Engineering topics
Publications and source records attributed to Cefola, P. J..
Kalman filtering techniques are combined with a semianalytical orbit generator to develop a sequential orbit determination algorithm. The algorithm is investigated for computational efficiency, accuracy, and radius of convergence by comparison with truth ephemerides and a Cowell special perturbations filters (GTDS). Test cases relevant to satellite navigation are examined.
Those factors which limit the usefulness of current implementations of the semianalytical approach are discussed. Numerical and analytical enhancements to the semianalytical approach are considered. A simple mathematical model is provided to estimate the computational speed of a semianalytical theory employing the suggested enhancements. The model can factor in current experience with semianalytical theories (integration stepsizes, quadrature orders, speed of recursive formulations, etc.) and the characteristics of the particular output requirement (observation span (or orbit determination interval), observation rate, observation model, etc.). Comparisons with numerical integration are suggested.
Analytical averaging techniques are used to expand the disturbing potential in the equinoctial coordinate frame by considering third body harmonics and zonal functions harmonics. General results are developed through applications of Legendre and associated Legendre polynomials and the Q sub nm functions for the gravitational potential.
Explicit expressions in equinoxial elements are derived for the n-th term in the disturbing potential for three-body and zonal harmonics. A method for recursive computation of the disturbing potential and the partial derivatives is proposed. Preliminary results for an arbitrary geopotential term are obtained.
Survey of averaging and multirevolution methods for long-term orbit prediction. A technical approach with the following features is recommended: (1) averaged variation-of-parameter equations, (2) analytical expressions for oblateness and third-body effects, (3) definite integrals for atmospheric drag and lunar effects (for long-period orbits), (4) nonsingular equinoctial element formulation, (5) multistep numerical integration processes, and (6) precise osculating-to-mean element transformation. Several orbital predictions illustrate the contribution of this technical approach to overall accuracy and efficiency. Future development of the analytical averaging method in nonsingular coordinates by automated manipulation of literal series is discussed.
Review of some results obtained in an effort to develop efficient, high-precision trajectory computation processes for artificial satellites by optimum selection of the form of the equations of motion of the satellite and the numerical integration method. In particular, the matching of a Gaussian variation-of-parameter (VOP) formulation is considered which is expressed in terms of equinoctial orbital elements and partially decouples the motion of the orbital frame from motion within the orbital frame. The performance of the resulting orbit generators is then compared with the popular classical Cowell/Gauss-Jackson formulation/integrator pair for two distinctly different orbit types - namely, the orbit of the ATS satellite at near-geosynchronous conditions and the near-circular orbit of the GEOS-C satellite at 1000 km.
The matrizant of the two-body problem is developed in terms of elements that are free from singularities for zero eccentricities and zero- and ninety-degree inclinations. Retrograde equinoctial elements eliminate the singularity for inclinations near 180 degrees, with only minor changes in the expressions for the matrizant. The 'single-averaged' variation-of-parameters equations for these elements are developed for third-body, oblateness, and drag effects. Higher order terms are included in the expansions for the third-body and oblateness potential. A computer program that uses these equations to predict orbital evolution is described. Numerical results are given for a near-circular orbit.
This paper investigates the equinoctial orbit elements for the two-body problem, showing that the associated matrices are free from singularities for zero eccentricities and zero and ninety degree inclinations. The matrix of the partial derivatives of the position and velocity vectors with respect to the orbit elements is given explicitly, together with the matrix of inverse partial derivatives, in order to facilitate construction of the matrizant (state transition matrix) corresponding to these elements. The Lagrange and Poisson bracket matrices are also given. The application of the equinoctial orbit elements to general and special perturbations is discussed.
Spacecraft trajectory initial conditions expressions as parameters dependent functions, applying results to Mars lander mission
Vector matrix second order sensitivity equation application to Mars entry guidance, performing numerical simulation of second order sensitivity guidance and tabulating results
Adaptive control for Mars entry based on sensitivity analysis
Vector-matrix differential equation for second order sensitivity coefficients applied to trajectory optimization, modeling and compensation and guidance and control
Adaptive control for Mars entry based on sensitivity analysis
Sensitivity guidance for entry into uncertain Mars atmosphere using onboard computer