ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
an optimization of the design and the definition of manufacturing tolerances for the most critical  dimensions.  Figure 5 presents the tolerance analysis results for the baseline design configuration of the linear actuator  and includes the roller screw’s radial run-out values based on the maximum measured values on commercial off-the-shelf roller screws. The prediction of the total run-out of the linear actuator lead at its  maximum extension is 16.1 microns which can only comply with the X-axis positioning accuracy  requirement by considering the lateral decoupling effect of the connection rod (i.e. minimum factor of 2.37 in Table 5). The predicted angular runout results of the lead screw is compliant with the requirements, in  particular considering the reduction of the angular deflection due to the connection rod’s bending  (minimum factor of 2.2).  In addition to the potential reduction of roller screw radial run-out on the custom SCM version, the static  misalignment effect can be reduced by a precise alignment of the flexible guiding system’s axis to the linear actuator control axis. This alignment feature has been integrated in the SCM FM design andcorresponding mounting procedures to further enhance the final positioning accuracies. It consists of the  alignment, using an optical measuring technique, and high precision shimming of the blade support frame, followed by the clamping of the compliant rod connection. This later being specially designed to avoid the transmission of coupling stress during assembly. Considering the criticality of the compliant rod connection design on the performances of the SCM’s  positioning performances, it has been decided to assess its impact on the flexible guiding system using breadboard validation tests.  Figure 5.  Final configuration tolerance analysis results The positioning control of the SCM (along Z) is performed by step counts of the stepper motor. With the  intention of allowing the initialization of the displacement control system, a high-precision switch (±1 μm) has been integrated in the SCM. It is triggered by the lower extremity of the lead screw as illustrated in  Figure 6 and Figure 7.  Another key design challenge of the linear Slit Change Mechanism being its repeatability over the  significant lifetime, careful attention has been given to the lubrication of the bearings and satellite roller screws. With the support of ESTL for the lubrication analyses (and for performing the lubrication of these  components), lead lubricated bearings and hybrid lead/grease lubrication have been selected for the  satellite roller screw. 37
Validation Test Program In order to assess the positioning performances of the Slit Change Mechanism in a step by step  approach, it has been decided to proceed with the following validation test strategy:  1. Characterization of the flexible guiding system performances over the full stroke 2. Assessment of the influence of the compliant rod connection on flexible guiding performances3. Performance measurements of the full mechanism on the SCM QM In this article, the test apparatus and results for the two derisking steps are presented and discussed. The initial breadboard model test results performed on the flexible guiding system are presented hereafter. They are followed by the characterization test results of the influence of the compliance rod connection on the slit positioning accuracy. The performance measurements for the full SCM mechanism are not  presented in this article. Flexible Guiding System Characterization Test Results The initial breadboard tests were performed on a full-scale model of the flexible guiding mechanism,  shown in Figure 8. These tests specifically aimed at measuring the geometrical positioning accuracy and repeatability over the full stroke for a limited number of cycles. The Z displacement was imposed using a  steel wire tensioning system. Displacements were measured using two different high precision Laser Displacement Meters and rotations using a high-resolution Laser Diode Sensor Vector Autocollimator. All X translation measurements were performed in a vertical configuration such to alienate gravity effects. The use of a traction system implied the need to perform +Z and –Z measurement tests separately and  combine the test results. Figure 8.  Breadboard model 1 test setup 39
Figures 9 and 10 present the slit carrier’s displacement and rotational test results in function of the  translation axis Z. High repeatability is observable between measured curves in each translation and rotation. All results are within the allowable limits with an exception for the R y(Z) tests results. Following  these tests, it has been observed that the tensioning cable fixation lug was slightly bent, introducing a perturbation moment around the Y axis for the –Z direction. For the +Z direction, the deviation from the  requirement have been attributed to the breadboard’s blade alignment system that has been further enhanced for the following breadboard tests and QM/FM design. Figure 9.  Slit carrier position in function of the Z in mm (biased by optical setup configuration) Figure 10.  Slit carrier rotation in function of the Z-translation in mm Z [mm]X(Z) [ µm] Z [mm]Y(Z) [ µm] Rx(Z) [arcsec] Ry(Z) [arcsec] Rz(Z) [arcsec] Z [mm] Z [mm] Z [mm] 40
Table 6 summarizes the initial breadboard model test results performed on the flexible guiding system presented above. Table 6.  Max. positioning inaccuracy & worst case repeatability results from Z-displacement Direction Position accuracy Repeatability Measured Required Measured Required X translation ±7.2 μm ±10 μm ±2.17 μm ±5 μm Y translation ±5.1 μm ±50 μm ±0.71 μm ±3 μm Rotation around X ±4.28 arcmin ±5 arcmin ±3.65 arcsec ±30 arcsecRotation around Y ±768 arcsec ±50 arcsec ±5.14 arcsec ±25 arcsec Rotation around Z ±2.60 arcmin ±10 arcmin ±0.11 arcsec ±5 arcmin Flexible Guiding System Performance T est Results with Compliant Connection The second set of breadboard tests were performed on a full-scale model of the flexible guiding  mechanism including the compliant connection rod and its clamping system. All design improvements  identified during the initial flexible guiding system tests, presented above, have been implemented in the  flexible guiding system design of the second Breadboard Model. This second set of tests specifically  aimed at measuring the geometrical positioning accuracy and repeatability over the full stroke for a limited  number of cycles (10). The Z-displacement was imposed using a micrometer in a vertical configuration. Displacements in X, Y and Z were measured using three high-precision Laser Displacement Meters.Rotations were measured using two high-resolution Laser Diode Sensor Vector Autocollimators. The test apparatus for these measurements, shown in Figure 11, included the use of an optical table with pneumatic vibration isolation system. Figure 11.  Breadboard model 2 test setup 41
Figure 12.  Breadboard model 2 flexible guiding system Figures 13 and 14 present the slit carrier’s displacement and rotational test results in function of the  translation axis Z for the second test series (10 full strokes). Very high repeatability is observable  between measured curves in each translation and rotation with a significant enhancement from the  previous test result presented in Figures 9 and 10. The second set of results is well within the allowable  limits with an exception for the R y(Z) test results. In the corresponding figure, the four slit positioning  locations are represented as circles. At these discrete locations where the positioning and repeatability requirements are applicable, they are just met. The origin of the behavior of the rotation around Y over  the full stroke can be attributed to either a misalignment of the flexible blades in the X direction or some dissymmetrical behavior of the flexible blades. Figure 13.  Slit carrier position in function of the Z-translation in mm 42
Figure 14.  Slit carrier rotation in function of the Z-translation in mm One potential enhancement of the flexible guiding system would be to connect both flexible blade ends  together as they act as free pivot points (at +X and –X sides). However, this would make the design much more complex, in particular considering that these two additional mechanical structures would need to  freely move along Z as well as to cope with the important launch environment and therefore be supported  too. In any case, the requirements for the rotation around Y have been relaxed during the project. The design of the flexible guiding system with a connection rod is therefore considered fully suitable for the application and in compliance to the stringent positioning and repeatability requirements. Table 7 summarizes the test results performed on flexible guiding system equipped with the compliant rod  connection (breadboard model 2). Table 7:  Max positioning inaccuracy & worst case repeatability test results from Z-displacement Direction Position accuracy Repeatability Measured Required Measured Required X translation ±1.0 μm ±10 μm ±0.14 μm ±5 μm Y translation ±8.5 μm ±50 μm ±2.36 μm ±3 μm Rotation around X ±0.69 arcmin ±5 arcmin ±2.08 arcsec ±30 arcsecRotation around Y ±67 arcsec ±50 arcsec ±1.21 arcsec ±25 arcsec Rotation around Z ±0.53 arcmin ±10 arcmin ±0.33 arcsec ±5 arcmin 43
Conclusions The validation test results have demonstrated the full adequacy of the flexible blade guiding system  implemented in SPICE’s Slit Change Mechanism in a stand-alone configuration. Further breadboard test  results, studying the influence of the compliant connection to the SCM linear actuator on an enhanced  flexible guiding system design have shown significant enhancements in the positioning accuracy and  repeatability of the selected flexible guiding system. Preliminary evaluation of the linear actuator design, including a detailed tolerance analyses, has shown the suitability of this satellite roller screw based mechanism for the actuation of the tested flexible guiding system and compliant connection. The presented development and preliminary testing of the high-precision long-stroke Slit Change  Mechanism for the SPICE Instrument are considered fully successful such that future tests consideringthe full Slit Change Mechanism can be performed, with the gained confidence, directly on a Qualification  Model. The selected linear Slit Change Mechanism design concept, consisting of a flexible guiding system driven  by a hermetically sealed linear drive mechanism, is considered validated for the specific application of the SPICE instrument, with great potential for other special applications where contamination and high  precision positioning are dominant design drivers. 44
(sync) cable provides redundancy for the dampers and deployment springs, allowing for system  deployment in the event of a spring or damper failure. The sync cable also helps maintain the ratio of the two hinge angles to approximately 2:1, as the elbow and shoulder hinges open 180 and 90 degrees respectively during deployment.  HGAS contains five mechanisms:  three Launch Restraint Mechanisms (LRM’s), a Lower Boom Assembly (LBA), and the gimbal assembly. The upper boom connects the LBA to the gimbal assembly as shown in  Figure 1. The entire HGAS assembly and associated mechanism are supported on an all-aluminum honeycomb plate. The LRMs and hinge line designs were developed from heritage Solar Dynamics Observatory hardware. Launch Restraint Mechanism The HGAS uses three LRMs  (known as the A, B, and C devices) to restrain the upper  boom and the gimbal assembly to the mounting plate prior to  deployment. Each LRM is composed of a latch rod, securing  the upper boom or gimbal  assembly to two spring-loaded jaws. Each jaw is attached to a  non-explosive actuator (NEA).  After firing, the NEAs release the latch rods allowing the system to deploy. As designed, LRM C  releases first, followed four seconds later by the LRMs A and B simultaneously. Deployment commences once all LRMs are released. Kick-off springs at LRM B and LRM C assist in separating the HGAS assembly from the LRMs.  Lower Boom Assembly The LBA, shown in Figure 2, is made up of the two deployment hinges; the elbow hinge and shoulder  hinge, connected via a 0.77-m (30-in) long lower boom. Each hinge incorporates a constant torquedeployment spring, a viscous fluid damper, and a potentiometer (for hinge angle telemetry).  Gimbal Assembly The two-axis gimbal assembly consists of two actuators using a stepper motor and harmonic drive gear  reduction in each axis. Both axes can rotate ±90° in 0.0075° increments. The gimbal is commanded using pitch (local X) and yaw (local Y) axis design orientation made up of a lower actuator bracket that connects  to the upper boom, a middle linkage bracket, and an antenna bracket mounted to the High Gain Antenna  (HGA). The gimbal assembly provides the HGA with the capability to access multiple TDRSS satellite communication links continuously throughout the GPM mission.Figure 1. HGAS Layout 46
Figure 2. Lower Boom Assembly, Key Components Hinge Spring and Damper Development and Sizing Resistance torque testing of the hinge system was performed to determine the minimum torque required  to deploy the system. A design deployment spring torque of 11.3 N-m (100 in-lbf) at both the elbow and  shoulder hinges was selected, leaving adequate torque margin at each hinge. The HGAS elbow hinge uses the same size and type viscous fluid damper as was already selected for  GPM’s solar array deployment system. Taking this known damping rate and design deployment spring  torques as givens, along with the intent for the HGAS to deploy at a 2:1 ratio, the shoulder hinge’s required damping rate was calculated. Since the damping varies strongly with temperature, both damper  temperatures are actively controlled with external heaters to 30.0°C ± 2.0°C (86°F ± 3.6°F) for flight.  Synchronization System A synchronization cable system was added to provide redundancy and control of the deployment path to  the mechanism. The elbow and shoulder hinges are connected via a stainless steel sync cable, which  helps to maintain the 2:1 intended hinge angle ratio. The pulleys connected to the hinges are sized for  this 2:1 ratio. The sync cable’s tension is maintained with a spring-loaded tensioning system that applies  additional tension via two spring-loaded tensioner arms. This tensioning system was sized primarily for thermal gradient cases, as there can be a worst case 65°C (149°F) gradient along the lower boom. Since  the ratio between elbow and shoulder hinges was designed as 2:1, a metric called the /g72/g79/g69/g82/g90/g3/g71/g72/g79/g87/g68/g3/g68/g81/g74/g79/g72 (/g507/g12 was defined as follows: /g507/g3/g32/g3/g72/g79/g69/g82/g90/g3/g68/g81/g74/g79/g72/g3– 2x shoulder angle       (1)  /g36/g3/g83/g82/g86/g76/g87/g76/g89/g72/g3/g507/g3/g76/g81/g71/g76/g70/g68/g87/g72/g86/g3/g87/g75/g72/g3/g72/g79/g69/g82/g90/g3/g76/g86/g3/g79/g72/g68/g71/g76/g81/g74/g3/g87/g75/g72/g3/g86/g75/g82/g88/g79/g71/g72/g85 during deployment/g17/g3/g36/g3/g81/g72/g74/g68/g87/g76/g89/g72/g3/g507/g3/g76/g81/g71/g76/g70/g68/g87/g72/g86/g3/g87/g75/g72/g3 elbow is trailing the shoulder. /g55/g75/g72/g3/g71/g72/g86/g76/g74/g81/g3/g76/g81/g87/g72/g81/g87/g15/g3/g68/g81/g71/g3/g87/g75/g88/g86/g3/g87/g75/g72/g3/g72/g91/g83/g72/g70/g87/g72/g71/g3/g89/g68/g79/g88/g72/g3/g82/g73/g3 /g507,/g3/g76/g86/g3/g93/g72/g85/g82/g3/g82/g85/g3/g81/g72/g74/g79/g76/g74/g76/g69/g79/g92/g3/g82/g73/g73/g3 /g93/g72/g85/g82/g3/g87/g75/g85/g82/g88/g74/g75/g82/g88/g87/g3/g87/g75/g72/g3/g71/g72/g83/g79/g82/g92/g80/g72/g81/g87/g17/g3 Gravity Negation System The gravity negation (g-negation) system enables a vertical deployment of the HGAS in a 1-g environment. Such deployments are required to characterize and qualify the HGAS system prior to flight. The HGAS deployment system qualification included three spacecraft integrated deployments; immediately after initial integration to spacecraft, after thermal vacuum testing (T-Vac), and after vibration  testing. Due to packaging constraints, accommodations for other deployable systems, and configuration  47
As it could not be actually measured, the stow bias was the least certain value in the ADAMS model. To  finalize calibration of the model, the stow bias was iterated until the latch time of both hinges closely matched the latching (i.e. fully deployed) times in the most current g-negated tests. The hinges would not latch at the exact same time due to compliance in the sync cable, allow/g76/g81/g74/g3/g87/g75/g72/g3/g86/g80/g68/g79/g79/g3/g507/g3/g68/g81/g74/g79/g72/g3between the two hinges.  The final calibration to the  ADAMS model was done  using the post T-Vac, gnegated, deployment data.  The hinge angles of the post  T-vac test and the  corresponding calibrated ADAMS model predictions are compared in Figure 6. The 3 N (0.7 lbf) of stow bias used to calibrate the model compares favorably to the 2 to 4.4 N (0.5 to 1.0 lbf) stow bias estimated by the test engineers prior to the test. This calibrated model became the baseline ADAMS model. This baseline model, run as a “on-orbit” (i.e. zero gravity  and g-negation system  model elements disabled) would be used to predict onorbit performance of the flight configuration HGAS.  Potential Deployment Interference After the HGAS subsystem was integrated to the spacecraft, a potential deployment interference issue  between the outboard actuator of the gimbal assembly and LRM C was discovered (see Figure 7). The potential interference was discovered while doing an analysis run with the HGAS ADAMS model for  another purpose (i.e. non-baseline), not to check for interferences between the gimbal assembly and  LRM C.  The predicted interference was verified with a run of the baseline ADAMS model. It was noticed that the  HGAS gimbal assembly penetrated into the rectangular solid representing LRM C (again, see Figure 7).  While constraint (e.g. hinges) locations are based off the CAD model and thus accurately located in the HGAS ADAMS model, not all the geometry is accurately portrayed, so the ADAMS model, by itself, cannot be relied on to check for interferences.  The ADAMS model results for elbow and shoulder hinge angles were inserted into the CAD model to  check for interferences. The CAD model also predicted an interference of the outboard actuator with LRM  C (see Figure 8). In the event this interference led to a failed deployment by jamming the outboard actuator against LRM C, the core spacecraft would lose its high bandwidth communication ability and would have to rely on a much slower data rate. This would delay completing or possibly compromise the completion of mission objectives.Figure 6. Calibrated ADAMS Model Predicted Hinge Angles  Compared to G-Negated Test Values /g3 52
Figure 7. Potential Impact of Gimbal Sect ion with LRM C (image from ADAMS model) This discovery led to close  examination of the HGAS deployable and the following revelation: The potential interference issue is rooted in a fundamental trait of the design. It is  /g78/g76/g81/g72/g80/g68/g87/g76/g70/g68/g79/g79/g92/g3/g76/g81/g71/g72/g87/g72/g85/g80/g76/g81/g68/g87/g72 with no  set deployment path. The  ramifications of this were not  understood during the design phase or even early in the testing phase, as it was assumed that the sync cable would keep the elbow  and shoulder hinges’ rotation  locked in a tight 2:1 ratio. In reality, compliance in the synchronization cable allows some independent rotation of the elbow hinge relative  to the shoulder hinge. The ratio tends towards 2:1, but isn’t  necessarily 2:1. This kinematically indeterminate system resulting from the sync cable compliance  allows the following to influence the HGAS deployment path:  /g120/g42/g76/g80/g69/g68/g79/g18/g36/g81/g87/g72/g81/g81/g68/g3/g44/g81/g72/g85/g87/g76/g68/g29 The relatively large gimbal assembly and antenna inertia resists movement  and thus the elbow end tends to lift off the deck first. Due to the geometry of the boom, this lifting of  the elbow drives the gimbal assembly towards the non-deploying LRM C structure on the HGAS deck (see Figure 7). /g120/g54/g87/g82/g85/g72/g71/g3/g40/g81/g72/g85/g74/g92/g3/g76/g81/g3/g47/g53/g48/g54/g29 Impact energy from the firing of the launch restraint mechanisms and the  force of the kick-off springs at each LRM can compound the issue. The LRM release order plays a  significant role in the deployment path. Figure 8. CAD Model Used to Veri fy Interference Predicted by  Baseline Adams Model 53
The parallel development of a dynamic computer model of this complex deployable and its g-negation  system proved indispensable. It allowed detection of the interference condition and informed the solution to the issue. It saved cost and schedule as much of the work was conducted offline while the spacecraft continued its integration and test program. As the HGAS subsystem hardware was integrated, the fidelity of the model was increased. As ground tests were performed, the model was calibrated. In the end, it was a very valuable tool that allowed the investigation and understanding of a complex deployable in a way that would have been difficult without it. Acknowledgements The authors would like to acknowledge the all team members that were involved in the conception,  development and delivery of HGAS to GPM as well as those that assisted in the issue resolution and  editing of this paper. Specific acknowledgements to: Chris Strickland (Product Development Lead), Tim  Pike (Designer/Engineer), Chris Mathews (Mechanical Technician), Danny Grove (Designer), Andrew Lea (Test Engineer), Bryon Stepp (Stress Analyst), John Tota (Quality Assurance), Wahid Zewari (Engineer), Minh Phan (Chief Engineer), Rodger Farley (Senior Staff Engineer), and Mark McGinnis (Reviewer). 58
In-Orbit Performance of th e MWRI Scanning Mechanisms Manfred Schmid*, Miao Jun**and Yu Shuang ** Abstract Scanning Equipment supporting the Millimeter Wave Radiometer Instrument (MWRI) are flying in a sunsynchronized orbit of 850-km altitude with an inclination of 98.8 /g113on the FY-3 meteorological satellite (FY  = Feng Yun, Wind and Cloud). MWRI is a linearly polarized, ten-channel passive Radiometer; it measures precipitation and water clouds, sea ice, snow/water equivalent, drought and flood index, land temperature and soil moisture.  Following the FY3-A, the FY3-B Satellite was launched in autumn 2010. Since that time, the Scanning Equipment was continuously operated. During the last three and a half years in orbit, the Scanning Mechanism has executed about 65 million revolutions, while the Scan Compensation Mechanism (SCM) -  used for momentum compensation - has already successfully executed more than one billion revolutions. During the commissioning phase of the instrument and during the first operation phase, random torque  spikes, which manifested themselves as a motor current increase, were observed in the Scan Drive  Mechanism, whereas the Scan Compensation drive operated nominally from the beginning. The result of the root cause investigations performed in order to isolate the issue, and the consequences  for the follow-on MWRI equipment which was successfully launched by end of September 2013 (now  flying on the FY 3-C Spacecraft), are discussed.   Figure 1. FY3 Spacecraft and Instrument Configuration Introduction The scanning equipment consists of the Scan Drive Mechanism (SDM), a physically independent Scan  Compensation Mechanism (SCM), and the Scan Drive Electronics (SDE). The conically scanning MWRI  instrument with a mass of 60 kg is continuously operated by the Scan Drive Mechanism (SDM). The  *Airbus Defence & Space (formerly Astrium GmbH Satellites), Friedrichshafen, Germany **ASES (Aerospace System Engineering), Shanghai, China 59