ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
420 sheet is attached to the sub-chassis at one of the an tenna attachments, which avoids the addition of a  new nut and bolt and so keeps weight down.   Operation and explanation of the retention method

The retention system is based on the enclosing sheet’s ability to generate pressure on the folded antenna. Given that the sheet shares an attachment with one antenna, it will come out over this and then  go a complete turn, around the profile until all the ant ennas are covered. A traction force is applied to the  sheet tangent to the line of the profile, which turns into pressure on the antenna and keeps them pressed in against the sub-chassis profile. Once the sheet has enough force for it alone to keep the antenna folded, it can be tied by the retention thread, which will keep it closed until triggered. This thread is joined  to the enveloping sheet using a double knot, one on the internal side of the sheet and one on its external side, to prevent it coming loose after deployment has taken place.   Materials selection criteria

The enveloping sheet needed a material that was flexible but also resistant to traction and that, furthermore, had good dissipation properties for electr ostatic energy to avoid build ups that could  discharge onto any of the antenna and have unwanted effects. A sheet of polyvinyl acetate was chosen. 

Deployment mechanism and burner circuit  Deployment

Antenna deployment was produced by releasing the pr essure they were under by burning the nylon  thread keeping the enclosing sheet up against them. On ce released, the recovery of the elastic force  retained in them would place them in the correct position for sending and receiving data. If the antennas are not rigid enough they will take more time to stabilize the oscillation produced during deployment. It could even be the case that insufficient rigidity would leave them in an incomplete curved position due to the rotation of the orbiting satellite of roughly 1 rpm (0.00265 rad/s), which would greatly decrease their  properties. To solve this problem, a steel was sought that enabled work in the cold without becoming too  hard and that could absorb large amounts of energy wi thout suffering plastic deformation, that is, with a  large elastic limit, which led to a steel alloy with chrome and nickel being chosen. 

Figure 5. High speed recording of antenna deployment   NASA/CP-2010-216272
421 However, tests carried out on a 5-mm-wide flat metal band, showed that this configuration was not totally  valid:  it responded well to folding but was too limp during deployment and, once deployed, took too long  to stabilize and lost position when even slightly acted upon.   We found the solution to this problem by giving the metal strip a curvature, in such a way that it was slightly arched in a plane perpendicular to the di rection of the antenna. This modification reduced  deployment stabilization time by 55% and notably improved resistance to losing its position through external activity.  However, it was observed that when tightening the bo lt on the convex part against a flat plane (the subchassis), the antenna was deformed and its angle out from the shielding edge was no longer 45º but approximately 60º. To solve this problem it was decided to change the flat surface of the sub-chassis for  one with the same curvature as the concave part of the antenna and use a washer with a flat side (in contact with the bolt) and a curved one (in contact with the antenna) in order to avoid deformation and so maintain the outward angle of 45º. 

Figure 6. Antenna attachment area Figure 7. Antenna attachment elements  

Burner circuit

The burner circuit is the set of elements that brea k the thread keeping the antenna folded at the moment  of deployment. It is made up of a ceramic resistance  and a power circuit to the battery and includes a  switch to check deployment, which is triggered by the storage thread. Both the electrical resistance and the deployment switch are housed in a receptacle specially located to get them nearer to the area where the antennas are tied with the retention thread. This  minimizes thread length and so prevents possible  knots or unwanted tangles.  A 9-Ω, 0.125-W ceramic bodied resistance is used. One idea was to use Nichrome wire but this had the  problem of heating its whole volume, which could damage the polymeric sub-chassis, whereas the ceramic electrical resistance focuses the heat source much more, which, moreover, favors retention thread burning. 

Figure 8, Figure 9 and Figure 10. Receptacle where the Burner circuit is housed.  NASA/CP-2010-216272
423   Figure 11.  First mechanism detail 

In the double deployment (Figure 12) case, two frames  (for mass reduction requirements) are deployed  successively. The first mechanism is the same desc ribed above, and an extra mechanism is added with a  harder envelope limitation. The mechanism that unfolds the extra panel consists of two aluminum hinges,  each one glued to one side frame. All the parts are assembled by a shaft with two steel rotation springs,  and blocked in both sides.  Both deployable systems are fixed in stowed configurat ion by a nylon yarn. This nylon wire attaches the  assembly formed by the movable panels, and the lateral shear panel of structure. So each PDM with its corresponding shear panel can be integrated in the Xatcobeo structure as an individual assembly. To  start the movement of the mechanism there are two py roelectric elements that cut the nylon yarn setting  free the panel when commanded. Using two pyroelectr ic elements, a redundant configuration could be  achieved. When the yarn is cut, all the rotation spri ngs are free to start the deployment of the system. 

Figure 12.  PDM1/PDM2 design 

Two main requirements are issued for PDM: it shall deploy only once and shall send the result of the  deployment action. To detect the deployment in the single system, power telemetries are used, and in the double a contact switch is designed.  Flat spring blocking system

One of the new advances in this design is the introducti on of a flat spring to block the first deployment. It  prevents micro vibrations induced movements in a s pace environment and serves at the same time as  the helical springs as an additional blocking system . The blocking induced by this system consists in  NASA/CP-2010-216272
424 fitting the tip of the flat spring into a 1-mm depth slot  machined in rotating shaft head. The tip of the flat  spring produces a friction force on t he head of the shaft, acting as a damping for the rotating movement.  The lateral force induced in the shaft is derived from a MSC/Nastran non linear static analysis of the flat  spring using different thicknesses and materials. Th e values obtained are listed in Table 1. From the  analysis, a vertical movement above the equilibrium position is observed when the tip of the spring is out of the slot, so this implies to make a thicker slot in  shaft head. Other option to be taken into account is to  manufacture the shaft head as a half cylinder instead of a slotted cylinder. 

Figure 13.  Deformation of flat spring 

Table 1.  Lateral force for different spring thickness and materials 

Spring material Th ickness (mm) Force (N)  Steel 0.3 1.813  Steel 0.4 7.29  Steel 0.5 13.57  Al 0.4 1.319  Al 0.5 3.525   Lubrication and materials selection

Being a one–shot mechanism avoids lubrication, with the only matter of use different materials in contact  areas, and also a test campaign to minimize single  failures. The damping of the rotation movement is  achieved by: friction between metallic parts in the lug (A l) and shaft (steel) assembly; lateral contacts of  plastic washers with principal hinge and lugs; and by fr iction between tip of flat spring and head of rotating  shaft. The damping of the movement that requires fr iction goes against the requirement of reducing heat  transfer amount between deployed parts and main body of the satellite. So the material of the friction  washer will be one that allows friction but with low th ermal conductivity. And in the lug – shaft area, the  design tolerances will be adjusted to atta in low contact area. 

Simulation  A set of simulations have been run to dimension the sp rings and select the values for the springs that will  be used in flight model. The kinematic and dynami c simulations have been developed using MSC/Adams  program.  PDM 1 simulation

In the first simulation, the model is mounted with differ ent sets of helical torsion springs and different flat  springs looking for the lower impact force in the mechanical limits. Prior to the simulation, the model needs to be calibrated to define the characteristics of t he movable parts and the contacts presented in it.   NASA/CP-2010-216272
425 The mobile part consists of four solar cells, a FR4 board in which panels are fixed, the shaft, the hinge,  the bolts that fix it to the board and the wiring of the solar arrays. The mass amount considered in the  mobile parts is 40 grams.  There are two contacts in the model (Figure 14):   • The first contact between the hinge and the two blocks  in the lateral shear panel of the satellite is of  impact kind. It’s defined by:   o Stiffness of the contact. The two blocking el ements are the same so only one contact has  been considered in the calculations to reduce com putation time. So the stiffness of that part  has been increased to simulate the stiffness of the other blocking element. The stiffness is  the one obtained in a beam with flexion load. The value considered is 21185 N / mm  o Penetration depth: 0.01 mm  • The second contact between the head of the shaft and the flat spring. The contact has first a friction  phase and when the plate enters in the slotted area of the head of the shaft it has an impact phase.  The characteristics are:  o Impact phase:   Stiffness of the contact. The stiffness is the one obtained in the tip of the flat spring  with flexion load. The value considered is 414 N / mm   Penetration depth: 0.5 mm  o Friction phase. It’s between the tip of flat sp ring (steel or aluminum) and the steel head of the  shaft. It’s defined as a Coulomb friction force.   Static friction coefficient: 0.6   Dynamic friction coefficient: 0.4 

Figure 14.  PDM1 / PDM2 Simulation model     The movement is obtained by a revolute joint between the mobile parts and the grounded parts (lateral shear panel of the satellite), located in the s haft axis, and free to rotate around shaft axis.   The parameters that will be used to look for the lo wer impact force with the analyses are the stiffness  coefficient of the torsion spring (one or two acting in parallel) and the lateral force induced in the shaft by  the flat spring. The stiffness coeffici ents of the torsion spring calculated fo r different wire thickness, spring  diameter, number of loops, and number of spring is are listed in Table 2. 

   NASA/CP-2010-216272
426 Table 2.  Stiffness coefficient for helical torsion spring 

Wire diameter  (mm) Spring diameter  (mm) Number  of turns Number of  springs Stiffness coefficient  (N mm / rad)  0.4 2 5 2 16  0.4 2 5 1 8  0.4 2 10 2 8  0.5 3 10 1 6.51  0.4 2.5 5 1 6.4  0.4 2.5 10 2 6.4  0.4 3 5 1 5.333  0.4 3 10 2 5.333  0.4 2 10 1 4  0.4 2.5 10 1 3.2  0.4 3 10 1 2.667 

In the first set of simulations a constant lateral fo rce of 7.29 N is used with different spring stiffness  values: 16, 8, 6.4, 5.33, 2.667 N-mm / rad. The re sults obtained are shown in Figure 15 and Table 2. 

Table 2.  PDM1 simulation results obtained with spring stiffness variation 

Spring stiffness Force of first  impact (N) First impact  time (s) Number of  measurable impacts End time (s)  16 850 0.12 4 0.6  8 400 0.15 3 0.67  6.4 350 0.17 3 0.69  5.33 300 0.19 3 0.75  2.667 200 0.26 3 0.9 

Figure 15.  Impact force for different spring stiffness 

NASA/CP-2010-216272
427 The end time is taken to be when constant impact force is obtained. Observe that the movement is slower  when reducing spring stiffness, and also the impact force decreases.   In the second set of simulations, the stiffness coe fficient has been maintained in 6.4 N-mm / rad and the  lateral force has changed with the values provided by t he use of different spring material and thicknesses.  The results obtained are presented in Table 3 and Figure 16. 

Figure 16.  Impact force in for different lateral forces 

The lateral force has small influence in impact force. When lateral force is small, impact force can increase, and impact time of last rebounds also increa ses; but when it’s bigger, the friction phase is more  important and reduces impact force and impact time, and the flat spring enters in head slot before and  helps to finish the movement. 

Table 3.  PDM1 simulation results obtained with lateral force variation 

Lateral force (N) First impact force (N) Third impact force (N) Third impact time (s)  13.57 340 44 0.61  7.29 380 60 0.66  3.525 365 53 0.64  1.813 350 46 0.625   PDM 2 simulation

In the double deployment mechanism, a similar simula tion has been held. In this case only the springs  have been considered in the optimization of the problem.  There are two groups of mobile parts involved in  the movement: one group consists of the first frame with its shaft, the principal  hinge, and its secondary  hinges, with a mass of 8 grams; and the second group is formed by the second frame, its hinges and the  secondary shaft, with a mass of 6 grams.  There are three contacts in the mechanism (Figure 14):  • The first contact between the hinge and the two blocks  in the lateral shear panel of the satellite is of  impact kind. It’s the same contact defined for PDM1.  • The second is between the hinges of the two mobile frames. It’s also of impact kind and defined by:  o Stiffness of the contact. The value considered is 1000 N / mm  o Penetration depth: 0.1 mm  NASA/CP-2010-216272
428 • Third contact between the curved ends of second frame and fixed board over lateral shear panel. It  has two phases, one of impact and one of friction. The characteristics are:  o Impact phase:   Stiffness of the contact. The value considered is 10000 N / mm   Penetration depth: 0.1 mm  o Friction phase. It’s between the tip of alum inum frame and the FR4 board. It’s defined as a  Coulomb friction force.   Static friction coefficient: 0.3   Dynamic friction coefficient: 0.1  Two revolute joints have been included in each rota tion shaft to allow movement between mobile parts. 

The parameters used during optimization process are t he stiffness of principal (k1) and secondary (k2)  springs. The results obtained are shown in Table 4. 

Table 4.  PDM2 simulation results obtained with k1 and k2 spring stiffness variation 

k1 k2 Contact 1 Force (N) Cont act 2 Force (N) End time (s)  16 8 180 670 0.4  8 4 110 340 0.45  6.4 3.2 110 --- 0.5  3.2 2.667 102 220 0.55  2.667 2.667 62 --- 0.75   As happens with PDM1, when stiffness is reduced the movement is slower and contact force decreases. 

Conclusions  Antenna deployment

When this work was begun, the Xatcobeo project management was counting on the possibility of introducing a commercial deployment system weighing more than 100 g. Over time, this cost in weight became unacceptable. So work was started on in-house  development with the final aim of dramatically  reducing the mass needed for the antenna deployment mechanism. It was decided to use polymeric materials and titanium instead of metals such as aluminum or steel, and the deployment and retention  systems were unified around a single support component, which in turn would be the electrical insulation for the antenna. All these decisions had the same aim:  reduce weight without compromising the reliability  and stability of the mechanism or the assembled whole.  By selecting polyamide 6 as the structural material for the sub-chassis and covering it with zinc- or silverbased dissipative paint, the requirements for the behavio r of materials in the working environment were  fulfilled.  The choice of attachment elements made of titanium in  M2 metric instead of steel in M3 reduced the final  weight to just 20% of the initial weight. Furthermore, by making the sub-chassis from polyamide and not aluminum, its mass was reduced by 58.67%.   The final result of development is that the assembly made up of the sub-chassis, attachment elements (bolts, nuts and washers), electrical resistance, switch, heat-stable enclosing sheet and nylon retention  thread had a total weight of 12 g.  The reliability shown in the tests;  the great lightness of  the assembly achieved using low density  materials; the large internal space leaving room for so lar panels, cameras or other installed devices in the  mechanism’s installation face; its robustness, which makes it easy to handle and simple to assemble; and  its low cost and short manufacture ti me, thanks to the Selective Laser Sintering rapid prototyping process;  NASA/CP-2010-216272
429 all this, and all that has been explained above means that we have, in conclusion, achieved an ideal  mechanism for the storage and deployment of antenna not  only for picosatellites such as the CubeSat 1U  developed in the Xatcobeo project, but al so for other types of picosatellites.   PDM

Small satellites require small mechanisms. But from  miniaturization appear new problems such as  manufacturing difficulties; in PDM it was to build sh afts with the required length, small diameter and the  adequate tolerances, which will lead to make shorter broken shafts instead the of full length used now.   A deployment system with a mass of 45 grams us ing boards of a 1.6-mm thic kness was completed, and  also it’s possible to reach a double deployment with 55 grams using 0.8-mm boards (that in our case is  not applicable due to mass restrictions). It’s achiev ed the desired goals of having a system in a 6.5-mm  envelope protruding from the sate llite and allowing easy integratio n in such tight dimensions.    A flat spring retention system has been designed to av oid in flight induced micro vibrations and act as an  additional blocking system. The flat spring used should be the one that produces the greater lateral force to induce damping in the movement, but not too thi ck to provoke plastic deformation in the bended area.  The selection of springs for PDM1 and PDM2 leads to look for springs with small stiffness: more wire turns, less wire diameter, higher spring diameter and to use only one spring.   References   1. N.K. Pramanik, R.S.Haldar, Y.K.Bhardwaj, S.Sabharwal, U.K.Niyogi, R.K.Khandal, "Radiation  processing of Nylon 6 by e-beam for improved properties and performance." Radiation Physics and  Chemistry  78 (2009) 199–205  2.     Clovis Pinto, Leonardo G. Andrade e Silva,  "Study of ionizing radiation on the properties of  polyamide 6 with fiberglass reinforcement”, Radiation Physics and Chemistry  76 (2007) 1708–1710  3.   Waldir Pedro Ferro, Leonardo Gondim de Andrade  e Silva. "Ionizing radiation effect studies on  polyamide 6.6 properties" Radiation Physics and Chemistry  71 (2004) 267–269  4.  A. Toorian, K. Diaz, S.Lee "The CubeSat Approach to Space Access." IEEE Aerospace Conference  Proceedings,  art. no. 4526293   5.   M.W.R. Alger, K.D. "A design and development of Pico- and Femto-satellites." International Journal  of Manufacturing Research  3 (1), pp. 48-79   6.   Swartwout, M., Kitts, C., Twiggs, R., Kenny, T., Ray Smith, B., Lu, R., Stattenfield, K., Pranajaya, F.   “Mission results for Sapphire, a student-built satellite” Acta Astronautica  62 (8-9), pp. 521-538  7.  Peter Fortescue, John Stark, Graham Swinerd. “Spacecraft Systems Engineering”.  8.  Joseph E. Shigley, Charles R. Mischke. “Diseño en ingeniería mecánica”      NASA/CP-2010-216272
432 many more waiting for launch opportunities [6] cove ring many applications, from breakthrough science  measurements to technology proving missions.   CubeSats have very important limitat ions, affecting their breath of use. Most importantly, they have limited  power, and must maintain a low mass to conform to the strict CubeSat specifications, which enable their simple integration and thus launch. Several technology developers are now addr essing these restrictions  through deployables, but their success is very limited after launch [7], [8], and no deployable CubeSat solar array of this type exists.  In this paper, we describe the de sign, construction, testing, and future considerations for XSAS. We begin  with design requirements, carry a design through modeling and prototype testing, continue with a  discussion of the design, modeling and testing results of this innovative design, and end the paper with a description of ongoing work, which includes testing in microgravity through NASA’s Reduced Gravity Student Opportunities Program.   Design Overview   The XSAS concept is a complete and modular powe r system composed of tw o major subsystems: a  multi-panel, double-sided, deployable solar array and a power management and storage bus (power bus). While stowed, the solar array and power bus fit with in a CubeSat’s 1U (10x10x 10 cm) and 0.5U (10x10x5  cm) volume respectively. Once released, the array passi vely deploys in an “accordion style” to increase  its surface area. The power bus r egulates and stores the energy from the solar array and interfaces  directly with the payload through a single universal connection.  This makes XSAS a modular plug-andplay “space battery” applicable for small CubeSats. The current design supports 16 stowed solar panels  and accommodates payloads of up to 1.5U (10x10x 15 cm) completing the maximum 3U (10x10x30 cm)  CubeSat size. Figure 1 displays a comparison of a standard CubeSat geometry with an early 14 panel stowed and deployed XSAS in tegrated 3U CubeSat. 

Figure 1. Dimensions of a Standard 3U CubeSat and 3U CubeSat with XSAS 

Previous 3U CubeSat were design and flown with a 5.0-kg upper mass limit. However, in August 2009  CalPoly requirements reduced this constraint to 4 kg [5]. As a result, the XSAS design presented in the  paper is based on the latter requirement. Table 1 lists the mass budget for each subsystem.      3U CubeSat Standard Geometry - Length: 34cm - Array Area: 197 cm2  3U CubeSat with Stowed XSAS - Length: 34cm  3U CubeSat with Deployed 14 Panel XSAS - Length: 171cm - Array Area: 1051 cm2  NASA/CP-2010-216272