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A branched nanostructure is synthesized. A porous material, with pores having a diameter of approximately 1 .mu.m or less, is placed in a vacuum. It is irradiated with an electron beam. This causes a trunk to grow from the porous material and further causes branches to grow from the trunk.
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.
National Aeronautics and Space Administration s (NASA) Goddard Space Flight Center (GSFC) in Greenbelt, MD, designed, built, tested, and launched the Lunar Reconnaissance Orbiter (LRO) from Cape Canaveral Air Force Station on June 18, 2009. The LRO spacecraft is the first operational spacecraft designed to support NASA s return to the Moon, as part of the Vision for Space Exploration. LRO was launched aboard an Atlas V 401 launch vehicle into a direct insertion trajectory to the Moon. Twenty-four hours after separation the propulsion system was used to perform a mid-course correction maneuver. Four days after the mid-course correction a series of propulsion maneuvers were executed to insert LRO into its commissioning orbit. The commission period lasted eighty days and this followed by a second set of thruster maneuvers that inserted LRO into its mission orbit. To date, the spacecraft has been gathering invaluable data in support of human s future return to the moon. The LRO Attitude Control Systems (ACS) contains two thruster based control modes: Delta-H and Delta-V. The design of the two controllers are similar in that they are both used for 3-axis control of the spacecraft with the Delta-H controller used for momentum management and the Delta-V controller used for orbit adjust and maintenance maneuvers. In addition to the nominal purpose of the thruster modes, the Delta-H controller also has the added capability of performing a large angle slew maneuver. A suite of ACS components are used by the thruster based control modes, for both initialization and control. For initialization purposes, a star tracker or the Kalman Filter solution is used for providing attitude knowledge and upon entrance into the thruster based control modes attitude knowledge is provided via rate propagation using a inertial reference unit (IRU). Rate information for the controller is also supplied by the IRU. Three-axis control of the spacecraft in the thruster modes is provided by eight 5-lbf class attitude control thrusters configured in two sets of four thrusters for redundancy purposes. Four additional 20-lbf class thrusters configured in two sets of two thrusters are used for Lunar Orbit Insertion maneuvers. The propulsion system is one the few systems on-board the LRO spacecraft that has built in redundancy. The Delta-H controller consists of a Proportional-Derivative (PD) controller with a structural filter on the thrusters and a Proportional controller on the reaction wheels. The PD control that employs the thrusters is used for attitude and rate control. The Proportional controller on the reaction wheels is used for commanding the wheels to a new momentum state. The ground commands used for the Delta-H controller are the system momentum vector, reaction wheel momentum, maximum expected command time, and which set of attitude control thrusters to use. The ability to command both the system momentum vector and reaction wheel momentum in the Delta-H controller provides both a capability and an additional source of operator error. Large angle slews via the Delta-H controller is achievable via this commands because these commands are used for the exit mode criteria. Setting these commands to non-consistent values prevents the mode from exiting nominally.
The Lunar Reconnaissance Orbiter (LRO) launched on June 18, 2009 from the Cape Canaveral Air Force Station. LRO, designed, built, and operated by the National Aeronautics and Space Administration (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. To date, the Guidance, Navigation and Control (GN&C) subsystem has operated nominally and met all requirements. However, during the early phase of the mission, the GN&C Team encountered some anomalies. For example, during the Solar Array and High Gain Antenna deployments, one of the safing action points tripped, which was not expected. Also, the spacecraft transitioned to its safe hold mode, SunSafe, due to encountering an end of file for an ephemeris table. During the five-day lunar acquisition, one of the star trackers triggered the spacecraft to transition into a safe hold configuration, the cause of which was determined. These events offered invaluable insight to better understand the performance of the system they designed. An overview of the GN&C subsystem will be followed by a mission timeline. Then, interesting flight performance as well as anomalies encountered by the GN&C Team will be discussed in chronological order.
Safety-critical distributed flight control systems require robustness in the presence of faults. In general, these systems consist of a number of input/output (I/O) and computation nodes interacting through a fault-tolerant data communication system. The communication system transfers sensor data and control commands and can handle most faults under typical operating conditions. However, the performance of the closed-loop system can be adversely affected as a result of operating in harsh environments. In particular, High-Intensity Radiated Field (HIRF) environments have the potential to cause random fault manifestations in individual avionic components and to generate simultaneous system-wide communication faults that overwhelm existing fault management mechanisms. This paper presents the design of an experiment conducted at the NASA Langley Research Center's HIRF Laboratory to statistically characterize the faults that a HIRF environment can trigger on a single node of a distributed flight control system.
The Solar Dynamics Observatory (SDO) was designed and built at the Goddard Space Flight Center, launched from Cape Canaveral on February 11, 2010, and reached its final geosynchronous science orbit on March 16, 2010. The purpose of SDO is to observe the Sun and continuously relay data to a dedicated ground station. SDO remains Sun-pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitude control system (ACS) is a single-fault tolerant design. Its fully redundant attitude sensor complement includes sixteen coarse Sun sensors (CSSs), a digital Sun sensor (DSS), three two-axis inertial reference units (IRUs), and two star trackers (STs). The ACS also makes use of the four guide telescopes included as a part of one of the science instruments. Attitude actuation is performed using four reaction wheels assemblies (RWAs) and eight thrusters, with a single main engine used to provide velocity-change thrust for orbit raising. The attitude control software has five nominal control modes, three wheel-based modes and two thruster-based modes. A wheel-based Safehold running in the attitude control electronics box improves the robustness of the system as a whole. All six modes are designed on the same basic proportional-integral-derivative attitude error structure, with more robust modes setting their integral gains to zero. This paper details the final overall design of the SDO guidance, navigation, and control (GN&C) system and how it was used in practice during SDO launch, commissioning, and nominal operations. This overview will include the ACS control modes, attitude determination and sensor calibration, the high gain antenna (HGA) calibration, and jitter mitigation operation. The Solar Dynamics Observatory mission is part of the NASA Living With a Star program, which seeks to understand the changing Sun and its effects on the Solar System, life, and society. To this end, the SDO spacecraft carries three Sun-observing instruments: Helioseismic and Magnetic Imager (HMI), led by Stanford University; Atmospheric Imaging Assembly (AIA), led by Lockheed Martin Space and Astrophysics Laboratory; and Extreme Ultraviolet Variability Experiment (EVE), led by the University of Colorado. The basic mission is to observe the Sun for a very high percentage of the 5-year mission (10-year goal) with long stretches of uninterrupted observations and with constant, high-data-rate transmission to a dedicated ground station to be located in White Sands, New Mexico. These goals guided the design of the spacecraft bus that will carry and service the three-instrument payload. Overarching design goals for the bus are geosynchronous orbit, near-constant Sun observations with the ability to fly through eclipses, and constant HGA contact with the dedicated ground station. A three-axis stabilized ACS is needed both to point at the Sun accurately and to keep the roll about the Sun vector correctly positioned with respect to the solar north pole. This roll control is especially important for the magnetic field imaging of HM I. The mission requirements have several general impacts on the ACS design. Both the AIA and HMI instruments are very sensitive to the blurring caused by jitter. Each has an image stabilization system (ISS) with some ability to filter out high frequency motion, but below the bandwidth of the ISS the control system must compensate for disturbances within the ACS bandwidth or avoid exciting jitter at higher frequencies. Within the ACS bandwidth, the control requirement imposed by AIA is to place the center of the solar disk no more than 2 arc sec, 3 , from a body-defined target based on one of the GTs that accompany the instrument. This body-defined target, called the science reference boresight (SRB), was determined from the postlaunch orientation of the GTs by averaging the bounding telescope boresights for pitch to get a pitch SRB coordinate, and by averaging the bounding boresights for yaw toet the yaw SRB coordinate. The location of this SRB in the 0.5-deg field-of-view for each GT then becomes the central target for each telescope; one GT is selected for use as the ACS controlling guide telescope (CGT) at any given time. Fine Sun-pointing is effected based on this SRB for all three instruments when the Sun is within the linear range of the CGT. In addition to limiting jitter, HMI science requires averaging several observations, making the instrument sensitive to low frequency motion that induces differential motion between each observation. This requires the spacecraft attitude to be stable about the roll axis to approximately 10 arcsec over a ten-minute period. Instrument calibrations require that the spacecraft point the SRB up to 2.5 degrees in pitch and yaw away from the center of the Sun, placing the Sun outside the field-of-view of the guide telescopes. In such instances, when the GTs cannot provide the definitive target for the ACS, on-board attitude determination combined with ephemeris prediction of the Sun direction must provide the definitive target. EVE is capable of observing the Sun with less dependence on attitude control. However, the ground data processing needs for calibrations result in the most strict attitude knowledge requirements for the mission: [35,70,70] arcsec, 3 , of knowledge with respect to the center of the solar disk. In addition to driving the ACS sensor selection, the knowledge requirements, which have their effect primarily during Inertial mode calibrations, drive the accuracy requirements for the solar ephemeris. The need to achieve and maintain geosynchronous orbit (GEO) drove the need for high-efficiency propulsive systems and appropriate attitude control. The main engine provided high specific impulse for the maneuvers to attain GEO, while the smaller ACS thrusters managed the disturbance torques of the larger engine and provided the capability for much smaller adjustment burns on orbit. SDO s large solar profile means that solar radiation pressure is a large torque disturbance, and the momentum buildup from this disturbance and the GEO altitude drives the ACS to use thrusters to manage vehicle momentum. The demanding data capture budget for the mission, however, requires SDO to avoid frequent thruster maneuvers, while concerns about on-orbit jitter restrict the maximum desired wheel speeds desired from the RWAs. The plan for on-orbit wheel speed and momentum management will be discussed as well as what is now being done in operation after the jitter environment was characterized. The SDO ACS hardware complement is single-fault tolerant. Two main processors carry virtually identical copies of the command and data handling and ACS software, and two identical attitude control electronics (ACE) boxes carry Coldfire processors with contingency ACS software and other hardware interface cards; the ACE structure allows reaction wheels to be commanded by the Sun-pointing Safehold independent of the Mil Std 1553 data bus. The sixteen Adcole CSSs are grouped into primary and backup sets of eight sensors, each set providing the ability to calculate a sun vector. Each set of eight eyes provides full 4 -steradian coverage. The Adcole DSS comprises an optics head and a separate electronics box providing a 1553 data interface. The electronics box is mounted inside the Faraday cage created by the spacecraft bus module. The DSS head with its 32- deg square FOV is mounted on the instrument module with its boresight along the spacecraft X axis, nearly aligned with the Sun during observations. Adcole has designed the DSS calibration parameters so that the accuracy is 0.24 arcminutes within 10 deg of the boresight, and diminishes to 3 arcminutes as the Sun moves towards the edges of its FOV . This DSS calibration scheme provides higher accuracy attitude determination over the range of the instrument calibration maneuvers.
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.
This slide presentation reviews the design, method of operation, and testing of a regenerable Volatile Organic Compound (VOC) control system that characterizes properties of sorbents used in separation technologies.
A control system strives to maintain the correct alignment of a laser beam in an interferometer dedicated to measuring the displacement or distance between two fiducial corner-cube reflectors. The correct alignment of the laser beam is parallel to the line between the corner points of the corner-cube reflectors: Any deviation from parallelism changes the length of the optical path between the reflectors, thereby introducing a displacement or distance measurement error. On the basis of the geometrical optics of corner-cube reflectors, the length of the optical path can be shown to be L = L(sub 0)cos theta, where L(sub 0) is the distance between the corner points and theta is the misalignment angle. Therefore, the measurement error is given by DeltaL = L(sub 0)(cos theta - 1). In the usual case in which the misalignment is small, this error can be approximated as DeltaL approximately equal to -L(sub 0)theta sup 2/2. The control system (see figure) is implemented partly in hardware and partly in software. The control system includes three piezoelectric actuators for rapid, fine adjustment of the direction of the laser beam. The voltages applied to the piezoelectric actuators include components designed to scan the beam in a circular pattern so that the beam traces out a narrow cone (60 microradians wide in the initial application) about the direction in which it is nominally aimed. This scan is performed at a frequency (2.5 Hz in the initial application) well below the resonance frequency of any vibration of the interferometer. The laser beam makes a round trip to both corner-cube reflectors and then interferes with the launched beam. The interference is detected on a photodiode. The length of the optical path is measured by a heterodyne technique: A 100- kHz frequency shift between the launched beam and a reference beam imposes, on the detected signal, an interferometric phase shift proportional to the length of the optical path. A phase meter comprising analog filters and specialized digital circuitry converts the phase shift to an indication of displacement, generating a digital signal proportional to the path length.
NH3 is a degradation product of SA9T, a solid-amine sorbent developed by Hamilton Sundstrand, that is continually emitted into the gas stream being conditioned by this sorbent. NH3 offgassing rates were measured using FTIR spectroscopy using a packed bed at similar contact times as offgassing tests conducted at Hamilton Sundstrand and at the Ames Research Center. The bed was challenged with moist air at several flow rates and humidities and NH3 concentration of the effluent was measured for several hours. The NH3 offgassing rates in open-loop testing were calculated from the steady state outlet NH3 concentration and flow rate. NH3 offgassing rates from SA9T were found to be influenced by the contact time with the adsorbent (flow rate) and by the humidity of the inlet gas stream, which are consistent with previous studies. Closed-loop vacuum-swing adsorption cycling rates verified that NH3 offgassing continues when a constant source of water vapor is present.
Life support systems for growing plants in microgravity should strive for providing optimal growing conditions and increased automation. Accurately tracking soil moisture content can forward both of these aims, so an attempt was made to instrument a microgravity growth module currently in development, the VEGGIE rooting pillow, in order to monitor moisture levels. Two electrode systems for a capacitance-based moisture sensor were tested. Trials with both types of electrodes showed a linear correlation between observed capacitance and water content over certain ranges of moisture within the pillows. Overall, both types of the electrodes and the capacitance-based moisture sensor are promising candidates for tracking water levels for microgravity plant growth systems.
NASA's Commercial Crew and Cargo Program (CCP) is stimulating efforts within the private sector to develop and demonstrate safe, reliable, and cost-effective space transportation capabilities. One initiative involves investigating the use of commercial electronic parts. NASA's CCP asked the NASA Engineering and Safety Center (NESC) to collect data to help frame the technical, cost, and schedule risk trades associated with electrical, electronic and electromechanical (EEE) parts selection and specifically expressed desire of some of the CCP partners to employ EEE parts of a lower grade than traditionally used in most NASA safety-critical applications. This document contains the outcome from the NESC's review and analyses.
Controlled environment chambers (e.g. growth chambers, SPAR chambers, or open-top chambers) are useful for measuring plant ecosystem responses to climatic variables and CO2 that affect plant water relations. However, data from chambers was found to overestimate responses of C fluxes to CO2 enrichment. Chamber data may be confounded by numerous artifacts (e.g. sidelighting, edge effects, increased temperature and VPD, etc) and this limits what can be measured accurately. Chambers can be used to measure canopy level energy balance under controlled conditions and plant transpiration responses to CO2 concentration can be elucidated. However, these measurements cannot be used directly in model development or validation. The response of stomatal conductance to CO2 will be the same as in the field, but the measured response must be recalculated in such a manner to account for differences in aerodynamic conductance, temperature and VPD between the chamber and the field.
The effect of gravity on the growth and development of plants has been the subject of scientific investigation for over a century. The results obtained in space to test specific hypotheses on gravitropism, gene expression, seed formation, or growth rate are affected by both the primary effect of the microgravity and secondary effects of the spaceflight environment. The secondary effects of the spaceflight environment include physical effects arising from physical changes, such as the absence of buoyancy driven convective mixing, altered behavior of liquids and gases, and the environmental conditions in the spacecraft atmosphere. Thus, the design of biological experiments (e.g. cells, plants, animals, etc.) conducted in microgravity must account for changes in the physical forces, as well as the environmental conditions, imposed by the specific spaceflight vehicle and experimental hardware. In addition, researchers must become familiar with other aspects of spaceflight experiments: payload integration with hardware developers, safety documentation and crew procedures, and the logistics of conducting flight and ground controls. This report reviews the physical and environmental factors that directly and indirectly affect the results of plant science experiments in microgravity and is intended to serve as a guide in the design and implementation plant experiments in space.
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Measurement and feedback control of nutrient solutions in plant root zones is critical to the development of healthy plants in both terrestrial and reduced-gravity environments. In addition to the water content, the amount of fertilizer in the nutrient solution is important to plant health. This typically requires a separate set of sensors to accomplish. A combination bulk moisture and salinity sensor has been designed, built, and tested with different nutrient solutions in several substrates. The substrates include glass beads, a clay-like substrate, and a nutrient-enriched substrate with the presence of plant roots. By measuring two key parameters, the sensor is able to monitor both the volumetric water content and salinity of the nutrient solution in bulk media. Many commercially available moisture sensors are point sensors, making localized measurements over a small volume at the point of insertion. Consequently, they are more prone to suffer from interferences with air bubbles, contact area of media, and root growth. This makes it difficult to get an accurate representation of true moisture content and distribution in the bulk media. Additionally, a network of point sensors is required, increasing the cabling, data acquisition, and calibration requirements. measure the dielectric properties of a material in the annular space of the vessel. Because the pore water in the media often has high salinity, a method to measure the media moisture content and salinity simultaneously was devised. Characterization of the frequency response for capacitance and conductance across the electrodes was completed for 2-mm glass bead media, 1- to 2-mm Turface (a clay like media), and 1- to 2-mm fertilized Turface with the presence of root mass. These measurements were then used to find empirical relationships among capacitance (C), the dissipation factor (D), the volumetric water content, and the pore water salinity.
Radiation use efficiency (RUE) has been a key parameter for developing simpler models of crop growth and yield. A great deal of effort has gone into measuring RUE in the field and in verifying its validity for predicting crop growth. However, a lack of data on responses of RUE to elevated C02 has resulted in the use of empirical relations that may lead to overestimates of crop yield to C02 enrichment. A new approach relating RUE to canopy quantum yield is presented that offers a more theoretical basis for estimating RUE responses to elevated C02.
The atmospheric composition inside spacecraft during long duration space missions is dynamic due to changes in the living and working environment of crew members, crew metabolism and payload operations. A portable FTIR gas analyzer was used to monitor the atmospheric composition within the Deep Space Habitat (DSH) during the Mission Operations Test (MOT) conducted at the Johnson Space Center (JSC). The FTIR monitored up to 20 gases in near- real time. The procedures developed for operating the FTIR were successful and data was collected with the FTIR at 5 minute intervals. Not all the 20 gases sampled were detected in all the modules and it was possible to measure dynamic changes in trace contaminant concentrations that were related to crew activities involving exercise and meal preparation.