ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
DMC-CFESAT (launched  2007)No snubbers. 908-mm panel  length. Original HDRM and Hinges DMC-UK2 (launched 2009) Lower snubbers only. 482-mm panel length.  Original HDRM and Hinges. EXACTVIEW-1 (launched  2012)Lower and upper snubbers. 641-mm panel length. Original HDRM and Hinges. Orbital Test Bed (launch expected 2015) No snubbers. 908-mm panel length. New HDRM evolution and  original Hinges. Figure 3. Evolution of the SSTL deployable Solar Array History of HDRS Original Development for DMC-CFESAT Original Hinge Development for DMC-CFESAT Initial concepts were based on the Surrey Rigid Array program, which took place at a similar time and  used tape springs for a 2-panel 1-yoke system with ~1.2 x 1.5 m panels for ~1-kW power generation (see Figure 4). The tape springs provide simple low part count mechanisms but due to their nature during 180°  rotation, there are points of negative torque margin, and when coupled with the high retarding torques of the flexprint power harness, a positive torque margin cannot be guaranteed. Initial testing proved that this  could be overcome by giving the array an initial push-off, however, there was permanent deformation  (creases) exhibited in the tape springs and although they still worked, it was deemed not suitable for flight  systems which require multiple deployments, so a new hinge concept was undertaken.  Figure 4.  Surrey Rigid Array in deployed position The subsequently developed hinge system design is relatively simple, utilizing two hinges very similar in  design; one fixed and one free to slide on its shafts to manage the material thermal expansion mismatch  from the CFRP solar panel and the aluminum body of the spacecraft. The constraints of limited volume  and mass for the hinges meant that a damped system could not be implemented. Instead, the system  HDRMs SnubbersHinges 423 HDRM  s Snubbers
The HDRM unit is predominantly bespoke SSTL design in terms of the separation interfaces and  housings; however, the central release actuator is a purchased unit (Frangibolt and Shape Memory Alloy  (SMA)). The basic principle of the Frangibolt relies on a pre-compressed SMA tube and a titanium bolt  with a special notched section which is positioned through the cylinder and restrained at either end. When heated up to a temperature of ~/g28/g19/g219/g38, the SMA undergoes a phase change returning to its original size thus developing enough tensile force to fracture the notched section of the titanium Frangibolt. Redundant heaters wrapped around the SMA are used to apply the heat which only requires unregulated DC power from 10 – 100 V to activate, and with temperature sensors built-in this makes a very simple mechanism with no real moving parts. The Frangi-SMA actuator was selected because it was cost effective, could be reset in-house, and energized via low-power switches. The Frangibolt is thus designed to fracture in a controlled fashion at a specific tensile load, which occurs  because of a stress concentration region via a notch – the frangible portion of the bolt. The notch  however limits the load capacity of the Frangibolt and in turn, of the HDRM. In order to ensure that the  Frangibolt performs as intended, tight control over the applied preload must be established in order to absorb all clearances so that the subsequent SMA expansion ~0.1 mm is successful in fracturing the bolt, while not overloading the Frangibolt.   There were a few design requirements on the HDRM over the aforementioned. One requirement was that  the parts of the cup-cone should require no match-machining of parts or special alignment impact to the spacecraft structure or arrays. Instead, the cup-cones simply had to ‘bolt-on’ over multiple interfaces which could create up to 0.5 mm of axial and lateral alignment error. To do this, the HDRM was designed to include multiple parallel and tapered shims in the form of a fixed swash plate to allow for parallel and  angular misalignments. However, the shimming capacity (parallel and angular) were never used and always set at nominal as the spacecraft assembly errors were minor during the flight builds. The actuator also had to be easily re-set without dismantling the spacecraft or misaligning previously set HDRM and panel hinges. Figure 6 shows a cross section of the Frangi-HDRM assembly.  Figure 6. Cross section of the Hold  Down and Release mechanism (HDRM) 425
during the detailed investigations several lessons learned were made which were to be later implemented in the flight designs:   /g120Radii under some of the mating interfaces could have caused a gap if they interfered so  tolerances were tightened and 100% optical inspection was made on all Frangibolts and mating parts. /g120There was a hexagon broached section in the end of the Frangibolt used to stop rotation with an  Allen key while the nut torque was applied during assembly. It was realized that this section could  be too long and extend into the working section of the bolt under the main nut and thus changing the local stiffness of the bolt possibly causing elongation of the bolt during extension of the Frangibolt. This was tested and found to be inconclusive although during one tensile test the  broached section did fail at a low load and thus it was removed from the design – subsequently  changed to a male hex feature.   /g120Upon disassembly of the setup, a slither of metal was found in the nut and the thread of the bolt  was damaged. It was thought that debris could have prevented the nut from being correctly  torqued up thus insignificant preload being applied which when the SMA went through its  elongation it was not restrained and could extend / gap. The latter concern coupled with the fretting required indicated that accurate knowledge of the preload in  the frangible section of the Frangibolt was needed and so a new setup (Figure 8) was defined in a tensile  test machine. Here a known load ~7 kN is applied and then a locking nut installed to 1 N-m thusguaranteeing the preload in the assembly is greater than the newly defined 6.5 kN required on the front.  Figure 8. Frangibolt preloading set up Challenges Experienced with HDRM Actu ator Selection on Later SSTL-100 Missions DMC-UK2 The HDRS successfully deployed all four DMC-CFESAT solar arrays in orbit following launch in 2007.  Following this success, the HDRS was employed on another SSTL-100 platform type, DMC-UK2. For this  mission, the deployable panel was significantly smaller than the original DMC-CFESAT (Figure 3), therefore direct use of the HDRS was deemed feasib le. While the HDRS had sufficient load capacity to  cope with the panel’s induced forces during launch, the launch environment was harsher and solar array  panel deflections were a concern. In order to avoid possible damage to the solar cells, then snubbers  427
were incorporated into the design to limit the array deflection. DMC-UK2’s single solar array deployed inorbit in 2009 - a successful application of the heritage HDRS design. Problems Experienced During EXACTVIEW-1 Environmental Test Campaign The load capacity was suitable for the baseline two platforms – DMC-CFESAT & DMC-UK2 – however  problems arose during the third platform (EXACTVIEW-1). For this mission, the panel size grew. While still smaller than the original DMC-CFESAT panels, the relative distance from the hold down to the top of  the panel increased and a different launcher was selected late in the program, which imparted higher  loads into the structure. Subsequent analysis showed a coupling of a spacecraft mode at ~280 Hz  (Figure 9) with the deployable panel so two additional upper corner snubbers were incorporated into the  design, which helped, but the design margins at the hold down against gapping were low to negative. As  this was late into the spacecraft EVT program with all hardware manufactured and built onto the spacecraft, the spacecraft was taken for a low-level vibration transmissibility test which confirmed the  analysis / negative margins and lead to a speedy re-design. Figure 9. EXACTVIEW-1 at vibration t est with HDRM sine loading shown All the interfaces were fixed and there was no time to procure / implement another hold down device with  a higher load capacity, so the preloading of the nut was investigated as preloading either side of the actuator using a torque wrench is known to have a large scatter. The torque scatter means that the resulting preload in the bolted joint could vary by as much as ±30% due to friction on the thread and washer. In order not to exceed the Frangibolt’s 6.5-kN load capacity, a nominal preload of 4.55 kN would be necessary and at the lower end of the scatter would result in a preload of 3.19 kN. At this lowest  tolerance preload, the resulting margin of safety against HDRM gapping was negative (gapping is the ability of the mating release interfaces to separate temporarily), this could lead to:  • HDRM pushing additional loads onto the hinges, and the hinge bearings being overloaded - thus  increasing friction and risk of deployment failure • During dynamic vibrations, fretting on the release interface could occur, which could lead to cold  welding preventing the panel from deploying • Gapping hold down portions could in turn induce additional lateral loads and bending moments on  the Frangibolt which could cause it to fracture prematurely and deploy during launch. It was not possible to implement a load cell or load washer into the design so the HDRM housing was redesigned and instrumented with strain gauges to monitor deflection to better determine the preload reducing the scatter to ±10% (see Figure 10). As a result of this method, the nominal preload could be increased to 6 kN, however, this alone was only just sufficient to avoid the potential for gapping so a  further change was made to achieve more favorable margins of safety by employing the true cup-cone interface instead of the previous flat-on-flat friction contact used on CFE-SAT and DMC-UK2. Shear Tests 428
were conducted, that showed the true cup-cone interface could support approximately 25% increased  lateral forces for the same preload than a flat-on-flat friction contact (see Figure 11).  Cone with strain gauges installed Figure 10. EXACTVIEW-1 HDRM Modifications Figure 11. Deployment interface shear test setup co mparing flat-flat interface with cup-cone  interface, for a nominal preload of 6075 N  The cup-cone interface and reduced preload scatter was sufficient to change the baseline negative  design margins to positive for the EXACTVIEW-1 mission, whose panel successfully deployed in-orbit in  2012. While the campaign was successful, it highlighted the design’s load capacity limitations and the criticality of the preload method. The man effort now associated with using this design due to the different assembly and calibration steps had increased the overall unit price of the system. When also considering the increased analysis non-recurring effort which would be required on each future program due to the  Frangi-HDRM’s sensitivities, a recommendation was made by the material review board to implement a  different higher load capacity solution on the next mission. The material review board’s analysis suggested a preload greater than 12 kN was required to give confident margins. It is also worth noting that DMC-UK2 and EXACTVIEW-1 platforms provided the benefit over the DMCCFESAT craft such that the HDRM and hinges are mounted directly to a single fixed body-mounted panel  interface, allowing the HDRS, static and deployable panels to be assembled and tested separate to the rest of the satellite structure (see Figure 2). Whereas, on DMC-CFESAT, the hinges mount to one  spacecraft interface, and the HDRM to another driven by the craft’s need for an alternative payload bay  Detailed cross section of the  modified HDRM assembly 429
HDRM at 0° deployment HDRM at 3° deployment Figure 12. Separation Nut and HDRM  arrangement and arcing effect of panel deployment profile Table 1. HDRM and Actuator Requirements Parameter Magnitude Notes Hold down actuator  load capacityMinimum load  capacity 15 kNThe selected Separation Nut is rated to loads <33 kN, hence providing margin even given ±30% torque scatter HDRM mass < 1.0 kg Operating voltage 20 – 33 VDC Operating current < 3 A Traditionally actuated by a low-power switch. Survival temperature-70ºC to +80ºC Based on preliminary mission analysis. Low Shock << 3000 g Physical Envelope Compatible with  heritage interfacesThe redesigned HDRM shall have the same spacecraft  interfaces to both the static and deployable panels. The  overall volume of the HDRM shall be largely unchanged,  and compatible with existing craft layouts. Due to the complex way in which the separation nut releases the spigot, the separation nut actuator  provides the limitation that it cannot be refurbished in-house and needs to be sent back to the supplier, in contrast to the original actuator where the frangible bolt could merely be replaced. This external  refurbishment limitation was outweighed by the confidence obtained through SSTL’s existing use of the separation nut and large margin on load capacity. The other limitation was that the separation nut actuator choice was also ITAR-restricted which limits the applications and slows the procurement process. However, there is a significant lack of European cost-effective, high-load capacity, low-shock  and proven hold down actuators, so there was no alternative at this time. Retraction and Non-Retracting Mechanisms Due to the spigot arcing hang-up effect described in Figure 12, it was foreseen necessary that the released spigot shall be withdrawn through the cup bore, prior to significant panel movement in order to  prevent the possibility of hang-up. The most elegant way to do this is to incorporate a retraction spring which pulls the released spigot through the deployable panel immediately upon deployment to ensure thatthe spigot is out of the way and will not clash. This is done using a compression spring mounted between  the cone and stud to act as a spigot retraction spring. The nature of the NEA's function means that the  deployment is low-shock; however, the retracting spring on the spigot will impart some low-magnitudeshock which shall be absorbed by a simple O-ring, instead of the previous honeycomb core material. CupCone Fixed Panel (craft) Deployed Panel (array)Separation nut  housingSpigotClash C 431
Initially, it was conceived that a retraction cap would be integrated onto the cone assembly to ‘catch’ the released spigot and stud shaft during deployment. Figure 13 shows that the retraction cap for a fullretraction system would protrude significantly above the panel surface with two notable drawbacks:  the likelihood of cell-shadowing in-orbit and the restriction of launch vehicle accommodation due to growth of craft footprint. The cell shadowing effect was deemed low risk since the small diameter (30 mm) protrusion above the array surface would only cause significant shadowing at large angles of incidence (60°), at which point the cell’s efficiency is significantly reduced and little power produced anyway. As an alternative to the retraction spring, it was also conceived feasible that the spigot could merely be  provided with a degree of freedom (rather than rigidly constrained to the cone), and thus without retraction, the spigot would clash with the cup during deployment but rotate and drag along the surface, rather than snagging and preventing deployment. Therefore, 3 configurations were envisaged:  full  retraction, partial retraction, and non-retracting (see Figure 13 ). The non-retracting variant has the  advantage that the cone can be kept low profile without a retraction cap, but the mechanism details are  less elegant. The degree of freedom is provided by a spherical washer interface, and the amount of spigot deflection needed to drag across the cup is <3°. Cap protrusion above panel = 35 mm Significant cell shadowing and accommodation concernsCap protrusion above panel = 28 mm Slight cell shadowing and accommodation concernsNo protrusion above panel. Increased risk of spigot hang-up .  Figure 13. Partial-retraction, Full-retrac tion and Non-retraction variants of HDRM HDRM Variant Breadboard Engineering Testing The 3 HDRM design variants were taken forward to  breadboard engineering testi ng in order to assess  which one to select and evolve to the final mechanism. For this testing, a simple separation nut spigot was used instead of a complete separation nut actuator, and rapid-prototype additive-layer manufactured plastic parts represent the cup, cone and retraction cap. This low-cost and quickly compiled test setup was invaluable for providing some initial indications of design success. All testing was done on the already existing sidewall deployment jig, with qualification model (QM) hinges  and panels. Two sized panels were used:  the smaller DMC-UK2 QM with short offset between the hinge  rotation axis and HDRM position, worsening the arcing effect; and the larger DMC-CFESAT QM panel.  The deployment tests were conducted in two stages:  initially a ‘light’ retraction spring was used (k =  0.189 N/mm) to reduce risk of damage to the plastic; then the testing was repeated with a ‘hard’ spring (the intended design, k = 0.98 N/mm). To give confidence, three deployment tests were done for each  variant across several deployment scenarios, varying the hinge spring torque and the push-off spring  presence to mimic different failure scenarios. The breadboard engineering tests provided the following conclusions: 432
Stowed: Released: Figure 15. Sep- Nut-HDRM Retraction Mechanism Overview Separation nut and carrier  assembly inserted – bayonet  interface of cup and nut carrier alignSeparation nut and carrier  rotated and bayonet with cup now engaged to support tensile load Figure 16. Separation nut carrier and cup bayonet load transfer Conclusions The revised Sep-Nut-HDRM assembly weighs <1 kg and supports up to 33-kN preload. The original  Frangi-HDRM design weighed 750 g and could only support 7-kN preload. The significant increase in load capacity far outweighs the small increase in mass; and allows for a very generic HDRM design for  future applications such as larger/multi-panel arrays or deployable boom/antenna. The QM Sep-NutHDRM has been built and is about to undergo hard-deployment testing. The decision has been taken that vibration testing is not necessary since the QM HDRM design load capacity is far larger than the intended  use, and thus the assembly does not experience any live load above its own preload. Instead, only  thermal hard deployment testing shall be undertaken on the QM HDRM to prove the design is fit for flight  while also covering the torque scatter possibility to avoid instrumentation and reduce recurring flight build costs. Subsequently, 4 FM HDRMs shall be built and integrated onto the craft during summer 2014. Retraction spring (compressed) Load-bearing  M10 nut Titanium  studSpigot shoulderCone element O-Ring and Spigot shoulder collisionRetraction  spring  (extended) elem m CupSeparation nut  and carrier  assemblyTool (not flight) 434
the TPS and Deployment Mechanism may be used on a future mission with similar requirements to the  original NPOESS mission. The Deployment Mechanism was successfully built and tested in 2012.Currently an engineering model with increased functionality is being fabricated, and the assembly will be  tested in the spring of 2014. Figure 1.  TSIS Rendering Table 1.  TPS Key Driving Requirements Requirement Method Precise tracking of the sun Gimbaled optical bench Stable thermal environment for optical bench Material selection, subsystem layout Steady-state pointing error of TPS with respect to sun  center <64 arcsec /g20/g305FSS (Fine Sun Sensor), precision  actuators, control system Dynamic pointing error of TPS with respect to sun center /g31/g25/g20/g3/g68/g85/g70/g86/g72/g70/g3/g20/g305Precision actuators, control system Operational Temperature Range -20C to +40C Material selection & platings, design Survival Temperature Range -30C to+50C Material selection & platings, design Natural frequency of TSIS during launch >50 Hz Launch latch design, structural design Natural frequency of TSIS on orbit >6 Hz Gimbal design, deployment mechanism TPS shall be operable in any orientation in 1G environment Gimbal design, deployment mechanism Uncompensated momentum contribution from TPS <±0.5  N-m per axisPrecision actuators, control system,  structural design 436
Thermal Pointing System Design The TPS was designed to meet the original NPOESS requirements summarized in Table 1. The natural  frequency at launch must be >50 Hz and this requires the TPS to have both stowed and deployed configurations. In the stowed configuration, the TPS is locked down to three launch towers with a 2-2-2 kinematic mount. The three attachment points are fastened with ground resettable non-pyro separation  nuts. The kinematic restraint uses a vee-spherolinder-cone design (Figure 8) to accommodate launch loads. The TPS is moved from the stowed to the deployed configuration by the Deployment Mechanism  (the subject of this paper). Figure 2 illustrates the phases that the TPS goes through after the separation  nuts are fired and it is deployed by the Deployment Mechanism. Figure 2. TPS Deployment Phases    Deployment Mechanism Design Overview The Deployment Mechanism incorporates a unique design that utilizes a variety of moving mechanical  components to drive the TPS from stowed to deployed and back again. A stepper gear motor drives the system that includes a ball screw, gears, bearings, flexures, welded bellows and position sensors. The  mechanism has a 75 /g1014range of motion between the stowed and deployed positions. Redundancy has  been included where it is feasible with redundant motor windings and positions sensors. Front and aft views of the Deployment Mechanism engineering model are shown in Figure 3.  437
Figure 3.  Front and Aft Views of th e Deployment Mechanism Engineering Model Deployment Mechanism Function The Deployment Mechanism is driven by a 45°, 2-phase stepper motor with an integral 3-stage, 384:1  planetary gear head and a detent brake. The output shaft of the gearmotor attaches to a 2:1 pinion-spur  gear reduction as shown in Figures 4 & 7. The spur gear is attached to a ball nut which is housed inside a titanium bracket and is supported by thin section duplex bearings at the top and a radial thin section bearing at the base. The ball nut rotates to advance a 6.35-mm (.25-in) pitch ball screw shaft with 25.4 cm (10 in) of total travel. The top end of the ball screw shaft is attached to a titanium yoke with  flexures. A stiff linear spring along the axis of the ball screw shaft at the interface joint with the yoke provides an over travel gage during the stowing operation. The yoke is attached to the TPS mast and pivots on radial bearings at each fork end. The TPS mast is supported by 3 rdaxis bearings at its base.  The 3rdaxis bearings include 80-mm (3.15-in) bore duplex back-to-back pair angular contact bearings on  one side and an 80-mm bore single radial bearing on the other side. The single radial bearing is preloaded with a beryllium copper flexure.  When the stepper motor drives the ball nut, the mast is driven between the stowed position at 15 /g892and the  deployed position at 90 /g892. Optical end of travel sensors indicate when the mechanism is in either the  stowed or deployed positions. A stop tower with a stiff linear spring that is compressed when the unit is in  the deployed position is used to preload the system. Detent magnets integral to the stepper motor are  used to prevent backdriving the system when it is deployed. To prevent the ball screw from over traveling,  a non-jamming hard stop engages slightly beyond the nominal deployed position. Figure 4 shows a  section view of the Deployment Mechanism CAD model. 438