ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
297   Figure 3.  Spiral Orbit Tribometer used to evaluate  the relative lubricant life of various alloys under  simulated bearing conditions. 

The plates were initially clean and only the ball wa s lubricated. For the present tests in the boundary  regime, the ball is first weighed and then lubricated by dripping a dilute solution of the lubricant, ~ 1 mg of  lubricant per ml of hexane solvent, onto the ball rota ting on a small bench lathe. The ball is reweighed  after evaporation of the solvent and the lubricant char ge is obtained from the weight difference. About 20  mg (24 nl) of lubricant, a multiply alkylated cycl opentane (MAC) designated by the trade name Pennzane  2001A, were used on each ball. Tests were run here at a system load of 130 N (30 lb), which resulted in a track width of ~0.4 mm and corresponds to a Hertz pressure of 1.06 GPa. The test conditions result in operation in the boundary lubricati on regime. Considering the small quant ity of lubricant present in the  tribosystem, test lifetimes on the order of a few tens of thousands are typical for good tribological  materials like 440C and past experience with the SOT i ndicate its relevance in mi micking the conditions in  a space ball bearing application [6].   Results and Discussion 

Under these conditions considered representative of  precision bearings, performance (life and friction)  was comparable to that observed with 440C bearing balls. Figure 4 shows friction for tests of 60NiTi balls running against 440C plates (left graph) and Nitinol 60 pl ates (right graph) in the SOT. Typical lubricant  lifetimes were approximately 30,000-60 ,000 revolutions indicating a slow lubricant consumption rate. In  one test using both 60NiTi plates and ball, the rig was permitted to run well beyond the complete consumption of lubricant. In that case, friction climbe d to a coefficient of around 0.35 but was stable and  no galling tendencies were observed.  In comparison, SOT tests of balls sputter coated with  pure, thin titanium films degrade the lubricant,  experience high friction and seizure within a few dozen revolutions. Unlike traditional titanium alloys, the  60NiTi alloy performs well and is a good candidate for bearings and mechanical components and is impervious to the corrosion issues facing traditional bearing materials especially during storage before launching.   NASA/CP-2010-216272
298   Figure 4.  SOT results for Nitinol 60 ball running against 440C (left) and Ni tinol 60 plates (right)  showing comparable behavior.  

Summary 

This research effort has identified 60NiTi alloy as a promising candidate material for bearing and  mechanical component applications. 60NiTi, when appropriately processed and fabricated, is  dimensionally stable, hard, wear resistant, non-galli ng, and tribochemically benign in the presence of  liquid lubricants. This behavior is in stark contrast to conventio nal alloys that contain such large amounts  of the metal titanium. It is believed that the good tr ibological performance under oil lubrication observed  for 60NiTi may extend to the entire NiTi family of allo ys since they all share si milar phase constituents and  basic atomic level bonding. The tribochemistry of 60NiTi and its metallurgical relatives is under further  study.  The identification of a viable bearing material that is non-magnetic, electrically conductive, hardenable,  displays favorable tribochemistry and is non-corros ive is a major research finding. No other bearing  material yet discovered has such a broad combination of properties. While it is clear that near term niche applications such as aerospace bearings and gears ex ist, many non-obvious applications are also likely  to present themselves. These include wear resistant, corrosion proof knives and cutters, electric machine  structural and dynamic components, high performance fasteners, valve components and many others.  Clearly, much more research will be required to understand the 60NiTi material  and its metallurgical  relatives. The relationships between mechanical and physical properties, atomic structure and micro-scale ordering and surface chemical interactions remain to be investigated. Nonetheless, the NiTi  metallurgical system clearly has engineering potential well beyond shape memory alloy applications that  has only begun to be exploited.   References 

 1.  S.A. Shabalovskaya: “Surface, Corrosion and Biocompatibility Aspects of Nitinol as an Implant  Material”, Bio-Medical Materials and Engineering, volume 12, 2002, 69-109.  2.  Buehler, W.J., J.V. Gilfrich, and R.C. Wiley:  “Effect of Low-Temperature Phase Changes on the  Mechanical Properties of Alloys near Composition Ti Ni”, Journal of Applied Physics, volume 34,  number 5, May 1963, pages 1475-1477.   3. Bueher, W.J.: ”Intermetallic Compound Based Materi als for Structural Applic ations”, Proceedings of  the 7th Navy Science Symposium: Solution to NAVY Problems Through Advanced Technology, May  NASA/CP-2010-216272
299 14-16, 1963, U.S. Naval Aviation Medical Center, ensacola, Florida, voulme 1. Office of Naval  Research, Arlington, VA, May 1963.  4. W.J. Buehler: “NITINOL Re-Examination”, White Oak Laboratory Alumni Association Leaf, Winter  2006, volume VIII, Issue I, www.wolaa.org.  5. R.W. Roberts and R. S. Owens;  Titanium Lubrication, Nature, volume 200, October 26, 1963, pages  357-358.  6.  Jones, W.R.Jr., Pepper, S.V., Wheeler, D.R., Jansen, M.J., Nguyen,  Q.N., and Schroer, A., “The Effect  of Stress and TiC Coated Balls on the Lifetime of  a Perfluoropolyalklether Using a Vacuum Rolling  Contact Tribometer”, Tribology Transactions, volume 43(2000), 4, 685-688.  7. M.N. Gardos, H.H. Hong, and W.O. Winer: “The Effect of An ion Vacancies on the Tribological  Properties of Rutile (TiO2-x), Part II: Experimental  Evidence”, STLE Tribology Transactions, volume  33, number 2, April 1990, pp. 209-220.  8.  M.D. McNeese, D.C. Lagoudas, and T.C. Pollock ; “Processing of TiNi from Elemental Powders by  Hot Isostatic Pressing”, Materials Science and Engineering A280 (2000), pp. 334-348.  9. K. Otsuka and X. Ren: “Physical metallurgy of Ti-Ni-based shape memory alloys,” Prog. Mater. Sci.,  vol. 50, pp. 511-678, 2005.  10. W.R. Jones, Jr., S.V. Pepper, M. J. Jansen, Q.N.Nguyen, E.P. Kingsb ury, S. H. Loewenthal, and R. E.  Predmore: “A New Apparatus to Evaluate Lubric ants for Space Applications-The Spiral Orbit  Tribometer (SOT), NASA TM-2000-209935. 

   NASA/CP-2010-216272
301 Gossamer Technology to Deorbit LEO Non-Propulsion Fitted Satellite 

C. Dupuy* and O. Le Couls** 

Abstract   Since 2004, CNES has decided to apply the end of life Code of Conduct rules to debris mitigation. Originally drawn up by the main European space agencies, it contains basic rules to be applied in space in order to limit the increase of orbital debris. In low Earth orbit, the rule is to limit in-orbit lifetime to 25  years after the end of the operational mission, or else to transfer to a graveyar d orbit above 2000 km. In  order to follow these instructions, a task force was set up in 2005 to find the best way to implement them on MICROSCOPE and CNES microsatellite family (MYR IADE). This 200-kg spacecraft should be  launched in 2014 on a 790-km high circular orbit. Without  targeted action, its natural re-entry would occur  in 67 years.   Two strategies to reduce this time period were compared:  propulsive maneuvers at the end of the  mission or the deployment of large surfaces to increase  significantly the ballistic coefficient. At the end of  the trade off, it was recommended:  ¾ For the non-propulsive system fitted satellites , to use passive aerobraking by deployment of  added surface,  ¾ For satellites having propulsive subsystem in bas eline for mission purposes, to keep sufficient  propellent and implement specific maneuvers.   The poster gives an overview of the process that  led to the development of a deployable aerobraking  wing using a lightweight aluminized Kapton membrane and an inflatable aluminum laminate boom.  • The main requirements   • The trade off among various aerobraking solutions  • The development plan   This technology presents a very attractive potential and it could be a first step in using of inflatable  technology on spaces vehicles, before to deal with others more exigent applications.     MICROSCOPE, orbital configuration MICROSCOPE, deorbit configuration 

Figure 1                                                    * Centre National d’Etudes Spatiales, Toulouse France   ** EADS Astrium, Saint Médard en Jalles, France 

Proceedings of the 40th Aerospace Mechanisms Symposium, NASA Kennedy Space Center, May 12-14, 2010  NASA/CP-2010-216272
302 The MICROSCOPE Spacecraft  

MICROSCOPE is based on CNES’ MYRIADE micro-sate llite product line which started in 1998. Its  scientific objective is to test the Equivalence Princi ple with an accuracy of 10 -15, about three orders of  magnitude better than the accuracy of the present on ground experiments. Th e space mission exploits  the Earth as the gravitational source , the very quiet in orbit environment and the possibility of a very long  free fall motion for the integration of the measurement.  The T-SAGE instrument is composed of two differential electrostatic accelerometers  (see Figure 2). Each  one includes two cylindrical and concentric test masses . The attitude, as well as the atmospheric and  thermal drag of the satellite, are actively controlled in such a way that the satellite follows the two test masses in their pure gravitational motion. The test-mass motions, within the highly stable silica instrument frame, are also servo-controlled using very accura te capacitive position sensing and electrostatic  actuators. The relative position of the two masses is thus maintained motionless and the fine measurement of the control force l eads to the test of the Equivalenc e Principle with the expected 10-15  accuracy. The mission is obviously extremely sensitiv e to microperturbations such as structural microcracking and fluid motions. It was apparent very early in the design that no liquid propulsion system could be envisaged. The satellite drag compensation and attit ude control involves a spec ific propulsion system.  For the first time, FEEP thrusters have been chosen but due to technical problems a cold gas solution is  now being investigated. The satellite will be on a sun-synchr onous polar orbit at 790-km altitude with an  ascending node at 6 h or 18 h. This leads to a nominal orbit where natural re-entry would take place in 67  years (see Figure 3). At 200 kg, MICROSCOPE is heav ier than common MYRIADE satellites. Its general  architecture also presents several particularities:  ¾ the payload is in the middle of the spacecraft,  ¾ the body is wider (870 x 790 section instead of 600 x 600),  ¾ the solar array is based on two symmetrical sing le-panel wings instead of  the standard MYRIADE  two-panel wing solar array. 

Figure 2: T-SAGE instrument     De-Orbiting MICROSCOPE: Options and Constraints   The dominant factor of orbit erosion is the atmosp heric drag, which depends linearly on the surface to  mass ratio (S/m) of a given spacecraft. The deceleration due to drag can be written as:  γ = ½ ρ S/m C d V² , where:  - ρ is the local atmospheric density,  - Cd is the coefficient of drag, in pr actice between 2 and 3, assumed to be equal to 2.6 in the rest of  this study,  - V is the orbital velocity.   The surface S is the cross section perpendicular to the veloci ty vector (see figure 5). After the end of its  operational life, the attitude of the satellite will not be controlled. Since for mo st of the decay phase  (above 650 km) there is no obvious nat ural self-pointing, we consider that all altitudes would be equally  NASA/CP-2010-216272
303 likely. The average apparent surface  of MICROSCOPE is evaluated at 2.7 m2. With a mass of 200 kg,  this leads to :  S/m = 0.0135 m2/kg. 

With the DTM94 atmospheric model, using the Codior software, which is a semi-analytical tool for long  term extrapolations developed by SERGA for CNES for sun-synchronous or low orbits, we find that MICROSCOPE’s natural re-ent ry would occur 67 years after decommi ssioning (see Figure 3), which is  assumed to take place on January 1 2017.   Figure 3: MICROSCOPE’s orbital decay without  special maneuver Figure 4: MICROSCOPE’s orbital decay with 6m 2 drag area 

Since MICROSCOPE is below 800 km, the Code of Condu ct option to transfer it to a graveyard orbit  higher than 2000 km at the end of its life does not seem feasible within the mass and volume allowed for  the spacecraft. We will theref ore focus our efforts on ens uring that re-entry will oc cur in less than 25 years  after the mission’s end. Considering its relatively small mass, uncontrolled re-entry is acceptable. Two  main strategies were identified:   - “Drag raise”: increase of natural orbit erosion so that or bital decay from the original orbit will take  less than 25 years,  - “Propulsion”: maneuver to go down to an orbit where life-time is less than 25 years.    The only option that has been studied in detail for “drag raise” increase, is to deploy surfaces around the spacecraft to raise natural aerodynamic drag. Other concepts could have been involved (deployment of  tethers to use magnetic forces for braking), but they  were quickly dismissed as impractical for a small  satellite. Figure 4 shows that the r equired 25 years re-entry would occur if the S/m is set to 0.03 m²/kg,  which calls for a mean drag area of 6 m². This result corresponds to a beginning of de-orbiting at the start of 2017. This means an addition of 3.3 m² of drag surface to the operational configuration of MICROSCOPE. 

Figure 5: Drag area calculation  NASA/CP-2010-216272
304 Drag Raise: Review of Concepts 

To begin with, a target shape for the deployed surfaces was selected. The ideal system would create additional drag homoge neously over all possible attitudes of the spacecraft. In theory, one may try to  approach the shape of a sphere.   

Sphere Diameter ≥ 2.7 m  / Volume > 10 000 liters  / Deployed total area  > 23 m²  / Mass budget > 30 kg  /      Yoke necessary to link with satellite 

Dihedral  ☺ easier to accommodate and   complement the surface provided by  satellite & solar array panels  ☺ Mass < 15 kg   ☺ Dihedron could be done with serial of   flat panel arrangement or with boom  associated with deployed membrane  film.  

Cone ☺ Same advantages than dihedron   arrangement  . Difficult to provide a perfect cone shape   without  discontinuities  / Not necessary for uncontrolled re-entry 

 ÆThe dihedral shape was finally chosen !   To build the wings, two families of technologies were quickly identified. The first is based on low weight flat panels. Because of the restrictions on volume in packed confi guration, only 2 wings of 3 panels (0,5  m² each) can be accommodated. This only leads to a total mean drag area of 4 m², 2. m² short of the  6. m² target. 

 Based on sandwich panels with carbon fiber sheets, t he total mass of this package including holding and  deployment devices has been evaluated at 6 kg. Even though de-orbiting with this system would take  around 40 years without margin from the nominal orbit, we disregarded it for the more innovative second solution family.  The second family of technologies is based on unfolded Aluminum / Kapton membranes with deployable  structures. Figure 6 describes various conc epts that were assessed for deployment.  NASA/CP-2010-216272
305   Figure 6: Technological trade off 

• Tensegrity  structures, made from cables and struts to which a state of pre-stress is imposed that  imparts tension to all cables. This solution was assessed to be complicated for this type of  application, and not easily accessible.   • Deployable masts based on multi-coil Stacer tube, such as those used on Demeter. Here, the fact that the mast rotates thro ugh deployment is unfavorable.  • Doubly curved membrane stiffened  by thin-walled collapsible ribs such as the Collapsible Rib  Tensioned Surface (CRTS) reflector devel oped by ESA. Here the packaging and the  demonstration that micro-cracking would not  happen during the mission, seemed difficult.  • Bi-stable masts  such as those originally developed by DLR for solar sails, made of two laminated  flexible -shaped sheets bonded at the edges to form a tubular shape. The main disadvantage is in  their significant mass for a limited (compared to so lar sail) area such as in this application,  • Inflatable mast  made of aluminum laminate. This is the solution that was finally selected because  of its light weight and its efficient packaging especially in term of volume. It is presented in more detail. The total mass of two inflatable wi ngs offering 3.7 m² of drag area is 12 kg. 

Figure 7: wings with inflatable mast 

NASA/CP-2010-216272
306 De-Orbiting Wing Technology 

The de-orbiting configuration is shown in Figure  7. Two 5-m-long wings made of aluminized Kapton  membrane (100 g/m2 density) have been deployed by a central inflatable mast. The actual total mean  drag surface becomes 6.7 m².  The total provisional mass, with margin, of a two-wing package is 12 kg (5 kg per wing and 2 kg for the inflation system), including thermal control. The materials have been selected for their endurance to ultraviolet exposition, and to atomic  oxygen aggression. They can also cope with high temperatures which  are expected when the membrane is in full sun.   IDEAS General Architecture 

The IDEAS system is composed of several sub-systems (see Figure 8), each of  them ensuring one or  several specific functions.  

Figure 8:  Product tree  

The hold down sub-system function is to mainta in IDEAS in folded config uration and to ensure the  unlocking before deployment. The inflation sub-system  function is to ensure the boom deployment and  rigidization. It is also ensuring the venting of the boom in folded configuration during launcher phase, and  in deployed configuration after rigidization. The MLI (M ulti Layer Insulation) sub-system function is to  ensure the thermal protection of IDEAS system from t he external environment, in stowed configuration.  The wing’s function is to ensure the deployment of the aerobraking su rfaces for IDEAS operating life. The  two wings have the same definition and are compos ed of one boom and two membranes. The boom is  ensuring the deployment and the rigidization to maintain  the wing in deployed conf iguration (with inflation  sub-system), and the two membranes are ensuring the aerobraking function.    IDEA S Inflation  sub-system MLI  (Thermical  insulation ) Wing 1  Sail 1 Sail 2 (idem ) Boom  Deployment  control Inflatable  mastMembrane Upper arm Tensionning  system Wing  2 (idem 1)  Structure Hold down  mécanism   NASA/CP-2010-216272
307 The inflated boom of the selected technological solution has to be rigidized after deployment to ensure a  correct mechanical behavior of the deployed surface without maintaining pressure. Several solutions  have been analysed:  • solvent evaporation  • Sub-Tg  • thermal polymerisation  • photochemical polymerization  • metallic laminate yielding   The criteria considered are power need, long in-orbit storage consequences (premature rigidization risk) and feasibility complexity. The retained solution is the metallic laminate yielding.   Presentation of the Laminate Technology   The principle of folding, deployment and rigidizati on of the IDEAS boom is described in the following  illustration.  

Figure 9:  Boom folding, deployment and rigidization concept 

The rigidization process chosen, “metallic laminate yiel ding”, consists of applying to the boom material a  sufficient deformation - applied by internal pressure – to suppress geometrical defect created by yielding  of the material during folding. Once the defects are suppressed, the mechanical behavior of the boom is ensured by its own stiffness. The boom material used to ensure this function is polyimide/aluminum/polyimide laminate. 

Figure 10: Laminate definition 

The aluminum layer ensures the mechanical behavior a fter rigidization. The Kapton internal layer ensures  the protection of the aluminum layer from internal mechanical “aggressions” and the tightness during boom inflation. The kapton external layer ensures the protection of the aluminum layer from external mechanical “aggressions” and has a thermal protection function. The main advantages of this solution are the good material stability during stor age phases and the use of  the same source of  energy as for the  deployment (no electrical power is necessary to rigidize).  Laminate aluminum boom technology is useful to in orbit low-stress-loaded structures. For the IDEAS  wings at altitude up to 320 km, aerodynamic s pressure is estimated below 0.42 E -3 Pa. This is negligible  face to dynamics load due to uncontrolled effect. Spin speed could reach 10 °/s.  NASA/CP-2010-216272
308 A test campaign has been performed, and demonstrates the good behavior of 1-meter long boom, after  deployment and rigidization (2). For a 4,6-meter long boom with 160-mm diameter, the maximum flexure load is 6 Nm and maximum compression load 60 N.  

Figure 11: 3-meter wing breadboard 

Conclusion 

The choice of the passive aerobraking to respect the IADC code of conduct, is validated, for the Myriade  microsatellite family without propul sion system. The development of this equipment is linked to the  MICROSCOPE project. The deorbiting progress wi th deployable aerobraking appendage is under CNES  patent (3).   In 2007, inflatable aluminum laminate boom and deployable membrane have been tested in microgravity environment during a 0g flight test campaign. A 3- meter wing breadboard has been manufactured and  tested on ground by Astrium Space Transportation.   Technological solution for deployment and rigidization are chosen and validations are well advanced. Currently the project is ending of the preliminary des ign phase and the PDR review is scheduled for the  beginning of 2011. The achievement of the de-orbiting sub-system development is foreseen end of 2013,  with the delivery of the flight model for MICROSCOPE satellite.  IDEAS project is, more than the development of a product dedicated to MICROSCOPE, the development  of a concept and its associated technology that can have other potential others space applications.   References 

1.  2007Aluminum laminate technology for micro-sate llite Deorbiting aerobrake system“ – O. LE COULS,  D. LACOUR (EADS Astrium) , C. DUPUY (CNES) – 3 rd European Workshop on Inflatable Space  Structures, ESTEC, October 12th 2006  2.  The innovative deorbiting aero brake system “IDEAS”: The gossam er technology for Micro-satellite  deorbiting – 10th European Conference on Spacecra ft Structures, Materials & Mechanical Testing,  Berlin 10-13 September   3.  Satellite airbrake devic e Patent n° WO2007096428   NASA/CP-2010-216272