ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
484   Figure 2: The SpaceX Docking System  conducting a joint interface acceptanc e test; soft capture ring  shown in the deployed state 

The SCS is deployed forward of Crew Dragon via a hexapod –  a minimally constrained, six -legged  arrangement of two- force struts. As the soft capture ring/petal elements make contact with the IDA, forces  travel down the rigid struts and are converted into torques by a moment arm connected to a passive rotary  spring damper , also referred to as attenuator arm. The rotary spring damper reacts the contact torque with  caged helical compression springs and a rotary eddy current damper. The soft capture system passively   attenuates the contact loads and allows the soft capture ring to move in six degrees of freedom (6- DoF),  permitting it to align with the passive IDA interface. Assuming sufficient alignment is achieved, the soft  capture latch pawls jump over corresponding  striker plates on the IDA and soft capture is achieved. Two  plungers on each soft capture petals  are depressed when capture is achieved, triggering a change in  microswitch state. If capture is not achieved, Crew Dragon will perform its failed capture response and  safely retreat from the ISS .   

Following successful soft capture and cessation of relative motion, the soft capture ring is retracted via  three linear dual -wound actuators. As the linear actuators are driven up they cause the rotary spring  damper s to rotate down and thereby return the soft capture ring to the stowed height. Following soft capture  ring retraction, twelve hard capture hooks , derived from the NASA Docking System  hook design,  connect  to corresponding passive hooks on the IDA. The hard capture hooks each contain a cam mechanism that  draws the hooks down and compresses a seal that runs around the circumference of the docking system.  Ready to hook switches provide the indication that the hooks are in a position to successfully latch onto the  IDA passive hooks. Finally, two umbilical dual-wound actuators mate power and data connectors to ISS.  Following a successful hard mate operation, the vestibule between Dragon and ISS is pressurized and the  hatches opened to allow for transfer of crew and cargo.  

To undock, the passive capture latches are moded to a “ready to release” state  such that they no longer  provide load retention capability . The umbilicals are retracted and the hard capture hooks opened, freeing  Dragon from the ISS and allowing GNC to take control to move Dragon away from the ISS . Contingency  release mechanisms, Non- Explosive  Actuators (NEAs) on the soft capture latches and F rangibolts  on the  hard capture hooks, are present in the event that release via the nominal actuators fails .   

Notably, the hard capture hooks are also utilized to retain the nosecone during launch and re- entry. This  protects docking system and other nearby components for reuse.  

The SpaceX docking system is a highly performing and efficient spacecraft mechanism. Its weight is  approximately  4% of the Dragon 2 mass. It has low power draw and  was designed and qualified for fiveflight reusability with minimal maintenance.
485 The SpaceX docking system is a well understood mechanism, which has been computer modeled,  simulated, and test correlated to a high degree. The system has been modeled in NASA’s Trick software  framework, so that this simulation can easily interface with other NASA docking simulations . At many levels  of individual components (springs/dampers), sub- systems (soft capture latches, hard capture hooks, rotary  attenuator arms), and full docking syst ems, the simulation has been correlated to match test data from both  SpaceX and NASA testing. The correlation of the simulation extends to a wide variety of mass properties  in the Crew Dragon class, at a wide range of contact conditions and velocities.  

Figure 3:  The SpaceX Docking System Operations Overview  

Figure 4: SpaceX Docking System Subassembly Overview
487  

Figure 5:   Attenuator Arm ATP Correlation, Test Flight  Example Arm Data vs. Simulation  

Sub-System Lessons Learned – Latch Acceptance Testing   The soft capture passive latch model was derived from a high- fidelity dynamics model, based upon CAD  geometry of the latch hardware. The Trick model of the latch did not attempt to simulate every possible  degree of freedom but used lookup table data to emulate the nominal behavior of the latch when  compressing during docking. The latch model included effects due to the unique geometry of the striker,  which is not smooth, and produces force spikes during nominal docking. These effects can be seen below  in Figure 6. An important lesson learned was the effect that the significant striker geometry had on l atching  dynamics, as well as the overall complexity of the latch design and dynamics.  

Figure 6:  Latch Forces vs. Time, Latch ATP Test data and Simulation data  

System Modeling, Correlation, and Lessons  Learned  

System Modeling and Correlation   The Sub- System ATP testing was critical to successfully modeling the SpaceX Docking System behavior  when tested at the SDTS facility at JSC . The Six Degree of Freedom Test System facility at NASA’s  Johnson Space Center enables  soft capture, attenuation, and retraction testing of docking systems  in a  controlled environment .
488 Correlation of the simulation to SDTS testing could only be accomplished due to integration of the simulation  with JSC’s simulation of the facility . The facility has several nuances, including controller and actuator  effects, which when accounted for, vastly improve the accuracy of the docking simulation. A comparison of  the measured forces at the docking plane and the relative vehicle motion, for a representative docking  condition, is s hown in Figure 7.  

Figure 7:  Simulation correlation to SDTS test data, Forces and Displacements vs. Time  

Two system test campaigns were performed on the docking system, an initial campaign in 2015 and a  second campaign in 2018 with the final flight de sign iteration  implemented. During the first campaign, we  learned that the sub- system modeling of the attenuator arms and latches were inadequate. While they could  capture the overall behavior of the docking system, there were too many cases which did not correlate  properly, many of which were directly due to properly modeling the non- linear behaviors of the attenuator  arms . There were also missed captures, which were directly due to improper tolerances of both the latch  and striker hardware, but this promp ted the creation of latch models of higher fidelity in order to understand  the latch behavior.  

Several lessons learned came from the model correlation during the second test campaign . Significant ly  more sub- scale testing was performed on docking system components, which allowed exact spring rates,  preloads and damping rates to be modeled and thus leaving only system level variables as unknowns. This  higher fidelity modeling of the attenuator arms resulted in much better correlation to a wide verity of test  cases, including those that specifically did not correlate as well during the first test campaign. A major  lesson learned was that misalignments of the Soft Capture System ring at deployment do not  drive docking  dynamics . Due to the passive nature of the docking system, and the relatively small pitch and yaw inertias  of the Dragon 2, significant misalignments of the soft capture r ing at deployment were negligible and did  not have a strong impact on capture success . This highlights that there is little benefit to high levels of  accuracy in SCS ring position, which can drive complexity for systems with Linear Electro- Mechanical  Actuators, at the kinetic energies seen in Dragon 2 docking.  

A critical finding from testing was that the capture switches could be triggered or cycled  while the  mechanism was captured, under heavy shear/bending load. This effect could be replicated with the  simulation,  after updates to the simulation using a more advanced capture switch model . The effect results  from an interaction between the specified free play  between the latches and strikers, in addition to the  location of the capture switches, nominally on the edge of the radius of the passive soft capture system   ring. This  led to an increase in a flight software persistence value, specifically the number of consecutive
489 control cycles where the capture switch is not pressed to initiate the failed capture response, to account for  this affect and prevent false negatives during the mission.      The high accuracy of the simulation meant that properly modeling the gravity off -loader became significant .  To better replicate the on- orbit dynamics, the SCS ring has an off -loader which reacts the force of gravity  from the SCS ring. Howev er, this somewhat simple device can have significant frictions and stictions which  become difficult to model for each unique run. A lesson learned from SDTS testing is to improve the  repeatability of the gravity off -loader and fully characterize its behavi or to aid in model correlation.  

Flight  Reconstruction   The Demo 1 test flight mission was reconstructed with the docking simulation  and the flight telemetry. The  simulation had good correlation to the flight  contact and attenuation dynamics, as well as the retraction and  hard capture dynamics . Figure 8 compares the Flight  Telemetry of the Attenuator Arm Encoders to the  simulated encoders for capture.  

Figure 8:  Flight Attenuator Arm Telemetry vs. Simulation vs. Time  

For undocking, there were two principal sources used for correlation, the Ready to Hook switches, and the  calculated Dragon velocity . The ready to hook switch states were used to verify the hook dynamics and  overall Dragon separation state. The calculated Dragon velocity was used to ve rify the forces separating  the Dragon from the ISS . Notably, there appeared to be little to no seal stiction (expected given the se al  lubrication), and the latches  provided nearly no restraining  force, resulting in significant departure velocity  when the hard capture hooks cleared. Properly modeling the hooks and their motion, in order to determine  when there is clearance for departure, is a critical lesson to modeling undocking.  

Qualification Testing  Failures and Lessons Learned  

There were many lessons learned across the various test campaigns, several of which manifested as test  failures . These early testing lessons learned  led to design updates and delta qualification testing before the  successful Demo 1 flight test  and ultimately a more reliable system . A subset of these failures  and their  corrective actions are outlined below.
490   Figure 9: (Left): The SpaceX undergoes dynamic testing on the NAS A JSC Six -Degree- of-Freedom  Dynamic Test System. (Right): The SpaceX Docking System prepared to enter thermal vacuum testing to  perform system level qualification tests.  

TVAC Harnessing Failure   During system level thermal vacuum testing, there was a loss of signal from the controllers  during the soft  capture deploy sequence.  To troubleshoot, the vacuum chamber was opened,  and it was determined that  the connectors to the soft capture latch controllers had sheared. Subscale testing was performed,  and it  was determined that the ov erwrap material of the harness, a rubber -like material , was stiffening the harness  in vacuum, which caused it to shear. By replacing this overwrap material  with a more flexible  material , this  condition was eliminated. This provides a valuable lesson learned in harnessing material selection,  especially in dynamic mechanisms  operating at a wide temperature range.  

Figure 10: Harness Connector Sheared during Deployment
492 failure; CT scann ing of both vendor and internal  assembled components can be utilized as a very  informative and robust quality control tool .     

Figure 12: Dislodged Retaining Ring  

Harness Failure   During an acceptance test, an intermittent connection was noted on one of the power and data buses. The  problem was isolated to one connection. The controller was removed and tested and no anomalies noted,  which led to additional investigation of the harnessing. It was found that ‘wiggling’ the harness could cause  the bus to disconnect . The harness was CT scanned to identify the root cause: improper insulation on one  wire allowed the cabling to ‘bird cage’, which could then short against the connector . The insulation was  not properly adhering to the cable, which would allow the insulation to pull off the cable and induce this  condition.  CT scanning provides extremely  clear indications of this root cause. This issue could have been  easily written off as a single workmanship issue; however, through CT scan it was  identified that this was  not an isolated incident but cable adhesion suscept ibility;  and through further testing the faulty cable lots   were identified and contained. This is another example where CT scanning can be a valuable investigation  tool when failures occur.  

     Figure 13: CT Scan images showing ‘bird caging’ of cable  which had its insulation improperly adhered  and pulled up. This allowed the cable to short against the connector housing
493 Lessons Learned: Systems Interactions  

SpaceX benefits from having a highly integrated team of software, avionics, operations, GNC and  mechanisms engineers. This integration is critical to the successful design and continued operation of the  SpaceX docking system. The close interactions between flight software and the mechanical team, for  example, allowed for  very robust testing of both hardware and software: the flight software algorithms, fault  detection methods, voting schemes and responses were used for all system level testing. This allowed  early validation that the software responded as expected, includi ng in off -nominal cases . This close  interaction also allowed validation of configurable parameters, and the accurate modeling of flight software  response in simulation allowed small changes in these parameters to be validated. For example, a key  parameter in the docking configuration is the persistence value to declare soft capture and loss of capture  from the soft capture switches . By using flight software during 6- DoF testing at JSC, and then adjusting this  configurable value in the docking simulation, this parameter could be fine -tuned for the desired response .  The close interaction between the simulation and GNC allowed rapid iteration based on expected contact  conditions and docking parameters.  

The integration between avionics teams and hardware teams was also critical to the rapid and successful  development of the SpaceX docking system. Controllers were developed specifically for the docking system  application, and close coordination during subassembly and system level testing allowed for the fine- tuning  of controller firmware. With a dynamic system, harness routing to the soft capture latches and switches is  critical; a misrouted harness could impede motion and have drastic impacts on mechanism success. Close  coordination with harnessing engineers, inc luding involving them in system level testing so they could better  understand the expected range of motion, allowed for a successful harness design.  

Conclusions  

The SpaceX  docking system is a custom implementation of the N ASA Docking System architecture . It is  the first American designed and manufactured docking system to be flight proven in decades . The SpaceX  docking system leverages a semi -passive system to effect docking capture, attenuation, and alignment  before retraction to hard captur e. It is effective for a broad range of contact conditions, is low power , light  weight, and cost effective.  The system has been extensively  simulated,  tested, and inspected to ensure  successful operation throughout a wide range of possible contact conditions.   

Figure 13: The Demo 1 Crew Dragon Capsule Docked to the ISS Forward IDA in March 2019.
494 References  

1. McFatter, J., K. Keiser, and T. Rupp , “NASA Docking System Block 1: NASA’s New Direct Electric  Docking System Supporting ISS and Future Human Space Exploration” Proceedings  of the 44th  Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16- 18, 2018  

2. NASA Docking System Interface Definitions Document (IDD), JSC 65795, Revision J
500  Approved for public release. OTR 2020- 00294    Figure 3: Typical MMA test sequence at unit level  

Torque/Force Margin Approach   In this discussion , references to “torque margin” also apply to linear motion, with “force” simply replacing  “torque.”    

This subject may be the biggest change in the new MMA standard, and one that will likely take the most  adjustment with  respect to the traditional approach for experienced MMA engineers . This new approach to  calculating torque margin was first widely discussed around the time of the previous AIAA MMA standard  release, but the aforementioned AIAA deadline prevented any further committee discussions on the matter.  Furthermore, this approach has already been published in its basic form in NASA’s MMA standard,  Reference 5, but there were some modifications and accommo dations made in this  AIAA MMA Standard  update.  

The basic premise was that there are at least three different types of resistive forces that MMAs encounter  – those that are relatively predic table and repeatable , those that are  variable and  harder to predict initially  or over life, and those that are purely the result of an induced acceleration. Examples of the first kind, or  “fixed” resistances, are a return or fail -safe spring, fluid pressure in a valve, etc . Examples of the second  kind, or “variable” resistances,  are coulomb fri ction, cable harness bending stiffness, etc . Examples of the  third kind, or “acceleration- depende nt” resistances, are those produced by rotary motion, launch vehicle  accel eration, etc.  

In the new torque margin calculation approach, which resembles the method in which stress or mass  margins are calculated, each resistive torque carries with it a different uncertainty factor. Traditionally, with  a torque margin requirement of 100%, all resistive torques were assigned an uncertainty factor of 2.0. In  the new  approach, the uncertainty factor is a function of the “variable” or “fixed” nature as described above,  as well as  of the method by which it s value is obtained, such as by analysis, testing on an engineering  model or qualification article , or testing on an  actual flight unit.  Figure 4 shows the minimum torque/force  margin factors used in this new approach.
501  Approved for public release. OTR 2020- 00294    Figure 4: Minimum Torque/Force Margin Factors  

Before discussing the new calculation method and equation(s), it is also worthwhile to mention that there  are now five different types of torque/force margins  identified , with the requirements of each needing to be  satisfied as appropriate for a given MMA or subsystem:   • Static Torque/Force Margin –  a measure of the excess torque available to overcome resi stance to  impending motion   • Constant Velocity Torque/Force Margin – a measure of the excess torque available to maintain  motion   • Holding Torque/Force Margin – a measure of the excess torque available to maintain position in  the presence of external disturbances   • Dynamic Torque/Force Margin –  a measure of the excess torque available to accelerate a body by  a given amount   • Stepper Motor Margin –  a measure of the excess pull -in torque at the drive rate available to  overcome friction loads seen at the motor  

The ne w single  equation used to calculate the torque/force margin is shown in Equation 1:  

   Torque  Margin  =Τavail  ΣΚfixΤfix +ΣΚvarΤvar+ΣΚ accΤacc −1                         (1)  where  - Tavail is the minimum available torque generated by the driving or holding component (e.g., spring,  motor).   - Tfix terms are the individual maximum “fixed” resisting torques that are not strongly influenced by  effects of friction, temperature, cycles, etc. (e.g.,  motor detent torque, vehicle maneuver -induced  torque, return spring torque, unbalanced pressure load limited by relief mechanisms).   - Tvar terms are the individual maximum “variable” resisting torques whose values may change with  environmental conditions a nd cycles (e.g., friction torque, wire harness torque due to flexing or long  term set).    - Tacc is the torque required to achieve the specified acceleration of the driven component.   - Kfix, Kvar, and K acc are the fixed, variable, and acceleration torque/force margin factors applied to  each individual resisting torque/force term in Equation 1 prior to summation.  

As mentioned earlier, torque/force margins are now calculated in the same manner as stress, strength, and  mass margins . As such, the requirement for st atic torque margin is no longer 100%, but that it be positive   (> 0) , after the appropriate uncertainty factors are applied. All terms in Equation 1 may not be used in all  MMA applications, and further guidance for each type of margin is provided in the standard.