ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
375 Earth Shield Deployment Mechanism (ESDM)  As TIRS is a cryogenic instrument, it must reject a large amount of heat to keep the focal plane array at  the operational temperature of 43K. In order to keep the radiator to a reasonable size, a 2.5-m2  deployable earth shield panel is used to block albedo. This earth shield is stowed at launch, and is rotated & locked into position by the ESDM. The ESDM is shown in Figure 3. 

Figure 3. The ESDM deploys the large Earth Shield panel from the stowed position flush with the  radiators 90 degrees to the deployed position. The left figure shows the panel stowed. The right  figure shows the panel fully deployed.  Cryo-Cooler Launch Lock (CCLL)  TIRS is actively cooled by a Stirling cycle cryo-cooler mounted beneath the instrument. This cooler is a source of jitter and is separated from the instrument by a passive vibration isolation system consisting of damping flexures. The flexures are too soft to survive the launch environment without a launch lock. The CCLL mechanism constrains the cryo-cooler supporting structure during launch and is released on orbit prior to instrument operations.   Scene Select Mechanism  The SSM is a single axis, precision mirror positioning mechanism, capable of 3 µradian stability. It can be driven in either direction for unlimited rotations. The rotating mirror is dynamically balanced over the spin axis, and does not require launch locking. Several configurations were traded before the SSM flight architecture was finalized 2. The mechanism is shown in Figure 5. The SSM met or exceeded the driving  requirements defined in Table 1.
376   Figure 4. The CCLL is located within the TIRS instrument. The instrument side panels are removed  to show the launch lock and the cryo-cooler keel structure that it restrains.  Table 1. SSM Driving Requirements  Requirement Value  Mass 33 lbm (15 kg)  Power 6 W average  Mirror Size 8.15 x 5.87 in (207 x 149 mm) elliptical  Knowledge ± 9.7 µradians over 34 minutes  Stability ± 9.7 µradians over 2.5 seconds  Duty Cycle 100%  Thermal Operational 0 / +20°C stable to ±1°C  Thermal Survival -50° / +40°C  Lifetime 3.25 years on orbit  Redundancy A/B side block redundancy  Operational Cadence Stare nadir for 30-40 minutes  Rotate 120°in <2 minutes to space view  Stare for ~30 seconds,   Rotate 120°in <2 minutes to blackbody view  Stare for ~30 seconds  Rotate to 120° in <2 minutes to nadir view 

Structure  The SSM will operate at a nominal 7-10°C; the scene mirror it rotates will be near 0°C. To minimize  heater power, low thermal conductivity titanium 6Al4V was chosen for the primary structural material. The 8.15 x 5.87 in (207 x 149 mm) scene mirror is made of an aluminum 6061-T6511 extrusion with a gold optical surface coating; the back is bare. The mirror is secured to a titanium mount with three aluminum 7075-T6 flexures to minimize thermal deflection of the mirror surface. The flexures are pinned to the mount then float bonded in place with Stycast 2850 epoxy to the mirror to reduce assembly-induced moment loading of the mirror. The mirror mount is bolted and liquid pinned to the SSM rotor shaft. The shaft is suspended in the bearing housing with a pair of angular contact ball bearings. The bearing pair is mounted within the bearing housing, which serves as the main housing of the SSM. Three radial legs extend from the housing and provide the mounting interfaces to the instrument. The encoder code disk
377 and motor rotor are bolted and liquid pinned to the shaft. The motor stator mount houses the stator, and  attaches to the non-mirror end of the SSM. The motor stator mount is a separate part from the bearing housing so that it can be shipped to the motor vendor for stator build up independent of the rest of the SSM. Similarly, the mirror mount is a separate part from the shaft to allow concurrent, independent assembly. The motor mount, mirror mount, and rotor shaft are finished in Tiodize type II for corrosion protection. The bearing housing was not Tiodized; it was left bare since it is almost entirely covered with strip heaters and other thermal control components which require non-anodized surfaces to minimize the generation of thermal hot spots.  

Figure 5. Internal detail of the SSM is shown in this quarter section view.  Several features were introduced to facilitate handling, reduce risk, and ease testing. Lift points and shipping interface points were independent of the flight interface points. This protected the precision flight interface from possible damage and wear from the numerous shipping and handling operations prior to installation in the instrument. A quartz optical cube was bonded to the motor housing to provide a boresight reference, facilitating SSM alignment and shimming into the TIRS instrument. Mirror pointing and stability measurements were referenced to this fiducial.
378 Baffles  A fixed aluminum baffle is fitted over the mirror end of the SSM, attached to the bearing housing. An  aluminum rotating baffle, attached to the shaft, protects the back of the scene mirror and the mirror mount from the cold space view and hot blackbody view thermal loads, and reduces stray light introduced into the telescope.   Actuator  The SSM is driven by a frameless, brushless, direct-current “zero cog” motor with redundant windings and damping coils. The 2-phase, 48-pole, moving magnet motor has an air gap of 0.4 mm, and a motor constant of 1.79 N-m/amp (253 in-oz/amp).   Bearings  The SSM uses a duplex pair of back-to-back mounted, 70-mm bore angular contact ball bearings with a 25-degree contact angle. The bearings are bonded in place to ensure stability at the µradian level. They are hard preloaded to 50 lbf (222 N), with the preload calculated to drop to 25 lbf (111 N) at the operational cold temperature. The bearings use tri-cresyl phosphate coated 440C balls in a phenolic cage, with 440C races. Pennzane lubricant is utilized; specifically an oil + grease slurry (50% each by weight) of Nye 2001 ultra-filtered synthetic oil and Nye Rheolube 2000 grease. The oil is also vacuum-impregnated into the retainer. Lubricant retention is provided by labyrinth seals sized for < 5% mass loss over the mechanism lifetime. Lubricant surface migration is prevented by the application of Nye-bar surface barrier coatings within the labyrinth seals. Since twice the estimated bearing lifetime (including margin) was only 100,000 cycles, a full flight fidelity life test was not performed. Instead, a life test using only a pair of bearing cartridges, with identical preload and thermal conditions, was executed successfully. 

Figure 6. Optical encoder is s hown in this section view mount ed on the SSM with the motor  removed.  Encoder  The SSM uses a 22-bit, pseudo-absolute optical encoder, which is an incremental unit that emulates an absolute encoder with a software counter and index pulses. The encoder has redundant read heads and drive electronics cards shown in Figure 7. The 125-mm code disk is fixed with liquid pinning to the rotating shaft. The read heads are fixed to the bearing housing. The encoder is shown in Figure 6. The
379 motor stator housing fits over the top of the read heads, encapsulating them; an integral contamination  shield protects the encoder from any potential particulates coming from the motor area. The encoder drive electronics are housed in a dedicated enclosure, which is mounted to the cover of the motor stator housing. The encoder design has some flight heritage as it is based on the pseudo-absolute architecture from the National Polar-orbiting Operational Environmental Satellite System (NPOESS) Preparatory Project (NPP) satellite Cross-track Infrared Sounder (CrIS) instrument. 

Figure 7. The optical encoder is shown mounted on the SSM with the motor removed.     Scene Select Mechanism Development Issues   Conceptually, the SSM is a very straightforward mechanism:  it is single-axis, direct-drive, without a launch lock. However, a challenging stability requirement, a complex thermal gradient, a custom digital servo-controller, and most of all a compressed schedule complicated the development. Many obstacles were encountered and overcome; the most significant technical issues are described below. Schedule and management issues also occurred but have been described previously 1.   The importance of functional breadboards  At the beginning of the development effort, the control engineers required hardware that could be used to develop the arc-second-level servo controller be available much earlier than we could expect to receive the actual motors and encoders. The hardware would need to have at least a 22-bit encoder and a similar motor to be useful to the development.  A commercial 24-bit absolute encoder was procured from the same vendor that was to deliver the flight unit. These units were available on a short delivery. The digital output from this unit was identical to that which would be delivered for flight; i.e., from a telemetry standpoint, it was electrically identical. A spare 2-phase, frameless brushless DC motor that was similar in performance to the flight motor was fortuitously available. A commercially available back-to-back duplex (DB) bearing pair equivalent to the flight design was quickly procured.
380 A simple aluminum bearing housing and spindle were fabricated, and the bearings installed with a spring  preload. The frameless motor was installed at one end of the shaft, and the encoder code disk and read head installed and aligned at the other end. While this configuration did not exactly mimic the SSM (which has the motor and encoder on the same end of the shaft), this was adequate for initial controls development.  

Figure 8. The SSM functional breadboard used a leftover frameless motor from another program  mounted with commercial bearings and an industrial 24-bit encoder to demonstrate positional  stability that exceeded the SSM requirements in  a laboratory environment. Stability measurements  were within 2.5 arcseconds with less than an arcsecond of noise.  This breadboard unit was available for use over 6 months before the encoders were delivered. This allowed the control electronics team to build a set of functional breadboard electronics much faster than if they would have waited for the flight hardware. An optical cube was bonded to the shaft, and an autocollimator was used to measure the stability and repeatability of the shaft position. This allowed 24-bit positioning stability and knowledge to be demonstrated in the laboratory. This was an important technical milestone which gave program management confidence that the SSM controller could be delivered on schedule. The unit, and some sample performance data is shown in Figure 8. 

Lesson Learned: Create a functional breadboard using components that can be procured quickly as early in the program as possible. These components do not need to be expensive. It will facilitate breadboard-level controls development, demonstrate performance in a laboratory setting, and potentially identify technical issues early. 

Critical assembly procedures should be practiced to reduce risk  In order to meet the arc-second-level stability requirement, critical bearing shaft and bearing housing surfaces had to be ground to tight tolerances and the bearings epoxied in place. Then the critical encoder code disk mounting surface on the shaft and the read head mounting surfaces on the bearing housing are ground true to the as-assembled spindle axis. These operations occurred at the encoder vendor which
384 to build up to a visibly detectable level during the instrument test. It should be noted that the film did not  degrade the science performance of the instrument or the performance of the SSM during the thermal vaccum testing. The film collected on the scene mirror and the first surface of the cold telescope lens; the two coldest surfaces closest to, and below the bearings. There was no detectable Pennzane contamination above the bearings near the warmer encoder code disk and encoder read head optics.The contamination was left in place as it did not affect the instrument and was not worth the cleaning risk. 

Figure 9. After insturment level thermal vacuum , a smudge was detected on the scene mirror and  a haze of pennzane was discovered on the mirror surface. The figure shows the surface of the  scene mirror as viewed from the exterior of the instrument.  Lesson Learned: While the contamination did not affect the instrument or mechanisms performance, a bake out would have prevented it. If the instrument operated in different wavelegths or a different lubricant was used, science performance could have been adversely affected. The bake out should have been done after spindle assembly but before installation of the optical encoder and scene mirror to prevent contamination of the main optical surface, code disk or read head optics. It was again confimed that contamination builds up on the coldest surfaces close to the outgassing points, and even the best labyrinth seal, properly coated with a barrier film, will not prevent the escape of outgassed lubricant. Note that even with a bake out, lubricant outgassing will never be completely eliminated, only significantly reduced.  Lesson Learned: Visible and high-resolution photographic inspection of critical mechanisms optical surfaces must occur before, during (if possible) and after critical development events if possible. Proper photographic documentation did occur at the instrument level but at the mechanisms component level was of inadequate resolution to be useful for surface contamination detection.
385 Earth Shield Deployment Mechanism  The ESDM consists of a pair of spring-driven hinges and a release/restraint mechanism which allow a  large earth shield panel to rotate 90 degrees from the stowed to deployed position. The large, 2.5-m2  earth shield is a composite panel with carbon fiber facesheets and aluminum honeycomb core; it is shown in Figure 10. The overall architecture has all active elements housed on the instrument side of the hinge line; no harness is passed through the hinges. Triangular mylar “wing” sheets, used to close out the ends of the earth shield panel, were folded between the earth shield panel and the radiator panel. These wings were pulled along and unfolded as the earth shield rotated to its full deployed position. The major requirement was the instrument/spacecraft uncompensated momentum requirement; no more than  158 N-m of uncompensated torque and 25 Nms of momentum could be generated by this deployment. The driving requirements for the ESDM are summarized in Table 2.  Table 2. ESDM Driving Requirements  Requirement Value  Earth Shield Dimensions 78 x 50 x 0.625 inch (1.98m x 1.27m x 15.8 mm)  Earth Shield Mass 24.9 lbm (11.3 kg)  Deployment Angle 90 degrees  Deployment Time ≤ 60 seconds  Uncompensated Momentum ≤ 25 Nms  Uncompensated Torque ≤ 158 Nm 

Figure 10. The Earth Shield shown in the stowed configuration during TIRS instrument integration.
386 Damped Actuator  The two ESDM latching hinges were based upon the deployable solar array panel hinges used on the  Lunar Reconnaissance Orbiter (LRO) spacecraft. They were powered by stowed energy, using torsion springs. The torsion springs (one in each hinge on the hinge line) were sized to provide positive deployment torque margin even if one of them failed. Each spring hinge contained a locking pawl which engaged when the deployment angle reached 90 degrees. This prevented any type of backdriving, and positively locked the panel in the deployed position. In order to accommodate thermal expansion and contraction of the panel, one of the hinges had axially floating bearings that were compliant to small linear deflections along the hinge line axis.  

Figure 11. The Earth Shield hinge system. Two spring-powered hinges drive the deployment. One  hinge mounts a rotary damper to significantly reduce the angular velocity while maintaining high  torque margin. The other hinge  mounts a potentiometer to pr ovide position feedback. The  strongback beam allows the hinges to maintain relative alignm ent when the Earth Shield   is removed.  The torque generated by the drive springs had to meet the torque margin requirement specified in the GSFC Gold Rules; it also had to meet the one failed spring success criteria. This resulted in the springs being strong enough to deploy the earth shield with enough velocity to exceed the uncompensated momentum requirement. In order to not exceed this requirement, a viscous fluid damper was used to reduce the deployment velocity, without significant reduction of the deployment torque margin. The heritage LRO viscous damper was utilized.  Due to the large size of the earth shield panel, it is removed from the instrument before shipment to the off-site spacecraft integration facility. The panel would be re-attached after the instrument was integrated to the spacecraft. To ease detach/re-attachment hinge alignment issues, and to facilitate testing, an intermediate stiff beam structure, or “strongback”, was introduced between the hinges and the earth shield panel. The hinges attached the strongback to the instrument, and the earth shield rode on the strongback. The earth shield could be easily removed from and reattached to the strongback without disturbing the hinge alignment. This configuration allowed the ESDM to be tested without disturbing the hinge alignment, and without the flight earth shield (using an equivalent inertia simulator).
387 Sensor  A potentiometer provided coarse position. The same model of potentiometer that was successfully flown  on the LRO solar arrays was used. Release/Restraint Device  A single-point, command-releasable restraint device was used to prevent the spring-powered hinges from deploying the earth shield. Two compliant snubbers were used to share launch loads with the hinges and the release restraint device.   The device used was an Ejector Release Mechanism (ERM). The model flown is rated for 1000 lbf  (4.4 kN) of holding force. It utilizes a shape-memory alloy which, when heated, releases a threaded cap which is held in place with a ball lockup type of lock. It releases with very low shock. It is manually resettable, and no hardware change out is required between testing and flight; the same hardware that is tested is flown. The ERM-1000 is shown in Figure 12. 

Figure 12. The ERM-1000 release restraint device shown on the left with the release cap captive,  locked to the ERM. The figure on the right shows the ERM during component thermal testing with  the release cap ejected after actuation.  The stock version of the ERM was modified slightly to work with the TIRS structure and electronics. The mounting flange, usually at the base of the device on the opposite end from the released cap, was moved to the same end as the released cap. This mechanical modification had heritage from other programs and was deemed low risk. Electrically, a 10-ohm in-line resistor was added to the heater circuit to better match our source current.  The ERM device is mounted to the instrument structure, and protrudes through the instrument radiator. The releasable cap attaches to the earth shield though a spring-loaded, bolt-catcher type compliant retraction interface, called the Coupler Retraction Mechanism, shown in Figure 13. This ensures that the cap clears the ERM completely upon release, and that the cap does not protrude beyond the interior surface of the earth shield, as this would slightly degrade the radiative performance of the shield. The mechanical contact between the structure and the earth shield at the lock down/release point is a matched radius cup and ball. This interface constrains all translations but is compliant to all rotations. To avoid the possibility of cold welding at this interface, dissimilar materials were used. The radiator side ball is aluminum 6061T6 with a type III hard anodized finish. The cone is titanium 6Al4V with a Tiodize® type II (Teflon® impregnated) finish. The ball and cup both have a large clearance hole through them that the ERM releasable cap and bolt catcher extend through. The cup and cone effectively isolate the ERM from earth shield moment and shear loads at the release point, allowing it to be loaded in pure tension.
388   Figure 13. A section view through the ERM and C oupler Retraction Mechanism. The inset figure  on the upper right shows the deploy ed configuration: The compr ession spring is fully expanded  with the ERM release cap pulled into the catch housing.  A kick-off load was introduced into the stowed earth shield by bending it slightly. This was done by  increasing the deflection at the restraint point, effectively pulling in at the center of the panel, reacting it against the snubbers and hinges. Due to this preloading, the panel was “bowed-in” when stowed. This bowing provided a kick-off load when the panel sprang back upon release, and obviated the need for dedicated kick-off springs.  Performance Testing Results  The ESDM underwent several deployment tests, with the velocity through most of the range varying between 2 and 5 degrees per second. The ESDM was deployed under ambient conditions, and in vacuum at room temperature and at -5°C. The thermal vacuum test is shown in Figure 14. Following acoustic and vibration tests, it was deployed again. All tests were successful and the results are shown in Figure 15. As is expected, the deployment time increased at cold temperature as the damper viscosity
389 increased. The analytical predictions were too fast by about a factor of 2 due to incorrect assumptions  about the drag caused by the unfolding wings, which turned out to be much higher than predicted.  

Figure 14. The Earth Shield and ESDM shown deploying within the thermal vacuum chamber after  the successful cold vacuum deployment test. The drag torque from the wing close out sheets  proved to be variable and made estimation diffi cult. Note the reflection of the chamber door and  photographer's camera flash on the inner surface of th e deployed earth shield  in the right figure. 

Figure 15. ESDM deployment analysis and test data. The ambient analytical prediction is the blue  dots on the left, the ambient test data is the green, blue, and red traces, and the purple trace is the  cold vacuum deployment data. Some of the non-linear behavior of the damper is shown in the  purple (cold vacuum) test trace near t = 3 seconds. Also, the variable drag from the wings  unfolding is shown on the blue (post evironmental test) trace near t = 8 seconds.  0102030405060708090 0 1 02 03 04 05 06 0Deployment Angle, Degrees Time, SecondsESDM Deployment Angle vs. Time Analysis Ambient T est Ambient Vacuum T est Cold Vacuum Test Ambient Post Vibration &  Acoustics T est