Carlos M. Roithmayr

Achieving Climate Change Absolute Accuracy in Orbit

The Climate Absolute Radiance and Refractivity Observatory (CLARREO) mission will provide a calibration laboratory in orbit for the purpose of accurately measuring and attributing climate change. CLARREO measurements establish new climate change benchmarks with high absolute radiometric accuracy and high statistical confidence across a wide range of essential climate variables. CLARREOs inherently high absolute accuracy will be verified and traceable on orbit to Systeme Internationale (SI) units. The benchmarks established by CLARREO will be critical for assessing changes in the Earth system and climate model predictive capabilities for decades into the future as society works to meet the challenge of optimizing strategies for mitigating and adapting to climate change. The CLARREO benchmarks are derived from measurements of the Earths thermal infrared spectrum (5–50 μm), the spectrum of solar radiation reflected by the Earth and its atmosphere (320–2300 nm), and radio occultation refractivity from which...

Attitude and Orbit Control of a Very Large Geostationary Solar Power Satellite

Design of an attitude- and orbit-control system is presented for a 3.2 x 3.2 km geostationary solar-array platform with an area-to-mass ratio of 0.4 m 2 /kg. The proposed control-system architecture utilizes electric thrusters for integrated attitude and orbit eccentricity control by counteracting, simultaneously, attitude-disturbance torques and a large orbital perturbing force. Significant control-structure interaction, possible for such a very large flexible structure, is avoided by employing a low-bandwidth attitude-control system. However, a concept of persistent-disturbance accommodating control is utilized to provide precision attitude control in the presence of dynamic modeling uncertainties and persistent external disturbances.

INTEGRATED ORBIT, ATTITUDE, AND STRUCTURAL CONTROL SYSTEM DESIGN FOR SPACE SOLAR POWER SATELLITES

The major objective of this study is to develop an integrated orbit, attitude, and structural control system architecture for very large Space Solar Power Satellites (SSPS) in geosynchronous orbit. This study focuses on the 1.2-GW “Abacus” SSPS concept characterized by a 3.2 × 3.2 km solar-array platform, a 500-m diameter microwave beam transmitting antenna, and a 500 × 700 m earth-tracking reflector. For this baseline Abacus SSPS configuration, we derive and analyze a complete set of mathematical models, including external disturbances such as solar radiation pressure, microwave radiation, gravity-gradient torque, and other orbit perturbation effects. The proposed control system architecture utilizes a minimum of 500 1-N electric thrusters to counter, simultaneously, the cyclic pitch gravity-gradient torque, the secular roll torque caused by an offset of the centerof-mass and center-of-pressure, the cyclic roll/yaw microwave radiation torque, and the solar radiation pressure force whose average value is about 60 N.

CLARREO Approach for Reference Intercalibration of Reflected Solar Sensors: On-Orbit Data Matching and Sampling

The implementation of the Climate Absolute Radiance and Refractivity Observatory (CLARREO) mission was recommended by the National Research Council in 2007 to provide an on-orbit intercalibration standard with accuracy of 0.3% (k = 2) for relevant Earth observing sensors. The goal of reference intercalibration, as established in the Decadal Survey, is to enable rigorous high-accuracy observations of critical climate change parameters, including reflected broadband radiation [Clouds and Earths Radiant Energy System (CERES)], cloud properties [Visible Infrared Imaging Radiometer Suite (VIIRS)], and changes in surface albedo, including snow and ice albedo feedback. In this paper, we describe the CLARREO approach for performing intercalibration on orbit in the reflected solar (RS) wavelength domain. It is based on providing highly accurate spectral reflectance and reflected radiance measurements from the CLARREO Reflected Solar Spectrometer (RSS) to establish an on-orbit reference for existing sensors, namely, CERES and VIIRS on Joint Polar Satellite System satellites, Advanced Very High Resolution Radiometer and follow-on imagers on MetOp, Landsat imagers, and imagers on geostationary platforms. One of two fundamental CLARREO mission goals is to provide sufficient sampling of high-accuracy observations that are matched in time, space, and viewing angles with measurements made by existing instruments, to a degree that overcomes the random error sources from imperfect data matching and instrument noise. The data matching is achieved through CLARREO RSS pointing operations on orbit that align its line of sight with the intercalibrated sensor. These operations must be planned in advance; therefore, intercalibration events must be predicted by orbital modeling. If two competing opportunities are identified, one target sensor must be given priority over the other. The intercalibration method is to monitor changes in targeted sensor response function parameters: effective offset, gain, nonlinearity, optics spectral response, and sensitivity to polarization. In this paper, we use existing satellite data and orbital simulation methods to determine mission requirements for CLARREO, its instrument pointing ability, methodology, and needed intercalibration sampling and data matching for accurate intercalibration of RS radiation sensors on orbit. We conclude that with the CLARREO RSS in a polar 90° inclination orbit at a 609-km altitude, estimated intercalibration sampling will limit the uncertainty contribution from data matching noise to 0.3% (k = 2) over the climate autocorrelation time period. The developed orbital modeling and intercalibration event prediction will serve as a framework for future mission operations.

Fuel-Optimal Transfers Between Coplanar Circular Orbits Using Variable-Specific-Impulse Engines

Minimum fuel circular-to-circular low-thrust orbital transfers with prescribed transfer time and with the thrust always directed parallel to the velocity vector are developed for a variable-specific-impulse system. The changeable specific impulse serves as the control variable. The power input is assumed to be constant, and the resulting thrust magnitude is inversely proportional to the chosen specific impulse. With these engine specifications, and under the assumption that the thrust magnitude always remains small compared to the gravitational forces, it is shown that the optimal specific impulse profile increases linearly with time. The associated optimal thrust magnitude is proportional to the vehicle mass; that is, the thrust per unit of vehicle mass remains constant. The mass fraction obtained with variable specific impulse is shown to be greater than or equal to that obtained with constant specific impulse; the advantage can be significant for orbital transfers with large changes in the semimajor axis.

Gravitational moment exerted on a small body by an oblate body

The present demonstration of a method for obtaining vector-dyadic expressions of the gravitational moment about a bodys center-of-mass proceeds through the derivation of an expression for the gravitational moment exerted by an oblate spheroid. The contribution of the earths oblateness to the gravitational moment exerted on a body has been numerically evaluated for a greatly simplified illustrative case; this contribution is noted to be significant by comparison with such other external moments as those exerted by aerodynamic forces.

Modeling Multibody Stage Separation Dynamics Using Constraint Force Equation Methodology

This paper discusses the application of the constraint force equation methodology and its implementation for multibody separation problems using three specially designed test cases. The first test case involves two rigid bodies connected by a fixed joint, the second case involves two rigid bodies connected with a universal joint, and the third test case is that of Mach 7 separation of the X-43A vehicle. For the first two cases, the solutions obtained using the constraint force equation method compare well with those obtained using industry- standard benchmark codes. For the X-43A case, the constraint force equation solutions show reasonable agreement with the flight-test data. Use of the constraint force equation method facilitates the analysis of stage separation in end-to-end simulations of launch vehicle trajectories

Co-spin with symmetry axis stabilization, and de-spin for asteroid capture

Consideration is given to attitude control associated with capturing a free-flying asteroid using an axisymmetric spacecraft. Asymptotically stable controllers are designed to align the spacecraft axis of symmetry with a line of descent that is fixed in the asteroid, and to eliminate all relative angular velocity before capture takes place. For the special case of an axisymmetric asteroid, an analytical expression is presented for the torque required to maintain alignment of the axes of symmetry of the spacecraft and the asteroid. After the asteroid is securely captured, the angular velocity of the rigid composite body relative to an inertial frame is arrested; we present a controller that is asymptotically stable and stays within specified thrust limits.

Application of Constraint Force Equation Methodology for Launch Vehicle Stage Separation

A = axial location of the solid rocket booster (SRB) reference point in Space Shuttle main engine (SSME) plume coordinate system, ft ax, ay, az = acceleration components along body axes (excluding components due to gravity), ft=s CA = isolated (freestream) axial force coefficient Cm = isolated (freestream) pitching moment coefficient CN = isolated (freestream) normal force coefficient Cn = isolated (freestream) yawing moment coefficient CY = isolated (freestream) side force coefficient Cl = isolated (freestream) rolling moment coefficient eA, eB = unit vectors in body A and body B, respectively Fp = plume impingement force, lb Fx, Fy, Fz = aerodynamic forces in axial, lateral, and normal directions, lb F CON A , F CON B = joint constraint force vector for body A and body B

An Argument Against Augmenting the Lagrangean for Nonholonomic Systems

Although it is known that correct dynamical equations of motion for a nonholonomic system cannot be obtained from a Lagrangean that has been augmented with a sum of the nonholonomic constraint equations weighted with multipliers, previous publications suggest otherwise. An example has been proposed in support of augmentation and purportedly demonstrates that an accepted method fails to produce correct equations of motion whereas augmentation leads to correct equations; this paper shows that in fact the opposite is true. The correct equations, previously discounted on the basis of a flawed application of the Newton-Euler method, are verified by using Kanes method and a new approach to determining the directions of constraint forces. A correct application of the Newton-Euler method reproduces valid equations.