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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2005

System Design and Performance of the Two-Gyro Science Mode For the Hubble Space Telescope

For fifteen years, the science mission of the Hubble Space Telescope (HST) required using at least three of the six on-board rate gyros for attitude control. Failed gyros were eventually replaced through Space Shuttle Servicing Missions. The tragic loss of the Space Shuttle Columbia has resulted in the cancellation of all planned Shuttle based missions to HST. While a robotic servicing mission is currently being planned instead, controlling with alternate sensors to replace failed gyros can extend the HST science gathering until a servicing mission can be performed, and also extend science at HST s end of life. Additionally, sufficient performance may allow a permanent transition to operations with less than 3 gyros (by intentionally turning off working gyros saving them for later use) allowing for an even greater science mission extension. To meet this need, a Two Gyro Science (TGS) mode has been designed and implemented using magnetometers (Magnetic Sensing System - MSS), Fixed Head Star Trackers (FHSTs), and Fine Guidance Sensors (FGSs) to control vehicle rate about the missing gyro input axis. The development of the TGS capability is the largest re-design of HST operations undertaken, since it affects several major spacecraft subsystems, the most heavily being the Pointing Control System (PCS) and Flight Software (FSW). Additionally, and equally important, are the extensive modifications and enhancements of the Planning and Scheduling system which must now be capable of scheduling science observations while taking into account several new constraints imposed by the TGS operational modes (such as FHST availability and magnetic field geometry) that will impact science gathering efficiency and target availability. This paper discusses the systems engineering design, development, and performance of the TGS mode, now in its final stages of completion.

Prior, Michael↗
NASA NTRS2005

Attitude Control System Design for the Solar Dynamics Observatory

The Solar Dynamics Observatory mission, part of the Living With a Star program, will place a geosynchronous satellite in orbit to observe the Sun and relay data to a dedicated ground station at all times. SDO remains Sun- pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitude control system is a single-fault tolerant design. Its fully redundant attitude sensor complement includes 16 coarse Sun sensors, a digital Sun sensor, 3 two-axis inertial reference units, 2 star trackers, and 4 guide telescopes. Attitude actuation is performed using 4 reaction wheels and 8 thrusters, and a single main engine nominally provides velocity-change thrust. The attitude control software has five nominal control modes-3 wheel-based modes and 2 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. The paper details the mode designs and their uses.

Starin, Scott R.↗
NASA NTRS2005

Wilkinson Microwave Anisotropy Probe (WMAP) Attitude Estimation Filter Comparison

The Wilkinson Microwave Anisotropy Probe (WMAP) spacecraft was launched in June of 2001. The sensor complement of WMAP consists of two Autonomous Star Trackers (ASTs), two Fine Sun Sensors (FSSs), and a gyro package which contains redundancy about one of the WMAP body axes. The onboard attitude estimation filter consists of an extended Kalman filter (EKF) solving for attitude and gyro bias errors which are then resolved into a spacecraft attitude quaternion and gyro bias. A pseudo-linear Kalman filter has been developed which directly estimates the spacecraft attitude quaternion, rate, and gyro bias. In this paper, the performance of the two filters is compared for the two major control modes of WMAP: inertial mode and observation mode.

Harman, Richard R.↗
NASA NTRS2002

Deep Space 1 flight experience: adventures on an ion drive (AAS 02-072)

This paper provides a brief overview of the DSI attitude control subsystem, shares a few lessons-learned, and describes some of the many daunting challenges faced by our tiny flight team during the course of the mission. Special focus will be given to the nuances of flying a spacecraft with ion propulsion, our nick-of-time rewrite of the attitude determination software after the failure of the on-board star tracker in late 1999, and DSl's subsequent successful flyby of comet Borrelly on September 22, 2001.

SEP ion propulsion DS1 flight experience↗
NASA NTRS2006

Deep impact attitude estimator design and flight performance

In July, 2005 Deep Impact successfully impacted comet Tempel-1 and imaged the crater evolution. To achieve this, an onboard autonomous navigation system had to estimate the trajectory and maneuver the Impactor spacecraft into the comet's path. A similar system, operating on the Flyby spacecraft, calculated the impact location for imaging. At the heart of each autonomous navigation system is the attitude and gyro bias estimator using star trackers and gyros. The encounter geometry and timing placed stringent demands on attitude bias magnitude and drift calling for sequenced tuning of the respective filters. This paper describes the attitude and gyro bias estimator algorithms, the estimator flight performance and particular challenges overcome by the team leading up to a spectacular encounter.

Tempel-1↗
NASA NTRS2004

On-orbit performance testing of the Pointing Calibration & Reference Sensor for the Spitzer Space Telescope

We present the on-orbit performance results of the Pointing Calibration and Reference Sensor (PCRS) for the Spitzer Space Telescope. A cryogenic optical (center wavelength 0.55 mu) imager, the PCRS serves as the Observatory's fine guidance sensor by providing an alignment reference between the telescope boresight and the external spacecraft attitude determination system. The PCRS makes precision measurements of the positions of known guide stars; these are used to calibrate measurements from Spitzer's star tracker and gyroscopes to obtain the actual pointing of the Spitzer telescope. The PCRS calibrates out thermomoechanical drifts between the 300 K spacecraft bus and the 5.5 K telescope. By using only 16 pixels, the PCRS provides high precision centroiding with extremely low (`64 mu W) power dissipation, resulting in minimal impact to Spritzer's helium lifetime. We have demonstrated that the PCRS meets its centroiding accuracy requirement of 0.14 arcsec 1-sigma radial, which represents about1/100 pixel centroiding. The Spitzer Space Telescope was launched in 25 August, 2003 and completed its In-Orbit Checkout phase two months later; the PCRS has been operating failure free ever since.

Spitzer↗
NASA NTRS2007

Gravity Probe B Timing System and Roll Phase Determination

An oven-controlled crystal oscillator at 16.368 MHz provides clock signals to all GP-B components and synchronizes the data collection, transmission and processing. The sampled science data signals are stamped with the vehicle time, a counter of the 10Hz data strobe divided down from the clock. The GPS receiver supplies an external reference for time transfer between the vehicle time and coordinated universal time. Ground and space flight tests show the time transfer error is within 1 microsecond. The time latency between the effective sample time of science signals and the stamped vehicle time is verified to 1 ms in the ground tests. The GP-B satellite is controlled to roll with a period of 77.5 sec about an axis along the direction to the guide star to average out the disturbance torques fixed to the body of the satellite and reduce the gyroscope readout noise. The roll phase is determined on the ground to high accuracy with the telemetry data from two star trackers and used in the data analysis to separate the drifts of gyroscope spin axes in the orbital plane and perpendicular to the orbital plane. The flight data shows that the roll phase is controlled to within 40 arcsec with a measurement uncertainty of 7 arcsec.

Jie, Li↗
NASA NTRS1990

Submillimeter Wave Astronomy Satellite Conceptual Design Review

The SWAS instrument consists of the moving telescope assembly, the interface baseplate assembly, and the thermal control housing. The movin g telescope structure holds the primary mirror, secondary mirror, and receiver front end in precise alignment. This structure also carries the receiver cold plate radiators, which view cold space both directl y and through reflection off the primary mirror. The moving telescope assembly attaches to the interface baseplate with an open pivot frame through which flex leads from the receiver front end also pass. Two flexure-mounted linear actuators tilt the moving telescope assembly u p to +/-3 deg in two axes. The interface plate serves as a mounting for these actuators as well as the acousto-optic spectrometer, the inst rument control electronics, the star tracker, and the balance of the receiver components. The thermal control housing attaches to the inte rface baseplate around its edge. The top cylindrical part of this hou sing serves to shade the receiver cold plate radiators from Sun and Earth influence. The bottom "D" shaped portion of the housing is therma lly isolated from the top and forms the electronlcs radiator. Heat is conducted from the Interface baseplate into this lower portion of the shell. The radiator geometry and conduction paths create a very stab le thermal environment in the center of the interface baseplate where the AOS and receiver IF amplifiers are heat sinked. The outer surface of the thermal control housing is covered with fused silica second s urface mirrors that reflect visible light and radiate infrared energy . The instrument is mounted to the spacecraft through four thermally- isolating titanium flexures an the underside of the interface basepla te.

Source record↗
NASA NTRS2007

Advanced Optical Technologies for Space Exploration

NASA Langley Research Center is involved in the development of photonic devices and systems for space exploration missions. Photonic technologies of particular interest are those that can be utilized for in-space communication, remote sensing, guidance navigation and control, lunar descent and landing, and rendezvous and docking. NASA Langley has recently established a class-100 clean-room which serves as a Photonics Fabrication Facility for development of prototype optoelectronic devices for aerospace applications. In this paper we discuss our design, fabrication, and testing of novel active pixels, deformable mirrors, and liquid crystal spatial light modulators. Successful implementation of these intelligent optical devices and systems in space, requires careful consideration of temperature and space radiation effects in inorganic and electronic materials. Applications including high bandwidth inertial reference units, lightweight, high precision star trackers for guidance, navigation, and control, deformable mirrors, wavefront sensing, and beam steering technologies are discussed. In addition, experimental results are presented which characterize their performance in space exploration systems.

Clark, Natalie↗
NASA NTRS2007

Guaranteeing Pointing Performance of the SDO Sun-Pointing Controllers in Light of Nonlinear Effects

The Solar Dynamics Observatory (SDO) mission is the first Space Weather Research Network mission, part of NASA s Living With a Star program.1 This program seeks to understand the changing Sun and its effects on the Solar System, life, and society. To this end, the SDO spacecraft will carry three Sun-observing instruments to geosynchronous orbit: 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. Links describing the instruments in detail may be found through the SDO web site.2 The basic mission goals are 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. These goals guided the design of the spacecraft bus that will carry and service the three-instrument payload. At the time of this publication, the SDO spacecraft bus is well into the integration and testing phase at the NASA Goddard Space Flight Center (GSFC). A three-axis stabilized attitude control system (ACS) is needed both to point at the Sun accurately and to keep the roll about the Sun vector correctly positioned. The ACS has four reaction wheel modes and 2 thruster actuated modes. More details about the ACS in general and the control modes in particular can be found in Refs. [3-6]. All four of SDO s wheel-actuated control modes involve Sun-pointing controllers, as might be expected from such a mission. Science mode, during which most science data is collected, uses specialized guide telescopes to point accurately at the Sun. Inertial mode has two sub-modes, one tracks a Sun-referenced target orientation, and another maintains an absolute (star-referenced) target orientation, that both employ a Kalman filter to process data from a digital Sun sensor and two star trackers. However, this paper is concerned only with the other two modes: Safe Hold (SH) and Sun Acquisition (SA).

Starin, Scott R.↗
NASA NTRS2007

Attitude Sensor and Gyro Calibration for Messenger

The Redundant Inertial Measurement Unit Attitude Determination/Calibration (RADICAL(TM)) filter was used to estimate star tracker and gyro calibration parameters using MESSENGER telemetry data from three calibration events. We present an overview of the MESSENGER attitude sensors and their configuration is given, the calibration maneuvers are described, the results are compared with previous calibrations, and variations and trends in the estimated calibration parameters are examined. The warm restart and covariance bump features of the RADICAL(TM) filter were used to estimate calibration parameters from two disjoint telemetry streams. Results show that the calibration parameters converge faster with much less transient variation during convergence than when the filter is cold-started at the start of each telemetry stream.

O'Shaughnessy, Daniel↗
NASA NTRS2007

Fault Detection and Correction for the Solar Dynamics Observatory Attitude Control System

The Solar Dynamics Observatory is an Explorer-class mission that will launch in early 2009. The spacecraft will operate in a geosynchronous orbit, sending data 24 hours a day to a devoted ground station in White Sands, New Mexico. It will carry a suite of instruments designed to observe the Sun in multiple wavelengths at unprecedented resolution. The Atmospheric Imaging Assembly includes four telescopes with focal plane CCDs that can image the full solar disk in four different visible wavelengths. The Extreme-ultraviolet Variability Experiment will collect time-correlated data on the activity of the Sun's corona. The Helioseismic and Magnetic Imager will enable study of pressure waves moving through the body of the Sun. The attitude control system on Solar Dynamics Observatory is responsible for four main phases of activity. The physical safety of the spacecraft after separation must be guaranteed. Fine attitude determination and control must be sufficient for instrument calibration maneuvers. The mission science mode requires 2-arcsecond control according to error signals provided by guide telescopes on the Atmospheric Imaging Assembly, one of the three instruments to be carried. Lastly, accurate execution of linear and angular momentum changes to the spacecraft must be provided for momentum management and orbit maintenance. In th~sp aper, single-fault tolerant fault detection and correction of the Solar Dynamics Observatory attitude control system is described. The attitude control hardware suite for the mission is catalogued, with special attention to redundancy at the hardware level. Four reaction wheels are used where any three are satisfactory. Four pairs of redundant thrusters are employed for orbit change maneuvers and momentum management. Three two-axis gyroscopes provide full redundancy for rate sensing. A digital Sun sensor and two autonomous star trackers provide two-out-of-three redundancy for fine attitude determination. The use of software to maximize chances of recovery from any hardware or software fault is detailed. A generic fault detection and correction software structure is used, allowing additions, deletions, and adjustments to fault detection and correction rules. This software structure is fed by in-line fault tests that are also able to take appropriate actions to avoid corruption of the data stream.

Starin, Scott R.↗
NASA NTRS2010

Crater Identification Algorithm for the Lost in Low Lunar Orbit Scenario

Recent emphasis by NASA on returning astronauts to the Moon has placed attention on the subject of lunar surface feature tracking. Although many algorithms have been proposed for lunar surface feature tracking navigation, much less attention has been paid to the issue of navigational state initialization from lunar craters in a lost in low lunar orbit (LLO) scenario. That is, a scenario in which lunar surface feature tracking must begin, but current navigation state knowledge is either unavailable or too poor to initiate a tracking algorithm. The situation is analogous to the lost in space scenario for star trackers. A new crater identification algorithm is developed herein that allows for navigation state initialization from as few as one image of the lunar surface with no a priori state knowledge. The algorithm takes as inputs the locations and diameters of craters that have been detected in an image, and uses the information to match the craters to entries in the USGS lunar crater catalog via non-dimensional crater triangle parameters. Due to the large number of uncataloged craters that exist on the lunar surface, a probability-based check was developed to reject false identifications. The algorithm was tested on craters detected in four revolutions of Apollo 16 LLO images, and shown to perform well.

Hanak, Chad↗
NASA NTRS2007

Estimating the Inertia Matrix of a Spacecraft

A paper presents a method of utilizing some flight data, aboard a spacecraft that includes reaction wheels for attitude control, to estimate the inertia matrix of the spacecraft. The required data are digitized samples of (1) the spacecraft attitude in an inertial reference frame as measured, for example, by use of a star tracker and (2) speeds of rotation of the reaction wheels, the moments of inertia of which are deemed to be known. Starting from the classical equations for conservation of angular momentum of a rigid body, the inertia-matrix-estimation problem is formulated as a constrained least-squares minimization problem with explicit bounds on the inertia matrix incorporated as linear matrix inequalities. The explicit bounds reflect physical bounds on the inertia matrix and reduce the volume of data that must be processed to obtain a solution. The resulting minimization problem is a semidefinite optimization problem that can be solved efficiently, with guaranteed convergence to the global optimum, by use of readily available algorithms. In a test case involving a model attitude platform rotating on an air bearing, it is shown that, relative to a prior method, the present method produces better estimates from few data.

Acikmese, Behcet↗
NASA NTRS2010

Solar Array Disturbances to Spacecraft Pointing During the Lunar Reconnaissance Orbiter (LRO) Mission

The Lunar Reconnaissance Orbiter (LRO), the first spacecraft to support NASA s return to the Moon, launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle. It was initially inserted into a direct trans-lunar trajectory to the Moon. After a five day transit to the Moon, LRO was inserted into the Lunar orbit and successfully lowered to a low altitude elliptical polar orbit for spacecraft commissioning. Successful commissioning was completed in October 2009 when LRO was placed in its near circular mission orbit with an approximate altitude of 50km. LRO will spend at least one year orbiting the Moon, collecting lunar environment science and mapping data, utilizing a suite of seven instruments to enable future human exploration. The objective is to provide key science data necessary to facilitate human return to the Moon as well as identification of opportunities for future science missions. LRO's instrument suite will provide the high resolution imaging data with sub-meter accuracy, highly accurate lunar cartographic maps, mineralogy mapping, amongst other science data of interest. 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. This controller combines the capability of fine pointing with on-demand large angle full-sky attitude reorientation. It provides simplicity of spacecraft operation as well as additional flexibility for science data collection. A conventional suite of ACS components is employed in the Observing Mode to meet the pointing and control objectives. Actuation is provided by a set of four reaction wheels developed in-house at NASA Goddard Space Flight Center (GSFC). Attitude feedback is provided by a six state Kalman filter which utilizes two SELEX Galileo Star Trackers for attitude updates, and a single Honeywell Miniature Inertial Measurement Unit (MIMU) to provide body rates for attitude propagation. Rate is computed by differentiating accumulated angle provided by the MIMU. The Observing Mode controller is required to maintain fine pointing while a large fully-articulated solar array (SA) maintains its panel normal to the solar incidence. This paper describes the disturbances to the attitude control resulting from the SA articulation. Observing Mode performance in the presence of this disturbance was assessed while the spacecraft was in an initial elliptical low altitude orbit during the commissioning phase, which started about two weeks after launch and lasted for 90 days. LRO demonstrated excellent pointing performance during Observing Mode nadir and inertial attitude target operations during this phase. Transient LRO attitude errors observed during commissioning resulted primarily from three sources, Diviner instrument calibrations, RW zero crossings, and SA articulation. Even during times of considerable disturbance from SA articulation, the attitude errors were maintained below the statistical attitude error requirement level of 15 arc-sec (3 sigma).

Calhoun, Philip↗
NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO) Thruster Control Mode Design and Flight Experience

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.

Hsu, Oscar C.↗
NASA NTRS2010

Launch and Commissioning of the Lunar Reconnaissance Orbiter (LRO)

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.

Shah, Neerav↗
NASA NTRS2011

Orion Relative Navigation Flight Software Analysis and Design

The Orion relative Navigation System has sought to take advantage of the latest developments in sensor and algorithm technology while living under the constraints of mass, power, volume, and throughput. In particular, the only sensor specifically designed for relative navigation is the Vision Navigation System (VNS), a lidar-based sensor. But it uses the Star Trackers, GPS (when available) and IMUs, which are part of the overall Orion navigation sensor suite, to produce a relative state accurate enough to dock with the ISS. The Orion Relative Navigation System has significantly matured as the program has evolved from the design phase to the flight software implementation phase. With the development of the VNS system and the STORRM flight test of the Orion Relative Navigation hardware, much of the performance of the system will be characterized before the first flight. However challenges abound, not the least of which is the elimination of the RF range and range-rate system, along with the development of the FSW in the Matlab/Simulink/Stateflow environment. This paper will address the features and the rationale for the Orion Relative Navigation design as well as the performance of the FSW in a 6-DOF environment as well as the initial results of the hardware performance from the STORRM flight.

D'Souza, Chris↗