Jean-Francois Hamel

Linearized Dynamics of Formation Flying Spacecraft on a J2-Perturbed Elliptical Orbit

A linearized set of equations of relative motion about a J 2 -perturbed elliptical reference orbit is developed. This model uses analytical relations that are well suited for onboard applications. The inclusion of the J 2 perturbation in a simple analytical model can lead to formation flying guidance and control algorithms that make use of the natural J 2 -induced relative motion to perform maneuvers instead of constantly compensating for this perturbation. The model uses the linearized differential drift rate of mean orbit elements to predict the impact of the J 2 perturbation on relative osculating spacecraft motion. It analytically provides the relative motion in Hill coordinates at any given true anomaly using only the initial osculating relative orbit elements and the initial orbit elements of the reference trajectory. A linear time-varying state-space form of the model is also presented. Simulation results show that relative motion prediction remains accurate over several orbits.

Improvement to the analytical predictor-corrector guidance algorithm applied to mars aerocapture

Introduction O NE of the state-of-the-art technologies considered to reduce the cost of planetary exploration is aerocapture. This technique allows the reduction of fuel cost for planetary insertion by using atmospheric drag to decrease the total orbital energy of the vehicle. It consists in a reduction of velocity from a hyperbolic orbit or highly elliptical orbit to a low-altitude near-circular planetary orbit. It has previously been demonstrated that aerocapture would be beneficial for human exploration of Mars.1 The purpose of an aerocapture maneuver is to bring the vehicle from given atmospheric entry conditions to desired atmospheric exit conditions. The desired exit conditions are typically expressed as a given apoapsis radius of the unperturbed orbit once the vehicle is out of the atmosphere. This apoapsis radius is chosen to minimize the velocity impulse that is required to reach the final mission orbit. Up to now, several types of algorithms, such as the analytical predictor-corrector,2−6 the energy controller,4,7 the numerical predictor-corrector,4,8−10 and the terminal point controller6,11 have been developed, considering only the vehicle bank angle as control parameter. As shown in Fig. 1, the authors classify these algorithms in three main categories: the analytical algorithms, the numerical algorithms, and the predefined-trajectory algorithms. Firstly, the analytical predictor corrector (APC) and the energy controller are part of the first category. These algorithms make certain assumptions that lead to an analytical guidance solution to the exit conditions for the current vehicle state. Secondly, the numerical predictor corrector numerically integrates the remaining part of the trajectory to predict the atmospheric exit conditions from the current position and updates the commanded bank angle for the remaining part of the trajectory. It is therefore part of the second category. Finally, the terminal point controller, part of the third category, uses a predefined optimal trajectory. In this case, the vehicle tries to remain on the optimal trajectory at any moment in time.

Neighboring Optimum Feedback Control Law for Earth-Orbiting Formation-Flying Spacecraft

Thispaperdevelopsafeedbackcontrollawthatguaranteesneighboringfuel-optimalityofthereconfigurationofa formation of Earth-orbiting formation-flying spacecraft. It aims for the case in which a specific formation is to be achieved at a specific true anomaly. It guarantees neighboring fuel-optimality of such a reconfiguration maneuver, assuming that the formation evolves in the vicinity of an uncontrolled reference trajectory. It is in the semi-analytic form,as onlyone time-varyinggain matrixneedstobecomputed beforethe maneuver. Itallows afuel consumption/ formation accuracy tradeoff with the selection of only one scalar gain. Simulations compare the performance of this controller with the linear-quadratic regulator and the mean orbit elements controller in the context of a 1 km size formation reconfiguration. Simulations show that this neighboring optimum controller can perform the maneuver with better accuracy while spending as much or less propellant than the other controllers.

Feature Matching Navigation Techniques for Lidar-Based Planetary Exploration

When landing on planetary bodies it is desired to determine accurately the velocity and the position of the spacecraft relative to a selected target position on the surface of the body. This paper responds to that requirement by proposing Lidar-based advanced navigation techniques based on feature matching. Some of the techniques proposed are inspired from conventional 2D and 1D correlation techniques while others are taking advantage of technologies already validated in space and known as star constellation matching algorithms. After presenting the concepts of each algorithm, a realistic validation scenario based on a Mars landing reference mission is presented with comprehensive simulation results. At the end, analyses of the advantages and drawbacks are presented.

TICFIRE: a far infrared payload to monitor the evolution of thin ice clouds

The TICFIRE mission concept developed with the support of the Canadian Space Agency aims: 1) to improve measurements of water-vapour concentration in the low limit, where cold regions are most sensitive and 2) to determine the contribution of Thin Ice Clouds (TIC) to the energy balance and the role of their microphysical properties on atmospheric cooling. TICFIRE is a process-oriented mission on a micro-satellite platform dedicated to observe key parameters of TIC forming in the cold regions of the Poles and globally, in the upper troposphere. It locates cloud top profiles at the limb and measures at nadir the corresponding upwelling radiance of the atmosphere directly in the thermal window and in the Far Infrared (FIR) spectrum over cold geographical regions, precisely where most of the atmospheric thermal cooling takes place. Due to technological limitations, the FIR spectrum (17 to 50 μm) is not regularly monitored by conventional sensors despite its major importance. This deficiency in key data also impacts operational weather forecasting. TICFIRE will provide on a global scale a needed contribution in calibrated radiance assimilation near the IR maximum emission to improve weather forecast. TICFIRE is therefore a science-driven mission with a strong operational component. The TICFIRE payload consists of two instruments; the main one being a Nadir-looking multiband radiometer based on uncooled microbolometer technology and covering a large spectral range from 7.9 μm to 50 μm. The secondary one is an imager that performs Limb measurements and provides cloud vertical structure information. This paper presents the key payload requirements, the conceptual design, and the estimated performance of the TICFIRE payload. Current technology developments in support to the mission are also presented.

Overview of Japan Canada Joint Collaboration Satellites (JC2Sat) GNC Challenges and Design

The Japan Canada Joint Collaboration Satellites – Formation Flight (JC2Sat-FF) project is a joint project between the Canadian Space Agency (CSA) and the Japan Aerospace Exploration Agency (JAXA). This paper presents an overview of JC2Sat mission and discusses unique GNC challenges and solutions with focus on attitude control subsystem (ACS) and relative navigation subsystem (RNS). Preliminary simulations are presented to validate that the design requirements are satisfied by proposed design techniques. Nomenclature Ad = desired geocentric separation latitude prop g sat ( ) a R = satellite gravitational acceleration including only the two-body and J2 perturbation accelerations B = earth magnetic field vector expressed in satellite body coordinates hy e = angular momentum error about the satellite pitch axis h = satellite angular momentum vector expressed in satellite body coordinates h % = satellite angular momentum error vector expressed in satellite body coordinates d h = desired angular momentum vector expressed in satellite body coordinates dw h = desired angular momentum about the satellite pitch axis i = inclination I = satellite moment of inertia matrix expressed in satellite body coordinates 1 2 , , k k k = control gains vo ro k k , = constant gains vr rr k k , = constant gains m = magnetorquer dipole moment vector expressed in the satellite body frame max m = magnetorquer limit vector r = an upper-bound on the covariance of the GPS absolute or relative position error

Performance Assessment of the Drag-Based Formation Control for the JC2Sat Mission

ठThe paper describes an innovative method for the performance assessment of drag-based formation flying control techniques. The method aims at avoiding time-consuming longduration closed-loop simulation campaigns with high fidelity simulators. The methodology quickly provides statistical insight into the formation control performance, thus supporting quick design iterations and systems engineering trade-off analyses. The paper reviews the drag-based guidance and control algorithm proposed for the JC2Sat mission, presents the methodology developed to assess performance, discusses the application of this methodology in the context of the JC2Sat mission and shows typical analysis results.

Pseudo-Doppler Velocity Navigation for Lidar-Based Planetary Exploration

When landing on a planetary body, the knowledge of the Lander velocity relative to the surface is required in order to ensure a safe touchdown. A conventional Doppler radar can provide this measurement. However, it has been shown in previous studies that Lidar mappers, with their ability to view the landing area in three dimensions, are ideal sensors for hazard-avoidance landing. In order to minimize the power, mass and volume of on-board sensors, this paper proposes to use the Lidar data to estimate the Lander relative velocity. An innovative Lidar-based Pseudo-Doppler velocity determination algorithm is presented with the purpose of determining velocity in conditions of high relative velocity, condition in which techniques based on feature matching may fail because of the distortions created by the Lander motion. The techniques proposed here could estimate the velocity during and at the end of the parachute phase when landing on Mars. The algorithm is described and validated by simulation.

Fuel-Equivalent Relative Orbit Element Space

This paper presents a new tool to analytically perform the guidance for reconfiguration of formation flying spacecraft. The technique consists in mapping the relative orbit elements into a fuel-equivalent space where similar displacements correspond to an equivalent fuel consumption. The minimal-fuel maneuver problem is consequently translated into a simple geometric problem in the fuel-equivalent space. The theory is applied to two well-known formations: the J 2 -invariant formation and the projected circular formation. The use of the fuel-equivalent space leads to very simple solutions for the most fuel-efficient way to attain both formations.

An uncooled mid wave and thermal infrared payload for fire monitoring

The Platform for the Observation of the Earth and for in-orbit Technology Experiments (POETE) mission concept has been developed to help overcome the scientific and socio-economic issues associated with forest fires. The proposed mission is based on a series of two highly autonomous and agile microsatellites, allowing for 3 to 7 visits per day. Each satellite payload includs a VIS-NIR instrument and a MWIR-TIR instrument. The two instruments combined provide for 6 spectral channels spanning from the visible to the thermal infrared for fire monitoring, retrieval of quantitative fire parameters (such as effective fire temperature, area and radiative energy release), and land surface temperature measurement. The MWIR-TIR instrument concept is a pushbroom scanner filter radiometer with on-board radiometric calibration capabilities. Its all-reflective three-mirror input optics delivers a 400-m GSD at an altitude of 700 km, relaying the scene signal to detectors based on INOs microbolometer technology for detection in four spectral channels centered at 3.8 μm, 8.8 μm, 10.5 μm and 12.0 μm. This paper presents an overview of the key mission requirements and derived sensor level requirements. A description of the conceptual design of the MWIR-TIR payload of POETE is given along with estimates of key performance parameters.