ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
577 Twist Capsule   The harness is a complex  subsystem in these mechanisms due to the huge amount of lines to be tranferred  from the stator to the rotor. For this purpose, in azimuth axis, a goose- neck coil supports  a metal foil which  goes from the external fixed part to the cable drum (internal rotating part). The foil alternates areas for  engaging (meshing) and areas for fixation  by dedicated brackets. This configuration uses the transfer of  inner drum to external housing by bending and requires some means  to guarantee that the harness does  not slide with respect to the inner drum when unwrapping from the internal drum to the external one. In  order to avoid sliding in the internal drum and get proper alignment in the external support, meshing teeth  and slots were implemented in the current twist capsule and support elements. In order to guarantee the  initial allocation of the twist capsule foil with respect to the teeth pattern,  the gap at the slots on the extreme  positions is  reduced to match exactly the dimension of the teeth.  

Several iterations on breadboards were done to refine the concept, and the cable attachment before and  after the twist capsule were validated by test . The torque behavior of the twist capsule creates a single  direction torque which helps to absorb gaps and backlash. The f inal proposed solution is the one  implemented in Solar Orbiter and Bepi -Colombo HGAMA that was  successfully qualified. The arrangement  of the cables in the twist capsule intends to isolate the noisy signals from sensitive ones. In this sense the  inductosyn sensor signals from the azimuth stator disk are separated from the rotor excitation and the high  current motor lines.  

For the elevation twist capsule, only two  twisted -shielded pair  cables (inductosyn excitation main and  redundant) shall travel from stator to rotor. In  this twist capsule, spiral configuration is used , with cable  configured as  a long spiral that is wrapped more in the internal drum or the external one, depending on the  shaft position. The total number of turns shal l be selected carefully  in order to provide the required motion  without undesired cable entrapment . The surfaces of the internal drum in contact with the  harness shall be  soft and corner -free in order to allow sliding between coils.  

Figure 5. Harness arrangement (left) and azimuth twist capsule (right)  

Hold Down and Release M echanism (HDRM)   The MGAMA is fixed to the spacecraft panel during launch environment by means of two HDRMs. The  HDRM 1 is located close  to the APM and the HDRM 2 just below the ARA. In this HDRM the movable plate  is merged with the ARA bracket and the fixed support is separated into t hree different pieces. The central
578 one only supports the N on-Explosive Actuator (NEA)  and the cone, whereas the other two support a pair  of spheres (female) each. In both the NEA support and the sphere support there are blades that provide  high stiffness in lateral  and vertical  axis and some flexibility in longitudinal  direction (boom axis) to absorb  thermoelastic distortions.  

By integrating the HDRMs with the boom and ARA , a low mass  was achieved  and considerable reduction  of loads transferred to the contact surfaces. The combination of spheres and central cone and proper  selection of the distance between these components was done to optimize strength and stiffness of the  complete subsystem.  Figure 6 present s the HDRMs location in the MGAMA.  

Figure 6. HDRMs location in the MGAMA  

Figure 7 shows the release process. When the rod of the NEA is released, the spring raises the bolt up to  the point where the rod makes contact with the upper surface of the hole in the central cone, acting as an  end stop for the bolt+rod assembly. The rod remains inside this cavity  to avoid any clash of the movable  part with the fixed assembly during deployment.  

Figure 7. Release sequence
579 Mechanism analysis  

One of the key analysis  for the proper dimensioning of the actuator is the torque budget. For this purpose,  the combination of the angular speed  and acceleration of the spacecraft together with the motion profile  characteristics shall be considered. By the calculation of the angular speed (Eq.1) and acceleration ( Eq.2)  of the antenna movable part and also the acceleration of its center of gravity  (Eq.3), the overall f orce (Eq.4)  and moment (Eq.5) on the APM can be obtained.  

Figure 8 shows  the axis convention system together with a 3D plot of the torque as a function of azimuth  and elevation angle.  

Figure 8. Axis convention system and torque in both axis as a function of AZ -El angles  

The performance analysis has been executed in M ATLAB/Simulink with the Dynamic Performance Model,  which includes a multibody definition of the antenna with SimMechanics and the main non- linearities and  disturbances (e.g. friction, twist -capsule or backlash).
580 The Dynamic Performance Model  is a mathematical model which allows, at least, the following analysis:   • Microvibration analysis.   • Mathematical description for AOCS compatibility verification.   • Multibody dynamic model.   • Simulation of MGAMA behavior.   Figure 9 shows the structure of main blocks of Dynamic Performance Model . 

Figure 9. Top -level simulator view  

In Figure 10 disturbance results in the motion phase are presented.  

Figure 10. Disturbance results in motion phase  

APM testing   Life Model  test campaign   For the qualification of the poi nting mechanism, a dedicated Life M odel was manufactured and assembled.  This unit was submitted to the following tests , from which only the most significant will be presented:   • Holding  and unpowered  torque test   • Accuracy, resolution and repeatability  test  • Stall test   • Vibration test   • TVAC cycling  and life test   • Microv ibration test  

For the functional test , a dedicated test set -up was used with a magnetic strip sensor and a torque cell  (see  left picture in Figure 11) . There is another picture of the micro- vibration test set -up on the right side of   Figure 11.   52 54 56 58-0.6-0.4-0.200.20.4 time [s]Torque [N·m]      52 54 56 58-0.500.5 time [s]Force [N]
581 .        Figure 11. Test s et-up for APM functional (left) and micro -vibration (right) test  

The accuracy, resolution and repeatibility test  was performed in a differential sequence, making small  movements , which were always a discrete number of steps  (on the order of 0.1º) and monitoring the  commanded and the real movement. T he error curve obtained show very good values with accuracy o n the  order of 0.02º and some wave error across the angular range.  

Figure 12 shows  the accuracy test results for Juice L ife Model .  • The orange curve plots the difference between the magnetic sensor and the inductosyn.   • The blue curve plots the difference between two consecutive samples measured by the  inductosyn  • The red curve represents the difference between two consecutive samples measured by the  magnetic sensor.  

Figure 12. APM L ife Model  accuracy curve before life test  

During the micro- viration test the exported torques and forces were measured to characterize the  mechanism (Figure 13) .
582   Figure 13. APM L ife Model exported torques and forces during microvibration test  

The vibration test (sine and random) was performed without significant issues, with the exception of some  connector loose bolts which were not  properly tightened  (Figure 14) . 

Figure 14. APM L ife Model  Vibration test set -up for Y and Z axis  

Life test and thermal cycling was performed in a TVAC chamber  between the temperature range -50ºC  to  +150ºC . The movement of the mechanism was monitored during the entire  test. The lifecycle is composed  of 2600  short -amplitude movements (1º) and 750 long- amplitude move ments (360º), making a total  of   2.2 million revs at  the motor level. Figure 15 shows the temperature profile  and the test set-up in the vacuum  chamber.
584 The final status of  the internal components of the APM was very good in general terms. Even if s ome local  depletion of MoS 2 was found in some areas , the general status of the teeth flanges was acceptable with no  pitting or surface degradation evidence . Figure 17 shows  a picture of the final status of the pinion,  antibacklash wheel and main gear teeth.  

Figure 17. Gears status after life test  

Protoflight model  test campaign   After the succesful test campaign with the Life Model , the Protoflight model  assembly, integration and test   was performed. As usual, functional , vibration and TVAC test s were performed and there were no significant  issues. Some pictures are shown in Figure 18.  

Figure 18. APM Protoflight model vibration test (left) and TVAC test (right)  

Another key point in the APM test  campaign was the magneti c characterization. For this test, a special  facility located at Airbus Friedrichshafen was used. Magnetic deperm ing was performed with a magnetic  moment measurement before and after to confirm that the reduction was effective.  After this process, a  total magnetic moment of 136 mA/m2 (combined value in the 3 ax es X Y Z) was obtained. Magnetic  susceptibility was also mea sured. Figure 19 shows  a picture of this test.
585   Figure 19. APM Protoflight model magnetic characterization test  

The MLI installation on the APM was done after magnetic testing . The areas in which the rotor and the  stator MLI are close  to each other  shall be covered with extreme precautions to avoid any clash or  mechanical blockage. The use of li ps or ribs to separate physi cally the MLI blankets is  strongly  recommended in these areas. Other  point s to keep strictly under control are the gaps between  MLI blankets,  especially if they can clash or get in contact with the spacecraft panels or other subsystems.  Figure 20  shows  a photograph of the MLI installed on the APM Protoflight model . 

Figure 20. APM Protoflight model covered with MLI  

The installation of the mechanism in the final antenna assembly requires a proper  fitting of the mechanical  interfaces with respect to the structural parts (boom with APM output ) and also the radio frequency flanges   from the waveguides. Once performed, the following tests were performed at antenna level:   • Functional tests, with deployment of the mechanism at clean room conditions   • Radio frequency tests at antenna level   • Vibration tests (sine)   • Acoustic test   • EMC tests   • Radio frequency tests   See in Figure 21 the deployment (left) and the acoustic test (right) at antenna level.
586      

Figure 21. MGAMA Functional (left) and Acoustic Test (right)  

Conclusions  

During the development and qualification phas e of the JUICE MGAMA, a two -axis pointing mechanism for  high accuracy applications and severe thermal environments has been developed. A qualification test  campaign has been performed, from which the following conclusions have been extracted:   • Combination of wet lubrication in the areas t hat can be sealed and dry lubrication in the areas that  cannot  is a valid approach, if the temperatures on the oil can be maintained in the temperature  range between -40ºC and +150ºC .  • For proper behavior of t he dry lubrication it is very important to contr ol the degradation of the MoS 2  layer by avoiding excessive pitting stress  on it. Better to avoid this lubrication in gears with very  high torques in relation to the pitch of the teeth.   • The MLI on mechanisms can help to control the temperatures and reduce the heat flux onto the  spacecraft, but the MLI definition and installation shall be done carefully with a detailed design of  the blanket’s  configuration. Improvisation during mechanism MLI installation is not allowed.   • For mechanisms with magnetic cleanliness requirements, deperm  is mandatory. However, some  misbalance of the magnets of the motors and the stainless steel present in some components may  increase the minimum magnetic moment above  the required value if this is below 150 mA/m2.  • When the mechanism size and weig ht is comparable to the panel to which it is attached, the  acoustic test in replacement of random is the best approach to avoid overstressing the components  in a non- realist ic manner.  

References  

1. P. Campo, A. Barrio “Development Of A High Temperature Antenna Pointing Mechanisms For Bepi - Colombo Planetary Orbiter”. ESMATS 2013.   2. P. Campo, A. Barrio, F. Martin  “Testing of Bepi Colombo Antenna Pointing Mechanism”. ESMATS  2015.   3. J. Vázquez, I. Pinto, I. Gabiola, I. Ibargoyen, F. Martín “Antenna Pointing Mechanisms for Solar Orbiter  High and Medium Gain Antennas”. ESMATS 2015