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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 37 records · Page 2

Equinoctial orbit elements - Application to artificial satellite orbits.

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.

Cefola, P. J.↗

Simulation of a navigator algorithm for a low-cost GPS receiver

The analytical structure of an existing navigator algorithm for a low cost global positioning system receiver is described in detail to facilitate its implementation on in-house digital computers and real-time simulators. The material presented includes a simulation of GPS pseudorange measurements, based on a two-body representation of the NAVSTAR spacecraft orbits, and a four component model of the receiver bias errors. A simpler test for loss of pseudorange measurements due to spacecraft shielding is also noted.

Hodge, W. F.↗

Astrodynamics. Volume 1 - Orbit determination, space navigation, celestial mechanics.

Essential navigational, physical, and mathematical problems of space exploration are covered. The introductory chapters dealing with conic sections, orientation, and the integration of the two-body problem are followed by an introduction to orbit determination and design. Systems of units and constants, as well as ephemerides, representations, reference systems, and data are then dealt with. A detailed attention is given to rendezvous problems and to differential processes in observational orbit correction, and in rendezvous or guidance correction. Finally, the Laplacian methods for determining preliminary orbits, and the orbit methods of Lagrange, Gauss, and Gibbs are reviewed.

Herrick, S.↗

Orbital resonances in the solar system

Orbital resonances are defined as any system of two or more satellites (including planets) orbiting the same primary and whose orbital mean motions are in a ratio of small whole numbers. Known orbital resonances in the solar system are identified, including those involving Jupiter's satellites Io, Europa, and Ganymede; Saturn's satellites Mimas and Tethys, Enceladus and Dione, and Titan and Hyperion; Saturn's ring gaps and Mimas; various asteroids and Jupiter; and the planets Neptune and Pluto. The stability of orbital resonances is examined, the origin of orbital commensurabilities is investigated, and a simple model of the simplest kind of eccentricity-type resonance is outlined. A method is described by which tides carry a noncommensurate pair of satellites into a stable libration, and current ideas concerning the formation of the gaps in Saturn's rings and the asteroid belt are discussed. Various approaches to the analysis of orbital resonances are laid out and illustrated. Three two-body commensurabilities in Saturn's satellite system are analyzed numerically.

Peale, S. J.↗

A solution of the variational equations for elliptic orbits in rotating coordinates

For elliptic reference orbits, formulas are given for the perturbation state transition matrix of the two-body problem. The formulas relate perturbations expressed in a local vertical rotating coordinate system and are valid for motion in the linear neighborhood of reference orbits with e in the range of 0 to 1. The elements of the state transition matrix are expressed in terms of natural parameters (horizontal and radial velocity, radius, eccentricity, true anomaly, etc.) at the initial and final points. In addition to the general form, a simplified version, valid for small eccentricity orbits, is given.

Jones, J. B.↗

Post-Newtonian gravitational radiation from orbiting point masses

General formulas are derived which describe the gravitational radiation at large distances from a system of bodies whose sizes are small compared with their separations. The calculation is carried out through post-Newtonian order within general relativity. More explicit formulas are derived for two-body systems, and detailed results are presented for circular orbits, gravitational bremsstrahlung, and head-on collisions.

Wagoner, R. V.↗

Pines method for circular and elliptical orbits

The Pines method for trajectory integration is derived from the equations of two-body motion. The concept of the perturbation derivative is explained and a formal definition is presented. An algorithm for applying the Pines method to numerical integration of trajectories is included.

Semar, C.↗

TBERR - Two-body error analysis program

Two-body error analysis computer program to evaluate resulting state vector and covariance matrix for orbiter after one coast and one burn maneuver

Muhonen, D. P.↗

An analytical state transition matrix for an orbit perturbed by oblateness

A new analytical state transition matrix is presented. This transition matrix contains the two-body terms and the secular and short periodic terms due to the J2 oblateness perturbation. The matrix is derived from a satellite theory which uses a set of 8 canonical elements with the true anomaly as the independent variable. This theory was chosen because of its accuracy and concise formulation. It presents no problems concerning the accuracy of the mean motion, which is typical for classical satellite theories. Numerical experiments show the J2 transition matrix to be superior to the simple two-body matrix.

Mueller, A. C.↗

Some three-body numerical solutions for low-thrust orbiter missions.

A ?velocity at the sphere of influence' method and an asymptotic matching method of patching together two-body low thrust solutions are compared to a number of three-body numerical results for outer planet orbiter missions. The two patching methods compare well with the numerical three-body results and do not depend on any particular choice for the size of the sphere of influence. The results apply, in a strict sense, only to the operational mode used-high-thrust terminal retro into orbit. Low-thrust spirals are not considered in the three-body analysis. The terminal low-thrust phase of thrusting is almost entirely reverse to the velocity vector during the planet centered phase of the trajectory. This may lead to important simplifications for low-thrust guidance and navigation procedures.

Mackay, J. S.↗

Application of MACSYMA to first order perturbation theory in celestial mechanics

The application of MACSYMA to general first order perturbation theory in celestial mechanics is explored. Methods of derivation of small variations in the Keplerian orbital elements are developed. As an example of the methods, the small general relativistic perturbations on the two-body Newtonian motion, resulting from the rotation of the central body, are developed in detail.

Anderson, J. D.↗

Three-Dimensional Lunar Mission Studies

Some three-dimensional lunar trajectories have been calculated by integration of the equations of motion of the classical restricted three-body problem of celestial mechanics. The calculations have been used for analysis of several aspects of lunar flight including requirements for achieving lunar impact and for establishment of a close lunar satellite. The allowable errors in initial conditions for lunar missions are strongly dependent on the values of the initial injection velocity and the injection angle. There can be large differences in results obtained from two-dimensional analyses (in which the vehicle trajectory is assumed to remain always in the earth-moon plane) and those obtained from three-dimensional analyses. Some of the accuracy tolerances can be fairly well estimated by use of a two-body analysis which considers the inclination of the plane of the vehicle trajectory to the earth-moon plane. Satisfactory orbits for a relatively close lunar satellite can be obtained with accuracies in the initial conditions approximately equal to those required for lunar impact.

Michael, William H., Jr.↗

Computation of solar perturbations with Poisson series

Description of a project for computing first-order perturbations of natural or artificial satellites by integrating the equations of motion on a computer with automatic Poisson series expansions. A basic feature of the method of solution is that the classical variation-of-parameters formulation is used rather than rectangular coordinates. However, the variation-of-parameters formulation uses the three rectangular components of the disturbing force rather than the classical disturbing function, so that there is no problem in expanding the disturbing function in series. Another characteristic of the variation-of-parameters formulation employed is that six rather unusual variables are used in order to avoid singularities at the zero eccentricity and zero (or 90 deg) inclination. The integration process starts by assuming that all the orbit elements present on the right-hand sides of the equations of motion are constants. These right-hand sides are then simple Poisson series which can be obtained with the use of the Bessel expansions of the two-body problem in conjunction with certain interation methods. These Poisson series can then be integrated term by term, and a first-order solution is obtained.

Broucke, R.↗

Program manual for HILTOP, a heliocentric interplanetary low thrust trajectory optimization program. Part 1: User's guide

A performance-analysis computer program, that was developed explicitly to generate optimum electric propulsion trajectory data for missions of interest in the exploration of the solar system is presented. The program was primarily designed to evaluate the performance capabilities of electric propulsion systems, and in the simulation of a wide variety of interplanetary missions. A numerical integration of the two-body, three-dimensional equations of motion and the Euler-Lagrange equations was used in the program. Transversality conditions which permit the rapid generation of converged maximum-payload trajectory data, and the optimization of numerous other performance indices for which no transversality conditions exist are included. The ability to simulate constrained optimum solutions, including trajectories having specified propulsion time and constant thrust cone angle, is also in the program. The program was designed to handle multiple-target missions with various types of encounters, such as rendezvous, stopover, orbital capture, and flyby. Performance requirements for a variety of launch vehicles can be determined.

Mann, F. I.↗

A Passive Gravitational Attitude Control System for Satellites

It is shown how the gravity-gradient effect may be utilized to design a long-lived, earth-pointing satellite attitude control system which requires no fuel supplies, attitude sensors or active control equipment. This two-body system is provided with a magnetic hysteresis damper which effectively damps out oscillations (librations) about the local vertical. The long rods, which must be extended in space from coiled up metal tapes, provide the required large moments of inertia and possess adequate rigidity and sufficient strength to endure the rigors of the extension process. The system is compatible with the requirements of multiple satellite launchings from a single last-stage vehicle. Analysis indicates that the gravitational torques are sufficient to keep the disturbing effects of solar radiation pressure, residual magnetic dipole moments, orbit eccentricity, rod curvature, eddy currents, and meteorite impacts within tolerable limits. It is believed that the high-performance, earth-pointing system described and analyzed in this paper represents an essential step in the development of high-capacity communications satellites requiring long life.

GRAVITY↗

Preliminary Survey of Retrograde Velocities Required for Insertion Into Low-Altitude Lunar Orbits

Closed lunar orbits are envisaged in lunar mission programs. The study described herein was undertaken to obtain an appreciation of the relevant fuel consumption requirements. The retrograde impulses necessary for establishing the orbits were assumed to occur at the point of closest approach of the main earth-moon trajectory; this point, designated as the arrival position, was restricted to a lunar altitude of 5,000 nautical miles or less. The orientation of the arrival position vector relevant to any coplanar radius vector is not constrained, however, and similarly the scalar value of the arrival velocity is unrestrained. Since the arrival altitude is restricted to 5,000 nautical miles or less, the perturbing accelerations of the earth and sun are sufficiently small that the vehicle and moon essentially comprise an isolated two-body system; this is discussed in the report. Retrograde velocities are determined for any required pericynthion position. If the pericynthion orientation requirement is relaxed then a smaller retrograde velocity is in some cases possible. A comparison between minimum retrograde velocities and retrograde velocities necessary for stipulated pericynthion positions is given. Arrival velocities are correlated with feasible earth departure conditions. The equations developed for determining retrograde velocities for desired pericynthion positions are considered useful for estimating essential data for the preliminary planning of lunar missions. Some graphical representation is included herein for immediate familiarization with possible conditions.

Jenkins, Morris V.↗

Spinning black hole scattering at $ \mathcal{O} $(G 3 S 2 )(G 3 S 2 ): Casimir terms, radial action and hidden symmetry

We resolve subtleties in calculating the post-Minksowskian dynamics of binary systems, as a spin expansion, from massive scattering amplitudes of fixed finite spin. In particular, the apparently ambiguous spin Casimir terms can be fully determined from the gradient of the spin-diagonal part of the amplitudes with respect to S 2 = −s(s+1)ħ 2 , using an interpolation between massive amplitudes with different spin representations. From two-loop amplitudes of spin-0 and spin-1 particles minimally coupled to gravity, we extract the spin Casimir terms in the conservative scattering angle between a spinless and a spinning black hole at $ \mathcal{O} $(G 3 S 2 ), finding agreement with known results in the literature. This completes an earlier study [Phys. Rev. Lett. 130 (2023), 021601] that calculated the non-Casimir terms from amplitudes. We also illustrate our methods using a model of spinning bodies in electrodynamics, finding agreement between scattering amplitude predictions and classical predictions in a root-Kerr electromagnetic background up to $ \mathcal{O} $(α 3 S 2 ). For both gravity and electrodynamics, the finite part of the amplitude coincides with the two-body radial action in the aligned spin limit, generalizing the amplitude-action relation beyond the spinless case. Surprisingly, the two-loop amplitude displays a hidden spin-shift symmetry in the probe limit, which was previously observed at one loop. We conjecture that the symmetry holds to all orders in the coupling constant and is a consequence of integrability of Kerr orbits in the probe limit at the first few orders in spin.

Classical Theories of Gravity↗

On the matrizant of the two-body problem

Matrizant of two body problem or integrable dynamic system for applications to computing perturbations and differential correction of orbits with observations

Broucke, R. A.↗