ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
126 Test Results 

Friction Tests   Figure 8 shows the WAGM set up for friction tests. Calibrated weights totaling 360.26 N (80.99 lb) were  suspended from the load cable. The pressure in t he AGM cylinder was adjusted to “float” the load such  that the vertically moving carriage of the AGM was stationary and within its working stroke. The air  

Figure 8: WAGM supporting 360.26-N (80.99-lb) payload 

pressure was then increased very slowly and the pressure P u (psig) was noted at which the load just  began to move upward. The pressure was then reduced very slowly and the pressure P d (psig) was noted  at which it just begun to move downward. The friction force referred to the payload end of the load cable was then calculated according to: 

N 2 / ) P P ( A Ld u p f−=            (1)  where:   L f  = friction force in lb    Ap = piston area in in2 (6.157 in2 (39.72 cm2)    N  = displacement multiplier ratio of pulley system (4.0:1)  The piston pressure in psig was read from the di gital panel meter of the WAGM control console (Figure  9). The pressure sensing channel had previously been calibrated against a primary pressure standard.  Motion was detected via the digital readout from the carriage position sensor. A displacement change of at least 0.0508 mm (0.002 in) (twice the least-count of the meter) was taken to indicate that friction had been overcome and the carriage had moved.    NASA/CP-2010-216272
127  

Figure 9: Piston pressure in psig (lower left) and carriage position from center in inches (upper  left) on the control console. Tip-tilt sensor ou tputs in milliradians are displayed on the other  two readouts. 

The test was repeated numerous times at payloads of 360.26 N (80.99 lb) and 249.06 N (55.99 lb).  Results are shown in Table 1. Each cell of the table contains two numbers. The upper number is a  pressure, either P u or P d. The lower number is the carriage position at which the pressure was recorded,  in inches from the vertical center of travel. 

Table 1: Friction test results 

   Pu k Pa (psig)Pd kPa (psig)Friction N (Ibf )Pu  kPa (psig)Pd kPa (psig)Friction N (Ibf ) Carriage Z cm  (in)Carriage Z cm  (in)Carriage Z cm  (in)Carriage Z cm  (in) 391.23 (56.743) 390.48 (56.635) 0.369 (0.083) 278.71 (40.424) 279.48 (40.535) -0.378 (-0.085) 18.331 (7.217) 18.326 (7.215) 0.813 (0.320) 0.826 (0.325) 390.55 (56.644) 389.31 (56.464) 0.618 (0.139) 278.62 (40.410) 279.49 (40.536) -0.431 (-0.097) -10.419 (-4.102) 0.815 (0.321) 0.823 (0.324) 390.55 (56.644) 391.01 (56.711) -0.231 (-0.052) 278.62 (40.410) 279.50 (40.538) -0.436 (-0.098) -10.419 (-4.102) -10.414 (-4.1) 0.815 (0.321) 0.833 (0.328) 390.43 (56.772) 390.48 (56.634) 0.472 (0.106) -10.411 (-4.099) -10.416 (-4.101) 391.55 (56.789) 390.48 (56.634) 0.529 (0.119) -10.409 (-4.098) -10.416 (-4.101) 391.52 (56.785) 389.94 (56.556) 0.783 (0.176) -10.409 (-4.098) -10.416 (-4.101)Payload = 360.26 N (80.99 lb) Bearing pressure = 210 kPa (30.5 psig)Payload = 360.26 N (80.99 lb)Bearing pressure = 210 kPa (30.5 psig) NASA/CP-2010-216272
128 As usual in friction testing, there is considerable scatter in the results. However, a typical value for vertical  friction referred to the payload end of the load cable was 0.445 N (0.1 lb). A lower payload value resulted  in less scatter but did not necessarily reduce the fric tion force. This is probably because a certain portion  of the friction comes from cable flexing rather than bearing friction in the pulleys.   In addition to friction, the uplift force accuracy is lim ited by the resolution of the ultra-precision regulator.  By test it was found that, with practice, the pressu re could be changed in increments as small as 69 Pa  (0.010 psi). This resulted in a force resolution of 69 Pa (0.01 psi) x A p / 4 = 0.67 N (0.015 lb) due to the  pressure regulator. The uplift force accuracy is the combination of the friction and the force resolution due  to the pressure regulator, or + 0.53 N (+ 0.12 lb).   Friction tests were also conducted on the AGM alone, without the 4x displacement multiplier. Figure 10 shows the AGM configured for this test. The 4x mult iplier mechanism is removed, as is the lower plate  that normally supports two pulleys of the multiplier. A load of 711.7 N (160 lb) is suspended directly from the lower cross-member of the AGM carriage. The piston pressure is then adjusted to an equilibrium value just sufficient to “float” the load. Piston pres sure and carriage position are monitored via the digital  panel meters of the control console (Figure 9). The test method then proceeds in the same way as the aforementioned friction test with the displacement  amplifier. It was found that, even with the ultraprecision pressure regulator, the piston pressure could not be adjusted accurately enough to maintain a perfect equilibrium. The carriage would always rise or fall very slowly. This is typical behavior for an AGM.  It indicates that friction, if there is any at all, is  less than the product of the piston area and the pressure  regulator resolution. For a pressure resolution of 35-69 Pa (0.005-0.010 psi) (typical for this type of ultraprecision regulator), the friction force is less than 138  mN (0.031 lb) at the offloader carriage or less than  36 mN (0.008 lb) at the payload end of the load cable when using the 4x multiplier. For practical purposes, the friction referred to the payload end of the load cable due to the AGM itself can be taken as zero.  

Figure 10: Friction test of AGM alone, without 4x displacement multiplier      NASA/CP-2010-216272
129 Proof Load Testing   Figure 11 shows the WAGM set up for proof load test ing. The goal was to demonstrate that the WAGM  could support a load of 711.7 N (160 lb) at the pay load end of the load cable without sustaining visible  damage and retain all functionality. There was no requirement for the AGM to float the proof load nor was  there a requirement for the X-Y stages to be moving while supporting the proof load.    For the test, the XY stages were both positioned at their outermost locations. That is, the Y stage was positioned as far as possible from the tower and the X stage was positioned at either travel limit. The  AGM was depressurized and the supply pressure to the air bearings was set to its normal value of 206.842 x 10 3 Pa (30 psig). The tower was extended high eno ugh so that a weight pan could be hung  from the payload end of the load cable without the pan touching the floor. At least 711.7 N (160 lb) of weights were placed on the weight pan and the entire system was examined for signs of distress. The  weights were left in place for a minimum of five minutes and then removed.   Figure 11 shows the WAGM with 716.1 N (160.99 lb) of certified weights hung from the payload end of the load cable. The XY stage was at one outboard “corner” of its range of  motion, per the test  requirement. No damage was visible or otherwise a pparent during or following the proof load test, nor  was any effect on performance detected during the remaining tests.  

Figure 11: WAGM undergoing proof load test  

X-Y Tracking Accuracy   The customer requested that the offload force resoluti on be less than 89 mN (0.02 lb) in all three spatial  directions. Friction tests and review of the pressure regulator resolution showed that 534 mN (0.12 lb) of  vertical force resolution was possible. In the horizon tal degrees of freedom this requirement is a function  of the payload weight and the angular tracking error when the payload is moving at constant velocity. For  a maximum payload weight of 355.8 N (80 lb) and a maximum payload speed of 1.59 mm/s (0.0625 in/s)  NASA/CP-2010-216272
130 this translated into less than 0.25 milliradian allowable steady state angular deviat ion of the cable guide  tube, an aggressive goal to be sure.   CSA conducted a so-called ‘box test’ to determine the steady state angular trac king error at the maximum  payload speed. A box test exercises the robotic arm th rough its entire range of motion in all three spatial  directions. Prior to start, a dummy weight was hung from the load cable and the tip tilt sensors were  zeroed. The WAGM custom software generates a log of  pressure, position, and angle at a 2-Hz sampling  frequency. This sampling frequency is sufficient because as mentioned previously the tip-tilt sensor signals are fed through anti-alias f ilters before digitization by the cont roller card, whereupon the signals  are then further digitally low-pass filtered onboard the controller card at a cutoff frequency of 1 Hz. The data logged over the entire box test trajectory were plotted against time in order to determine the steady state x-y tracking accuracy as a function of robot arm speed.   The upper panel of Figure 12 shows both the raw sig nals of X-Z angle and Y-Z angle taken directly from  the WAGM software log, as well as the signals post-p rocessed in MATLAB with zero phase shifting digital  filtering at 1 Hz for better visualization of the mean  underlying angular errors. The lower panel of Figure  12 shows the trajectory of the payload in all three coordinate directions. The Z-trajectory in the lower panel reflects the length of the load cable from the un iversal joint to the pick point on the payload. Figure  12 shows that the steady state angular tracking error is a function of t he payload speed. This is because  the feedback control loop utilizes only proportional control (see Figure 7). Integral control was not utilized because it was found that the integr al gain required to have any significant effect within the deployment  time made the system unstable.   

Figure 12: Angular tracking error vs. payload speed during a box test 0 500 1000 1500 2000 2500 3000 3500 4000 4500-40-30-20-10010203040milliradiansDSM Box Test Sep. 3, 2009: Entire Trajectory Ki = 0, Kp = 5.3, Max DSM Speed = 0.15 in/s Yellow Stanchions + Roll Poles    X-Z Accel (mrad) Y-Z Accel (mrad) FiltFilt X-Z Accel (mrad) FiltFilt Y-Z Accel (mrad) 0 500 1000 1500 2000 2500 3000 3500 4000 45002030405060708090100110120 Time (secs)Distance (inches)    X (inches) Y (inches) Z (inches)Max S.S. Tracking Error = 10 mrad Max DSM Speed = 0.15 in/sDSM Speed = 0.08 in/sS.S. Tracking Error = 5.3 mradNo observable hunting NASA/CP-2010-216272
131 Figure 12 shows that at a payload speed of 3.81 mm/s (0.15 in /s) the steady state tracking error  approached 10 milliradians. Close to the contractual maximum payload speed of 1.6 mm/s (0.0625 in/s)  the steady state tracking error was around 5 millir adians. For a maximum payload weight of 355.8 N (80  lb), this corresponds to a lateral drag force of just 1.779 N (0.4 lb).   Lessons Learned   Free play in the stages of the COTS telescoping tower presented some significant challenges during  development. Free play around the vertical axis of  the telescoping tower produced a yaw mode that  interacted unfavorably with the horizontal active tracking system. Free play in t he roll direction gave rise  to a sudden shift of the load platform as the AGM traversed over center in  the X-direction.     To mitigate the control engagement of the yaw mode two anchor stanchions and a set of guy cables were added to stiffen the tower in yaw. To eliminate load platform shifting two telescoping poles were installed beneath the tines of the load platform to stiffen th e tower in roll. The customer deemed the set-up  repeatability of these stiffening elements insuffici ent however, and so another stiffening solution was  devised that lent itself to gr eater repeatability of setup.   

Figure 13: Pair of triangulated roll poles to stiffen telescoping tower of industrial lift   The solution implemented was a pair of telescoping t ubes that triangulate the structure (Figure 13). The  telescoping tubing retracts and extends automatically wi th the raising and lowering of the load platform.  NASA/CP-2010-216272
132 The two sections of each tube are clamped together onc e the load platform is at the desired height. The  setup is straightforward and repeatable with the use of a torque wrench to pre-load the triangulating  members.   Refashioning COTS equipment for use in an application fo r which it was not originally intended should be  approached with caution. Such use often depends on characteristics of the COTS equipment that are not usually important (free play of the tower in yaw and roll) and may therefore not be well controlled in design and manufacture. This is the risk that comes with the cost savings motivating the use of COTS  equipment. In hindsight, the problems with the towe r could have been solved at the design stage by  simply using two towers, one lifting at each rear corn er of the load platform. More attention to tower  stiffness could potentially allow t he X-Y control system to be stable at higher gains and possibly with  integral gain, thus giving higher tracking accuracy.   Conclusions  

WAGM is the first proven three-DOF gravity offloadi ng system of its kind to deliver high fidelity zerogravity simulation over large displacements with very low set-up time. Its very simple interface to the  payload suggests that it could be used for zero-g simu lation with a wide variety of deployable structures  and other uses as well. Potential applications of WAGM include:  • Ground testing of deployable spacecraft structures  • Pick and placement of heavy, sensitive equipm ent (e.g., during spacecraft installation)  • Gravity offloading of human subjects for injury rehabi litation or to study the effects of zero-gravity  conditions on the musculoskeletal system 

With its improvements over existing technologies in zero-gravity simulation fi delity, set-up time, and  extended range of motion, it is easy to envision many  applications for WAGM at significant cost savings  to the user in terms of set-up time, te sting time, and accurate system validation. 

References 

[1] Kienholz, D., Crawley, E., Harvey, T., “V ery Low Frequency Suspension Systems for   Dynamic Testing,” Proc. 30th Structures, Structural Dynamics, and Materials Conference, Mobile, AL,  April 1989   [2] Kienholz, D., “A Suspension System for Si mulating Unconstrained Boundary Conditions,”   Proc. 12th Int’l Modal Analysis C onf., Honolulu, HI, February, 1994   [3] Tanimoto, R., Moore, A., MacDonald, D., Thom as, S., Murray, A., Polanco, O., Agnes, G.,   “Model and Test Validation of Gravity Offload System,” Proc. 48 th   AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Material s Conf., Honolulu, HI,  April, 2007    [4] Corey, R. L. and Pidgeon, D. J., “Electric Propulsion at Space Systems Loral,” IEPC-2009-  270, 31st International Electric Propulsion Conference, MI, 2009    NASA/CP-2010-216272
133 Gimbals Drive and Control Electronics D esign, Development and Testing of the  LRO High Gain Antenna and Solar Array Systems 

Boris Chernyakov* and Kamal Thakore** 

Abstract   Launched June 18, 2009 on an Atlas V rocket, NASA’s Lunar Reconnaissance Orbiter (LRO) is the first step in NASA’s Vision for Space Exploration program and for a human return to the Moon. The spacecraft  (SC) carries a wide variety of scient ific instruments and provides an ex traordinary opportunity to study the  lunar landscape at resolutions and over time scal es never achieved before. The spacecraft systems are  designed to enable achievement of LRO’s mission requi rements. To that end, LRO’s mechanical system  employed two two-axis gimbal assemblies used to driv e the deployment and articulation of the Solar  Array System (SAS) and the High Gain Antenna Sy stem (HGAS). This paper describes the design,  development, integration, and testing of Gimbal Cont rol Electronics (GCE) and Actuators for both the  HGAS and SAS systems, as well as flight testing during the on-orbit commi ssioning phase and lessons  learned.     Introduction 

In January 2004, the President of the United States  unveiled the Vision for Space Exploration which  charted the path for the return of humans to the Moon and deep space. The first milestone in that plan was an unmanned lunar orbiter to be launched in late 200 8. By late 2004, that first mission had become  the Lunar Reconnaiss ance Orbiter, and as 2005 began, development had st arted at NASA’s Goddard  Space Flight Center.  With a goal launch date of end of 2008, the mission had only four years to be developed from concept to  launch. In addition, the primary mission required that  the spacecraft remain in a nominal 50-km polar  mapping orbit for a minimum of one year, collecting da ta over the entire lunar surface under all possible  lighting conditions; imposing challeng ing constraints on SC design. Sched ule, cost, mission, and reliability  requirements manifested themselves in every aspect of the LRO design.     To meet the needs of a discover y-class mission with an accele rated development schedule, the  development, fabrication and testing of the LRO GCE and Actuator su bsystems was inevitably fastpaced. Schedule needs of the overall mission require d that the GCE/Actuator subsystems provide  reliable, environmentally tested flight hardware within 15 months of specification release; just in time for  SC integration and test (I&T) activities. Figure 1 sh ows the layout of the integrated spacecraft, with fully  deployed HGAS and SAS.  This paper will discuss the overall design, developm ent, integration, testing (ground and flight), and  lessons learned for the GCE and Actuator subsystems of both the HGAS and SAS.                                                           

* ATK Space Systems and Services, Beltsville, MD  ** NASA Goddard Space Flight Center, Greenbelt, MD 

Proceedings of the 40th Aerospace Mechanisms Symposium, NASA Kennedy Space Center, May 12-14, 2010  NASA/CP-2010-216272
134              Figure 1.  LRO Spacecraft with Deployed Solar Arrays and High Gain Antenna 

Design 

As described in the Introduction portion of this  paper, the Gimbal Control Electronics and Actuator  subsystems of the LRO are intended to enable the primary mission objectives, and were required to meet cost, reliability and schedule constraints imposed by t he class and short developm ent time of the overall  mission.   On LRO, the function of the GCE is to control and drive dual-axi s redundant actuat ors on the SAS and  HGAS. The Solar Array System GCE drive enables me chanical rotation of 90° on one actuator and 180°  on the other, allowing the Solar Arra ys to track the sun in two axes and provide a reliable power source  for the Orbiter. The HGAS GCE subsystem drives  180° rotation on both axes, enabling the high-gain  antenna to point toward Earth whenever it is in view (providing maximum time for data downlinking and ground system communications). Two redundant, dual  disk, incremental en coders – one on the input  (fine) and one on the output (coars e) – are integral to the actuat or assembly for both the HGAS and SAS  GCE subsystems. These encoders provide fine and coarse incremental position sensing with respect to a unique home/reference position located at the center of travel. They also provide a logical state change  for each physical motor cardinal step taken, thereby providing an output resolution equal to the cardinal  step size. A pictorial view of the LRO actuators and the internal lay out of the fine and coarse encoders  are shown in Figure 2. HGAS and SAS gimbal  systems can be seen in Figure 3. 

HIGH GAIN ANTENNA (40 W KaTx, 100 Mbps)INSTRUMENT MODULE  (6 instruments, 460 Gbits /day) SOLAR ARRAY (2000 W BOL, 80 AH Battery)PROPULSION MODULE (898 kg N2H4)AVIONICS PANEL (SpW /1553, 412 GbitsStorage)SPACECRAFT BUS (Modular Honeycomb Design) 2 mX ZY LRO Orbiter Characteristics Mass  (CBE) 1916 kgDry: 1018 kg, Fuel: 898 kg (1313 m/sec) Orbit Average Bus Power 647 W @ Beta 0 Data Volume, Max Downlink rate 461 Gb/day, 100Mb/sec Pointing Accuracy, Knowledge 60, 30 arc-sec NASA/CP-2010-216272
135    Figure 2.  LRO HGAS and SAS gimbals actuators 

                        SAS Gimbals             HGAS Gimbals  

Figure 3.  LRO S/C in the EMI chamber with a removed SA and fully deployed HGA  

The GCE/Actuator subsystems cons isted of two engineering development units (EDU), three flight boxes,  two commercial actuators, and five flight actuators. Each electronics box provides a fully redundant two-axis control or drive for the thr ee-phase harmonic drive actuators.    These gimbal assemblies, including the Gimbal Control Electronics (GCE) and harmonic drive actuators  used for the SC deployables were developed by a team comprised of engineers from Goddard Space Flight Center (GSFC), Alliant TechSystems (ATK),  Broad Reach Engineering (BRE) and Sierra Nevada  Corporation (SNC).  Actuator Selection

In order to meet reliability and schedule constraints, the project leveraged ava ilable technologies and inproduction hardware, and strove for overall simplici ty of design wherever pos sible. To ensure that  reliable, environmentally tested GCE/Ac tuator subsystems were delivered in time for SC I&T activities, the  project selected SNC Actuators, identical to those developed for t he Solar Dynamic Observatory (SDO)  spacecraft.  These actuators had been extensively tested duri ng SDO High Gain System integration, with  performance characteristics shown to meet the ba sic requirements for LRO’s mission. The actuator  interfaces were also well understood, allowing subs ystem and interfacing system designs to proceed  rapidly. In addition, these actuators provided the internal redundancy required by LRO’s mission.  NASA/CP-2010-216272
136 Although the HGAS and SAS inertias differ by a factor of fifty, the analysis s howed that the actuators  could easily achieve the required torque margins for both systems with optim ized motor drive control  electronics. Given that the SNC Ac tuators met the needs of both the HGAS and the SAS, and could take  advantage of the development and testing for SDO, t hese actuators were selected for use on LRO over  more optimized but more schedule intensive actuators.    Drive Electronics Selection/Design

Due to the wide range of operational constraints imposed by a lunar mission, such as a challenging thermal environment and the need for extensive ground testing in a 1g environment, a drive electronics  system with multiple set points was selected to provide flexibility.   In addition, it was found that the drive system nee ded to allow micro-stepping with multiple resolution  options for micro-step rates. Accurate tracking of th e sun on two axes required that the solar arrays be  driven at a rate of 7 pulses per seco nd (pps), and their large inertias added the risk that spacecraft jitter  may have caused violation of point ing budgets. To minimize the potent ial effects of SAS tracking on  spacecraft stability, the GC E assembly (Figure 4) was designed to provide the capability for microstepping at multiple resolutions.    

Figure 4.  LRO Gimbal Control Electronics Box      GCE Design

To take advantage of the inherent internal redundancy of the SNC actuators, and to improve the reliability  of two critical SC control subsystems, the GC E subsystem was designed as a 100% redundant  electronics box. The GCE subsystem is comprised of tw o identical controllers, two identical Motor Driver  Boards, two identical DC to DC Converter Boards, two backplanes and the chassis. The spacecraft  provides the subsystem a 31-Volt Direct Current (VDC) nominal voltage and Primary/Redundant Side Enabled pulsed discrete commanding to the subsystem. The GCE communicates with the SC via the MIL-STD-1553B bus. A block diagram of the GCE is shown below in Figure 5.  NASA/CP-2010-216272