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NASA NTRS2002

Reduced Toxicity Fuel Satellite Propulsion System Including Catalytic Decomposing Element with Hydrogen Peroxide

A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply for consumption in an axial class thruster and an ACS class thruster. The system includes suitable valves and conduits for supplying the reduced toxicity propellant to the ACS decomposing element of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits for supplying the reduced toxicity propellant to an axial decomposing element of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits for supplying a second propellant to a combustion chamber of the axial thruster, whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.

Steven J Schneider↗

NASA NTRS2002

Particle Segregation in a Flowing Suspension Subject to High-Gradient Strong Electric Fields

The widespread use of electro-hydrodynamic devices and processes emphasizes a critical need for developing a comprehensive predictive theory capable of improving our fundamental understanding of the behavior of a suspension subject to an AC electric field and shear, and of facilitating the design and optimization of such devices. The currently favored approach to the qualitative interpretation of the AC field driven manipulation of suspensions is based on a model which considers only the force exerted on a single particle by an external field and neglects the field-induced and hydrodynamic interparticle interactions both being inversely proportional to the interparticle distance raised to the power three. On the other hand, the purpose of the field-induced separation is to concentrate particles in certain regions of a device. This clearly raises the fundamental question regarding the extent to which we can neglect these slow decaying electrical and hydrodynamic collective interactions and rely on the predictions of a single-particle model. Another important issue that still remains open is how to characterize the polarization of a particle exposed to a strong electric field. The presentation will address both these questions. Experiments were conducted in a parallel-plate channel in which a 10(exp -3) (v/v) suspension of heavy, positively polarized Al2O3 spheres was exposed to an AC field under conditions such that the field lines were arranged in the channel cross-section perpendicular to the streamlines of the main flow. To reduce the effects of the gravitational settling of the particles, the channel was slowly rotated (4 rpm) around a horizontal axis. Following the application of a high-gradient strong AC field (approx. kV/mm), the particles were found to move towards both the high-voltage (HV) and grounded (GR) electrodes and to form 'bristles' along their edges.

Acrivos, Andreas↗

NASA NTRS2002

A Proposal to Investigate Outstanding Problems in Astronomy

During the period leading up to the spectacular launch of the Space Shuttle Columbia (STS-109) on 1 March 2002 6:22 am EST, the team worked hard on a myriad of tasks to be ready for launch. Our launch support included preparations and rehearsals for the support during the mission, preparation for the SMOV and ERO program, and work to have the science team's data pipeline (APSIS) and data archive (SDA) ready by launch. A core of the team that was at the GSFC during the EVA that installed ACS monitored the turn-on and aliveness tests of ACS. One hour after installation of ACS in the HST George Hartig was showing those of us at Goddard the telemetry which demonstrated that the HRC and WFC CCDs were cooling to their preset temperatures. The TECs had survived launch! After launch, the team had several immediate and demanding tasks. We had to process the ERO observations through our pipeline and understand the limitations of the ground based-based calibrations, and simultaneously prepare the EROs for public release. The ERO images and the SMOV calibrations demonstrated that ACS met or exceeded its specifications for image quality and sensitivity. It is the most sensitive instrument that Hubble has had. The ERO images themselves made the front page of all of the major newspapers in the US. During the months after launch we have worked on the SMOV observations, and are analyzing the data from our science program.

Ford, Holland↗

NASA NTRS2003

Reduced Toxicity Fuel Satellite Propulsion System Including Plasmatron

A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply for consumption in an axial class thruster and an ACS class thruster. The system includes suitable valves and conduits for supplying the reduced toxicity propellant to the ACS decomposing element of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits for supplying the reduced toxicity propellant to an axial decomposing element of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits for supplying a second propellant to a combustion chamber of the axial thruster. whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.

Schneider, Steven J.↗

NASA NTRS1946

Effect of Fuel Composition, Engine Operating Variables, and Spark-Plug Type and Condition on Preignition-Limited Performance of an R-2800 Cylinder

The preignition characteristics of the R-2800 cylinder, as effected by fuel consumption, engine operating variables, and spark plug type and condition, were evaluated. The effects on preignition-limited performance of various percentages of aromatics (benzene, toluene, cumene, xylene) in a base fuel of triptane were investigated. Two paraffins (triptane and S + 6.0 ml TEL/gal) and two refinery blends (28-R and 33-R) were preignition rated. The effect of changes in the following engine operating variables on preignition limit was determined: inlet-air temperature, rear spark plug gasket temperature, engine speed, spark advance, tappet clearance, and oil consumption. Preignition limits of the R-2800 cylinder using Champion C34S and C35S and AC-LS86, LS87, and LS88 spark plugs were established and the effect of spark plug deterioration was investigated. No definite trends in preignition-limited indicated mean effective pressure were indicated for aromatics as a class when increased percentages of different aromatics were added to a base fuel of triptane. Three types of fuel (aromatics, paraffins, and refinery blends) showed a preignition range for this cylinder from 65 to 104 percent when based on the performance of S plus 6.0 ml TEL per gallon as 100 percent. The R-2800 cylinder is therefore relatively insensitive to fuel composition when compared to a CFR F-4 engine, which had a pre-ignition range from 72 to 100 percent for the same fuels. Six engine operating variables were investigated with the following results: preignition-limited indicated mean effective pressure decreased, with increases in engine speed, rear spark plug gasket temperature, inlet-air temperature, and spark advance beyond 20 F B.T.C. and was unaffected by rate of oil consumption or by tappet clearance. Spark plugs were rated over a range of preignition-limited indicated mean effective pressure from 200 to 390 pounds per square inch at a fuel-air ratio of 0.07 in the following order of increased resistance to preignition: AC-LS97, AC-LS88, Champion C358, AC-LS86, and Champion C34S. Spark plug deterioration in the form of cracks in the porcelain had been broken away from the center electrode and were retained in the spark plug cavity, the preignition limit was decreased as much as 57 percent. When the broken pieces had been removed, the preignition limit increased from that of the undamaged porcelain as the weight of removed porcelain was increased.

Pfender, John F.↗

NASA NTRS1991

Characterization of Asymmetry in Magnetoacoustic Emission Burst by Numerical Processes

It has been well known that the pattern of the magnetoacoustic emission (MAE) burst observed during the sweep over one half-cycle of the hysteresis loop becomes asymmetric depending on the strength of the magnetic domain wall-defect interaction and the state of residual stresses in a ferromagnet. The ascending asymmetry due to the former has been observed at a very low frequency (.7 Hz) of applied AC magnetic field at a given amplitude. The descending asymmetry due to uniaxial compressive stress has been typically observed at the AC applied magnetic field frequency of 20 Hz. The physical interpretation of both types of asymmetry has been well established. It is, however, necessary to perform investigations of the dependence of asymmetry on externally controlled parameters such as the amplitude and frequency of the AC applied magnetic fields. The purpose of the present study is therefore to devise a mathematical means that describes the degree of asymmetry of the MAE burst and apply this scheme to investigate the AC magnetic field amplitude dependence of the asymmetry.

Namkung, M.↗

NASA NTRS1996

Isokinetic and isometric strength-endurance after 6 hours of immersion and 6 degrees head-down tilt in men

PURPOSE: To determine weight (water) loss levels for onset of muscular strength and endurance changes during deconditioning. METHODS: Seven men (27-40 yr) performed maximal shoulder-, knee-, and ankle-joint isometric (0 degree.s(-1) load) and isokinetic (60 degrees, 120 degrees, 180 degrees.s(-1) velocity) exercise tests during ambulatory control (AC), after 6 h of 6 degrees head-down tilt (HDT; dry-bulb temp. = 23.2 +/- SD 0.6 degrees C, relative humidity = 31.1+/- 11.1%) and after 6 h of 80 degrees foot-down head-out water immersion (WI; water temp. = 35.0 +/- SD 0.1 degree C) treatments. RESULTS: Weight (water) loss after HDT (1.10 +/- SE 0.14 kg, 1.4 +/- 0.2% body wt) and WI (1.54+/- 0.19 kg, 2.0 +/- 0.2% body wt) were not different, but urinary excretion with WI (1,354 +/- 142 ml.6 h(-1)) was 28% greater (p < 0.05) than that of 975 +/- 139 ml.6 h(-1) with HDT. Muscular endurance (total work; maximal flexion-extension of the non-dominant knee at 180 degrees.s(-1) for 30 s) was not different between AC and the WI or HDT treatments. Shoulder-, knee-, and ankle-joint strength was unchanged except for three knee-joint peak torques: AC torque (120 degrees.s(-1), 285 +/- 20 Nm) decreased to 268 +/- 21 Nm (delta = -6%, p < 0.05) with WI; and AC torques (180 degrees.s(-1), 260 +/- 19 Nm) decreased to 236 +/- 15 Nm (delta = -9%, p < 0.01) with HDT, and to 235 +/- 19 Nm (delta = -10%, p < 0.01) with WI. CONCLUSION: Thus, the total body hypohydration threshold level for shoulder- and ankle-joint strength and endurance decrements is more than 2% body weight (water) loss, while significant reduction in knee-joint muscular strength-endurance occurred only at moderate (120 degrees.s(-1) and lighter (180 degrees.s(-1)) loads with body weight loss of 1.4-2.0% following WI or HDT, respectively. These weight (water) losses and knee-joint strength decrements are somewhat less than the mean weight loss of 2.6% and knee-joint strength decrements of 6-20% of American astronauts after Skylab flights to 84 d.

NASA Center ARC↗

NASA NTRS2005

Line of Sight Stabilization of James Webb Space Telescope

The James Webb Space Telescope (JWST) builds upon the successful flight experience of the Chandra Xray Telescope by incorporating an additional LOS pointing servo to meet the more stringent pointing requirements. The LOS pointing servo, referred to in JWST as the Fine Guidance Control System (FGCS), will utilize a Fine Guidance Sensor (FGS) as the sensor, and a Fine Steering Mirror (FSM) as the actuator. The FSM is a part of the Optical Telescope Element (OTE) and is in the optical path between the tertiary mirror and the instrument focal plane, while the FGS is part of the Integrated Science Instrument Module (ISIM). The basic Chandra spacecraft bus attitude control and determination architecture, utilizing gyros, star trackers/aspect camera, and reaction wheels, is retained for JWST. This system has achieved pointing stability of better than 0.5 arcseconds. To reach the JWST requirements of milli-arcsecond pointing stability with this ACS hardware, the local FGCS loop is added to the optical path. The FGCS bandwidth is about 2.0 Hz and will therefore attenuate much of the spacecraft ACS induced low frequency jitter. In order to attenuate the higher frequency (greatet than 2.0 Hz) disturbances associated with reaction wheel static and dynamic imbalances, as well as bearing run-out, JWST will employ a two-stage passive vibration isolation system consisting of (1) 7.0 Hz reaction wheel isolators between each reaction wheel and the spacecraft bus, and (2) a 1.0 Hz tower isolator between the spacecraft bus and the Optical Telescope Element (OTE). In order to sense and measure the LOS, the FGS behaves much like an autonomous star tracker that has a very small field of view and uses the optics of the telescope. It performs the functions of acquisition, identification and tracking of stars in its 2.5 x 2.5 arcminute field of view (FOV), and provides the centroid and magnitude of the selected star for use in LOS control. However, since only a single star is being tracked at any time within the FGS FOV there is only tip and tilt information; rotation about the FGS LOS will not be sensed. The FGCS uses the FSM to move the guide star within the FGS FOV and place the centroid of the guide star at any desired position within the FGS focal plane. Using this architecture allows the FGCS to correct the low frequency LOS jitter that is induced by the spacecraft ACS in pitch and yaw, and achieve the milli-arcsecond pointing stability required by JWST. The less stringent ISIM FOV roll performance will be provided solely by the ACS, using the spacecraft gyros and star trackers. Since the FSM is in the optical path, the pointing stabilrty of a science object in any of the instruments will be similar to that of the guide star LOS.

Meza, Luis↗

NASA NTRS2004

Jupiter Icy Moons Orbiter (JIMO) Electrical Systems Testbed

The Jupiter Icy Moons Orbiter (JIMO) mission will send a spacecraft to explore three of Jupiter s moons (Callisto, Ganymede, and Europa), all of which show evidence of containing vast subterranean oceans beneath their icy surfaces. The evidence of these oceans was discovered by Galileo, and the moons are believed to have the three essential ingredients for life: water, energy, and the necessary chemical elements. Galileo has shown that melted water on Europa has been in contact with the surface of the moon in geologically recent times, and may still lie relatively close to the surface. This project will also introduce a revolutionary new form of electric propulsion powered by a nuclear fission reactor. This electric propulsion is called ion propulsion. It was used on a previous mission called Deep Space 1, proving that ion propulsion works for interplanetary travel. Since JIMO will be traveling farther from the sun, solar power will be difficult to supply the electric energy demanded by the mission. Therefore a nuclear reactor and a thermo-electric converter system will be necessary. Besides making the trip to three of Jupiter's moons - one after the other - a realistic possibility, this new form of power and propulsion opens up the rest of the outer solar system for future exploration. JIMO will fulfill its goals by exploring Europa first, with subsequent trips to the moons Callisto and Ganymede in order to provide comparisons key to understanding the evolution of all three. In order to ensure the stability and proper preparation of the electrical system on JIMO, the High Power AC Power Management and Distribution (PMAD) Test Bed is being developed. The testing on.this AC PMAD will consist of electrical performance verification of candidate power system components. Examples of these components are: high power AC switchgear, high power ACDC converters, AC power distribution units, DC power distribution units, etc. Throughout the course of the summer the over-current control circuit for the five different size relays will be constructed and tested. This circuit will sense the current input to the spacecraft loads and automatically switch off power if the current is too high. Once the circuit is verified to function properly and all necessary modifications have been made, a circuit schematic and board layout will have to be drawn using OrCAD, and the circuits will have to be built.

Trapp, Scott J.↗

NASA NTRS2012

Integrated Simulation Design Challenges to Support TPS Repair Operations

During the Orbiter Repair Maneuver (ORM) operations planned for Return to Flight (RTF), the Shuttle Remote Manipulator System (SRMS) must grapple the International Space Station (ISS), undock the Orbiter, maneuver it through a long duration trajectory, and orient it to an EVA crewman poised at the end of the Space Station Remote Manipulator System (SSRMS) to facilitate the repair of the Thermal Protection System (TPS). Once repair has been completed and confirmed, then the SRMS proceeds back through the trajectory to dock the Orbiter to the Orbiter Docking System. In order to support analysis of the complex dynamic interactions of the integrated system formed by the Orbiter, ISS, SRMS, and SSMS during the ORM, simulation tools used for previous ‘nominal’ mission support required substantial enhancements. These upgrades were necessary to provide analysts with the capabilities needed to study integrated system performance. Prevalent throughout this ORM operation is a dynamically varying topology. In other words, the ORM starts with the SRMS grappled to the mated Shuttle/ISS stack (closed loop topology), moves to an open loop chain topology consisting of the Shuttle, SRMS, and ISS, and then, at the repair configuration, extends the chain topology to one consisting of the Shuttle, SMS, ISS, and SSRMS/EVA crewman. The resulting long dynamic chain of vehicles and manipulators may exhibit significant motion between the Shuttle worksite and the EVA crewman due to the system flexibility throughout the topology (particularly within the SRMS/SSRMS joints and links). Since the attachment points of both manipulators span the flexible structure of the ISS, simulation analysis may also need to take that into consideration. Moreover, due to the lengthy time duration associated with the maneuver and repair, orbital effects become a factor and require the ISS vehicle control system to maintain active attitude control. Several facets of the ORM operation make the associated analytical efforts different from previous mission support, including: (1) the magnitude of the SRMS handled payload (Le., Orbiter class), (2) the orbital effects induced on the integrated system consisting of the large Shuttle and ISS masses connected by a light flexible SRMS, (3) long duration environmental consequences due to the lengthy operational times associated with the maneuver and repair of the TPS, (4) active attitude control (as opposed to free drift) interacting with the SRMS and SSRMS manipulators (also due to the length of the maneuver and repair), (5) relative dynamics between the EVA crewman and thc worksite influenced by the extended flexible topology. In order to meet these analysis challenges, an ORM simulation architecture was developed leveraging upon numerous pre-existing simulation elements to analyze the various subsystems individually. For example, core manipulator subsystem simulations for both the SRMS and SSRMS were originally combined to provide the dual-arm dynamics topology simulation (in the absence of orbital dynamics and vehicle control). This capability was later merged with the simulation used to analyze SRMS loading with a heavy payload in the orbital environment with an active payload control system (in this case, the ISS Attitude Control System (ACS)), configured for the ORM. The resulting worksite dynamics simulation, based off of the modified ORM simulation, provided the extended topological chain of vehicles and manipulators, while taking into account the orbital effects of both the Shuttle and ISS (as well as its ACS). Verification and validation (V&V) of these integrated simulations became a challenge in itself. A systematic approach needed to be developed such that integration simulation results could be tested against previous constituent simulations upon which these simulations were built. General V&V categories included: (1) core orbital state propagation, (2), stand-alone SRMS, (3) stand-alone SSRMS, (4) stand-alone ISS ACS, (5) integrated Shuttle, SRMS, ISS (with active ACS) in the orbital environment, and (5) dual-arm SRMS/SSRMS dynamics topology. Integrated simulation V&V run suites were created and correlated to verification runs from subsystem simulations, in order to establish the validity of the results. This paper discusses the simulation design challenges encountered while developing simulation capabilities to mirror the ORM operations. The paper also describes the incremental build approach that was utilized, starting with the subsystem simulation elements and integration into increasing more complex simulations until the resulting ORM worksite dynamics simulation had been assembled. Furthermore, the paper presents an overall integrated simulation V&V methodology based upon a subsystem level testing, integrated comparisons, and phased checkout.

Thermal Protection System↗

NASA NTRS2007

Observing Mode Attitude Controller for the Lunar Reconnaissance Orbiter

The Lunar Reconnaissance Orbiter (LRO) mission is the first of a series of lunar robotic spacecraft scheduled for launch in Fall 2008. LRO will spend at least one year in a low altitude polar orbit around the Moon, collecting lunar environment science and mapping data to enable future human exploration. The LRO employs a 3-axis stabilized attitude control system (ACS) whose primary control mode, the "Observing mode", provides Lunar Nadir, off-Nadir, and Inertial fine pointing for the science data collection and instrument calibration. The controller combines the capability of fine pointing with that of on-demand large angle full-sky attitude reorientation into a single ACS mode, providing simplicity of spacecraft operation as well as maximum flexibility for science data collection. A conventional suite of ACS components is employed in this mode to meet the pointing and control objectives. This paper describes the design and analysis of the primary LRO fine pointing and attitude re-orientation controller function, known as the "Observing mode" of the ACS subsystem. The control design utilizes quaternion feedback, augmented with a unique algorithm that ensures accurate Nadir tracking during large angle yaw maneuvers in the presence of high system momentum and/or maneuver rates. Results of system stability analysis and Monte Carlo simulations demonstrate that the observing mode controller can meet fine pointing and maneuver performance requirements.

Calhourn, Philip C.↗

NASA NTRS2007

Observing Mode Attitude Controller for the Lunar Reconnaissance Orbiter

Calhoun, Philip C.↗

NASA NTRS2005

Human Systems Integration: Unmanned Aircraft Control Station Certification Plan Guidance

This document provides guidance to the FAA on important human factors considerations that can be used to support the certification of a UAS Aircraft Control Station (ACS). This document provides a synopsis of the human factors analysis, design and test activities to be performed to provide a basis for FAA certification. The data from these analyses, design activities, and tests, along with data from certification/qualification tests of other key components should be used to establish the ACS certification basis. It is expected that this information will be useful to manufacturers in developing the ACS Certification Plan,, and in supporting the design of their ACS.

Source record↗

NASA NTRS2008

The HD 163296 Circumstellar Disk in Scattered Light: Evidence of Time-Variable Self-Shadowing

We present the first multi-color view of the scattered light disk of the Herbig Ae star HD 163296, based on coronagraphic observations from the Hubble Space Telescope Advanced Camera for Surveys (ACS). Radial profile fits of the surface brightness along the disk's semi-major axis indicates that the disk is not continuously flared, and extends to approx.540 AU. The disk's color (V-I)=1.1 at a radial distance of 3.5" is redder than the observed stellar color (V-I)=0.15. This red disk color might be indicative of either an evolution in the grain size distribution (i.e. grain growth) and/or composition, both of which would be consistent with the observed non-flared geometry of the outer disk. We also identify a single ansa morphological structure in our F435W ACS data, which is absent from earlier epoch F606W and F814W ACS data, but corresponds to one of the two ansa observed in archival HST STIS coronagraphic data. Following transformation to similar band-passes, we find that the scattered light disk of HD 163296 is 1 mag arcsec(sup -2) fainter at 3.5" in the STIS data than in the ACS data. Moreover, variations are seen in (i) the visibility of the ansa(e) structures, in (ii) the relative surface brightness of the ansa(e) structures, and in (iii) the (known) intrinsic polarization of the system. These results indicate that the scattered light from the HD 163296 disk is variable. We speculate that the inner disk wall, which Sitko et al. suggests has a variable scale height as diagnosed by near-IR SED variability, induces variable self-shadowing of the outer disk. We further speculate that the observed surface brightness variability of the ansa(e) structures may indicate that the inner disk wall is azimuthally asymmetric. Subject headings: circumstellar matter - stars: individual (HD 163296) - planetary systems: formation - planetary systems: protoplanetary disks

Wisniewski, John P.↗

NASA NTRS2008

Continued Development of the Advanced Stirling Convertor (ASC)

The Advanced Stirling Convertor (ASC) is being developed under contract with the NASA Glenn Research Center (GRC) and is supported by NASA s Science Mission Directorate for potential use in future radioisotope power systems having significantly increased efficiency and higher specific power compared to the current thermoelectric systems. An ASC with a lower temperature (approx.650 C) Inconel heater head is currently being substituted into the DOE/Lockheed Martin Advanced Stirling Radioisotope Generator (ASRG) program with a predicted convertor efficiency of 34 percent (AC electrical out to heat input ) at a temperature ratio of 2.7 and is expected to deliver approximately 75 W(sub ac). Continued development of the higher temperature (approx.850 C) version using existing materials and fabrication techniques in the hot portions is reported on here. The higher temperature ASC is expected to have 38 percent efficiency (AC electrical out to heat input) at a temperature ratio of 3.1 and is expected to deliver approximately 88 W(sub ac). The high temperature ASC also has approximately 30 C higher rejection temperature, which allows for further reduced system mass because of the reduced radiator size. Six higher temperature and hermetically sealed convertors are being built under this effort for extended life testing at GRC.

Wong, Wayne A.↗

NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO) Guidance, Navigation and Control (GN&C) Overview

The National Aeronautics and Space Administration s (NASA) Lunar Reconnaissance Orbiter (LRO) launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle and into a direct insertion trajectory to the oon. LRO, which was designed, built, and operated by the NASA Goddard Space Flight Center in Greenbelt, MD, is gathering crucial data on the lunar environment that will help astronauts prepare for long-duration lunar expeditions. The mission has a nominal life of 1 year as its seven instruments find safe landing sites, locate potential resources, characterize the radiation environment, and test new technology. To date, LRO has been operating well within the bounds of its requirements and has been collecting excellent science data images taken from the LRO Camera Narrow Angle Camera of the Apollo landing sites appeared on cable news networks. A significant amount of information on LRO s science instruments is provided at the LRO mission webpage. LRO s Guidance, Navigation and Control (GN&C) subsystem is made up of an onboard attitude control system (ACS) and a hardware suite of sensors and actuators. The LRO onboard ACS is a collection of algorithms based on high level and derived requirements, and reflect the science and operational events throughout the mission lifetime. The primary control mode is the Observing mode, which maintains the lunar pointing orientation and any offset pointing from this baseline. It is within this mode that all science instrument calibrations, slews and science data is collected. Because of a high accuracy requirement for knowledge and pointing, the Observing mode makes use of star tracker (ST) measurement data to determine an instantaneous attitude pointing. But even the star trackers alone do not meet the tight requirements, so a six-state Kalman Filter is employed to improve the noisy measurement data. The Observing mode obtains its rate information from an inertial reference unit (IRU) and in the event of an IRU failure, the rate data is be derived from the star tracker, but with degraded pointing performance. The Delta-V control mode responsibility is to maintain attitude pointing during the cruise trajectory, insertion burns and lunar orbit maintenance by adjustments made to the spacecraft s velocity magnitude and vector direction. The ACS also provides for a thruster based system momentum management algorithm (known as Delta-H) to maintain the system and wheel momentum to within acceptable levels. In the event an anomaly causes the LRO spacecraft to lose the ability to maintain its current attitude pointing, a Sun Safe mode is included in the ACS for the purpose of providing a known power and thermally safe coarse inertial sun attitude for an indefinite period of time, within the manageable limits of the reaction wheels. The Sun Safe mode is also the initial spacecraft control mode off of the launch vehicle and provides for a means to null tip-off rates immediately after separation. The nominal configuration is to use the IRU for rate information in the controller. In the event of a gyro failure a gyroless control mode was developed that computes rate information from the CSS data.

Garrick, Joseph↗

NASA NTRS2010

Integration and Testing of the Lunar Reconnaissance Orbiter Attitude Control System

Throughout the Lunar Reconnaissance Orbiter (LRO) Integration and Testing (I&T) phase of the project, the Attitude Control System (ACS) team completed numerous tests on each hardware component in ever more flight like environments. The ACS utilizes a select group of attitude sensors and actuators. This paper chronicles the evolutionary steps taken to verify each component was constantly ready for flight as well as providing invaluable trending experience with the actual hardware. The paper includes a discussion of each ACS hardware component, lessons learned of the various stages of I&T, a discussion of the challenges that are unique to the LRO project, as well as a discussion of work for future missions to consider as part of their I&T plan. LRO ACS sensors were carefully installed, tested, and maintained over the 18 month I&T and prelaunch timeline. Care was taken with the optics of the Adcole Coarse Sun Sensors (CSS) to ensure their critical role in the Safe Hold mode was fulfilled. The use of new CSS stimulators provided the means of testing each CSS sensor independently, in ambient and vacuum conditions as well as over a wide range of thermal temperatures. Extreme bright light sources were also used to test the CSS in ambient conditions. The integration of the two SELEX Galileo Star Trackers was carefully planned and executed. Optical ground support equipment was designed and used often to check the performance of the star trackers throughout I&T in ambient and thermal/vacuum conditions. A late discovery of potential contamination of the star tracker light shades is discussed in this paper. This paper reviews how each time the spacecraft was at a new location and orientation, the Honeywell Miniature Inertial Measurement Unit (MIMU) was checked for data output validity. This gyro compassing test was performed at several key testing points in the timeline as well as several times while LRO was on the launch pad. Sensor alignment tests were completed several times to ensure that hardware remained on a rigid platform.

Simpson, Jim↗

NASA NTRS2011

Solar Dynamics Observatory Launch and Commissioning

The Solar Dynamics Observatory (SDO) was launched on February 11, 2010. Over the next three months, the spacecraft was raised from its launch orbit into its final geosynchronous orbit and its systems and instruments were tested and calibrated in preparation for its desired ten year science mission studying the Sun. A great deal of activity during this time involved the spacecraft attitude control system (ACS); testing control modes, calibrating sensors and actuators, and using the ACS to help commission the spacecraft instruments and to control the propulsion system as the spacecraft was maneuvered into its final orbit. This paper will discuss the chronology of the SDO launch and commissioning, showing the ACS analysis work performed to diagnose propellant slosh transient and attitude oscillation anomalies that were seen during commissioning, and to determine how to overcome them. The simulations and tests devised to demonstrate correct operation of all onboard ACS modes and the activities in support of instrument calibration will be discussed and the final maneuver plan performed to bring SDO on station will be shown. In addition to detailing these commissioning and anomaly resolution activities, the unique set of tests performed to characterize SDO's on-orbit jitter performance will be discussed.

O'Donnell, James R., Jr.↗