ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
308 travel is approached with the hinge radius essentially going to infinity  when fully deployed, and the linear  translation from the constrained radius needs to be free to move . The testing result s in a torque versus  angle plot for each hinge  similar to the one shown in Figure 10. The output is characteristic of a tape s pring  hinge, with a relatively flat output while the hinge unbends  and a peak as the transverse radius, R 1,  unflattens . The initial torque is actually not zero, but this is a test effect due to the backlash in the test  equipment . The final torque is actual ly zero, because the hinge springs are fully released, however because  the hinge springs are unbuckled, there is significant resistance to stowing which is used to resist the onorbit disturbance torques . The fully deployed hinge with unbuckled tape spring s is a structure  with the  stiffness of the tape springs to react the external loads . 

Second- Generation Design   No significant changes are planned for the new design. The fixture requires redesign for the new hinge  interfaces but not for the t est methodology which was successful and cost effective.  

Figure 9.  Hinge Torque Output Test 

Deployment Test  

The Original Design   The hinges installed into the mechanism are about 1750 mm by 1500 mm as seen in a deployed mechanism in Figure 11. This resulted in a relatively large deployment fixture.
309    Figure 10.  Torque versus Angle  

Figure 11.  Hinge Installed on Mechanism  

The first deployment fixture used during qualification is shown in Figure 12. This deployment fixture  performance was good and with the base,  was mobile and adjustable. The size of the fixture was limiting  however . When testing one qualification or prototype unit, the size of the deployment fixture wasn't a  significant issue . When testing multiple units per shipset, the size of the fixture quickly dominated the  volume of the Ambient and Thermal Vacuum deployment floor space.   Hinges  Hinge  Unbending  Secondary  Buckle  Nonlinearity    Transverse  Radius Peak
310   Figure 12.  Origi nal Deployment Fixture  

Based on the original deployment fixture, a redesigned fixture was built . This fixture was designed to nest  and minimize the required floor space when using multiple fixtures . The redesign did achieve the goal of  being able to test an entire shipset of the mechanism in the Thermal Vacuum chamber at one time, allowing  one chamber run for each shipset build, resulting in a significant cost savings . The redesigned fixture is  shown in the vacuum chamber with the door open in Figure 13. Multiple fixtures can be seen behind the  first fixture and mechanism.  

Figure 13.  Deployed Hinges in Large Thermal Vacuum Chamber  

Additionally , changes were made to the test telemetry  collection . Support equipment fittings were made to  simplify the installation of the hardware on the deployment fixture which reduced the installation and  checkout time. The coupling to the telemetry was initially an issue due to the translation and center of  rotation issues previously discussed. The redesign resulted in a tab which was fixed to the flight hardware,
311 so once the coupling was set up and initially adjusted, there was often no further adjustment required for  each installation of flight hardware. This greatly reduced the time during setup during testing and resulted  in consistent test results . Test results from one unit are shown in Figure 15. Note that the shorter times are  in vacuum and are insensitive to temperature. The longer times are in air due to the damping of the air  resistance during deployment.  

Figure 14.  Deployment Test and Deployment Test Telemetry 

Figure 15.  Mechanism Deployment in Vacuum and Air  

Some additional deployment testing was performed during qualification. To investigate the impact of the  alignment during installation, testing was performed on the qualification unit with off nominal rotations and translations in the hinges . Testing was performed with an induced twist or rotation in the hinge  including  the confi guration shown in Figure 16. Deploymen t tests in this and similar off -nominal configurations  demonstrated that there was no significant  difference in the deployment of  the mechanism  with the twist  when compared to the baseline. While every effort was made to properly align the hinges during installation,  this testing was able to determine the sensitivity to alignment, and allow credible engineering assessment of the ins tallation process.   Telemetry Coupling   Vacuum   Air
312   Figure 16.  Hinge Performance Verified with Induced Twist  

Second- Generation Design   The redesigned hinge was for a deployable shield, approximately 2000 mm by 500 mm, stowed in a height  of approximately 50 mm . The mechanism is shown in Figure 17, with the stowed mechanism shown in  Figure 18. 

Figure 17.  Deployed Redesigned Hinge, Installed on Proposed Shield  

Figure 18.  Installed Redesigned Hinge
313 In the original design, the size of the hardware drove the logistics and test costs  and the test configuration  was a driver in the hardware design. Based on this knowledge , the 2000- mm dimension of the flight panel  was shortened to approximately 500 mm . The other dimensions were left unchanged, allowing a simple  change to reduce the size of the test article. By maintaining the weight of the test panel equal to the flight  panel, the moment of inertia about the deployment axis was left effectively unchanged . This resulted in the  test article shown in Figure 19. This resized test artic le could be tested in component -sized facilities, and  moved without large transportation equipment.  

Figure 19.  Redesigned Hinge, Installed i n Proposed Test Article on Notional Fixture  

Ultimately as the vehicle design matured, this panel was not needed to meet the mission requirements, and  this deployable was removed from the baseline design. Conceptually, this test method is sound, and as  other missions require shields, baffles, or sails, this method can be utilized to provide lightweight deployable  mechanisms  at a reasonable cost.  

Lessons Learned    These  hinges are not traditional hinges . The tape spring hinges are undamped,  have relatively low torque  output , and do not have a fixed center of rotation. This can be further complicated if the curvature of the  stowed hinge is not controlled as this can introduce additional “off -nominal” lateral translation, complicating  the release and deployment . The relatively low torque output, the moving center of rotation and the  possibility of lateral motion increases the difficulty of deployment testing by increasing the possibility of inadvertently constraining the mechanism during deployment testing.    The initial design of mechanism required the entire structure for testing the hinges . With an undamped  hinge, testing with a representative inertia is needed to achieve flight -like deployment dynamics during test .  The resulting large lightweight structures  presented a number of logistical challenges .    The large size of the mechanism required significant floor space for manufacturing.    The large size of the structure required large and more expensive environmental test facilities . The test  facilities used  are typically used for vehicle- sized e quipment, rather than component -sized hardware. The  large size introduced significant air damping during ambient deployments and required large test fixtures .  Large test fixtures require more space for storage and operation.    While the large lightweight flight hardware was easy enough to handle, transportation was difficult due to  the size of the shipping containers . The containers were large and heavy , due to the container material  needed to package the hardware.
314 Reduci ng the size of the test article for manufacturing and acceptance testing , reduces  the cost and  schedule of the build and test, and makes  smaller  test facilities feasible for use in testing. The support  equipment  for smaller hardware is easier to handle and store.  

When a similar design was proposed, the lessons learned from the first program were incorporated into the  new design to reduce the size of the test article. By maintaining the flight -like inertia,  the testing maintains  flight -like dynamics . The integrated deployable is tested for assembly, integration, and normal operation on  the vehicle. Because the flight installation will be tested on the vehicle anywa y, there are no additional  hidden c osts as a result of this approach , resulting in successful acceptance testing with all performance  verified at a lower cost.
315 Design and Development of NEA Scout Solar Sail Deployer Mechanism  

Alexander R. Sobey* and Tiffany Russell Lockett * 

Abstract    The 6U (~10 cm x 20  cm x 30  cm) cubesat Near Earth Aste roid (NEA)  Scout 1, projected for launch in  September 2018 aboard the maiden voyage of the Space Launch System , will utilize a solar sail as its main  method of propulsion throughout its ~3- year mission to a Near Earth Asteroid. Due to the extreme volume  constraints levied onto the mission, an acutely compact solar sail deployment mechanism has been designed to meet the volume and mass constraints , as well as provide enough propulsive solar sail area  and quality  in order to achieve mission success. The des ign of such a compact system required the  development of approximately half  a dozen prototypes in order to identify unforeseen problems , advance   solutions , and build confidence in the final design product . This paper focuses on the obstacles  of  developing a solar sail deployment mechanism for such an application and the lessons  learned from a  thorough development process. The lessons presented will have significant applications  beyond the NEA  Scout  mission , such as the development of other deployable boom mechanisms and uses for gossamer - thin films in space.  

Introduction    The NEA Scout  solar sail design comes as a successor to two 3U (~10  cm x 10  cm x 30  cm) cubesats:   the  NASA Marshall Space Flight Center  developed solar sail NanoSail -D 2 and the Planetary Society solar sail  Light Sail-A3 (Light Sail-B to be launched in 2016). Both spacecraft flew as technology demonstration  missions in Low Earth Orbit :  NanoSail -D in 2010 (Figure 1, left ) and Light Sail-A in 2015 (Figure 1, right ).  These two cubesats represent pathfinders on the way to utilizing solar sail propulsi on in order to achieve  science mission s, such as  the primary science objective for NEA Scout:  image and characterize a Near  Earth Asteroid. This mission would not ordinarily  be poss ible with a 6U cubesat , however, NASA has taken  an interest in applying cubesat form factors, methodologies, and risk to perform cost effective interplanetary science missions.  Solar sail technology is a key to enabling that capability 4. While it is conce ivable for  a 6U  cubesat mission to reach a NEA with co nventional chemical propulsion, both the number of targets and the  launch window would be tightly constrained. By utilizing solar sail propulsion, intercepting a large number of targets in virtually any  launch window is made possible. Cubesats are typically deployed as a secondary  payload, and therefore have little to no control over changes in launch schedule and must remain flexible.  

Figure 1.  NanoSail -D 10- m2 Sail (left) and LightSail -A 32 -m2 Sail (right)  

                                                            * NASA Mar shall Space Flight Center , Huntsville AL  

Proceedings of the 43rd Aerospace Mechanisms Symposium, NASA Ames Research Center, May 4-6, 2016
316 NEA Scout Configuration  As a 6U, interplanetary cubesat, NEA Scout will address strategic knowledge gaps of near earth asteroids.  The spacecraft accommodates an imager, star tracker, reaction wheels, avionics, power system,  communications, and a reaction control system in addition to the solar sail subsystem  (Figure 2) . Volume  is a premium within the fixed  constraints of the 6U cubesat form factor  (~10 cm x 20  cm x 30  cm). Mass ,  14 kg total,  is a demanding constraint as well. Solar sail acceleration is a function of sail area and spacecraft  mass. To reach a target asteroid within 2.5 years and meet the 14-kg mass restrictions of the Space Launch  System cubesat deployer, the total s ail area needed to produce enough propulsion was calculated to be   86 m2, deployed with four booms at 6.8 m of length each.  

Figure 2.  NEA Scout Flight Configuration as of September 2015  

The solar sail subsystem (Figure 3) consists of a  single 86- m2 colorless polymer (CP1), 2.5- micron thick  aluminized sail, sail spool assembly, four Elgiloy  (stainless steel alloy variant)  Triangular Rollable and  Collapsible (TRAC)5 booms at 6.8 m each, a gear-driven boom deployer assembl y, a stepper motor, a  motor controller board, and a sensor suite . The deployer design is based on the successful Nanosail -D  deployer system with exception to the addition of a stepper motor  which provides a slower, controlled  deployment , two boom deployers  instead of one , sensor feedback, and a single sail design on an oblong  spool . The sail spool assembly mount s atop of the boom deployer assembly. The sail spool is allowed to  freely rotate about a center post  as the sail deploys. The center post doubles as a channel for the wire  harness and cabling from the reaction control thrusters at the forward (sun facing) portion  of the spacecraft  to the avionics box in the aft  as well as providing structural support between the two halves .  

Figure 3.  NEA Scout 86 -m2 deployed solar sail  (left) and sail spool and boom   deployer assembly (right)  The boom deployers displayed in Figure 4 consist of two boom spools, each with two booms per spool. Separating the booms into two spools is necessary due to the boom length requirement and the stowed  volume constraint. The four booms each exit the deployer every 90 degrees. Each boom spool consists of a center hub, which the booms mount to, a thin top flange (shown in orange) , and a geared bottom flange  (shown in gray). The top flange primarily  provides contact friction from the spool to the boom during  deployment; therefore, only needs to be thick enough to avoid deflections into the top plate (shown in blue).
317 Due to packaging requirements , part thickness  are kept minim al, specifically in long axis (3U direction) of  the cubesat. The bottom geared flange remains thick to provide adequate gear contact with the center  pinion gear, which is directly connected to the motor/gearbox.  

Figure 4.  NEA Scout Boom Deployer Model Stowed   An earlier concept of the boom deployer consisted of  two modular deployers which mount ed onto the bus  and connected through a center module containing the motor. One of the early concept dep loyer modules  can be seen in Figure 5  both i n Computer Animated Design and as a physical prototype . The benefits  included ease of manufacturing and assembly. The concept was abandoned for the single base plate  design primarily due to alignment concerns. Wi th a single plate, it became significantly simpler  to mount  the three gears with minimal backlash while at the same time guaranteeing teeth would not bind during the  large temperature fluctuations experienced during the early  phases  of the mission. Furthermore, a single  base plate allowed for easier installation of the booms and great er alignment precision.    The 6.8 -m-long TRAC booms typically have a slight bend upwards toward the weld side. This misalignment  along the spine of the boom must be accounted for during installation of the boom. Each boom is to be  installed in an orientation that minimized gravity effects (hanging downward). During installation, the tip of the boom is to be located at the desired plane  perpendicular to the  long axis (3U direction) before the boom  is bolted/clamped at the root. By doing this for all 4 booms, the final plane of the sail can be controlled within acceptable angular limits. Finally, a single base plate allowed for load to be carried primarily though the  plate itself instead of through the bus interface. This allows a mass reduction of the interface.  Due to this  design,  the primary load path of the spacecraft is though the baseplate, which considers a single structure  appealing.  

    Figure 5. Early Concept: Modular Design of Deployers    From the c ross-section view in Figure 6, the inside of the deployer can be viewed. In this view , the boom s  spools are shown as translucent in order for  the boom clamps (shown in yellow) can be seen. The clamps  attach each boom to the center hu b (shown in brown) with two 100- degree countersunk screws. A lso in  Figure 6  on the right , both the clamp and the hub themselves are rounded near the top to allow for the
318 boom to flare out at the base and add stiffness to the boom section nearest the deployer. Both analysis and  testing have shown this flare necessary to achieve the highest boom buckling performance.  

The spring- loaded boom arms, shown in green in Figure 6, are used to contain the boom spool during predeployment as well as deployment. Torsion springs are located at each arm and place pressure onto the boom spool at the Rulon J PTFE rollers. The necessity of these arms and their function is discussed in  greater detail under the ‘Design Challenges’ section. Rollers on the backside of the arms serve to help  guide the booms out during deployment and reduce friction. The backside rollers do not place pressure  onto the boom spool directly.  

The boom tip standoff allows for the sail to be attached to the boom slightly above the boom. This standoff  is able to tuck in closely to the deployer in order to maintain the tight volume requirement. The boom tip standoff also serves as a hard stop, not allowing the boom tip to retract further into the deployer. It was  noted during the vibration testing of Light Sail-A, that the boom tips would retract into the deployer slightly.  The retraction was not a great amount, but enough to possibly cause a failure. It was suggested by the Light Sail-A team  to add a hard stop at the boom tips. The boom tip standoff serves this function.  

Figure 6.  Cross -Section NEA Scout Boom Deployer Model Stowed (left) and Boom Attachment to  Hub (right)  In order to minimize volume, Rulon J PTFE flanged sleeve bushings are used in place  of bearings  for the  boom -spool interface. Two bushings contact both the top and bottom of the boom spool at center race. An  example of this bushin g can be viewed in Figure 7 shown in black (note the specific bushing in Figure 7 is  standard PTFE and not the Ru lon J variant). Both bushings sit on the top and bottom of the post. Only the  bottom bushing is present in Figure 7. These bushings  both significantly reduce friction and allow for tight  alignment of the spools. As with several aspects of this design, volume constraints and form factor are the  design drivers. Similar, but larger bushings  are used for the sail spool.
319   Figure 7.  Baseplate with Posts and Bottom Bushing Mounted   The burn wire mechanism , shown in Figure 8,  allows for the boom deployers to be locked down during  launch and up until deployment  of the sail . The mechanism itself is only a slight modification on the  NanoSail -D burn wire mechanism that served the same purpose. The mechanism locks down one of the  two spool geared flanges. By locking down one of the flanges, the entire geared system is unable to rotate.  

Figure 8.  Burn Wire Mechanism (left) and Early Prototype Burn Wire Mechanism (right)   The geared flange (shown in Figure 9) is machined with a spoke pattern with sixteen recesses. These  recesses allow the gear to be locked down at 22.5- degree intervals. The spring loaded lever (shown as  gray in Figure 8) has a cylinder mounted to it  (not shown) . This cylinder fits into any one of the sixteen  recesses of the spoke pattern when the gear is to be locked down.  When locked down,  a monofilament  wire of 50-lb-test (220- N) Honeywell Spectr aline is tied off to the spring- loaded lever in order to keep the  cylinder tightly pressed into the recess. Once the  Spectraline is cut, the spring- loaded lever swings open   pulling  the cylinder out of  the recess into  the gear ’s channel, allowing the spool to spin freely. In order to  cut the Spectraline, two Nickel -Chromium wire heaters are added in series to the Spectraline (one being  the primary heater and the other functioning as  a redundant heater). The heater  is a coiled Nickel- Chromium  wire mounted into a ceramic sleev e. When enough current is run through the heater, in a matter of seconds,  the Spectraline is effectively cut allowing the spring- loaded lever to fall into the open position. Ignoring minor  dimensional adjustments, the burn wire mechanism remains similar to the NanoSail -D mechanism with the  addition of a microswitch on the lever to provide feedback when the lever has opened.   Bushings
320   Figure 9.  Spool Gear Machined with Spoke Pattern for Launch Lock  

Design Challenges  

Blooming   NanoSail -D, LightSail , and NEA Scout  utilize  Triangular Rollable and Collaps ible (TRAC)  booms originally  developed and patented by the Air Force Research Laboratory  (AFRL ). NeXolve (Huntsville, AL)  currently  has the design license for manufacturing and is on contract to produce the engineering development unit  booms for NEA Scout (Figure  10). As the sail for each mission grew  10 m2, 32 m2, and 86 m2 respectively,  the boom length also grew :  2.2 m, 4 m, and 6. 8 m respectively. At larger lengths, new complications arose  during deployment. For example, due to the strain energy developed while spooling,  TRAC booms slip past  one another during deployment, causing the boom wraps  to expand radially  and create a gap between the  central hub and the first boom wrap.  This reaction is referred to as ‘blooming’ and leads to complications   during deployment . If not controlled properly during dep loyment, ‘blooming’  can lead to suboptimal  deployme nt and possible failure (Figure 11). 

Figure 10.  Boom Deployer Prototype with flight- like TRAC booms   Both NanoSail -D and Light Sail-A addressed issues with ‘blooming,’ therefore the problem was identified  early in the design. Early attempts at creating a MSC Adams multibody dynamic simulation solution proved futile as the forces inside of the deployer were difficult to quantify. These f orces include: strain energy in  the boom, torsion on the boom arms, contact friction of the arm rollers on the boom, friction between the boom spool flanges and boom wraps, and friction between subsequent boom wraps. It was evident early  in the design phas e that prototypes would need to be developed in order to understand and control  ‘blooming.’ With the aid of fused  deposition 3D printing and machined parts, several prototypes were built ,  tested, and iterated upon.