ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
The size of the test article required a shock fixture large enough that it would have numerous modes below  the 10-kHz upper frequency bound for the test. The short impact duration by nature also excites these  modes, resulting in amplification of the input shock at high frequencies. To address the adverse effects of  numerous lightly damped resonances in the frequency band of interest, Moog designed viscoelastic  damping into the shock fixture to control the amplification of these modes and to aid in tuning and controlling  the shock spectrum. The addition of the viscoelastic damping in the design of the shock fixture helped  mitigate high frequency over-tests due to fixture resonances. This fixture used is shown in Figure 7and has  a 0.6 x 0.6 meter (24 x 24 in) surface for integration to test articles with a grid pattern of fastener holes.  Figure 5. Various Shock Bullets  The final technique used for shaping the spectrum for this test was the use of a second fixture to absorb a  significant portion of the momentum transmitted to the main shock fixture during the impact. Transferring  this momentum from the primary shock fixture to the secondary fixture limits the net displacement of the  primary shock fixture and consequently the low frequency shock levels that are otherwise an inherent part  of many mechanical impact shock tests. This technique was developed at Moog starting in 2010 and was  fine tuned over several years to develop a shock test process to provide controlled and repeatable tuning  options with standard fixtures. The availability of standard shock fixtures enabled the CSADA re-test to be  performed with a lesser schedule impact despite the unique demands on SRS shaping. The shock test  setup used for the CSADA is shown in Figure 7.  Shock Test Results  At the start of the CSADA shock effort, the entire shock setup, including a mass simulator for the CSADA  mounted to the Lockheed Martin-provided vibration fixture, was used to tune the shock environment to best  match the required shock dose while limiting the low frequency portion of the SRS to a significantly lower  dose than generated in the previous shock test. After suitable tuning results were achieved, an Engineering  Development Unit (EDU) was installed into the test setup and two shocks were applied to it while highspeed video captured the resulting motion at the slip ring assembly. The review of this high-speed video  data was used to verify that the shock environment no longer resulted in sufficient excitation of the slip ring,  as evidenced by significantly reduced brush motion. After verifying the suitability of the shock environment  with the EDU, the qualification unit was installed and tested.  The CSADA was subjected to three shock events in each of the axes tested (Y and Z). See Figure 6 for  axis definition. The control accelerometer time histories for one of the 3 shocks in the Z axis is shown in  Figure 8. From these time histories, the highly repeatable nature of the shock event can be seen. The pulse  duration of approximately 5 milliseconds matches well with the type of duration expected in a flight shock  event.  NASA/CP—2018-219887 297
The shock response obtained from the original (failed) shock test is shown in Figure 9. Note the large  degree of over-test in the low-to-mid frequency range. The control accelerometer shock response spectra  for the Z direction shocks using the improved test setup are overlaid in Figure 10. The highly repeatable  nature of the shock test method is even more apparent here where the shock dose as a function of  frequency is shown to be nearly identical across all three shocks over the entire frequency band from  100 Hz to 10,000 Hz. The highly repeatable nature of the shock events allowed this test to be run very close  to the nominal input level to further limit the shock dose to the CSADA without significant risk of an out of  tolerance shock event. The Figure 10 SRS overlay also demonstrates the ability of this shock setup to  achieve very low shock levels at low frequencies near 200 Hz where the slip ring assembly showed  susceptibility to severe over test.  For the Y axis test, similar repeatability from one shock to the next, short pulse duration, and low shock  levels at low frequencies were also achieved. The final shock testing of the CSADA resulted in a successful  qualification of the unit and avoided a repeat of the slip ring assembly failure observed from the initial  method of shock testing.   Figure 9. Original (Failed) Shock Test Data versus Input Level, MPE, P95/50, and Vibe Data NASA/CP—2018-219887 299
Figure 10. CSADA Z Direction SRS Overlay  Shock Test Guidelines  At the outset of any shock test effort, it is important to work through the physical implications of the shock  test specification and what it means in terms of the type of event to be simulated. The shape of the shock  spectrum, distribution of energy as a function of frequency, and the total pulse duration can provide vital  information about the type of environment to be simulated. Significant low frequency energy and long  duration pulse widths often correspond to events where thrust is being applied, such as the ignition shock  pulse of a missile. Simulating these environments is a significantly different effort than a high frequency  dominated pyrotechnic event like a spacecraft separation event. Where a significant change in velocity of  the device is to be simulated, similar velocity change should be introduced in the test. Drop tests, etc. often  work very well in these scenarios. Where little or no velocity is to result from the shock event, efforts should  be taken to create a shock environment that minimizes the velocity change and net displacement in the  test. Keeping in mind the type of real world event being simulated helps guide the test engineer through the  process of setting up a good method of replicating the shock. The entire series of events to result in the  simulation of the shock should then be reviewed to verify suitable capability and address challenges. The  required force levels to achieve the shock must be within the excitation system’s range. In the case of  Moog’s lab this equates to pressures at which the pneumatic gun can be fired with high repeatability.  The desired shape of the excitation pulse into the test fixture should be reviewed and initial assessments  made about how to best approximate it. For Moog, this involves looking through the available combinations  of bullets, strike plates, and sacrificial target materials. If an excitation shape is required that is significantly  different from those available, the customer is engaged about needs for custom designs to be created prior  to proceeding through test.  NASA/CP—2018-219887 300
The effects of fixture modes should be assessed for how they can be used to help shape the shock  environment and their effects mitigated when not beneficial to achieving the desired SRS. For certain  environments, it is highly beneficial to introduce fixture modes at certain frequencies. Similarly having  methods to introduce damping into fixture modes is often beneficial in controlling the amplification of the  modes and keeping the SRS within specification at high frequencies. Moog often uses finite element  analysis to estimate the effects of fixture modes and introduces modal damping through methods like  viscoelastic constrained layer damping and particle damping.  Loss of Motor Stability during TVAC  During cycle 20 of 27 TVAC cycles, the CSADA motor stalled at the end of a hot dwell. The stall occurred  when the motor was driven at its “normal” rate (1 rev per day on the output) at its maximum voltage (74 V).  Troubleshooting determined that loss of step stability had caused the motor to stall. This was verified by  driving the motor at 74 V and once stalled, changing the motor drive voltage to 25 V, which resulted in  immediate rotation without any loss of steps. It was also shown that the CSADA would step stably with drive  voltages up to 71 V, and it would quickly lose step stability when driven at 72 V or higher.  Loss of Step Stability:  The CSADA is driven by Lockheed Martin’s common 70-V stepper motor driver. This motor driver provides  voltage excitation for approximately 60 ms each time a motor step is needed. At the normal rate, each  60-ms pulse occurs approximately every 2 seconds. The motor driver includes a diode network across each phase that allows motor current to flow if the Back-ElectroMotive Force (BEMF) voltage is below approximately -8 volts. This provides some additional damping to help the motor damp out after the short excitation pulse. Figure 11. Telemetry Data Showing Loss of Step Stability  Stepper motor stability and dynamic stability analysis are well covered by references 1 & 2. It was  determined that the primary contributors for the CSADA loss of step stability (an example of which is shown  in Figure 11) were the extremely low and variable detent cogging torque and the relatively low motor BEMF  constant. This issue was somewhat masked at ambient and cold by the high detent hysteresis torque  NASA/CP—2018-219887 301
provided by the motor, which, combined with internal friction, provided sufficient damping for stable  operation. After the unit was run-in during the initial TVAC cycles, the slight decrease in bearing and  harmonic drive friction from run-in and the reduction in bearing and harmonic drive viscous drag at hot,  combined with the slight reduction in motor detent hysteresis torque at hot, removed enough damping from  the motor for it to become unstable when driven with an excitation voltage above 72 V.   Stability Model Issues  Clearly, based on the TVAC failure, the original step stability analysis was not sufficient. For the CSADA, a  “heritage” stability model had been utilized, which up to this point had demonstrated a good track record. It  is important to note that the heritage stability model was designed for 4-phase unipolar stepper motors  driven at 70 V by the same motor driver. For conservatism, all combinations of the key parameters were  analyzed at all combinations of minimum and maximum (see reference 1 for stability model methodology  and background). The minimum bearing friction and damping values analyzed were zero, so a loss of  stability was not expected. The model included the dynamic properties of the harmonic drive, output shaft,  and solar array, being modeled as a multiple mass-spring torsional system. After the failure occurred,  digging into the model revealed a serious flaw – the diode circuit in the motor driver was not modeled  correctly. The diode circuit only allows current to flow when the BEMF voltage generated by any motor  phase is less than -8 volts (as depicted in Figure 12). The model, as designed, allowed current to flow if the  BEMF voltage generated by the motor was above +8V or below -8V, so the damping provided by the diode  circuit in the stability model was double what would actually be provided.   Figure 12. Zener Diode I-V Curve Used to Aid in Stability Model Construction Note: CSADA drive electronics include blocking diodes that prevent current from flowing in the  forward bias direction. Once the stability model was corrected, it was validated by inputting the measured or estimated values for  the qual unit that went unstable during TVAC, and comparing the predicted results to the TVAC failure  (Figure 14). The corrected model was able to reproduce the failure, and even showed the same step pattern  as the failure. Furthermore, the threshold voltage for stability with the corrected model was 70 volts, which  is very close to the actual threshold voltage. Predicted BEMF plots were also compared to actual measured  BEMF profiles (Figure 13), and the BEMF profiles matched well.   NASA/CP—2018-219887 302
Figure 13. Stepper Motor Voltage Plot Showing Clipped BEMF Voltage at -8 Volts Figure 14. Corrected Stability Model, Showing the Same Behavior as Original TVAC Failure Once the stability model was corrected, it was clear that the original motor design could not be salvaged.  With the stability model corrected, it also became apparent that the motor suppliers request to reduce the  required resistance, allowing more current to flow through each phase, would not address the stability  problem since this approach would further reduce the motor BEMF constant. In contrast, the motor needed  to have a higher BEMF constant in order to extract more damping from the motor drive electronics diodes,  and a correspondingly higher motor resistance could be tolerated. At this point, the motor supplier’s parallel  development effort to improve detent consistency was not showing signs of positive progress, and the  magnet wire size was already at the lower threshold allowed by internal Lockheed Martin design standards.  Therefore, the effort with this motor supplier was terminated in favor of a new motor design that would  provide a much higher BEMF constant, and was expected to provide more consistent detent properties. 56.8 56.9 57 57.1 57.2 57.3 57.4010203040506070Motor Phase Voltage Plot Time, secPhase VoltagePhase 1 Phase 2 Phase 3 Phase 4 0 20 40 60 80 100 120-10-8-6-4-202 time (sec)Position (steps) NASA/CP—2018-219887 303
Insight Instrument Deployment Arm (IDA) Motor Winding Anomaly  Instrument Deployment Arm (IDA) Hardware Description  The IDA for the InSight Project is utilized to place instruments onto the Mars surface after landing of the  spacecraft. A rendering of the lander on the surface of Mars is shown in Figure 1. The IDA is identified in  the picture and has four axes of motion that utilize the motors under investigation.  Figure 1. InSight Lander Shown on the Surface of Mars  The IDA is exposed directly to the environment of Mars and experiences large changes in temperature from  day to night. This thermal cycling can have deleterious effects on assemblies composed of materials with  different coefficients of thermal expansion. To demonstrate compliance to the motor’s thermal cycle  requirements, a thermal cycle test was performed on three units. To establish design margin, the number  of test thermal cycles is much larger than the quantity of cycles during the mission. This type of thermal  cycle testing is considered a qualification test and the hardware is not flight worthy after going through the  testing. For brushed motor assemblies, the adhesively bonded windings on the rotor and the winding  terminations to the commutator are the primary thermal cycle sensitive items. There are also the bonds  between the magnets and the housing and other ancillary areas that are also stressed during the thermal  cycle testing. One of the three test motors, serial number 041 is shown in Figures 2, 3, and 4.  Instrument  Deployment Arm  NASA/CP—2018-219887 308
Figure 2. Motor Serial Number 041  Figure 3. Brush Assembly  Motor Diameter is 30 mm  Figure 4. View of the Shaft End with the Commutator and Winding Terminations Shown  Thermal Cycle Test Setup  Three motors were selected for thermal cycle testing and placed in the thermal chamber. The motors are  not rotated during the thermal cycle testing to enable accurate measurements of the winding resistance.  The instrumentation was connected to read the winding resistance of the motors as shown in Figure 5 while  the temperature was cycled from hot to cold. The temperature of various locations on the motor assemblies  and voltage drop across the windings (resistance) was recorded continuously during the test. The  resistance of the windings varies directly with the temperature of the test article. The expected appearance  of the recorded data for the windings would show the resistance of the windings rising and falling smoothly  with the temperature of motor as shown in Figure 6.  NASA/CP—2018-219887 309
Figure 5. Test Instrumentation Figure 6. Typical Shape of Winding Resistance  versus Time Showing Smooth Variation with  Temperature Cycles  The Anomaly Appears  The resistance recorded for one of the three motors looked mostly as expected for the first few thermal  cycles of the testing. Some measurement noise was present in the data traces. Two of the motors, however,  did not produce the results expected. The traces instead appeared as in Figure 7. The blue curve is the  temperature and the grey, orange, and yellow curves are the motor winding resistances.  Figure 7. Motor Resistance and Temperature versus Time Plot for Cycle #3  The orange curve is motor serial number 041, the grey curve is motor serial number 042, and the yellow  curve is motor serial number 051. The winding resistance of all of the motors shows electrical noise during  the region of low temperature dwells. The orange motor data performed as expected. The noise within the  data was unexpected and a little troubling, but the most significant feature of the grey and yellow curves is  the sudden change in apparent resistance of the windings during the transitions to and from the cold  temperatures. During the cold transition, the winding resistance should smoothly lower and at no time  should it indicate an increase or sudden jump up in resistance. After 38 temperature cycles, serial number  Temperature/Resistance Time S/N 041S/N 051 S/N 042 Temperature NASA/CP—2018-219887 310
041 adopted the same signature as the other two motors, with a sudden increase of winding resistance as  the temperature was lowered through the cold portion of the cycle (see Figure 8).  Figure 8. Failure of Serial Number 041 at Temperature Cycle #38  To determine the source of the strange increases in resistance, the motor was disassembled and inspected  at high magnification. The motor windings are bonded together and to the rotor core laminations with an  epoxy. During visual inspection under high magnification, some indications of cracks in the epoxy were  detected. The images do not indicate if the cracks are the source of the winding resistance anomalies, but  they are certainly potential locations for trouble. Some of the images taken are shown in Figures 9 and 10.  The cracks indicated in the images were not present in the motor winding epoxy prior to the thermal cycle  testing.  Figure 9. Winding/Rotor Epoxy  Cracks Indicated by Lines and Arrows Figure 10. Winding/Rotor Epoxy -  Crack in Epoxy Shown in Red Oval None of the inspection methods had indicated the actual source of the resistance changes. Some locations  were implicated by the presence of cracks in the epoxy coating, but none of them showed any signs of a  feature that would cause an increase in resistance as the temperature is reduced during the cold  temperature portion of the thermal cycles. The solder joints at the commutator were inspected as well and,  while some of the joints had workmanship issues, none of the connections showed any type of failure that  would support the test data. The brush block/commutator/shaft assembly was X-Ray inspected while  installed in the motor and the brush springs (cantilever arms) showed a proper preload on the brushes as  indicated by the deflection of the spring arm (see Figure 11).  Initial Failure  of S/N 04 1 NASA/CP—2018-219887 311
The next inspection method used was a  Computed Tomography X-Ray ThreeDimensional inspection technique, or “CT  Scan”. The motor assembly was placed on a  plate in the CT Scanner machine that rotates  about a vertical centerline. The motor is  incrementally rotated a small amount and for  each image slice that is taken. Several  thousand image slices are taken, each from a  different position and angle, around the motor.  The multiple slices are then assembled into  the three-dimensional computer model  represented in the figures that follow.  Overall CT Scan views of the rotor assembly  are shown in Figures 12, 13, and 14. Figure 12  shows the rotor assembly with the individual winding wires clearly visible. Figure 13 is a cross-section of  the rotor with one slot on the rotor exposed and the wires within the slot visible in fine detail. Figure 13 is a  view of the brush assembly end of the shaft with the brushes and cantilever arms clearly visible.  The CT Scan inspection method provides high resolution three-dimensional computer models that can be  interrogated in locations within the assembly that are otherwise not possible to inspect. Since no  disassembly is required, the in-situ state of all components can be viewed. Addtionally, any failure condition  that may be disturbed by disassembly can be captured without risk of losing the failure evidence in the  process of removing intervening material.  Close inspection of the CT Scan model revealed the source of the winding failures to be coincident with the  crack sites of the epoxy located on the opposite end of the rotor from the commutator assembly and winding  wire terminations. Figure 15 shows a close-up view of the end of the rotor where the wires exit the slots  and there is indication of damage to the winding wires, identified by the red arrows. Focusing on the areas  of apparent damage, the culprit is finally found to be broken winding wires at the exact locations of the  epoxy cracks. The wire can be seen to have necked down prior to failing, indicating a tensile failure.  Figure 11. X-Ray Image of the  Brush/Commutator Assembly Figure 12. Rotor  AssemblyFigure 13. Section of Rotor  Showing Wires in the SlotFigure 14. Brushes, Commutator, and  Cantilever Arms NASA/CP—2018-219887 312
Further investigation of other epoxy  crack locations revealed several  failed wires and many stretched  wires with necked down crosssections. The necking would  account for resistance changes in  the first few cycles after the epoxy  crack propagated to the location of  the wires. The sudden increases of  winding resistance in the test data  were the result of several broken  regions moving back together as  the temperature was raised. While  the winding resistance might have  appeared nominal at a higher  temperature, the winding wire  damage points would certainly not  have supported a high enough load  current to produce the required  torque output from the motor in the  IDA application.  Figure 16. View Showing Several Locations of Failed Wires. The Necking Down of the Breaks Indicates  Tensile Failures. The Wire Breaks Follow the Cracks in the Epoxy Coating.  The wire diameter in the image is 80 microns.  Figure 15. Magnified View of the CT Scan Model Showing  the End Turns and the Slots in the Rotor Laminations NASA/CP—2018-219887 313