The Experts below are selected from a list of 72 Experts worldwide ranked by ideXlab platform

Anthony J. Calise - One of the best experts on this subject based on the ideXlab platform.

  • Optimization of Launch Vehicle Ascent Trajectories with Path Constraints and Coast Arcs
    Journal of Guidance Control and Dynamics, 2001
    Co-Authors: Peter F. Gath, Anthony J. Calise
    Abstract:

    This paper describes improvements made to a hybrid analytic/numerical algorithm for optimization of launch vehicle trajectories. Modifications are described which improve the accuracy of the solution. In addition, the algorithm has been extended to include path constraints for normal force and angle of attack, and the terminal constraints have been generalized to allow optimal attachment to a target orbit defined by inclination, apogee Radius and Perigee Radius. Singularities that occur for circular orbits and for equatorial orbits are identified. Finally, necessary conditions are derived and applied for the introduction of coasting arcs that are typically required for a great variety of missions.

  • Further improvements to a hybrid method for launch vehicle ascent trajectory optimization
    18th Applied Aerodynamics Conference, 2000
    Co-Authors: Anthony J. Calise, Sunil Tandon, Daniel Young, Sungwan Kim
    Abstract:

    This paper describes improvements made to a hybrid analytic/numerical algorithm for optimization of launch vehicle trajectories. Modifications are described that improve the speed and accuracy of the solution. A free time formulation is addressed in which coast arcs are optimized as a part of the ascent trajectory. The insertion of a coast arc along the final orbit allows the terminal constraints to be expressed in terms of the conditions at Perigee, which significantly simplifies the expressions for both the constraints and the transversality conditions. In addition, argument of Perigee for the final orbit has been added as a terminal constraint, which is useful in considering ascent to a geo-synchronous transfer orbit. Results are given for several realistic launch scenarios that illustrate the solution process, and assess the efficiency of the numerical procedure. Professor, Fellow AIAA, Email: anthony.calise@ae.gatech.edu t Senior Engineer/Scientist, Member AIAA, Email: sunil.tandon@boeing.com * Senior Manager, Member AIAA, Email: daniel.h.young2@boeing.com § Senior Engineer/Scientist, Senior Member AIAA, Email: sungwan.kim@boeing.com Copyright © 2000 by the American Institute of Aeronautics and Astronautics. All rights reserved. Nomenclature A Aerodynamic force in axial direction Ae Nozzle exit area a Semi-major axis aa Axial acceleration max Maximum axial acceleration ce Nozzle exit velocity E Orbital energy e Eccentricity H Hamiltonian h Orbital angular momentum i Inclination m Mass M Mach number N Aerodynamic force in normal direction Q Position costate vector P Velocity costate vector/primer vector pa Atmospheric pressure pv Collocation variables for velocity state qR Collocation variables for position costate qv Collocation variables for velocity costate R Radius vector from Earth center Re Earth Radius ra Apogee Radius rp Perigee Radius T Thrust Tvac Maximum available vacuum thrust V Inertial velocity vector a Angle of attack P Angle between final Radius and velocity vector e Constraint multiplier ATM Mass costate l American Institute of Aeronautics and Astronautics (c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. p, Gravitational constant co Schuler frequency C0p Argument of Perigee lb Unity vector along body axis If Unity vector perpendicular to orbit plane 1N Unity vector pointing north ln Unity vector perpendicular to body axis IP Unity vector along primer vector 1R Unity vector along Radius vector lv Unity vector along velocity vector

Vladimir S. Aslanov - One of the best experts on this subject based on the ideXlab platform.

  • Tether-assisted re-entry capsule deorbiting from an elliptical orbit
    Acta Astronautica, 2017
    Co-Authors: Vladimir S. Aslanov, Alexander S. Ledkov
    Abstract:

    Abstract The problem of a payload delivery from an elliptical orbit to Earth surface by a space tethered system is considered. This operation includes three stages: tethered motion, orbital flight of the payload to the border of the atmosphere, and its descent into the atmosphere. The third stage is critical for the whole mission, because the payload can be damaged as a result of thermal or dynamic loads. Thus, the first stage should provide the transition of the payload into a trajectory with the permissible loads. The objective of the paper is to determinate the conditions of the payload separation from the tether, which provide minimization of dynamical or thermal loads during the atmospheric motion or minimization of a Perigee Radius of the payload orbit. The mathematical models for the each stage and equations for loads calculation are presented in the paper. The series of numerical calculation shows that eccentricity of the satellite orbit significantly affects the trajectory of the capsule with the payload in the atmosphere. The location of separation points on elliptical orbits and angles of the tether deflection that provide a transfer of the capsule to the trajectory with minimum Perigee Radius, dynamical pressure, heat flux, and total heat were found. In contrast to the case of a circular orbit, the separation of the capsule at the local vertical does not provide a transfer into an orbit with a minimum Radius of Perigee. It was observed that the increase in eccentricity can leads to the decrease of the Radius of Perigee and the total heat, but at the same time to growth of the dynamic and thermal loads. The Radius of Perigee of the re-entry capsule cannot be considered as a single indicator of efficiency of the tether deployment control law. The set of thermal and mechanical constraints must be taken into consideration at the stages of the tether control law and the re-entry capsule design developing.

  • Swing principle in tether-assisted return mission from an elliptical orbit
    Aerospace Science and Technology, 2017
    Co-Authors: Vladimir S. Aslanov, Alexander S. Ledkov
    Abstract:

    Abstract The problem of a tether-assisted payload return from an elliptical orbit is considered in this study. In contrast to the existing works devoted to this issue, the article deals with a tether length control that provides a transfer of the payload into a descent trajectory from the tether rotation mode. Application of the swing principle for the tether control is investigated. The simplified mathematical model of the space tethered system is developed. It is shown that the stable limit cycle could exist under the considered control. The approximate analytical solution for this cycle is obtained. The stability of this solution is studied by the Lyapunov's theorems. The optimal control, which provides transfer of the payload into the descent trajectory with minimum Perigee Radius, is found as a result of the simulation series. It is shown that the tether should occur several turns before the payload separation. For example, in the YES-2 experiment, it is demonstrated that proposed control makes it possible to perform a payload return using a tether of considerably shorter length. The main conclusion of the paper is that the proposed scheme of the payload deorbit is more effective than the classical static or dynamic tether deployment schemes.

  • Deployment, Dynamics, and Control of a Tether-Assisted Return Mission of a Reentry Capsule
    Rigid Body Dynamics for Space Applications, 2017
    Co-Authors: Vladimir S. Aslanov
    Abstract:

    This chapter is devoted to a tether-assisted reentry capsule return mission. The mathematical model of the space tethered system is given. Simplified equations are obtained for the case of the deployed tether. Analytic solutions for satellite motion around its center of mass under the influence of the gravitational moment and tether tension force are given for the cases of slow change of tension force and tether deflection from the local vertical, of small oscillations of the satellite axis around the tether line, of the satellite oscillations near local vertical. Evaluation of microaccelerations on board the satellite is carried out using these solutions. A particular case of the satellite oscillations with a vertical elastic tether in a circular orbit is considered. The bifurcation diagram is constructed, and the hetero- and homoclinic separatrix trajectories are determined. Melnikov's method is used to study the satellite chaotic behavior near separatrices under the action of the periodic tether tension force. It is shown that elastic tether oscillations can result in appearance of chaotic modes of spacecraft motion. The condition of chaos existence, which allows to determine the measure of damping sufficient for prevention of chaotic modes, is obtained. The influence of mass, geometric, and elastic characteristics on the form of phase portrait and on the value of periodic disturbance, caused by oscillations of the elastic vertical tether, is studied. Influence of the orbit eccentricity on the space tether system motion is considered in this chapter. The control law that allows to increase the deflection angle of the tether from the local vertical is offered. The control law is based on a principle of a swing. An approximate analytic solution for the deflection angle from the local vertical is obtained for the control law, and stability of the tether vertical position under the proposed control is investigated. The problem of a payload delivery from an elliptical orbit taking into account atmospheric stage of reentry is considered. The tether deployment should provide the transition of the payload into a trajectory with the permissible loads. The conditions of the payload separation from the tether, which provide minimization of dynamic or thermal loads during the atmospheric motion or minimization of a Perigee Radius of the payload orbit, are studied.

Loucks Mike - One of the best experts on this subject based on the ideXlab platform.

  • Arcus Mission Design: Stable Lunar Resonant HEO for X-ray Astronomy
    2018
    Co-Authors: Carrico John, Dono Perez Andres, Policastri Lisa, Loucks Mike, Plice Laura
    Abstract:

    The Arcus mission, proposed for NASA's 2016 Astrophysics Medium Explorer (MIDEX) announcement of opportunity, will use X-ray spectroscopy to detect previously unaccounted quantities of normal matter in the Universe. The Arcus mission design uses 4:1 lunar resonance to provide a stable orbit for visibility of widely-dispersed targets, in a low background radiation environment, above the Van Allen belts for the minimum two-year science mission. Additional ad-vantages of 4:1 resonance are long term stability without maintenance maneuvers, eclipses under 4.5 hours, Perigee Radius approximately 12 Re for data download, and streamlined operational cadence with approximately 1 week or-bit period

  • Arcus Mission Design: Stable Lunar-Resonant High Earth Orbit for X-Ray Astronomy
    2018
    Co-Authors: Carrico John, Dono Perez Andres, Policastri Lisa, Plice Laura, Loucks Mike
    Abstract:

    The Arcus mission, proposed for NASA's 2016 Astrophysics Medium Explorer (MIDEX) announcement of opportunity, will use X-ray spectroscopy to detect previously unaccounted quantities of normal matter in the Universe. The Arcus mission design uses 4:1 lunar resonance to provide a stable orbit for visibility of widely-dispersed targets, in a low background radiation environment, above the Van Allen belts for the minimum two-year science mission. Additional ad-vantages of 4:1 resonance are long term stability without maintenance maneu-vers, eclipses under 4.5 hours, Perigee Radius approximately 12 Re for data download, and streamlined operational cadence with approximately 1 week orbit period

Alexander S. Ledkov - One of the best experts on this subject based on the ideXlab platform.

  • Tether-assisted re-entry capsule deorbiting from an elliptical orbit
    Acta Astronautica, 2017
    Co-Authors: Vladimir S. Aslanov, Alexander S. Ledkov
    Abstract:

    Abstract The problem of a payload delivery from an elliptical orbit to Earth surface by a space tethered system is considered. This operation includes three stages: tethered motion, orbital flight of the payload to the border of the atmosphere, and its descent into the atmosphere. The third stage is critical for the whole mission, because the payload can be damaged as a result of thermal or dynamic loads. Thus, the first stage should provide the transition of the payload into a trajectory with the permissible loads. The objective of the paper is to determinate the conditions of the payload separation from the tether, which provide minimization of dynamical or thermal loads during the atmospheric motion or minimization of a Perigee Radius of the payload orbit. The mathematical models for the each stage and equations for loads calculation are presented in the paper. The series of numerical calculation shows that eccentricity of the satellite orbit significantly affects the trajectory of the capsule with the payload in the atmosphere. The location of separation points on elliptical orbits and angles of the tether deflection that provide a transfer of the capsule to the trajectory with minimum Perigee Radius, dynamical pressure, heat flux, and total heat were found. In contrast to the case of a circular orbit, the separation of the capsule at the local vertical does not provide a transfer into an orbit with a minimum Radius of Perigee. It was observed that the increase in eccentricity can leads to the decrease of the Radius of Perigee and the total heat, but at the same time to growth of the dynamic and thermal loads. The Radius of Perigee of the re-entry capsule cannot be considered as a single indicator of efficiency of the tether deployment control law. The set of thermal and mechanical constraints must be taken into consideration at the stages of the tether control law and the re-entry capsule design developing.

  • Swing principle in tether-assisted return mission from an elliptical orbit
    Aerospace Science and Technology, 2017
    Co-Authors: Vladimir S. Aslanov, Alexander S. Ledkov
    Abstract:

    Abstract The problem of a tether-assisted payload return from an elliptical orbit is considered in this study. In contrast to the existing works devoted to this issue, the article deals with a tether length control that provides a transfer of the payload into a descent trajectory from the tether rotation mode. Application of the swing principle for the tether control is investigated. The simplified mathematical model of the space tethered system is developed. It is shown that the stable limit cycle could exist under the considered control. The approximate analytical solution for this cycle is obtained. The stability of this solution is studied by the Lyapunov's theorems. The optimal control, which provides transfer of the payload into the descent trajectory with minimum Perigee Radius, is found as a result of the simulation series. It is shown that the tether should occur several turns before the payload separation. For example, in the YES-2 experiment, it is demonstrated that proposed control makes it possible to perform a payload return using a tether of considerably shorter length. The main conclusion of the paper is that the proposed scheme of the payload deorbit is more effective than the classical static or dynamic tether deployment schemes.

Plice Laura - One of the best experts on this subject based on the ideXlab platform.

  • Arcus Mission Design: Stable Lunar Resonant HEO for X-ray Astronomy
    2018
    Co-Authors: Carrico John, Dono Perez Andres, Policastri Lisa, Loucks Mike, Plice Laura
    Abstract:

    The Arcus mission, proposed for NASA's 2016 Astrophysics Medium Explorer (MIDEX) announcement of opportunity, will use X-ray spectroscopy to detect previously unaccounted quantities of normal matter in the Universe. The Arcus mission design uses 4:1 lunar resonance to provide a stable orbit for visibility of widely-dispersed targets, in a low background radiation environment, above the Van Allen belts for the minimum two-year science mission. Additional ad-vantages of 4:1 resonance are long term stability without maintenance maneuvers, eclipses under 4.5 hours, Perigee Radius approximately 12 Re for data download, and streamlined operational cadence with approximately 1 week or-bit period

  • Arcus Mission Design: Stable Lunar-Resonant High Earth Orbit for X-Ray Astronomy
    2018
    Co-Authors: Carrico John, Dono Perez Andres, Policastri Lisa, Plice Laura, Loucks Mike
    Abstract:

    The Arcus mission, proposed for NASA's 2016 Astrophysics Medium Explorer (MIDEX) announcement of opportunity, will use X-ray spectroscopy to detect previously unaccounted quantities of normal matter in the Universe. The Arcus mission design uses 4:1 lunar resonance to provide a stable orbit for visibility of widely-dispersed targets, in a low background radiation environment, above the Van Allen belts for the minimum two-year science mission. Additional ad-vantages of 4:1 resonance are long term stability without maintenance maneu-vers, eclipses under 4.5 hours, Perigee Radius approximately 12 Re for data download, and streamlined operational cadence with approximately 1 week orbit period