Joseph B. Mueller

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.

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.

Avoidance Maneuver Planning Incorporating Station-Keeping Constraints and Automatic Relaxation

Space debris is a growing concern for the sustained operation of our satellites. The population in space is continually increasing, both on a gradual basis as new satellites are placed on orbit and in sudden bursts, as evidenced with the recent collision between the Iridium and inactive Cosmos spacecraft. The problem is most severe in densely populated orbit regimes, where many operational satellites face a sustained presence of close-orbiting objects. In general, the frequent occurrence of potential collisions with debris will have a negative impact on mission performance in two important ways. First, repeated avoidance maneuvers diminish fuel and thus reduce mission life. Second, excursions from the nominal orbit during avoidance maneuvers may violate mission requirements or payload constraints. It is therefore important to consider both fuel minimization and station-keeping objectives in the avoidance planning problem. In this paper, we formulate the avoidance maneuver planning problem as a linear prog...

Onboard Planning of Collision Avoidance Maneuvers Using Robust Optimization

As the amount of space debris accumulates, and as more future missions involve multiple close-orbiting spacecraft, the risk of collision is becoming a more prominent concern in both mission design and operations. The recent collision between the Iridium and inactive COSMOS spacecraft has shown demonstrated this growing problem, with NORAD already tracking over 200 pieces of new debris in just the rst few weeks since the event. Spacecraft with the capability to detect potential collisions, then plan and enact avoidance maneuvers can successfully mitigate this risk. A particular challenge with avoiding space debris is the large degree of uncertainty in the relative state of the objects. In this paper, we present a practical method for the onboard planning of collision avoidance maneuvers. The relative orbit dynamics are modeled as a discrete, linear time-varying system that models both circular and eccentric orbits. Minimum-fuel avoidance maneuvers are planned by using a robust linear programming (LP) technique. The original non-linear, non-convex avoidance constraints are transformed into a time-varying sequence of linear constraints, and the navigation uncertainty is applied in a worst-case sense. The resulting maneuver can be solved eciently as an LP with no integer constraints, and can guarantee collision avoidance with respect to bounded navigation uncertainty.

Decentralized Formation Flying Control in a Multiple-Team Hierarchy

Abstract: In recent years, formation flying has been recognized as an enabling technology for a variety of mission concepts in both the scientific and defense arenas. Examples of developing missions at NASA include magnetospheric multiscale (MMS), solar imaging radio array (SIRA), and terrestrial planet finder (TPF). For each of these missions, a multiple satellite approach is required in order to accomplish the large‐scale geometries imposed by the science objectives. In addition, the paradigm shift of using a multiple satellite cluster rather than a large, monolithic spacecraft has also been motivated by the expected benefits of increased robustness, greater flexibility, and reduced cost. However, the operational costs of monitoring and commanding a fleet of close‐orbiting satellites is likely to be unreasonable unless the onboard software is sufficiently autonomous, robust, and scalable to large clusters. This paper presents the prototype of a system that addresses these objectives—a decentralized guidance and control system that is distributed across spacecraft using a multiple team framework. The objective is to divide large clusters into teams of “manageable” size, so that the communication and computation demands driven by N decentralized units are related to the number of satellites in a team rather than the entire cluster. The system is designed to provide a high level of autonomy, to support clusters with large numbers of satellites, to enable the number of spacecraft in the cluster to change post‐launch, and to provide for on‐orbit software modification. The distributed guidance and control system will be implemented in an object‐oriented style using a messaging architecture for networking and threaded applications (MANTA). In this architecture, tasks may be remotely added, removed, or replaced post launch to increase mission flexibility and robustness. This built‐in adaptability will allow software modifications to be made on‐orbit in a robust manner. The prototype system, which is implemented in Matlab, emulates the object‐oriented and message‐passing features of the MANTA software. In this paper, the multiple team organization of the cluster is described, and the modular software architecture is presented. The relative dynamics in eccentric reference orbits is reviewed, and families of periodic, relative trajectories are identified, expressed as sets of static geometric parameters. The guidance law design is presented, and an example reconfiguration scenario is used to illustrate the distributed process of assigning geometric goals to the cluster. Next, a decentralized maneuver planning approach is presented that utilizes linear‐programming methods to enact reconfiguration and coarse formation keeping maneuvers. Finally, a method for performing online collision avoidance is discussed, and an example is provided to gauge its performance.

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.

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.

Unified GN&C system for the Space Rapid Transit launch vehicle

This paper presents the overall design of a guidance, navigation and control system for a novel two-stage to orbit launch vehicle. The Space Rapid Transit, or SRT, is being designed by Princeton Satellite Systems to address the most critical needs of the United States space program: namely the rapid, safe and economic delivery of payloads and astronauts into low Earth orbit. The launch of the vehicle progresses through four phases: 1) atmospheric ight with the aid of a ferry stage, 2) a boost phase to LEO, 3) orbit acquisition and rendezvous, and nally 4) reentry and gliding landing. In order to achieve accurate orbit insertion, the ferry stage must rst deliver the orbiter to a precise position and velocity at the correct time. Non-linear dynamic inversion techniques are used to design a guidance law for the powered ight phase. The boost phase uses a gimbaled main engine with classical feedback control to zero the disturbance torque. Once in orbit, the reaction control system utilizes 16 thrusters to perform attitude maneuvers and small orbit maneuvers for rendezvous. Relative orbit guidance is then performed by formulating a robust linear program that plans minimum-fuel maneuvers to achieve target states subject to avoidance constraints and initial state uncertainty. Another linear program is used with weighted slack variables to allocate pulsewidths to the thrusters so that the force and torque demands for orbit and attitude control are achieved as closely as possible, subject to the maximum thrust constraints. The overall design of the guidance and control systems for all ight phases is presented along with preliminary simulation results.