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NASA NTRS1980

Accuracy of aircraft velocities from inertial navigation systems for application to airborne wind measurements

An experimental assessment was made of two commercially available inertial navigation systems (INS) with regard to their velocity measuring capability for use in wind, shear, and long-wavelength atmospheric turbulence research. The assessment was based on 52 sets of postflight measurements of velocity (error) during a "Schuler cycle" (84 minutes) while the INS was still operating but the airplane was motionless. Four INS units of one type and two units of another were tested over a period of 2 years after routine research flights similar to air-linetype operations of from 1 to 6 hours duration. The maximum postflight errors found for the 52 cases had a root mean square value of 2.82 m/sec with little or no correlation of error magnitude with flight duration. Using an INS for monitoring ground speed during landway in a predicted high wind shear situation could lead to landing speeds which are dangerously high or low.

Rhyne, R. H.↗

NASA NTRS1997

Development of a Closed-Loop Strap Down Attitude System for an Ultrahigh Altitude Flight Experiment

A low-cost attitude system has been developed for an ultrahigh altitude flight experiment. The experiment uses a remotely piloted sailplane, with the wings modified for flight at altitudes greater than 100,000 ft. Mission requirements deem it necessary to measure the aircraft pitch and bank angles with accuracy better than 1.0 deg and heading with accuracy better than 5.0 deg. Vehicle cost restrictions and gross weight limits make installing a commercial inertial navigation system unfeasible. Instead, a low-cost attitude system was developed using strap down components. Monte Carlo analyses verified that two vector measurements, magnetic field and velocity, are required to completely stabilize the error equations. In the estimating algorithm, body-axis observations of the airspeed vector and the magnetic field are compared against the inertial velocity vector and a magnetic-field reference model. Residuals are fed back to stabilize integration of rate gyros. The effectiveness of the estimating algorithm was demonstrated using data from the NASA Dryden Flight Research Center Systems Research Aircraft (SRA) flight tests. The algorithm was applied with good results to a maximum 10' pitch and bank angles. Effects of wind shears were evaluated and, for most cases, can be safely ignored.

Whitmore, Stephen A.↗

NASA NTRS1999

1999 Flight Mechanics Symposium

This conference publication includes papers and abstracts presented at the Flight Mechanics Symposium held on May 18-20, 1999. Sponsored by the Guidance, Navigation and Control Center of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

John P. Lynch↗

NASA NTRS2011

Docking Offset Between the Space Shuttle and the International Space Station and Resulting Impacts to the Transfer of Attitude Reference and Control

The Space Shuttle does not dock at an exact 90 degrees to the International Space Station (ISS) x-body axis. This offset from 90 degrees, along with error sources within their respective attitude knowledge, causes the two vehicles to never completely agree on their attitude, even though they operate as a single, mated stack while docked. The docking offset can be measured in flight when both vehicles have good attitude reference and is a critical component in calculations to transfer attitude reference from one vehicle to another. This paper will describe how the docking offset and attitude reference errors between both vehicles are measured and how this information would be used to recover Shuttle attitude reference from ISS in the event of multiple failures. During STS-117, ISS on-board Guidance, Navigation and Control (GNC) computers began having problems and after several continuous restarts, the systems failed. The failure took the ability for ISS to maintain attitude knowledge. This paper will also demonstrate how with knowledge of the docking offset, the contingency procedure to recover Shuttle attitude reference from ISS was reversed in order to provide ISS an attitude reference from Shuttle. Finally, this paper will show how knowledge of the docking offset can be used to speed up attitude control handovers from Shuttle to ISS momentum management. By taking into account the docking offset, Shuttle can be commanded to hold a more precise attitude which better agrees with the ISS commanded attitude such that start up transients with the ISS momentum management controllers are reduced. By reducing start-up transients, attitude control can be transferred from Shuttle to ISS without the use of ISS thrusters saving precious on-board propellant, crew time and minimizing loads placed upon the mated stack.

Helms, W. Jason↗

NASA NTRS2010

Prospects for UT1 Measurements from VLBI Intensive Sessions

Very Long Baseline Interferometry (VLBI) Intensives are one-hour single baseline sessions to provide Universal Time (UT1) in near real-time up to a delay of three days if a site is not e-transferring the observational data. Due to the importance of UT1 estimates for the prediction of Earth orientation parameters, as well as any kind of navigation on Earth or in space, there is not only the need to improve the timeliness of the results but also their accuracy. We identify the asymmetry of the tropospheric delays as the major error source, and we provide two strategies to improve the results, in particular of those Intensives which include the station Tsukuba in Japan with its large tropospheric variation. We find an improvement when (1) using ray-traced delays from a numerical weather model, and (2) when estimating tropospheric gradients within the analysis of Intensive sessions. The improvement is shown in terms of reduction of rms of length-of-day estimates w.r.t. those derived from Global Positioning System observations

Boehm, Johannes↗

NASA NTRS2013

Mars Science Laboratory Entry Descent and Landing Simulation Using DSENDS

The most recent planetary science mission to Mars is Mars Science Laboratory (MSL) with the Curiosity rover, launched November 26, 2011 and landed at Gale Crater on August 6, 2012. This spacecraft was the first use at Mars of a complete closed-loop Guidance Navigation and Control (GN&C) system, including guided entry with a lifting body that greatly reduces dispersions during the Entry, Descent and Landing (EDL) phase to achieve a 25 km x 20 km landing error relative to the selected Gale Crater landing target. In order to confirm meeting the above landing criteria, high-fidelity simulation of the EDL phase is required. The tool used for 6DOF EDL trajectory verification analysis is Dynamics Simulator for Entry, Descent and Surface landing (DSENDS), which is a high-fidelity simulation tool from JPLs Dynamics and Real-Time Simulation Laboratory for the development, test and operations of aero-flight vehicles. DSENDS inherent capability is augmented for MSL with project-specific models of atmosphere, aerodynamics, sensors and thrusters along with GN&C flight software to enable high-fidelity trajectory simulation. This paper will present the model integration and independent verification experience of the JPL EDL trajectory analysis team.

Mars↗

NASA NTRS2013

Mars Science Laboratory Entry Descent and Landing Simulation Using DSENDS

The most recent planetary science mission to Mars was Mars Science Laboratory (MSL) with the Curiosity rover, launched November 26, 2011 and landed at Gale Crater on August 6, 2012. This spacecraft was the first use at Mars of a complete closed-loop Guidance Navigation and Control (GN&C) system, including guided entry with a lifting body that greatly reduces dispersions during the Entry, Descent and Landing (EDL) phase to achieve a 25 km X 20 km landing error relative to the selected Gale Crater landing target. In order to confirm meeting the above landing criteria, high-fidelity simulation of the EDL phase is required. The tool used for 6DOF EDL trajectory verification analysis is Dynamics Simulator for Entry, Descent and Surface landing (DSENDS), which is a high-fidelity simulation tool from JPLs Dynamics and Real-Time Simulation Laboratory for the development, test and operations of aero-flight vehicles. DSENDS inherent capability is augmented for MSL with project-specific models of atmosphere, aerodynamics, sensors and thrusters along with GN&C flight software to enable high-fidelity trajectory simulation. This paper will present the model integration and independent verification experience of the JPL EDL trajectory analysis team.

Burkhart, P. Daniel↗

NASA NTRS2015

Development of the Resource Prospector Planetary Rover

The Resource Prospector (RP) is an In-‐Situ Resource Utilization (ISRU) lunar rover mission under study by NASA. RP is planned to launch in 2020 to prospect for subsurface volatiles and to extract oxygen from lunar regolith. The mission will address several of NASA's "Strategic Knowledge Gaps" for lunar exploration. The mission will also address the Global Exploration Roadmap's strategic goal of using local resources for human exploration. The distribution of lunar subsurface volatiles drives the mission requirement for mobility. The spatial distribution is hypothesized to be governed by impact cratering with the top 0.5 m being patchy at scales of 100 m. The mixing time scale increases with depth (less frequent larger impacts). Consequently, increased mobility reduces the depth requirement for sampling. The target RP traverse will extend 1 km radially from the landing site to sample craters of varying sizes. Sampling craters with different ages will reveal possible volatile emplacement history. In 1 Ga, approximately 60-70 craters of 10 m diameter form per km2. Thus, the rover will need to sample at least ten of these craters, which may require a total traverse path length of 2-‐3 km. During 2014-2015, we developed an initial prototype rover for RP. The current design is a solar powered, four-wheeled vehicle, with hub motor drive, offset four wheel steering, and active suspension. Active suspension provides capabilities including changing vehicle ride height, traversing comparatively large obstacles, and controlling load on the wheels. All-wheel steering enables the vehicle to point arbitrarily while roving, e.g., to keep the solar array pointed at the sun while in motion. The offset steering combined with active suspension improves driving in soft soil. The rover's on-board software utilizes NASA's Core Flight Software, which is a reusable flight software environment. During 2015, we completed the initial rover software build, which provides low-level hardware interfaces, basic mobility control, waypoint driving, odometry, basic error checking, and camera services. Development of the prototype rover has enabled maturation of many of the subsystems to TRL 5. During the next year, we will conduct integrated testing of concepts of operation, navigation, and remote driving tools. In addition, we will perform environmental tests including radiation (avionics), thermal and thermal/vacuum (mechanisms), and gravity offload (mobility).

robotics↗

NASA NTRSDate not supplied

Near Rectilinear Halo Orbit Determination with Simulated DSN Observations

This paper presents the results of a high-fidelity simulation of spacecraft orbit determination in a near rectilinear halo orbit (NRHO). Others in the literature have examined this problem with linear covariance analysis, but the highly-nonlinear dynamics of this orbit challenge the assumptions underlying such analyses. The present work builds on similar analysis performed by other authors to contribute a fuller understanding of the operational requirements for NRHO navigation. The present work serves as a check to the assumptions of previous studies and an independent verification of those results. The results from the literature are extended by quantifying the space of orbital states from which a spacecraft with given control authority can safely return to the nominal path. Spacecraft state uncertainty estimates are evaluated as a function of time. Simulated range and range-rate measurements with the Deep Space Network (DSN) ground stations are used to model orbit determination accuracy. Orbit maintenance maneuvers are performed using both short-horizon and long-horizon stationkeeping targeting. Monte Carlo analysis of orbit determination and stationkeeping is performed. This paper quantifies the achievable state uncertainty with deep space network (DSN)-only range and range-rate observations. This paper also addresses requirements on the frequency of DSN observation periods and correlates ground contact frequency with navigation accuracy. The results of several related studies are presented and discussed: the effect of missing ground station passes, the effect of missing stationkeeping maneuvers, the sensitivity of the spacecraft state estimate to realistic error sources, and stationkeeping propellant budget.

Nathan L Parrish↗

NASA NTRS1992

Application of high-precision two-way ranging to Galileo Earth-1 encounter navigation

The application of precision two-way ranging to orbit determination with relatively short data arcs is investigated for the Galileo spacecraft's approach to its first Earth encounter (December 8, 1990). Analysis of previous S-band (2.3-GHz) ranging data acquired from Galileo indicated that under good signal conditions submeter precision and 10-m ranging accuracy were achieved. It is shown that ranging data of sufficient accuracy, when acquired from multiple stations, can sense the geocentric angular position of a distant spacecraft. A range data filtering technique, in which explicit modeling of range measurement bias parameters for each station pass is utilized, is shown to largely remove the systematic ground system calibration errors and transmission media effects from the Galileo range measurements, which would otherwise corrupt the angle-finding capabilities of the data. The accuracy of the Galileo orbit solutions obtained with S-band Doppler and precision ranging were found to be consistent with simple theoretical calculations, which predicted that angular accuracies of 0.26-0.34 microrad were achievable. In addition, the navigation accuracy achieved with precision ranging was marginally better than that obtained using delta-differenced one-way range (delta DOR), the principal data type that was previously used to obtain spacecraft angular position measurements operationally.

Pollmeier, V. M.↗

NASA NTRS1991

Autonomous reconfigurable GPS/INS navigation and pointing system for rendezvous and docking

This paper describes the results of an integrated navigation and pointing system software development effort sponsored by the NASA MSFC through a SBIR Phase 2 Program. The integrated Global Positioning System (GPS)/Inertial Navigation System (INS) implements an autonomous navigation filter that is reconfigurable in real-time to accommodate mission contingencies. An onboard expert system monitors the spacecraft status and reconfigures the navigation filter accordingly, to optimize the system performance. The navigation filter is a multi-mode Kalman filter to estimate the spacecraft position, velocity, and attitude. Three different GPS-based attitude determination techniques, namely, velocity vector matching, attitude vector matching, and interferometric processing, are implemented to encompass different mission contingencies. The integrated GPS/INS navigation filter will use any of these techniques depending on the mission phase and the state of the sensors. The first technique, velocity vector matching, uses the GPS velocity measurement to estimate the INS velocity errors and exploits the correlation between INS velocity and attitude errors to estimate the attitude. The second technique, attitude vector matching, uses INS gyro measurements and GPS carrier phase (integrated Doppler) measurements during a spacecraft rotation maneuver to determine the attitude. Both of these techniques require only one GPS antenna onboard to determine the spacecraft attitude. The third technique, interferometric processing, requires use of multiple GPS antennae. In order to determine 3-axis body attitude, three GPS antennae (2 no-coplanor baselines) are required.

Upadhyay, Triveni N.↗

NASA NTRS2004

In-Flight Assessment of a Pursuit Guidance Display Format for Manually Flown Precision Instrument Approaches

In-flight evaluations of a pursuit guidance display system for manually flown precision instrument approaches were performed. The guidance system was integrated into the RASCAL JUH-60A Black Hawk helicopter. The applicability of the pursuit guidance disp1aFs to the operation of Runway Independent Aircraft (RIA) is made evident because the displays allow the pilot to fly a complex, multi-segment, descending, decelerating approach trajectory. The complex trajectory chosen for this in-flight assessment began from a downwind abeam position at 110 knots and was hand-flown to a 50 ft decision altitude at 40 knots using a rate-command/attitude-hold plus turn-coordination control system. The elements of the pursuit guidance format displayed on a 10-inch liquid crystal display (LCD) flat panel consisted of a flightpath vector and a "leader" aircraft as the pursuit guidance element. Approach guidance was based primarily on carrier-phase differential Global Positioning System (GPS) navigation, and secondarily on both medium accuracy inertial navigation unit states and air data computer states. Required Navigation Performance (RNP) concepts were applied to the construction of display elements such as lateral/vertical deviation indicators and a tunnel that indicated to the pilot, in real-time, the performance with respect to RNP error bounds. The results of the flight evaluations of the guidance display show that precise path control for operating within tight RNP boundaries (RNP 0.007NM/24ft for initial approach, RNP 0.008NM/19ft for intermediate approach, and RNP 0.002NM/9ft for final approach) is attainable with minimal to moderate pilot workload.

Moralez, Ernesto, III↗

NASA NTRS2017

Precision Asteroid Astrometry

Among the methods used to guide spacecraft to their destinations, Optical Navigation (OpNav) remains an effective option. OpNav makes use of star fields and small body ephemerides to precisely locate spacecraft. To facilitate accurate OpNav, the small body ephemerides must be constantly updated; asteroid orbits accumulate errors each year of a few milliarcseconds. Through extended exposures, taken with strategic offsets, a least-squares solution can be found that determines updated ephemeris data. This updated data can also be used by occultation astronomers to gain further information about the small bodies, including their size and shape. Using the 24-inch telescope at the Caltech Table Mountain Observatory (TMO), we capture two or more 180 second exposures of each target. These images, combined with a file for the predicted background star field and two reference files, are then processed through a series of scripts and programs. Starting with a prediction file and two to five exposures of the asteroid, the data is processed. This original data is about 32MB per observation. Once the data are reduced to only Right Ascension and Declination for each target, the data are ready for delivery. This consists of text only, and for each target takes about 80 bytes; this resulting data reduction is about five orders of magnitude. This method produces observed positions that are refined by about 12 milliarcseconds, a refinement that is accomplished almost nowhere else. The occultation observations that are facilitated by the ephemerides being refined also produce results that are not possible in any other way from ground-based observations.

Dial, Jason↗

NASA NTRS2019

GOES I-M: A Retrospective Look at Image Navigation and Registration (INR), Jitter and Lessons Learned

The Geostationary Operational Environmental Satellite (GOES) I-M series of spacecraft was the second generation of United States meteorological observational platforms in geosynchronous orbit. They served as the principal Earth- viewing observational platforms for continuously monitoring dynamic weather events from the mid-1990s and into the 21st century. This paper will look back at the program framing key system attributes of the mission, which necessitated a multi-layered development approach to meet stringent meteorological instrument Line-of-Sight (LoS) pointing and pointing stability requirements. The overall approach involved understanding, correcting, and avoiding pointing errors across a broad frequency range including what would typically be called dynamic interaction and jitter. Background information will be provided covering the mission architecture and program drivers. The systems solution for man- aging and mitigating the deleterious influences of on-board disturbances in order to meet the challenging instrument LoS pointing and jitter requirements will be described, along with the ‘first of its kind’ Image Navigation and Registration system. A broad look back at the lessons learned that emerged from the GOES I- M experience will be presented, with the intent of capturing general and specific insights for developers of future missions having stringent payload instrument pointing requirements. These discussions will touch on such critical aspects as defining jitter and related pointing requirements, the importance of early system architectural decisions, understanding and reducing on-board disturbances, the balance of test and analysis, and the imperative for maximizing on-orbit operational flexibility in order to accommodate unexpected dynamic interactions.

Sudey, John↗

DOE OSTI2020

A Chip-Scale Oscillation-Mode Optomechanical Inertial Sensor Near the Thermodynamical Limits

Modern navigation systems integrate the global positioning system (GPS) with an inertial navigation system (INS), which complement each other for correct attitude and velocity determination. Here, the core of the INS integrates accelerometers and gyroscopes used to measure forces and angular rate in the vehicular inertial reference frame. With the help of gyroscopes and by integrating the acceleration to compute velocity and distance, precision and compact accelerometers with sufficient accuracy can provide small–error location determination. Solid–state implementations, through coherent readout, can provide a platform for high performance acceleration detection. In contrast to prior accelerometers using piezoelectric or capacitive readout techniques, optical readout provides narrow–linewidth high–sensitivity laser detection along with low–noise resonant optomechanical transduction near the thermodynamical limits. Here an optomechanical inertial sensor with an 8.2 µg Hz –1/2 velocity random walk (VRW) at an acquisition rate of 100 Hz and 50.9 µg bias instability is demonstrated, suitable for applications, such as, inertial navigation, inclination sensing, platform stabilization, and/or wearable device motion detection. Driven into optomechanical sustained–oscillation, the slot photonic crystal cavity provides radio–frequency readout of the optically–driven transduction with an enhanced 625 µg Hz –1 sensitivity. Measuring the optomechanically–stiffened oscillation shift, instead of the optical transmission shift, provides a 220× VRW enhancement over pre–oscillation mode detection.

77 NANOSCIENCE AND NANOTECHNOLOGY↗

NASA NTRS1978

Determination of navigation FDI thresholds using a Markov model

A method for determining time-varying Failure Detection and Identification (FDI) thresholds for single sample decision functions is described in the context of a triplex system of inertial platforms. A cost function consisting of the probability of vehicle loss due to FDI decision errors is minimized. A discrete Markov model is constructed from which this cost can be determined as a function of the decision thresholds employed to detect and identify the first and second failures. Optimal thresholds are determined through the use of parameter optimization techniques. The application of this approach to threshold determination is illustrated for the Space Shuttle's inertial measurement instruments.

Walker, B. K.↗

NASA NTRS2013

Doppler Lidar Sensor for Precision Navigation in GPS-Deprived Environment

Landing mission concepts that are being developed for exploration of solar system bodies are increasingly ambitious in their implementations and objectives. Most of these missions require accurate position and velocity data during their descent phase in order to ensure safe, soft landing at the pre-designated sites. Data from the vehicle's Inertial Measurement Unit will not be sufficient due to significant drift error after extended travel time in space. Therefore, an onboard sensor is required to provide the necessary data for landing in the GPS-deprived environment of space. For this reason, NASA Langley Research Center has been developing an advanced Doppler lidar sensor capable of providing accurate and reliable data suitable for operation in the highly constrained environment of space. The Doppler lidar transmits three laser beams in different directions toward the ground. The signal from each beam provides the platform velocity and range to the ground along the laser line-of-sight (LOS). The six LOS measurements are then combined in order to determine the three components of the vehicle velocity vector, and to accurately measure altitude and attitude angles relative to the local ground. These measurements are used by an autonomous Guidance, Navigation, and Control system to accurately navigate the vehicle from a few kilometers above the ground to the designated location and to execute a gentle touchdown. A prototype version of our lidar sensor has been completed for a closed-loop demonstration onboard a rocket-powered terrestrial free-flyer vehicle.

Amzajerdian, F.↗

NASA NTRS1973

Satellite orbit determination

A historic account of the activities of the Satellite OD Group during the MM'71 mission is given along with an assessment of the accuracy of the determined orbit of the Mariner 9 spacecraft. Preflight study results are reviewed, and the major error sources described. Tracking and data fitting strategy actually used in the real time operations is itemized, and Deep Space Network data available for orbit fitting during the mission and the auxiliary information used by the navigation team are described. A detailed orbit fitting history of the first four revolutions of the satellite orbit of Mariner 9 is presented, with emphasis on the convergence problems and the delivered solution for the first orbit trim maneuver. Also included are a solution accuracy summary, the history of the spacecraft orbit osculating elements, the results of verifying the radio solutions with TV imaging data, and a summary of the normal points generated for the relativity experiment.

Jordan, J. F.↗