Andrew Heaton

Orbital Express Advanced Video Guidance Sensor

In May 2007 the first US-sponsored fully autonomous rendezvous and capture was successfully performed by DARPAs Orbital Express (OE) mission. For the following three months, the Boeing ASTRO spacecraft and the Ball Aerospace NEXTSat performed multiple rendezvous and docking maneuvers to demonstrate some of the technologies needed for satellite servicing. MSFCs advanced video guidance sensor (AVGS) was a near-field proximity operations sensor integrated into ASTROs Autonomous Rendezvous and Capture Sensor System (ARCSS), which provided relative state knowledge to the ASTRO GN&C system. AVGS was one of the primary docking sensors included in ARCSS. This paper provides an overview of the AVGS sensor that flew on orbital express, a summary of the AVGS ground testing, and a discussion of AVGS performance on-orbit for OE. The AVGS is a laser-based system that is capable of providing bearing at midrange distances and full six degree- of-freedom (6-DOF) knowledge at near ranges. The sensor fires lasers of two different wavelengths to illuminate retro- reflectors on the long range target (LRT) and the Short Range Target (SRT) mounted on NEXTSat. The retro- reflector filters allow one laser wavelength to pass through and be reflected, while blocking the other wavelength. Subtraction of one return image from the other image removes extraneous light sources and reflections from anything other than the corner cubes on the LRT and SRT. The very bright spots that remain in the subtracted image are processed to provide bearing or 6-DOF relative state information. AVGS was operational during the Orbital Express unmated scenarios and the sensor checkout operations. The OE unmated scenarios ranged from 10 meters to 7 kilometers ending in either a docking or a free-flyer capture. When the target was pointed toward the AVGS and in the AVGS operating range and field-of-view (i.e. along the approach corridor of the NEXTSat), the AVGS provided full 6-DOF measurements. The AVGS performed very well during the sensor check-out operations, effectively tracking beyond its 10-degree Pitch and Yaw limit-specifications. AVGS also provided excellent performance during the unmated operations, effectively tracking its targets, and showing good agreement between the SRT and LRT data. The AVGS consistently exceeded the tracking range expectations for both the SRT and LRT. During the approach to re-mate in scenario 3-1 recovery the AVGS began tracking the LRT at 150 m, well beyond the OE specified operational range of 120 meters, and functioned as the primary sensor for the autonomous rendezvous and docking. For all scenarios, the AVGS was used while ASTRO was in the approach corridor to NEXTSat, and during close proximity operations and docking.

Orbital Express Advanced Video Guidance Sensor: Ground Testing, Flight Results and Comparisons

Orbital Express (OE) was a successful mission demonstrating automated rendezvous and docking. The 2007 mission consisted of two spacecraft, the Autonomous Space Transport Robotic Operations (ASTRO) and the Next Generation Serviceable Satellite (NEXTSat) that were designed to work together and test a variety of service operations in orbit. The Advanced Video Guidance Sensor, AVGS, was included as one of the primary proximity navigation sensors on board the ASTRO. The AVGS was one of four sensors that provided relative position and attitude between the two vehicles. Marshall Space Flight Center was responsible for the AVGS software and testing (especially the extensive ground testing), flight operations support, and analyzing the flight data. This paper briefly describes the historical mission, the data taken on-orbit, the ground testing that occurred, and finally comparisons between flight data and ground test data for two different flight regimes.

Orbital Express AVGS Validation and Calibration for Automated Rendezvous

From March to July of 2007, the DARPA Orbital Express mission achieved a number of firsts in autonomous spacecraft operations. The NASA Advanced Video Guidance Sensor (AVGS) was the primary docking sensor during the first two dockings and was used in a blended mode three other automated captures. The AVGS performance exceeded its specification by approximately an order of magnitude. One reason that the AVGS functioned so well during the mission was that the validation and calibration of the sensor prior to the mission advanced the state-of-the-art for proximity sensors. Some factors in this success were improvements in ground test equipment and truth data, the capability for ILOAD corrections for optical and other effects, and the development of a bias correction procedure. Several valuable lessons learned have applications to future proximity sensors.

The Advanced Video Guidance Sensor: Orbital Express and the Next Generation

The Orbital Express (OE) mission performed the first autonomous rendezvous and docking in the history of the United States on May 5-6, 2007 with the Advanced Video Guidance Sensor (AVGS) acting as one of the primary docking sensors. Since that event, the OE spacecraft performed four more rendezvous and docking maneuvers, each time using the AVGS as one of the docking sensors. The Marshall Space Flight Centers (MSFCs) AVGS is a near- field proximity operations sensor that was integrated into the Autonomous Rendezvous and Capture Sensor System (ARCSS) on OE. The ARCSS provided the relative state knowledge to allow the OE spacecraft to rendezvous and dock. The AVGS is a mature sensor technology designed to support Automated Rendezvous and Docking (AR&D) operations. It is a video-based laser-illuminated sensor that can determine the relative position and attitude between itself and its target. Due to parts obsolescence, the AVGS that was flown on OE can no longer be manufactured. MSFC has been working on the next generation of AVGS for application to future Constellation missions. This paper provides an overview of the performance of the AVGS on Orbital Express and discusses the work on the Next Generation AVGS (NGAVGS).

NanoSail:D Orbital and Attitude Dynamics

NanoSail-D unfurled January 20th, 2011 and successfully demonstrated the deployment and deorbit capability of a solar sail in low Earth orbit. The orbit was strongly perturbed by solar radiation pressure, aerodynamic drag, and oblate gravity which were modeled using STK HPOP. A comparison of the ballistic coefficient history to the orbit parameters exhibits a strong relationship between orbital lighting, the decay rate of the mean semi-major axis and mean eccentricity.

POSE algorithms for automated docking

POSE (relative position and attitude) can be computed in many different ways. Given a sensor that measures bearing to a finite number of spots corresponding to known features (such as a target) of a spacecraft, a number of different algorithms can be used to compute the POSE. NASA has sponsored the development of a flash LIDAR proximity sensor called the Vision Navigation Sensor (VNS) for use by the Orion capsule in future docking missions. This sensor generates data that can be used by a variety of algorithms to compute POSE solutions inside of 15 meters, including at the critical docking range of approximately 1-2 meters. Previously NASA participated in a DARPA program called Orbital Express that achieved the first automated docking for the American space program. During this mission a large set of high quality mated sensor data was obtained at what is essentially the docking distance. This data set is perhaps the most accurate truth data in existence for docking proximity sensors in orbit. In this paper, the flight data from Orbital Express is used to test POSE algorithms at 1.22 meters range. Two different POSE algorithms are tested for two different Fields-of-View (FOVs) and two different pixel noise levels. The results of the analysis are used to predict future performance of the POSE algorithms with VNS data.

Strong Solar Radiation Forces from Anomalously Reflecting Metasurfaces for Solar Sail Attitude Control

We examine the theoretical implications of incorporating metasurfaces on solar sails, and the effect they can have on the forces applied to the sail. This would enable a significant enhancement over state-of-the- art attitude control by demonstrating a novel, propellant-free and low-mass approach to induce a roll torque on the sail, which is a current limitation in present state-of-the-art technology. We do so by utilizing anomalous optical reflections from the metasurfaces to generate a net in-plane lateral force, which can lead to a net torque along the roll axis of the sail, in addition to the other spatial movements exhibited by the sail from solar radiation pressure. We characterize this net lateral force as a function of incidence angle. In addition, the influence of the phase gradients and anomalous conversion efficiencies characteristics of the metasurfaces are independently considered. The optimum incidence angle that corresponded with the maximum net lateral-to-normal force ratio was found to be −30° for a metasurface exhibiting 75% anomalous conversion efficiency with a phase gradient of 0:71k0.

Near Earth Asteroid (NEA) Scout Solar Sail Contingency Trajectory Design and Analysis

Exploratory missions to investigate accessible Near Earth Asteroids (NEAs) can benefit from leveraging dynamics associated with a solar sail-based spacecraft. As a part of this effort, NEA Scout is a solar sail mission designed to propel a 6U CubeSat by harnessing solar radiation pressure from the Sun. The spacecraft will be launched as a secondary payload on NASA’s Space Launch Vehicle (SLS) Exploration Mission One (EM-1). As the launch of EM-1 has recently been rescheduled for December 2019, alternative target NEAs are identified. Additionally, solar sail-based trajectories for NEA Scout mission also need to be reevaluated. In this paper, high-fidelity trajectories for the NEA Scout mission are investigated for varying launch dates. Furthermore, feasible trajectory solutions are presented for multiple candidate asteroids.

Solar Torque Management for the Near Earth Asteroid Scout CubeSat Using Center of Mass Position Control

A novel mechanism, the Active Mass Translator (AMT), has been developed for the NASA Near Earth Asteroid (NEA) Scout mission to autonomously manage the spacecraft momentum. The NEA Scout CubeSat will launch as a secondary payload onboard Exploration Mission 1 of the Space Launch System. To accomplish its mission, the CubeSat will be propelled by an 86 square-meter solar sail during its two-year journey to reach asteroid 1991VG. NEA Scout’s primary attitude control system uses reaction wheels for holding attitude and performing slew maneuvers, while a cold gas reaction control system performs the initial detumble and early trajectory correction maneuvers. The AMT control system requirements, feedback architecture, and control performance will be presented. The AMT reduces the amount of reaction control propellant needed for momentum management and allows for smaller capacity reaction wheels suitable for the limited 6U spacecraft volume. The reduced spacecraft mass allows higher in-space solar sail acceleration, thus reducing time-offlight. The reduced time-of-flight opens the range of possible missions, which is limited by the lifetime of typical non-radiation tolerant CubeSat avionics exposed to the deep-space environment.

Propulsion Technology Assessment: Science & Enabling Technologies to Explore the Interstellar Medium

As part of a larger effort led by the Keck Institute for Space Studies at the California Institute of Technology, the Advanced Concepts Office at NASAs George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar systems heliopause, approximately 100 AU from the center of the sun, within 10 years and within a 2025 to 2035 launch window. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature (MaSMi) Hall thruster, a solar sail and an electric sail. Two aerial densities, and thus characteristic accelerations, 0.426 mm/s(exp 2) and 0.664 mm/s(exp 2), were analyzed for the solar sail option in order understand the impact of near and long term development of this technology. Similarly, two characteristic accelerations, 1 mm/s(exp 2) and 2 mm/s(exp 2), were also analyzed for the electric sail option in addition to tether quantities of 10 and 20, respectively, and individual tether length of 20 km. A second analysis was conducted to determine what existing solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming an earth departure characteristic energy capability provided by a Space Launch System (SLS) Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist (E-Ju), and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist (E-Ju-Su-Sa). The Oberth maneuver would need to be performed very close to the sun, wherein this study assumed a perihelion distance of approximately 11 solar radii, or 0.05 AU, away from the surface. The heat shield technology required to perform this type of ambitious maneuver was assumed to be similar to that of NASAs Solar Probe Plus mission, which is slated to launch in July 2018. With respect to a SLS Block 1B earth departure characteristic energy capability of 100 km(exp 2)/s(exp 2) for the E-Ju trajectory option, results indicated that compared to having no advanced propulsion system onboard, both the MaSMi Hall thruster and solar sail options subtract approximately 8 to 10 years from the total trip time while the electric sail outperforms all options by subtracting up to 20 years. With respect to an average kick stage velocity capability of 2.5 to 3.5 km/s at perihelion, the most sensitive segment of the E-Ju-Su-Sa trajectory option, results indicated that both the MaSMi Hall thrust and solar sail options only subtract 1 to 3 years from the total trip time whereas the electric sail again outperforms all other options by subtracting up to 5 years. In other words, if the Technology Readiness Level of an electric sail could be increased in time, this propulsion technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so even faster. Completing such an ambitious mission in that short of a timespan would be very attractive to many as it would be well within the average career span of any of those involved.