ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
469 Design and Test of the Orion Crew Module Side Hatch  

Lance Lininger* and Kyle Gotthelf**  

Abstract  

The Orion spacecraft is a critical component of NASA’s Artemis program to establish a permanent human  presence on the lunar surface and further enable future crew ed missions to Mars. As a result of traumatic  lessons learned from the Apollo and space s huttle programs, NASA has established thorough crew safety  requirements that were particularly challenging to implement for the Orion Crew Module side hatch.  This  paper describes the background and evolution of the side hatch design, the features used to address the  crew safety requirements, the testing challenges in preparation for the Artemis  1 mission, and lessons  learned.    

Introduction  

Due to four decades of hard- learned crew safety lessons (Apollo 1 command module fire, space s huttle  Challenger, and space s huttle Columbia), NASA established requirements to make Orion the safest and  most reliable human- rated spacecraft ever built . However,  these re quirements  when coupled with the long  Orion mission durations, introduced some significant challenges with the side hatch design (see Figure 1  for the location of the side hatch on the Orion Crew Module) .   

Figure 1.   Orion Crew Module  

Specific side hatch changes from Apollo/ space s huttle include (also see Table 1):   • Introduction of the NASA -STD-5017 mechanisms design standard for crew ed missions (based on  AIAA- S-114)  • More conservative structural/force/torque margins   • Implementation of redundancies (where possible)   • Higher factor on mechanism life cycle verifications (4X for crewed missions)   • Longer mission duration   • Higher cabin pressure   • Lower leak rate  

* Lockheed Martin Space, Littleton,  CO;  lance.r.lininger@lmco.com   ** Honeybee Robotics, Longmont, CO  

Proceedings of the 45th Aerospace Mechanisms Symposium, NASA Johnson Space Center, 2020
470 Table 1. Side Hatch Key Driving Requirements Comparison, Apollo vs. O rion  Requirement  Apollo  Orion   Cabin Pressure  44.8 kPa 106.9 kPA   Max Leak Rate  ~2.3 kg / day  (estimated)  0.68 kg / day  Crew member size/strength  Average M ale range  1% female to 99% Male   Seal Redundancy  Single seal  Redundant Seals   Hatch Size  0.64 m2 0.79 m2  Hatch Opening Cycle Life  Records uncertain  1180  cycles   Mission Duration  14 days  > 21 days   NASA -STD-5017 (AIAA -S-114  derivative) applicability  Didn’t exist  Yes 

As a starting point, t rade studies were conducted to identify the design solutions that were most likely to  meet  these requirements.  Past crew hatch designs were examined from Gemini, Apollo, space s huttle, and  ISS. After numerous reviews and discussions, it was determined that a highly modified version of the Apollo  hatch was best suited for Orion. Figure s 2 and 3 show the Apollo and Orion side hatch designs with key  components.   

Figure 2.  Apollo Side Hatch
471   Figure 3.  Orion Artemis 1 Side Hatch   

Orion Side Hatch Design Details  

The side hatch contains the following key components:   • Aluminum structural panel : An initial trade was performed on a flat vs. curved structure (Apollo  used a curved structure). Due to the more stringent sealing requirements, it was decided that a f lat  hatch structure would be used on Orion (to improve sealing tolerances and reduce cost).  It was  also theorized at the time this would simplify the latch train kinematics (latches wouldn’t have to  transmit loads over a three- dimensional  curved surface), although this turned out to be a negative  aspect of the design later in development testing.   • Latches:  Two latch trains ( “A” train and “B” train, 17 latches total) are used to compress the  perimeter seals and retain the hatch to the spacecraft structure. Due to increased strength margins  and higher pressures required on Orion, the latches were scaled- up significantly  over Apollo (see  Figure 4). The latches are over -center mechanisms , linked together through a series of linkages  and bell cranks . Ideally, as hatch pressure loads increase this should drive the latches f arther over  center  into the hardstops (see Figure 5). Rigging of the latches is critical. I n a perfectly rigged hatch,  all the latches would be overcenter and contacting the hardstops simultaneously . However , even  with designed- in adjustable rigging features, it is extremely difficult to achieve the desired hardstop  conditions due to the many part tolerances and the interactive nature of the latch linkages.
473 • Gearbox:  Honeybee Robot ics m ulti-stage gear system used as the interface for both ground and  flight crews to actuate the latch trains  from inside and outside the Crew Module (Figure 7). A key  gearbox feature is the Latch Restraint Arm (LRA), which is a spring- loaded structural element that  engages the root of both latch trains when the latches are in the final over -center position ( Figure 8).  This acts as a redundant structural load path to prevent inadvertent unlatching.  The gearbox  include s an LRA switch that arms  and disarms  the spring mechanism which places  the LRA into  the latched position (see Figure 9  for the original switch design used on the side hatch development  unit).  The Arm/D isarm Switch is meant to serve two primary functions. One of the needed functions  is to ‘disarm’ the LRA from the interior of the Crew Cabin so the astronauts can open the side hatch  via the Interior Latch Actuation Handle (ILAH). Disarming the LRA is done by rotating the switch in  the counterclockwise  direction. The act of disarming the LRA can also be achieved from the exterior  of the Hatch via ground crew actuating the Exterior Latch Actuation Mechanism (ELAM).  The other  primary function of the switch is  to ‘arm’ the LRA by rotating the switch in the clockwise  direction.  This is meant to be performed prior to closing the hatch. Therefore, the act of arming the  mechanism is meant to be an ‘action’ and not a ‘state’. Arming the switch will allow the LRA to  ‘engage’ and lock the latch train but this will only happen when the latches are actuated to the fully  closed position.  When the LRA engages it is meant to ‘kick’ the switch back to its Neutral position.  

Figure 7. Gearbox Overview
475 • Counterbalance assembly:  Moog Inc. hydraulic/pneumatic actuator which provides the motive  force to open the hatch in both normal and emergency egress conditions  once the hatch latch trains  are disengaged (Figure 10). 

Figure 10. Counterbalance Assembly  

Initial Design and Analysis  

The Artemis 1 mission (which is uncrewed) did not originally include a  full fidelity side hatch but rather a  structural simulator  for the flight . It was decided late in the Artemis 1 design s chedule to add a full fidelity  side hatch to reduce technical risk prior to Artemis 2 (first crewed mission) . This decision accelerated the  side hatch schedule by approximately one year which introduced obvious programmatic challenges. Due  to these schedule constraints the design team had to be judicious on the level of detail in the Finite Element  Model ( FEM ) and kinematic analyses. In order to meet the Critical Design Review and the development  testing milestones, t he initial analysis of the hatch design was limited to standard FEM/stress modeling and  rigid-body kinematic studies of the latch trains which all showed positive results with significant margins.  There was a plan to do a mor e in-depth flex -body analysis to examine the complete interactions of the  structure and latch trains, h owever this couldn’t be completed  in time. The CDR was held successfully (with  many independent NASA and LM reviewers) and so after n umerous meetings wit h management , it was  agreed the team should proceed into development testing to  validate the design.   

Development Testing  

As typical for a mechanical system of this complexity, the team started tests at a subcomponent level.  Numerous structural and life tests were conducted on the latch mechanisms, including individual latches  and a small connected set of 4 latches. The static load test of a single latch is summarized in Figure 11.  This test was intended to be an ultimate load test, and the actual capability was at least  2X the anticipated  strength (the test was stopped due to limitations  of the test fixture) . This gave the team confidence that the  latches were very structurally robust to react the pressure loads and would easily meet the more stringent  NASA structural requirements.  The test also validated the over -center action of the latch (at an individual  latch level).
476   Figure 11. Individual Latch Pressure Simulation Static Load Test   After various subcomponent tests, a full -size flight -like development hatch was built and a development test  program was implemented. The initial plan for development testing included:   • Functional testing   • Proof -pressure testing (at 1.5 x nominal mission pressure)   • Vibration testing (at pressure)   • Thermal cycle testing  

The initial functional testing of the hatch development unit  showed no anomalous behavior on the latch   trains (all latches functioned smoothly and correctly with the gearbox load inputs ). The engineering team  did note the “feel” on the gearbox LRA switching mechanism changed as the system was worn in, but the  switc h did successfully move to the armed position and periodic borescope inspections showed the LRA  continued to engage as designed. The development hatch was slated for use on the Orion Crew Module  Structural Test Article  and the hatch delivery was behind schedule, so there was intense programmatic  pressure to complete the hatch- level development testing as soon as possible. As a result, the team went  into proof pressure testing with minimal instrumentation and no cameras  (believing this test would be easily  completed because of the Apollo similarity plus the  successful subcomponent testing and analysis ). Due to  the schedule constraints, t he test team did NOT independently verify the LRA engagement via borescope   for this first test and relied completely on the LRA switch  being in the armed position.  Strain gauges were  placed on key areas of all latches and a single LVDT was placed on the exterior of the hatch structure.  Figure 12 shows the proof pressure pre- test configuration.  

The target proof pressure in the test was 161 kPa (1.5 times the crew cab in nominal pressure for the  appropriate margin). As the test team slowly increased pressure in stages to 21  kPa, 69 kPa, and 109 kPa,  there were no indications of issues . However, at ~90% of proof pressure, a loud noise was heard from the  containment chamb er. There were several small windows looking into the room, but no test team members  were looking at the hatch directly when the noise was heard. The pressure system was immediately shut  down. When the room was opened, the hatch had been completely “libera ted” from the proof pressure  fixture (Figure 13 ).
478 Development Test Failure Investigation   The investigation began at once, starting with a fishbone failure diagram. The team arrived at more than 50  possible root cause contributors . An initial look at the latch strain gauges did show non- linearity near the  point of failure (see Figure 14 ). The latches shouldn’t behave this way for two key reasons, namely that  they are supposed to act as over -center mechanisms, and that the engaged LRA should be a backup  structural path to prevent latches from backdriving. These were the leading root cause theories, but it was  evident t hat more analysis was needed along with additional testing with improved instrumentation.  

Figure 14 . B-Train Latch Strain Gauge Data f rom Test Failure   The originally planned flex -body analysis was given new urgency. Numerous meetings were held with both  NASA and L ockheed Martin expert analysts, and several models were built using various software tools  (MSC ADAMS , MSC NASTRAN, ABAQUS).  The best (although limited) results came from the ADAMS  model  (Figure 15) . In this case, the  CAD model  was used to generate a complete 17 latch kinematic  assembly.  Flexibility was added to key members that reacted high loads through the use of 1- D springs  where stiffness values were  analytically derived from the FEM. Latches were modeled using a bushing  element at the r oller contact region.  Latch stiffness values were derived from the latch subcomponent  testing.  Finally, the structure was modeled using FEM derived enforced displacements at each latch  location.  This model was continually updated as the development testing continued, however full correlation  was never achieved. The model could show general trends but it wasn’t able to duplicate specific loads or  deflections  of the latch trains as seen in test ing. This was due to the hundreds of parts and joints, each with  different friction and stiffness values. When combined with the complex pressure- induced deflections of the  hatch structure, it became clear that the model would be of limited use.  Therefore,  the team continued to  focus on an empirical approach.
479   Figure 15 . ADAMS Model of Hatch Mechanisms   A design of experiments study was implemented to isolate all variables which resulted in nine additional  pressure tests. Many more strain gauges were added to the prototype, particularly in the gearbox, on the  linkages connecting the latches, and on the hatch structure. More LVDTs were added plus  numerous GoPro cameras inside and outside the pressure vessel to monitor latch movements.  

The first retest was intended to be a repeat of the initial test conditions, but with slower pressure application  and more careful monitoring of all instrumentation. The gearbox LRA was confirmed to be engaged via  borescope for this first ret est. Interestingly, the hatch reached proof pressure and the latch non- linearities  didn’t recur. This immediately suggested that the LRA had not been engaged for the first test, despite the  LRA switch being in the correct position.    

A second retest was performed, this time with the LRA  purposely  in the disengaged position. The latch  strain gauges showed the same non- linearity as in the initial test. Furthermore, the hardstop gaps on most  latches increased as the pressure increased (indicating the lat ches were not acting as traditional overcenter  mechanisms).    

Later tests focused on the rigging procedure for the latch positions. In the original failed test, small gaps  were allowed at the latch hardstops (again with the belief that pressure should onl y drive the latches into  the hard stops). The rigging procedure was updated to reduce and mostly eliminate the hard stop gaps.  Subsequent proof pressure testing with the improved rigging was successful, even with the gearbox LRA  purposely disengaged, which demonstrated the LRA could still function as a redundant hatch retention  feature.  

Both the analytical model and the detailed test instrumenta tion supported a fundamental conclusion that as  pressure increases on the hatch, it bows the flat structure like a balloon which caused the distance between  latches to decrease. Since the linkages connecting the latches are rigid, this defeated the over -center  mechanisms, allowing them to release (see Figure 16 ).
480     Figure 16 . Latch Movement Due to Hatch Pressure   It was also noted the test configuration had several components purposely left out as they weren’t deemed  relevant (hinges, thermal isolator brackets, and counterbalance guiderail parts were not installed). These  components slightly stiffen the hatch structure and would have decreased pressure deflections (following  the Test Like You Fly philosophy is important).   At the conclusion of the investigation the root cause was determined to have 3 primary contributors:   1. Hatch structural deformations under pressure- induced latch and linkage loads that  counteracted the latch overcenter loads, which caused the latches to move back  “undercenter.”   2. Inadequate latch train rigging   3. The gearbox LRA  was not engaged despite the LRA switch appearing to function  correctl y. This was due to a design issue found within the switch mechanism.   

Gearbox LRA Switch Mechanism   In hindsight, it was fortuitous that the LRA switch wasn’t functioning properly. If it had functioned, the LRA  would have prevented the latch trains from backdriving, thus giving false confidence in the latch train rigging  procedure and essentially eliminating fault tolerance of the hatch retention system.  Nevertheless, it was  necessary to understand root cause of the LRA switch issue and resolve it.   It is important to note that the LRA is not directly coupled to the Arm/Disarm S witch. Therefore, the LRA  could engage without noticeable motion of the switch due to complicated cam features inside the ELAM not  operating as intended. Corrective actions to address these issues were implemented in the A rtemis 1   design.  The switch functionality did not change except that the Neutral state was removed as it was not  needed and only added confusion. The main improvement came by moving the arm/disarm mechanism  from the ELAM and into the Gearbox Lid.  This allowed for a direct coupling between the LRA and a n added   ‘Status  Indicator’ that was used to show engagement  status.  The status indicator  displays green when the  LRA is engaged and red  when disarmed (see Figure 1 7). Table 2 summarizes the relationship between the  switch state, status indicator, and LRA condition.  It is important to note that this improvement can only be  verified by personnel  inside the Crew Module. Since the Artemis 1 mission is uncrewed, an additional layer  of redundanc y was implemented via a lockout pin installed from the outside of the closed hatch which  guarantees a locked position of the LRA. Further improvements for LRA  status ha ve been implemented in  the A rtemis 2  design to give grou nd crew and astronauts  definitive  feedback on proper engagement.
481 Table 2. Artemis 1 Gearbox LRA Switch States   Switch State  Status Indicator  LRA Condition   “Arm”  Green  Engaged   “Arm”  Red Ready -to-engage   “Disarm”  Red Disengaged  

Figure 17. Improved Artemis 1 LRA Switch Design  

Artemis 1 Side Hatch   Due to the time required to troubleshoot the proof -pressure anomaly, there was insufficient time to complete  the hatch development test program. The Artemis 1 flight hatch was already well into manufacturing and  assembly  while the pressure anomaly investigation was concluding, so it was determined that corrective  actions from the anomaly investigation would be implemented immediately on the flight hardware. The  improvements to the gearbox and rigging procedure were completed and in late 2018 the Artemis 1 s ide  hatch successfully passed proof -pressure testing and a combined pressure/vibration test  (a protoflight  approach was utilized as there was insufficient time and resources for a dedicated qualification unit) . All  instrumentation results from these tests met the pass -fail criteria established from the development test  anomaly investigation. The hatch has been successfully integrated on the Artemis 1 vehicle and thermal  vac testing of the hatch is being conducted as part of the spacecraft level testing at t he time of writing this  paper.
482 Lessons Learned   Specific lessons learned (and relearned) during the Orion Side Hatch development :   • Accurate modeling of complex mechanisms with many parts (especially where flex -body effects  exist) may not be possible  or feasible. Use analysis results cautiously and t est early.   • Beware of using pre- existing designs without thoroughly understanding the effects of your  requirements (the Apollo- like latch trains were more sensitive to pressure- deflections from the new  flat Orion hatch structure). Additional independent analysis in the early design stages could have  identified the  pressure deflection  failure mode.   • Think ahead to possible failure scenarios during testing and ensure that sufficient instrumentation  is used to h elp identify and characterize the failure modes should they occur. Still/video cameras  should be used even for simple structural tests.   • For complex mechanical systems, beware that successful subcomponent tests can impart  false  confidence for higher level assembly tests.   • Ensure that rigging (adjustment) procedures for complex assemblies  are well documented with  clear pass/fail criteria on critical dimensions and features.   • Ensure that the method for verifying the state of mission- critical mechanical functions  is reliable  (the gearbox latch restraint arm state was not properly established during hatch development  testing).  

Conclusion  

The Orion side hatch completed a development test program that identified deficiencies in the design which  have been corrected for the first uncrewed  Artemis  1 mission. The flat structure design (originally thought  to be an improvement over the Apollo curved structure design for seal leak rates and latch mechanism  performance) introduced new pressure- deflection problems that would have reduced system reliability  without correction. The gearbox inspection and hatch rigging procedures were updated which resulted in a  successful completion of the Artemis 1 side hatch protoflight testing, and successful delivery/integration  into the spacecraft.  

Acknowledgements  

• Lockheed Martin Orion Me chanisms team   o Paige Carr, Bill Chastain, James Beat, John Lawlor, Doug Harrison   • NASA Orion Mechanisms Team   o Joseph Anderson, Brent Evernden, Oscar Guzman  

References  

• Evernden, B, Guzman, O., “ ORION - Crew Module Side Hatch: Proof Pressure Test Anomaly  Investigation”, NASA Technical Reports Server (2018)   • Walkover, L, Hart, R, Zosky, E, “The Apollo Command Module Side Access Hatch System,” Aerospace  Mechanisms Symposium, 4 (1969), 157- 167
483 Design, Development, Testing, and Flight of the Crew Dragon Docking System  

Jaret Matthews , Caitlin Driscoll* , Edward Fouad*, Andrew Welter *, Marc Jar mulowicz**   and Jessica Ipnar**  

   Abstract  

On March 3, 2019 10:51 U TC, the SpaceX Crew Dragon Demo 1 spacecraft successfully docked to the  International Space Station (ISS) on the Node 2 Forward International Docking Adapter (IDA) . The SpaceX  developed docking system successfully accommodated the misalignments and relative rates between Crew  Dragon and IS S, and ultimately attenuated the kinetic energy following contact . This was the first  successful  docking of a Commercial Crew program vehicle to the International Space Station. The successful Demo 1  flight was the culmination of years of development . This paper describes  the SpaceX docking system, and  a sample of the lessons learned through simulation, test and inspection.  

Figure 1: The Demo 1 Crew Dragon System on Approach to ISS  

Introduction  

The SpaceX docking system is comprised of several mechanisms . The soft capture system (SCS) allows  for Crew Dragon to compensate for misalignments with ISS resulting from navigation and control errors and  provides the initial (soft) attachment to the station. The SCS also is responsible for  bringing the capsule to  rest with respect to ISS, decelerating it from an initial approach speed at contact with ISS of 8 cm/s. The  SCS consists of a ring with three coarse alignment petals, each with embedded passive soft capture  latches . The SC S petals and passive soft capture latches are derived from the NASA Docking System  design.  

* Space Exploration Technologies,  Hawthorne, CA;  caitlin .driscoll @spacex.com   ** Dynamic Concepts, I nc., Huntsville, AL;  mjarmulowicz@dynamic -concepts.com  

Proceedings of the 45th Aerospace Mechanisms Symposium, NASA Johnson Space Center , 2020