E. Glenn Lightsey

Efficient and optimal attitude determination using recursive global positioning system signal operations

In this paper, a new and efficient algorithm is developed for attitude determination from Global Positioning System signals. The new algorithm is derived from a generalized nonlinear predictive filter for nonlinear systems. This uses a one time-step ahead approach to propagate a simple kinematics model for attitude determination. The advantages of the new algorithm over previously developed methods include: it provides optimal attitudes even for coplanar baseline configurations; it guarantees convergence even for poor initial conditions; it is a non-iterative algorithm; and it is computationally efficient. These advantages clearly make the new algorithm well suited to on-board applications. The performance of the new algorithm is tested on a dynamic hardware simulator. Results indicate that the new algorithm accurately estimates the attitude of a moving vehicle, and provides robust attitude estimates even when other methods, such as a linearized least-squares approach, fail due to poor initial starting conditions.

Global Positioning System Integer Ambiguity Resolution Without Attitude Knowledge

Inthispaper,anewmotion-basedalgorithmforglobalpositioningsystemintegerambiguityresolutionisderived. Thealgorithmrepresentstheglobalpositioningsystem sightlinevectorsinthebodyframeasthesumoftwovectors, onedependingonthephasemeasurementsandtheotherontheunknownintegers.Thevectorcontainingtheinteger phases is found using a procedure developed to solve for magnetometer biases. In addition to a batch solution, this paper also provides a sequential estimate, so that a suitable stopping condition can be found during the vehicle motion. The newalgorithm has several advantages: it doesnot requirean a prioriestimateofthevehicle’ s attitude; it provides an inherent integrity check using a covariance-type expression; and it can sequentially estimate the ambiguitiesduring thevehiclemotion. Itsonly disadvantageisthatit requiresatleastthreenoncoplanar baselines. The performance of the new algorithm is tested on a dynamic hardware simulator.

Real-Time Navigation for Mars Missions Using the Mars Network

A NASA Mars technology program task is developing a prototype, embedded, real-time navigation system for Mars final approach and entry, descent, and landing using the Mars Network’s Electra ultrahigh frequency transceiver. The Mars Network is ideally situated to provide spacecraft-to-spacecraft navigation via the Electra ultrahigh frequency transceiver, which is a versatile telecommunications payload that is capable of providing autonomous on-orbit, real-time trajectory determination using two-way Doppler measurements between a Mars approach vehicle and a Mars Network orbiter. A set of analyses based on the 2010 encounter at Mars between the MarsScienceLaboratory and theMarsReconnaissanceOrbiter demonstrate that the navigation system is capable of achieving a 300m or better atmosphere entry knowledge error and that the resulting technology is a key component to enabling pinpoint landing. The development approach, software design, and test results from an engineering development unit are presented.

A real-time kinematic GPS sensor for spacecraft relative navigation

The concept and prototype implementation of a spaceborne relative navigation sensor based on a pair of GPS receivers is presented. It employs two individual receivers exchanging raw measurements via a dedicated serial data link. Besides computing their own navigation solution, the receivers process single difference measurements to obtain their mutual relative state. The differential processing provides a high level of common error cancellation while the resulting noise is minimized by appropriate use of carrier phase measurements. A prototype relative navigation sensor making use of the above concepts has been built up based on the GPS Orion 12 channel L1 receiver and qualified in hardware-in-the-loop tests using a GPS signal simulator. It provides a relative navigation solution with representative r.m.s. accuracies of 0.5 m and 1 cm/s, respectively, for position and velocity. For ease of use the relative state is provided in a co-moving frame aligned with the radial, cross-track and along-track direction. The purely kinematic nature of state estimation and the small latency make the system well suited for maneuvering spacecraft, while the minimalist hardware requirements facilitate its use on microsatellite formations.  2002 Editions scientifiques et medicales Elsevier SAS. All rights reserved. Zusammenfassung Die vorliegende Arbeit stellt das Konzept und die Prototypimplementierung eines raumgestutzten Relativnavigationssensors vor, der auf einem Paar von GPS Empfangern basiert. Es werden zwei separate Empfanger eingesetzt, die Rohmessungen uber eine dedizierte serielle Schnittstelle austauschen. Neben der Berechnung der eigenen Navigationslosung prozessieren die Empfanger Einfachdifferenzen dieser Messungen, um gegenseitig ihre Relativposition und Geschwindigkeit zu bestimmen. Die differentielle Prozessierung gewahrleistet dabei in hohem Mas die Ausloschung gemeinsamer Fehleranteile, wahrend der Rauschanteil durch die geeignete Nutzung von Tragerphasenmessungen minimiert wird. Der Prototyp eines derartigen Relativnavigationssensors wurde auf Basis eines GPS Orion 12 Kanal L1 Einfrequenzempfangers aufgebaut und mit Hilfe eines GPS Signalssimulators getestet. Er liefert eine Relativnavigationslosung mit reprasentativen Genauigkeiten von 0.5 m und 1 cm/s fur Position und Geschwindigkeit. Zum einfachen Gebrauch wird die Relativnavigationslosung in einem mitbewegten Koordinatensystem dargestellt, das entlang der radialen Richtung, der Vorwartsrichtung und der Bahnnormalen orientiert ist. Die rein kinematische Natur der Zustandsschatzung und die geringe Latenzzeit machen das System gut geeignet fur Raumfahrzeuge mit aktivem Antrieb. Zusatzlich erleichtern die minimalistischen Hardwarevoraussetzungen den Einsatz in Formationen von Mikrosatelliten.  2002 Editions scientifiques et medicales Elsevier SAS. All rights reserved.

Robust spacecraft attitude determination using global positioning system receivers

This work presents the development of a new attitude determination system based on global positioning system signals. This new algorithm utilizes signal-to-noise-ratio measurements from canted antennas to produce three-axis attitude solutions. These solutions are then used to determine the integer ambiguities for double-difference carrier-phase measurements. The two measurement types are processed through an extended Kalman filter to produce attitude solutions. The algorithm is tested using both hardware-in-the-loop orbit simulations and static rooftop data. These experiments demonstrate the speed of the integer resolution process and the accuracy of the resulting carrier-phase attitude solutions. Further tests demonstrate that the canting of the antennas has little or no effect on the double-difference carrier-phase measurements for antenna baselines of 1 m. The algorithm is modified to include magnetometer measurements. Simulation demonstrates how this addition further improves the integer resolution. The results of the rooftop and simulation experiments demonstrate that this new algorithm can quickly and accurately resolve the integer ambiguities associated with carrier-phase attitude determination. Once the integers are resolved, the resulting algorithm generates solutions of less than 0.5 deg in accuracy. To conclude this work, some possible future work is discussed.

Three-axis attitude determination using global positioning system signal strength measurements

An alternative method to the standard carrier-phase algorithm for deriving three-axis attitude solutions from global positioning system (GPS) signal sources is investigated. This new method uses signal-to-noise ratio (SNR) measurements from two or more canted antennas and a knowledge of each receiving antennas gain pattern to generate pointing vector solutions. These vector solutions are then converted to a three-axis attitude solution. The method has the advantage of requiring no complicated initializing procedures such as integer ambiguity resolution, and it can generate three-axis solutions with as few as two antennas. A solution is produced whenever a minimum of three GPS satellites are in view, regardless of the vehicles orientation at that time or at any time previously. The performance of this SNR approach is investigated using a Kalman filter to derive solutions on a satellite whose attitude generally remains fixed in the local frame. The performance of the algorithm is evaluated while the canting angles between the antennas and the filter values are varied. In addition, self-calibrating and self-scaling options are explored to reduce the algorithms dependence on any outside knowledge of the vehicles attitude during an in-flight calibration process. The study also examines the performance of the solution under the expected error conditions of an inaccurate calibration and sky blockage. The results demonstrate that the estimate is relatively insensitive to these expected errors. Some general conclusions are drawn about the performance and sensitivity of the developed algorithm.

Discretized constrained attitude pathfinding and control for satellites

A general purpose guidance, navigation, and control algorithm is developed for satellites with three degree-of-freedom rotation maneuverability. The algorithm is capable of meeting multiple pointing constraints autonomously by using a combination of constrained attitude pathfinding and control elements. The unit sphere is discretized into a graph using an icosahedron-based pixelization subroutine. An admissible path between attitude keep-out zones is found with the A* pathfinding algorithm. The trajectory is followed using a rate and torque constrained quaternion feedback controller. The algorithm is capable of running in real time on a low-power flight computer. An embedded version of the algorithm has secured flight opportunities on two student-built three-unit CubeSats. Both sets of mission requirements are satisfied with the same three-unit CubeSat attitude control system, demonstrating the algorithm’s versatility as a general purpose controller. The autonomy provided by the constrained control algori...

INSPIRE: Interplanetary NanoSpacecraft Pathfinder in Relevant Environment

The INSPIRE project would demonstrate the revolutionary capability of deep space CubeSats by placing two nanospacecraft in Earth-escape orbit. Prior to any inclusion on larger planetary missions, CubeSats must demonstrate that they can operate, communicate, and be navigated far from Earth – these are the primary objectives of INSPIRE. Spacecraft components, such as a JPL X-band radio and a robust watchdog system, would provide the basis for future high-capability, lower-cost-risk missions beyond Earth. These components should enable future supplemental science and educational opportunities at many destinations. The nominal INSPIRE mission would last for three months and achieve an expected Earth-probe distance of 1.5x10 km (dependent upon escape velocity as neither spacecraft will have propulsion capability). The project would monitor onboard telemetry; operate, communicate, and navigate with both spacecraft; demonstrate cross-link communications; and demonstrate science utility with an onboard magnetometer and imager. Lessons learned from this pathfinder mission should help to inform future interplanetary NanoSpacecraft and larger missions that might use NanoSpacecraft components.

Expected EDL navigation performance with spacecraft to spacecraft radiometric data

Pinpoint landing (defined for the purpose of this discussion as landing within 1km of a preselected target) is a key Advanced Entry, Descent and Landing (EDL) technology for future Mars landers. Key scientific goals for Mars exploration, such as the search for water and characterization of aqueous processes on Mars, the study of mineralogy and weathering of the Martian surface and the search for preserved biosignatures in Martian rocks, require placing landers at pre-defined locations of greatest scientific interest. The capability to land within 1 km of a pre-defined landing site will improve safety and enable landing within roving range of sites of scientific interest while avoiding hazardous areas. A critical component of the closed-loop guidance, navigation and control (GN&C) system required for pinpoint landing is position and velocity estimation in real time. Spacecraftto-spacecraft navigation will take advantage of the UHF link between two spacecraft (i.e. to an orbiter from an approaching lander for EDL telemetry relay) to build radiometric data, specifically the total count carrier phase of the Doppler shifted 2-Way coherent UHF signal, that are processed to determine position and velocity in real time. The improved onboard state knowledge provided by spacecraft-to-spacecraft navigation will reduce the landed position error and improve the performance of entry guidance. Results from the first of two years planned for this effort are documented here, including selection and documentation of prototype algorithms that will go forward into flight code along with analysis results used to define the algorithm set.

GPS/INS Kalman Filter Design for Spacecraft Operating in the Proximity of the International Space Station

The next generation reusable launch vehicle will operate in flve ∞ight phases: ascent, on-orbit, proximity operations, re-entry and landing. Navigation during each of these ∞ight phases presents unique challenges. The Space Shuttle addresses these challenges by the use of a number of navigation sensors. However, an integrated GPS/INS navigation system may be able to meet the navigation requirements of all ∞ight phases. Integrated GPS/INS systems have been built and demonstrated for the ascent, re-entry and landing phases and their performance is known. However, the same cannot be said for integrated GPS/INS systems for the on-orbit and proximity operations ∞ight phases. Therefore, this research examines the performance of GPS/INS navigation during a rendezvous with the ISS. Error models for INS and GPS navigation sensors operating in the vicinity of the ISS have been developed. The GPS error model includes the efiects of GPS signal blockage and multipath near the ISS. These error models have been used to develop an integrated GPS/INS extended Kalman fllter. A simulation of the fllter has been developed and the flrst test case shows position errors of less than 1 meter and velocity errors of less than 0.02 m/s in each axis. The second test case shows the position errors grow to approximately 47.5 meters and the velocity errors grow to 0.13 m/s during a ten minute GPS outage. Finally, the third test case shows that the GPS/INS EKF performance is not signiflcantly degraded by GPS signal blockage due to the ISS during a simulated rendezvous with the ISS.