ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
specially profiled bracket (B) to which a hinge spring hook was attached and a plastic wedge (C) attached  to the structure side wall. The wedge element is also used as a soft stop component. To prevent violent door bouncing and subsequent satellite damage, the latching system is used. The latch works with the  leaf spring (D) which, when pressed by the moving door deflects until the latch pin (E) slips into the hole in the leaf spring. Thus the door is caught, preventing the rebound and stays wide open at an angle of  110 deg. A microswitch is used to verify and confirm door opening. Hold Down & Release Mechanism (HDRM) To hold the door during take-off, a very reliable HDRM was used, which is released by the Dyneema cord  melting system. This type of lock and release mechanism has been tested during the ESA Rosetta mission preparation of the MUPUS instrument. In DRAGON, the mechanism, shown in Figure 6, has  been scaled to meet higher loads.   Figure 6.  DRAGON Hold Down & Rel ease Mechanism principle of operation shown in stowed (left) and released (right) position When stowed, the doors are held by latches mounted to the tensioned elastic flat spring and adjusted by  tightening an M10 screw that changes the position of an adjusting element with respect to a structural wall. Latches are held in place by the levers constructed of carbon fiber tubes (to decrease its mass and  inertia). At the ends of the levers, flat springs are provided with attachments for the Dyneema string. The  use of these flat springs guarantees the mechanism not to loosen during Dyneema creep. Therefore, it is  possible to use a string with a small diameter (0.5 mm) that is easy to melt with little power. Melting of the  Dyneema string releases the levers rotated by torsion springs, which allow the latches to slip from the door brackets and thus letting the doors open. The clamp ratio between shear pins and Dyneema string for this mechanism is 200:1.  Test and Qualification The Dragon deployer passed a full series of tests, beginning with functional and ending with  environmental. Functional tests have been carried out on the existing test bench located at SRC, to  actively simulate the lack of gravity and were recorded by high-speed camera. A system, shown in  Figure 7, uses a spring that generates constant torque and is able to compensate the impact of gravity  acting on the suspended object. The system relies on feedback from a force sensor that is turned on by  the PID controller that sends a signal to a DC motor, which adjusts the tension of the spring tape by  491
appropriate rotation. During the test, a spacecraft mass model mock-up was used. The mock-up suspended on a rope was ejected from the deployer located below the test stand. Motion of the satellite  during ejection is quite a fast phenomenon. Therefore, setting parameters of the PID controller play an  important role in simulating zero gravity. Three tests were conducted with different settings of control parameters. In an ideal case, the satellite  ejection speed is increased in accordance with the characteristics of the force generated by the conical  spring. The satellite separates from a pushing platform after the pushing action stops. From that moment,  the spacecraft should move at the same final speed. Tests results shown in Figure 8 are presented in the form of plots of time vs. speed and speed vs. displacement. They were obtained from the analysis of  high-speed camera images.  Figure 7.  Functional test stand to act ively simulate the lack of gravity Triangular marks on the plots indicate the time when a pushing platform reached its end position and thus  the point at which the satellite begins free flight. Square marks indicate both time and location when the  satellite passes remaining protruding elements of the deployer and is considered to be outside of it.  492
Figure 8. Satellite ejection speed graphs  Results from test 1 and 3 best fit the theoretical movement conditions of the spacecraft in zero gravity. Test 2 is characterized by a sudden drop in speed of the satellite after separation which indicates that the  mass was not supported adequately. From the time of the doors’ release, the pushing platform supporting the satellite moves 150 mm in about 0.3 sec to reach an end speed slightly less than 1 m/s. The speed limit was set at 1.5 m/s and the expected value, which did not take into account motion resistance, was  1.2 m/s. The test demonstrated that assumptions have been met, and a video recorded with a high-speed  camera showed that the ejection takes place very smoothly and doors open much faster than the  spacecraft is ejected. Figure 9. Deployer qualification tests:  a) thermal-vacuum; b) vibration The next functional test was performed in the thermal-vacuum chamber shown in Figure 9a. The time of  Dyneema string melting and mass model ejection was less than 1 sec at -20°C and a 0.0004-Pa vacuum  level. There was no destruction or change of resistance of the heating element. DRAGON passed all  required vibration and shock tests. No change was observed in resonant frequencies after sine and  random vibration. The resonant frequencies in the X and Y axes (the base plane axes) were 205 Hz and  278 Hz in Z axis (along which spacecraft ejection occurs). Light ringing could be seen coming from the deployer main spring. 493
Finally, a fit check was performed on a launch vehicle. DRAGON was attached to the bracket on the  adapter ring between the rocket and the main satellite. A deployment signal was sent from the launch  vehicle and the deployer opened safely. Telemetry with sensor confirmation of deployment was sent to  the rocket computer. Lessons Learned There was not enough time to design a new suitable and reliable separation device or order it from an outside source. In the deployer, a scaled hold down & release mechanism from MUPUS instrument was used. This approach has been proven as reliable. A general lesson connected with this issue is to  carefully analyze a scaled device that will be used in a new application. We dedicated a lot of attention to  redesign with this new mechanism so it could withstand higher loads. What we did not notice originally was that the MUPUS mechanism levers were arranged in a position towards the operator, while in  DRAGON that position was just opposite. The use of components to mount the Dyneema string in a way similar to that used in MUPUS made arming of the mechanism very difficult. So two new redesigned  components had to be manufactured. To cut costs and to speed up the deployer production, all rotating joints used commercially available  PEEK flanged sleeve bearings and polished stainless steel (SS) shafts. This combination was also used  for latches made from stainless steel pressed against PEEK lever blocks and sliding across PEEK door  brackets. Tests showed that such a combination is sufficient for applications working in not too low nor too high temperatures. Two utility issues have sprung up during the deployer testing with the Heweliusz spacecraft. Both concern  satellite safety. DRAGON is not a tightly closed container but has a structure with many open spaces that could expose a spacecraft to danger. Also, deployer doors have open cutouts to save mass. This causes  the possibility of damage in case of an unintended fall of some piece on spacecraft solar cells during  locating pins set up. To avoid it, suitable ground-use covers were designed to protect the satellite. In  Figure 10, the satellite integration procedure is shown on the left and the deployer with a safety cover on  the right.  DRAGON, different from other separation ejectors, is equipped with a pushing platform. Satellite structure  rails are not supported. This results in a possibility of hitting satellite parts, for example solar cells, during  satellite insertion into the deployer. To prevent this, four wheel assemblies were added in every corner with a 1-mm gap between the spacecraft and every wheel. Figure 11 shows wheel assemblies that also  help to provide a very stable low friction guiding during satellite deployment. 494
Conclusions A nanosatellite orbital separation device was designed, manufactured, assembled and successfully tested. The overall dimensions of the deployer are 282x320x287 mm and its mass is 7.2 kg. It is able to  deploy a small cuboid BRITE-PL Heweliusz spacecraft of dimensions 200x200x200 mm and mass of 7 kg. Commercially available standard orbital deployers are usually developed with particular spacecraft in  mind. DRAGON separation device was developed with a target to satisfy preexisting specific constraints. The mass ratio of DRAGON deployer to satellite is about 1. In the author’s opinion, such a system should  have a mass ratio less than 0.8, preferably 0.5, and that is possible to achieve with a closed-type  deployer and structure optimization. Figure 12 show graphs where masses of aforementioned popular deployers and Dragon are plotted. In the deployer mass vs. payload mass graph (left ) markers that  represent standard American-made container-type ejectors are joined by a line (CalPoly P-POD 3U and  NASA WWF 6U). If we take this line as a determinant of representative orbital deployers, it turns out that  the ejector designed for 7 kg payload should have a mass of about 5 kg. Figure 12.  Deployers mass vs. payload mass that they are able to eject General conclusion:  it is very hard to design, manufacture, integrate and test a space device in two  months. It is even more challenging during the holiday period which was the case of DRAGON development. But all this was possible with a well-motivated team that was using solutions already proven and previously tested during development of the space instruments in SRC PAS. References 1. Dobrowolski M., Grygorczuk J. “Lock and release mechanism for the CHOMIK penetrator device and its tribological properties”  /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g23/g20/g86/g87/g3/g36/g48/g54 ; NASA/CP-2012-217653 pp. 29-37 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), pp50-53. 3. Toorian A., Blundell E., Puig Suari J., Twiggs R. “Cubesats as responsive satellites” /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3 /g22/g85/g71/g3/g53/g72/g86/g83/g82/g81/g86/g76/g89/g72/g3/g54/g83/g68/g70/g72/g3/g38/g82/g81/g73/g72/g85/g72/g81/g70/g72/g3/g21/g19/g19/g24/g3 AIAA-RS3 2005-3001 4. Thompson L. D. "Development of a NASA 6-U Satellite" /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g21/g24/g87/g75/g3/g36/g81/g81/g88/g68/g79/g3/g36/g44/g36/g36/g18/g56/g54/g56/g3 /g38/g82/g81/g73/g72/g85/g72/g81/g70/g72/g3/g82/g81/g3/g54/g80/g68/g79/g79/g3/g54/g68/g87/g72/g79/g79/g76/g87/g72/g86 SSC11-X-8 /g3 496
the SIMPLE boom [4], which demonstrated effective functionality. The design was adapted for this  application. Further details on the tape spring design are described in a later section below.   a) b) Figure 1.  Membrane optic prototype support st ructure in a) deployed and b) stowed  configurations.  Each tape spring is subjected to a compression load at full deployment to react the design tension in the membrane optic. To eliminate the need of a rotation lockout mechanism for the hub assembly at full deployment, the tape spring neutral axis must be coincident with the shaft axis (see Figure 3). This prevents an induced moment at the hub axis from the eccentric loading, which may result in hub rotation if the moment is greater than the tape spring bending stiffness. The same principle determines the  required location for the mounting of the adjacent tape spring end, also shown in Figure 3. The  intersection of the tape spring neutral axes is coincident with the membrane loading vector, which bisects  the angle formed by tape spring neutral axes and is normal to the edge of the membrane end tab.  a) b)   Figure 2. a) Exploded and b) Assembled renderings of top frame corner hub.  The tape spring is located within the central hub to minimize the overall strain in the transition region from  the clamped end to the final roll diameter. This mounting position, in addition to constraining the tape  spring roll axis to be concentric with the hub shaft axis, results in interference between the tape spring  and the hub rotation axis in the fully deployed state. The first option considered, illustrated in Figure 4,eliminates this interference with a U-shaped fixture fastened to the hub disk. This allows the tape spring  499
to flatten during rolling. The assembly center of gravity was designed to be coincident with the rotation  axis as to not induce dynamic imbalance loads during deployments. This design was not selected since the overall width of the hub fixture resulted in the system dimensions to exceed the 1U CubeSat volume  requirement. Figure 3.  Top frame tape spring neutral axes locations. Figure 4. Tape spring hub mechanism design concept requires no clearance trimming Figure 2b) shows the selected design of trimming the tape spring end to clear the hub shaft. The tape  spring clearance cut geometry has a significant effect on the deployment torque and despite only a small reduction in deployed buckling strength [7]. A tight clearance cut, shown in Figure 5a), will maximize these parameters. Fabricated prototypes with this trim profile were observed to have a high failure rate in the radius cut region during rolling of the tape spring. This configuration could be successfully rolled if the tape spring is initially flattened. Sufficient available torque and buckling margins allowed for a more gradual taper cut, shown in Figure 5b), which provided a robust solution capable of repeated stow-deploy  cycles without special handling requirements. If the fixture dimensions can be accommodated, it is  preferred to have a tape spring with constant cross section, i.e. no clearance trimming. Constraint/Release Mechanism In the stowed configuration, the rotation degree of freedom (dof) of each tape spring hub must be  constrained until deployment is initiated. The tape spring laminate was designed such that containment fixtures are not required to prevent the tape spring roll from radially billowing out in the stowed state. However, if the free end of the tape spring is held fixed, but the roll rotation dof is unconstrained or opened, the tape spring is in an unstable state and the tape spring roll will begin to rotate. The constraint and release mechanism components to control the deployment for all six tape spring rolls are shown in Figure 6. To prevent binding and have a synchronous deployment, this constraint must be released from  all six hubs simultaneously.  500
a) b) Figure 5.  Two example tape spring end clearance cuts.  Figure 6.  Constraint  mechanism components.  The boom tape spring hub fixtures are fastened to the base plate equally spaced 120º apart. Each boom  fixture is paired with a top frame tape spring hub assembly, which is positioned above the vertical  extension. A cylindrical rod controls the deployment of the pair by vertical (Z-axis) motion. The deployment control rods are mechanically coupled via a rigid ring. For this prototype unit, deployment is  initiated by manual actuation of the deployment synchronizer ring. Details for the interface of the deployment control rod and the boom tape spring roll is shown in Figure 7. The tape spring hub half shaft protrudes through the fixture wall. A notched, thin cylindrical disk with a slip fit-sized hole is placed on the  shaft. The slip disk also contains a radial tapped hole for a setscrew that mechanically couples/decouples  the slip disk from the hub shaft. During tape spring rolling, the setscrew is loosened such that the slip disk  can remain stationary. The deployment control rod contains a protruding pin. The rod is actuated vertically such that pin fits within the tape spring notch. Once the tape spring is rolled to the desired length, the setscrew is fastened down. The boom tape spring roll is now locked or constrained. Actuating  the control rod vertically down clears the pin from the disk notch, which releases the constraint allowing  the tape spring to deploy.    501
a) b) Figure 7.  Boom tape spring c onstraint design in the a) lo cked and b) released states. A similar methodology is implemented to control the deployment of the top frame tape spring. The  operation and design features of the top frame constraint mechanism are best explained by describingthe structure stow procedure. The structure starts in the fully deployed state shown in Figure 1a). The top frame tape springs are simultaneously rolled until the relative distances between the hubs approximate  the stowed state shown in Figure 1b). The top frame is temporarily constrained using tape or Velcro applied circumferentially around the assembly. The tape spring booms are then simultaneously rolled until  the extending hub shaft (Figure 8b) can be placed into the respective slip disks. The temporary top frame  constraint is removed to allow manipulation into the slip disks. The boom tape springs are further rolled such that the top frame hubs are preloaded against the vertical hard stop (Figure 8a). The fixture prevents  the top frame from deploying at this stage. Each top frame tape spring is then tightened around the hub disk and held, while setscrews in the constraint fixture (Figure 8a) are fastened to lock the rotation. All tape springs are now fully constrained. Actuating the deployment control rod clears the constraint fixture  from the tape spring hub shaft (Figure 8b). The height of the vertical hard stop is set such that the top  frame and boom constraints are released simultaneously.  Figure 8. Top frame hub cons traint mechanism design in the a)  locked and b) released states. Flexible Composite Components The flexible composites described for deployable structure applications are traditional stiff matrix  composites with traditional aerospace-grade resin systems (977-2 and 8552 epoxies, cyanate esters, etc.). The laminate total thickness is kept relatively small such that it can be subjected to large bending  strains without failure. The mechanics of flexible FRP composite have been analyzed and tested  502
reduce the degradation, but this results in a high-energy system that results in an unacceptable shock  event upon full deployment. Excessive degradation can lead to deployment failure. A simple means to  control deployment rate and avoid shock loads has not been developed. Strain energy deployable  structures often employ complex mechanisms that control deployment paths and rates using housings, motors and fluid based dampers. These devices can double the system mass, negating any improvements in structural performance and increasing the system and launch costs. A simple materialbased solution is currently being investigated, where a CFRP laminate with discrete /g72/g79/g68/g86/g87/g76/g70/g3 laminas  provides deployment force and /g89/g76/g86/g70/g82/g72/g79/g68/g86/g87/g76/g70/g3 laminas to passively control the deployment rate of the  structure. The challenge here is to tailor the net viscoelastic response for deployment rate control. The  proposed idea can be thought of as a spring (provided by the low creep elastic lamina) in parallel with a damper (provided by the viscoelastic lamina). Telescope Structure Tape Spring Design From previous experience, a composite laminate consisting of an S-2 glass unidirectional fiber sandwich  by an Astroquartz plain weave layers at 45º bias has shown good roll stability property with high resistance to creep. The unidirectional lamina was AGY 933-AA-750 S-2 Glass fiber pre-impregnated with  Patz Materials & Technolgies PMT-F7 with 10% by weight 3M nanosilicate. The JPS Style 525  Astroquartz II plain weave was also pre-impregnated with the PMT-F7 epoxy. Material properties are  listed in Table 1. The effective orthotropic properties for the laminate are presented in Table 2.  Table 1.  Tape spring material properties Property (units)Astroquartz II/PMT-F7  Plain WeaveS2 Glass/PMT-F7  Unidirectional E1(GPa) 21.7 56.9 E2(GPa) 21.7 16.8 G12(GPa) 3.4 6.1 /g8112 0.11 0.27 t (mm) 0.1016 0.0762 Table 2. Tape spring laminate properties Property (units)Astroquartz II/PMT-F7  Plain Weave E1(GPa) 21.7 E2(GPa) 21.7 G12(GPa) 3.4 /g8112 0.11 t (mm) 0.254 The stress in the membrane was a critical design parameter. The membrane design stress level of  68.9 kPa (10 psi) was chosen, which translates to the corner pull-point load of approximately 0.5 N. The  load-carrying members in the structure are illustrated in the side view of one corner location in Figure 11. The position of point Ain Figure 11 was considered fixed for the initial calculations, which sets the  vectors of the membrane force, Fm, and the positioning lanyard force, Fl. The vector of the spring force,  Fs, and decoupling spring stiffness, ks, were varied in a Matlab script to calculate the forces in the top  frame and boom tape spring members and their respective lengths. This study gave the position  tolerance of point Ato maintain Fmwithin ±0.05 N and the required tape spring cross-section radius to resist buckling with margin. A commercial off-the-shelf extension spring with stiffness of approximately  505