ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
330 there was a need,  without disturbing the test verified harness bundle loops and restraint s across both hinge  joints. There was a G -10 spacer, coated with BR -127 ESD, placed under the shoulder hinge to thermally  isolate the deployed boom from the spacecraft.  See Figure 1 for MAG boom description, and Figure 2 for  deployed boom and umbra.  

   Figure 1. MAG Boom Assembly  Description        Figure 2. Deployed Boom and Umbra  

TiNi FC3 frangibolt s were used as the release actuator s for both +Z lock and –Z lock. Each lock assembly  used a Viton pad as a shock damper. There was also a preloaded spring plunger at the +Z lock to provide  a push- off force at the separation interface during deployment. To mainta in Mag boom magnetic  cleanliness, the plunger was installed on the launch lock housing that mounted on the spacecraft. There  was no spring plunger at the –Z lock since the outward boom was preloaded at the –Z lock. The boom  spring- back force after frangibolt firing was  much higher than the push off force from the plunger. On orbit,  after firing the +Z lock and –Z lock  in sequence manually , the Mag boom swung away  from the spacecraft  and deploy ed toward the bottom of the spacecraft . It went  through a transient phase of over -travel and  under -travel and settled  to the final position. The final Mag boom position was controlled by the hinge  holding torque.  Similar ly design ed hinges were  used on multiple previous APL missions with excellent ,  successful heritage  [1]. Figure 3 illustrate s the Mag boom deployment sequence .     

Figure 3. MAG Boom Deploy ment  Sequence
331 Ground Test Verification  

In the early program phase, an Engineering Model (EM)  boom was built to demonstrate the design concept,  the hinge performance, the harness routing scheme, the boom deployment transient  behavior  and the  deployment settling time. Multiple standalone in-air deployment test s on an air-bearing table sucessfully  demonstrated  the Mag boom performance. Qualification level environmental tests performed using the EM  boom validated  the structure and thermal design margins , and also demonstrated compliance with the  alignment requirement s. Figure 4 lists the EM boom test sequence, i ncluding hinge test, bond joint strength  test, boom segment bend test, boom assembly tests, as well as alignment verifications.  

Figure 4. EM Boom Test Sequence  

The F light Model (FM) boom was built under extreme control for maintaining magnetic cleanliness. Only  magnetic -clean  certified tools we re allowed in the designated Mag boom work area. Acceptance level  environment tests validate d the FM boom workmanship. Standalone in-air deployment test ing using  flight  frangibolt actuators verified the boom deployment sequence and deployed boom alignment . Figure 5 shows  the FM Mag boom standalone deployment test. Figure 6 lists the FM boom test flow.  

Figure 5. Flight Boom Deployment Test
332   Figure 6. FM Boom Test Flow  

After integrating to the spacecraft, the Mag boom was included in some spacecraft level tests , including  pop-and-catch followed by  partial walk out, mass property, vibration, acoustic, thermal balance, and full  walk out tests. Figure 7 lists the Mag boom related spacecraft level test flow.  

Figure 7. FM Boom Related Spacecraft Test Flow  

The spacecraft full boom walk out test was performed to verify that the full boom range movement, plus  over travel motion, was free of obstr uction from Multi- Layer Insulation ( MLI) blankets , harnesses, tapes,  and other surrounding  components. With the spacecraft lifted in the air and the boom positioned below  the  spacecraft, the test  introduced a major safety concern. The final test approach was to maneuver the boom  using extension rods so operators c ould avoid working under the suspended load. The test procedure was  rehea rsed several times using the EM boom and a  spacecraft mock up. T he final full walkout test was  successfully performed on the spacecraft with no abnormal findings . See Figure 8.
333              EM Boom Practice   Spacecraft on GSE D olly    FM Boom Test  

Figure 8. Boom Full Walkout Test on Spacecraft  

Dynamic Analysis of the Mag  Boom Deployment  

Deployment Analysis Background   The Mag boom  assembly mechanical deployment was demonstrated and verified at various phases, from  the engineering model stage to flight model stage. The critical verification was the final , flight -like integration  level, flight model stage where all flight components including harness, harness connectors, harness  holdowns,  thermal blankets and micro -meteoroid shielding  were assembled onto the Mag boom, along  with mass simulators in place of flight instruments. Test verification of deployment was the confirmation of  the two Mag boom  arms swinging out and settling into their design positions, which were validated with  precision alignment measurements.  

In the preliminary design phase, a boom- deployment analytical model was built using Adams multi -body  dynamics software. The model  was not very detailed in complexity, but was  sufficient to provide an  indication whether or not the dual -arm Mag boom concept would work for the Program. The deployment  simulation based on this model  provided a level of confidence that the sweep of the arms would not impact  the spacecraft, and that the torque at the hinges would be adequate to deploy the Mag boom to its intended  position.  

Another Adams model was built  in the later phase, after the final flight -like integration and testing phase, to  closely represent the final flight deployment test behavior of the Mag boom. The objectives of generating a  test-verified model were to observe the dynamic effect of deployment of the Mag boom  on the spacecraft,  and to have the capability of running various scenarios of spacecraft maneuvers in orbit and other dynamics  related predictions or validations. The Adams model can provide a validation, or at least an indication, of  the dynamic measurements received from the spacecraft.
334 Description of Adams Model   The primary Mag boom assembly geometry was imported into the Adams model. The analytical model  consisted of the shoulder hinge, the inner boom, the elbow hinge, the outer boo m, the primary launch lock  and snubber, the secondary launch lock, and the four instruments that mount on the outer boom ( Figure  9).  Harness and blankets were included as smeared masses on the two boom segments. The flight -like Mag  boom  that was installed on the deployment, gravity -negated table was weighed so that the Adams model  mass could be adjusted to match. All the components in the Adams model were treated as rigid bodies,  each having its own mass moment -of-inertia and center -of-mass. Since the Mag boom in the deployment  test was oriented with hinge axis being vertical, and the booms were supported on the deployment table  with compressed air bushings, it was assumed in the Adams model that gravity was negated.  

Figure 9. Adams Multi -Body Dynamics Model of Mag boom  

There were three boundary constraints in the Adams model. The shoulder hinge bracket which tied the  Mag boom to the spacecraft was fixed in 6 degrees -of-freedom (DOF). The primary launch- lock bracket on  the inner boom, close to the elbow hinge, held the other stowed end of Mag boom in 6 DOF. The secondary  launch- lock bracket at the tip of the Mag boom was also fixed in 6 DOF. These launch -lock constraints were  timed to be removed in the solver in order to release and deploy the Mag boom. A preload force was applied  at the snubber between the booms when the outer boom folded over the inner boom in the stowed  configuration. The force was derived from a flexible Nastran static analysis of the Mag boom . This force  was a timed reaction between the two booms, and as the outer boom moved apart from the inner boom,  the force went to zero. Hard stops in the rotations of the shoulder and elbow hinges were modeled as stiff  impact -contacts.  

To provide the work  required to open each hinge, a spring torque was applied at the shaft of the hinge. This  torque was not defined as a typical Adams built -in torsion spring with a stiffness and a preload. Because of  the nature of the “saloon- door” hinge, which was sprung i n both directions, an equation had to be derived  to apply this reversing torque that was based on the angle of the hinge brackets. There was a minimum  torque designed into each hinge of approximately 2.26 N ·m, where the minima of the torque curve was  situa ted, and that minima corresponded to the angle of the settling position of each boom. A damper was  designed into the hinge to reduce the overtravel of the hinges and to absorb some energy before the hinges  impacted their hardstops (Figure 10). The componen ts of the damper were included in the multi -body  simulation, and to account for the activation of the damper, a rotational spring- damper function was  modeled that applied the damping only when the damper brackets had rotated with respect to each other.  Friction was also included in the hinge modeling.
335   Figure 10. Components of a Mag Boom  Hinge as Modeled in Adams  

Comparison of Deployment Test and Simulation   Spring torque, damping, boom kick -off preload and contact stiffness -penetration were adjusted and a  simulation that had a close representation of the deployment behavior and duration was progessively  developed and established. The main correlation goals for the Adams model were to generate the same  deployment shapes, and to complete the deployment clos e to the same duration as the deployment test.  Figure 11 shows the Mag boom  on the deployment platform and the Adams model prior to actuation of the  release mechanism. Figure 12 through Figure 15 capture the major deployment shapes of the Mag boom.  Figure 16 shows the final position of the Mag boom  after release and settling from its oscillation.  

The simulation appeared to capture most of the dynamic behavior, in particular the overtravel of each hinge,  the trajectory of each arm due to the multiple joints  of the boom, and the max swing during oscillation. It  was noted, while tuning the parameters, that the time of the shapes in the analysis tended to be quicker in  the early part of the deployment, and then ended up taking a longer time to settle. The presumption was  that the Adams model did not take into account the air resistance on the Mag boom surface, the friction  between the air bearing and the platform, and the additional weight of the air bearings and hoses, all of  which would increase the response t ime in the simulation. At the end, where the settling time took longer  in the simulation, the lack of air resistance was not impeding the oscillation of the Mag boom. Overall, the  Adams simulation was deemed sufficient to  describe the dynamic behavior of the Mag boom deployment  as it closely followed what was seen in the test. As in the test, the simulation showed that the sweep and  reach of the arms would not cause any impact events with the spacecraft.  

Figure 11. Mag boom Test Configuration and Adams Model Prior to Release
336   Figure 12. Release Kickoff, Test at 0.5 sec and Analysis at 0.3 sec  

Figure 13. Elbow Hinge Overtravel, Test at 3.0 sec and Analysis at 1.2 sec  

Figure 14. Shoulder Hinge Overtravel, Max Mag boom  Swing, Test at 12.5 sec and Analysis at 9.9 sec  

Figure 15. Mag boom Max Rebound Swing, Test at 21.0 sec and Analysis at 19.0 sec  

Figure 16. Mag boom Settled, Test at 70 sec and Analysis at 166 sec.
337 Prediction of On- orbit Deployment Effect on Spacecraft   The Adams model of the Mag boom  was integrated with a representative  PSP spacecraft , matching the  geometry , mass, moments of inertia, and center of mass of the spacecraft  bus (Figure 17). The objectives  were to observe how the analytical deployment affected the attitude of the spacecraft, and whether  the  behavior of the Mag boom and the spacecraft would couple together to cause an impact scenario.  

Figure 17. Mag boom Integrated on Spacecraft  

Figure 18. Mag Boom Deployment and Effect on Spacecraft  

Figure 18 illustrates the Mag boom deployment and its effect on the spacecraft. As the Mag boom unfurled,  the spacecraft showed very  small changes in attitude. When the fully outstretched boom swept toward the
338 rear of the spacecraft, there was a larger moment offset from the spacecraft, which caused it to rotate off  its original attitude. The largest angle off the spacecraft axis was seen when the Mag boom  reached it s  maximum overtravel after sweeping past the spacecraft axis .  

The angle was determined to be 12 deg, and after approximately 12 oscillations, the Mag boom motion  damped out, and the spacecraft attitude settled at 7.6 deg ( Figure 19). It was found that the spacecraft  angular velocity during the deployment was at a max of 2.6 deg/sec (0.045 rad/sec).  

Figure 19. Attitude of Spacecraft Off-Axis was Max 12 deg and Settled at 7.6 deg, while Max Angular  Rate was 2.6 deg/s  

On-orbit Deployment and Final Positi on Verification  

On-orbit deployment of the magnetometer boom began on Flight Day 2, just after 24 hours post launch.  Many factors influenced the decision to deploy the boom so quickly after orbit insertion including the  blockage of thrusters that needed to be commissioned prior to a trajectory correction maneuver one week  after launch,  and diminishing data rates as the vehicle traversed farther from Earth. The on- orbit deployment  verification consisted of three significant parts : confirmation of the fracturing of the intermediate frangibolt  prior to detonating the primary frangibolt for boom motion, confirmation of the primary frangibolt fracturing  causing boom motion, and final verification of the position of the magnetometers.   Intermediate Frangibolt Detection   Frangibolt heating for the Mag boom Hinge Release +Z Tall Side on the primary side was initiated with  temperature and timing cut -offs to prevent continuous current draws in the case of a failure. Proper  actuation of this frangibolt would not result in boom motion, as the actual deployment consists of rupturing  two frangibolts. Even with the nominal plan of heating the frangibolt through both the primary and secondary  paths, it was desirable to confirm the breaking of the first frangibolt prior  to releasing the second frangibolt.   Without motion of the magnetometer boom, it was unlikely that the science magnetometer on the boom  could be used for verification of the fracturing of this first frangibolt. Looking at other possible sensors  capable of detecting such an event, PSP is equipped with a Northrop Grumman Scalable Space Inertial  Reference Unit. This inertial measurement unit consists of four low noise hemispherical resonator gyros as  well as four accelerometers, enabling sensing of both rotati onal and translational motion of the spacecraft.  Additionally, each set of four sensors is arranged in a tetrahedral, allowing for redundant sensing of events;  that is, a glitch on a single sensor can be distinguished from an actual event that would be obs ervable on  multiple sensors [ 2].
339 As shown in Figure 20, when the temperature of the frangibolt reached its fracture temperature near 70°C  at time 75 seconds, an impulsive disturbance was observed across all four gyros. This is indicative of the  frangibolt breaking and imparting its energy on the spacecraft. W hile this is not a guarantee of proper  actuation, this disturbance level is significantly above the observed quiescent noise floor and the close  proximity to the stated fracture temperature is a strong indication. The absence of a disturbance when the  redundant current path was exercised reinforces that conclusion or leads one to a double fault scenario,  seeing both paths would have not properly activated the frangibolt.  

Figure 20. +Z Mag Frangibolt Actuation, Gyro Rate Time Histories  

During ground test ing, some activation signatures were sufficiently small that they were unobservable in  the time histories directly. For those cases, the method of using shock spectrums proved to be sufficient.  By analyzing how the frequency content in the gyro measurement s changed over time, one could discern  the impulsive event as a release of energy across a broad spectrum of frequencies. The on- orbit activation  of the +Z Mag frangibolt is shown in Figure 21.  The in -flight signature was similar to ground test spectrums.  

Figure 21. +Z Mag Frangibolt Actuation, Gyro Rate Shock Spectrum   With positive confirmation of the +Z Tall Side frangibolt released, it was deemed safe to proceed with the  full boom deployment.
340 Second Frangibolt Detection  The fracturing of the –Z Short S ide frangibolt for the magnetometer boom hinge release occurred in a similar  fashion to the +Z Tall Side, without the need to scrutinize the gyro data to determine if the deployment  transpired. Like many spacecraft , PSP was mass and power constrained, leading to the selection of small  reaction wheels. While the size of the wheels is suitable for nominal operations, the rate of change of the  momentum of the deploying boom was sufficient to overwhelm the torque capability of the wheels. Instead  of engaging thrusters to counteract the momentum of the deploying boom, since the vehicle was in a benign  orbital location from a pointing standpoint, the spacecraft was allowed to be pulled off attitude by the boom  deployment and autonomously return to the desired atti tude. This sequence of events is displayed in Figure  22.  Figure 22. Attitude Response to Magnetometer Boom Deployment   At time 30 seconds, the boom begins to deploy and brings the spacecraft off sun- point, peaking at  11 degrees off its initial attitude. Over the next 60 seconds, the boom motion dampens out, following which  a slow recovery slew imparted by the reaction wheels is performed. The complete time from initial boom  motion to restoration of attitude was on the order of 180 seconds.  

Deployed Boom Position Verification   The FIELDS Instrument Suite is a set of 5 voltage and 3 magnetic sensors designed to measure DC and  fluctuating magnetic and electric fields, plasma wave spectra, the PSP’s floating potential, and solar radio  emissions. To achieve this goal, the varying sensors are located across the spacecraft, with four located  on the magnetometer boom. These f our consist of two fluxgate magnetometers, a search coil  magnetometer and a voltage sensor (V5 Sensor) [ 3]. To survive solar thermal  inputs and to function  properly, these sensors must be maintained in the shadow of the spacecraft. During commissioning, an  activity was planned to verify the coarse positioning of the boom prior to the first solar encounter to ensure  the safety of the instruments. It is important to note that while the V5 Sensor is a voltage sensor, it is also  photosensitive and can be used to determine if solar input is incident on the detector.   To accomplish the verification, vehicle slews moving the Sun out from behind the Thermal Protection  System and down each optical axis of the seven sun sensors were performed. Locations of the sun sensors  and V5 Sensor are shown in Figure  23. While the slews were designed to illuminate the sun sensors, they
341 were not sufficiently large to illuminate a correctly positioned magnetometer boom, nor the bar of the V5  Sensor. Nominal shadow patterns are presented in Figure24. Absence of a Sunlight signature on the V5  Sensor during the slewing profile would indicate the boom was positioned correctly in the umbra, behind  the spacecraft. Temperature data was also monitored during the slews. Confirmation of the proper attitude  profile was performed using star trackers, gyros, and sun sensors.  

Following the on- orbit execution of the slews, Gu idance and Control analysts and FIELDS scientists  reviewed the data collected and determined that the coarse alignment of the magnetometer boom was  nominal, inside the umbra of the Thermal Protection System. This has now been corroborated by three  separate solar encounters where no ill effects of solar intrusion on the boom have been noted.  

Figure 23. Placement of the V5 Sensor and Sun Sensors  

Figure 24. Shadowed region at the V5 Sensor
342 Conclusions  

The Parker Solar Probe thermal protection system umbra created a unique alignment reqirement for the  MAG boom. Multiple G -negated boom standalone deployment tests with both EM boom and flight boom  validated boom alignment accuray and repeatbility. The ADAMS model analysis provided a prediction on  boom deployment motion, angles, and settling time in space. The on- orbit boom position verfication using  V5 instrument, solar limb sensors, and slewing the spacecraft confirmed a successful boom deployment to  the nominal position.  

Lessons Learned  

• Prior to PSP, all APL saloon -door style hinge deployables were deployed with one hinge line at a  time. Deploying two hinge lines at the same time had not been previously performed. With PSP bi - fold Mag boom experience, it is proven that deploying a Mag boom with  two hinge lines at the same  time is feasible.   • Add a position indicator at each hinge joint to provide a positive indication of the final deployed  position.   • Plan to provide as many blanket grounding points as possible. For the Mag boom, there were 34  groun ding point s available. However, for the two Flexgate magnetometers alone , there were 64  ground lugs to be grounded. Needless to say, there were boom blankets and SCM blanket ground  lugs that need to be attached also. It was a big effort to figure out how t o attach all the blanket  ground lugs.   • Always include safety engineers in the deployment test planning at the beginning. For the full boom  walk out test on the spacecraft, there were a lot of safety concerns. Final test strategy was approved  by both APL and GSFC safety engineers.   • Future efforts to correlate an Adams model to a test setup should include ground- based and  ambient characteristics. Such as air resistance, frictions and drags from ground support equipment   sources.   • Ground test s should include all teams responsible for in-flight monitoring.  This will allow the teams  first-hand knowledge on how data was collected and how to  properly  interpret the data that feeds   into the analysis . 

Acknowledgments  

We would like to acknowledge the support of the APL PSP project management and that of our sponsor,  the NASA Goddard Space Flight Center , along with APL Space Exploration Sector line management. We  would also like to thank mechanical engineers of NASA Goddard Space Flight  Center 5430 br anch who  loaned us the GSFC air -bearing tables for  deployment  testing of the PSP Mag boom.  

References  

1. W. Cheng, et al, Solar Array and High Gain Antenna Deployment Mechanisms of the STEREO  Observatory, AIAA 2007- 6096 (2007)   2. Northrop Grumman. (2016). Scalable SIRU Family.  Retrieved from Northrop Grumman:  https://www.northropgrumman.com/MediaResources/MediaKits/Satellite/Documents/SIRU_Family.pdf   3. S. D. Bale, et al, “The FIELDS Instrument Suite for S olar Probe Plus. Measuring the Coronal Plasma and  Magnetic Dield, Plasma Waves and Turbulence, and Radio Signatures of Solar Transients,” Space Sci.  Rev. 204, 49– 82 (2016).