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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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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Catalyzed Combustion of Bipropellants for Micro-Spacecraft Propulsion

This paper addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on micro-spacecraft. These spacecraft are planned to have a mass less than 50 kilograms with attitude control estimated to be in the 10 milli-Newton thrust class. These combustors are anticipated to be manufactured using Micro Electrical Mechanical Systems (MEMS) technology and are expected to have diameters approaching the quenching diameter of the propellants. Combustors of this size are expected to benefit significantly from surface catalysis processes. Miniature flame tube apparatus is chosen for this study because microtubes can be easily fabricated from known catalyst materials and their simplicity in geometry can be used in fundamental simulations for validation purposes. Experimentally, we investigated the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter microtubes, with special emphases on ignition processes in fuel rich gaseous hydrogen and gaseous oxygen. Flame thickness and reaction zone thickness calculations predict that the diameters of our test apparatus are below the quenching diameter of the propellants in sub-atmospheric tests. Temperature and pressure rise in resistively heated platinum and palladium microtubes was used as an indication of exothermic reactions. Specific data on mass flow versus preheat temperature required to achieve ignition are presented. With a plug flow model, the experimental conditions were simulated with detailed gas-phase chemistry, thermodynamic properties, and surface kinetics. Computational results generally support the experimental findings, but suggest an experimental mapping of the exit temperature and composition is needed.

Schneider, Steven J.↗
NASA NTRS2003

An overview of the pointing control system for NASA's Space Infra-red Telescope Facility (SIRTF)

This paper discusses the pointing control system for NASA's Space Infra-Red Telescope Facility (SIRTF). After an introduction to the SIRTF mission and telescope, an overview is given of the pointing control hardware, pointing control architecture, pointing requirements and capability, attitude constraints and commanding, attitude observers and required calibrations.

pointing control attitude estimation calibration↗
NASA NTRS2009

Extended Kalman Filter for MMS State Estimation

The Magnetospheric MultiScale Mission is a four spacecraft formation flying mission designed to study the Earth s magnetosphere. The spacecraft fly in highly elliptical orbits, forming a tetrahedron at apogee. Each spacecraft spins at 3 RPM and is equipped with a star scanner, slit sun sensor, and accelerometer. The purpose of this work is to develop an Extended Kalman Filter to simultaneously estimate the attitude, angular velocity, angular acceleration, and center of mass of each spacecraft.

Markley, F. Landis↗
NASA NTRS2004

Bayesian-Augmented Identification of Stars in a Narrow View

An algorithm for the identification of stars from a charge-coupled-device (CCD) image of a star field has been extended for use with narrower field-of-view images. Previously, the algorithm had been shown to be effective at a field of view of 8 degrees. This work augments the earlier algorithm using Bayesian decision theory. The new algorithm is shown to be capable of effective star identification down to a field of view of 2 degrees. The algorithm was developed for use in estimating the attitude of a spacecraft and could be used on Earth to help in the identification of stars and other celestial objects for astronomical observations.

Clouse, Daniel↗
NASA NTRS2025

Pose Estimation in the Geometric Algebra G (3, 0, 1)

Using techniques from Geometric Algebra, methods for solving attitude determination problems are extended to the problem of simultaneously estimating attitude and location, i.e. pose, from direct or indirect measurements of geometric objects including points, lines, and planes. Wahba’s problem is generalized for the solution of initial pose. The multiplicative extended Kalman filter is generalized for the maintenance of pose solutions. Farrankopf’s model for biases in rotation rate sensors is generalized to biases in twist rate measurements. A multiplicative extended Kalman filter for pose in the presence of such biases is developed.

Sensor Bias Modeling↗
NASA NTRS1988

Feasibility of using GPS measurements for OMV attitude update

This paper presents the results of a feasibility study to determine whether the measurements from GPS satellites can be used to estimate the OMV attitude to an accuracy comparable to the onboard sun sensors, i.e., better than 0.5 degree in each axis. The results documented in this paper demonstrate that OMV attitude can be estimated to an accuracy of 0.1 - 0.5 degree in each axis by processing GPS measurements in an onboard integrated, 17-state GPS/inertial navigation filter. The result is particularly significant for missions of short duration burns where accurate attitude information is needed to minimize guidance and control errors. It is shown that the GPS attitude technique described in this paper can be easily implemented in the current OMV navigation system design. Results reported in this paper are expected to support the goal of developing a fault-tolerant guidance, navigation, and control system offering an improved total navigation performance for the OMV.

Upadhyay, Triveni N.↗
NASA NTRS1975

Advanced Earth Observation System Instrumentation Study (aeosis)

Various types of measurements were studied for estimating the orbit and/or attitude of an Earth Observation Satellite. An investigation was made into the use of known ground targets in the earth sensor imagery, in combination with onboard star sightings and/or range and range rate measurements by ground tracking stations or tracking satellites (TDRSS), to estimate satellite attitude, orbital ephemeris, and gyro bias drift. Generalized measurement equations were derived for star measurements with a particular type of star tracker, and for landmark measurements with a multispectral scanner being proposed for an advanced Earth Observation Satellite. The use of infra-red horizon measurements to estimate the attitude and gyro bias drift of a geosynchronous satellite was explored.

White, R.↗
NASA NTRS2000

Chandra X-Ray Observatory Pointing Control System Performance During Transfer Orbit and Initial On-Orbit Operations

The Chandra X-ray Observatory (CXO, formerly AXAF) is the third of the four NASA great observatories. It was launched from Kennedy Space Flight Center on 23 July 1999 aboard the Space Shuttle Columbia and was successfully inserted in a 330 x 72,000 km orbit by the Inertial Upper Stage (IUS). Through a series of five Integral Propulsion System burns, CXO was placed in a 10,000 x 139,000 km orbit. After initial on-orbit checkout, Chandra's first light images were unveiled to the public on 26 August, 1999. The CXO Pointing Control and Aspect Determination (PCAD) subsystem is designed to perform attitude control and determination functions in support of transfer orbit operations and on-orbit science mission. After a brief description of the PCAD subsystem, the paper highlights the PCAD activities during the transfer orbit and initial on-orbit operations. These activities include: CXO/IUS separation, attitude and gyro bias estimation with earth sensor and sun sensor, attitude control and disturbance torque estimation for delta-v burns, momentum build-up due to gravity gradient and solar pressure, momentum unloading with thrusters, attitude initialization with star measurements, gyro alignment calibration, maneuvering and transition to normal pointing, and PCAD pointing and stability performance.

Quast, Peter↗
NASA NTRS1993

Quaternion normalization in spacecraft attitude determination

Attitude determination of spacecraft usually utilizes vector measurements such as Sun, center of Earth, star, and magnetic field direction to update the quaternion which determines the spacecraft orientation with respect to some reference coordinates in the three dimensional space. These measurements are usually processed by an extended Kalman filter (EKF) which yields an estimate of the attitude quaternion. Two EKF versions for quaternion estimation were presented in the literature; namely, the multiplicative EKF (MEKF) and the additive EKF (AEKF). In the multiplicative EKF, it is assumed that the error between the correct quaternion and its a-priori estimate is, by itself, a quaternion that represents the rotation necessary to bring the attitude which corresponds to the a-priori estimate of the quaternion into coincidence with the correct attitude. The EKF basically estimates this quotient quaternion and then the updated quaternion estimate is obtained by the product of the a-priori quaternion estimate and the estimate of the difference quaternion. In the additive EKF, it is assumed that the error between the a-priori quaternion estimate and the correct one is an algebraic difference between two four-tuple elements and thus the EKF is set to estimate this difference. The updated quaternion is then computed by adding the estimate of the difference to the a-priori quaternion estimate. If the quaternion estimate converges to the correct quaternion, then, naturally, the quaternion estimate has unity norm. This fact was utilized in the past to obtain superior filter performance by applying normalization to the filter measurement update of the quaternion. It was observed for the AEKF that when the attitude changed very slowly between measurements, normalization merely resulted in a faster convergence; however, when the attitude changed considerably between measurements, without filter tuning or normalization, the quaternion estimate diverged. However, when the quaternion estimate was normalized, the estimate converged faster and to a lower error than with tuning only. In last years, symposium we presented three new AEKF normalization techniques and we compared them to the brute force method presented in the literature. The present paper presents the issue of normalization of the MEKF and examines several MEKF normalization techniques.

Deutschmann, J.↗
NASA NTRS1983

The Integrated Sensor System Data Enhancement Package

The purpose of the Integrated Sensor System (ISS) Data Enhancement Package (DEP) is to improve the accuracies of the data obtained from the inflight tests performed on aircraft. The DEP is a microprocessor-based, flight-qualified electronics package that assimilates data from a Ring Laser Gyro (RGL) system, a standard NASA air data package, and other inputs. The DEP then processes these inputs in real-time to obtain optimal estimates of the aircraft velocity, attitude, and altitude. These estimates can be passed to the flight crew, downlinked, and/or stored on a mass storage medium. The DEP is now being built for the NASA Dryden Flight Research Center. Completion is anticipated in early 1984. A primary use of the ISS/DEP will be for the collection of quality data for the estimation of aircraft aerodynamic coefficients, including stability derivatives, using system identification methods. Initial anticipated applications will be on the AV-8B, F-14, and X-29 test aircraft.

Trankle, T. L.↗
NASA NTRS2021

Analysis of Mars 2020 Perseverance Entry, Descent, and Landing Attitude Initialization Performance

The Mars 2020 Perseverance rover successfully landed at Jezero Crater on February 18, 2021. To perform this feat, the entry, descent, and landing navigation filter must be initialized with an estimate of the spacecraft state, which includes attitude from the cruise attitude control subsystem. Accuracy of this initial attitude estimate has an impact on important metrics such as touchdown velocity. This paper presents a post-landing reconstruction of the attitude initialization error budget, showing that the requirement of 0.15 deg ($3\sigma$, per axis) was met. Each error source is described and analyzed, including flight telemetry where possible. Analysis of the error budget shows that it is driven by systematic sun sensor errors, star-scanner-to-sun-sensor alignment stability, and inertial-measurement-unit-to-sun-sensor alignment stability. Finally, a contingency plan to initialize the navigation filter in the event of a star scanner failure is presented. While this plan was not needed in flight, results indicate that the attitude initialization requirement could have still been met in this off-nominal scenario.

Lo, Kevin D.↗
NASA NTRS1992

A visual display aid for planning rover traversals

An interactive graphical planning system has been developed, which allows a human operator to design and check traversals (cross-country paths) for a planetary rover vehicle. The display provides the operator with necessary information about the terrain and indicates violations of operational or dynamic constraints on the rover. The operator can select different kinds of two-dimensional maps as well as a perspective view of the rover environment to plan the traversals. An experiment has been carried out to determine the ability of the operator to estimate the rover attitude in a large variety of situations. It turned out that the estimation error is highly dependent on the rover attitude itself. This result can be used to determine a vertical scale for the perspective representation of the terrain which avoids an underestimation of dangerous rover attitudes.

Bernard, Herbert F.↗
NASA NTRS2003

Multiplicative Versus Additive Filtering for Spacecraft Attitude Determination

The absence of a globally nonsingular three-parameter representation of rotations forces attitude Kalman filters to estimate either a singular or a redundant attitude representation. We compare two filtering strategies using simplified kinematics and measurement models. Our favored strategy estimates a three-parameter representation of attitude deviations from a reference attitude specified by a higher- dimensional nonsingular parameterization. The deviations from the reference are assumed to be small enough to avoid any singularity or discontinuity of the three-dimensional parameterization. We point out some disadvantages of the other strategy, which directly estimates the four-parameter quaternion representation.

Markley, F. Landis↗
NASA NTRS2005

ATTITUDE FILTERING ON SO(3)

A new method is presented for the simultaneous estimation of the attitude of a spacecraft and an N-vector of bias parameters. This method uses a probability distribution function defined on the Cartesian product of SO(3), the group of rotation matrices, and the Euclidean space W N .The Fokker-Planck equation propagates the probability distribution function between measurements, and Bayes s formula incorporates measurement update information. This approach avoids all the issues of singular attitude representations or singular covariance matrices encountered in extended Kalman filters. In addition, the filter has a consistent initialization for a completely unknown initial attitude, owing to the fact that SO(3) is a compact space.

Markley, F. Landis↗
NASA NTRS2018

Mission Summary of Cassini Spacecraft Guidance and Control Hardware Health and Performance

The Cassini-Huygens mission ended on September 15, 2017, after nearly two decades in ight. The well-designed Cassini spacecraft had robust hardware that permitted two extended missions, lasting nine years longer than the expected prime mission. At the end of the mission, the Attitude and Articulation Control Subsystem (AACS) was using two pieces of redundant back-up of hardware, one reaction wheel and the hydrazine thruster branch, due to hardware anomalies earlier in the mission. The back-up hardware performed nominally through the rest of mission. The prime reaction wheels at the end of the mission had reached more than 130% of the consumable limit for number of revolutions. No thruster on either thruster branch accumulated more than 45% of the consumable limits. The inertial reference unit slightly exceeded the pre-launch requirements on bias error, but as the software continuously estimated this value in ight, the attitude estimation was not adversely a ected. The star trackers performed nominally, and though there was a spacecraft anomaly in 1998 related to the star trackers, the origin was not in hardware itself. The Sun sensors and accelerometer both performed as expected and met all requirements throughout the mission. Ultimately, the lifetime of the Cassini spacecraft was not limited by hardware performance. Planetary protection requirements necessitated the end of the mission as the spacecraft's propellant reserves depleted, and Cassini plunged into Saturn's atmosphere with a healthy attitude control system.

Stupik, Joan↗
NASA NTRS1975

Use of known landmarks for satellite navigation

The use of known landmark measurements along with an onboard attitude reference system for estimation of both the ephemeris and attitude of an earth orbiting satellite is discussed. Simulation results are presented which show that, although known landmarks alone do not provide sufficient information to bound the uncertainty in attitude and ephemeris, they are useful in combination with other navigation aids, such as star tracking, ground tracking, or satellite tracking. An extended Kalman filter is used for optimal estimation. The result of using a smoothing algorithm is also shown.

White, R. L.↗
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 NTRS2005

Hubble Space Telescope Angular Velocity Estimation During the Robotic Servicing Mission

In 2004 NASA began investigation of a robotic servicing mission for the Hubble Space Telescope (HST). Such a mission would require estimates of the HST attitude and rates in order to achieve a capture by the proposed Hubble robotic vehicle (HRV). HRV was to be equipped with vision-based sensors, capable of estimating the relative attitude between HST and HRV. The inertial HST attitude is derived from the measured relative attitude and the HRV computed inertial attitude. However, the relative rate between HST and HRV cannot be measured directly. Therefore, the HST rate with respect to inertial space is not known. Two approaches are developed to estimate the HST rates. Both methods utilize the measured relative attitude and the HRV inertial attitude and rates. First, a nonlinear estimator is developed. The nonlinear approach estimates the HST rate through an estimation of the inertial angular momentum. The development includes an analysis of the estimator stability given errors in the measured attitude. Second, a linearized approach is developed. The linearized approach is a pseudo-linear Kalman filter. Simulation test results for both methods are given, including scenarios with erroneous measured attitudes. Even though the development began as an application for the HST robotic servicing mission, the methods presented are applicable to any rendezvous/capture mission involving a non-cooperative target spacecraft.

Thienel, Julie K.↗