ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 15. LARES SSEP Qualification Model on the vi bration shaker for lateral test on left and for  vertical test on right Figure 16.  Accelerometer position and vibration test reference system /g55/g72/g86/g87/g3/g86/g72/g84/g88/g72/g81/g70/g72/g3 On this model, the following tests have been done: 1. Dimensional controls 2. Mass measurement 3. Vibrations 4. Functional tests at the most critical temperature (found during the tests performed on the  breadboards) in climatic chamber 5. TVAC test on electrical components only 6. Separation test /g55/g72/g86/g87/g86/g3/g53/g72/g86/g88/g79/g87/g86/g3/g3 In the following figures, the results of low sine excitation before and after full-level sine for the  accelerometers positioned on the four hinges (M1-M4, see Figure 16) and on LARES satellite (M5, see  Figure 16) have been reported for the X direction (that is the vertical one directed as the launcher thrust  axis). 269
Figure 17. Low sine pre, X direction In Figure 18, the results of low sine excitation for the same accelerometers after full-level sine are  reported.  Figure 18. Low sine post, X direction The previous data show that after a full-level sine a little structure stabilization is present, as revealed by  the slight difference between the resonance frequency before and after full level sine test (about 3%). Final value of the first natural frequency on X direction is 230 Hz, greater than the 140 Hz (requirement). Figure 19. Full sine test, X direction00.30.60.91.21.5 10 100 1000HzgMP 1 MP 2 MP 3 MP 4 MP 5 Control 00.30.60.91.21.5 10 100 1000 HzgMP 1 MP 2 MP 3 MP 4 MP 5 Control 0246810 0 50 100 150 200 250 300 35 0 HzgM1 M2 M3 M4 Control M5Vertical axis = 230 Hz X axis = 211.2 HzX axis = 237.7 Hz 270
The full sine test showed a lower value of the first frequency, demonstrating a certain dependence of this  parameter with sine amplitude. The low sine test performed after this full sine confirmed that no damage  occurred in the mechanism.  /g55/g72/g86/g87/g3/g38/g82/g81/g70/g79/g88/g86/g76/g82/g81/g86/g3 /g131The Vibration Test has been performed on the SSEP QM and was successfully completed. /g131The satellite was in flight configuration and all the success criteria have been satisfied. /g131The test confirmed the FEM prediction (the SSEP behaves like a 1 DoF system) /g120Loads acting on SSEP during tests are less than expected.   /g120In the two lateral directions, the same behavior has been measured, in particular 142 Hz first  frequency (70 Hz was the requirement) measured during low level sine, while 130 Hz first frequency  measured during the full level. The frequency shift on full level sine is usual when part of a preload structure. Visual inspection did not  identify any damage or physical degradation on the item under test. SSEP QM Separation tests /g51/g68/g86/g86/g3/g38/g85/g76/g87/g72/g85/g76/g68/g3 /g131Test conditions according to specification requirements (worst temperature distribution on mechanism  was applied); /g131“Clean” separation without any problem; /g131Separation with an ejection speed 0.6 < v < 0.9 m/s; /g131Correct activation of the end of stroke switch; /g131Acquisition of all the test data (temperatures, video recording, NEA activation parameters, ejection  mechanism strains). Figure 20. LARES QM Separation Test configuration Figure 21. LARES QM Separation filmed sequence – 5 27 ms of elapsed time from the first to the  last photogram 271
/g55/g72/g86/g87/g86/g3/g70/g82/g81/g70/g79/g88/g86/g76/g82/g81/g86/g3 /g131The Separation Test was performed on the SSEP QM + LARES DM and was successfully completed. /g131The SSEP was in flight configuration and all the primary success criteria have been satisfied.  /g131Visual inspection did not identify any damage or physical degradation on the item under test (no  damage occurred after vibration test). /g131All the additional controls, measurement and investigations did not show evidence of any degradation  or loss of performance of any of the mechanical components. During the Launch Phase, the on-board camera mounted on the Support Subsystem has shown the  nominal performance of the SSEP subsystem.  Figure 22. Picture of LARES Satellite release on orbit  Lessons Learned The ‘lessons learned’ from the positive results obtained in the activities performed on LARES Program  SSEP is the confirmation of the importance of: /g131developing a design oriented to the integration feasibility; /g131using an “over-sensorized” model (in this case SSEP QM) to verify the sensitivity of the integration  and alignment mismatches vs performance requirement in order to minimize the influence of operator errors;  /g131adopting a suitable optimized model philosophy that can allow deep investigations of the critical  areas; /g131focusing strong effort in the testing activity at breadboard level to correlate in advance the  mathematical models in order to minimize the technical risk (the BB model testing phase has been  included in the timeframe of CDR). The importance of the above choices becomes more evident in those programs characterized by very  stringent and challenging requirements such as LARES.  LARES successfully  released on orbit SSEP 272
Testing Orion’s Fairing Separation System Henry Martinez*, Chris Cloutier*, Heber Lemmon*, Daniel Rakes*, Joe Oldham* and Keith Schlagel* 

Abstract Traditional fairing systems are designed to fully encapsulate and protect their payload from the harsh  ascent environment including acoustic vibrations, aerodynamic forces and heating. The Orion fairing separation system performs this function and more by also sharing approximately half of the vehicle  structural load during ascent. This load-share condition through launch and during jettison allows for a substantial increase in mass to orbit.   A series of component-level development tests were completed to evaluate and characterize each  component within Orion’s unique fairing separation system. Two full-scale separation tests were  performed to verify system-level functionality and provide verification data. This paper summarizes the fairing spring, Pyramidal Separation Mechanism and forward seal system component-level development  tests, system-level separation tests, and lessons learned. Introduction The fundamental components of the Orion Multi-Purpose Crew Vehicle fairing are similar to heritage  systems as shown in Table 1. However, addition of the load share between the fairings and the Service  Module requires a unique architecture and presents challenges requiring substantial testing in order to validate the design can successfully jettison loaded fairings and ensure the safety of Orion astronauts. Table 1. Fairing System Comparison Vehicle Segments Vertical Separation Hori zontal Separation Separation Energy Orion MPCV Trisector Frangible joint Frangible joint & PSM Mechanical springs Ariane 5 (1) Bisector Thrust Rail Frangible joint Thrust rail Atlas V 400 (2) Bisector Pyro Sep Bolts V-Clamp Spring thrusters  Atlas V 500(2) Bisector Thrust rail Sep bolt Thrust rail Delta IV (3) Bisector Thrust rail Sep bolt Thrust rail Delta IV Heavy(3) Trisector Thrust rail Sep bolt Thrust rail Pegasus (4) Bisector Sep bolt / clamp band Frangible joint Gas thrusters Saturn V/Skylab (5) Quad Thrust rail Latchpins Thrust rail Extensive development and qualification testing was and is currently being performed on each of the  fairing components. This paper focuses on the early development testing for the fairing Pyramidal  Separation Mechanism, spring assembly, and forward seal as well as results from the more recent system-level fairing separation testing. *Lockheed Martin Space Systems Company, Denver, CO /g51/g85/g82/g70/g72/g72/g71/g76/g81/g74/g86/g3/g82/g73/g3/g87/g75/g72/g3/g23/g21/g81/g71/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/g3 273
Orion Fairing Design Overview The Orion fairing system, referred to as the Spacecraft Adapter Jettisoned Panels shown in Figure 1, provides protection for the Service Module (SM) on the pad and during ascent as well as carries approximately half of the vehicle structural loads up through jettison. During the Orion Exploration Flight  Test -1 (EFT-1) mission, the fairings are jettisoned under approximately 0.2g of thrust at an altitude of approximately 1830 m (600,394 ft). Figure 1. Orion Configurati on (Image Courtesy: NASA) The fairing jettison system shown in Figure 2 is comprised of composite tri-sector segments which are  severed by frangible joints and pyrotechnic separation bolts, then jettisoned with springs and hinges. The  frangible joints are the primary load path between each of the fairing panels along the horizontal and vertical seams. Each frangible joint contains a single core charge initiated by Explosive Transfer Line at opposite ends for redundancy.  Figure 2. Fairing Components After severance of the frangible joints, each panel is released by two redundantly actuated pyrotechnic  separation bolts between cup-cone interfaces called Pyramidal Separation Mechanisms (PSMs). ThePSM transfers load from the inner load path (SM) to the outer load path (fairing panels) and allows  274
clearance of these load bearing surfaces during deployment. Made from titanium with CANADIZE® coating, the PSM incorporates variable angle surfaces with shallow angles taking shear loads and  steeper angles providing clearance. Selection of the primary load bearing angle involved a careful  balance of minimizing axial force induced into the pyrotechnic separation bolt and frictional resistance during separation. Once the PSMs are released, the fairings are pushed away with titanium spring actuators which rotate the  panels about hinges along the lower separation plane. The Push-Off Springs are located near the top of  the panels adjacent to the PSM and provide initial deployment force for the fairing panels. There are two  springs per panel sized to provide positive margin in the event of a spring coil fracture per NASA human rating requirements.  The forward seal system is a circumferential seal comprised of an elastomeric p-shaped seal compressed  by titanium flexures called airstream deflectors. Together they create an environmental seal between the  top of the fairing panels and prevent air ingestion during ascent while allowing for relative deflections  between the sealing surfaces. The seal system also acts as a conductive path between the avionics ring and the fairing panels.  Component Testing Component-level development testing evaluated the individual components of the system and provided  valuable information allowing for mass and performance optimizations. Load testing was performed on  the Pyramidal Separation Mechanisms to ensure the material and coatings could endure the high loads and cyclic wear during ascent and still separate. Spring testing was performed to validate design margins  and evaluate performance under the stressing ascent vibration environment. The seal system was cycled  in order to validate life requirements. Each of these tests provided insight into the performance of each  component against its most stressing environment and in most cases required some modification to  ensure all performance requirements would be met during qualification. Pyramidal Separation Mechanism Testing Tests were conducted to evaluate the performance of the cup-cone design to bear load and the effectiveness of the lubricant. Figure 3 shows a cross section of the flight design. Figure 3. PSM Cross Section 275
There were three rounds of initial PSM development testing that focused on validation of the ramp angles,  material and coating performance. The first test identified refinement of the load bearing surfaces and manufacturing tolerances. The second test validated the refinements and changes learned from initial  testing and to predict slippage in the joint. The third test was used to verify that the material and coatings would survive the load and life requirements.  Test fixtures were designed to support the PSM halves and mount to a servo-hydraulic load frame. The  fixture geometry was designed such that the prescribed ascent loads and moments could be achieved via a single load-axis. Two fixture sets were machined, one for a maximum axial case and the other for a maximum lateral case. Modifications were made to these basic designs to include both forces and  moments when the predicted loads changed. In all, the test setup included 10 Linear Displacement  Transducers (LDT), an instrumented Strainsert bolt (first and second test), four strain gages mounted on  an inert flight-like bolt (third test), a main load cell to control the applied load levels, and an LDT to  monitor overall applied displacements. The first test revealed the critical load bearing surfaces were not engaging properly. In order to correct  this, shims were fabricated of variable thicknesses and placed between the two nodes to determine the proper gap. Changes were also made to the interfacing radii for each of the nodes and to the primary critical load bearing surfaces to minimize contact stress between the two halves. The second testconducted validated the changes made to the design. Slippage was more predictable with smoother  hysteresis deflection curves. While the test showed heavy burnishing there was no galling and the parts easily separated. Figure 4 shows the wear differences between the first and second test configurations.  Figure 4. Wear Characteristics of Initial and Updated Design Concepts 276
The third test series used a flight-like inert separation bolt which was instrumented with four uniaxial strain  gages as well as a new test fixture. This test evaluated the preload operation and measured the resultant  axial and moment loads in the separation bolt over the increased life cycle count based on updated analysis. The cycle count increase was based primarily on updated ground cycle analysis which was typically low load, but had to be accounted for. Data from this round of testing (shown in Figure 5) was also used to determine load when the PSM slipped. This final test confirmed the design was robustenough to survive the loads at twice the number of life cycles and no further modifications were required.  Figure 5. PSM Load versus Displacement Spring Assembly Testing A “Test Like You Fly” approach was followed for spring testing with initial wear-in/run-in followed by random vibration then multiple functional tests to verify design margin. Figure 6 shows the general cross section of the spring assembly and its various components. Figure 6. Fairing Spring Assembly 277
Initial wear-in/run-in was performed to burnish close running fit areas and evaluate potential friction losses  within the assembly by measuring the force versus displacement over the working height of the spring.Test data showed minimal losses (1% or less) compared to the individual spring coil testing performed by the vendor.  Random vibration test levels initially started out as benign but as dynamic modeling matured the levels  more than doubled. This gave cause for concern of excessive loading of the spherical bearing within the assembly which allows for proper spring alignment during deployment. Under random vibration testing,  the spring surged in a smooth sinusoidal motion along the coil axis for the most part but at certain levels  the peak displacements were significantly amplified and chaotic. This motion induced higher than  expected loads into the spherical bearing. Another noteworthy event was the amount of off-axis displacement resulting in cyclic impacting of the coil with its adjacent assembly hardware. At qualification levels the titanium spring damaged the TUFRAM ®- coated aluminum parts and generated debris which was deemed unacceptable. Three options were  selected and tested; bare aluminum, aluminum with hard anodize coating, and 304 CRES stainless steel.Ultimately, the soft 304 CRES stainless steel performed the best by absorbing the impact energy. Posttest results had the appearance of peening but without the effect of creating debris. Figure 7 shows post  vibration test results of the original anodized aluminum and the final 304 CRES design.   Figure 7. Anodized Aluminum and 304 CRES Vibration Test Results Another issue noted was ground bonding paths. The use of a spherical bearing and a dry film lubricant in  the race did not allow for a clean path resulting in intermittent grounding.  Functional testing was conducted after vibration testing to be more flight like with regard to environment  exposure. The functional test verified energy margins and hard stop functionality. No issues were found  during functional testing. In fact, the functional test fixture was modified to accommodate lanyard  connector testing using the same development springs. Overall, the development springs have become a  great test asset with some being functioned as many as 15 times with no detrimental damage or wear.  278
Forward Seal Testing The forward seal system shown in Figure 8 is comprised of two components; a specially designed three segmented elastomeric seal and titanium aero deflectors.  Figure 8. Forward Seal System The aero deflector is essentially a flexure capable of protecting the elastomeric seal from the harsh  ascent thermal environment while compressing it against the sealing surface and providing a ground  path. The shape of the aero deflector was complex due mainly to the large design deflections (up to 1.3 cm (0.5 in) between sealing surfaces) and required a combination of conventional machining and wire  Electrical Discharge Machining (EDM) to maintain the tight manufacturing tolerances. The original  avionics ring surface tested was aluminum with electroless nickel which provided a smooth finish withhigh hardness and excellent wear resistance without any lubricant. However, application of the  electroless nickel to the large avionics ring surface was found to be prohibitive due to cost and schedule  constraints. Brushed-on nickel coating was eventually chosen. Figure 9 shows the wear on the base  material after 2000 cycles of both electroless nickel and brushed-on nickel coatings with no lubricant. From these initial tests it was determined that the addition of lubricant was required if brushed-on nickel plating was to be used.  Figure 9. Nickel Coating Wear after 2000 Cycles 279
Numerous tests followed trying to find a lubricant that would provide the necessary wear resistance with  the brushed-on nickel.  Eventually, the trade study was narrowed to five lubricants; Nye Lubricants Inc. special formula paste, Braycote®602EF, Krytox®, RheolubeTM 2004, and RheolubeTM2004 with MOS 2. The first test conducted  was an oven test to evaluate the resistance to controlled heat of 66 /g113C (150 /g113F) for a short duration (to  simulate ascent heating). While several of the lubricants shown in Figure 10 did not remain in place, there  is some concern that the test was not representative of flight.  Figure 10: Controlled Oven Test Each of the five lubricants are currently planned to undergo a 30-day environmental exposure test at the  Kennedy Space Center. In addition, a final round of wear testing will be performed. After these tests are complete, the final lubricant selection will be made. System Testing Overview Due to the dynamic nature of fairing jettison induced by the load share event as well as the complex  interaction between mechanisms, a system-level test was performed to verify each of the components worked together under flight-like separation conditions. The series of tests designed to verify performance  and gather data for the analytical model included test fixture characterization, panel-level testing and fullscale separation tests. Multiple separation tests were required to understand the sensitivity of the system to load and thermal conditions.  Test Fixture In order to achieve flight-like boundary conditions, dynamic response and load share, a special test fixture  was designed to simulate the Service Module. The fixture incorporated adjustable features allowing the overall structure to be simplified for manufacturability and cost while ensuring proper axial stiffness and dynamic response. The primary stiffness tuning feature was a simple spring plate located at each of the six load path interfaces. The second method for tuning of the structure was adjustability of the load head interface.  Structural stiffness and modal testing of the fixture was performed to help with tuning of the stiffness,  adjusting load share, and anchoring the analytical model. Stiffness testing included thirteen load cases to  evaluate the axial stiffness of the fixture and PSM interface. Instrumentation of the structure included 156  280
strain gage channels, 28 displacement measurements and 6 load cells in addition to input load and  displacement. Most loads were input using hydraulic actuators tied down to the floor or the reaction frames. Figure 11 shows the test fixture including a frame used to mount instrumentation.  Figure 11. Test Fixture Stiffness Testing Moment loads were applied using pulleys and offloading weights in order to simplify operations. Each  load case was applied three times while displacement, load and strain gage data was recorded. The displacement and strain data was used to independently evaluate the stiffness and predict load share. While strain and displacement results were within 3% of each other, the overall initial stiffness prediction  was off by ~25%. It was found that interactions between interfaces at the load head and spring plate were more complex than initially modeled. Capturing these differences in the finite element model was crucial in correlating test data to model predictions. Ultimately, results were still within the tuning capability of the fixture and no design changes were required.  After initial axial tuning, local bending and torsion load cases applied to the PSM interface provided  additional information which led to further refinement of the stiffness and load input interface. The  adjustable features of the fixture were used to obtain loading of both the inner structure and the fairings within acceptable design parameters. In order reduce schedule, the final load share was adjusted to within 10% of the predicted value with load biased higher to the fairings to ensure a conservative test. After tuning of the stiffness, modal testing of the fixture was performed to capture flexible body modes in  axial, bending and torsion with 129 degrees of freedom over 65 grid points. Ten target modes were initially identified over the frequency range of interest (0-50 Hz) and eleven were excited during testing.Results were then used to refine the finite element model. Similar to a fixture evaluation for vibration or  acoustic testing, fixture stiffness characterization is recommended for dynamic separation tests which rely  on proper boundary conditions and loads. Even with advances in finite element modeling, complex  boundary conditions can be difficult to predict without some level of testing. Adjustability within the Orion fairing test fixture accounted for uncertainties in the predicted stiffness while allowing for simplification of the overall structure to reduce cost and fabrication schedule. In addition, the characterization provided insight into potential model uncertainties with the flight structure and provided valuable information for  correlating the test fixture model. Characterization of the flight Service Module is currently scheduled for  early 2014. 281