Michael A. Paluszek

Development of an aerodynamic model and control law design for a high altitude airship

Lighter-than air vehicles are an attractive solution for many applications requiring a sustained airborne presence. The buoyancy force provides an energy-free form of lift, offering a non-traditional approach to long-duration missions for which traditional aircraft are not well-suited. Potential applications include roving or hovering surveillance and communication utilities for both military and commercial use, and a variety of remotesensing instruments for the scientific community. In particular, the Missile Defense Agency plans to utilize unmanned airships at high-altitudes to provide a long-duration missile defense presence around the coast-line of the United States. Operated at 70 kft, each of these “high altitude airships” will fly above all regulated air-traffic for several months to years, will reside in a steady atmospheric regime, and will utilize solar energy to provide all required power. Two key objectives for this type of mission are that the unmanned airship have exceptionally long endurance, and that it operate with a sufficiently high-level of autonomy. In order to achieve these objectives, a robust guidance and control system is required, capable of auto-piloting and controlling the airship under an extremely wide range of atmospheric and wind conditions. The successful design of such a system first requires an accurate model of airship dynamics across its expansive flight envelope, and a representative model of the expected disturbances. The dynamics of an airship are markedly different from traditional aircraft, with significant effects from added mass and inertia, and a much higher sensitivity to wind. In this paper, a typical airship configuration is first sized to meet energy balance and mass constraints. The geometry of this configuration is then used to develop a general aerodynamic model for the airship. The equations of motion with added mass and inertia are developed, and the open-loop dynamics are analyzed across a range of flight conditions. Finally, control laws are designed for a single operating condition, and the closed-loop performance is presented across a range of velocities.

Robust Attitude Control Systems Design for Solar Sails, Part 1: Propellantless Primary ACS

A robust attitude control system (ACS) architecture is developed for near-term solar sail missions. The proposed sailcraft ACS consists of a propellantless primary ACS and a microthruster-based secondary ACS. The primary ACS employs two ballast masses running along mast lanyards for pitch/yaw trimming and thrust vector control. It also employs roll stabilizer bars at the mast tips for quadrant tilt control. The secondary ACS utilizes lightweight pulsed plasma thruster (PPT) modules mounted at the mast tips. Such a microPPT-based secondary ACS can be employed for attitude recovery maneuvers from various off-nominal conditions, including tumbling, that cannot be handled by the propellantless primary ACS. The overall simplicity, effectiveness, and robustness of the proposed ACS architecture are demonstrated for a solar sail flight validation experiment of a 40-m sailcraft proposed in a 1600-km, dawn-dusk sun-synchronous (DDSS) orbit. The proposed ACS architecture will be applicable with minimal modifications to a wide range of future solar sail flight missions with varying requirements and mission complexity. An overview of the state-of-the-art microPPT technology, PPT requirements for solar sails, and pulse-modulated control design and simulation results are presented in the companion paper (Part 2).

Control-Moment Gyroscopes for Joint Actuation: A New Paradigm in Space Robotics

Manned spacecraft will require maintenance robots to inspect and repair components of the spacecraft that are accessible only from the outside. This paper presents a design of a novel free-flying maintenance robot (known as a MaintenanceBot.) The MaintenanceBot uses Control Moment Gyros (CMGs) for manipulator arm and attitude control. This architecture provides high authority control in a compact low power package. Relative position and attitude determination is accomplished with an RF system supplemented by a vision system at close range. When not docked to the manned vehicle (which must be done periodically to refuel and recharge batteries or when the manned vehicle performs orbit changes) the MaintenanceBots fly in formation using a cold gas thruster system and formation flying algorithms that permit dozens of MaintenanceBots to coordinate their positions. The use of CMGs is a prominent feature of this design. An array of CMGs can exchange angular momentum with the spacecraft body to effect attitude changes, as long as certain mathematical singularities in the actuator Jacobian are avoided. The proposed maintenance robot benefits dramatically from the dynamics and control of a multibody robotic arm whose joints are driven by CMGs. In addition to high power efficiency, another advantage of this concept is that spacecraft appendages actuated by CMGs can be considered reactionless, in the sense that careful manipulation of the CMG gimbal angles can virtually eliminate moments applied to the spacecraft body. This paper provides a preliminary design of the MaintenanceBot. Analysis of the formation flying and close maneuver control systems is included. Simulation results for a typical operation is provided.

Robust Attitude Control Systems Design for Solar Sails, Part 2: MicroPPT-based Secondary ACS

A secondary attitude control system (ACS) of utilizing tip-mounted, lightweight pulsed plasma thruster (PPT) is developed for near-term solar sails. Such a secondary ACS can be employed for attitude recovery maneuvers from various offnominal conditions, including tumbling, that cannot be handled by a propellantless primary ACS (described in the companion paper). The microPPT-based ACS can also be employed for the spin stabilization of sailcraft as well as as a backup to the conventional ACS of sail carrier spacecraft prior to sail deployment as well as during pre-flight sail checkout operation. An overview of the state-of-the-art microPPT technology, PPT performance requirements for solar sails, and pulse-modulated PPT control design and simulation results are presented in this paper.

Propellantless AOCS Design for a 160-m, 450-kg Sailcraft of the Solar Polar Imager Mission

An attitude and orbit control system (AOCS) is developed for a 160-m, 450-kg solar sail spacecraft of the Solar Polar Imager (SPI) mission. The SPI mission is one of several SunEarth Connections solar sail roadmap missions currently envisioned by NASA. A reference SPI sailcraft consists of a 160-m, 150-kg square solar sail, a 250-kg spacecraft bus, and 50-kg science payloads, The 160-m reference sailcraft has a nominal solar thrust force of 160 mN (at 1 AU), an uncertain center-of-mass/center-of-pressure offset of ±0. 4m ,and a characteristic acceleration of 0.35 mm/s 2 . The solar sail is to be deployed after being placed into an earth escaping orbit by a conventional launch vehicle such as a Delta II. The SPI sailcraft first spirals inwards from 1 AU to a heliocentric circular orbit at 0.48 AU, followed by a cranking orbit phase to achieve a science mission orbit at a 75-deg inclination, over a total sailing time of 6.6 yr. The solar sail will be jettisoned after achieving the science mission orbit. This paper focuses on the solar sailing phase of the SPI mission, with emphasis on the design of a reference AOCS consisting of a propellantless primary ACS and a microthruster-based secondary (optional) ACS. The primary ACS employs trim control masses running along mast lanyards for pitch/yaw control together with roll stabilizer bars at the mast tips for quadrant tilt (roll) control. The robustness and effectiveness of such a propellantless primary ACS would be enhanced by the secondary ACS which employs tip-mounted, lightweight pulsed plasma thrusters (PPTs). The microPPT-based ACS is mainly intended for attitude recovery maneuvers from off-nominal conditions. A relatively fast, 70-deg pitch reorientation within 3 hrs every half orbit during the orbit cranking phase is shown to be feasible, with the primary ACS, for possible solar observations even during the 5-yr cranking orbit phase.

Design and Simulation of Sailcraft Attitude Control Systems Using the Solar Sail Control Toolbox

A MATLAB-based software tool, called the Solar Sail Control Toolbox, has been developed for solar sail analysis, design and simulation. The toolbox includes functions for multibody dynamic modeling, attitude control systems design, thrust vector control design, orbit analysis, solar sail mission analysis, thermal analysis and power subsystem analysis. Sailcraft models can be created, analyzed and simulated in the toolbox without the need to deal with any other software tools. This paper demonstrates the toolbox by showing how a gossamer sailcraft with a moving mass actuator and spreader bar control system is modeled and simulated. The toolbox couples attitude and orbit dynamics into the same simulations. The user can choose from several different attitude dynamics models, including specially developed multibody models for the moving mass and gimbaled boom control systems, and from several orbit models including point mass, n-body and non-spherical earth. The moving mass, gimbaled boom, and non-spherical earth dynamics are developed in this paper. The moving mass model permits any number of moving masses that are constrained to have one degree of translational freedom. The gimbaled boom model includes two bodies with one two degree-of-freedom hinge following Hooker’s formulation. The interconnected rigid bodies have joints which allow only rotational motion. Both methods explicitly eliminate constrained degrees of freedom. The non-spherical Earth gravitational model employs the recursive non-singular (except r = 0) method of Mueller and Gottlieb. A disturbance modeling package is included which can be used independently or integrated with the solar sail simulations. Environmental disturbances are: optical disturbances including solar pressure radiation, earth albedo and earth radiation; magnetic and radio-frequency torques; thermal; and aerodynamics. This paper includes the mathematical formulations of the disturbances. It is also possible to include user models of components and disturbances, such as sail membrane models with center of pressure/center of mass offsets. Spacecraft models are created using a graphical computer-aided design package that is included in the toolbox. The component information defined in the CAD package is used to generate mass properties and for all disturbance calculations. Component-level data includes optical and thermal properties for the surfaces; mass, center-of-mass, and inertia; and magnetic dipoles, RF sources, etc. Additional information can also be stored with each component so that the CAD file serves as a database for all spacecraft model data. Simulations are script-based. Controllers are implemented digitally and are not part of the simulation right-hand-side.

AOCS Performance and Stability Validation for Large Flexible Solar Sail Spacecraft

Future solar sail missions, such as NASA’s Solar Polar Imager Vision, will require sails with dimensions on the order of 50-500 m. We are examining a square sail design with moving mass (trim control mass, TCM) and quadrant rotation primary actuators plus pulsed plasma thrusters (PPTs) at the mast tips for backup attitude control. Quadrant rotation is achieved via roll stabilizer bars (RSB) at the mast tips. At these sizes, given the gossamer nature of the sail supporting structures, flexible modes may be low enough to interact with the control system, especially as these actuators are located on the flexible structure itself and not on the rigid core. This paper develops a practical analysis of the flexible interactions using state-space systems and modal data from finite element models of the system. Torsion and bending of the masts during maneuvers could significantly affect the function of the actuators while activation of the membrane modes could adversely affect the thrust vector direction and magnitude. Analysis of the RSB and TCM dynamics for developing high-fidelity simulations is included. For control analysis of the flexible system, standard finite-element models of the flexible sail body are loaded and the modal data is used to create a modal coordinate state-space system. Key parameters include which modes to include, which nodes are of interest for force inputs and displacement outputs, connecting nodes through which external forces and torques are applied from the flex body to the core, any nominal momentum in the system, and any steady rates. The system is linearized about the nominal attitude and rate. The state-space plant can then be analyzed with a state-space controller, and Bode, Nyquist, step and impulse responses generated. The approach is general for any rigid core with a flexible appendage. This paper develops a compensator for a simple two-mass flex system and extrapolates the results to the solar sail. A finite element model of the 20 m solar sail by ATK Space Systems, recently validated in ground tests, is used to demonstrate the sail analysis approach.

Optical navigation system

exible navigation system for deep space operations that does not require GPS measurements. The navigation solution is computed using an Unscented Kalman Filter (UKF) that can accept any combination of range, rangerate, planet chord width, landmark and angle measurements using any celestial object. The UKF employs a full nonlinear dynamical model of the orbit including gravity models and disturbance models. The ONS package also includes attitude determination algorithms using the UKF algorithm with the Inertial Measurement Unit (IMU). The IMU is used as the dynamical base for the attitude determination algorithms. That is, the gyros model is propagated, not the spacecraft model. This makes the sensor a more capable plugin replacement for a star tracker, thus reducing the integration and test cost of adding this sensor to a spacecraft. The linear accelerometers are used to measure forces on the spacecraft. This permits accurate measurement of the accelerations applied by thrusters during maneuvers. The paper includes test results from three cases: a geosynchronous satellite, the New Horizons spacecraft and the Messenger spacecraft. The navigation accuracy is limited by the knowledge to the ephemerides of the measurement targets but is sucient for the purposes of orbit maneuvering.

System Design of a Reusable, Horizontal Take-Off/Horizontal Landing Two Stage to Orbit Vehicle

and horizontal landing allows for safe, rapid access to space. The end-to-end system design and preliminary trajectory optimization is presented. The feasibility of the concept is then demonstrated through a high-delity simulation.

Formations for Close-Orbiting Escort Vehicles

One concept for protecting and inspecting valuable space assets such as GPS satellites is the use of escort vehicles flying in formation with the asset. This paper presents a systematic analysis of nominal escort orbits and delta-V budgets for formation maintenance and asset inspection in LEO, GPS, and GEO orbits. Navigation accuracy and formation controller type are addressed as they affect the minimum size of the orbit and delta-V budget. The effects of differential drag, solar pressure, and the J2 gravitational harmonic are considered separately. Geometric constraints such as nadir-pointing payloads on the asset are also considered.The objectives of the escort vehicles are to maximize the protection space around the prime asset, maintain a safe separation distance from the prime and other escorts, avoid interference with the prime’s payloads, perform periodic inspections, and maximize the mission lifetime by minimizing the fuel consumption required for formation maintenance. Since the asset’s orbit is to be unchanged, the escort’s nominal orbits will be arrayed around the asset with the asset at the center of the formation. This is in contrast to a regular elliptical formation where all spacecraft are on the ellipse and will result in smaller separations for the same size ellipse. In order to achieve a large protection space and avoid the sensing/communication field of view of the prime, a passive formation is desired in which the escort orbits about the prime once each orbit, while oscillating back and forth in the cross-track direction. The cross-track amplitude involves orbital element differences which can potentially result in along-track drift due to the J2 perturbation. Selection of the element differences to avoid inclination differences serves to minimize the drift rate.The nominal escort geometry is an elliptical formation with a minimum separation distance of 1 km and a cross-track amplitude of about 0.707 km. The worst case secular drift rate is 13.2 m/hour, occurring at 42 degrees inclination in a 600 km LEO orbit. By defining the same formation with zero inclination difference and maximum difference in right ascension, the drift rate can be reduced to 2 mm/hour. For GEO orbits, even with the maximum inclination difference, the drift rate is less than 2.4 cm/hour. The drift rates scale linearly with the size of the escort’s relative ellipse. Delta-Vs to maintain these formations are on the order of 45-60 m/s/year. The contribution from J2 in LEO varies from less than 10 m/s to 350 m/s depending on the orbit parameters discussed above, but for many orbits can be compensated for with a semi-major axis difference. The contribution from drag (LEO only) is about 25 m/s. Solar pressure contributes about 30 and 43 m/s in LEO and GEO respectively. For both drag and solar pressure, the differential area was assumed to be 5 square meters, representative of a large space asset and a small escort vehicle.