ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 17. Change in ultrasonic TOF as a function of temperature for a 24-cm  (9.5-in) unloade d tension rod. Summary Ultrasonic TOF measurements were successfully implemented as a means for determining the preload of  installed bolt elements for Frangibolt®actuators. For each torqued bolt, the load was derived from the  measured load-induced ultrasonic TOF change using an average conversion factor determined from load-frame tests. The discrepancy between the ultrasonic and load cell preload readings for 6.3-cm (2.5-in) long, notched bolts nominally torqued to 13,344 N (3000 lb) load was less than ± 445 N (100 lb) or  ± 3%. Similar results were obtained for 24-cm (9.5-in) tension rods. The discrepancies observed likely  contained contributions from bolt-to-bolt variations in the actual length loaded, the bolt temperature and  bolt geometry. Improvement in the measurement fidelity could be achieved by calibrating the bolt elements individually as opposed to using an average calibration factor. The data suggest that a onepoint calibration at 8896-N (2000-lb) load would suffice for the 6.3-cm (2.5-in), notched bolts. Having to  load test each bolt, prior to tightening, however is probably not worth the marginal improvement that would be attained. A commercially available Bolt Tension Monitor with a 0.25-GHz sampling rate was shown to be sufficient  for the pre-load measurements. The accuracy gained through use of a system employing a 1.0-GHz  digital scope was insignificant. For the notched bolts, it was found that monitoring the earliest maximum (or minimum) in the echo from the bolt end gave the most reliable load measurement result. Broadband transducers in the 5 – 15 MHz range were found to work effectively. The observed ± 3% error obtained with the ultrasonic TOF methodology for setting the bolt element preloads is a significant improvement over the ± 25% error observed with straightforward tightening to a specified torque. Acknowledgement Support of The Aerospace Corporation’s Independent Research and Development Program is gratefully  acknowledged. 256
References 1. TiNi Aerospace, Inc., 2505 Kerner Ave., San Rafael, CA 94901. 2. Bikford, J. H. and Nassar, S. (Eds.) (1998). /g43/g68/g81/g71/g69/g82/g82/g78/g3/g82/g73/g3/g37/g82/g79/g87/g86/g3/g68/g81/g71/g3/g37/g82/g79/g87/g72/g71/g3/g45/g82/g76/g81/g87/g86 , New York, NY:  Marcel Deckker, Inc., Chapter 25, Table 2, p. 674. 3. Table showing Accuracy of Bolt Tensioning Methods,  http://www.roymech.co.uk/Useful _Tables/Screws/Preloading.html 4. Allen, D. R., W. H. B. Cooper, C. M. Sayers, and M. G. Silk (1982). /g55/g75/g72/g3/g56/g86/g72/g3/g82/g73/g3/g56/g79/g87/g85/g68/g86/g82/g81/g76/g70/g86/g3/g87/g82/g3 /g48/g72/g68/g86/g88/g85/g72/g3/g53/g72/g86/g76/g71/g88/g68/g79/g3/g54/g87/g85/g72/g86/g86/g72/g86 , Research Techniques in NDT, Vol. 6 (R. S. Sharpe, ed.). New York, NY:  Academic Press, pp 151-209. 5. Smith, J. F. and J. D. Greiner. “Stress Measurement and Bolt Tensioning by Ultrasonic Methods,”  Journal of Metals (July, 1980), pp 34-36. 6. Pao, Y. H., W. Sachse and H. Fukuoka (1984), /g36/g70/g82/g88/g86/g87/g82/g72/g79/g68/g86/g87/g76/g70/g76/g87/g92/g3/g68/g81/g71/g3/g56/g79/g87/g85/g68/g86/g82/g81/g76/g70/g3/g48/g72/g68/g86/g88/g85/g72/g80/g72/g81/g87/g86/g3/g82/g73/g3 /g53/g72/g86/g76/g71/g88/g68/g79/g3/g54/g87/g85/g72/g86/g86/g72/g86 , Vol XVII (W. P. Mason, ed.). New York, NY: Academic Press, pp 61-143. 7. Dakota Ultrasonics, 1500 Green Hills Road, #107, Scotts Valley, CA 95066 8. K. Salama and C. K. Ling, “The Effect of Stress on the Temperature Dependence of Ultrasonic  Velocity,” Journal of Applied Physics, 51, no. 3 (March 1980), pp. 1505-1509 All trademarks, service marks, and trade names are the property of their respective owners 257
LARES Mission: Separation a nd Retention Subsystem Alessandro Bursi*, Pierluigi Camilli*, Claudio Piredda**, Gianni Babini* and Elio Mangraviti*  Abstract As part of the Lares (LAser RElativity Satellite) mission, an all-Italian scientific mission launched with the  Vega maiden flight in February 2012, a mechanical separation and retention subsystem (SSEP) has been  developed to retain the LARES satellite during launch and release it in the final orbit. The design flow was  based on the identification of the driving requirements and critical areas to guide the trade-off, design, analysis and test activities. In particular, the SSEP had to face very high environmental loads and to minimize the contact areas with the satellite that had a spherical shape. The test activity overview is provided. Introduction The scope of this paper is to describe the activities done for the design of the SSEP (Separation  Subsystem) developed for the LARES program, give an overview of the test campaign performed, and  show the relevant main results underlining the acquired experience.  LARES Mission Main Objectives LARES (LAser RElativity Satellite) is an all-Italian scientific mission, developed for the Italian Space  Agency and successfully deployed in orbit by the European launch vehicle Vega on its maiden flight on  February 13, 2012. The main goal of the LARES mission is to take high accuracy measurements of the  Lense-Thirring effect, i.e., the dragging of inertial frame due to additional space-time curvature produced by the Earth’s angular momentum. Two secondary objectives were assigned to the mission:  to provide a  separation platform for additional payloads and to support the launcher qualification. In order to support  Vega qualification, the system included a stand-alone avionics system devoted to measure the launcher  payload area environmental conditions during flight and to provide video of launcher lift-off, stage separations (by using an external camera), and payload separation in the target orbit. LARES System Architecture and Subsystems The LARES system is composed of the following main elements: /g131The LARES satellite, a tungsten sphere of ~400 kg with 92 Corner Cube Reflectors embedded; /g131 The Separation and Retention Subsystem (SSEP) to retain the LARES satellite during the launch  flight and release it in the final orbit; /g131The Support Subsystem, the main structural interface with the launch vehicle upper stage and to  support all the subsystems of LARES system; /g131Avionics and Harness Subsystem, the subsystem to provide separation of the payloads, to assist the  launch vehicle in acquisition, to monitor key parameters inside the launcher fairing and the  separations of the launch stages during ascent by means of video cameras, and to download data to ground stations. /g131Eight secondary payloads. *CGS S.p.A., Milan Italy **Consultant 259
Figure 1. LARES system before launch The industrial contract for LARES development was given to the Compagnia Generale dello Spazio as  prime contractor, and a structure of industrial and universities subcontractors for the different subsystems. In particular, Rheinmetall S.p.A. satellite line of business (the line of business later merged with CGS),  designed, manufactured and tested the Separation Subsystem.   SSEP Design Driving Requirements and Critical Areas Identification The main goal of this kind of Subsystem was meeting the following challenging, severe and contradicting requirements:1. To be stiff with the LARES satellite installed (lateral first frequencies higher than 70 Hz, while vertical  frequency higher than 140 Hz); 2. To withstand high environmental loads (i.e., quasi static load of 15 g in each orthogonal directions  applied to the 400-kg payload);  3. To lock a spherical body tungsten made by minimizing the contact areas; 4. To be single point failure tolerant in terms of release function (degraded kinematical separation  conditions were acceptable); 5. To transmit negligible shock to the LARES during the separation The structural/stiffness requirements (number 1 ,2, 4 and 5) were specified by the launcher in order to be  as much conservative as possible as this was the VEGA launcher’s maiden flight. The third one was a  scientific requirement aimed not to alter the spherical shape of the satellite that for scientific purposes  shall be as perfect as possible so as to be modeled as a spherical body. To fulfill the above requirements,  the following critical areas have been identified:  1. Reliable, simple and failure tolerant design concept identification;  2. Suitable separation device identification /g198only non-explosive initiators have been considered;  3. Separable surface design optimization in terms of material selection (with the satellite side fixed),  geometry and preload; 4. Robust and repeatable alignment and integration process identification. Trade studies The first trade analysis for the concept selection was performed by taking into account the above listed  critical areas. Existing or similar applications have been used as starting point. Several configurations have been analyzed considering applicable requirements in terms mainly of mechanical environment,  260
The Equatorial Solution has four equatorial pin-pushers blocking the sphere. Each pin pusher provides a  preload by means of non-explosive actuators. During the flight phase, the pins are hardly pressed in the cavities by the devices; once the satellite has reached the separation altitude, the separation devices are activated by removing the preload on the pin-pushers and allowing the pin retraction via springs in order  to disengage the satellite. In Figure 3 and Figure 4, the preliminary design of this solution and the two pin  working positions are shown. Figure 3. Equatorial SolutionFigure 4. Pusher working positions The benefits of this solution were simplicity, stiffness and the failure tolerance because it is possible to  design the interface geometries between pin and satellite to allow the disengagement of the satellite even  if a separation device fails; but a problem exists with the loads acting on the pin when the structure is preloaded. The retracting spring on each pin had to be bigger than available space for accommodation.  Solution n. 4 (Hinge Solution) This solution is an upgrade of the equatorial one, because in it LARES engagement is ensured by four  hinges in order to avoid the friction problems in the contact points between LARES and the four pins. Thestructures must be compacted to improve the stiffness, the vibration behavior, and to reduce the total  mass, as shown in Figure 5.  Figure 5. Hinge Solution This solution collects all the advantages of the previous solutions by minimizing the disadvantages; for  this reason this has been selected as final configuration for LARES SSEP. The separation device geometrical position and number (4 actuators) required a very good simultaneity in the separation. This simultaneity requirement coupled with high strength and power budget boundaries lead to the usage of a NEA Electronics solution to initiate the LARES separation. In particular, the actuator had to withstand a maximum 195000-N static load applied in compression along the pin axis and had to separate properly  262
when subjected to a total compressive load of 148000 N along the pin axis. Then the release mechanism,  considering an infinitely rigid interface flange, had to guarantee a minimum axial stiffness (against the load direction) of 219 MN/m. Furthermore, the release mechanism allows an angular misalignment of rod corresponding to a 10° cone under the loads above specified. LARES SSEP Selected Configuration Summary The selected configuration for the SSEP is composed of the following elements:  /g131A supporting structure that guarantees lower mechanical interface toward the LARES System support  structure (white structure in Figure 1).  /g131An upper main mechanism that is the mechanical interface to the LARES satellite, including: /g190The interface pins to the satellite /g190The hinge system, which allows the retention of the satellite during the launch phase and its separation on orbit. /g190The initiator device that releases the hinge system /g190The switches indicating the opening of the first mechanism and the activation of the second one. /g131A support and preloading system /g131An ejection LARES mechanism. /g131An electrical interface to the avionics, including: /g190The switches lines /g190The initiators lines The functionality of the release mechanism is explained in the separation sequence shown in Figure 6, from locked configuration (on left) up to the satellite separation (on right):  Figure 6. Separation sequence (from left to right) and functional explanation When LARES system reaches its final orbital position and the satellite has to be released, an electrical  signal is sent to the actuators that disengage the pin pulled back by means of a retraction spring, thus  allowing the hinge to rotate. Hinge rotation is also driven by two spiral springs suitably preloaded. Once the LARES satellite is set free by the main separation mechanism, the compression spring of the  ejection system, in the zero-gravity condition, pushes the LARES satellite along the vertical direction. At  the end of compression spring elongation, the LARES satellite will have the nominal vertical speed of  0.75 m/s. The push-out spring has been designed to provide a force of 350 kg for the safety reason to  avoid LARES ejection during the operation on ground in case of accidental activation of the SSEP. An  electrical device will confirm the release operation. Support and preloading system Ejection mechanism de ploys the satellite with the re quired velocit y.The hinges are blocked and preloaded on this beam. Once the separation devices are  activated, this beam is retracted then the hinges are able to rotate making the satellite free 263
A section of the SSEP with the satellite installed and with the description of the preload path is shown in  Figure 7.

Figure 7. SSEP Functional Scheme The contact surfaces between SSEP and LARES satellite are: /g120LARES and support and preload system: Sphere on cone; the SSEP part is covered by Kapton, to avoid LARES surface damage. This contact area is so large that the Hertzian stress is not  critical (negligible); /g120LARES and hinges: sphere on sphere (cavity on satellite and pin on SSEP). The sphere  geometries have been optimized to have an acceptable level of Hertzian stress; in addition, to  avoid cold welding issues, Braycote lubricant has been used in between. Both caves on satellite  and hinge pins have spherical form. Integration approach In order to implement the theoretical preload philosophy that has a uniform vertical preload applied by the  support system and the hinge preloads acting symmetrically and balanced in the equatorial plane (the preload of two hinges on the same axis null the force contribution in this direction if the forces are the  same), an excellent alignment has to be reached during the integration. The SSEP by design was suitable to reach this theoretical condition, but the integration procedure had to be carefully tuned. In particular, the support and preload system has been instrumented with 4 strain gauges placed at 90° in  order to evaluate the preload distribution and uniformity on the whole bottom circumference, and each separation device rod has been instrumented too by using strain gauges to measure the force applied to  the separation device itself. Additional strain gauges were placed on the structure of the Qualification  Model to verify the sensitivity of the preload distribution with respect to the integration approach. Figure 8 shows three integration steps. The satellite sits on the support and preload system . A  vertical preload is applied to the satellite bottom part by  pressing this red item against the satellite surface (by  screwing the blue screws). This preload is needed to  avoid satellite gapping during vertical sine vibration  (equivalent static acceleration 15 g) The hinges are in place, engaging the satellite equatorial  cavities . By screwing this  element, the hinge rotates,  applying pressure to the satellite  on side A and on the separation  device shaft on side BSide A Side B 264
Figure 8. SSEP Flight Model integ ration with LARES Flight Model The detailed definition of the integration procedure and the decision to perform the shear pin holes  execution in an assembled configuration guaranteed the capability to reach a final integration fully  compliant to the requirements and with an extremely high repeatability.  Model philosophy and tests Taking into account the criticality of the subsystem and the requirements to satisfy, the following model  philosophy was adopted: /g131Two (2) Breadboards models /g131One (1) Qualification Model /g131One (1) Flight Model In particular, the Breadboards were to be representative of the two most critical elements - an arm of the  bond and the spring of separation / boost system. These models were used for the characterization of the  hinges from both the structural point of view (stiffness) and from the functional point of view (torque margin, deployment main parameters). The Qualification Model (QM) was representative of the flight model but the satellite was the LARES Demonstration Model (DM) that was identical to the flight for interfaces but it was the 6% heavier than the flight model to be conservative. Philosophy verification /g37/g85/g72/g68/g71/g69/g82/g68/g85/g71/g86 :/g3On these models, characterization tests were performed at the extremes of temperature in  a climatic chamber (operative range 10°C - 40°C). A cycle in thermal vacuum (non-functional) to verify the  compatibility of materials (lubricants) with the space environment was performed too.  /g52/g88/g68/g79/g76/g73/g76/g70/g68/g87/g76/g82/g81/g3/g48/g82/g71/g72/g79 : On this model the following tests were carried out: /g131dimensional test /g131mass measurement /g131vibration /g131separation test at the critical temperature (highlighted by the tests on the breadboard ) /g131thermal vacuum test exclusively for the electrical part /g41/g79/g76/g74/g75/g87/g3/g80/g82/g71/g72/g79 : On this model the following tests were carried out /g131dimensional test /g131mass measurement /g131functional acceptance tests at room temperature (postponing system-level environmental testing). Breadboard #1 tests /g50/g69/g77/g72/g70/g87/g76/g89/g72/g86/g3 /g131To demonstrate the functionality of the pin and cave interface (separable surfaces) under operative  load (the critical separable surface); /g131To demonstrate the functionality of the pin and cave interface under operative load with a load factor  of 1.5. /g3 /g3/g3 265
/g51/g68/g86/g86/g3/g70/g85/g76/g87/g72/g85/g76/g68/g3 /g131No plastic deformation on the cave or pin under operative load (measured both via optical method  and Coordinate Measurement Machine (CMM)); /g131No plastic deformation on pin or cave under operative load time 1.5 measured both via optical  method and CMM);  /g55/g72/g86/g87/g3/g68/g85/g87/g76/g70/g79/g72/g3/g76/g71/g72/g81/g87/g76/g73/g76/g70/g68/g87/g76/g82/g81/g3 /g13115-5PH H1025 Steel Pin; /g131THA-18N Tungsten Cave; Figure 9. Steel Pin on left and Tungsten cave on right /g55/g72/g86/g87/g86/g3/g86/g72/g84/g88/g72/g81/g70/g72/g3 1. Visual inspection with electronic microscope (magnifcation: 7.5x; 10x; 20x; 40x);  2. Dimensional check with Leitz machine;3. B/B mounting under compression machine; 4. Visual inspection with photos;5. Greasing pin and cave with Braycote 600;6. Slowly applying compression load (26612 N), compression speed equal to 0.0003 mm/s;7. B/B demounting under compression machine;8. Visual inspection with electronic microscope (magnification: 7.5x; 10x; 20x; 40x) 9. Dimensional check with Leitz machine; 10. B/B mounting under compression machine;11. Visual inspection with photos following grease removal;12. Greasing pin and cave with Braycote 600;13. Slowly applying compression load times 1.5 (39918 N), compression speed equal to 0.0003 mm/s; 14. Visual inspection with photos following grease removal; 15. B/B demounting from compression machine;16. Dimensional check with Leitz machine;17. Visual inspection with electronic microscope (magnification: 7.5x; 10x; 20x; 40x). To perform a perfect correlation on dimensions in the various performed tests, a measurement points map was used on the cave and pin such that in a ll the checks the same points were measured.  /g55/g72/g86/g87/g86/g3/g85/g72/g86/g88/g79/g87/g86/g3 A microscope analysis has been performed after the test to look for eventual problems due to plastic  deformation on the steel pin and tungsten cave. Figures 10-12 show the test set-up and the microscope steel pin photos. The microscope analysis and the measurements with the Leitz machine showed that no  plastic deformation occurred during the test. 266
Figure 10. BB1 Test set-up                 Figure 11. Pin at 10x after first test Figure 12. Cave at 20x after the test (left) and before (for comparison) Breadboard #2 tests /g50/g69/g77/g72/g70/g87/g76/g89/g72/g86/g29/g3 /g131To measure the stiffness along the three principal directions of the restraint mechanism; /g131To correlate the FEM analysis previously performed on the restraint mechanism with experimental  data; /g131To evaluate friction coefficient of the restraint mechanism;  /g131To demonstrate restraint mechanism strength under operative load; /g131To demonstrate restraint mechanism strength under operative load with a load factor of 1.5 /g51/g68/g86/g86/g3/g70/g85/g76/g87/g72/g85/g76/g68/g3 /g131Check of the complete structural integrity of the BB#2 after the tests /g131The value of the friction torque, before and after the tests, must be compatible with the ejection of the  LARES satellite /g55/g72/g86/g87/g3/g68/g85/g87/g76/g70/g79/g72/g3/g76/g71/g72/g81/g87/g76/g73/g76/g70/g68/g87/g76/g82/g81/g3 The test article was composed of the restraint mechanism subdivided in its constituent components: /g131Hinge /g131Latch /g131Pin /g131Preload bolt /g131Actuator Dummy All the components were completely representative of flight components, including material, tolerances  and dimensions. 267
Figure 13. BB2 test set-up                     Figure 14. Restraint Mechanism /g55/g72/g86/g87/g3/g86/g72/g84/g88/g72/g81/g70/g72/g3 1. Dimensional Check; 2. Visual inspection;3. Measurement of friction torque;4. Evaluation of restraint system stiffness; a) Measurement of vertical stiffness; b) Measurement of transverse stiffness c) Measurement of longitudinal stiffness  5. Measurement of strength at nominal max load;6. Visual Inspection; 7. Measurement of friction torque;8. Dimensional check;9. Measurement of friction torque;10. Measurement of the strength at nominal load times 1.5; 11. Visual Inspection12. Measurement of friction torque 13. Dimensional check  /g55/g72/g86/g87/g3/g53/g72/g86/g88/g79/g87/g86/g3 /g131The comparison between the dimensional checks performed with Leitz machine before and after  strength tests assures that no plastic deformation occurred on the restraint mechanism during all tests performed up to 1.5 times the qualification load. /g131Friction coefficient measured values were acceptable and compatible with the ejection of the LARES  satellite, and didn’t show variation during the performed tests.  /g131The obtained results clearly demonstrated the soundness of FE modeling of the entire SSEP  subsystem and relevant components (measured stiffness was 20% more than the predicted one in all the 3 directions). SSEP QM vibration test /g50/g69/g77/g72/g70/g87/g76/g89/g72/g86/g29/g3 /g131To determine the mechanical characteristics of LARES SSEP (in terms of natural frequencies, etc.) /g131To demonstrate the validity of the mechanical design with respect to the applicable mechanical  environment /g131To verify that no degradation or malfunction of LARES SSEP QM components occurs during test. /g131To verify that no preload loss occurs before and after testing /g3 /g55/g72/g86/g87/g3/g68/g85/g87/g76/g70/g79/g72/g3/g76/g71/g72/g81/g87/g76/g73/g76/g70/g68/g87/g76/g82/g81/g3 The LARES SSEP QM Model is shown in Figure 15.  268