ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 5.  Efficiencies of MWG, HDG and Maxon GP’s Series The gear can be operated up to that torque at full input speed. If that torque is exceeded (Fig. 3, red  dotted line), the output shaft will lose its torsional coupling and the gear will act as a safety clutch, unless the torque has decreased. No harm will result after a one-shot or repeated transgression of the maximal torque. Torsional operation is the most important task of gears. As seen in Fig. 3, the torsional stiffness is also  strongly related to mechanical backlash. Whereas most planetary gears (GP52C, GP62A) have nearly no stiffness in the region of zero torque transition, MWG and HD are nearly free of backlash and therefore can provide a significant amount of torsional stiffness. Although harmonic drive gears (HDG) are overall  much stiffer than MWG, one interesting phenomenon around the region of zero torque transition (Fig. 3,  right side) can be observed. The ‘Lost Motion’ in HDG leads to degradation of the torsional stiffness of these gears. In this region, the stiffness of HDG and MWG is approximately the same size. This can be  important with respect to resonant frequencies of coupled mechanisms or space applications. The mechanical backlash or hysteresis loss of planetary gears is considerably bigger than in HDG or  MWG. HDG shows a hysteresis loss less than 1 [arcmin], whereas MWG shows even less hysteresis due to its friction-free force transmission (Fig. 4). After introducing efficiency classes [11], it is easy to compare the efficiencies of various gears with  respect to their reduction ratio and operating conditions. MWG shows an energy conversion rate in  efficiency class 1 (Fig. 5) within different operating speeds. The planetary gear stages of Maxon GP-series define industrial standard quality (class 2 and 3). HDG can hit the efficiency class 1 but is restricted  to optimized operating conditions such as 40°C and low input speed. Lowering the operational temperatures to -10°C, the efficiencies of HDG can reach class 5. This results from the HDG lubrication needs. 180
Conclusion MWG is a gear designed to operate without contact between the force transmitting elements. It therefore  overcomes the problem of lubrication and related problems in a new manner. Temperature-dependentlubrication effects, problematic contact stresses on the meshing gearing parts, hardening procedures of  teethes and all related problems are no longer existent. The question of lifetime is transferred to the much  simpler question of proper dimensioning of bearings. MWG benefits from extended lifetime and from  extended operational temperature ranges with fairly constant mechanical properties.    Magnetic gears have less torsional stiffness than conventional, mechanically geared systems. Despite  this fact, the torsional stiffness of MWG is partly equal (HDG) or even higher (planetary gears) than inconventional gearing systems. The torque density of MWG is steadily increasing, also lightweight design applications benefit from that new technology. MWG offers a simple internal structure with only a few  components, and has the potential as high reduction, high precision gearing system.    MWG and its unique combination of features give way to new prospects in future space applications.  References 1. Space Tribology Handbook, 5th edition (v2), pp.41 ff. and pp.132 ff., ESR Technology Ltd., 2013 2. Puchhammer, G. (2013). Magnetic Gears – New Approaches in Aerospace Applications, ‘Workshop -  Electromagnetic Devices in Aerospace Applications’, ESA-W13, ESA, Noordwijk, The Netherlands 3. Perez-Diaz, J-L, Diez-Jimenez, E.; MAGDRIVE: ‘A Non-Contact Magnetic Superconducting Harmonic  Drive’, ESMATS 2013, ESA, Noordwijk, The Netherlands 4. Jørgensen, F. T., Andersen, T. O. and Rasmussen, P. O. (2008) The Cycloid Permanent Magnetic  Gear, IEEE Transactions on Industry Applications, Vol. 44, No. 6, November/December 2008, pp.  1659 - 1665 5. Hatch, G. P. (2010). Recent Developments In Permanent Magnet Gear Systems & Machines,  ‘Magnetics 2010 Conference’, Lake Buena Vista, Florida, USA 6. General Catalogue, www.harmonicdrive.de7. Maxonmotor Catalog, Program 2013/2014 181
Figure 1. 3U CubeSat with two Ro ll-Out De-Orbiting modules (RODEO™)  RODEO ™Overview RODEO™ is a “turn-key” self contained de-orbiting system with everything required to operate the system  included within each RODEO™ canister. This includes the composite roll-out boom fitted with robust thin  film drag sail material, and the operational electronics with a built-in timing system and batteries. The  mechanical system is extremely simple with a very low mechanical part count and complexity. The use of  a high specific stiffness carbon fiber reinforced composite boom minimizes system mass yet provides a very robust deployed structure. The electronic system is also simple, providing an intentional release circuit and a timed release circuit. The timed release will execute if end-of-life is detected by a lack of  power to the charging circuit for a preset amount of time. Figure 2. Stowed CubeSat RODEO™  Fundamentally the RODEO™ structure is easily scaled in length and width and has an extremely high  packaging efficiency (greater than 100:1) allowing for integration upon “under-utilized” portions of the  spacecraft, such as behind deployable solar panels. Multifunctional Composite Slit-Tube Boom The key technology for RODEO™ is the rolled composite boom. CTD has developed a wide variety of  deployable composite slit-tube booms that range from 12.7 mm (0.5 in) to 127 mm (5.0 in) diameter (see  Figure 3) and up to 23 meters (75 ft) in length. Common to all of these structures (including the RODEO boom) are robust composite slit-tube booms that can be flattened and rolled for storage (see Figure 4). This method of stowage and deployment is highly efficient, allowing large structures, in comparison to the  spacecraft, to be stored within very small volumes. The material and composite laminate properties of  rollable composite booms are highly tailorable and can be optimized to meet the specific requirements for a given application. Additionally, because the boom and drag sail are rolled together in one piece, the  mechanical system has a very low part count with minimal complexity. The RODEO™ boom is self deploying, meaning the stored strain energy of the packaged boom provides  the necessary deployment force. The design of such a boom is challenging since the composite must survive the mechanical strain of being stowed within a very small volume, while also having adequate  184
deployment torque and significant axial stiffness upon deployment. The multi-functionality of the boom  results in minimal parasitic mass, stowed footprint, and cost, making these booms attractive for a variety of applications, including within roll-out de-orbit devices. Figure 3. Wide range of deployable  slit-tube booms developed by CTD Figure 4. Deployable composite s lit-tube boom used for RODEO™  Concept of Operations Inside the RODEO™ housing, CTD has included a very simple electronic circuit board and battery. The  board only has two inputs and a ground wire. The first input is to initiate a commanded release by supplying spacecraft voltage. The second input supplies a very small trickle charge to the battery and  resets a timer circuit so that RODEO™ will remain stowed. However, if the spacecraft loses functionality and stops supplying the trickle charge, the timer circuit initiates and begins counting. After a predetermined period of time with no trickle charge supplied, the timer circuit executes an automatic  command to deploy the RODEO™ drag sail. Once the command to deploy is sent (either manually or via  the timer circuit) an internal hot wire will release the spring-loaded hinged door, and the RODEO™ deorbit wing will deploy. Deployment occurs via the single-degree of freedom composite roll-out boom that is restrained by the hinged door of the RODEO™ canister.  RODEO Sizing To demonstrate the benefits of RODEO™, CTD ran an orbital decay analysis using NASA’s DAS 2.0  software in order to identify the minimum surface area required to de-orbit each size of satellite. This RODEO™ sizing analysis has been summarized in Table 1 for a 25-year reentry timeframe, and is based  on a 700-km circular orbit. Table 2 provides estimates for volume and mass for the three RODEO™systems assessed.   RocketSat-8 X Flight Experiment RocketSat-8 was the eighth RocketSat student project at the Colorado Space Grant Consortium  (COSGC) since the program began in 2006. RocketSat-8 flew as part of the national program called  RockSat-X. This program allows university-level students to fly experiments on board a sub-orbital rocket  launched out Wallops Flight Facility each August. The flight reaches an altitude of approximately 160 km  and provides experiments with full access to the space environment including 30 to 180 seconds of stable microgravity. Experiments also have access to power and telemetry. There is a fee to fly on the flight,  which helps foster partnerships with industry, universities, and students. CTD partnered with COSGC to  develop the RODEO™ experiment, enabling the students to cover their required fee to fly. Students worked side by side with engineers from CTD in the development of RODEO™. The experience in developing RODEO™ with CTD enriched the experience for each of the students involved. 185
Table 1. Required RODEO sizes for 25-year reentry Table 2. Estimated volume and mass for three RODEO™ configurations Flight Configuration For the flight design, CTD selected a RODEO™ that would provide approximately 0.15 m2deployed  surface area, enough area to de-orbit a 3U CubeSat in a period of 25 years (see Table 1). This area is achieved using three pop-out “wings” from a central 127-cm diameter boom as pictured in Figure 5. The  wings are formed 120° apart to provide area in multiple planes. Structural “battens” bonded to the backbone of the boom and extending out to the edge of the Mylar are used to support the wings and  enforce the deployed position. These battens are flexible and allow the three wings to be wrapped around to one side of the boom and stacked for stowage, as can be seen in Figure 6. Figure 5. 3U CubeSat RODEO™ configuration 186
Figure 6. Stowage of 3U CubeSat RODEO™ model Figure 7. Deployment progressi on for 3U CubeSat RODEO The stowed RODEO™ is held constrained within an aluminum box with a hinged door (see Figure 6). When the door is released, RODEO™ deploys in less than 1 second (see Figure 7). The flight hardware  was fabricated using space-qualified materials. Multiple stowage/deployment cycles were completed in  CTD’s laboratory. The RODEO™ was found to deploy reliably without any deployment snags, or stalls.  Flight Testing On August 13, 2012 RockSat-8 was lofted to an altitude of approximately 160 km above the Atlantic  Ocean before landing via parachute about 145 km from the Wallops Flight Facility. The payload was  recovered and high definition video of the RODEO™ experiment was retrieved from the payload. The  video showed a complete, but off-nominal deployment. The deployment was slower than expected but  continued in short bursts until reaching full deployment. Images from the RODEO™ flight experiment are  pictured in Figure 8.  Figure 8. Images from RODEO™ flight experiment 187
This anomalous deployment was shown to be caused by moisture saturation in the composite boom. In  the dry environment of Colorado, the boom always deployed nominally. When the students from theCOSGC took their 2012 RockSat-8 payload, including RODEO™, to Wallops, located on the very humid  coastal region of Virginia, they found that the RODEO™ boom did not deploy properly in final system tests and checks prior to integrating the payload into the rocket. CTD determined that the high humidity  caused the composite to absorb moisture, plasticizing the resin and significantly reducing the stiffness and stored strain energy of the boom. The effect was duplicated in laboratory conditions and iscompletely reversible. The boom would deploy as expected after drying out. The moisture issue was discovered two weeks before payload integration. To mitigate the issue, desiccant was packed within the RODEO™ canister to keep it dry for several days during launch  preparation. Unfortunately, the launch was delayed by several weeks and the desiccant became  saturated, allowing RODEO™ to also become saturated with moisture. As a result, the boom did not  immediately deploy. However, the multiple starts and stops observed indicated that as the outer layers of the boom dried out, the boom was gradually able to reach full deployment. Moisture would not be an issue for a CubeSat or small satellite mission since the boom would have ample  time (months to years) to completely dry out in the hard vacuum of space prior to deployment. In addition, CTD has since incorporated a different moisture-resistant polymer matrix into the boom and has shown  that deployment is no longer affected by moisture. Conclusions The most responsible and proactive method of mitigating orbital debris caused by small and very small  spacecraft is to incorporate an effective de-orbit device. To this end, CTD has developed RODEO™; a family of ultra-lightweight de-orbiting devices. All RODEO™ devices are capable of de-orbiting both  functional and non-responsive satellites within NASA’s required timeframe. RODEO™ modules utilize  extremely lightweight, mass efficient composite slit-tube boom elements to deploy the drag film in acontrolled, reliable fashion. CTD recently worked with the Colorado Space Grant Consortium to demonstrate a RODEO™ deployment on a suborbital platform as part of the RocketSat-8 experiment. While high-definition video taken during the experiment showed that RODEO™ did eventually deploy, the deployment was offnominal. Moisture contamination was found to be the cause, which was only an issue for the sub-orbital  demonstration due to the very short duration of the flight. At this point, RODEO™ is ready for future  CubeSat missions. References 1. NASA Orbital Debris Program Office, http://orbitaldebris.jsc.nasa.gov 2. Iannotta, Becky and Makil, Tariq, “U.S. Satellite Dest royed in Space Collision,” space.com, February  11, 2009 3.NASA technical standard NASA-STD-8719.14, o btained from  http://www.hq.nasa.gov/office/codeq/doctree/871914.pdf 188
Deployment Methods for an Origam i-Inspired Rigid-Foldable Array Shannon A. Zirbel*, Brian P. Trease**, Spencer P. Magleby*and Larry L. Howell* Introduction The purpose of this work is to evaluate several deployment methods for an origami-inspired solar array at  two size scales: 25-meter array and CubeSat array. The array enables rigid panel deployment and  introduces new concepts for actuating CubeSat deployables. The design for the array was inspired by the origami flasher model (Lang, 1997; Shafer, 2001). Figure 1 shows the array prototyped from Garolite and Kapton film at the CubeSat scale. Prior work demonstrated  that rigid panels like solar cells could successfully be folded into the final stowed configuration withoutrequiring the panels to flex (Zirbel, Lang, Thomson, & al., 2013). The design of the array is novel and  enables efficient use of space. The array can be wrapped around the central bus of the spacecraft in the case of the large array, or can accommodate storage of a small instrument payload in the case of the CubeSat array. The radial symmetry of this array around the spacecraft is ideally suited for spacecraft that need to spin. This work focuses on several actuation methods for a one-time deployment of the array. The array is  launched in its stowed configuration and it will be deployed when it is in space. Concepts for both passive and active actuation were considered.  Figure 1.  (left) Folded form of the four-sided CubeSat array. (right) Deployed configuration Actuation Concepts Several methods of actuation have been explored, including a motor-driven perimeter truss, pneumatic  actuation, centripetal acceleration, stored strain energy, and thermal activation (with a shape memory plastic). Because of the size of the panels in the 25-meter array, the perimeter truss is desirable to  support the deployment motion. For the CubeSat array, a less bulky actuation method is preferable.  Pneumatic actuation, centripetal acceleration, stored strain energy, and thermal activation of a shape memory plastic were demonstrated at that scale. *Department of Mechanical Engineering, Brigham Young University, Provo, UT **Jet Propulsion Laboratory, Pasadena, CA /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g23/g19/g87/g75/g3/g36/g72/g85/g82/g86/g83/g68/g70/g72/g3/g48/g72/g70/g75/g68/g81/g76/g86/g80/g86/g3/g54/g92/g80/g83/g82/g86/g76/g88 /g80/g15/g3/g49/g36/g54/g36/g3/g42/g82/g71/g71/g68/g85/g71/g3/g54/g83/g68/g70/g72/g3/g41/g79/g76/g74/g75/g87/g3/g38/g72/g81/g87/g72/g85/g15/g3/g48/g68/g92/g3/g20/g23/g16/g20/g25/g15/g3/g21/g19/g20/g23 189
Motor-driven perimeter truss A scale model of an array and truss, shown in Figure 1, were prototyped to demonstrate the functionality  and interaction of the two. The truss was SLA-printed, the connecting flexures were 3D-printed in Nylon, and the array was prototyped from Garolite and Kapton. Astromesh, which has a motor-driven actuation, is flight-proven and would likely be used in the final design, should the array be selected for a flight  project. Since our objective is primarily to show the interface between the array and the truss, we have sought to imitate the deployment of the AstroMesh without motorizing the model. Figure 2.  Partially opened m odel of the array and truss Torsion springs, shown in Figure 3, were attached to the truss to bias the truss open; i.e., the springs are  deflected when the array and truss are stowed. This method of stored strain energy makes the array self-deploy; however, we’ve guided the deployment to mimic motorization. The goal was to demonstrate a fixed rate of extension for each of the six sectors of the array. There is some clearance in the joints that prevent the truss from being a single-degree-of-freedom system, and therefore must be actuated at several points to deploy synchronously. When the center of the array is fixed, the truss rotates around the fixed point about 1.5 times to deploy the array.   Figure 3.  Torsion springs were affixed to the truss to bias the truss open We considered two different attachment points on the array: directly on the panels or at the membrane  between panels. To avoid ripping the Kapton film, we opted to attach directly to the panels. The challenge with attaching to the panels is that they undergo a complex rotation from the deployed to stowed position.  190
The joint needs to undergo a 90° torsion as well as bend 90° to one side. We chose a serpentine flexure,  shown in Figure 4, to accomplish this rotation.  Figure 2.  Serpentine flexures enabled the compl ex motion needed to connect between the truss  and the array The primary drawback of this joint is that it doesn’t constrain any degrees of freedom. However, it does  allow the complex rotation and allows for some extension of the joint as the distance between the array and the truss isn’t necessarily constant, especially considering the rotation of the panels. Having a compliant joint that can accommodate that change in distance as well as complete the two-axis rotations  is beneficial. In fact, it enables the two to interface. The flexures are secured to the truss with screws. We opted to glue the flexures to the panels, although  they could also be pinned or bolted to the panels. There is some slight interference with the ends of the flexures and the final folded form of the array (i.e., the outermost panels cannot sandwich perfectly flat against each other). We tested a cable-driven actuation method on a portion of the bays from the perimeter truss, shown in  Figure 5. The cable-driven prototype has a good mechanical advantage when the input cable isperpendicular to the truss members, but the low transmission angle that occurs when the bays are fully  stowed results in binding if the actuation cable is pulled straight down. Figure 3.  Cable-driven actuation on a few bays from the square truss 191
Pneumatic actuation For pneumatic actuation, a plastic bladder was adhered to the back of the array and inflated with  compressed air. If the bladder were designed to be air tight, it would remain inflated after the array opened and provide a semi-stiff support for the array. Centripetal acceleration Centripetal acceleration is the easiest to implement, as it can be accomplished by spinning the satellite.  We demonstrated a concept for centripetal acceleration using a central torsion spring.  A torsion-spring deployer, shown in Figure 6, was prototyped to actuate the array. The array fits snugly  over the small hexagon to hold it in place as the spring is displaced. We have demonstrated a rapid deployment of the small arrays with the centripetal acceleration from the torsion spring, as it is released. Figure 4.  (left) Prototype of the torsion spring te st base with array. (right) Translucent rendering  of CAD model, showing the cavity for the torsion spring Stored strain energy Strain energy can be stored in the array in a variety of ways. One approach evaluated at the CubeSat  scale is to affix tape springs radially along each sector. The prototype array opened quickly, but did not  lay perfectly flat; therefore, a mechanism to lock the panels in their deployed configuration may be desirable.  Figure 5.  (left) Stowed configuration of the ar ray with radial tape springs. (right) Deployed  configuration 192