ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
© 2020 Lockheed Martin Corporation   374                          Figure 9 - Start of Blanket Deployment                  Figure 10 - Second Stage Deployment Complete  

Orbit Raising  

Lockheed Martin’s commercial satellite heritage has been to deploy the solar arrays once all orbit raising  maneuvers have been completed. With rigid solar arrays, power generation during this orbit raising phase  is possible because when stowed, rigid solar  arrays still have solar cells that can be illuminated by the Sun  (the entire outboard panel can generate power during orbit -raising). However, the MM Solar Array does not  have any exposed solar cells when stowed. This means that the MM Solar Array must be deployed as soon  as separation from the launch vehicle occurs. The result is that vehicle orbit raising maneuvers must be  done with 23-meter (75- foot) long deployed solar arrays on either side of the vehicle. The vehicle’s orbit  raising engine drove signi ficant stiffness and moment -loading design requirements into the Boom Assembly  and hinges , but the highest load was seen prior to the last engine firing when the vehicle was at its lightest  mass.
© 2020 Lockheed Martin Corporation   376 Cure Profile of Composites  

Another lesson learned involving the mast revolved around the cure profile of the mast. Masts had been  successfully produced for  the development phase, the qualification phase, and even the production phase  for two flight programs. But when production began on masts for the third flight program, a change was  made that had dramatic unintended consequences. In an effort to be more af fordable, the production team  changed the cure profile of the mast so the mast could be cured in a shorter amount of time. The change  consisted of a quicker ramp rate to the cure temperature and a lower cure temperature. Both changes were  acceptable per the tolerances on the engineering drawing. However, the tolerances on the drawing were  quite large ( 20°C) and no one had ever built a mast to these parameters before . The result was a brittle  mast that cracked when stowed, though the cracks did not present t hemselves until the third or fourth  stow/deploy cycles. By that time, the masts had been fully integrated into the Deployer Assemblies before  it was realized that they were discrepant and would have to be replaced. Cost and schedule impacts  resulted, along with a lengthy  failure review board investigation. The lessons learned in this instance are  many. First, cure profiles of composite structures should not be changed after they have been proven to  produce acceptable flight hardware. Or if they are changed,  then adequate delta- qualification tests should  be conducted to ensure the end product is still acceptable. Second, clear communication channels  between  production teams and engineering teams are essential . The production team made the change to the cure   profile without any buy -in from the engineering team . The engineering team would have denied the request  for the cure profile change if they had been consulted because they knew the cure profile was critical to this  type of composite structure. Third, the engineering team needs to ensure that whatever tolerances are  established for  the cure profile are qualified. The engineering team assumed  that the end product  would be  the same given one ramp rate versus another or given one cure temperature versus another. But the  engineering team never validated this assumption through tests. If the engineering team specifies a wide  range of ramp rates and cure temperatures, then the engineering team needs to ensure those ranges  produce acceptable hardware.  

      Figure 12 – Caution: even though these two composite cure profiles are within the tolerances, one  may produce acceptable hardware while the other may not
© 2020 Lockheed Martin Corporation   377 Hinge Springs   The next lesson learned involved  the springs used in the hinges. During the first phase of deployment of  the MM Solar Array, four hinge lines are deployed to position the MM Solar Array for deployment of the  Blanket Assembly. During qualificat ion of the hinges, the springs used in the hinges were procured from a  supplier that specialized in producing springs. But for the flight units, the procurement team changed the  supplier of the springs to a general machine shop to save cost. This machine s hop delivered discrepant  springs that were fabricated from the wrong material. But the discrepancy was not realized until after two  sets of hinges had been built, tested, and installed on the flight MM Solar Arrays. Major cost and schedule  impacts resulted and several work -arounds had to be implemented to meet critical vehicle need dates.  Some basic but important lessons were learned from this event. First, communication between the  engineering team and the procurement teams is  important. The responsible engineer was not consulted on  the supplier change for the springs. If he had been, he would not have approved the change. Second, if it  is important that a product like a spring be procured from a specific source, then engineering should create  a source- control drawing that specifies the approved suppliers.  

Hinge Studs  

Another lesson learned involving hinges revolved  around a part of the hinge called the stud. This  small machined part contains a threaded end  with an undercut feature. The callout on the  undercut was “R .015 undercut to minor dia. of  thd”, i.e.,  undercut w ith a radius of 0.015 inch  (0.38 mm) to the minor diameter of the thread.  Even though this part and parts like it had been  successfully fabricated for over 20 years, a  supplier used to produce studs for the MM Solar  Array hinges misinterpreted this undercut callout.  This supplier interpreted the callout to be  undercut 0.015 inch (0.38 mm) from the minor  diameter. The result was a neck on the part that  was too small  (see Figure 9 ). During assembly of  one of the hinges, the threaded part snapped off  when a nut was torqued to it. This discrepancy  was noted after multiple sets of hinges had been  built, tested, and integrated into flight MM Solar  Arrays. Major cost and schedule impacts resulted  and several work -arounds had to be  implemented to meet critical vehicle need dates.  In the end, the flight drawing was updated to  explicitly call- out the minor thread dimension and  show the undercut to that dimension. The lesson  learned here is that even if the part has been  successfully produced by multiple suppliers for   over 20 years, design engineers should always assess their drawings for any ambiguity because i f there is  any ambiguity  in the drawings, then someone may misinterpret it.  Figure 13 - Correct Hinge Stud Undercut Versus  Incorrect Undercut
© 2020 Lockheed Martin Corporation   378   Figure 14 - Old Hinge Stud Undercut Callout Versus New Callout  (dimensions in inches)  

Tension Mechanism  

The next set of lessons learned involve the Tension Mechanism, which was discussed in some detail in the  Deployment section. As discussed there, the Tension Mechanism contains  a cable wrapped around a spool .  One end of the cable is swaged to a threaded stu d while the other end of the cable is secured to the spool.  The spool is connected to a constant force negator spring so as a result, the negator spring is constantly  trying to retract the cable around the spool. During assembly of the Tension Mechanism and duri ng  installation of the Tension Mechanism on the MM Solar Array, care must  be taken to ensure that  uncontrolled retractions of the cable do not occur. If uncontrolled retractions occur, then the threaded stud  at the loose end of the cable can be uncontrollably slammed into parts of the Tension Mechanism and  damage the mechanism. Damage was most often seen in the form of broken strands in the cable. In order  to prevent these uncontrolled retractions, 3D -printed plastic parts were used as shop aids to restrain the  motion of the cable. The lesson learned is that proper tooling  is required to prevent unintended motion of spring- loaded  mechanisms. One other valuable  lesson learned regarding  the Tension Mechanism involved the decision to build a  dedicated qua lification unit. The Tension Mechanism design  is largely unchanged from a heritage mechanism that has  been performing nominally on- orbit for many years . Due to  this successful heritage, the design team was tempted to not  build a dedicated qual unit and jump straight into acceptance  build and test. However, it was suggested to the team that  since the product had not been built in over 15 years, it would  be prudent to build a dedicated qualification unit. This  decision turned out to be the right one because as  requirements evolved, the design changed and deltaqualification tests were required to qualify these changes.  The fact that a dedicated qualification unit existed made it  easy to conduct these delta -qualification tests.  

   Figure 15 - Tension Mechanism with  Restraining Tool  (red part)
© 2020 Lockheed Martin Corporation   379 Tooling  

One of the good lessons learned from the MM Solar Array experience was to get the production team and  the tooling team involved early in the program. When these teams are involved early, they can get a headstart on what tools will be needed to most effectively assemble and test the product. By working alongside  the design engineers, proper tooling can be designed and fabricated in time to support the flight builds.  Often, a program may get to a certain part of the build and find that they forgot to build a critical tool or  fixture, or a critical tool or fixture is not designed correctly. By engaging with the production and tooling  teams early, the MM Solar Array team avoided this common pitfall.  One of the most complex pieces of  tooling was the system used to offload the MM Solar Array during deployment tests , called the Rail Offload  System. The Rail Offload System had to meet several  challenging requirements. It had to accommodate  the entire 23- meter ( 75-foot) length of the MM Solar Array. It had to accommodate a lateral motion of about  3 meters ( 10 feet) to allow the hinges to open during the first stage deployment. It had to be as low friction  as possible to simulate zero gravity. And it had to offload each major subassembly in the MM Solar Array.  The Boom Assembly, Deployer Assembly, and Blanket Container Assembly were offloaded using one link  to the Rail Offload System, but the Blanket Assembly required each of the 30 panels in the Blanket  Assembly to be individually linked to the Rail Offload System. This was the biggest challenge because  when the Blanket Assembly is in its stowed configuration, the Blanket Assembly is less than 2.54- cm   (1-inch) thick. The challenge was to develop a solution to be able to package 30 panel offloads in such a  small space.  The solution had 3 major parts  to it. One was that thin offload trolley s were used. Two was  that two tracks were used, one to offload the odd- numbered panel s and one to offload the even- numbered  panels. By doing this, the offload trolleys for adjacent panels could nest within each other, thus saving  space in the stowed configuration. And three was that it turned out that the 30 offloads did not have to fit in  a 2.54- cm (1-inch) wide space. It was acceptable for the offload cables  to have a slight angle to them when  the Blanket Assembly was stowed as long as there was some compliance built into the offload cable.   

Figure 16  - Blanket Assembly Rail Offload System
© 2020 Lockheed Martin Corporation   381 Obtain Independent Perspectives  

During the first deployment of the fifth flight MM Solar Array, an anomalous behavior was noted at the very  end of the deployment. The behavior manifested as a sudden jerk of the blanket followed by ripples of the  blanket along its length. The deployment successfully completed to lock -out, but this behavior was not  nominal , so an investigation was started. The team investigated many possible causes and converged upon  what they thought was the likely root cause, i.e. , that a bracket was undersized and not allowing a bar to  be smoothly pulled out of this bracket. The team was  in complete agreement that this was the likely root  cause and was about to implement some trouble- shooting steps when it was suggested to the team that  they solicit some outside input from a group of independent observers. It was not long before the group of  independent observers were able to convince the team that an undersized bracket could not have caused  this anomaly. Further inspection of photographs of the hardware revealed the true root cause, i.e. , the  presence of a burr on the bracket. Inspection of the actual hardware confirmed  the presence of the burr on  this bracket and only this bracket. All other brackets had been properly deburred. During deployment, the  burr behaved like a hook and momentarily grabbed onto the bar. As the force of the deploy ment overcame  the burr, it suddenly released the bar, causing the sudden jerk in the deployment motion. The obvious  lesson here is to ensure all burrs are removed from interfaces that move relative to one another. But a more  important lesson is that schedu le pressures sometimes cause teams to want to quickly determine root  cause and move on to the next step. If this team had done that, the evidence of the burr would have been  destroyed and root cause may never have been determined. Teams must learn to take a breath and ensure  that root cause has been conclusively determined before moving to the next step in the process. Another  lesson learned is that it is always valuable to get an independent assessment of an anomaly from a trusted  expert.  The independent e xperts in this case were able to see the issue from a different perspective and  challenge the assumptions that the team made to determine the wrong root cause. The independent experts  were able to offer a more plausible root cause that upon further investi gation turned out to be the correct  root cause.  

Figure 19  – Burr on Bracket
© 2020 Lockheed Martin Corporation   382 Conclusion  

Many lessons were learned during the development, qualification, and production of the MM Solar Array .  The MM Solar Array has been successfully qualified to AIAA  S-111 for Solar Cells and AIAA  S-112 for Solar  Panels, including continuous monitoring of the solar cells during the 1.5X thermal life cycle test. All  mechanisms on the MM Solar Array have met the requirements of AIAA  S-114 for Moving Mechanical  Assemblies. Four Wings are currently flying on two commercial satellites that were launched in February  and April of 2019 and two more Wings are set to fly on another commercial satellite in early 2020. As its  name suggests, the Multi -mission Modular Solar Array stands ready to support multiple missions and is  looking forward to powering the next generation of high -powered Lockheed Martin satellites.   

Figure 20 - MM Solar Array Integrated to Vehicle
383 Solid -State Hinge Mechanism for Simple Panel Deployment System  

Thomas (TJ) Rose*, William (Brad) Hensley * and William Francis * 

Abstract  

A common actuation method for flip-out panel deployment is a traditional  torque hinge mechanism,  which  is successful largely due to its deterministic nature.  However , these mechanisms typically require bulky  support structures with high part counts. Another actuation method  for flip -out deployments is a solid- state  hinge mecha nism, or tape spring. Tape springs have been utilized for several satellite panel deployments  in the past, where the primary advantages of these systems are simple design and low part count. With the  space industry on the cusp of an industrialization surge , we have identified this integrated  design to provide  a distinct advantage over the complex traditional systems. This paper will discuss the development process  for the flight -qualified Roccor Panel Deployment System ( ROC- PDS)  with a specific focus on the analytical  and empirical methods developed for solid- state hinge  mechanism.  

Figure 1: Solid -State Hinge Mechanisms and Deployment system  

Introduction  

The ROC -PDS was developed at Roccor , from concept to flight qualification,  as an integrated solution for  utilization in large constellation applications. The defining advantage of this system is that the part count is  an order of magnitude less than its torque hinge competitor. As shown in Figure 3, this system is comprised  of a composite tube as the primary structure, with  integrated solid- state hinges. By integrating the hinges  in this way, the tube can take the place of the k ickoff spring, deployment torsional hinge, lockout  mechanism, lockout damper, while still serving as the primary deployed structure supporting the pa nel.  

Development of the ROC -PDS required material and structural qualification methods developed by Roccor  and the High Strain Composites ( HSC ) community. This effort  included the development of qualification  techniques for high strain composites, experi mental techniques for material performance measurement,  testing and analysis techniques for the design and validation of subsystems like the solid- state hinges, as  well as empirical and analytical methods for deployment kinematic predictions.  

New Space Qualification Challenges   Roccor has successfully qualified and delivered HSC products for space- flight customers including 1) a  radio- frequency precision furlable boom, and 2) a furlable antenna system; and is currently qualifying HSC  products for space- flight customers including 3) an FCC -certified deorbit device, and 4) a solar array 

* Roccor, Longmont, CO  

Proceedings of the 45th Aerospace Mechanisms Symposium, NASA Johnson Space Center, 2020
384 deployment system. Each mission’s development and qualification program followed a similar test and  analysis plan to that illustrated in Figure 2. Although these four missions were relatively risk tolerant, the  tasks labeled with caution symbols in Figure 2 were very costly, schedule- intensive, and performancelimiting due to a lack of industry accepted practices applying to HSCs. The lack of standards creates a  roadblock for the use of HSCs for high- value DoD or NASA missions , for which a new class of highperformance deployable structures would be enabling. Roccor views collaboration with the broader spaceflight community to develop industry -accepted thin- ply composite manufacturing methods, databases,  engineering tools, test and qualification methods as vital to its business model.  

Figure 2. Roccor’s  benchmark thin- ply composite development and qualification plan adapted from recent  Roccor TPC space- flight programs (steps labeled with a caution symbol are costly, difficult and limiting  due to the lack of applicable industry standards).  

ROC- PDS System  Architecture  

As shown in Figure 1, the ROC -PDS is comprised of a panel, stowage containment fixtures , a hinged boom  with end brackets, kinematic mounts, a hold down release mechanism, and two solar array drive actuators .  The ROC -PDS was developed at Roccor from concept to flight qualification as a low part count solution for  utilization in large constellation applications. The defining advantage of this system is that , by integrating  solid-state hinges into the tube,  the part count is an order of magnitude less than comparable mechanized  systems. By integrating the hinges in this way, the tube can take the place of the k ickoff spring, deployment  torsional hinge, lockout mechanism, lockout damper, and still serves  as the primary deployed structure  supporting the panel.
385   Figure 3: (Left) Stowed System configuration; (Center) Strain energy deployment simulation configuration;  (Right) Deployed system simulation configuration  

Development of Solid -State Hinge  

Closed cross section tubes containing solid-state hinges have been demonstrated in the past for the use  with several long antenna applications  (Ref. 1). Much of the groundwork for this technology was developed  by research facilities ( 2, 3, 4, 5, and 6) to characterize the hinge design features and performance. The  intended application was to package several antennas, approximately twenty meters long, into a compact  rectangular package. The deployment of this tube was mostly dependent on hinge torque, as there was   little to no mass or rotational inertia at the tip relative to a system such as the Roccor PDS. Utilizing a similar  tube and hinge architecture to package the Roccor PDS system would require an evaluation of hinge torque  required to deploy a massive panel, as well as its ability to stabilize the panel mass once deployed. To  assess the feasibility of the hinged boom for this application the deployed system was modeled in ABAQUS  (Figure 4).  

Figure 4: Modal Evaluation of locked - out Panel Deployment System  

Results of this study showed that the first fundamental frequency mode was rotation about the boom, which  cause the hinge tapes to shear relative to one another. Therefore, the two driving design requirements for  the hinge in the Roccor PDS were deployment  torque and shear stability. Optimization and validation of the  hinge to provide enough deployment torque and torsional stiffness was the primary challenge for Roccor in  the development of the panel deployment system.