ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 1. The Saturn V second stage moves over roads at KSC, using GSE hardware (Reference 1).  Figure 2. The Shuttle ET moves over the road at KSC (Reference 2).  However, the MPTS designed for SLS is different; its design is to accommodate multiple size articles. Figure  3 shows the initial concept of the MPTS where the intent is to support multiple hardware along with the  possibility to be used in future programs.   NASA/CP—2018-219887 460
Figure 3. The CS flight hardware is among the payloads that the MPTS will transport (Reference 3).  Current planned uses for the MPTS include the shipment of the SLS CS – NASA’s largest and heaviest  rocket – as well as Pathfinder – a simplified simulation of the CS that is similar in mass and length – Dynamic  Demonstration Unit, and three different STAs – each with their own distinctive mass and length. The various  payloads will also need to be transported to different locations, as shown in Figure 4. The MPTS will take  the three different STAs from Michoud Assembly Facility (MAF) in New Orleans, LA to Marshall Space  Flight Center (MSFC) in Huntsville, AL while the CS and Pathfinder will be shipped to Kennedy Space  Center (KSC) on Merritt Island, Florida via Stennis Space Center (SSC) in Hancock County, MS.  Figure 4. GSE is used for transportation at and to multiple centers, including KSC, MAF, MSFC, and SSC  (Reference 4).   The MPTS design has encountered some unique challenges. Primarily, with its use to support various size  payloads, a modular design was needed. Previous hardware, along with initial concepts, used a strongback,  NASA/CP—2018-219887 461
a structure connecting the front and back to form a single rigid unit. While including a strongback can  simplify design, it quickly became impractical for the MPTS design. Some reasons to eliminate a strongback  includes weight and maneuverability. Due to constraints of existing infrastructure such as roads and floor  loading, the design weight of the MPTS had to stay below certain limits, and the size necessary to produce  a rigid strongback structure to redirect loads away from the flight article were determined to exceed these  limits. Following Hurricane Katrina, a levee was added along the transport route to address flooding.  Surmounting this levee with something as long as the CS would have been difficult to accomplish with a  strongback structure, and the operation to accommodate the different lengths would be taxing.   The elimination of the strongback left the front and back end of the article free to operate independently.  To prevent an excessive load from occurring at payload interfaces, released degrees of freedom (DOF) are  included at strategic locations. This also serves to improve the predictability of loads at the interfaces.  Figure 5 shows released DOF at a high level.  Figure 5. The MPTS includes released DOF, to prevent excessive loads at payload interfaces.  In place of the strongback, a laser alignment guide on the transporter maintains the separation distance. A  monitoring system tracks relative motion of multiple released degrees of freedom. The combination of the  released degrees of freedom and monitoring system provides the means for the article to traverse  compound road contours, as illustrated in Figure 6. A remote-controlled drive system will shut down if set  travel limits are violated.   NASA/CP—2018-219887 462
Figure 6. The MPTS articulating over the contour of the road.  Design and Analysis Challenges  Designing the heaviest and largest rocket to date is not an easy task. Designing a transportation unit for  that rocket comes with its own challenges. With the increased size, the need for larger and stronger  hardware also increased. This shifts the selections of hardware, such as bearings for released degrees of  freedom or bolts in fastened joints, from commonly produced parts, to specially-sized items that are difficult  to procure.   Figure 7 shows an example of some catalog-available vendor hardware. The catalog lists a range of sizes  for a particular part. The larger load-rated items are in less demand, and therefore require longer lead time.  The increased size makes handling difficult, and special consideration is needed for even simple touch  tasks. The size also makes logistics a challenge; passing through doors that were adequately sized for  smaller articles designed in the past becomes a tight-fit operation. A balance between managing loads and  maintaining an optimum size was ultimately achieved.  Figure 7. The Crosby Group catalog of Cold-Tuff Heavy Lift Swivel Hoist rings (Reference 5).  NASA/CP—2018-219887 463
Following the egress from the facilities building, the next challenge is to travel on the road. Consider, for a  moment, the one-time transport operation of the shuttle Endeavour through the streets of Los Angeles from  the airport to the science center. Careful planning was performed in evaluating its travel route. Obstacles  were assessed and removed as needed along the travel path; adjustments were needed to avoid impacts.  Figure 8 shows a snapshot from a video of the shuttle as it travels from the airport to the science center.  Figure 8. Snapshot footage of shuttle Endeavor during its transport through Los Angeles (Reference 6).  A similar evaluation was performed in the design of the MPTS. Because this transport operation is meant  to be repeated for different articles, consideration of the terrain and any obstructions along its path must be  made. Factors include road grade, turns, signposts, overgrown flora, culverts, and bridges. Careful  consideration of the travel path was taken to ensure the MPTS, while transporting its cargo, is capable of  traversing the terrain without additional effort to modify the terrain or travel path.   This consideration led MPTS development to the challenge of evaluating the released degree of freedom  our transportation system provides. Numerous computer model simulations and dynamic analyses were  performed to ensure the range of motion provided by the system is adequate to support the path it needs  to travel. The simulations involved creating a 3D model of Saturn Boulevard at the Michoud Assembly  Facility in New Orleans, LA. The model was generated from data collected in a road survey, which includes  the grade created by the levee that was added after Hurricane Katrina. Traversing this levee would require  the unit to ascend and descend a steep grade, while making a turn and negotiating a bank in the road. This  all occurs while each end is individually managing the sectional contours of the road beneath it. The  evaluation was conducted using MSC Adams to ensure there is sufficient range to prevent the movements  from bottoming out and to avoid unintentional loading of the article being transported. This simulation  includes a flexible finite element model of our GSE hardware and the various payloads it supports, which  is connected by joint mechanisms. The flexible body also includes contact definition between mating  surfaces and accounts for dimensional tolerances that could develop in production. The maximum design  range of motion also required consideration. Excessive motion would result in a larger-than-necessary  structure that would create ingress/egress problems of wheel loading, and weight concerns on the road,  and over bridges and culverts. The maximum range of motion also necessitates evaluation from a loading  standpoint in order to ensure the shifting load does not overload a section of the hardware or article. With  the MPTS designed to support multiple sized articles, the evaluation was conducted on each unit to ensure  all the articles would be safely transported by a single MPTS design.  NASA/CP—2018-219887 464
Most importantly, the objective is to ensure the design does not impart loads that could potentially damage  the flight hardware during transport. The challenge stems from requiring the MPTS to be designed to the  standard factors of safety for ground equipment, but a lower factor of safety is used in the design of the  launch vehicle as a means of lowering the weight. Added to the complexity is how the loads are applied to  the vehicle. The design of flight hardware typically considers the launch loads as the driving load case.  However, the vehicle is typically transported horizontally. The loads induced by the MPTS are imparted in  a different orientation during the transportation loads as compared to the launch environment.   The vehicle design seeks to minimize the need for local reinforcements. Early discussion with the flight  vehicle engineer allowed us to determine the interface definition between the article and the MPTS. This is  to ensure suitable locations are provided on the flight hardware capable of supporting the transport load  without the need to add weight for localized reinforcements. The collaborative effort includes the  identification of interface locations along with the number and size of the contact area. Attachment brackets  are designed to connect to these interfaces. This minimizes the weight on the flight article, but also adds  the challenge of designing brackets that are light enough to handle, or are situated in a manner allowing  the use of support equipment. On-site installation also makes alignment and tolerance control difficult,  particularly when attempting to align multiple parts as shown in Figure 9. With various articles requiring  transport, the MPTS is designed to accommodate the different tolerances associated with aligning the  multiple attachment locations.  Figure 9. Illustrative depiction of multiple interface locations. Lessons Learned  Additive Manufacturing  Additive manufacturing processes, commonly referred to as 3D printing, were used to generate tangible  models of the MPTS. This offered a unique and straightforward approach to replicate plastic models of  individual components of the SLS transportation system. A picture of a model generated in the preliminary  phase of MPTS development is shown in Figure 10.   NASA/CP—2018-219887 465
Figure 10. Preliminary conceptual additive manufacturing model.  Creating models generated through additive manufacturing was an expedient way to understand and  demonstrate the interdependency of MPTS mechanisms. Released degrees of freedom at various locations  of the MPTS affect both local and global movement of the combined payload and transportation system.  The relative motion of the system was able to be simulated real time with a scaled down version of ground  support equipment. The model shown in Figure 11 demonstrates MPTS relative motion.  Figure 11. MPTS Additive Manufacturing Model (Top View)  NASA/CP—2018-219887 466
Figure 12. Vendor-reported coefficient of friction in relation to bearing pressure loads (Reference 7). The coefficient of friction documented by the vendor is based on this condition. An evaluation of our design  indicated our bearing pressure on the liner was below what was needed to achieve the desired friction. Our  worst case loading conditions, which includes higher level wind loads in combination with the worst case  operating, was considered in the selection. However, under normal operating conditions, the bearing  pressure is significantly less and resulted in a negative impact to our anticipated friction loads.   Advertised properties provided by vendors should be investigated early in the design development process.  Meanwhile, hardware strength evaluations should be conducted using enveloping values with additional  margin until the stated properties can be verified. With the understanding of the vendor’s required condition,  the MPTS was able to be modified, and the friction load was verified to be within acceptable range after the  changes were made.  Managing Interfaces and Parallel Development Efforts  To streamline the overall process, MPTS and payload development schedules included overlap. That is,  the GSE design began before the payload designs were finalized. This reduced risk for the overall program  because it allowed each side of the interface to take the other into account. If the GSE were developed  after the payload design was completed, the payload design runs the risk of not having an interface with  adequate strength to be transported, necessitating costly redesign effort. While concurrent development is  desirable, two key challenges arose: how to address mismatched Factors of Safety (FOS) requirements  and how to manage design changes in the payload.   GSE is typically required to meet higher FOS than the payload, and this was the case for the MPTS and its  payloads. Early in the design phase, questions arose regarding how to enable each side of the interface to  meet its requirements. GSE loads can and do drive flight hardware design , albeit at a local level in most  cases.   Mismatches were addressed in the following manner: for interfaces where the payload added an interface  solely to provide an attachment for the GSE, the interface was designed to allow the GSE to meet its  required factors of safety. These interfaces were determined by the payload in order to minimize the impact  to the payload. For interfaces where the interface was also used for flight, a less desirable situation, the  interface design still attempted to allow the GSE to meet its required factors of safety. However, in these  situations, there was often less flexibility to accommodate the GSE, and waivers to the GSE requirements  often came into discussion. In situations where reduced FOS were considered, proof load planning became  NASA/CP—2018-219887 468
NASA Docking System Block 1: NASA’s New Direct Electric Docking System  Supporting ISS and Future Human Space Exploration  Justin McFatter*, Karl Keiser** and Timothy Rupp+ Abstract  The NASA Docking System (NDS) Block 1, shown in Figure 1, is a key component of NASA’s vision for  space exploration. It is designed to provide capability for visiting vehicles to dock to the International Space  Station’s recently-installed International Docking Adapter ports. It is the first docking system to be  developed by NASA since the Apollo-Soyuz Test Project of the 1970’s. The NDS Block 1 includes  provisions for capture, structural attachment, power/data transfer, and undocking. It uses a direct-drive  electromechanical Stewart Platform capture system architecture, along with an innovative automated  control scheme to achieve an unprecedented level of performance and simplicity. Its design implements  the new International Docking System Standard, which will be a key enabler of diverse and flexible  exploration missions. NDS qualification was completed in 2017 to support a planned first flight in 2018 on  the Boeing CST-100 Starliner.  Figure 1. NASA Docking System Block 1  *The Boeing Company, Houston, TX ** The Boeing Company, Huntsville, AL +NASA Johnson Space Center, Houston, TX Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018 NASA/CP—2018-219887 471
orientation of the motion platform using the measured lengths of the six actuators. The lengths of the  actuators can then be adjusted to accomplish the desired motion of the platform. In these ground-based  systems, the calculation of the platform position and orientation, known as the forward kinematic solution,  must be performed numerically using a processor-based computer due to the lack of a closed form solution.  During the early phases of NDS development, the cost and schedule impact of developing a processorbased controller certified for the space environment were identified as a major program risk for architectures  that required it. This risk was one of the key discriminators that led to the selection of a SCS architecture  that does not require the forward kinematic solution.  Figure 4. Torque Motor Servo System from Apollo Soyuz study  The arrangement of the NDS linear actuators is such that the soft capture ring is coaxially aligned with the  tunnel whenever the six actuator lengths are the same. The NDS control architecture makes use of this fact  to simplify the commanding of the actuators during capture and ring alignment. During capture, the  controller calculates the average length of the six actuators at every instant of time and uses it as a point  of reference for commanding current to the motors. This logic gives the soft capture ring a central tendency  as it complies to accommodate vehicle misalignments. The simple coordination algorithms used in the SCS  allow for an implementation that does not require a processor-based controller. During soft capture  operations, the NDS relies solely on position and current feedback to control the six linear actuators as they  bring the vehicles together and engage the structural mating features of the two tunnels. Major components  of the NDS are shown in Figure 5.   Operational Sequence  The docking event is an automated sequence punctuated by several discrete events. During the approach,  the SCS is extended to the “ready to capture” position, where it awaits initial contact between the two  spacecraft. The GN&C system of the chaser vehicle ensures that initial contact occurs within the capture  envelope of the docking system. From initial contact onward, the chaser vehicle is in free drift and relies on  the docking system to capture and maneuver the vehicles to hard mate. This approach distinguishes NDS  from previous docking mechanisms, which all relied on the host vehicle thrusters to push the soft capture  features together after contact to achieve capture. The design of the NDS SCS has eliminated this need.  Qualification testing of the SCS in the summer of 2016 demonstrated 100% capture success for a robust  set of initial contact conditions, without the use of post-contact thrust.  NASA/CP—2018-219887 475
Figure 5. NASA Docking System Schematic  At contact, the forces between the coarse alignment guides cause the NDS linear actuators to be displaced  from the ready to capture position. This event signals the controller to initiate a maneuver known as the  Lunge. During Lunge, the soft capture system extends in a compliant manner to accommodate vehicle  misalignments and engage the soft capture latches, one of which is mounted in each of the three guide  petals on the soft capture ring. Each latch engages an opposing striker plate mounted behind the passive  side soft capture ring. The interface between the capture latch and the striker is shown in Figure 6.  Once the capture latches are engaged, sensors mounted in the soft capture ring indicate that capture is  achieved, signaling the system to begin attenuating the relative motion of the two vehicles. At this point, the  two soft capture rings are loosely held together by the capture latches such that the incoming vehicle cannot  escape. Once the time allocated for attenuation expires, the controller begins adjusting the lengths of the  actuators to equalize them, thereby aligning the two mating tunnels. With alignment complete, the linear  actuators retract, closing the gap between the two tunnels and engaging their shear features.  Once the two structural tunnels are in close proximity, a set of twelve structural hooks are driven, which  engage corresponding compliant passive hooks on the opposing tunnel. The structural hooks close the  remaining gap between the tunnels, compressing an elastomeric seal and creating an airtight structural  interface. With the hard part finished, motorized umbilical connectors can then be driven, allowing for power  and data transfer between the docked spacecraft. The major hard capture interface components are shown  in Figure 7.  NASA/CP—2018-219887 476
Figure 8. NDS 6DoF Docking Test  Table 1. NDSB1 Initial Contact Condition Limits  A test case used to validate the docking simulation was the case of an 18 t host vehicle docking to the ISS  Node 2 Forward docking port with an approach velocity of 4.5 cm/s and a Yaw misalignment of -5 deg about  the nadir axis. At initial contact the NDS is tilted towards Petal 1, located in the starboard direction of the  chaser vehicle, as shown in Figure 9. Initial contact occurs between the capture latch on Petal 1 and the  corresponding striker on the passive. Contact is detected when the SCS ring at the base of the petal  contacts the passive SCS ring. Struts 1 and 2, which are attached to Petal 1, are then driven backwards  during Lunge due to the contact at the SCS ring. During this time, the Lunge control law causes the other  four SCS struts to extend and engage the latches on Petals 2 and 3. The time history of strut lengths is  shown in Figure 10.   Initial Conditions Limiting Value Closing (axial) rate 3 to 6 cm/s Lateral (radial) rate up to 4 cm/s Angular rateup to 0.2 deg/s about closing axis up to 0.2 deg/s about any lateral axis Lateral (radial) misalignment up to 11 cm Angular misalignmentup to 5 deg about closing axis up to 5 deg about any lateral axis NASA/CP—2018-219887 479
Figure 9. Initial Contact at -5 deg Yaw  After the capture sensors are activated, the SCS transitions into Attenuate, where the actuators counter the  momentum of the chaser vehicle while limiting force output via current limiting. After the initial high load  event in Attenuate, the motion of the chaser is essentially arrested, with only small motion remaining within  the space between capture latches and the strikers. The measured and predicted motor currents for the   -5 deg Yaw case are shown in Figure 11. For the test program as a whole, the results showed good agreement with analytical predictions. The  docking simulation parameters were adjusted based on test results to match the performance observed in  the single-axis extreme misalignment cases. After completion of the model correlation activity, the results  of the randomly distributed contact cases were compared to the model. The results of this comparison,  shown in Figure 12, indicate good agreement between test results and predicted system performance.  To demonstrate safety and robustness, the system was tested with a number of simulated failure conditions  including missed captures, single string avionics failures, and failure of a single capture latch. In all cases,  the system was able to recover safely either by bouncing off without collision or by safely terminating the  docking attempt. To round out the test program, the flexibility of the system for other mission scenarios was  demonstrated by adjusting the control law gains to optimize for lighter vehicles. With this alternate  configuration, the system successfully docked a 15 t chaser vehicle to a 5 t payload with single-axis  misalignments of 5 deg and 11 cm. The ability to support light vehicles is key to exploration missions that  may involve docking a crew capsule to a Mars or Lunar surface lander or other small payload.  NASA/CP—2018-219887 480
Figure 12. 6DoF Test Results  Lessons Learned during 6DoF Testing  Successful engagement of the soft capture latches is indicated by the activation of a set of capture sensors  embedded in the active soft capture ring at the base of the coarse alignment petals. Each capture sensor  is activated by a simple spring-plunger mechanism that is depressed by contact with the opposing soft  capture ring. The arrangement of these sensors is shown in Figure 13.    Figure 13. Soft Capture Sensors  Both the button and shaft of the capture sensor are coated with a polytetrafluoroethylene (PTFE)  impregnated hard anodize. During development-level 6DoF testing in 2014, capture sensors began seizing  intermittently after approximately 400 docking cycles. This condition prevented the capture latches from  successfully engaging the strikers on the passive test article. The seizing was caused by damage to the  anodize coating on the button and shaft wear surfaces of the capture sensor, due to repeated impact against  the passive soft capture ring. Figure 14 shows the condition of the capture sensor plunger and housing at  NASA/CP—2018-219887 482