ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
An advantage of the proposed solution is a fully controlled, precise shape of the manufactured tubular  boom, homogeneity of the material deformation, no folding either bending of beryllium tape during production process, economical construction and easy control. A disadvantage is dependence of the manufactured tubular boom length on the device size (the longer the tape, the longer the device) and low production rate. At the moment the maximum length of the manufactured tape is 1.5 m. A separate issue worth paying attention to is the hardening of the final product – a process necessary for  obtaining the required stiffness and elastic properties of the final tubular boom. The proposed method (Figure 10) is the one of hardening with a thick-wall aluminium tube (1), inside of which there is a  segment of the tubular boom (2). The aluminium tube is reeled with a heating element (3) that heats it evenly. The tubular boom is loaded over a mandrel rod (4) and heated to the hardening temperature of 315°C (according to the heat treatment procedure provided by the BERYLCO 25 distributor). Numerous optimization tests were conducted using 25-cm prototypes of the bender and hardening heater  that allowed some parameters to be determined, i.e. the tape pulling strength, the tape travel time, the  tape leading system, the pressure force, the forming rod diameter, the mandrel rod diameter, etc. Based  on the results of a foregoing analysis, the target device, including two modules – the bender and hardening heater – was developed and the first series of 1.5-m long tubular booms was successfully  manufactured (Figure 11).  In space applications the tubular boom is to be stored on a reel. When deployed it is to form a cylindrical  tube. In order to confirm such a possible use of the manufactured tubular booms and to characterize them as well, the final product underwent a series of relevant validation tests carried out by GUTRONIC in  close cooperation with CBK PAN. The tests-proved the mechanical and resilient properties of the product met all requirements.13  2 4 Figure 10.  Hardening heater – pictoral view: 1 – aluminium tube, 2 – tubular  boom, 3 – heating element, 4 – mandrel rodFigure 9. Local bending – process scheme:  1 – tape, 2 – susceptible tie, 3 – prismatic base, 4 – forming rod,  5 – working area 518
Figure 11. First series of five hardened 1.5 m long tubular booms MAB functional tests MAB underwent a series of acceptance and qualification tests, all conducted at CBK PAN. Table 1 shows  the sequence of tests performed on the MAB flight model and their results.  Table 1. MAB Test Campaign No. Test Result Remarks 1. Functional test at room temperature (~20°C) passed 2.T-shock test (25 cycles of rapidly changing  temperature in the range of +60°C ÷ -25°C )passed 3. Functional test at room temperature (~20°C) passed 4. Vibration test passed 5. Functional test in Thermal-Vacuum Chamber (-6°C) failed 6. Functional test in Climate Chamber (-20°C) – Investigation of failure cause 7. Functional test at room temperature (~20°C) passed 8. Vibration test passed 9. Functional test in Climate Chamber (-10°C) Passed 10. Functional test in Thermal-Vacuum Chamber (-6°C) Passed Approved for flight As indicated in Table 1, the first functional test in the Thermal-Vacuum Chamber at -6°C ended up in a failure. A number of failure scenarios were considered in order to explain this result, however two causes  were most likely:  /g120 thermal shrinkage causing mechanical jam of the rotating components,  /g120failure of the Dyneema string melting – the force with which the string leaned against the resistor  could be too low, causing poor adhesion and resulting in ineffective temperature dissipation.  519
As a part of a thorough investigation test 6 was carried out. It showed that thermal shrinkage was the  cause of the detent deadlock and consequently the mechanism malfunction. At a low temperature (below  0°C) the detent axle could not rotate despite the proper electromagnet drive action. Remedial measures were taken: MAB was disassembled and the hole in the bearing bush was enlarged. The normal running fit became the easy running one. During MAB disassembly, an extra tension spring was added to  increase the lock & release mechanism reliability (showed in Figure 8). The main spring task was to generate an additional tension force for the Dyneema fiber to better wrap it over the heater.  Since the problem was successfully detected and eliminated, the set of tests was repeated. The device  passed all subsequent tests, achieving TRL 6. Figure 12 shows MAB during functional and vibration tests. Figure 12. MAB during: (top) functional test at room temperature, (bottom) vibration test. 520
Conclusions & Lessons Learned The prototype of the miniature antenna called MAB, driven by the modified escapement mechanism, has  been successfully developed together with the innovative technology of tubular boom manufacturing. The mechanism has been qualified through the dedicated test campaign, thereby proving correctness of the  novel deployment mechanism idea – the idea assuming a dual complementary action of the driving  system which simultaneously performs a role of a drive unit and a kind of moderator.  As with any mechanisms development, some important lessons were learned: /g120Dealing with the new type of tubular boom showed how delicate and troublesome the tape is. It  was very easy to damage it at every stage of assembly and testing. Once the project team had to cope with a tape crack (Figure 13) as a result of numerous winding and unwinding actions. It was  initialized by rugged edges caused by careless hole drilling. Working on MAB, the team got to  know the mechanical properties of the tubular boom. Figure 13. Tubular boom fatigue (marked in red) /g120 Thermal difficulties – in the case of such a miniature mechanism thermal phenomena is of a key  importance. During the first series of tests MAB failed the functional test at low temperature due  to the detent deadlock. It turned out that the normal running fit was too restrictive.  /g120 Assembly and manufacturing difficulties – the mechanism assembly and miniature components  manufacturing turned out to be very challenging.  A pinpoint accuracy was required. Furthermore, GUTRONIC company, for many years closely cooperating with CBK PAN ,gained a  complete and unique knowledge of manufacturing tubular booms of BERYLCO 25. It got to know  advantages and disadvantages of the developed method as well as the comprehensive characteristics of the final product. Now GUTRONIC specializes in manufacturing tubular booms for space applications and is able to improve its method, at the same time enriching the range of products with various tubular booms lengths, diameters and materials. Possible application The foreseen applications of the MAB-type deployment mechanism equipped with the miniature tubular  boom or another spring drive are as follows:  /g120compact electrical antennas, /g120ultra-light planetary manipulators (the torsional stiffness of structure can be strengthened by  combining booms into a pair), /g120dual action units used wherever the dynamic nature of spring drive introduces dangerous  vibrations and induces severe local stress in the structure that need to be mitigated and  521
smoothed (i.e. in hinges with a latching system commonly used to deploy booms with small sensors or measuring instruments, in solar panels systems, etc.),  /g120active dampers/actuators. As it was shown in this paper, MAB is a properly functioning and reliable instrument. Its miniature character predetermines the device to be applied particularly on board nano- and micro-satellites. Future development The already studied solution seems to be a perfect starting point for future space projects. The subject  aroused ESA's interest. Under Polish Industry Incentive Scheme, the agreement between ESA and  ASTRONIKA (spin-off company founded by a CBK PAN group of engineers), GUTRONIC and CBK PAN  was signed. Within the framework of the contract, a new generation of the electrical antenna will arise. The device will be significantly improved in comparison to MAB – it will be a fully functional, 2.5-m long  antenna designed on the basis of the new type Ø10-mm tubular booms. The existing one (Ø6 mm) does  not cover applications on small and medium satellites, representing a large number of present-day projects. Therefore, taking into account a more common use of tubular booms, production of Ø10-mmprofiles of thicker beryllium bronze strips seems to be favorable. References 1.  Matthews S. E.: "The Seven Most Incredible Telescopes In Existence. The Biggest, Smallest and Most  Robotic Stargazing Tools.” /g51/g82/g83/g88/g79/g68/g85/g3/g54/g70/g76/g72/g81/g70/g72 , 12 (2013). 2. Grygorczuk, J., Banaszkiewicz, M., Seweryn, K., Spohn, T.: “MUPUS Insertion Device for the Rosetta  Mission”. /g45/g82/g88/g85/g81/g68/g79/g3/g82/g73/g3/g55/g72/g79/g72/g70/g82/g80/g80/g88/g81/g76/g70/g68/g87/g76/g82/g81/g86/g3/g68/g81/g71/g3/g44/g81/g73/g82/g85/g80/g68/g87/g76/g82/g81/g3/g55/g72/g70/g75/g81/g82/g79/g82/g74/g92 , 1 (2007), 50-53. 3. Grygorczuk, J. et al.: “Prototype of Ultra-Light Planetary Manipulator – Design, Test and Simulations”.  /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g20/g21 /g87/g75/g3/g54/g92/g80/g83/g82/g86/g76/g88/g80/g3/g82/g81/g3/g36/g71/g89/g68/g81/g70/g72/g71/g3/g54/g83/g68/g70/g72/g3/g55/g72/g70/g75/g81/g82/g79/g82/g74/g76/g72/g86/g3/g76/g81/g3/g53/g82/g69/g82/g87/g76/g70/g86/g3/g68/g81/g71/g3/g36/g88/g87/g82/g80/g68/g87/g76/g82/g81/g3 /g11/g36/g54/g55/g53/g36/g3/g21/g19/g20/g22/g12/g15/g3 Noordwijk, the Netherlands.  4. Grygorczuk, J. et al.: “Advanced Mechanisms and Tribological Tests of the Hammering Sampling  Device CHOMIK”. /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g20/g23/g87/g75/g3/g40/g88/g85/g82/g83/g72/g68/g81/g3/g54/g83/g68/g70/g72/g3/g48/g72/g70/g75/g68/g81/g76/g86/g80/g86/g3/g68/g81/g71/g3/g55/g85/g76/g69/g82/g79/g82/g74/g92/g3/g54/g92/g80/g83/g82/g86/g76/g88/g80/g3 /g11/g40/g54/g48/g36/g55/g54/g3/g21/g19/g20/g20/g12/g15/g3 Constance, Germany.  5. Collective work: “Tubular spacecraft booms (extendible, reel stored)”. /g49/g36/g54/g36/g3/g86/g83/g68/g70/g72/g3/g89/g72/g75/g76/g70/g79/g72/g3/g71/g72/g86/g76/g74/g81/g3 /g70/g85/g76/g87/g72/g85/g76/g68/g3/g11/g74/g88/g76/g71/g68/g81/g70/g72/g3/g68/g81/g71/g3/g70/g82/g81/g87/g85/g82/g79/g12 , February 1971 522
the spacecraft’s orbit. For a simple spinner, the system’s angular momentum is simply its moment of  inertia about the axis of rotation times its spin rate. Assuming a spin rate of 0.8 Hz and a 500-km orbit, the  torque required to slew in this manner exceeds the capabilities of our magnetic torque rods by roughly  two orders of magnitude.1 Figure 1. MicroMAS Orbit and Swath Overlap2 Since MicroMAS does not possess enough torque to precess the entire angular momentum bias of a simple spinner, it must leave its angular momentum oriented in a constant direction, which is illustrated in Figure 2. Because of this arrangement, a simple spinner version of MicroMAS must exchange the  momentum of the spacecraft between internal storage devices (such as reaction wheels) to maintain  proper orientation of the radiometer payload as the spacecraft traverses its orbit. At point (c), the  momentum storage device would have to account for twice the angular momentum produced by the rotating system. For a simple spinner rotating at 0.8 Hz, this angular momentum would be an order of magnitude greater than the momentum storage capability provided to MicroMAS by the reaction wheels in  the MAI-400. 3Additionally, at points (b) and (d) in the orbit, the momentum storage devices must trade  between themselves the entirety of the spacecraft’s momentum four times during the course of a single rotation of the spacecraft, which would require a torque nearly five orders of magnitude greater than the maximum torque available to the MAI-400’s reaction wheels. 3No miniature reaction wheel set could  realistically transfer momentum fast enough between its wheels to keep a simple spinner aligned with the LVLH frame. 1 Figure 2. MicroMAS Maintaining LVLH Alignment with Constant Momentum1 524
For these reasons, MicroMAS was designed as a zero-momentum, dual-spinning spacecraft, in which a  purpose-designed “scanner assembly” mechanism would provide the actuating mechanical and electrical interface between the spinning payload module and the spacecraft bus fixed in the LVLH frame.Figure 3 provides a system overview of MicroMAS in its dual-spinning configuration. Figure 3. MicroMAS System Overview It was determined early in the design phase that to ensure mission success, the scanning mechanism  would need to possess a specific set of capabilities (presented in Table 1). This list of capabilities was then formalized into specific design requirements, against which the performance of the finalized  mechanism could be validated.   Table 1. Scanner Assembly Design Requirements Requirement Justification/Rationale Must fit within a 10 cm by 10 cm by 3.5 cm  volumeThis ensures that the mechanism fits within  the CubeSat cross-section and that all  components will fit within the length  available in a 2U bus structure Must be able to continuously operate in a  space environment for a nominal mission lifetime of one year.This is the only system onboard that  requires continuous mechanical (rather than electrical) operation The scanning assembly shall exert a maximum torque less than or equal to 7 mN-mMaximum torque available at the reaction  wheels at 0 rpm is 7.4 mN-m Must be capable of rotating a 1U, 1-kg payload at a nominal rate of 1 Hz (60 rpm)                        Requirement for data collection at target  altitude Must have angular position knowledge to a  precision higher than 0.1 deg (6 arcmin)Minimize error for geolocation accuracy Must have at least 8 electrical feed-lines Minimum number of lines required for  power/data transfer between bus and payload 525
Scanner Assembly Design The design of the MicroMAS scanner assembly (shown in Figure 4) incorporates both COTS and customdesigned hardware in order to satisfy the aforementioned design requirements. Commercial parts include a brushless DC motor (rotor/stator only); an incremental rotary encoder; a thin-section radial ball bearing;  and a miniature slip ring capsule. Custom parts were then designed to incorporate the COTS components  into a single assembly and provide the necessary mechanical interfaces between the bus and payload structures. The entire assembly fits within the 10 cm x 10 cm x 3.5 cm (3.937 in x 3.937 in x 1.378 in) volume constraints imposed by the design requirements. Figure 4. Cross-section view of Scanner Assembly COTS Component Selection The need for low-speed vacuum operation led to the selection of a zero-cogging brushless DC motor as  the basis of operation for the scanner assembly. A brushless motor was selected over conventionalbrushed motors because of the electrical arcing that can occur between the brushes during vacuum operation. Furthermore, the low iron content of the zero-cogging rotor allows for smooth rotation at the relatively low speed of 60 rpm. The selected component was an Aeroflex * Z-0250-050-3-104 rotor and  stator combination, due to its continuous torque capabilities, and large internal bore. Its continuous torque capability of 254 mN-m (36 in-oz) was well above the 7 mN-m operating requirement, leaving the system  with a large performance margin for future applications. The dimensions, while not the most compact of all the options surveyed, fit within the design space, though the 2.7-cm (1.060-in) height of the stator  necessitated that several components be mounted internal to the assembly.  To provide angular position knowledge, the MicroE **M1500V optical rotary encoder system was chosen,  for its vacuum operation capability and its available array of optical gratings of various diameters and  resolutions.5The selected grating has an outer diameter of 5.04 cm (2.0 in), which allows it to be placed  *Aeroflex Incorporated, Plainview, NY **MicroE Systems, GSI Group, Bedford, MA 526
Figure 5. 3-D Printed Scanner Assembly Mockup Engineering Design Unit The next unit fabricated was an Engineering Design Unit (EDU). The purpose of this assembly was to  verify mass properties of the custom components, assess whether any part posed manufacturing issues, and provide an initial functional unit for early testing and driver development. Due to component lead  time, mass mockups for the rotor and stator were used, and functionality was provided through the use of  a surrogate motor mounted externally to the assembly, as shown in Figure 6. At this point in the design  process, a mass mockup of the satellite – with the Engineering Design Unit scanner assembly joining bus and payload mockups – was subjected to random vibration testing at “qualification” level, as prescribed  by GSFC-STD-7000. 9Post-test inspection of the hardware showed that the glass grating disk did not  shatter or separate in any way from the hub to which it was mounted, quelling two concerns that the team had about using a glass grating. Additionally, the shaft still spun freely after the test, showing that the  bearing could survive the combined loading imposed by random vibrations and from stowage in a  CubeSat deployment device. *   Figure 6.  Engineering Design Unit with surrogate drive motor *MicroMAS will be deployed by a NanoRacks CubeSat Deployer (NRCSD), which can be seen at  <http://nanoracks.com/nanoracks-completes-flight-integration-cubesats-bound-orb1-iss/>. 528
Engineering Test Unit After incorporating these design changes, an Engineering Test Unit (ETU) was built utilizing flight-spare  COTS components (shown in Figure 7). This unit was used for full functional testing and validation, and also served as a baseline estimate for cost and schedule of the flight-model scanner assembly.  Figure 7. Engineering Test Unit scanner assembly Structural Validation Considering the lack of flight heritage of a dual-spinning CubeSat, and the criticality of the scanner  assembly to the MicroMAS mission, a broad series of tests were conducted to qualify the mechanism for successful performance in relevant environments. Structural tests and analyses were conducted to identify the most effective method for securing the rotor to the motor shaft, due to concern for loss of  contact during thermal expansion/contraction. Three mating methods were evaluated – epoxy, tolerance  rings, and mechanical press-fit. Though tolerance rings seemed most directly applicable for the task, a  lack of knowledge regarding their space heritage (and a lack of experience using them) resulted in the choice for epoxy securing the rotor and shaft in the flight-model assembly. Rotor/Shaft Interface Test Units Occurring in parallel with fabrication of the engineering test unit scanner assembly, a series of tests were  conducted to determine which method would be used to secure the rotor to the shaft in the flight unit. This trade was motivated by the concern that differential thermal expansion between the stainless steel rotor and the aluminum shaft would cause a loss of contact between the two parts at the low temperatures. Calculations were performed to determine the change in rotor and shaft diameters at the expected (cold) operational temperature, along with three temperatures below that operational point. The three classifications of interference fits are presented in Table 2.  Table 2. Diameter ranges for interference fits 10 Fit Classification SymbolMinimum Size  Above Basic Shaft  Diameter (mm) [in]Maximum Size  Above Basic Shaft  Diameter (mm) [in] Locational  Interference FitH7/p6 0.026 [0.0010] 0.042 [0.0016]  Medium Drive Fit H7/s6 0.043 [0.0017] 0.059 [0.0023]  Force Fit H7/u6 0.060 [0.0024] 0.076 [0.0030]  For the calculations that follow, a worst-case scenario was assumed where the rotor inner diameter was  the largest value allowed by the H7 fit class. Table 3 summarizes the material interference between the  shaft and the rotor diameters for various temperatures. 529
Table 3. Change in interference fit between sh aft/rotor under various te mperature differentials Temperature (°C) /g507/g55 (°C)Interference (mm)  [in]Fit  Classification 25 0 0.051 [0.0020] Medium Drive 0 -25 0.038 [0.0015]Locational  Interference -10 -35 0.033 [0.0013]Locational  Interference -15 -40 0.030 [0.0012]Locational  Interference -25 -50 0.025 [0.00099] Locational  Transition The results of the above calculations showed that although the press fit mating method would be  acceptable for the expected operating temperature range, it did not leave much margin for survival if the system got colder than expected. This led to a study of different mating methods and a series of tests to characterize the performance of the options under consideration.  /g3 The two additional mating methods considered were epoxy and tolerance rings. The selected epoxy was  Lord 3135, due to extensive institutional knowledge of the product and known flight heritage on the Chandra X-ray Observatory and the Suzaku satellite. 11A tolerance ring, shown in Figure 8, is a  corrugated metal strip specifically designed for mating concentric, cylindrical parts. Sample units were provided by USA Tolerance Rings. * Figure 8. Tolerance ring12 The tolerance ring is installed between the shaft and the bore, providing a radial spring force that  maintains alignment of the two parts. The rings are available in a wide variety of stock styles and sizes, making them ideal for nearly any combination of loading and geometry. However, the lack of internal experience with this product, coupled with no known flight heritage added two critical risks to this option. Table 4 summarizes the advantages and disadvantages of three methods that were considered. The press fit method was used in the engineering test unit, and was thus tested through the functional  tests the ETU scanner assembly was subjected to. The other two options, epoxy and the tolerance ring,  were tested separately. The purpose of this series of tests was to assess the strength of the rotor-shaft  interface provided by both methods. This was accomplished by creating two sets of surrogate shafts and rings, with dimensions identical to the flight components. The shaft was made from 6061-T6 aluminum alloy – the same material that the flight shaft would be made from – with a 9.9-mm wide by 0.127-mm deep (0.390 in by 0.005 in) circumferential groove cut into it, as depicted in Figure 9. The initial dimensions for the groove were established by combining required dimensions of both the tolerance ring  and the epoxy – the tolerance ring size set the groove width while the recommended epoxy bond line  *USA Tolerance Rings, Inc., Pennington, NJ 530