Leena Singh

Band-Limited Guidance and Control of Large Parafoils

* Principal member of the Technical Staff, Decision Systems Group, MS 15, Member AIAA. † Principal member of the Technical Staff, Aerospace Guidance and Control Group, MS 70, Member AIAA. ‡ Member of the Technical Staff II, Tactics, Guidance, and Control Group, MS 77. § Senior member of the Technical Staff, Cognitive Robotics Group, MS 77. ** Principal member of the Technical Staff, Manned Space Systems Group, MS 70, Member AIAA. †† Senior member of the Technical Staff, Vehicle and Robotics Group, MS 23, Member AIAA. ‡‡ Principal member of the Technical Staff, Navigation and Localization Group, MS 77. §§ Principal member of the Technical Staff, Tactical Systems Program Office, MS 79. *** Aerospace Engineer, NSRDEC, 15 Kansas Street, Senior Member AIAA. ††† Team Leader, NSRDEC, 15 Kansas Street, Member AIAA. 20th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar 4 7 May 2009, Seattle, Washington AIAA 2009-2981

Model predictive control in nap-of-Earth flight using polynomial control basis functions

This paper presents analysis using a new set of design tools to compute finite horizon optimal controls specifically for onboard model predictive control based trajectory synthesis. The optimizer must produce a finite-horizon control trajectory that will enable a UAV to track a low-altitude, high-speed (nap-of-Earth flight) reference trajectory. The optimizer must synthesize the finite-horizon controls based on a suitable fidelity plant model. This can rapidly become a high-dimensional, nonconvex optimization search, particularly if the dynamic model is nonlinear and the horizon long compared to the control bandwidth. To reduce the scope of the optimization problem we constrain the finite-horizon controls to a scaleable set of control basis functions (CBF). We also use these CBFs to identify a linear perturbation model around a nominal realizable trajectory. In this paper, we focus on polynomial CBFs such as Laguerre and Legendre. We compare results to a baseline of those obtained without any simplifying approximations, and to a repeating sequence of tent functions that were introduced in previous publications. Our analysis indicates that Laguerre polynomials pose a good choice of CBFs to design the optimal controls for NOE type of experiments; a fourth order Laguerre polynomial supplies 5 distinct and characteristic polynomial terms and is adequate for good tracking performance. The performance is equivalent to using 10 tent functions. However, since only 5 Laguerre polynomials need to be manipulated to form the optimal control, the optimization speed is greatly enhanced. It is significantly faster than solving a full order MPC problem.

Attitude determination and control system design for the CYGNSS microsatellite

This paper presents the development of the attitude determination and control system design of the Cyclone Global Navigation Satellite System spacecraft. The CYGNSS constellation consists of eight small satellite observatories in 500 km circular orbits at an inclination of 35 deg released from a single launch platform. Each CYGNSS spacecraft will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes with the objective to fundamentally improve gap-free coverage for hurricane forecast and monitoring. Realising this objective requires the spacecraft to accurately and reliably point its signal collection antennae in desired directions and hold its Earth relative attitude over long time durations to prescribed knowledge and point requirements. Indeed, the microsatellite ADCS regulates all spacecraft estimation and control functionality spanning detumbling; Sun acquisition and hold; pointing control and momentum-management over the micro-satellite lifetime to design requirements. This paper presents the ADCS hardware, software and algorithms used to control the spacecraft in all phases of CYGNSS operations and presents simulation based performance results of the closed-loop estimation and control systems.

Conical scanning approach for Sun pointing on the CYGNSS microsatellite

The small scales of area, volume, and power of small spacecraft, such as NASAs 25-kg Cyclone Global Navigation Satellite System (CYGNSS) satellites, constrain the number of independent subsystems that they can support. Consequently, small satellites often require novel approaches to execute the same mission functions that a larger satellite can easily perform with familiar sensor, actuator and algorithm options. In the case of CYGNSS, the spacecraft must execute a Sun acquisition and pointing phase but the actuator suite does not include 2-axis sun sensors or rate gyros; two measurements that seem like obvious inclusions for the Sun acquisition task. Instead, during Sun acquisition, the CYGNSS attitude control system uses a limited actuator and sensor set consisting of three magnetic torque rods, a three-axis magnetometer, and Sun incidence-angle measurements from three solar panel faces. This paper describes the sensing and control algorithms implemented in CYGNSS flight software to acquire and maintain Sun pointing with the available measurements and actuators. The Sun pointing algorithm uses a conical scanning approach based on traditional RF pointing and target-tracking systems, which consists of two key control loops: (1) a rate loop, which initiates a body spin about the solar-array face axis, and (2) a slower angle controller that tracks the array power gradients measured over the course of the fast spin. A slew toward the peak power eventually drives the solar panel face normal to spin in a cone centered about the Sun vector. The Sun acquisition process has a large convergence basin, is stable in the Lyapunov sense, and demonstrates excellent performance behavior in simulation.

L1 Adaptive Control Design for Improved Handling of the F/A-18 class of Aircraft

This paper presents the development and preliminary results obtained from the application of L1 adaptive control methods to the design of the inner-loop flight control system of variants of the F/A-18 aircraft, intended to emulate the handling qualities of the standard reference F/A-18 aircraft. The control method is designed to be robust to high frequency and unmodeled dynamics always present when a complex nonlinear system is approximated as a linear plant for design purposes. Results show the capability of the L1 adaptive controller to effectively manipulate the naturally lightly damped pitch and roll dynamics of the uncontrolled plant to the desired dynamics of the easier-to-control standard F/A-18 with bounded transient responses and steady state tracking errors over a range of operating conditions. The ability to synthesize controllers based on a set of desired reference dynamics allowing the onboard controller to synthesize suitable stabilising and error regulating control commands for the plant possesses many of the prerequisites needed to form the basis of a standard set of certifiable flight control software designs.

Modifications to Pose Estimation Algorithms that Allow for Near -Planar Feature Points

In the process of developing terrain rel ative navigation algorithms, different pose estimation algorithms were researched, tested and evaluated. The two best algorithms that fit our purposes were by Fiore, and by Ansar and Daniilidis . However, these algorithms required modifications for our app lication. Specifically, Fiore’s algorithm failed to handle feature points that are nearly planar. A modification to the algorithm presented herein allows for the seamless handling of planar, near -planar, and non -planar feature points without any addition al computational workload. Ansar and Daniilidis’ algorithm sporadically produced large errors in the presence of noise. Our modifications minimize this tendency with only a slight increase in the computational workload. I. Introduction N the process of de veloping t errain relative navigation (TRN) algorithms for planetary landing, several pose estimation algorithms were investigated. These algorithms take the identified features on the surface of the planet , and the corresponding features on the image and determine the position and attitude of the vehicle. Most pose estimation algorithm attempt to determine the camera matrix. Some methods are iterative and produce the ‘best’ solution, while others are linear and fast, but don’t produce as accurate resul ts . Two methods, one by Fiore [1] and a second by Ansar and Daniilidis [2], approach the problem in a different way. Neither algorithms attempt to find the camera matrix but instead find the ‘depths’ of the feature points. With the depth information, t he features in the 2D camera image are turned into 3D points in the camera’s frame of reference. With a simple absolute orientation algorithm, it is now possible to recover the position and attitude of the vehicle [3]. While implementing the two algori thms, problems developed . The algorithm given in Fiore’s paper works very well if the feature points do not lie in a plane. (If the feature points do lie in a plane, a second algorithm is presented that handles that specific case.) However, the algorith m does not work well when the points are ‘near -planar ’. Since planetary features look almost planar from orbit , but not quite, they fall in this range where the algorithm breaks down. This paper address that problem and provides a modification to the alg orithm that seamlessly handles non

Limitations of scaling momentum control strategies to small spacecraft

As a spacecraft becomes smaller, a number of physical effects scale both favorably and unfavorably for passive stabilization of the craft. Unfortunately, two separate quantities both scale unfavorably for the use of traditional spinning rotor actuators (e.g. reaction wheels, momentum wheels, control moment gyros) for momentum and attitude control. First, the dominant disturbance torques on small spacecraft in low earth orbit, aerodynamic drag and solar radiation pressure, both become relatively larger as spacecraft size decreases. Second, the effectiveness of spinning rotors reduces as the rotor inertia decreases with the square or the wheel radius. These two factors conspire to greatly reduce the effectiveness of rotor-based momentum control systems at small scales. This reduction requires small spacecraft designers to either devote a significantly larger mass fraction to momentum control or adopt alternative momentum control systems. In this study we examine this problem from two viewpoints. First, empirical data is used to find a relationship between spacecraft size and mass fraction devoted to attitude control. While the International Space Station can devote less than 1% of its mass fraction to momentum control effectors, GEO telecom spacecraft tend to need around 1–2% of available mass, and some CubeSats must devote greater than 50% of their mass fraction. Second, we derive an expression for the smallest spacecraft that can use a reaction wheel for effective momentum management. For reasonable assumptions, this lower limit is on the order of 1 cm length scale, which is in good agreement with the empirical trend. Finally, we list some alternative momentum management strategies and discuss how they apply to spacecraft at the smallest size: the centimeter scale ChipSat.