ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
137   Figure 5.  LRO GCE block diagram 

 The GCE design maximizes internal mode control capability while minimizing command and control interfaces with the flight software. Major communica tions and control functions are accomplished by two  Field Programmable Gate Arrays (FPGAs). One of thes e is located on the Controller Card to provide SC  command and telemetry interfaces, processing of crit ical and housekeeping GCE telemetry, operational  positioning controls, and automated positioning controls. The other FPGA is located on the Motor Driver  card, and controls the stepper motors’ comm utation sequences, micro-stepping capabilities,  programmable power set points, and closed loop current control.   To allow minimal power dissipation in the gimbals, and to provide flexibility in selection of an optimum  drive, the GCE design utilized a constant current driv e system with pulse width modulated control. The  system offers eight current set points ranging between 200 and 390 mA. Operational step rates are  commandable in the range of 0 to 67 cardinal steps/ sec for flight operation and 0 to 150 steps/sec for  laboratory and ground testing. The system also en abled optional micro-stepping, allowing commanding of  micro-step resolutions from 0 (1 cardinal step) to 6 ( 64 cardinal steps), to minimi ze the effect of tracking  motions on spacecraft jitter. Internal electronics cond ition the actuator motor and output optical encoder  signals, providing a closed loop control with  resolution of 0.0075° per cardinal step.   The GCE uses five distinct command modes to execut e gimbal control; the GCE mode state diagram is  shown in Figure 6. In addition to the basic command ing of the gimbals, the GCE also monitors gimbal  execution and positional accuracy by way of operational error flags; the Position Error and Step Error flags. During the initialization mode, the GCE sets the “Home” position to 0, and the Position Error flag is  checked every time the “Home” marker is crossed to verify that the current position is at 0 counts. Any  deviation from 0 indicates positional accuracy violat ion. The Step Error flag is generated if encoder  feedback indicates that the actuator failed to complete a command for the motor to take a step. These operational error flags ensure that the gimbals c an be accurately commanded and controlled. The GCE  also provides high resolution temperature monitoring for multiple PRT (platinum resistance thermometer)/Thermistor sensors for various LRO SC hardware components. 

NASA/CP-2010-216272
138 ELSEELSE ELSEState STOPState TRACK State GOState INITIALIZEState MOVETRACK command STOP command or EOT GO command STOP command or GO doneINITIALIZEcommandSTOP command or Initialize doneSTOP command orEOT or MOVE done MOVE commandPower Up ELSEELSE

Figure 6.  GCE Mode State Diagram 

Technical Problems and Solutions   As with all spaceflight hardware, design of the GCE/Actuator subsystems presented many technical challenges. Not only did the systems need to meet fli ght requirements, they also needed to be able to be  efficiently and thoroughly tested in all operational modes.  Re-circulating Current and Motor Current Measurement Distortion

One such challenge was presented by the need to provide an accurate motor current measurement in a pulse modulated system with multiple actuator drive current settings, while oper ating from an unregulated  power bus.  Specifics of the pulse width modulation control produc e a re-circulating current (supplied by the actuator  coils during dead time), bypassing a sense resistor. Figure 7 shows one possible re-circulating current condition, with all motor driver switches set on HIGH.  Current Sense 0.5 OhmDiff Amp  x10Low  Pass Filter Low  PassFilterSet PointComparatorPhase A Phase B Phase C ErrorTo FPGAFrom  FPGA From  FPGA From  FPGA From  FPGACurrent Flow (I) Current Flow (I/2)Current Flow (I/2)    Phase A - High Phase B - HighPhase C - High

Figure 7.  GCE shunt current monitoring, all motor driver switches on HIGH 

When all the motor driver switches are set on HIGH, the current is supplied by the actuator coils, recirculating to the power supply and therefore bypa ssing the sense resistor. The current then decays  based on the time constant of the actuator coils (t = 100mH/69W = 1.4ms), reducing the shunt current  measurement, and distorting the motor current value.   NASA/CP-2010-216272
139 To resolve the problems related to the re-circulating current, the FPGA measures the dead-time using  knowledge of the pulse width modulated (PWM) duty cy cles. The FPGA calculates the measured current  error by computing the amount of me asured current distortion (% of time the current is measured) from  knowledge of the PWM duty cycles. The commanded set point is then continuously adjusted by multiplying the power set point by the shunt measurement  distortion factor, reported in the telemetry as a  “power scale factor”. Figure 8 shows a simplified model of the current control mode with “dead-time compensation”.    

Figure 8.  GCE Current Control Mode   Power Return Electromagnetic Interference/Radio Frequency Interference

Another challenge was presented by th e design of the gimbal drive electronics power returns. These drive  electronics required that the primary (spacecraft) po wer returns be connected to the secondary returns,  and therefore must be isolated from the chassis. To satisfy LRO’s EMI/RFI requirements, low noise PRT  conditioning circuits were utilized and many ot her design considerations were implemented.   When the GCE/Actuator subsystem EDU was initially EMI tested, analysis showed significant EMI violations and excessive noise. To resolve the violat ions and decrease the noise, many modifications to  the EDU were made. These included redesign of the power distribution to the motor drive electronics,  implementation of constant current source circuitry for temperature monitoring, and a partial re-layout of printed circuit cards. Final EMI and functional test ing verified that these design changes successfully  resolved all EMI and noise related issues.   Component and System Level Tests   Systems level testing of the HGAS and SAS was particu larly critical to mission success. In order to  conduct multiple, and often simultaneous, tests, many  combinations of flight hardware, commercial  equivalents, and Ground Support Equipment configurations were identified, with specific harnesses fabricated.  Although all GCE subsystem testing was done with a dedicated lab-view setup, all system and SC level GCE testing was conducted with Integrated Test and Operations System (ITOS) driven test racks. ITOS  would be used in flight for commanding and telemetry m onitoring, and provided a flight like ground control  environment for testing. Use of t he ITOS test racks for testing also  allowed the team to develop and  become familiarized with ground station telemetry and control pages, used throughout system and Orbiter  integration and testing, and in flight. Two sample  GCE/Gimbal telemetry pages are shown below in  NASA/CP-2010-216272
141 Specialized test configurations allowed various detailed characterizations of subsystem elements to be  developed. This was particularly useful in the case of the HGAS actuators. Although the SAS flight  software (FSW) could be ground tested with flight gimbals without solar array panels attached, the HGAS FSW and tracking and control algorithms fully verifi ed in a 1G environment and was mostly conducted  with the set of commercial (but otherwise flight like) ac tuators. Therefore it was critical to characterize the  flight units prior to integrating them into the gimbals assembly.    The characterization of the GCE/Actuator components prior to flight was especially helpful in the identification and isolation of the source of one spec ific anomaly encountered during flight operations.  Figures 11 shows the results of te st characterizations for HGAS flight  actuator encoder responses when  operating at a 67 pps step rate in  clockwise (CW) and counterclockwise (CCW) directions. The distinct  differences of the signal waveforms illustrated here will be a ddressed in the Flight Operations section of  this paper.             

Figure 11.  LRO HGAS Actuator Encoder responses at a 67 pps step rate (Left: CCW, Right: CW)    To conduct the necessary HGAS Range of Motion tests with LRO’s flight antenna, a special sequence of coordinated motions and tracking rates were developed. These special sequences were intended to enable testing of the entire end-of-travel to end-of-trave l range of motions for all quadrants of the two-axis  gimbals while avoiding “zippering” in the 1G environment. These special motion sequences and tracking rates were developed from calculations and analysis of torque capabilities that allowed determination of  safe start and end positions, while enabling the ground test to exercise slew and tracking rates. Specific  current set points for each motion were established to produce a combination of single and dual gimbals motions for the test that would minimize the time required to complete any individual test. These motions were run numerous times during th e integration and environmental te sting of the spacecraft. The HGAS  Range of Motion Profile is shown in Figure 12.  NASA/CP-2010-216272
142   Figure 12.  HGAS Range of Motion test profile 

Lessons Learned 

Although numerous potential improvements to GCE/Ac tuator requirements, design and operation were  realized during the many hours of HGAS, SAS and SC testing, integration and operation, only a few of the most significant are discussed below.    Lack of loadable “Position Register” 

One potential improvement to the sy stem would be to have a pre-loadabl e position register. This would  simplify laboratory development and ground level testi ng, as well as eliminate some confusion when  gimbals are deliberately positioned to certain angles prior to system shut down. With a pre-loadable position register, on power up the GCE could be loaded with the default value. If knowledge of absolute  position is required for follow-up tests, the register could be pre-loaded with the value corresponding to an  actual position, and any further movement, even without  system initialization, would reflect absolute  gimbal position. In a 1G environment, the initialization is sometimes an undesirable mode, and can be a time consuming operation. Use of pre-loaded position  register can be utilized to save time and avoid  undesirable operation modes.  An alternative solution may also be the addition of non-volatile memory, where specific system parameters can be maintained at all times.   Acceleration/Deceleration Profiles

One of the original requirements imposed on th e GCE by the LRO Design Specification was that  acceleration (Accel) and deceleration (Decel) ramping profiles be internally g enerated. At the time the  Design Specification was developed (early in LRO’s project life), ACS tracking requirements and controls  were not fully defined, making it difficult to establish requirements for the Accel/Decel profiles. As a result, Y-axis Gimbal XZ-axis Gimbal 90 45 90 4590 45 9045 Gimbal Home Command 2 Command  4 Command  1Command 3Co mmand 6Command  5 START Position HGA Pointed at Floor NASA/CP-2010-216272
145 Effects of Micro-Stepping on Spacecraft Jitter   Further testing was conducted to study  the effect of various micro-step resolutions on SC jitter. The plots  in Figure 13 show the motor winding currents at 2 mi cro-step and 32 micro-step resolutions, and indicate  why some improvement of SC jitter at hi gher micro-step resolutions was expected. 

Figure 13.  Actuator winding currents at 2 (left) and 32 (right) micro-steps per cardinal step modes 

In addition, a Gimbal Step Resolution Experiment was conducted on the LRO Solar Array Gimbals to determine the effects of the different step resolution s on spacecraft Jitter and Attitude Error. The test  entailed running the Solar Array gimbals in track mode, using three different step rates or resolutions: cardinal, 16 micro-step, and 32 mi cro-step. The experiment was conducted  on consecutive orbits in the  sequence indicated in Table 1.     Table 1. Timeline for SA Gimbal step resolution experiment  Test Timeline  Time Event  2009-183-21:42:00 SA Y Gimbal stopped, commanded to cardinal stepping  2009-184-01:07:44 SA Y Gimbal stopped, commanded to micro step  with resolution of 32  2009-184-03:00:00 SA Y Gimbal stopped, commanded to micro step  with resolution of 16   To analyze the data, MATLAB was used to generate  a Power Spectral Density (PSD) plot of the  Spacecraft Body Rates for each of the three test ca ses. In order to obtain more information about the  difference between the spacecraft’s response to various gimbal micro-step resolutions, the time history of the Attitude Error and IRU rates was calculated. The vari ance in the data is shown in Table 2 for SC body  rates and Table 3 for Attitude Error.   Table 2. Variance for SC Body Rates for full data set  Body Rates (asec/sec) X Variance Y Variance Z Variance  Cardinal 27.4473 110.3259 8.2311  Micro 16 20.9319 113.7145 8.0349  Micro 32 22.2692 84.5873 7.4360 

Table 3. Variance for SC Attitude Error for full data set  Attitude Error  (asec) X Variance Y Variance Z Variance  Cardinal 16.1147 21.4523 4.8761  Micro 16 4.4508 13.0828 2.1641  Micro 32 2.5229 8.2906 2.2742   After analysis of the results, it was determined use of the 32 micro-step resolution significantly improves attitude and jitter performance, in comparison to use of cardinal stepping.      NASA/CP-2010-216272
146 Conclusion 

On-time hardware delivery, successful SC integrat ion and test, and six months of successful on-orbit  testing and operation have demonstrated that the GCE/Actuator team, and the overall LRO Project team,  used a viable approach for the rapid design, development and integration of critical spacecraft subsystems.   Acknowledgements 

The authors would like to recognize many individuals in strumental to the successful design, development,  testing, and integration of the LRO GCE/Actuator subsystems. The Pala Manhas, Praful Petal, Chris  Hodge, and the entire BRE GCE team not only provided the development and timely delivery of the gimbal control electronics, but also all nece ssary support during SC operation. HGAS and SAS  mechanical engineers Adam Matuszeski (GSFC), Greg Martins (GSFC), and Mike Hersh (Sigma) were all critical to mission success. SNC’s Jeff Moser and GSFC’s Joe Schepis provided essential support in  resolving all actuator related issues. ACS analyst Ge rardo Cruz provided analysis and evaluation of onorbit jitter and attitude stability tests. LRO System s Engineering support from Dave Everett (GSFC),  Michael Pryzby (ATK), and Steve Andrews (GSFC) provided indispensable guidance and oversight during development and critical support during environmental tests and orbital simulations. There are numerous others to mention that are also appreciated for their efforts. 

References 

Technical Memo: LRO SA and HGA Systems Torque Margins Analysis. Technical Memo: Step Stability Margin  Assessment, J. Schepis, GSFC  Technical Memo: Analysis of results from Solar Arra y Gimbal Step Resolution Experiment, G. Cruz-Ortiz 

NASA/CP-2010-216272
148                       Figure 1 – STIS Fastener Capture Plate                        Figure 2 – ACS Fastener Capture Plate  

                     Figure 3 – Handrail Removal Tool                                  Figure 4 – Clamp Removal Tool  

General Tool Description   A typical solution to the loose fastener problem would be to disengage, remove, and discard each  fastener one-by-one. This would involve multiple tool s used in concert, for example, grabbing the fastener  with one tool while it is being disengaged with a seco nd tool and then disposing of it into a trash bag or  third tool. Stowing the released fastener is another complicated operation during which there is risk of  losing the fastener. The tools invented simplify t he problem by reducing the number of tools operated  simultaneously and by removing the risky operation of stowin g the fastener.  With the new tools, each tool is permanently attach ed to its worksite by the operator. This allows the  operator to have both hands free. Each tool has clear windows with holes that line up with the fasteners  that are to be removed. These holes allow the oper ator to insert a screwdriver-type bit through the  window to unscrew each fastener. Behind each window is a gap, which creates a pocket or cavity for the  fastener to fall in to after it is unscrewed. In some ca ses, a compliant gel material fills in the gap to grab  the head of the released fastener. The holes in the window are small enough that the fastener cannot  escape, but large enough for the bit to pass through. This concept is pictured with the Handrail Removal Tool in Figure 5.      NASA/CP-2010-216272
150 Unique Tool Features   Some unique features of the technology that give it advantages over prior fastener removal methods are  listed below. All the listed features  were intended and developed to minimize task time, as well as  maximize the ease and confidence of the astronaut operator.  1) The action of disengaging fasteners is not autom ated but rather allows the operator full control  and feedback.  2) The fasteners can be completely disregarded after they are disengaged. They do not need to be  grabbed, stowed or cleaned up.  3) The tool can be attached in place just once and does not need to be handled or moved around  the work site. It then can be used as the handling interface for the removed hardware.  4) The loosened fasteners are positively captured at all times and cannot come free without ultimate  structural failure.  5) The tool minimizes particulate debris, even if a fastener were to be drilled out.  6) The tool can prevent a loosened fastener from incidental thread re-engagement with its mating  part.  7) The tool can keep the loosened fastener positioned for easy re-engagement if necessary.     Tool Design 

Tool Design Overview

The STIS Fastener Capture Plate tool is shown with  labeled components in Figure 7. The other fastener  capture tools created for the HST SM4 mission have si milar features but differing geometries and capture  different amounts of fasteners. 

Figure 7 – STIS Fastener Capture Plate Overview  The STIS Fastener Capture Plate attaches to the front  side of the STIS instrument, shown in Figure 8.  The attachment mechanism is not detailed here, but the Fastener Capture Plate is essentially anchored  at its four corners to the STIS Front Panel. The Front Panel, detailed in Figure 9, needs to be removed to gain access to the failed electronics board.  Handhold ZipNuts®  4x Depth  Indicator  CG Plate Cutter Windows x72  2x Soft Dock  & Drive-Off  Mechanism  Knob  Tether  Loop  Electrical  Receptacle  NASA/CP-2010-216272
152 The layout of the 111 fasteners holding the Front Panel  on can be seen in Figure 10. There are three  types of fasteners, two having washers, t he details of which can be found in Table 1.  Figure 10 – STIS Front Panel Fastener Layout 

Table 1 – STIS Front Panel Fastener Details 

#4 Torq-Set  Fasteners x39  #8 SHC Screw  Fasteners x12 #4 SHC Screw  Fasteners x56  12 12 56 56 39 Qty  NAS 620C8 0.179 ID #8 Reduced Diameter Washer NAS 1352N08LL8 0.164-32, 1/2 inch long #8 Socket Head Cap Screw NAS 620C4 0.125 ID #4 Reduced Diameter Washer NAS 1352E04LE7 0.112-40, 7/16 inch long #4 Socket Head Cap Screw NAS 1189E04T6K 0.112-40, 3/8 inch long #4 Torq-Set Screw Specification / Part  NumberSize (inches) Fastener Type  NASA/CP-2010-216272
153 When attached in place on STIS, the default position of the Fastener Capture Plate aligns its clear  windows and holes with the Front Panel fastener locations. The colors on the tool (red, white and blue) help the operator differentiate between the different types of fasteners. Each fastener also has an alphanumeric label. This allows an operations monitor to take note of what fasteners have and have not yet  been removed, so as to make sure none get omitted. These colors and labels can be seen in Figure 11. 

Figure 11 – STIS Fastener Capture Plate Fastener Labels    Tool Operation

Before the STIS FCP can be installed, the worksite  needs to be prepared. This involves removing a  handrail and a radiator clamp that block access to the Front Panel. Both of these items are held in place  with fasteners, which must also be disengaged, capt ured, and removed. This is accomplished with the  Handrail Removal Tool and the Clamp Removal Tool. T hese tools function with similar fastener capturing  technology as the Fastener Capture Plate. Removing the handrail leaves behind the handrail’s stanchions, seen in Figure 8. The handrail and clamp are removed and stowed in trash bags. Then anchor points must be installed onto th e worksite for attaching the FCP.  These anchor points are created by removing four fa steners from the Front Panel, one at each corner,  and replacing them with custom studs. Once they are in place, the FCP can be installed. Figure 12 shows  NASA/CP-2010-216272
154 FCP installation. First, the FCP is positioned in pl ace and the left side is lowered beneath the L-Handle.  Second, the right side can be lowered such that the stud s align with their mating holes in the FCP. Then  third, the four corresponding nuts ca n be tightened onto the four studs, essentially fastening the FCP in  place on top of the Front Panel. 

Figure 12 – STIS FCP Installation 

In order to tighten the four nuts, the operator may need to let go of the FCP. In microgravity, this could  allow the FCP to float freely and become misaligned. To prevent this, the FCP has a “soft dock” mechanism that holds it in place temporarily. This mechanism is essentially two spring-loaded plungers  that interface with the handrail stanchio ns at the worksite, shown in Figure 13. 

Figure 13 – STIS FCP Soft Dock Details 

Spring  Plunger  Soft Dock Feature  1 2 3 FCP L-Handle 2x ZipNuts®  2x Wingnuts   NASA/CP-2010-216272
155 Once the FCP is fully installed, the task of remo ving the remaining 107 fasteners and the Front Panel can  begin. Each fastener can be disengaged by a bit thr ough an Access Hole in the windows in the FCP. The  Access Hole will assist a lignment within specified tolerances of the screw head s once the FCP is installed  properly. Figure 14 shows a detailed view of the window design. Some windows may have multiple  Access Holes, and some Access Holes may have countersinks to help guide the bit into place.  

Figure 14 – Close-Up View of an FCP Window Component 

Figure 15 – Close-Up View of a Fastener Being Disengaged Through a Window  Bit  #4 Torq-Set Fastener Access Hole  Individual Window Countersink  NASA/CP-2010-216272