ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 5. Left: Example of a force vs frequency requirement. Right: Example of two frequency bands of a  maximum instantaneous torque versus time requirement.  Figure 5 also shows a maximum instantaneous disturbance versus time requirement. This type of  requirement is more appropriate for (but not limited to) mechanisms with short or intermittent duty cycle.  The type of requirement used is often dictated by instrument needs. In order to verify this requirement,  the time history signal (from analysis or test) needs to be split into different frequency contributions and to  be compared to the requirements in each frequency band, as illustrated in Figure 5.  Microvibration Verification  Which verification methodologies are appropriate?  Verification by test and analysis or a combination of both approaches may be more suitable in different  circumstances.  Verification by analysis can be very useful in the early stages of the design when there is no hardware  available yet. In addition, in some cases performing a test on the complete system is not feasible. For  example, assessing the induced microvibration by the coupled system of a solar array and a SADM is no  possible by test. In this case, analysis or a combination with test data at SADM level is more appropriate.  Lastly, verification by analysis may be required if no suitable test facility is available due to the required  sensitivity, frequency range or physical size of the test subject.  Verification by test is advisable when representative hardware is available, as well as a suitable test  facility. It can validate (part of) the analysis model. Testing is particularly useful to assess the impact of  environmental tests or different environmental conditions on the microvibration performance of the  mechanism.  How to perform verification by analysis?  Microvibration verification by analysis is typically performed in a two-step approach:  •mechanism level (source of disturbance) •system level The objective of the mechanism-level analysis is to estimate the expected exported torque/force between  the mechanism and a hard-mounted interface during operation, represented by F EXPORTED  in the diagram  (Figure 6). F IN represents internal mechanism forces/torques such as electric motor torque.   NASA/CP—2018-219887 196
Figure 6. Mechanism-level exported microvibration model  The objective of the system-level analysis is to estimate the expected performance of the disturbanceaffected equipment (e.g. pointing performance of optical equipment) when input mechanism disturbance  is applied at the relevant mechanism to spacecraft interface point. For such an analysis, either a blocked  force approach or coupled mechanism-spacecraft model can be used to simulate the input disturbance  (Figure 7).  Figure 7. System-level model. Right: Coupled mechanism & S/C model. Left: Blocked force model.  An example model developed in MATLAB/Simulink is shown in Figure 8. The standard building blocks of  such a microvibration model may consist in:  •Motor drive electronics oCurrent control oPosition control oOutput voltage to be applied on motor •Electric motor model oStepper, Voice coil, brushless DC etc. oCoil dynamics (R, L, back-EMF, detent, eddy current losses) oRotor dynamics •Mechanism components oGearbox (transmission error, backlash, friction) oOther non-linear force contributors: Slip-ring, Cable wrap •Structural parts oStructural parts with significant modes in frequencies of interest represented as state space matrix of flexible body exported from FEM tool oTypical parts modeled as FEM: Doors, Solar array, Scanner etc. NASA/CP—2018-219887 197
Figure 8. Example of a microvibration model in MATLAB/Simulink Verification by analysis is an integral part of the overall verification process at unit level and especially  important at system level, due to the complexity of performing system-level performance testing under  gravity conditions. As such, an effective way to accurately assess system-level microvibration  performance can be to perform simplified system-level tests, correlate the results with the simulation and  perform the detailed performance assessment using the correlated analytical model.   This process can provide an extremely accurate prediction of in-flight performance, as shown by the  example in Figure 9 of in-flight spacecraft jitter data versus analytical prediction performed during initial  commissioning of a spacecraft.   Figure 9. Comparison of inflight S/C jitter vs analytical estimate (Plotted as FFT)  Motor drive electronics Electric motor model Gearbox Flexible harness  Structural model  NASA/CP—2018-219887 198
How to perform verification by test?  Verification by test at mechanism level is typically performed on a dynamometer, often referred to as  Kistler table (after the brand name of the predominant supplier). To illustrate the capabilities and  limitations of this type of equipment, two facilities are discussed here: the Reaction Wheel  Characterisation Facility (RCF) (Figure 10) and the 6-DOF microvibration test facility (6dMVMS) (Figure  11). Even though the RCF is named after RWs, it can be and has regularly been used for other  mechanisms as well. Figure 12 shows the specification of both facilities. Test results using the relatively  new 6dMVMS have already been reported elsewhere [10].  Figure 10. The RCF facility at ESA [8]  Figure 11. The 6dMVMS facility at ESA [9]  NASA/CP—2018-219887 199
Figure 12. Test facility specifications. Left: RCF specifications. Right: 6dMVMS specifications [8] [9]  The RCF system is composed of the following subsystems [8]:  •Instrumentation of the facility and interconnection to the data handling system •Test table providing a rigid interface between the specimen and the instrumentation •Marble block acting as seismic reaction mass and hosting the test facility •Pneumatic isolation system to decouple from disturbing environmental ground vibrations •Vacuum bell equipped with a dry pumping system and a vacuum gauge •Signal conditioning, data acquisition and processing system The microvibration measurement system is composed of two platforms [9]:  •a measurement and excitation platform (lower platform) •a vibration isolation platform (top platform) Each platform is formed from a Minus-K passive isolation system and an array of sensors and actuators  which provide 6 DOF control and measurement. The measurement platform is a bespoke design, and has  two modes:  •measurement, where the platform measures the 6 DOF force/torque exerted on the platform; •excitation, where the platform subjects a specimen to a controlled 6 DOF microvibration environment. The platform is designed to allow measurements and excitation at low frequencies (< 0.03 Hz) and very  low amplitudes (μN).  The facilities both have unique capabilities missing in the other facility. The RCF can be equipped with its  vacuum bell to perform measurements in vacuum if required. The 6dMVMS can excite equipment and is  therefore uniquely capable of verifying microvibration susceptibility requirements by test. The frequency  range of both facilities is complementary, with only relatively small overlap, although an extension of the  frequency range for the 6dMVMS is currently under investigation.  There are several good practices to take into account when verifying microvibration requirements by test:  •Modes of the test set-up can influence the data. This usually limits the frequency range and mass of the unit under test. For the RCF, the modes of the empty table are higher than 1250 Hz. With a payload, the unit under test and its MGSE, of 10 kg this drops to ~1000 Hz and even lower for NASA/CP—2018-219887 200
7. Richardson, G., Seabrook, T.  "Ground and on orbit microvibration verification of the triplesat 1m resolution earth imaging spacecraft constellation."  ECSSMET 27-30 September 2016. 8. Wagner, M., Selle, J.  “Reaction wheel characterisation facility – Facility description.” 24 October 2013. 9. Wagner, M., Selle, J.  “6 degree of freedom micro-vibration test facility – Facility description.” February 2017. 10. Wagner, M., Selle, J.  “First real-life results of novel micro vibration measurement facility.” ESMATS, 20-22 September 2017. 11. Smet, G., Richardson G., McLaren S., Haslehurst, A.  "Managing reaction wheel microvibration on a high resolution EO small spacecraft." ESMATS, 25-27 September 2013. NASA/CP—2018-219887 204
Investigation of Bolt Preload Relaxation for JWST Thermal Heat Strap Assembly  Joints with Aluminum-1100 and Indium Gaskets  Andrew Bartoszyk*, Thomas Walsh**, Jody Davis*, Jason Krom+, Patrick Williams**, Jason Hylan* and Edgar Hemminger++ Abstract  Accurately predicting fastener preload relaxation in the James Webb Space Telescope (JWST) heat strap  assemblies is essential to insure adequate thermal performance during its mission lifecycle. The  mechanisms for preload relaxation in the strap joints include Al-1100 material creep, indium gasket flowout, and embedment of the joint faying surfaces. This report documents the results from a bolted joint  relaxation test, including analysis and curve fitting of the test data for predicting preloads five years after  initial torque application. The report also includes the derivation of a preload uncertainty factor enveloping  both torque/preload application scatter and expected preload relaxation at the end of mission life.  Introduction  The JWST thermal control system includes Al-1100 heat straps for controlling heat flow to and from  temperature critical instruments in order to meet instrument performance requirements and optimize  observatory science capability. Critical to heat strap performance is preload maintenance at the bolted joint  interfaces over mission lifetime. To improve thermal conduction through the bolted interfaces, Al-1100 and  indium gaskets are integral in many of the joints. Although these materials are common for improving  thermal conduction in aerospace bolted joints, Al-1100 creeps under load at room temperature and indium  gaskets can gradually flow out. Creep and flow-out will relax joint preload and therefore degrade joint  thermal conductance over time. Many of the heat strap joints also include threaded inserts in Al-1100,  resulting in high stresses in the parent material around the insert and exacerbating Al-1100 creep.  Figure 1 illustrates a representative flight joint for the JWST heat strap assemblies.  Figure 1. Representative Heat Strap Bolted Joint Interface  Because Al-1100 is relatively low in strength, compression of Al-1100 under washers and threaded insert  pull-out limit the design space for the initial bolt preload. This limited design space necessitates a rigorous  approach for predicting joint preloads in the heat straps throughout the joint life cycle.   *NASA Goddard Space Flight Center, Greenbelt, MD ** Stinger Ghaffarian Technologies, Greenbelt, MD +Sigma Space Corporation, Lanham, MD ++ Ares Corporation, Greenbelt, MD Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018  NASA/CP—2018-219887 205
In order to quantify preload loss over time, a comprehensive joint relaxation test was completed. The test  includes a set of representative flight-like joints, a control set, and various joint configurations in order to  compare preload loss between different drivers for joint relaxation. The different joint coupons were torqued  to the same nominal preload and the preload relaxation was measured over a period of eight months. The  test data was then processed in an Excel spreadsheet for future preload loss predictions. Uncertainty  factors were also derived for calculating max and min expected preloads, accounting for both scatter in  initial preload application and joint relaxation over time.  Test Description  Table 1 lists the joint configurations tested. For each group, with the exception of the control group, a set  of 3 coupons were tested with the bolts re-torqued 24 hours after the initial torque-up and a set of 3 coupons  were not re-torqued. For the control group, only 3 coupons were tested without any re-torqueing of the  bolts. The coupons included #10-32 A-286 fasteners, phosphor-bronze threaded inserts (2D length), and  lubrication of external/internal threads with Braycote. All coupons were torqued to a nominal preload of  1550 N (350 lbf).  Table 1. Coupon Test Configurations  All test coupons included a FUTEK piezo-electric load cell for preload monitoring. The data acquisition  system included the SENSIT Test and Measurement Software. The program was set to sample load cell  outputs every 0.1 second during initial torque-up for each joint. After initial torque-up, the load cells were  sampled every 10 minutes. Test data was downloaded on a daily basis for monitoring and analysis. A  separate load cell was put under a known weight and monitored for several months to confirm that the  load cell read-out did not drift over time.  Joint Relaxation Test Data Analysis  The physical mechanisms driving preload loss in a thermally conductive bolted joint with Al-1100 and an  indium gasket are varied and complicated, making it difficult to derive a simple mathematical model for  predicting joint relaxation. An alternative approach is to curve fit data from a joint relaxation test and  extrapolate the empirical model for predicting future preload loss. After curve fitting the test data with several  candidate functions, the natural logarithm function proved to be the most accurate for fitting the data across  the various coupon test configurations.  Configuration Description  Control Al-6061 top plate compressed against an Al-6061 bottom plate  Indium 0.005-in (0.13-mm) thick Indium foil between Al-6061 plates  1100C Al-1100 plate compressed against an Al-6061 bottom plate  1100T Al-6061 plate bolted into an Al-1100 plate, putting threads in tension  Foil Multiple layers of 0.002-in (0.05-mm) thick high-purity aluminum  (99.999% Al) in direct compression against an Al-6061 bottom plate  Flight Combines multi-layer aluminum foil in compression, 1100 plate in  compression, Indium in compression, 1100 in tension 60616061 60616061 Indium 60611100 1100 w/ inserts6061 6061Foil Foil 1100 w/ inserts1100 Indium NASA/CP—2018-219887 206
Using the method of least squares, the data was processed in Excel and the software’s Solver Tool was  applied for minimizing the sum of the chi square values (the square of the difference between the real data  and the predicted data). Since the data acquisition system sampled preload every 10 minutes for 32 joint  coupons and the test duration was more than 8 months, the spreadsheet was set up to sample the test  data every 24 hours. This minimized the processing time for each Solver Tool run and kept the spreadsheet  at a manageable size. A number of curve fits were also processed with higher analysis sampling rates in  order to check the nominal analytical sampling rate. Additionally, the coefficient of determination (R2) was  calculated for each curve fit to quantify how well the test data fit the curve. Figure 2 is a plot of the test data  and curve fit for the Flight-1 coupon.   Figure 2. Test Data and Curve Fit for Flight-1 Coupon  Joint Relaxation Test Results  Figure 3 plots the average preload predictions after 5 years, with 2-sigma error bars calculated from the  test data. The results from this test are specific to the joint configurations tested (materials, thicknesses,  configurations, initial nominal preload, etc.) and apply only to these specific configurations. Caution should  therefore be exercised when extrapolating the results to design deviations from the specific configurations  tested. As can readily be seen in Figure 3, the flight configuration (with 1 re-torque), is predicted to lose  ~25% of its initial preload (average + 2-sigma). Without the re-torque, the preload loss is >40%. For  comparison, the control configuration, a typical joint for structural applications in flight and ground support  hardware, is predicted to lose ~5% of its initial preload after 5 years. The test results also show that the  biggest drivers for joint relaxation are the Al-1100 in tension (inserts embedded in Al-1100) and the highpurity aluminum foil in compression. Joint relaxation from indium flow-out appears to benefit the most from  a single re-torqueing of the bolts (from ~25% to <10%).   354.7 280290300310320330340350 0 1000 2000 3000 4000 5000 6000Preload (lbf) Time (hrs) Test Data CF NASA/CP—2018-219887 207
Figure 3. Joint Relaxation Preload Loss Predictions (5 years out)  The plot in Figure 4 compares the curve fits for the different configurations. The plots only include the  coupons that were re-torqued, with the exception of the control coupons. Figure 4 shows the curve fits for  preload loss after approximately 6000 hours of testing (~8.3 months), the flight coupons were losing  ~0.25 lbf (1.1 N) of preload per week and the control group was losing <0.1 lbf (0.4 N) of preload per week.  Figure 4. Curve Fits for Preload Loss Over Time  0%5%10%15%20%25%30%35%40%45% Control 1100C Indium Foil 1100T Flight% Preload Loss Coupon Type% Preload Loss (5 years out) No Retorque 1 Retorque after 24 Hrs 280290300310320330340350360 0 1000 2000 3000 4000 5000 6000Preload (lbf) Time (hrs) Control Flight Indium 1100C 1100T Foil NASA/CP—2018-219887 208
Preload Uncertainty Factor  A torque/preload uncertainty factor (UF) is typically included in bolted joint analysis for calculating the  expected maximum and minimum preloads after torque application to a design nominal preload. The UF  for a given joint configuration can be recovered from torque/preload test data with the assumption that the  preload probability distribution function (PDF) is Gaussian. NASA-STD-5020 [1] recommends the recovery  and use of a B-basis (90% of population with 95% confidence) UF for bolted joint analysis.   In addition to the torque/preload UF, the calculated max/min preload must also account for preload loss  over time, especially when preload is critical for joint performance. A total UF, accounting for the scatter in  preload from both the initial torque-up of a bolt and relaxation over time, is derived in order to simplify the  final bolted joint analysis and retain the use of existing in-house analysis programs. Preload changes due  to temperature effects are not included in the UF since it is calculated separately.  In order to easily combine the preload scatter for the nominal installation preload and joint relaxation, the  PDF for preload loss from joint relaxation is assumed Gaussian and independent from the initial torque-up  preload PDF. This is a reasonable assumption if the torque/preload UF is low. The average relaxed preload  after torque-up and preload loss is simply the nominal initial preload minus the average predicted preload  loss across all coupons for a given configuration.   _ =  1−% 	 	(1)  Since the initial preload and preload loss are uncorrelated random variables and their PDFs are Gaussian,  the B-basis UF for the relaxed preload PDF is simply the RSS (root sum square) of the UF values for the  initial preload PDF and preload loss PDF.  = +(2)  A total UF− can then be derived for the minimum expected preload.   =  1−% (1−  ) (3)   =  1−% −+% 	 	 (4)   =  (1−) (5)  = % +−%  (6)  Figure 5 illustrates the relationships between UF− and the nominal preload.  Figure 5. UF− and Nominal Preload  UF+ is simply equal to the torque/preload UF, since the max preload for any given time is before joint  relaxation progresses.   =  (1+) (7)  =  (8)  NASA/CP—2018-219887 209
Validation and Test Results of the 2.5D Technology for Pancake Slip Rings  Mélanie Henry* and Sandro Liberatoscioli* Abstract  A slip ring assembly project has been performed to demonstrate that in future space mechanisms a slip  ring assembly based on 2.5D printed circuit board (PCB) technology can replace the currently used risky  and costly design. Four units with 2.5D technology have been manufactured, assembled and tested for  an ESA ARTES 5.2 project to validate this technology at a level of TRL5.  The 2.5D PCB technology slip rings have undergone all the mechanical and environmental tests without  showing any degradation. During the TVAC test, some lines have shown high dynamic resistance  performance. After investigation, it was confirmed that this degradation (noise) was linked to a  contamination problem.  The same level (even better) of performance is achieved with this type of technology compared to the  standard one.  2.5D Slip Ring Definition and Validation  The data and power transfer within space mechanisms, such as a Solar Array Drive Mechanism (SADM),  is in principle ensured by a contact technology. These components known as slip rings are of two types:  cylindrical and pancake. The pancake slip ring concept is the one discussed in this paper. The pancake  solution is made of electrical conducting tracks arranged on a flat disk as concentric rings centered on the  rotating shaft (Fig. 1). The planetary configuration allows optimization of weight, volume and electrical  features of the circuits.  The 2.5D PCB slip ring is a planetary slip ring composed of  two sub-assemblies:  •The stator assembly, the fixed part of the slip ring. •The rotor assembly, the rotating part of the slip ring. The contact between the stator and rotor is ensured by wire  brushes, soldered on the stator side on a PCB and  preloaded to be in contact with the track during the whole life  time.  Figure 1. 2.5D slip ring overview  The main challenge of the project was to manufacture the rotor sub-assembly in only one part (PCB  technology) to remove all the operations of gluing, machining, potting and soldering that increases the  manufacturing activities. Therefore, the novelties of this innovative technology are:  oTo apply a thickness layer of pure copper up to 500 µm based on PCB technology, this in order to manufacture a groove for brush guidance. oTo apply a hard gold plating on the tracks (according to a space-qualified RUAG process) oTo obtain the insulation barriers between tracks during PCB manufacturing operations oTo create a guidance groove in each track during PCB manufacturing operations oTo have a thermal drain inside the PCB to dissipate as much as possible the heat at the interface oTo choose the better candidate between different possibilities for the insulating barrier material *RUAG Schweiz AG, Nyon, Switzerland Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018 NASA/CP—2018-219887 211
In addition, special attention was focused on the stator sub-assembly to simplify the integration phase. A  dedicated stator PCB was designed to improve the thermal dissipation of the slip ring and reduce the  variation of current distribution into the wire brushes. The main modifications are showed in Figure 2 and  Figure 3.  Figure 2. Rotor sub-assembly improvement  Figure 3. Stator sub-assembly improvement  The main benefits due to these improvements are:  -Reduction of manufacturing process time (no gluing, no adjustment of wire brushes, direct soldering) by about 50% -Better reliability (controlled processes) -High thermal dissipation on both sides (rotor and stator) -Better current repartition into the wire brush (decrease of creep effect) -Possibility to be modular Fixation internal and external  only with screws (no potting).  NASA/CP—2018-219887 212