Relative Navigation Sensor Research Articles

LIDAR adaptive parameters selection and target pointing control for close-proximity space operations

This paper deals with advanced Guidance, Navigation and Control (GNC) functions required to enable autonomous and safe operations of a chaser spacecraft in close-proximity to an uncooperative space target, as in Active Debris Removal or On-Orbit servicing scenarios. Specifically, it presents an original approach to autonomously and adaptively select the field of view and resolution of a scanning LIDAR to improve both state estimation accuracy and computational efficiency of a LIDAR-based relative navigation system. In general, the correct operation of such system is also determined by the capability to keep the boresight axis of the relative navigation sensor aligned with the target geometric center. To address this task, an original control technique, based on the sliding-mode formulation and relying on a reduced attitude representation, is proposed. This control scheme is also compared to state-of-the-art PD and PID approaches in terms of pointing accuracy and control effort. The proposed techniques have been numerically validated in a simulation environment integrating the chaser attitude control and the LIDAR-based relative navigation functions in a closed-loop architecture. The simulation environment realistically reproduces the generation of LIDAR-based point clouds, and the spacecraft relative dynamics in close proximity. Results show that the adaptive selection of the LIDAR operational parameters improves the relative navigation performance, while the proposed sliding-mode control guarantees higher pointing accuracy than PD and PID control approaches.

Optimal Powered Descent Guidance Under Thrust Pointing Constraint

This paper proposes a method for solving a three-dimensional propellant-optimal powered descent guidance problem under a thrust pointing constraint, where the thrust direction is restricted. Incorporating a thrust pointing constraint is as essential as pursuing propellant optimality for modern powered descent guidance, particularly when a lander uses terrain relative navigation sensors. However, the theoretical knowledge of optimal solutions is limited. One of the paper’s contributions is a method for finding the optimal solution under an upward-pointing constraint with an active–inactive–active switching structure. Usually, there is a maximum–minimum–maximum structure for the optimal thrust magnitude. Another contribution is compact propellant-optimal powered descent guidance under a thrust pointing constraint. The guidance may find the solution by establishing analytical models to evaluate switching points from the switching functions and predict the final-state errors. Numerical simulations provide the solutions of lunar pinpoint-landing scenarios, in addition to demonstrating the excellent propellant optimality, smooth convergence, and fast execution.

Design and Analysis of Descent-to-Landing Navigation Incorporating Terrain Effects

This study investigated the ability of a navigation filter to process multiple terrain-based sensors, such as slant-range, slant-speed, and terrain relative navigation sensors, during a descent-to-landing scenario to estimate the state of a landing vehicle. The filtering technique leveraged was based upon a factorized form of the multiplicative extended Kalman filter, and terrain-based measurements were fused with star camera and inertial measurement unit sensor returns to estimate the position, velocity, and attitude of the landing vehicle. Monte Carlo simulations were carried out to assess the performance of the navigation filter along a lunar descent trajectory, as well as the consistency of the filtering solutions. A comprehensive error budget was constructed, followed by a sensitivity analysis, to investigate the behavior of the filter uncertainty with respect to specific error sources. A final analysis was performed, in which the filter was subjected to terrain model mismatch.

Safe trajectory design and pose estimation for target monitoring in GEO

This paper lies in the framework of mission scenarios, such as Active Debris Removal and On-Orbit Servicing, which require an active spacecraft (chaser) to orbit in close-proximity with respect to a space target. Specifically, these activities involve relative orbital maneuvers, such as monitoring, rendezvous and docking, in which the target-chaser distance ranges from a few tens of meters (depending on the target size) up to contact (in the case of docking). A critical challenge related to the realization of these maneuvers is the need to minimize the risk of collision, considering that the target is a non-cooperative object which may be characterized by uncontrolled rotational dynamics. This goal can be achieved by designing relative trajectories which satisfy specific constraints in terms of safety and stability, on one side, as well as by exploiting relative navigation technologies and algorithms which provide highly accurate estimates of the target-chaser relative motion parameters thus allowing to relax the control requirements. Both these aspects are addressed by this paper with focus on Geostationary Earth Orbits since they represent a particularly crowded orbital region in which the possibility to remove large debris and to extend the operative life of spacecraft, such as telecommunication ones, may have a significant scientific and economic benefit. Hence, an original method is presented to design safety ellipses for target monitoring around GEO targets, which, simultaneously, can provide optimal relative observation geometry for relative navigation (pose determination) using Electro-Optical sensors. The design approach is formulated in mean orbit parameters and it is based on a relative motion model relevant to two-satellite formations which includes the non-Keplerian perturbations due to secular Earth oblateness, as well as the possibility of considering targets moving along a small-eccentricity orbit. An example of trajectory design is shown considering a GEO target as test case. Given this trajectory, pose determination performance is also evaluated within a numerical simulation environment capable of realistically reproducing target-chaser relative dynamics, the operation of a scanning LIDAR selected on board the chaser as relative navigation sensor, and pose estimation algorithms based on the processing of 3D point clouds.

Overview and GNC design of the CubeSat Proximity Operations Demonstration (CPOD) mission

The CubeSat Proximity Operations Demonstration (CPOD) mission will demonstrate rendezvous, proximity operations, and docking with a pair of 3U CubeSats using miniaturized components and sensors. The goal of this mission is to develop small spacecraft technologies with game-changing potential and validate these technologies via spaceflight. Several new systems have been designed specifically for this program, including: next generation star trackers, next generation miniature reaction wheels, miniature cold-gas multi-thruster propulsion modules, a new relative navigation sensor suite, power management and distribution electronics based on flight proven designs, and intelligent software solutions hosted on multiple low-power Linux ARM processors. This paper will present an overview of the CPOD spacecraft and the mission Concept of Operations, as well as the design of the formation flying guidance, navigation, and control subsystem and algorithms.

A relative navigation sensor for CubeSats based on LED fiducial markers

Small satellite platforms are becoming very appealing both for scientific and commercial applications, thanks to their low cost, short development times and availability of standard components and subsystems. The main disadvantage with such vehicles is the limitation of available resources to perform mission tasks. To overcome this drawback, mission concepts are under study that foresee cooperation between autonomous small satellites to accomplish complex tasks; among these, on-orbit servicing and on-orbit assembly of large structures are of particular interest and the global scientific community is putting a significant effort in the miniaturization of critical technologies that are required for such innovative mission scenarios. In this work, the development and the laboratory testing of an accurate relative navigation package for nanosatellites compliant to the CubeSat standard is presented. The system features a small camera and two sets of LED fiducial markers, and is conceived as a standard package that allows small spacecraft to perform mutual tracking during rendezvous and docking maneuvers. The hardware is based on off-the-shelf components assembled in a compact configuration that is compatible with the CubeSat standard. The image processing and pose estimation software was custom developed. The experimental evaluation of the system allowed to determine both the static and dynamic performances. The system is capable to determine the close range relative position and attitude faster than 10 S/s, with errors always below 10 mm and 2 deg.

Design and validation of a GNC system for missions to asteroids: the AIM scenario

Deep space missions, and in particular missions to asteroids, impose a certain level of autonomy that depends on the mission objectives. If the mission requires the spacecraft to perform close approaches to the target body (the extreme case being a landing scenario), the autonomy level must be increased to guarantee the fast and reactive response which is required in both nominal and contingency operations. The GNC system must be designed in accordance with the required level of autonomy. The GNC system designed and tested in the frame of ESA’s Asteroid Impact Mission (AIM) system studies (Phase A/B1 and Consolidation Phase) is an example of an autonomous GNC system that meets the challenging objectives of AIM. The paper reports the design of such GNC system and its validation through a DDVV plan that includes Model-in-the-Loop and Hardware-in-the-Loop testing. Main focus is the translational navigation, which is able to provide online the relative state estimation with respect to the target body using exclusively cameras as relative navigation sensors. The relative navigation outputs are meant to be used for nominal spacecraft trajectory corrections as well as to estimate the collision risk with the asteroid and, if needed, to command the execution of a collision avoidance manoeuvre to guarantee spacecraft safety

Integrated Navigation System Design for Micro Planetary Rovers: Comparison of Absolute Heading Estimation Algorithms and Nonlinear Filtering.

This paper provides algorithms to fuse relative and absolute microelectromechanical systems (MEMS) navigation sensors, suitable for micro planetary rovers, to provide a more accurate estimation of navigation information, specifically, attitude and position. Planetary rovers have extremely slow speed (~1 cm/s) and lack conventional navigation sensors/systems, hence the general methods of terrestrial navigation may not be applicable to these applications. While relative attitude and position can be tracked in a way similar to those for ground robots, absolute navigation information is hard to achieve on a remote celestial body, like Moon or Mars, in contrast to terrestrial applications. In this study, two absolute attitude estimation algorithms were developed and compared for accuracy and robustness. The estimated absolute attitude was fused with the relative attitude sensors in a framework of nonlinear filters. The nonlinear Extended Kalman filter (EKF) and Unscented Kalman filter (UKF) were compared in pursuit of better accuracy and reliability in this nonlinear estimation problem, using only on-board low cost MEMS sensors. Experimental results confirmed the viability of the proposed algorithms and the sensor suite, for low cost and low weight micro planetary rovers. It is demonstrated that integrating the relative and absolute navigation MEMS sensors reduces the navigation errors to the desired level.

Noncooperative Rendezvous Using Angles-Only Optical Navigation: System Design and Flight Results

This paper presents system design and on-orbit results from the Advanced Rendezvous Demonstration using Global Positioning System and Optical Navigation (ARGON). ARGON has been conducted during the extended phase of the Prototype Research Instruments and Space Mission Technology Advancement (PRISMA) mission in April 2012. It represents one of the first rendezvous technology experiments using line of sight measurements from optical navigation and was motivated by the new generation of on-orbit-servicing and debris-removal missions which are discussed at national and international level. Its primary goal was to demonstrate the capability of an active servicer spacecraft to safely approach and rendezvous a non-cooperative passive client using angles-only optical navigation in a ground-in-the-loop fashion. To this end, a dedicated flight dynamics system has been developed for routine processing of the camera images collected on-board, for estimation of the relative orbit of the servicer with respect to the client vehicle, for maneuver planning and commanding. Despite the inherent difficulty to estimate the actual range to target through angles-only measurements and the constraints affecting the communication between ground-station and servicer, ARGON demonstrated an efficient and safe rendezvous from 30 km to the final hold point at 3 km mean separation selected before experiment start. This was possible due to the achieved relative navigation accuracy in combination with a guidance strategy based on the relative eccentricity/inclination vector separation method. As shown in the paper, the availability of independent and precise navigation information from carrier-phase differential Global Positioning System techniques gave the possibility to properly evaluate the achieved performance and cross-compare different relative navigation sensors after the conclusion of the technology demonstration.

Relative state estimation and observability for formation flying satellites in the presence of sensor noise

This paper presents an investigation of the relative state estimation and observability for two formation flying satellites using two different relative navigation sensor sets. The first set consists of a transmitter antenna on one satellite and a single receiver antenna on the other satellite to measure the inter-satellite range using a radiofrequency ranging signal. The second set uses three receiver antennas to measure multiple ranges, effectively providing angular information. In this paper, it is shown for the more complete sensor set that the estimation error in the relative state is a function of the pseudorange error, the inter-satellite distance, and the receiver antenna baselines. By varying these parameters, conditions are found for which the measurements obtained using the first sensor set result in relative state estimation and observability comparable to those obtained with the more complete sensor set. This offers potential reductions in cost and complexity for certain mission scenarios.

The HST SM4 Relative Navigation Sensor System: Overview and Preliminary Testing Results from the Flight Robotics Lab

The upcoming Hubble Space Telescope (HST) Servicing Mission 4 (SM4) includes a Relative Navigation Sensor (RNS) experiment which uses three cameras and an avionics package to record images, and estimate in real-time the relative position and attitude (aka “pose”) during the Shuttle capture and deployment of the telescope. RNS recently completed its third and final phase of testing at the Marshall Space Flight Center Flight Robotics Laboratory. This testing utilized flight spare cameras, engineering development unit avionics and flight pose algorithms to estimate the pose of a Hubble mockup mounted to the Flight Robotics Laboratory (FRL) Dynamic Overhead Target Simulator (DOTS). The mockup was moved through a variety of flight-like lighting conditions and trajectories. In this paper we present pose estimation results from the third phase of RNS FRL testing.

History of Space Shuttle Rendezvous and Proximity Operations

Space Shuttle rendezvous missions present unique challenges that were not fully recognized when the Shuttle was designed. Rendezvous targets could be passive (i.e., no lights or transponders), and not designed to facilitate Shuttle rendezvous, proximity operations, and retrieval. Shuttle reaction control system jet plume impingement on target spacecraft presented induced dynamics, structural loading, and contamination concerns. These issues, along with limited reaction control system propellant in the Shuttle nose, drove a change from the legacy Gemini/Apollo coelliptic profile to a stable orbit profile, and the development of new proximity operations techniques. Multiple scientific and on-orbit servicing missions, and crew exchange, assembly and replenishment flights to Mir and to the International Space Station drove further profile and piloting technique changes. These changes included new proximity operations, relative navigation sensors, and new computer generated piloting cues. However, the Shuttle's baseline rendezvous navigation system has not required modification to place the Shuttle at the proximity operations initiation point for all rendezvous missions flown.

Trajectory Tracking Controller for Vision-Based Probe and Drogue Autonomous Aerial Refueling

This paper addresses autonomous aerial refueling between an unmanned tanker aircraft and an unmanned receiver aircraft using the probe-and-drogue method. An important consideration is the ability to achieve successful docking in the presence of exogenous inputs such as atmospheric turbulence. Practical probe and drogue autonomous aerial refueling requires a reliable sensor capable of providing accurate relative position measurements of su‐cient bandwidth, integrated with a robust relative navigation and control algorithm. This paper develops a Reference Observer Based Tracking Controller that does not require a model of the drogue or presumed knowledge of its position, and integrates it with an existing vision based relative navigation sensor. A trajectory generation module is used to translate the relative drogue position measured by the sensor into a smooth reference trajectory, and an output injection observer is used to estimate the states to be tracked by the receiver aircraft. Accurate tracking is provided by a state feedback controller with good disturbance rejection properties. A frequency domain stability analysis for the combined reference observer and controller shows that the system is robust to sensor noise, atmospheric turbulence, and high frequency unmodeled dynamics. Feasibility and performance of the total system is demonstrated by simulated docking maneuvers of an unmanned receiver aircraft docking with the non-stationary drogue of an unmanned tanker, in the presence of atmospheric turbulence. Performance characteristics of the vision based relative navigation sensor are also investigated, and the total system is compared to an earlier version. Results presented in the paper indicate that the integrated sensor and controller enable precise aerial refueling, including consideration of realistic measurements errors, plant modeling errors, and disturbances.

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.