ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
460 Random Vibration Testing  Random Vibration tests were conducted to validate the structural design of the spacecraft and also to  validate the deployer mechanism design. The random vibration test would also be used to validate the worm drive locking feature. Figure 17 shows LightSail-1 on the vibration table at Cal Poly.  

Figure 17. LightSail-1 during random vibration testing at Cal Poly.  The structure and worm drive locking feature performed well. The worm drive locking feature was verified by observing that the spindle orientation had not changed after random vibration testing by lining up reference marks on the spindle and deployer top plate before random vibration testing. If the reference marks lined up after random vibration testing it confirmed that the spindle was held fixed by the worm drive, which is what occurred.  

Figure 18. Spindle reference marks before (left) and after (right) random vibration testing.  While the structure performed well and the worm drive locking feature was validated, it was observed that the booms were pulled inside the deployer ~10 mm during the random vibration test. It was determined that the boom wrap was allowed to balloon slightly due to some compliance in the flexure spring/tensioner assembly. This was solved by packaging the booms in such a way that the tensioner was completely bottomed out against a hard stop while the boom wrap was pulled tight. Additionally, a pin was added to the tip of each boom to act as a hard stop to not allow the boom to retract into the deployer.
462   Figure 21. Sail deployment test.  Cold Deployment Testing  Cold deployment testing consisted of cooling only the deployer and booms to a prescribed temperature  and then deploying only the booms while monitoring the drive motor current and comparing that with the current required to deploy the booms at room temperature. It was decided to focus on the deployer mechanism and booms, without the sail, because the sail behavior was not expected to change significantly over the temperature range being tested and because of the difficulty controlling moisture in a laboratory environment (without a vacuum chamber). The moisture in the system needed to be carefully controlled since the cold testing was conducted below freezing and the introduction of ice in the system could limit the performance of the deployer. The cold testing was performed by flowing cold nitrogen vapor into a container housing the deployer with small cutouts for the booms to pass through until the deployer assembly reached the appropriate temperature after which the booms were deployed. The drive motor current was monitored during cold deployment testing for comparison with room temperature  deployment data. The deployer was qualified down to -6 ˚C where it was well within the allowable motor  current margin.   Many full-scale deployment tests have been conducted on the engineering model hardware for  development. Full-scale "acceptance" deployment tests will be conducted on the flight hardware including before and after random vibe and TVAC. As of this writing the flight unit TVAC test has not been completed. These flight hardware tests are held to a minimum because of degradation of the sail material during re-folding of the sail after deployment testing. However up to 6 deployment tests on a set of sails is reasonable without significant degradation.
463 Conclusion  The result of work done on the LightSail-1 program is a fully qualified 32-m2 solar sail packaged in a  ~1.5U volume with a mass <3 kg (4.6 kg total spacecraft mass). The boom deployer assembly has a  deployed to post deployed ratio of 80:1. It is planned to launch LightSail-1 by the end of 2012. 

Figure 22. LightSail-1 fully deployed in the lab at Stellar Exploration.  Acknowledgments  The authors would like to thank The Planetary Society and its members for sponsoring LightSail-1, Louis Friedman and Jim Cantrell for hiring Stellar Exploration as the sail module developer and spacecraft integrator, and Bill Nye (The Planetary Society's Executive Director) for his continued support of the LightSail program. The authors would also like to thank Jeremy Banik of AFRL for his guidance and support during the development of LightSail-1.  References: Anderson, J.L. (11/29/11). NASA. NASA's NanoSail-D 'Sails' Home -- Mission Complete. Retrieved  January 3, 2012, from http://www.nasa.gov/mission_pages/smallsats/11-148.html  McInnes, C. R. (1999). Solar Sailing Technology Dynamics and Mission Applications. Chichester, U.K:  Praxis Publishing.  Montgomery, Edward E., Adams, Charles L. (2008, April). NanoSail-D. CubeSat Developers Workshop,  San Luis Obispo, CA.  The Planetary Society. Projects: LightSail-Solar Sailing. Retrieved January 3, 2012, from  http://www.planetary.org/programs/proj ects/solar_sailing/lightsail1.html
469 geometry, the hinge causes each of the panels to deploy 90 deg and to twist 45 deg in an axis normal to  the initial deployment axis.   

Figure 3 – Third generation actuator showing interconnect PC board, tell-tale harness, kickoff  spring and retainer, and upper interface assembly with “near spherical bearing”  Because of this and because of the small radius the plug travels during deployment, the secondgeneration actuator had the tendency to bind during deployment. The reduction of the cylindrical interface area between the cup and plug reduced the likelihood of binding due to any lateral load or misalignment during actuator release. Binding was also reduced by changes to the lead in angles on the outer diameters of the cup and plug. The taper on the inner diameter of the plug reduces the probability of the cup hanging on the kick-off spring during deployment. The change to a two-piece plug, consisting of an annular plug with a cap allows insertion of the plug without compressing the kick-off spring. This eliminates the possibility of binding and galling while setting the actuator for flight, greatly facilitating reassembly. Figure 4 shows some of the features of the third generation actuator. 

Figure 4 – Four views of the third generation device.   Electrical  Interface Board Kick-off Spring Dual cone & pin  2-piece cup  Kick-off  spring and retainer  Tell-tale  jumper Ridges reduce  thermal contact between cup and plug Kickoff spring & retainer Tell-tale jumper
471 Table 2- Test data from second generation actuator.  Test #  Vacuum  (Torr)  Current  (A)  Voltage  (V)  Power  (W)  Tamb  (°C)   Trel  (°C)   ΔT  (°C)  Δtime  (s)  ΔT/Δt  Energy  (W-s)  % Batt  Energy  Note  4 2.50E-01 2.00 5.60 11.2 19.1 94.0 74.9 147 0.51 1646 1.02% 5  11 2.50E-02 2.16 6.05 13.1 23.6 94.0 70.4 100 0.70 1306 0.81% 4  17 2.50E-02 2.16 6.05 13.1 23.6 93.6 72.0 100 0.72 1320 0.82% 4  12 2.50E-02 2.16 6.05 13.1 21.1 104.0 82.9 123 0.67 1607 1.00% 5  18 2.50E-02 2.16 6.05 13.1 19.3 91.0 71.7 115 0.62 1502 0.93% 4  5 2.50E-01 2.58 7.22 18.6 24.0 87.8 63.8 70 0.91 1305 0.81% 4  6 2.50E-02 2.60 7.28 18.9 18.5 92.0 73.5 68 1.08 1287 0.80% 5  7 2.50E-02 2.60 7.28 18.9 18.9 89.2 70.3 66 1.07 1249 0.77% 4  8 2.10E-02 2.60 7.28 18.9 19.5 100.3 80.8 69 1.17 1306 0.81% 5  9 2.10E-02 2.60 7.28 18.9 20.0 90.2 70.2 62 1.13 1174 0.73% 4  10 2.50E-02 2.60 7.28 18.9 23.5 91.0 67.5 61 1.11 1155 0.72% 4  13 2.50E-02 3.02 8.46 25.5 18.9 88.7 69.8 46 1.52 1175 0.73% 5  14 2.30E-02 3.02 8.46 25.5 19.9 80.2 60.3 39 1.55 996 0.62% 4  16 2.50E-02 3.02 8.46 25.5 23.8 80.6 56.8 38 1.49 970 0.60% 4  Not a “standard” power case.   Low power case   Nominal Power Case   High Power Case   Outlier  

We performed a preliminary analysis of only the separations performed in vacuum. A B-basis time to  separate was calculated for the whole data set, and each of the three battery state of charge cases. The  B-basis time to separate was calculated as follows:  T2σ = Mean + (2 * σ ) where T 2σ is the 2 sigma time to  separate, Mean  is the arithmetic mean of the times to separate, and σ is the standard deviation between  the times to separate. Table 3 lists the data for the four cases analyzed (all vacuum separations, nominal  power, high power, low power).  Table 3 - Second generation statistical results.  Data Set Mean Time  to Separate Standard  Deviation 2-sigma time to separate  All vacuum 78.86 33.04 144.94  Nominal (2.6A) 65.2 3.56 72.33  Low (2.16A) 109.5 11.45 132.39  High (3.06A) 41 4.36 49.717 

The lessons learned from the 2nd generation testing were:   This version of the actuator was too difficult to set for flight. The operator was required to precisely  align the plug in the cup and then to insert it into the cup with enough force to overcome the preload  spring. Having to push the plug into the cup with force makes the actuator too susceptible to galling.   Performance degrades over time because of galling produced during resetting operations. Both high  pressure surfaces needed to be hard coated.   The actuator was susceptible to racking when coupled with the reference mission solar array  engineering model.
472  The amount of energy required for release needed to be further minimized for the 3rd generation  design.   The tell-tale function of the actuator was determined to be reliable through all testing.     Installation of the plug is critical to the consistent operation of the actuator. We determined to make  insertion easier and more repeatable by a redesign that eliminated the preload spring force during  insertion.   Use the highest possible insertion temperature during assembly.  Third Generation   The first 4 units fabricated were fabricated to select which actuator coatings should be employed for follow-on use of the actuators fabricated for flight. Actuator SN002 featured Chemical Conversion Coating (MIL-C-5541, Class 3) coating on the aluminum cup part and Tiodize, Type II (AMS-2488, Type 2) on the titanium plug part. For SN002 (Engineering Model), we conducted 7 separations, and then we retired the unit.  Actuator SN004 featured Hard Coat Anodize (MIL-A-8625, Type 3) on the aluminum cup part and Tiodize, Type II (AMS-2488, Type 2) on the titanium plug part. We conducted 6 separations, the self-actuation test, the static load test, and 2 more actuations. The third generation tests used an engineering model CubeSat solar array to represent the correct boundary conditions for the device. Figure 6 depicts the solar array wing and test fixture. The SN004 actuator had much better performance, and its   tolerances and coating combination were selected for the final design.  We characterized performance of the SN004 actuator in the following sequence:   A single functional test at ambient conditions to verify workmanship (test 4.1). Time and delta-T were  recorded for all functional tests.   Two separations in vacuum at nominal power (tests 4.2 and 4.3).   A separation in vacuum at high power (4.5).   A separation in vacuum at low power (4.6).   A self-actuation test to determine the maximum temperature the actuator can be exposed to without  spontaneously actuating (4.7). Temperature at separation was recorded.   A static load test to verify that the actuator can withstand launch loads. Time and load were recorded  (4.8).   Two separations in vacuum at nominal power to verify that the device was not affected by the static  loads test (4.9 and 4.10). 

Figure 6 – The solar array wing and test fixtur e duplicated the boundary conditions for flight  actuator testing.  

Substrate Stiffener Actuator PlugHinge Assembly  Antenna Blade,  Captured by Solar Array Actuator Cup Assembly
473 The test at ambient conditions verified that workmanship was acceptable. The two tests at ambient  temperature in vacuum showed the expected reduction in delta-T and time to separate. These results also showed (despite the small number of tests) a tight data grouping, showing that the device is repeatable. The tests at 3 A and 2.16 A showed that, as expected, the time to separate is inversely proportional to input current, when the tests are run at the same initial temperature. Somewhat surprisingly, the delta-T for the 2.16 A case was lower than those of the 2.6-A cases. Figure 7 shows the test stand inside the vacuum chamber.  Figure 7 – Test stand, actuator and CTERA shown in vacuum chamber.   Following the tests at 2.16 A and 3 A, we conducted the self-actuation test. The purpose of the self-actuation test was to determine the temperature at which the actuator would release a deployable structure while unpowered. The actuator was required to reach a temperature higher than 50°C before self-actuating to demonstrate a 10°C temperature margin. The ambient pressure temperature cycling chamber was set to ramp temperature at a rate of 1°C per minute, starting at room temperature, which was approximately 19°C. We monitored the tell-tale circuit until the device actuated at 65.9°C. The highest predicted temperature for the CubeSat prior to solar array deployments is 40°C, which gives a margin of 25.9°C on self-actuation.  Following the self-actuation test, we conducted a static loads test. The purpose of the static load testing  was to determine if the selected coatings and tolerances of the serial number 004 actuator can withstand launch loads after 7 separations. Note that the interference fit for the actuator tested, SN 004, is significantly smaller than those of the flight devices (SN 011-018). Since holding capability is a function of the interference fit, the flight actuators are able to react more load than the SN 004 device. The fixture used for static load testing is depicted in Figure 8. Tension on the actuator was increased in ¼ load increments and held for at least 30 seconds until the final load of 160 N (36 lb) was reached. Load was maintained at or above 156 N (35 lb) for 5 minutes. Following the static loads test, the CTERA was cycled two more times at ambient temperature and nominal power in vacuum (tests 4.9 and 4.10). Results from these two tests indicated that the electrical and thermal performance of the device wasn’t affected by the static loads test. Results from tests 4.2, 4.3, 4.9 and 4.10 are tightly grouped. The raw data from the tests are presented in Table 4
474 Figure 8 – The CTERA withstood 1.5 times the maximum expected flight load for 5 minutes.   Table 4 - Non-flight third generation actuator data, SN 004. Test #  Vacuum  (Torr)  Current (A)  Voltage (V)  Power (W)  Tamb (°C)   Trel (°C)   ΔT  (°C)  Δtime (s)  ΔT/Δt  Energy (Ws)  % Batt  Energy  Interference  Fit  Note  4.1 ATM 2.60 7.28 18.9 19.5 71 51.5 33 1.56 625 0.39% 0.0003    4.2 2.50E-02 2.60 7.28 18.9 18.7 45.0 26.3 18.5 1.42 350 0.22% 0.0003    4.3 2.50E-02 2.60 7.28 18.9 18.5 44.0 25.5 18 1.42 341 0.21% 0.0003    4.9 2.50E-02 2.60 7.28 18.9 18.4 47.8 29.4 17 1.73 322 0.20% 0.0003 1  4.10 2.50E-02 2.60 7.28 18.9 18.5 46.0 27.5 17.5 1.57 331 0.21% 0.0003 1  4.5 2.50E-02 3.02 8.46 25.5 18.5 39.8 21.3 12 1.78 306 0.19% 0.0003    4.6 2.50E-02 2.16 6.05 13.1 18.8 42.2 23.4 26 0.90 340 0.21% 0.0003    4.4 7.55E+02 2.60 7.28 18.9 UN K UN K UN K UN K UN K UN K UNK 0.0003 2  4.7 7.55E+02 0.00 0.00 0.0 18.9 65.9 47.0 NA NA NA NA 0.0003 3  4.8 7.55E+02 0.00 0.00 0.0 19.0 NA NA NA NA NA NA 0.0003 4  1 After static load test.  2 Actuated while adjusting current delivered to the resistor.  At ambient.  No data recorded.  3 Will not self actuate until ~66°C. SV will not go above 40°C.  Adequate margin.  4 Static Load Test.  Requirement (with test factor 1.5) was 143.5N (32.25#).  Held above 157N (35.3#) for  5 minutes. 

When the results from SN 004, with AMS-2488, Type 2 coating for titanium and hard coat anodize coating  for the aluminum were compared to those for SN 002, with the same titanium finish and chemical conversion coating (MIL-C-5541, Class 3) of the aluminum, the SN 004 device was the clear winner.   SN 002 was tested first, and we concluded testing after the first 7 cycles because the data were poorly  grouped, the performance of the device was perceptibly deteriorating, and we observed increasing damage to the interfacing surfaces of the aluminum and titanium parts. SN 002 raw data are not presented due to space restrictions.   Once the final design was selected, we fabricated eight flight units and conducted the following tests in  the following order to validate the design. The flight mechanisms (SN 011 through 018) underwent the following testing to further validate the design and workmanship, and the raw data from these tests are presented in Table 5:
476 We recommended to our team that 2 sigma durations be programmed to assure that timed separations  are successful. Table 6 lists the data for the four cases analyzed (2.6 A at ambient temperature in vacuum, 2.16 A at -20°C in vacuum, ~2.15 A during thermal balance testing at 0°C, ~2.25 A during thermal cycling in ambient pressure at 50°C).  Table 6 – Rudimentary statistical analysis of CTERA temperature, time and energy data  Vacuum, 2.6A, 20°C Vacuum 2.16 A, -20°C    mean std dev X + 2 σ   mean std dev X + 2 σ  Time 54.0 7.2 68.4 Time 130.1 7.4 144.9  dT 65.4 6.6 78.5 dT 101.7 15.0 131.6  Energy 1022.1 136.0 1294.1 Energy 1699.9 96.6 1893.1  Vacuum, Thermal Balance, 2.15A,  0°C Ambient, Thermal Cyclling, 2.25A, 50°C    mean std dev X + 2 σ   mean std dev X + 2 σ  Time 75.0 4.1 83.2 Time 65.8 12.4 90.6  dT 87.0 4.32 95.6 dT        Energy 1292.8 52.8 1398.4 Energy 941.6 194.8 1331.2 

We had already confirmed that actuation in vacuum reduces the energy required to actuate (see the  difference between test 4.1 and tests 4.2-4.10). Looking further at the data, we noticed a correlation between input power and time to separate, and a correlation between initial temperature and time to separate. The following several figures pictorially represent relationships between initial temperature, delta-T, time, power level, and the different interference fits of the various actuators.  The left chart of  Figure 9 shows that release time is directly related to the energy expended, as one would expect. It also shows a correlation between the power applied to the device, illustrated here by various current levels, and time to release. The faster energy is applied to the device (more power), the less time and energy it takes to operate. The right chart of Figure 9 shows the relationship between release delta-T and power. As power increases, release delta-T decreases. It appears as the power applied to the high CTE part is increased, its rate of expansion is higher and less heat is applied to the low CTE part. Since less heat is applied to the low CTE part, it expands less and more slowly, resulting in a lower delta-T before the parts separate.  

Figure 9 – Strong relationships between energy, delta-T, and time to separate
477 The left chart of Figure 10 shows the effect the interference fit has upon the delta-T required for the parts  to separate. As the analytical model suggests, the looser fit requires less delta-T for the parts to separate. The right chart of Figure 10 shows the effect the interference fit has upon the energy required to separate. The high CTE part doesn’t need to expand as much to free the low CTE part, which takes less energy. Additionally, when the interface pressure between the two parts is lower, less conductive heat transfer results between the two parts, so less energy is wasted heating the low CTE part. We initially expected less scatter in these charts, but realized static friction effects cause scatter (torque vs. preload tests exhibit this scatter, for example), and that minute surface imperfections have an appreciable affect on results, since the interface pressures are quite high in this design. These two charts also show two other relationships. The first, which we already discussed, is between input power and the time, energy, and delta-T required for separation. The second, discussed in the following paragraph, is the relationship between the initial temperature and the time, energy and delta-T required for separation.  Figure 10 – Relationships between interference fit, delta-T and energy required to separate  The left chart of Figure 11 shows the relationship between the initial temperature and the energy required to separate. The right chart of Figure 11 shows the relationship between actuation delta-T and the initial temperature. Since the parts were manufactured at 20°C, they will have a tighter interference fit with a lower initial temperature and will have a looser interference fit as they have a higher initial temperature.    Figure 11 – Relationships between initial temperature, and energy to actuate   Testing of the flight parts has thus far confirmed the following:    A single CTERA has been shown to not degrade in performance over 10 cycles. We intend to  perform further life testing when given the opportunity.
479 A Nichrome Burn Wire Release Mechanism for CubeSats  

Adam Thurn*, Steve Huynh*, Steve Koss*, Paul Oppenheimer*, Sam Butcher*,   Jordan Schlater** and Peter Hagan+ 

Abstract   The nichrome burn wire release mechanism uses a nichrome burn wire which when activated heats up and cuts through a Vectran tie down cable allowing the deployable on the satellite to actuate. The release mechanism was designed from scratch with the goals to make it small, inexpensive, simple, reliable and easy to use by anyone including student-run University CubeSat projects. The release mechanism, shown in Figure 1, utilizes a two saddle design with compression springs to apply a spring stroke and force to the nichrome wire to thermally cut through the Vectran cable when heated. Through a test program and using a design of experiments (DOE) approach it was determined that the applied current to the nichrome wire and the diameter of the nichrome wire were the key parameters governing successful performance. To activate the nichrome wire a constant current of 1.60 ± 0.05 amps is applied to ensure a successful and reliable cut. The tight tolerance constant current source is necessary in order to reliably: 1.) thermally cut the cable and 2.) prevent overheating failure of the nichrome wire to allow the mechanism to be reusable for many actuations without replacing the nichrome filament. The tight tolerances on the current prevent failure of the nichrome wire from overheating under too much current in a vacuum while also providing adequate thermal margin to cut through the Vectran tie down cable in air which requires more current than in vacuum. The burn wire release mechanism has been tested in air at room temperature and in vacuum at temperatures as low as -50°C and as high as 70°C on two different Vectran cable thicknesses. The release mechanism has shown to have cut times ranging from 2.4 to 7.2 seconds under these operating conditions. The burn wire release mechanism has 400 firings in component and system level testing without a single failure. The mechanism has been qualified for flight on the TEPCE (Tether Electrodynamic Propulsion CubeSat Experiment) program to release two carpenter tape deployments and a stacer and tether deployment system. 

Figure 1. Assembled burn wire release mechanism                                                              * Naval Research Laboratory, Washington DC  ** NRL Co-op Student from the University of Cincinnati, Cincinnati, OH  + NRL Co-op Student from Northeastern University, Boston, MA 

Proceedings of the 41st Aerospace Mechanisms Symposium, Jet Propulsion Laboratory, May 16-18, 2012