ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
/g47/g68/g88/g81/g70/g75/g3 /g41/g76/g91/g68/g87/g76/g82/g81/g3 /g48/g72/g70/g75/g68/g81/g76/g86/g80 : During launch, ERA is configured in its so-called “Charlie Chaplin”  configuration (see Fig. 7) and is attached to the MLM by means of six launch fixation mechanisms.  Basically, each mechanism consists of one or two EVA-driven hooks to tie down and release ERA from  the mounting seats. The hooks are adjustable. The engagement pins on ERA are fixed. The mounting  seats are made to fit to minimize the backlash. The preload in the hooks is 500 to 1000 N. Gaps between ERA and its support are introduced in order to allow for easy assembly, to cope with thermal expansion, and to ensure proper release. These attachment points are located at the elbow joint (supports 3 and 4),  on the wrist electronic boxes (supports 2 and 5), and on the roll joints (supports 1 and 6). At the Launch BP, the End Effectors are supported in all directions. Figure 9.  Launch interface points for the MLM Figure 10. ERA integration fit checks on the MLM 329
EVA Compatible Mechanisms, Design Details

Figure 11. Location of ERA manual overrides /g48/g68/g81/g88/g68/g79/g3/g50/g89/g72/g85/g85/g76/g71/g72/g86/g3/g76/g81/g3/g87/g75/g72/g3 /g40/g81/g71/g3/g40/g73/g73/g72/g70/g87 /g82/g85/g3/g54/g88/g69/g86/g92/g86/g87/g72/g80 :An EVA manual override is implemented in the EES  Actuator Unit, the Integrated Service Tool (IST) and also for the Torque Force Sensor (TRS) Rigidization  Mechanism (TRM). The manual overrides for the Actuator Unit and for the TRM will be blocked duringlaunch to secure the rigidization mechanism is in the closed position. Just after launch during installation  of ERA in orbit, the launch lock has to be de-blocked. For de-blocking, the override tool has to be inserted into the 7/16-inch hexagonal receptacle of the Actuator Unit and of the TRM manual override interface. By  this first insertion, a pin located within the hexagonal receptacle will be pushed backwards into the deblocked position. The blocking device ensures that the pin will stay within this de-blocked position during all the following in-orbit operations.  The manual override inlets and position indicators are covered by MLI in order to prevent sun trapping  and heat leaks. Before actuating the EVA override, the cosmonauts have to fold back the related parts of the MLI, which then will be secured in the back folded position. /g48/g68/g81/g88/g68/g79/g3/g50/g89/g72/g85/g85/g76/g71/g72/g86/g3/g76/g81/g3/g87/g75/g72/g3/g48 /g68/g81/g76/g83/g88/g79/g68/g87/g82/g85/g3/g45/g82/g76 /g81/g87/g3/g54/g88/g69/g86/g92/g86/g87/g72/g80 /g29/g3Each hinge line in the joints contains an override  mechanism. The special tooling can be inserted at the locations at the central drive axis. For the pitch and  yaw and elbow hinges, the backdriveability is achieved by inserting the pin mechanism in the central hole,  directly engaging the electrical motor unit, hence at the input axis. This means that no backdriving of the spur gear is required. For the roll axis, the accessibility of this hinge prevents this approach. A special engage-disengage mechanism ensures that the gear train is by-passed; hence no back driving of this  gear train is needed. As the ERU hinges can also be dismounted from the tube, special bolts are used to allow decoupling.  These 6 bolts are accessible via a tool. Handrails are required for momentum equilibrium, when the  astronaut is turning these bolts. Again, special tooling is foreseen.    330
The objective of the stopping distance test is to demonstrate that the maximum stopping distance for any  part of ERA is maximum 15 cm at a maximum tip velocity of 10 cm/s. The test has been performed with and without a payload grappled to the arm. The measured stopping distances were in the range 4-7.5 cm. /g3        Figure 8. Moving flat floor disc on la rge flat floor surface (double decoupled)

/g37/g82/g82/g86/g87/g72/g71/g3/g48/g82/g71/g68/g79/g3/g54/g88/g85/g89/g72/g92 /g29/g3The aim of the boosted modal survey test was the qualification of the interfaces  between ERA and the launcher in terms of proto-flight limit loads. Modes measured during the low-level modal survey test were selected in order to achieve qualification loads at the interfaces by exciting these modes on a high excitation level. Not all predicted qualification levels could be reached since the behavior of the test structure was extremely non-linear. The source of the non-linearity was investigated by additional test runs. The non-linearity was caused by interface conditions, which depend on the loads  applied on the ERA. Moreover, the dynamics of ERA were characterized by hammering noises at high excitation levels. The elbow interfaces could be completely qualified during the Boosted Modal Survey  testing. Qualification was not feasible for some components of the other interfaces especially for the axial components. Additional static strength tests were performed to complete the qualification. From the boosted modal survey test it could be deduced that the modal parameters like the mode shapes and Eigen frequencies found in the modal survey test do not apply for ERA under high overall launch loads. Thus, they are not directly applicable for the validation of the finite element model. However, qualification was feasible for the lateral components. Qualification for the axial components was expected to be difficult to achieve and turned out to not be feasible. In order to qualify these components, a separate  static test was required and performed.   

/g41/g76/g81/g76/g87/g72/g3 /g40/g79/g72/g80/g72/g81/g87/g3 /g36/g81/g68/g79/g92/g86/g76/g86 : Control performances such as tracking and positioning must rely on good  modeling of the arm dynamics. This part is mainly determined by the flexibility in the limbs and the  controlled joint behavior. Extensive finite elements analysis with the NASTRAN tool forms the basis for providing model data for the limbs in a multi-body dynamics model within ERA Simulation Facility. For final validation of the modeled flexible behavior of the arm, the real arm will be kicked during test and the position is measured. During this test the joints are not active, and the brakes will be on. The controlled  behavior of the joints is mainly driven by the properties of the gearbox. The final achieved backlash  including any hysteresis effects was less than 1 mrad. /g3 332
GPM Solar Array Gravity Ne gated Deployment Testing Jonathan Penn*, Chris Johnson*, Jesse Lewis**, Trevin Dear** and Alphonso Stewart** Abstract NASA Goddard Space Flight Center (GSFC) successfully developed a g-negation support system for use  on the solar arrays of the Global Precipitation Measurement (GPM) Satellite. This system provides full  deployment capability at the subsystem and observatory levels. In addition, the system provides  capability for deployed configuration first mode frequency verification testing. The system consists of air  pads, a support structure, an air supply, and support tables. The g-negation support system was used to  support all deployment activities for flight solar array deployment testing.  Introduction Global Precipitation Measurement (GPM) is a satellite developed in partnership between NASA and the  Japanese Aerospace Exploration Agency (JAXA) to advance scientific understanding of the Earth’s water cycles and help forecast extreme climate events.  GPM is the successor to the Tropical Rainfall Measuring Mission (TRMM) that was launched in 1997. Just  as for TRMM, GPM depends on solar arrays to deploy once in orbit to provide power. GPM requires a significantly larger amount of power than its predecessor and was designed with a Solar Array Deployment and Drive System (SADDS) that provides approximately 1.75 times more energy. This larger and more efficient array consists of two wings each with four panels per side. Figure 1.  Rendering of GPM in Orbit           Figure 2.  Walk Out of Flight GPM Solar Array One of the challenges at NASA GSFC has been to negate the effects of gravity during deployment testing  of the SADDS since it cannot support its own weight in a 1-g environment. The ability to “g-negate” allows  the engineers to verify the SADDS deployment capability before and after environmental tests and again  prior to launch when fully integrated to the spacecraft. The solar arrays are required to deploy in less than  300 seconds. A consistent time within 10% of baseline performance is required to evaluate performance  changes. At full deployment, all hinge hardstops must be contacting and all hinge latches engaged. The  *Stinger Ghaffarian Technologies, Greenbelt, MD  **NASA Goddard Space Flight Center, Greenbelt, MD /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g23/g21/g81/g71/g3/g36/g72/g85/g82/g86/g83/g68/g70/g72/g3/g48/g72/g70/g75/g68/g81/g76/g86/g80/g86/g3/g54/g92/g80/g83/g82/g86/g76/g88/g80/g15/g3/g49/g36/g54/g36/g3/g42/g82/g71/g71/g68/g85 /g71/g3/g54/g83/g68/g70/g72/g3/g41/g79/g76/g74/g75/g87/g3/g38/g72/g81/g87/g72/g85/g15/g3/g48/g68/g92/g3/g20/g23/g16/g20/g25/g15/g3/g21/g19/g20/g23 335
concept was very simply a flat level surface big enough for the deployment area, it was required to be  modular to facilitate movement within Goddard and to the launch facility at JAXA. The deployment table was an assembly of several tables. Each table was a 1.22 m (4 ft) by 2.44 m (8 ft) aluminum face sheet honeycomb panel on an aluminum frame and legs. Unfortunately, these tables had warped and worn over time and use. Alternative tables were researched but facility compatible materials and weight limits for mobility restricted selection. Cost was an additional constraint restricting table replacement and  refurbishment. It was, in result, determined to be most suitable to proceed with the heritage tables and air  pads and tailor both for GPM’s needs.    Figure 4.  Solar Array Deployment Path Design and Development Once the decision was made to use the air pads and deployment tables, the next step was to design  additional assemblies that would adapt this hardware to the GPM Solar Array’s geometry. This section details how the air pads and deployment tables were optimized and interfaced to the GPM Solar Array to create a working g-negation system.   Figure 5.  Air Pad Figure 6.  Set of Four Deployment Tables Air pad Development To evaluate air pad capabilities, weights were loaded onto structures supported by the same number of  total air pads expected to be used on a flight Solar Array Wing and the air flow was opened fully. The  Air Supply  Hose Support  Structure Airpad Table  Surface Table  Surface 337
intent of this was to compare the maximum capability of the air supply with the weight of the Solar Array  Wing. Unfortunately a problem was encountered as the weight was increased.   Simply increasing the airflow was not sufficient to increase the load capability because the air pads would  start to chatter at high frequency once the weight reached a certain threshold. The air pads were acting as air motors, storing and releasing energy from their plenums causing them to rise and fall. The time delay in filling and emptying the plenums caused a time lag between the air flow into the plenums and the  resulting pressure buildup below the air outlets. This pressure buildup inside the plenum increased the  flow rate through the air outlets, which in turn increased the air pressure below the outlets, increasing the upward force on the load, raising the load, which lowered the pressure below the pad. Figure 7. Air Pad Capability Test This chatter was unacceptable as it would introduce friction preventing ground deployment of the Solar  Array Wing. The solution was counterintuitive: the team blocked half of the air outlet holes on the bottom of the airpads and eliminated the chatter. This moved the major pressure drop to the air outlets. With the absolute pressure ratio at these outlets now above the critical ratio (0.528 for air), the air velocity through the air outlets rises to Mach 1, and no downstream air pressure changes below the air pads can affect the flow rate. With the pneumatic feedback loop broken, no instability is possible. Once this design fix was implemented it was determined that our facilities could supply enough air to lift the weight of a Solar Array Wing with a 40% margin. Figure 8. Air Pad Cross-Section Air padAir Supply  HosesPlatform supported  by air pads Weights Air Inlet Air Pad  BodySpherical Bearing Output Air Hole 338
Panel and Boom Supports A support structure was designed to ensure that the Solar Array Wing had sufficient support to prevent  tipping but did not become over-constrained and bind up during deployment. Four panel supports were  built that each captured the lower Ground Support Equipment (GSE) attachment inserts of its respective  mating panel, shown in Figure 9. Two of the panel supports also had a strut extending to an upper GSE  attachment location to stabilize the array against tipping, shown in Figure 10. An additional support was also designed for the boom. The supports were optimized for an even load distribution among 19 air pads, shown in Figure 12. In some cases two or three airpads were placed directly next to one another to  support a higher load and were coupled with a universal joint to avoid over constraint. Mockup panels  were used to evaluate system performance prior to installing to the ETU assembly.  Figure 9.  Panel 1 Support  Figur e 10. Panel 1 Mounted to Support Figure 11.  G-Negation System in stalled to –Y Solar Array Wing Air Pad   (4 Places ) Attachments to Panel Bottom   (3 Places ) Upper Attachment to Panel  Frame +Y Solar  Array Wing  (Stowed) -Y Solar Array Wing  (Deployed)  Deployment  Tables (13x)  Handrails GPM  Spacecraft Boom  Boom Support Air pads  (19 Places)  Panel Support  (4 Places)  Panel 1  Panel 1  Support Lower GSE  Attachment Upper GSE  Attachment  339
Figure 12.  G-Negation System Air Pad Locations Deployment Tables Problems were experienced immediately on the first tests performed with mockup Solar Array Panels  attached to panel supports even though the air pad capability testing proved that the air flow was sufficient to carry a much higher mass. The deployment tables, being over 20 years old, had spent  considerable time in storage and warping had occurred on their 1.22 m (4 ft) by 2.44 m (8 ft) honeycomb panel top surfaces. Getting the seams of two tables aligned well was in some cases difficult, and getting  all the seams aligned with the full array of 16 tables was not possible. The resulting height differences along the seams of adjacent tables resulted in increased friction or even blocked Air Pad travel. The  problem was compounded when multiple air pads attempted to traverse seams at the same time. This  table misalignment is illustrated in Figure 13.  Figure 13.  Table Misalignment (Height Differential) Much time and effort was spent optimizing the alignment between tables, however, the team reached a  point where any improvement achieved at one location was at the expense of another. It was decided at this point to experiment with laying various materials across the tables to bridge the seams and reduce  the effect of the misalignments. It was found that one particular vinyl flooring material vastly improved  performance. The reason this particular vinyl material helped so much was not yet fully understood and  this topic is discussed in greater depth later in this paper. With the addition of the table surface cover to the table assembly, table leveling and alignment was  revisited. With the table surface cover reducing the misalignment effect at the table-to-table seams, the  need for near perfect seam alignment was no longer necessary. Instead, a “best fit” approach to lining up  the seams was employed. First, each of the 12 main tables was placed in the appropriate relative  1 2  3  4 5 6 7  8 9 11  12 13  14 15  16  17 18  19 10  0.127 mm (0.005 in) < 2.36 mm   (0.093 in)  < 0.635 mm (0.025 in ) Table Table Air Pad  340
successful deployment testing with the mock up panels, the team felt that the development stage was  complete. The team was ready to implement the G-Negation System on the Flight Solar Array Wing. Figure 15.  Testing with Mockup Panels               Figure 16.  Tables with Vinyl Covering                 Figure 17.  Qual Solar Array Test    Adaptation to the Cleanroom In order to utilize the g-negation system on the flight Solar Array Wings, it was necessary to replace the  vinyl surface cover with a low outgassing substitute. Another requirement was the need to dissipate electrostatic charge created by airflow out of the air pads. The particular vinyl used during development testing had specific properties that made it work successfully. This vinyl had a softer lower layer that conformed to the height difference between two tables and a harder layer on top that provided a good ride surface for the air pads, seen in Figure 18. 

Figure 18. Vinyl Surface Covering    Vinyl Layer 1: Soft  Vinyl Layer 2: Hard Table Height Misalignment  Table  Table   342
After extensive research only two alternatives were found with similar properties to the soft layer of vinyl.  The first, Viton Foam, was prohibitively expensive; the amount of material needed to cover the table surface would have cost over $100k. The second, Poron 4701-30, a material made by Roger’s Corporation that is commonly used in aerospace applications was adopted, and large rolls were  purchased. In order to replace the hard layer of vinyl, another material was chosen to place on top of the Poron 470130. The material was 0.254-mm (0.010-in) thick layer of Melinex 515. This material was selected because  it could be supplied with a vapor-deposited aluminum (VDA) coating that dissipated the electro-static charge created by the air pads. The combination of these two materials provided a unique solution of soft and hard material properties to properly absorb floor imperfections while simultaneously being able to  support the air pads.  Each material was ordered and delivered in factory available sheet sizes. In order to lay out each material  on the table, the factory rolls needed to be cut into specific sizes and shapes and placed side by side to fit the table. This was done offline. Once the materials were cleaned and rolled for transport into the cleanroom assembly began. The cut Poron 4701-30 strips were laid out such that their seams did not overlap the table seams. The  Poron 4701-30 provided the first layer to cover the table top. The Melinex 515 strips were then laid out such that their seams did not overlap the Poron 4701-30 seams. The Melinex 515 seams were then taped using 3M40 tape to prevent air pressure from lifting the strips between the seams. The 3M40 tape was chosen because it has the benefit of dissipating electrostatic build up between any seams of the Melinex515 or defects in the material. It should be noted that the Poron 4701-30 strips were not taped between the seams to prevent tearing due to the fragility of the material. The layout can be seen in Figure 19.  Figure 19. Air Pad Gliding on Melinex 515 and Poron 4701-30 Figure 20. G-Negation Table Layers 343
The finished solar array flooring product is a 20 x 25 foot surface made up of thirteen 8 x 4 foot tables.  The tables are first covered by 2.36-mm (0.093-in) thick Poron 4701-30 foam and then topped by 0.254mm (0.010-in) thick Melinex 515 coated with vapor deposited aluminum and taped together with 2-mmthick 3M40 ESD tape.  Figure 21. G-negation Table with Melinex 515 and Poron 4701-30 Figure 21 shows the completed solar array deployment table surface. The picture portrays the surface  such that bubble effects can be seen, however these are an optical illusion due to the reflectivity of the VDA coating on the Melinex 515. Touching the material verifies that there will be no air bubbles in the Poron and Melinex 515 layers once allowed to settle for 24 hours. Minor remaining air bubbles that exist can be pushed out by the test article during testing.    Overall this solution for g-negation worked very well for the largest rigid solar array system built at  Goddard Space Flight Center. G-Negation Deployment Syst em Installation Challenges While the initial development testing of the air pad system was successful of supporting a higher mass  than required, careful implementation of the system with the Solar Arrays was necessary to achieve required performance. The two characteristics used to assess the performance of the solar array  deployment system are the total deployment time and deployment path. The deployment time and path  varied with characteristics of the g-negation installation specifically in the areas of hose routing and the dolly alignment to the g-negation tables. During initial testing of the solar array design while working on the protoflight boom assembly, it was  found that the deployment time varied greatly with changes in the pitch angle, /g84 p, between the deployed  boom axis and the deployment tables. Initial testing showed any angle outside -0.10 /g113to +0.10/g113, at the  shoulder hinge, produced deployments that were longer than expected. Likewise, small changes in the  roll angle, /g84R, of the shoulder hinge provided variable deployment paths.  The offset in pitch was evident upon visual inspection of the boom hardstops after the deployment  finished. A gap at the upper hardstop and contact at the lower hardstop indicates that the pitch angle is greater than zero and the array is attempting to deploy against gravity. A gap at the lower hardstop and  contact at the upper hardstop indicates a pitch angle less than zero and that the array is attempting to  deploy into the g-negation tables and binding the air pads. Adjusting the pitch angle within ±0.10 /g113, and as  close to zero as possible, closed the gap at the hardstops and provided a significantly faster deployment. 344