ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
557   Figure 3. Sun’s View of TRMM SADA at high beta angle  

An initial TRMM/XTE SADA rotary actuator qualification unit partial  1000- hr life test revealed that no signs  of deterioration could be detected by any of the tests designed to monitor health (threshold voltage, output  torque) . But post -test inspection revealed significant amount of wear in the bearing to flexspline interface  and the gear teeth. The 304L flexspline inner diameter had galled with the 440C bearing outer race, the  inner race had slipped down the wave generator plug, and the Pennzane  2000 oil was no longer present  [1]. As detailed in Reference 1, a series of changes were made and then a combined TRMM/XTE SADA  life test that covered 1.1x full TRMM life (11 million output degrees  of travel) was  successfully completed.  

The TRMM –Y side array, was  always on the sun,  or warm side (TRMM performed routine yaw maneuvers  to keep one s pacecraft side toward the sun) and experienced temperatures greater than li fe test for  substantial periods, while the +Y SADA on the cold side of the spacecraft, operated within predicted limits.   Unfortunately, there was  no operational way to reduce temperatures as the major heat input was from direct  sun on the –Y SADA output flange and via conduction from the inboard boom and shoulder hinge. The  excessive temperature resulted in increased (exponential) vaporization of l ubricants, which are required for  good lubrication of bearings and harmonic drive for proper actuator life. The original 1000 -hour life test had  proven that internal degradation may not b e apparent from telemetry, i.e., there may be no warning until  catastrophi c failure so  SADA could halt at some random position without notice. In September 2002 the   –Y SADA showed first indication of impending failure as it was stuck during slew -to-feather  for ~ 2 minutes .  Prior to jamming, the SADA was also missing steps every 14 seconds during normal tracking (Figure 5) .  The TRMM team then decided to discontinue sun tracking and park the – Y SADA at the feathered horizontal  position as a precaution to avoid the possibility of that array becoming stuck in a non- prefer red position,  which could hav e resulted in excessive drag .  

Figure 4. –Y SADA operated hotter than expected during mission
558   Figure 5. –Y SADA Anomaly: Prior to jamming for ~2 min, actuator was missing steps during tracking   This lack of sun tracking with the - Y solar array  led to slightly less available pow er, but still allowed sufficient  power for nominal operations of all working instruments . In 2003 , the TRMM team had to decide whether  the +Y SADA could reliably be expected to continue operating normally until fuel considerations forced  controlled reentry to be initiated. An anomaly investigation resulted in thorough lubricant loss, thermal and  life analyses being performed. Key aspects taken into account included the fact that the most cri tical  temperatures are at the harmonic drive wave generator bearing / flex spline  / circular spline interface, and  the outboard motor bearing. Thermal analysis assumed thermal conductivity value across ball bearings  taking advantage of oil meniscus spanning each ball/race interface. As lubricant is drive n off, meniscus  supports less heat flow, creating higher thermal resistance and thus higher temperatures.  

Figure 6. TRMM SADAs Lubricant Loss Analysis  

This can lead to a runaway condition where balls eventually run dry. It is then possible that the –Y SADA  lost lubricant and the dry running may have generated particles . Stainless steel wear particles generated  can become ferro-  magnetic due to work hardening. They can t hen be attracted to motor magnets and may  clog the small air gap between coils and magnets . Without torque multiplication at the output, the motor  may permanently jam. Another source of possible failure was the precipitation of lead naphthenate high-
559 pressure additive, which was detected to some extent in the life test . This effect may be more related to  solubility and wear cycles.  The actuator missed  steps every ~14 seconds corresponding  to every 180° of  motor rotation which may indicate lead precipit ate at the interface of wave- generator bearing and flex spline  that engages twice per motor revolution as each lobe passes by .  

Figure 7. +Y SADA Life Prediction          Life analysis predicted that the +Y TRMM SADA c ould reliably be expected to continue operating normally  until fuel considerations forced controlled reentry to be initiated. Analysis predicted that the 1.1x full TRMM  life (11 million output degrees  of travel) would be reached in March of  2006 and that 16 million  degrees  would be ac complished by July of 2010. The +Y drive operated well within temperature limits and did not  experience the same problems.  

However, because of the situation with the –Y side solar array, the power subsy stem required  special  attention during State Of C harge (SOC) periods, especially during periods of low Beta angle. Eventually  the +Y SADA operated successfully for a total of  approximately  21 mi llion output degrees of travel , until the  spacecraft was decommissioned on April 8, 2015.  

The GPM Solar Array Drive Assembly  

The GPM Solar Array Deployment and Drive System consists of two independent wings, each of which  comprise five major elements. Shown in Figure 8, these elements include: a) Solar panel assembly  consisting  of a rigid panel that supports the solar cells with electrical wiring (cell, diode board, associated  circuitry, temperature sensor) and coarse sun sensor on outboard panels ( quantity 4); b) Deployment boom  assembly connecting the panel assemblies to the SADA output flange. There is one boom assembly per  wing. The hinges within the boom assembly provide the deployment force to position the panel in  the  required orientation. The booms  also support the harnesses trav eling between the panel assembly  and the  SADA; c) Restraint  / Release (R/R) mechanisms securing the panel assembly and boom to the spacecraft  in the stowed configuration. There are five Restraint/Release mechanisms per wing: four for the panel  assemblies and one for the deployable boom assembly; d) Panel Hinges providing the deployment force  for eac h panel assembly. There are t wo panel hinges per hinge line; e) S olar Array Drive Assembly  (SADA)   which rotates the boom and panel assemblies in order to track the sun through each orbit and provide   means to transfer the power generated by the solar arrays  and the sensor telemetry signals to the  observatory.
560   Figure 8. The GPM Solar Array Deployment and Drive System  

The design of the GPM SADA was driven mostl y by structural, thermal, and life / Number of cycles  considerations : structural as it had to help meet the 0.6  Hz minimum fundamental frequency for the largest  Solar A rrays every built at NASA Goddard with a inertial load of 255 kg- m2; thermal as it had to ensure the  actuator was kept within its acceptable temperature limits to prevent the issues that caused the TRMM – Y  SADA anomaly we previously described; and life as it had a requirement of 17,500 cycles ~ 8,750,000°.  

The GPM Solar Array Drive Assembly consists of 4 main subsystems: electro- mechanical (rotary actuator),  structural  (actuator output & Harmonic Drive ( HD), spacecraft and hinge interface plates), thermal  (radiators , hinge interface plate,  and thermal tube), and harness management (cable wrap) . The rotary  actuator  designed and manufactured by Sierra Nevada Corporation’s (SNC) Durham, NC facility, is  compris ed of a three- phase redundant stepper motor with a 200:1 reduction , a Size 25 T -Cup harmonic  drive gearbox , with position feedback provided by integral redundant coarse and fine optical encoders. The  actuators provide positioning within the SADA. The GPM SADA actuator has two hard stops limiting travel  to ±125° and it uses a SuperDuplex™ thin section output bearing lubricated with Pennzane 2001- 3PbNp  oil and 2000- 3PbNp grease. GPM’s structural, thermal, and life requirements were considerably more  severe than previous NASA missions that had used this same type of SNC rotary actuator (SDO, LRO) .  

Figure 9. The GPM Solar Array Drive Assembly
563 Interface Plate (HIP) on the other hand is mounted on the top of the actuator output and it rotates with it .  The HIP provides an interface between the actuator and the shoulder hinge. I t blocks the sun from the  actuator output, and it also provides some bending stiffness . The design of  the thermal radiator from the  structural point of view was driven by the minimum fundamental frequency requirement of 100 Hz.  SADA  structural design was both simple and effective. The solar array loads travel through the shoulder hinge   and the HIP to the out put of the actuator , and then out to the lower bus structure through the SIP. The cable  wrap inertial loads are transferred to the SIP via the cable wrap housing. The following load cases were  analyized using finite element analysis:  22 Gs in X,Y,Z (Mass Acceleration Curve loads); 2050 N shear,  440 N axial, 290 N -m moment (Combined Qual Limit Loads); 55°C and - 15°C (thermal stress) . Stress  analysis showed positive margins of safety on yield and ultimate for all components.  

Figure 11. Cross Section View of SADA Finite Element Model   A superb thermal design was of critical importance for the succesful operation of the SADA  which  had to  handle the thermal load due solar cell power, motor, and solar impingement . On top of the previously  described issues experienced during the TRMM mission, the GPM SADA thermal design was also driven  by lessons learned from the LRO mission . During the LRO observatory level testing it was discovered that  the rotary actuators were ex ceeding their temperature limits as the power loses ( P=I2R) within the solar  array harness were very significant and dumping heat direclty into the actuator. Modifications to the LRO  SADA during and after thermal vacuum testing, such as adding radiators an d removing M ulti-Layer  Insulation  blankets still did not resolve the issues . This resulted in LRO having to make major changes to  the on- orbit solar array operations to ensure functionality of the actuators . Furthermore, GPM had close to  two times the solar array harness power as LRO and 3 times the mission life. Due to this , the GPM SADA  underwent a significant design effort to accommodate the motor power, harness heat ing, and extreme  environment . 

Table 2. Key thermal requirements  

The actuator design could not be modified due to contract with the vendor . Thermal design was then  focused around the rest of the SADA components . Coatings and tapes were applied on the Hinge Interface  Plate, Cable Wrap Housing / end plate, spools, as shown in Figure 12. A large radiator was attached to the
564 edge of the Spacecraft  Interface Plate as shown in Figure 9. The cable wrap was modified to absorb / reject  the heat generated within the harness:  an aluminum thermal tube extends into the actuator quil shaft to  isolate the harness from the quill shaft; the Cable Wrap housing is directly attached to the Interface Plate  instead of the actuator’s accessory flange for direct heat path to the radiator; a 3- piece Titanium spool is  used to isolate the Cable Wrap bearings from harness heat, minimize bearing gradients, and minimize  thermal expansion loads into the Cable Wrap bearings; heat straps going from the Cable Wrap housing to  the lower bus structure were added to help stabilize any transient temperature swings, and provide a stable,  well-characterized heat sink; the Cable Wrap be arings use Pennzane lubrica nt instead of Braycote. This  was consistent with the actuator bearings to ensure no cross contamination of lubricants . Also, Pennzane  has more restrictive temperature limits but longer life than Braycote.  

Figure 12. GPM SADA Thermal Design  

The Cable W rap (CW)  acts as the harness management device. It allows electrical & power signal wires to  pass across the rotational joint with no relative harness motion external to the housing. The CW Mechanism  is driven by actuator motion / rotation of the Solar Array . Both are connected using an Oldham type coupling  with the female portion attached to the actuator quill shaft, and the male side attached to the cable wrap  spool subassembly . This coupling design allows for positional, l inear, and angular misalingments between  the CW and the Actuator and transmits torque without placing side loads to either shaft . Invidivual wires  are sewn together to form two belts that spiral around a central rotating spool . This ribbon portion of the  harness is clamped to both the rotating spool and the CW  housing as shown in Figure 13. By “mechanically  grounding” the ri bbon harnessing to the housing and rotating spool, in addition to “mechanically grounding”  the round bundle to the S/A boom, all motion of the harness occurs within the Cable Wrap housing as  “winding & unwinding” in a spiral motion about the rotating spool  [4]. 

Figure 13. GPM Cable Wrap Transversal Section [4]
569 is well controlled by the thermal system.  This is much better than the 31°C on the –Y TRMM SADA that  resulted in the anomaly .  

Temperatures of the cable wrap are also within the nominal desired values . The maximum Hinge Interface  Plate temperature (closest to actuator output flange) is 35°C below the 40°C requirement indicating that the  thermal design is working properly , and unlike on the TRMM – Y SADA, the GPM SADA ac tuators have   good lubrication of bearings and harmonic drive for proper actuator life.  

.  

Figure 18. GPM SADA Performance Plots
571 Lessons Learned  

For scenarios where a high actuator bending stiffness is required,  the use of SuperduplexTM bearings  with  a high contact angle and preload mounted with a very tight slip fit  can be advantageous . A thorough  harmonic drive accelerated life -testing  program wi th different combinations of circular splines and flexible  splines materials can help select the right combination that provides the desired actuator life. Ensuring  proper lubrication of the actuator bearings and harmonic drive is critical to the life of t he SADA. Performing  a detailed thermal analysis including all heat sources such as internal power, I2 R Joule heating (including  harness going through the quill shaft), environmental loads, etc., is critical to the success of the design.  This is because lubricant temperatures need to be within their operating ranges, otherwise critical  components such as the harmonic drive wave generator bearing / flex spline / circular spline interface and  the outboard motor bearing would prematurely fail . These heating loses can be very significant dumping  heat directly into the actuator and potentially increasing lubricant depletion. To this effect, the use of a  ‘thermal tube’ in the quill shaft helps conduct the majority of the heat away fr om the actuator . Having a  robust thermal design with oversized radiators, actuator output sunshades, and redundant heaters is always  beneficial . Testing of different Solar Array  tracking profiles that allow maximization  of the amount of  feathering helps reduce the number of propulsion burns required to maintain the SC in the desired altitude  as well as reduce the SADA number of cycles and wear on the actuator .  

Acknowledgements  

The succes s of the GPM SADA would have never  been possible without the outstanding support of NASA  Goddard Space Flight Center’s  electro- mechanical,  mechanical, and thermal  engineering branches ’  personnel  and supporting  contractors . The authors would like to thank everyone for their remarkable efforts .  Special recognition  to Sierra Nevada Corporation’s (SNC) Durham, NC facility  for being so accommodating  and cooperative.  

References  

1. Farley, R. E. and Ngo, S. “Development of the Solar Array Deployment and Drive System for the XTE  Spacecraft”, Proceedings of the 29th Aerospace Mechanisms Symposium, NASA Johnson Space Flight  Center, May 1995, pp. 268- 282.  2. Habibvand, A., “Super Duplex Bearing vs. Simplex Duplex Bearing Technical Memorandum”, Industrial  Tectonics Bearings Corporation, Rancho Dominquez, CA, May 1998.   3. Mobl ey, J. and Parker, J. “Harmonic DriveTM Gear Material Selection and Life Testing”, Proceedings of  the 41st Aerospace Mechanisms Symposium, Jet Propulsion Laboratory, May 2012, pp. 39- 52.  4. Ward, J., “GPM SADA Cable Wrap  Mechanical Design Details ”, NASA GSFC, Greenbelt, MD, July  2009 .   5. Patano, S. et al, “ GPM Orbital Maintenance Planning and Operations in Low Solar Activity  Environment ”, 2018 SpaceOps Conference, Marseille, France, May 2018.
573 Two-Axis Antenna Pointing Mechanism Qualification f or Juice Mission   Dual -Band Medium Gain Antenna   

Jorge Vázquez*, Mikel Prieto* , Jon Laguna*  and Antonio Gonzalez *     

Abstract  

JUICE  is an ESA mission travelling to Jupiter to get information from the Jovian environment. The Medium  Gain Antenna will be used in the Telemetry, Tracking and Command Subsystem of the satellite and serve  as the communication antenna d uring close -to-sun phases. It is  a dual band antenna (X and Ka band) with  a  two- axis high accuracy pointing mechanism, which serves also as a deployment actuator. The antenna  is fixe d to the spacecraft panel by means of two hold- downs. In this paper , a subsyste m description and  brief summar y of the qualification campaign is presented, along  with main lessons learned and test  conclusions.  

Introduction  

The Medium Gain Antenna Major Assembly (MGAMA) is a medium- gain dual -axis steerable antenna  assembly that provi des the main uplink and downlink communications between the JUICE spacecraft and  Earth.  The MGAMA will work on the following bands:   • A 2-way X -Band link   • A 2-way Ka -Band link   Based on BEPI -Colombo and SolO heritage, the main  challenges of the design are due to the nature of the  mission. Some of them are:   • High temperature range in the Antenna Pointing Mechanism ( APM ), due to the combination of hot  environment during cruise phase of the mission (up to +150ºC in the APM) and very cold  environment in the science phase ( -90ºC in some parts of the mechanism).   • Dedicated output shaft support with a double full duplex bearing and a dedicated flexible coupling  configuration to get high torsional stiffness and adequate structural strength without overloading  internally the gearhead motor .  • High accuracy required, on the order of 0.02 º, with an Inductosyn sensor for fine position feedback  in the rotation axis. For the reference acquisition,  travel -to-end-stop strategy is  implemented in the  electronics.   • High solar radiation and contamination requirements, with a dedicated thermal protection shield  design based on an  MLI blanket .  • Very stringent heat flux requirement from the antenna to the spacecraft.   • Low mass.   • Need of 358º angular range for elevation axis.  

* Sener Aeroespacial, Getxo, Spain   jorge.vazquez@aeroespacial.sener  

Proceedings of the 45th Aerospace Mechanisms Symposium, NASA Johnson Space Center, 2020
574 Design Description   Main characteristics of the antenna are presented in Table 1.   Table 1. Medium Gain Antenna Main characteristics   Transmit and R ecieve  bands  X band and Ka band (dual)   Peak Gain  (X-band)  28.6 dB i Uplink - 30 dB i Downlink   Peak Gain (Ka -band)  39.8 dB i Uplink - 39.3 dB i Downlink   RF Power handling  50W (X band) - 20W (Ka Band)   Main Reflector diameter  510 mm   Antenna type  Axially Displaced Ellipse Antenna   RF transmission element  Waveguide   Deployment angle  180º  Azimuth pointing range  210º  Elevation pointing range  358º  Mechanism temperature range  -40º +150ºC   Accuracy (both axis)  0.02º   Sensor type  Inductosyn   Stepper motor type  Permanent Magnet  2-Phase 6 -Pole  Power conssumption (per axis)  < 5W at ambient temperature  Total weight (including mechanism)  27.4 kg 

The main components of the antenna are presented in Figure 1.  

Figure 1. Medium Gain Antenna  3D view with main components
575 • The ARA (Antenna Reflector Assembly), represented in purple in Figure 1, is in charge of  supporting the radiating elements of the antenna, and includes  the 510-mm Main Reflector, Sub  Reflector, ARA Bracket, Feed and struts for  supporting the Sub Reflector.   • The Boom assembly is composed of the structural components that grant support of the ARA and  connect ARA with the APM, allowing the ARA pointing and deploying.   • The APM is responsible for the pointing and deploying of the MGAMA components and contains  the only permanent interface with the spacecraft .  • The HDRM 1 connects the output of the APM to the spacecraft panel  to protect the mechanism  from excessive loading when in launch conditions. The ARA is fixed by HDRM2.   • The Radiofrequency Chain is composed of two waveguide sections (X -band and Ka -band  waveguides) in charge of connecting the feed in the ARA with the rotatory joint in the APM. They  are supported to the boom by means of calibrated blades . 

Pointing Mechanism   The pointing mechanism from MGAMA is a dual -axis gimbal providing azimuth and elevation steering  capability. The azimuth axis is driven by the gearhead motor (GHM)  geared to a rotating bracket supporting  the elevation actuator which  is linked to the MGAMA boom. Both are based on stepper motors with  planetary reducers geared to the corresponding output brackets. An integrated X /Ka dual-band and dual - axis Rotary Joint Assembly (RJA) routes the RF energy through the APM in both transmit  and receive   directions.  The APM components are shown in Figure 2.      Figure 2. APM 3D view  (left) and APM Protoflight model (right)  

The actuator is based on a gearhead motor  connected to a pinion  with an elastic coupling in between to  isolate the low speed stage from bending moments and shear forces . The pinion is actuating on the wheel  (ouput shaft) where an antibacklash gear has been incorporated to eliminate play between the teeth.
576 The pinion,  main gear and antibacklash wheel are made of non- magnetic material , with dry lubrication  (sputtered MoS 2) and hardening coating on the external surface to increase the pitting resistance. In the  inside of the gearhead motor, as it can be sealed and temper ature is conveniently controlled by means of   dedicated heaters, wet lubrication has been selected (Braycote 601). This combination of dry and wet  lubrication made it possible to get adequate tribological behavior of the mechanism components while  avoiding lube loss due to evaporation or wet lubricant issues in very cold conditions.  

The anti -backlash wheel  consists of an outer geared ring which is connected to the center shaft by means  of 6 calibrated blades. These blades allow a deformation of 0.4º (0.5 mm at the teeth)  between the central  and outer part of the wheel . The center shaft is connected to the APM main shaft  whereas the outer ring is   preloaded to the pinion to get the backlash reduction.  

In Figure 3 there is a 3D view of the spur gear and the anti -backlash design.  

Figure 3. Main Gear 3D view  

The feedback of the position is obtained by means of an inductosyn sensor. The Inductosyn transducer is  a position sensor made of two plates with square wave patterns printed on them, which become the primary  (excitation in rotor) and secondary (sin/cos patterns in the stator) of an electrical transformer. The phase  between the input and output varies in the relative movement, resulting in a 180º phase change for a  displacement of a coil angular range (360º/Number of poles).  

Figure 4. Inductosyn sensor