in  the release mechanism, resulting in wear leading to galling inside the release mechanism. An overview of  each failure will be provided, along with a summary of the failure investigation and recovery process. Finally,  Lessons Learned from each of the major issues and the overall development of the Capture Latch will be  presented.  Introduction  The Capture Latch is an indeterminate mechanism in the NDSB1 tasked with providing the initial connection  (soft capture) between space vehicles during a docking procedure. The Capture Latch maintains the  connection between space vehicles during initial contact, vehicle alignment, and vehicle hard capture. Once  hard capture is completed the capture latches are released and the NDS SCS is stowed. Each NDS has  three latches, one on each petal of the NDS as shown in Figure 1.  Figure 1. NDS with 3 Capture Latches  *The Boeing Company, Huntsville, Alabama ** NASA Johnson Space Center, Houston, Texas Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018  NASA/CP—2018-219887 485
The CLA is a derivative design of the latch on the docking system used in the Apollo–Soyuz Test Project  (ASTP) [1]. The ASTP latch design was further refined for follow-on docking systems including the APAS95 docking system used in the Space Shuttle [2]. There were several critical upgrades in this design revision  that were previously absent. They include the ability to release under significantly larger loads, meeting  requirements set forth in NASA-STD-5017 (design for minimum risk), and an automated secondary release.  Design Requirements  The primary CLA requirements were derived from system performance characteristics and needs. The  primary requirement was to be able to release under a specified tensile load quickly enough to ensure two  vehicles could detach during a failed docking attempt without collision between spacecraft or damage to  the NDS.  The secondary release requirements were derived for conditions under which the capture latch experienced  a failure in the primary drive train and had to be released very quickly. The pre-capture force of the  mechanism had to be limited to reduce the amount of forward vehicle velocity needed to push the latches  out of the way and complete a capture. The latch is designed for a relatively short mission life of 214 days  on-orbit with up to 4 dockings per mission.  Design Overview  Figure 2 shows the main components for the capture latch mechanism. They are:  1. Motor: Provides the nominal actuation for the mechanism. The motor is a dual wound brushless motor with redundant Hall-Effect Device (HED) position sensors. Each string of the motor is driven by a separate controller. 2. Latch Pawl: Latching feature that reacts load from Passive Striker to attain capture between mating docking systems. Can be positioned to latch (Ready to Capture) or release (Ready to Release) using the motor. 3. Transmission/Linkage System (internal, not shown): Transmits torque from the motor to the Latch Pawl and retains the pawl in desired position. 4. Secondary Release Mechanism: Provides for secondary release in the event of a nominal drive system failure. The mechanism contains a Non-Explosive Actuator (NEA) that is activated with a simple on/off power supply. The mechanism also contains a compressed spring that provides the force necessary to extend the Secondary Release Mechanism and drive the latch into the released mode. 5. Passive latch striker plate: This is a simplified representation of the stationary latch interface hardware on the passive docking system. Figure 2. Latch Nomenclature  NASA/CP—2018-219887 486
The capture latch mechanism has three main configurations or modes, as shown in Figure 3 and Figure 4:  Ready to Capture, Ready to Release, and secondary release. The Ready to Capture position (Figure 3.1)  is the nominal mode that allows for the latch to capture and hold the passive striker on the opposing docking  system. The ready for release (Figure 3.4) and secondary release (Figure 4) modes allow for the release  of the striker plate. The Ready to Release mode is the nominal release position achieved via the motor  drive system. This is the primary release mode utilized while the latch is operating nominally. The secondary  release mechanism allows for the release of the passive striker plate in the event the latch drive system is  rendered inoperable (whether due to motor failure, jamming, binding, seizing, etc.). The secondary release  mechanism is operated via a Non-Explosive Actuator (NEA) that releases a spring-operated push rod. Once  the secondary release mechanism is fired the system is in an unrecoverable released state and the latch  can no longer perform docking.   A full nominal capture and release operation is shown in Figure 3. The operational steps are as follows:  Step 1: Latch is set to the ready to capture mode.   Step 2: During docking the incoming passive latch depresses the latch pawl.  Step 3: The latch pawl passively snaps over the striker plate.   Step 4: To release the motor and drivetrain reposition internal latch linkages to a Ready To Release  configuration.  Step 5: The latch pawl is then pushed and rotated out of the way by the passive latch striker plate.  Step 6: The motor and drivetrain reconfigure the internal latch linkages into the Ready To Capture  mode. The latch is ready for another nominal docking operation.  1. Ready to Capture Mode 2. Capturing Passive Latch 3. Passive Latch Captured 4. Remode to Ready to Release Mode 5. Passive Latch is Released 6. Return to Ready to Capture Figure 3. Nominal Capture and Release Sequence  NASA/CP—2018-219887 487
1. Passive Latch Captured 2. Secondary Release (NEA) Activated  3. Passive Latch is Released 4. CLA is Permanently in Release Mode  Figure 4. Secondary Release Sequence  A secondary release operation is shown in Figure 4. The operational steps are as follows:  Step 1: Latch is in the ready to capture mode and has captured the passive latch.  Step 2: After a failure of the primary drive system, the secondary release mechanism is activated,  and the NEA releases energy stored in the compression spring causing the secondary release  mechanism to extend.  Step 3: The latch pawl is then pushed and rotated out of the way by the passive latch striker plate.  Step 4: The capture latch is now permanently in the secondary release mode.  Summary of Development/Qualification/Acceptance Testing  Development Testing  The purpose of the Capture Latch Development Test was to reduce technical risk associated with the  assembly and test of the flight design configuration by conducting testing early in the project. The  development unit was built to the same design configuration as the flight Capture Latch. Some variations  from the flight design were permitted to facilitate a quick test. First, the NEA in the Secondary Release  Mechanism was a development article and the spherical bearing at the base of the NEA was not swaged  in place. Second, the Capture Latch Motor and Controller were also development articles. Lastly, the cover  for the controller was a 3D-printed metal casing, instead of the flight machined part.  The development test consists of the following tests: 1. Run-In, 2. Functional, 3. Random Vibration  (qualification levels/durations), 4. Thermal Vacuum (TVAC), 5. Primary Release, 6. Secondary (NEA)  Release, and 7. Static (Ultimate) Load. Run-In testing consists of five functional cycles, where the Capture  Latch is cycled between Ready to Capture and Ready to Release modes. Functional testing starts with the  Capture Latch in Ready to Capture mode. A test fixture is used to simulate docking and verify that the latch  engages a passive latch striker. The latch is then commanded to Ready to Release mode and the test  NASA/CP—2018-219887 488
Figure 10. Load Washer Light Microscroscopy The investigation into the galling covered many aspects of the design and test for the NEA and Capture  Latch, but immediately focused on random vibration testing as the most likely source. It was proposed that  free play in the secondary release mechanism along the pin axis could have permitted the secondary  release mechanism to move excessively during vibration testing, causing the spool to move relative to the  load washer. With enough motion, this could remove all the dry film lubrication coating, and allow galling  between the two surfaces. The design of the mechanism included free play along the pin axis to facilitate  assembly under worst case tolerance conditions. Additionally, the NEA was not preloaded. To test this  theory, the Development Capture Latch was reassembled to be subjected to a repeat of the random  vibration test and secondary release test. During the testing, high-speed video was used to capture the  dynamics of the secondary release mechanism movement.   This test was informative, showing that the motion of the mechanism was extremely violent, subjecting it to  nearly continuous shock loads as the mechanism shifted between the hard stops on either side of the free  play. This motion was especially evident during testing in the axis parallel to the pins. Following the  completion of the vibration testing, the unit was subjected to a Secondary Release Test, which it passed.  The NEA was disassembled, and similar wear was observed. Although the failure could not be duplicated,  the excessive wear was duplicated, and the test was considered successful.  Based on these results it was determined that the mechanism should be redesigned to remove free play.  The design improvement was implemented by installing flat washers and Belleville washers on the pins to  remove all gaps, with a slight preload, as shown in Figure 11. Additionally, the NEA was given a small  preload. The development capture latch was rebuilt with the redesigned configuration and subject to the  same set of tests. The redesigned configuration showed no signs of the galling/excess wear. The  Qualification Capture Latch was then rebuilt per the new configuration and subjected to random vibration  testing and a secondary release test. The unit passed and was shown to have no excessive wear during a  post-test inspection.  NASA/CP—2018-219887 495
Figure 11. CCLR Detail Lessons Learned  Avoid Loosely Constrained Parts   Loosely constrained parts can be damaged when subjected to vibration or dynamic loading events. Avoid  loosely constrained parts whenever possible.  Thermal Stresses In Potted Parts  Thermally induced stress in potted parts due to differences in CTE between the potting material and the  supported parts can be significant.   Fully Address Failures During Development Testing  Not fully addressing or investigating failures during development testing can allow design defects to  propagate into the flight design and production.  Watch the Test Whenever Possible  The Test Engineers/Technicians may not see the same things that a designer would. When the design  engineers observed the vibration test for the first time, it was very clear that the motion induced by the test  was unacceptable. If they had observed the development random vibration test, the issue may have been  addressed much sooner in the design.  Use Caution with Commercial Off The Shelf (COTS) Parts  Using COTS parts in a design results in a less rigorous verification of the part performance and suitability  for use in the desired application. When using COTS parts, or tailored COTS parts, the designer is  responsible for ensuring that all critical interfaces between the COTS part and the assembly are compatible.  NASA/CP—2018-219887 496
Conclusions  After overcoming the difficulties discussed in this paper, the Capture Latch Assemblies for the NDSB1 were  successfully designed, qualified, and manufactured/tested. A total of twelve flight assemblies – representing  four NDSB1 units – have been completed. A number of lessons were learned during this process which will  be applied to future CLA production and development as part of the next version of the NDS – the NDS  Block 2, which will be used with the Orion and Deep Space Gateway programs.  References  1. Swan, W, “The Apollo –Soyuz Test Project Docking System,” 10th Aerospace Mechanism Symposium , NASA TM 33-777, Pasadena, CA., 1976, pp. 29, 36. 2. Cook, J., Aksamentov, V., Hoffman, T, & Bruner, W. “ISS Interface Mechanisms and their Heritage”, AIAA Space 2011, pp. 24. 3. Dick, B., Oesch, C., and Rupp, T., “Linear Actuator System for the NASA Docking System,” 17th European Space Mechanisms and Tribology Symposium. NASA/CP—2018-219887 497
Figure 1. A typical high-speed canned motor pump.  These complicating factors aligned in an unfortunate way on the International Space Station in 2010 and  led to the premature failure of one of two active thermal control pumps used to maintain heat balance within  the habitable volume. A failure investigation on this pump indicated that extreme wear of the bearings was  the root cause due to operation beyond the hydrodynamic load capacity of the bearing. A complete  description of this failure and early operational recovery efforts has been documented in [13].   This failure triggered an extensive study of the wear rates for this bearing material in relevant environments.  An experimental approach was taken for this investigation due to the high rubbing speed and flooded, liquid  condition at which this wear occurs. A new experimental apparatus was designed, constructed and  operated to generate a database of relevant wear rate coefficients suitable for predicting wear life of highspeed canned motor pumps in operating microgravity conditions. The apparatus, known as the Extreme  Environment Tribometer (EET), includes an operating temperature range of -68°C to 65°C (-90° to 150°F),  pressure range from 6.9 to 2585 kPa (1 to 375 psia), and rubbing speeds up to 25 mps (81 fps). It was  designed to operate in 100% anhydrous ammonia and therefore will also accommodate both the new class  of designer heat transfer fluids with anti-corrosion additives as well as the traditional heat transfer fluids  such and propylene glycol and water.   Wear in Hydrodynamic Bearings  Hydrodynamic bearings, such as those commonly used in canned motor pumps, experience wear at both  high-speed and low-speed operation. Low-speed wear is primarily experienced during startup and  shutdown of the rotating system. Figure 2 has been reproduced from [13] to demonstrate this phenomenon.  At low rotational speeds the shaft contacts the static bearing until the mixed lubrication regime is reached.  As speed is increased further, beyond the mixed lubrication region, the bearings operate in a non-contacting  or hydrodynamic mode. In this mode, no wear occurs and the surfaces are separated by a thin layer of  lubricant, or working fluid in the case of a canned motor pump. In the low-speed phase of operation, solid  surface contact occurs and a wear condition may exist. Typically, the static bearing material is selected to  tolerate this wear while the shaft in the bearing location is prepared with a hard coating having a very low  surface roughness. This low-speed wear condition is a well understood phenomenon and wear coefficient  data of various bearing and shaft coating combinations is readily available. Even if the data is not readily  available, wear coefficient data at low speeds can be obtained from a test program conducted on a large  number of commercially available pin-on-disc or block-on-ring tribometers. In fact, in the early phases of  the work described herein, low-speed wear coefficient data was generated to gain an understanding of the  wear couple at hand. NASA/CP—2018-219887 500
High-speed wear on hydrodynamic bearings is a unique phenomenon and can only occur when the load  capacity of a bearing is exceeded at high speed. This situation occurred on the cooling pumps of the  International Space Station (ISS) in 2010 and triggered the premature shutdown of one of the pumps.  Although in this situation the weight of the rotor was not present due to operation in a microgravity  environment, secondary forces aligned to overcome the non-contacting load capacity of the bearing leading  to severe and unexpected wear. The phenomenon of high-speed wear has not been studied in the  application to hydrodynamic bearings, and this phenomenon is unique to niche applications such as  microgravity operation. Wear mechanisms and material wear coefficients in this region require a test-based  approach to quantify the functional life of a bearing which experiences these conditions.     The Extreme Environment Tribometer (EET)  Design  A survey of commercial and custom designed research tribometers was conducted in order to identify a  test rig that could meet the needs of the current test program [14-20]. No such apparatus was identified,  which triggered a new tribometer design effort. The EET was designed with the following characteristics,  which were required to conduct high-fidelity wear coefficient measurements in conditions that matched  those in the ISS cooling pump bearings:  •Pin-on-disc configuration with modular running gear and test chamber capable of block-on-ring or hydrodynamic testing. •Capable of high rubbing speeds that spanned the front and rear journal bearing design space. •Flooded liquid operation at the rubbing contact points with liquids that are unfriendly to many materials, specifically anhydrous liquid ammonia. •Vacuum and elevated pressure capabilities to enable test operation with liquids that are not thermodynamically stable at standard temperature and pressure (STP). •Constant temperature operation during test operations. The final design of the EET is shown in Figures 3 and 4. The design consists of a pressure vessel that  houses a vertical, rotating shaft supported on dry, unlubricated rolling element bearings. The use of  unlubricated bearings was paramount to the successful creation of the wear coefficient database, since  contamination of the test fluid, anhydrous ammonia in this case, by the oils or greases in the rolling element  bearings would invalidate the data applicability to the ISS wear conditions. Additionally, the high pH of the Figure 2. The Stribeck Curve for the Hydrodynamic Bearing.  NASA/CP—2018-219887 501