ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
79 solution implemented during detailed design phase was to  add shear pins and corresponding hole / slot at  the gear / output shaft interface, as well as to increase  the size of the screws al lowing higher preload, and  therefore increasing the sliding threshold. The lesson  here is to always allow enough contingency overall  to cope with likely increased requirements and to re cognize the benefit of s hear pins which are less  dependent on the bolted joint to prevent sliding.   Harness Routing

The overall azimuth movement range is 120°. This mo vement is composed of long strokes at beginning of  life (deployments during ground tests, then in-space deployment after up to 6 years ground storage), and then subsequent small movements over the lifetime ( 15 years in-orbit). These small movements are daily  continuous for Radio Frequencies (RF) sensing applications.  The harness design has been chosen mainly due to the volume constraints requirements, and include a relaxation loop (left picture of Figure 4). Design of the relaxation loop has been greatly accelerated and  proven to be a good concept due to the use of rapi d prototyping early in the project. However, both RAs  being identical, they have the same harness length, leading to different  connectors positions. A tooling  bracket is used at ADTM Mk2 level to ensure each RA  harness achieves the same positions consistently.  Integration at ADTM Mk2 level of two RAs proved difficult to consistently achieve these positions, due to small variations of the cabling inside the harness. Ha rness bending / routing reflects these variations, and  influences significantly the end position of the connec tors. A solution for this was to implement longer  wires with temporary connectors at RA level for tests,  and cut to length at ADTM Mk2 level then wire the  final flight connectors.   Qualification   Qualification Status

At the time of writing this document, initial functi onal and post mechanical (i.e. vibration & shock) tests  have been performed on the qualification model and some flight models (FMs). Functional tests consist of  RA torsional stiffness measurements (including backlash  threshold verification), resistances + insulation +  grounding measurements, axes perpendicularity a nd coning measurements, full sweep motion over the  complete angular range of both RAs (for step size  accuracy and repeatability), minimum start-up voltage  (for torque margin analysis) and hysteresis performance across the operating range.  Thermal Balance has been performed on the qualif ication model and thermal model correlation  performed. Thermal Vacuum tests have been performe d on both qualification model and the first four  FMs. The qualification model will continue to final functional tests and life test sequence. 

   Figure 6: ADTM Mk2, Left = Vibrations and Sho cks Setup, Right = Thermal Balance Preparation   Coning and Stiffness results

The requirements for coning and stiffness measurements have been met; Figure 7 shows a few pictures representative of the performance achieved.  NASA/CP-2010-216272
80         Figure 7: Left = Coning measurement, Right = Stiffness + Backlash measurements 

Tests Lessons Learned   Step size Accuracy

The step size accuracy is one of the key functional parameters of the ADTM Mk2 when used for RF sensing missions. The performance of the closed control loop system is indeed improved when the  dispersion of the step size is low.  The knowledge of the step size accu racy is therefore fundamental and  has to be characterized during qualification and acceptance tests.  To measure all the steps sizes (which are nominally 0.0025° /step) with the required resolution, a 22-bit  encoder (resolving down to 360°/2^22 ≈ 0.000086°, which is 3.4% of the nominal step size) mounted on  the output shaft of the monitored axis was selected. Measurements are made during a full sweep cycle of  the 120° angular range in functional tests, an example is shown below in Figure 8.  

Figure 8: Encoder Step Size Reading during Motion  NASA/CP-2010-216272
81 Zooming on Figure 8, some repeatable frequenci es seem to occur, as seen in Figure 9. 

Figure 9: Encoder Step Size Reading during Motion, Zoomed In 

Further investigating this in an Excel spreadsheet and using the discrete Fourier Transform by FFT on the  encoder and ADTM Mk2 potentiometer readings, t he plots shown in Figure 10 were produced. 

       Figure 10: Fourier Transform of step size over 4096 consecutive steps,  left = Encoder Fourier transform, right = Potentiometer Fourier transform   The Fourier transform on potentiometer data does not yi eld exploitable results ( aberration from the FFT  algorithm, no frequency detected in the noise of t he potentiometer reading), but the FFT on the encoder  data does make the frequencies very easy to read.  Further thoughts on the root cause lead to calculat e all the ADTM gearboxes tooth to tooth frequencies,  which converted back to output step size are t he following: 1.125, 8.889, 54, 360, 2400steps.   The FFT analysis tool in MS Excel is limited by the ma ximum of 4096 samples. This leads to a frequency  analysis only possible in the range 2 to 2048 steps (if taking 4096 consecutive steps), with a very rough  resolution in the higher steps frequencies and a good resolution in the lower steps frequencies. So in the identified frequencies above, the 1. 125 and 2400 steps one won't be visible. Note:  to increase resolution  on higher steps frequencies, it is possible to anal yze 16384 steps modulo 4 (i.e., 4096 samples), but this  method cannot be extended indefinitely as we will t hen hit the maximum mechanical range of the ADTM  Mk2.  NASA/CP-2010-216272
82 In the left plot of Figure 10, only 8.889 and 360 can be identified, 54 does not appear. 

       Figure 11: Fourier Transform, Zoom-In   However, this does not explain the much higher amp litude frequencies found in the left plot of Figure 10.  Looking now at the encoder itself, it returns a 22-bit position with normal binary numbers. Calculating the  relationship between bits number and the equivalent  steps length, the frequencies shown in Table 2  appear.   Table 2: Relationship between Encoder bits and Steps Frequency  Encoder Bit Bit Changes over 360° Angle per Bit Change Steps per Bit Change  22 4194304 8.58307E-05 0.034332275  21 2097152 0.000171661 0.068664551  20 1048576 0.000343323 0.137329102  19 524288 0.000686646 0.274658203  18 262144 0.001373291 0.549316406  17 131072 0.002746582 1.098632813  16 65536 0.005493164 2.197265625  15 32768 0.010986328 4.39453125  14 16384 0.021972656 8.7890625  13 8192 0.043945313 17.578125  12 4096 0.087890625 35.15625  11 2048 0.17578125 70.3125  10 1024 0.3515625 140.625  9 512 0.703125 281.25  8 256 1.40625 562.5  7 128 2.8125 1125  6 64 5.625 2250  5 32 11.25 4500  4 16 22.5 9000  3 8 45 18000  2 4 90 36000  1 2 180 72000 

NASA/CP-2010-216272
83 

Figure 12: Fourier Transform, Zoom-In bis 

Looking back to Figure 10, some of these frequencies do appear in the readings. In particular, in Figure 12, the amplitude seen at the positions corresponding to the 11th bit (70.31 steps) and 13th bit (17.58)  are the major contributors to the er ror recorded by the encoder. Thus it seems that most of the error  recorded by the encoder during functional tests is actually generated by the encoder itself. Looking in  detail at the encoder technology used, it appears that only the first 11 bits are real mechanical bits, and  the rest of the bits are interpolated using sine / cosi ne signals in the encoder. Therefore, it would appear  that the interpolation algorithm used has some edges effects that are very significant compared to the data read resolution.  There is a lesson learned here: we should not sele ct high precision measuring equipment only based on  claimed accuracy by the supplier, but should also qualify such equipment beforehand. In this case, the supplier claimed a much finer absolute accuracy than w hat we are actually seeing using the FFT function,  and this is repeatable on all the encoders we have bought  for mass production tests. At the time of writing  this document, alternative measurement equipments ar e being investigated, but we now have a better  understanding of our current test equipment and may be able to correct data after calibration, although this is an on-going activity as well.  Despite the disturbances of the step size accuracy  introduced by the test encoder, the ADTM mk2  performances have been proven to be compliant with the targeted flight applications.   Thermal Vacuum (TVAC) Tests with an Inertia

When testing the qualification model in a TVAC configuration (after thermal balance), it was initially planned to attach the pitch output shaft to a rigid in ertia to simulate the worst case antenna load. To  achieve this, a 90 kg-m 2 inertia was to be used outside the chamber, with a shaft linking it to the ADTM  unit via a feed-through on the top of the TVAC chamber. The main interest for this inertia was to be able to measure hysteresis, and preliminary tests proved that the pass-though seal had too large an influence  on the results to be of any use, as shown in Figure 13.  NASA/CP-2010-216272
84 

Figure 13: Inertia Test 

On the Figure 13, the inertia wa s loaded to apply ~25-Nm load on the ADTM shaft, and then released  free, the idea being to make a quick verification of the frequency response of the setup. The green line is the inertia movement read by the encoder, and the other lines are the prime and redundant potentiometer  monitoring the SMG movements. This test enables us to measure a few things:  3 First the backlash measured on the potentiometers when loading the inertia corresponds to the  SMG backlash, which is apparent when the applied load on the RA output shaft is compensating  the spring motor anti-backlash device (>7 Nm).  The backlash measured on the SMG here is  ~0.01°, or 4 steps.  3 The other interesting information from this plot is the hysteresis of the i nertia: after the test, it  settled about 4 steps away (according to the encode r) from its start position. So here again, the  lesson learned is that to measure sensible data on  very small movement, it is necessary to really  think through all the possible causes of friction that  would influence significantly the results (here:  passing through a TVAC chamber seal).  3 And finally, zooming in the data it is possible to extract the length of a cycle of the inertia during  motion, which corresponds to the natural frequency of the system. Here it was measured to be at  ~1.15 Hz. 

Conclusion   The ADTM Mk2 has been a challenging program. New mass production test equipment has been  developed and is still being improved to cope with the higher demand (more than double from Mk1  versions). Qualification and flight productions have been and still are running in parallel successfully, thanks to the good cooperation between the prime customer and the ADTM team.   First qualification test results were as expected and demonstrated the robustness of ADTM mk2 design.  Some lessons have been learned along the way, such as design improvement rationale and also test equipment selection and calibration, such as encoder accuracies. Combination of analyses and units tests results are being collected to allow family data to be built up and presented (as is being requested more and more by our customers). Production is on its way!  Even if today the ADTM Mk2 fulfils the most stringent requirements for the current flight missions, the  anticipated evolution of these missions continues to require better performances like a smaller step size  NASA/CP-2010-216272
87 Scanning Mechanism of the FY -3 Microwave Humidity Sounder 

Manfred Schmid*, Li Jing** and Christian Hehr* 

Abstract 

Astrium GmbH Germany, developed the scanning e quipment for the instrument package of the MicroWave Humidity Sounder (MWHS) flying on the FY-3 meteorological satellite (FY means Feng Yun, Wind  and Cloud) in a sun-synchronized orbit of 850- km altitude and at an inclination of 98.8 °. The scanning  mechanism rotates at variable velocity comprising several acceleration / deceleration phases during each  revolution.  The Scanning Mechanism contains two output shafts, each rotating a parabolic offset Antenna Reflector.  The mechanism is operated in closed loop by means of redundant control electronics. 

Introduction 

MWHS is a sounding radiometer for measurement of global atmospheric water vapour profiles. An Engineering Qualification Model was developed and qualified and a first Flight Model was launched early  2008. The system is now working for more than  two years  successful in orbit.   A second Flight Model of the Antenna Scanning Mechani sm and of its associated control electronics was  built and delivered to the customer for application on the follow-on spacecraft that will be launched by the end of 2010.    Instrument Description 

The operating frequencies of MWHS include both th e atmospheric transparent window at 150 GHz and  the water vapor absorbing lines around 183.31 GHz. The 150-GHz radiometer has two polarizations (V  and H) and the 183.31 radiometer has three channels (from ±1 to ±7 GHz around 183.31 GHz).  The instrument comprises four major functional  units, namely the Antenna Scanning Mechanism (ASM)  with Antenna Reflectors and Feed Horns, the Moto r Control Drive Electronics (MCDE), the Front-end  Receiver with Receiver Electronics and the Instrument  Electronics Unit. An Instrument block diagram is  shown in Figure 1.   The scanning equipment (shown in the center of Fig. 1) includes the Scanning Mechanism, the Antenna Reflectors and the Feed Horns, all mounted to a base-plate on top of the instrument.                                                     * EADS Astrium GmbH Satellites , Friedrichshafen, Germany  ** Centre for Space Science and Applied Research, CAS, China   Proceedings of the 40th Aerospace Mechanisms Sym posium, NASA Kennedy Space Center, May 12-14, 2010  NASA/CP-2010-216272
88  

Figure 1. Instrument Block Diagram 

The arrangement of the MWHS instru ment on the spacecraft is shown in Figure 2. The Scanning Unit is  mounted on top of the instrument electronics compartment  (shown in Figure 2 on the right side). Figure 3  depicts a typical picture taken by MWHS, s howing the global brightness temperature. 

                     Figure 2.  MWHS Instrument Configuration on FY-3 Spacecraft 

    NASA/CP-2010-216272
89   Figure 3.  Global Brightness Temperature taken by MWHS 

Key Requirements   Major mechanism functional performance characterist ics are the scan velocity constancy, the scanning  time accuracy, the pointing accuracy (calibration), an d the capability to accelerate/decelerate the Antenna  Reflectors within a minimum time in order to ma ximize the observation time during each scan period.   Switch-over capability between three different scan ning modes is provided. There is the nominal scan  mode with its variable speed, a constant velocity mode without hot/cold calibration, and the possibility to  stop the scan at each arbitrary position. The equipment orbit life requirement of 3 years leads to about 36  million mechanism revolutions at nominal operation.  A summary of the key requirements is shown in Table 1.   Table 1.  Key Requirements 

Requirement Value  Orbit Altitude 836 km  Swath Width 2700 km  Spatial resolution (Nadir)  15 km  Power consumption < 18W  Power Supply DC: 27V, ±12V, +5V  Mass 15 kg (overall, incl Base Plate and thermal H/W)  Environmental temperature -30° ∼ +55°C  Life >Three Years  Viewing Angle (Earth Scan) +/-53.35°   Scanning speed error (during earth scan) 1 %   Pointing Accuracy <0.1 deg  Scanning Period  2.667 s ± 50 ms  Scan profile Scheme 1:  Nominal Scan profile  Scheme 2:  constant speed in 2.667 s period  Scheme 3:  stop at any position   Operation Frequency  150 and 183 GHz  NASA/CP-2010-216272
90 The nominal operational instrument rotation speed is variable (hot calibration, cold calibration, earth scan  and motion between these three angular ranges). T he earth scan takes 1.71 s for 106.7 deg. of scan  angle.   

Figure 4. Scan Profile 

Scan Mechanism Overview and Configuration 

The Antenna Scan Mechanism uses a Base Plate on top of which the Scan Drive Unit is mounted. The Scan Drive Unit comprises a brushless DC Motor with  redundant stators and a redundant 17-bit optical  encoder for closed loop velocity, respectively position control. The controller is FPGA based and included  into the Motor Control Drive Electronics (MCDE). The motor carries two output shafts. One Antenna Reflector is mounted to each of the output shafts in order to allow simultaneous and synchronous rotation  of both reflectors. The incoming signal is focused via the two reflectors to the focal planes of the overall three Feed Horns mounted to support structur es on either end of the Base Plate. 

Some important design features are:  • The mechanism is equipped with two pairs of pre-loaded ball bearings in face to face  arrangement. In combination with the direct dr ive, the output motion is free of backlash and  hysteresis.  • A very good scan speed performance (low speed erro r) is achieved by the controller using sine  commutation and by mechanical optimization of the motors for low detent torque.  • The mechanism is designed for high dynamic performance (high reflector acceleration) and  energy efficiency. This implies minimization of the moment of inertia of the rotating masses.  • The flexible control electronics interface allo ws fine definition and rotational adjustment of the  reflector mechanical zero position even on the spacecraft level.   The rotating mass’ moment of inertia (<0.007 kg-m 2) was minimized, not only to avoid a momentum  compensation drive in the mechanism which would have increased the mechanism complexity, but also  to minimize for power consumption and to provide ma ximum reflector acceleration capability, leading to a  maximum of useful earth observation time. This goal was achieved by designing light weight antenna  reflectors from aluminium with an attached lightweight thermal shield ("barrel"), rotating together with the  reflectors.  NASA/CP-2010-216272
91 Since the Drive Unit carries the rotating Antenna Reflec tors, it has to be aligned with high accuracy with  respect to the Feed Horns (Front Ends) mounted to both sides of the Drive Unit. Due to the fact that the  base plate thickness was limited to 20 mm maximum,  special attention had to be paid to limit the  mechanical and thermal deflection of the Base Plate especially in combination with the attachment  interface on instrument side. After specifying the inte rface planarity on Instrument side, the Al base plate  was designed as a milled st ructure with stiffening ribs , closed by a thin Al cover plate attached to its  bottom by means of screws and dowel pins in order  to achieve a box-type structure providing good  bending but also adequate torsion stiffness at low sensitivity against thermal gradients.    Figure 5 shows the arrangement of the Scan Drive Un it and Antenna Reflectors on the Base Plate. The  Front Ends with Antenna Horns are mounted to each end of the Base Plate.  

Figure 5.  Top View of Antenna Scan Mechanism (ASM) 

Figure 6. Scan Drive Mechanism Complete (ASM). 

Drive Unit 

The Drive Unit is built up with a titanium Housing and a steel Shaft interfacing the two Antenna Reflectors.  The material combination was chosen to optimize for minimum mass and good thermal match at  temperature extremes. The Optical Encoder for motor commutation and position feedback is mounted to one side of the Drive Unit Housing. The Drive Unit is m ounted via its footprint to t he structural Base Plate. 

Drive Unit  Antenna Reflector Antenna Reflector Front End  NASA/CP-2010-216272
92 Two pairs of preloaded thin ring angular contact ball bearings with a contact angle of 20° in face to face  arrangement are used in the ASM. The outer bearing diameter is about 50 mm. The number of balls is 24 per bearing and the ball diameter is 3.175 mm. Bear ing material is stainless steel 440 C equipped with  phenolic resin cages, lubricated with Fomblin Z 25. The bearing preload is 650 N. 

The brushless DC motor is designed for a maximum output torque of 2.3 Nm at 27 V. The motor has  independent stators so to provide redundancy and to allow adjustment of the two stators in rotational  direction with respect to each other. This feature wa s included in order to further optimize the system in  view of the high required velocity constancy. By rota tional adjustment of the redundant stator with respect  to the main stator, an additional minimization of the overall detent torque could be realized.  

Motor commutation and position feedback to the contro ller is provided by an optical 17-bit encoder in  redundant configuration. 

Figure 7.  Drive Unit (Drive Module) 

Figure 8. Drive Unit with Reflectors and Thermal Baffle (AIT protection mounted) 

 Scanning shaft interface provides Reflector alignment  provision Scanning  shaft Alignment  cube  Drive Unit Housing  Encoder   Drive Unit  interface to  Base Blate  NASA/CP-2010-216272
94 cylinder and stiffening ring were manufactured and connected by electron beam welding in order to  simplify the manufacturing process. The other end of t he baffle was attached via screws to a circular plate  attached between Drive Unit and Antenna Reflectors (see Figure 8).    During vibration testing, a failure of the electron weld seam was identified, so that a re-design of the  thermal baffle was necessary. In order to solve th e issue, it was decided to manufacture the Baffle  structure in a sophisticated manufacturing process from one piece of aluminium so to avoid any welding seams and the critical welding process (Figure 9). The subsequently repeated vibration test was successfully passed without any problem.   

Figure 9. Thermal Baffles Manufactured as Monolithic Structure (Originally Welded)    Minimization of Motor Detent Torque

According to the requirement, a maximum scan speed erro r of 1% shall realized. In order to achieve this  value, the motor Control Electronic commands the Br ushless DC motor in sine control. However, due to  the fact that the motor detent tor que plays also an important role in  achieving good velocity stability,  additional measures to reduce the detent torque were identified. The mechanism contains two independent motor stators for redundancy reasons. An overall detent torque of up to ±25 Nmm was expected on basis of analysis for the two stators. It wa s found that by rotation of one of the two stators by  half a stator slot pitch (corresponding to 2.5 deg rota tion), the detent torque could be significantly reduced  to a maximum of ±15 Nmm. This design feature was consequently included into the design to further  optimize for minimum velocity disturbances.   NASA/CP-2010-216272