Search NASA⌕ Search

SEARCH · Search NASA

Results for “Trajectory Design”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Design and Optimization of Steering Laws for Geocentric Solar Sailing

Solar sail technology has the potential to be a practical method of propulsion for future deep space Cubesats and small satellites. Although in general, deep space, heliocentric flight is the optimal regime for solar sailing, many near-term solar sail flight demonstrations have, or will occur in Earth-centered orbits. This is particularly true for low-cost, smallsat-based solar sail technology demonstration flights, such as the Planetary Society LightSail 1 and 2, and the upcoming NASA Advanced Composite Solar Sail System (ACS3). ACS3 will be the first spaceflight application of NASA’s newest smallsat deployable composite boom technology, and NASA’s first practical solar sail. Launch of ACS3 is currently anticipated for late 2021 or early 2022. In this paper, we will describe the challenges of operating solar sailing spacecraft in low Earth orbits, where, depending on initial orbital altitude and inclination, disturbances due to aerodynamic drag, eclipsing can be of the same order or greater than forces caused by solar radiation pressure. A notional smallsat-class solar sailing vehicle based on the ACS3 solar sail will be used as a reference for trajectory design and optimization studies. As a starting point we will assume initially circular orbits, and a locally optimal steering law designed to maximize the rate of change in semi major axis, similar to that described by McInnes. This will be generalized to orbits with arbitrary initial inclinations and non-zero eccentricities, and to account for atmospheric drag and eclipsing effects. Optimal steering laws for maximizing rate-of-change in orbit inclination and eccentricity will also be explored. Finally, we will examine potential for Earth-centered solar sailing mission applications beyond technology demonstrations, including: continuous observations from inside the Earth’s magnetotail; augmented geostationary capabilities from displaced geostationary orbits; and continuous polar observation for climate science investigations from pole-sitter orbits.

W. Keats Wilkie↗
NASA NTRS1989

Guidance, steering, load relief and control of an asymmetric launch vehicle

A new guidance, steering, and control concept is described and evaluated for the Third Phase of an asymmetrical configuration of the Advanced Launch System (ALS). The study also includes the consideration of trajectory shaping issues and trajectory design as well as the development of angular rate, angular acceleration, angle of attack, and dynamic pressure estimators. The Third Phase guidance, steering and control system is based on controlling the acceleration-direction of the vehicle after an initial launch maneuver. Unlike traditional concepts, the alignment of the estimated and commanded acceleration-directions is unimpaired by an add-on load relief. Instead, the acceleration-direction steering-control system features a control override that limits the product of estimated dynamic pressure and estimated angle of attack. When this product is not being limited, control is based exclusively on the commanded acceleration-direction without load relief. During limiting, control is based on nulling the error between the limited angle of attack and the estimated angle of attack. This limiting feature provides full freedom to the acceleration-direction steering and control to shape the trajectory within the limit, and also gives full priority to the limiting of angle of attack when necessary. The flight software concepts were analyzed on the basis of their effects on pitch plane motion.

Boelitz, Frederick W.↗
NASA NTRS2007

Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

An overview of the aerothermodynamic environments definition status is presented for the Mars Science Laboratory entry vehicle. The environments are based on Navier-Stokes flowfield simulations on a candidate aeroshell geometry and worst-case entry heating trajectories. Uncertainties for the flowfield predictions are based primarily on available ground data since Mars flight data are scarce. The forebody aerothermodynamics analysis focuses on boundary layer transition and turbulent heating augmentation. Turbulent transition is expected prior to peak heating, a first for Mars entry, resulting in augmented heat flux and shear stress at the same heatshield location. Afterbody computations are also shown with and without interference effects of reaction control system thruster plumes. Including uncertainties, analysis predicts that the heatshield may experience peaks of 225 W/sq cm for turbulent heat flux, 0.32 atm for stagnation pressure, and 400 Pa for turbulent shear stress. The afterbody heat flux without thruster plume interference is predicted to be 7 W/sq cm on the backshell and 10 W/sq cm on the parachute cover. If the reaction control jets are fired near peak dynamic pressure, the heat flux at localized areas could reach as high as 76 W/sq cm on the backshell and 38 W/sq cm on the parachute cover, including uncertainties. The final flight environments used for hardware design will be updated for any changes in the aeroshell configuration, heating design trajectories, or uncertainties.

Edquist, Karl T.↗
DOE OSTI2011

Space Charge Correction on Emittance Measurement of Low Energy Electron Beams

The goal of any particle accelerator is to optimize the transport of a charged particle beam along a set path by confining the beam to a small region close to the design trajectory and directing it accurately along the beamline. To do so in the simplest fashion, accelerators use a system of magnets that exert approximately linear electromagnetic forces on the charged beam. These electromagnets bend the beam along the desired path, in the case of bending magnets, and constrain the beam to the desired area through alternating focusing and defocusing effects, in the case of quadrupole magnets. We can model the transport of such a beam through transfer matrices representing the actions of the various beamline elements. However, space charge effects, produced from self electric fields within the beam, defocus the beam and must be accounted for in the calculation of beam emittance. We present below the preliminary results of a MATLAB code built to model the transport of a charged particle beam through an accelerator and measure the emittance under the influence of space charge effects. We demonstrate the method of correctly calculating the emittance of a beam under space charge effects using a least square fit to determine the initial properties of the beam given the beam size measured at a specific point after transport.

43 PARTICLE ACCELERATORS↗
NASA NTRS2017

Automated Sensitivity Analysis of Interplanetary Trajectories for Optimal Mission Design

This work describes a suite of Python tools known as the Python EMTG Automated Trade Study Application (PEATSA). PEATSA was written to automate the operation of trajectory optimization software, simplify the process of performing sensitivity analysis, and was ultimately found to out-perform a human trajectory designer in unexpected ways. These benefits will be discussed and demonstrated on sample mission designs.

trajectory design↗
DOE OSTI2021

Directional locking and the influence of obstacle density on skyrmion dynamics in triangular and honeycomb arrays

In this work, we numerically examine the dynamics of a single skyrmion driven over triangular and honeycomb obstacle arrays at zero temperature. The skyrmion Hall angle θ sk , defined as the angle between the applied external drive and the direction of the skyrmion motion, increases in quantized steps or continuously as a function of the applied drive. For the obstacle arrays studied in this work, the skyrmion exhibits two main directional locking angles of θ sk = –30° and –60°. We show that these directions are privileged due to the obstacle landscape symmetry, and coincide with channels along which the skyrmion may move with few or no obstacle collisions. Here we investigate how changes in the obstacle density can modify the skyrmion Hall angles and cause some dynamic phases to appear or grow while other phases vanish. This interesting behavior can be used to guide skyrmions along designated trajectories via regions with different obstacle densities. For fixed obstacle densities, we investigate the evolution of the locked θ sk = –30° and –60° phases as a function of the Magnus force, and discuss possibilities for switching between these phases using topological selection.

75 CONDENSED MATTER PHYSICS, SUPERCONDUCTIVITY AND↗
DOE OSTI2019

Modelling Curvilinear Beamline Effects on Beam- and Spin-Dynamics in the Fermilab Muon g-2 Storage Ring.

A model of the Fermilab Muon g-2 Experiment storage ring is developed with the BMAD simulation library for beam- and spin-dynamics studies. Particular focus is given to the electrostatic quadrupoles, the plates of which follow the curvature of the particle design trajectory of the magnetically uniform storage ring. To model these, standard electric multipoles and a novel method based on cylindrical harmonics are investigated and found to accurately match the electric field maps produced for the bulk and fringe regions of the curved quadrupoles respectively, despite these methods being specifically for straight beamlines. The toroidal multipole description is then developed, to include curvature and longitudinal field variations (to model the fringe field regions). Various combinations of the different descriptions (including use of the field maps) are used for the bulk and fringe fields of the quadrupoles, and the toroidal multipoles are found to provide the best agreement with the measured tunes of the storage ring. Two integrators are also developed, based on the method of Wu, Forest and Robin. The first of these describes motion in a uniform magnetic dipole field with arbitrary electric field, and produces results comparable to those of the adaptive-step fourth-order Runge–Kutta integrator implemented in BMAD, suitable for tracking over thousands of turns. The second is a generalisation of the first to arbitrary magnetic vector potentials superimposed on that of the uniform dipole, with arbitrary electric fields, and was also found to agree with the Runge–Kutta integrator. Simulations of antimuon storage for 2000 turns of the storage ring were performed, where losses were studied along with the evolution of beam polarisation. The storage losses, important for reducing the systematic error in the measurement of the muon g-2, were found to be largely due to collimator shape. The correlation between momentum and spin precession frequency reflect the effects expected from the electric field correction in the T-BMT equation, on the order of 10 -6 . Particle injection was also studied, for various injection angles, quadrupole strengths and strengths of injection kick, and in the cases where experimental data was provided, the model was found to be consistent with the experiment.

43 PARTICLE ACCELERATORS↗
DOE OSTI2022

Synchronizing ESR and HSR for collisions at IP6 and IP8

Bunch circulation times in the ESR and HSR must be identical, whatever the electron and hadron energies. A particular bunch in either ring can only have collisions at either IP6 or IP8, but not at both. Synchronization is achieved by co-ordinating the design trajectory circumferences of the two rings, and by controlling the HSR and ESR path lengths from IP6 to IP8. Radial shift control of the HSR circumference is summarized. The conditions for synchronization are presented. The impact of operating HSR at low proton energy with 4 inner and 2 outer arcs is assessed. Changes of the ESR circumference due to the use of super-bends at the lowest electron energies are evaluated. The effect of J x -tuning on the ESR circumference and its compensation by an HSR radial shift are discussed. Finally, the potential effects of using a Ring Cooler for hadron cooling are considered.

43 PARTICLE ACCELERATORS↗
NASA NTRS1973

Low-thrust mission risk analysis, with application to a 1980 rendezvous with the comet Encke

A computerized failure process simulation procedure is used to evaluate the risk in a solar electric space mission. The procedure uses currently available thrust-subsystem reliability data and performs approximate simulations of the thrust sybsystem burn operation, the system failure processes, and the retargeting operations. The method is applied to assess the risks in carrying out a 1980 rendezvous mission to the comet Encke. Analysis of the results and evaluation of the effects of various risk factors on the mission show that system component failure rates are the limiting factors in attaining a high mission relability. It is also shown that a well-designed trajectory and system operation mode can be used effectively to partially compensate for unreliable thruster performance.

Yen, C. L.↗
NASA NTRS1973

Low-thrust mission risk analysis.

A computerized multi-stage failure process simulation procedure is used to evaluate the risk in a solar electric space mission. The procedure uses currently available thrust-subsystem reliability data and performs approximate simulations of the thrust subsystem burn operation, the system failure processes, and the retargetting operations. The application of the method is used to assess the risks in carrying out a 1980 rendezvous mission to Comet Encke. Analysis of the results and evaluation of the effects of various risk factors on the mission show that system component failure rates is the limiting factor in attaining a high mission reliability. But it is also shown that a well-designed trajectory and system operation mode can be used effectively to partially compensate for unreliable thruster performance.

Yen, C. L.↗
NASA NTRS1973

Interaction of the solar wind with the outer planets.

The hypersonic analog for the interaction of the solar wind with Jupiter, Saturn, Uranus, Neptune, and Pluto is used to provide estimates of shock shapes and locations, as well as average magnetosheath and/or ionosheath properties for these planets. Several representative spacecraft flyby trajectories (designed for outer-planet 'Grand Tour' simulations) are superimposed upon a series of figures in order to provide estimates of potential plasma and field parameters which may be encountered. Consideration is given first to the possibility that several of these planets have intrinsic magnetic fields and, secondly, to the interaction of the solar wind directly on the ionosphere should there actually be no intrinsic field. Saturn and Pluto are chosen as examples of this latter case.

Dryer, M.↗
NASA NTRS1974

Investigation of new radar-data-reduction techniques used to determine drag characteristics of a free-flight vehicle

An investigation was conducted of new techniques used to determine the complete transonic drag characteristics of a series of free-flight drop-test models using principally radar tracking data. The full capabilities of the radar tracking and meteorological measurement systems were utilized. In addition, preflight trajectory design, exact kinematic equations, and visual-analytical filtering procedures were employed. The results of this study were compared with the results obtained from analysis of the onboard, accelerometer and pressure sensor data of the only drop-test model that was instrumented. The accelerometer-pressure drag curve was approximated by the radar-data drag curve. However, a small amplitude oscillation on the latter curve precluded a precise definition of its drag rise.

Woodbury, G. E.↗
NASA NTRS1974

Data handling and control of the Mariner Venus Mercury '73 spacecraft

Data handling and control subsystems are described for the MVM73 spacecraft, which demonstrated the feasibility of multiple-planet gravity-assist missions. In spite of strict budgetary constraints, MVM73 carried out experiments utilizing television, an IR radiometer, a UV spectrometer, a fluxgate magnetometer, a charged particle telescope, a modulation-demodulation unit, an X-band and S-band dual frequency communications system, a steerable high-gain antenna, a time-sharing digital articulation and pointing system, and moveable solar panels. The trajectory design and flight data system are described, and the systems governing articulation and pointing and attitude control are discussed.

Scull, J. R.↗
NASA NTRS1978

Viking first encounter of Phobos - Preliminary results

Viking Orbiter-1 (VO-1) made a series of close flybys of the Martian satellite Phobos in February and May 1977. A description is presented of the results obtained during the flybys in February. The flyby geometries for the encounter period in February are shown in a graph. The trajectory design gave flybys on the illuminated side of Phobos within 80 to 300 km during the entire encounter period. The primary encounter observations of Phobos included visual and infrared imaging as well as radio tracking of VO-1 while it was under the gravitational influence of Phobos. Visual imaging was obtained from two narrow-angle television cameras. Infrared observations were obtained from an infrared thermal mapper. Radio data included S- and X-band Doppler and ranging data to VO-1 with a 10-second Doppler count. Assuming for Phobos a volume of 500 + or - 900 cu km, a mean density of 1.9 + or - 0.6 g/cu cm is obtained for it on the basis of the processed data.

Tolson, R. H.↗
NASA NTRS1979

The first libration-point satellite - Mission overview and flight history

On August 12, 1978, a scientific spacecraft called International Sun-Earth Explorer-3 (ISEE-3) was launched towards the interior sun-earth libration point, L1. The spacecraft was placed into a 'halo orbit' around the L1 point on November 20, 1978, thus becoming the first libration-point satellite. During its 100-day transfer trajectory, ISEE-3 lingered in a region where the gravitational effects of the sun and the earth are comparable, leading to some interesting tradeoffs concerning the maneuver strategy for halo-orbit insertion. Following orbit insertion, stationkeeping maneuvers were required to maintain the delicate equilibrium in the halo orbit. Details are presented for all of the velocity change maneuvers that were executed prior to the completion of the first halo orbit on May 14, 1979. Orbit selection, trajectory design, and the scientific objectives of the ISEE-3 mission are also discussed.

Farquhar, R. W.↗