Mission Planning for Optical Satellite’s Constant Surveillance to Geostationary Spacecraft

2021 ◽  
Author(s):  
Haitao Zhang ◽  
Zhi Li ◽  
Weilin Wang ◽  
Hao Wang ◽  
Yasheng Zhang
Keyword(s):  
Trajectory Optimization ◽  
Initial Phase ◽  
Situational Awareness ◽  
Initial Position ◽  
Mission Design ◽  
The Sun ◽  
Orbital Inclination ◽  
Orbital Maneuvers ◽  
Phase Interval ◽  
Geometric Relationship

Abstract For a mission to constantly watch geostationary (Orbital inclination isn’t 0, GEO) spacecraft by an optical satellite during a whole fly-around cycle, study the relative position relationship between the two and sun during fly-around mission; design the trajectory of the optical satellite, on which, the optical satellite keeps facing to the spacecraft in the direction opposite the Sun. Firstly, for constant surveillance to geosynchronous (Orbital inclination is 0) spacecraft, study from the Keplerian orbit elements, analyze its geometric relationship with the sun and the optical satellite. Then calculate the initial phase interval that meets the requirements of the mission. Compared with Clohessy-Wiltshire equation (CW equation), this method is more concise and the spatial physical meaning is clearer. However, the orbital inclination of GEO spacecraft is usually not 0. Secondly, taking GEO spacecraft with 1° inclination as an example, calculate the initial phase interval of the mission. Thirdly, select an initial phase in the initial phase interval, and design the fly-around trajectory based on CW equation. Lastly, the optical satellite’s position when it receives the mission is initial position, and the position when the fly-around mission starts is final position. The optical satellite’s approach trajectory is summarized as spacecraft's Lambert trajectory optimization. Take the time of two orbital maneuvers as optimization variables, and the fuel consumption as optimization objective. Optimize the plan of orbital maneuvering. The total pulse thrust velocity required for orbital maneuver after optimization in the example is 18.2514m/s, which is highly feasible in engineering. This method can be used for space situational awareness and in-orbit services of GEO spacecraft.

Simulation of 5DOF Model for Arial Grasping

2015 ◽  
Vol 772 ◽  
pp. 381-387
Author(s):  
Haniyeh Rashidi Fathabadi ◽  
Afshin Banazadeh ◽  
Fariborz Saghafi
Keyword(s):  
Trajectory Optimization ◽  
Vertical Plane ◽  
Initial Position ◽  
Degree Of Freedom ◽  
Mathematical Representation ◽  
Additional Mass ◽  
Additional Degree ◽  
Air Vehicle ◽  
Euler Approach ◽  
Aerial Hunting

This study presents dynamic modeling and simulation of an air vehicle consisting of a body, gripper and a claw. This model is inspired from birds’ aerial hunting, while considering the extra degree of freedom associated with the claw. For a manipulator like a gripper, additional degree of freedom creates more flexibility for grasping. The main contribution of this paper focuses on the development of a model that is suitable for trajectory optimization in grasping phase. Mathematical representation of the system is developed based on the Newton-Euler approach in MATLAB-Simulink environment, considering the motion in vertical plane. The dynamic behavior of the system is evaluated by simulation in variety situations and sensitivity analysis is carried out to determine and characterize the parameters having the most and least effects on grasping. It is shown that the initial position of the gripper and the claw as well as the additional mass that is added to the system in grasping phase make considerable changes in the dynamics that necessitates the use of the control system. In addition, smooth trajectories and controls are obtained by adding friction to the system in order to avoid dynamic divergence.

scholarly journals Simultaneous Five Color (UBVRI) Polarimetry of VV Puppis

1987 ◽  
Vol 93 ◽  
pp. 203-203
Author(s):  
V. Piirola ◽  
A. Reiz ◽  
G.V. Coyne
Keyword(s):  
Circular Polarization ◽  
Linear Polarization ◽  
Wavelength Dependence ◽  
Position Angle ◽  
Orbital Inclination ◽  
V Band ◽  
Highly Active ◽  
Phase Interval ◽  
In The Beginning ◽  
Bright Phase

AbstractObservations of linear and circular polarization in five colour bands during a highly active state of VV Puppis in January 86 are reported. A strong linear polarization pulse with the maximum in the blue, PB ≈ 22%, is observed at the end of the bright phase when the active pole is at the limb and a weaker secondary pulse, PB ≈ 7%, is seen in the beginning of the bright phase, when the active pole reappears. Strong positive circular polarization is also observed in the blue and the ultraviolet, РU ≈ PB ≈ 18%, PV ≈ 10% during the bright phase. The circular polarization reverses the sign in the B and V bands during the faint phase and a negative polarization hump is seen when the active pole crosses the limb. The circular polarization in the V band reaches the value PV ≈ −10% at the hump, after which it remains near PV ≈ −5% during the faint phase. This is probably due to radiation coming from the second, less active pole and accretion thus takes place onto both poles. The wavelength dependences of the positive and negative parts of the circular polarization curve are different and no polarization reversal is seen in the U band. The position angle of the linear polarization is well determined during a large portion of the cycle, especially in the V band, thanks to the activity from both poles. A best fit to the position angle curve, taking into account also the duration of the positive circular polarization phase interval ΔΦ = 0.40 (in the V band), yields the values of orbital inclination i = 78° ± 2° and the colatitude of the active magnetic pole ß = 146° ± 2°. The relatively good fit to the position angle data indicates that the simple dipole model is nearly correct in the case of VV Puppis. Some wavelength dependence is, however, seen in the position angle curves, especially in the I band where the slope Δθ/ΔΦ at the main pulse is considerably smaller than in the other bands. The shape of the position angle curves changes also in the blue and the ultraviolet around the middle of the bright phase. This is probably due to optical thickness effects as the side of the accretion column which is toward the observer changes near this phase.

scholarly journals Solar System limits on gravitational dipoles

2020 ◽  
Vol 495 (4) ◽  
pp. 3974-3980
Author(s):  
Indranil Banik ◽  
Pavel Kroupa
Keyword(s):  
Solar System ◽  
No Value ◽  
Small Region ◽  
Initial Position ◽  
The Sun ◽  
Current Form ◽  
Modified Newtonian Dynamics ◽  
Upper Limit ◽  
Gravitational Dipoles ◽  
Isolated Point

ABSTRACT The gravitational dipole theory of Hadjukovic (2010) is based on the hypothesis that antimatter has a negative gravitational mass and thus falls upwards on the Earth. Astrophysically, the model is similar to but more fundamental than Modified Newtonian Dynamics (MOND), with the Newtonian gravity $g_{_\mathrm{ N}}$ towards an isolated point mass boosted by the factor $\nu = 1 + \left(\alpha /x \right) \tanh \left(\sqrt{x}/\alpha \right)$, where $x \equiv g_{_\mathrm{ N}}/a_{_0}$ and $a_{_0} = 1.2 \times 10^{-10}$ m s−2 is the MOND acceleration constant. We show that α must lie in the range 0.4–1 to acceptably fit galaxy rotation curves. In the Solar System, this interpolating function implies an extra Sunwards acceleration of ${\alpha a_{_0}}$. This would cause Saturn to deviate from Newtonian expectations by 7000(α/0.4) km over 15 yr, starting from known initial position and velocity on a near-circular orbit. We demonstrate that this prediction should not be significantly altered by the postulated dipole haloes of other planets due to the rather small region in which each planet’s gravity dominates over that of the Sun. The orbit of Saturn should similarly be little affected by a possible ninth planet in the outer Solar System and by the Galactic gravity causing a non-spherical distribution of gravitational dipoles several kAU from the Sun. Radio tracking of the Cassini spacecraft orbiting Saturn yields a 5σ upper limit of 160 m on deviations from its conventionally calculated trajectory. These measurements imply a much more stringent upper limit on α than the minimum required for consistency with rotation curve data. Therefore, no value of α can simultaneously match all available constraints, falsifying the gravitational dipole theory in its current form at extremely high significance.

The Solaris Solar Polar Mission

2020 ◽  
Author(s):  
Donald M. Hassler ◽  
Jeff Newmark ◽  
Sarah Gibson ◽  
Louise Harra ◽  
Thierry Appourchaux ◽  
...  
Keyword(s):  
Remote Sensing ◽  
Solar Physics ◽  
Polar View ◽  
Ecliptic Plane ◽  
Vantage Point ◽  
Mission Design ◽  
Future Research ◽  
The Sun ◽  
Gravity Assist ◽  
Solar Observations

<p>The solar poles are one of the last unexplored regions of the solar system. Although Ulysses flew over the poles in the 1990s, it did not have remote sensing instruments onboard to probe the Sun’s polar magnetic field or surface/sub-surface flows.</p><p>We will discuss Solaris, a proposed Solar Polar MIDEX mission to revolutionize our understanding of the Sun by addressing fundamental questions that can only be answered from a polar vantage point. Solaris uses a Jupiter gravity assist to escape the ecliptic plane and fly over both poles of the Sun to >75 deg. inclination, obtaining the first high-latitude, multi-month-long, continuous remote-sensing solar observations. Solaris will address key outstanding, breakthrough problems in solar physics and fill holes in our scientific understanding that will not be addressed by current missions.</p><p>With focused science and a simple, elegant mission design, Solaris will also provide enabling observations for space weather research (e.g. polar view of CMEs), and stimulate future research through new unanticipated discoveries.</p>

scholarly journals Multiphase Trajectory Optimization of a Lunar Return Mission to an LEO Space Station

2021 ◽  
Vol 2021 ◽  
pp. 1-15
Author(s):  
Luyi Yang ◽  
Haiyang Li ◽  
Jin Zhang ◽  
Yazhong Luo
Keyword(s):  
Trajectory Optimization ◽  
Space Station ◽  
Low Earth Orbit ◽  
Hybrid Optimization ◽  
Target Space ◽  
Mission Design ◽  
Optimization Scheme ◽  
Initial State ◽  
Bank Angle ◽  
Entire Trajectory

Lunar exploration architecture can be made more flexible and reliable with the support of a low-Earth orbit (LEO) space station. This study therefore evaluated a proposed hybrid optimization scheme to design the entire trajectory of a reusable spacecraft starting from trans-Earth injection (EI) at the perilune and ending at an LEO space station. As such a trajectory has multiple constraints and multiple dynamical models, it is divided into the trans-Earth phase, aerocapture phase, and postatmospheric phase. The optimization scheme is performed at two levels: sublevel and top level. At the sublevel, two novel pseudo rules are proposed to optimize the trans-Earth trajectory so that it satisfies the coplanar constraints of the space station. Then, in the aerocapture phase, the bank angle is optimized to satisfy the mission constraints, and in the atmospheric phase, the one-impulsive maneuver is performed and optimized to insert the spacecraft into the target space station orbit. The multiple phases are connected to each other by boundary conditions where the terminal state of the previous phase is transformed into the initial state of the following phase. At the top level, the vacuum perigee height is selected as the mission design variable based on problem characteristics analysis and a hybrid optimization scheme is conducted to minimize the total velocity increment. The simulation results demonstrate that the proposed hybrid optimization method is effective for the design of an entire trajectory with acceptable velocity cost which is less than that in the previous study. The coplanar constraints of the space station and other mission constraints in each phase are also satisfied. Furthermore, the proposed trajectory design method is shown to be applicable to a reusable spacecraft returning to an LEO space station parked in any arbitrary orbital plane.

scholarly journals Three-Apogee 16-h Highly Elliptical Orbit as Optimal Choice for Continuous Meteorological Imaging of Polar Regions

2011 ◽  
Vol 28 (11) ◽  
pp. 1407-1422 ◽  
Author(s):  
Alexander P. Trishchenko ◽  
Louis Garand ◽  
Larisa D. Trichtchenko
Keyword(s):  
Spatial Resolution ◽  
High Energy ◽  
Optimal Choice ◽  
Elliptical Orbit ◽  
Polar Regions ◽  
Orbital Inclination ◽  
Orbital Parameters ◽  
Orbital Maneuvers ◽  
Orbital Periods ◽  
Optimum Solution

Abstract A highly elliptical orbit (HEO) with a 16-h period is proposed for continuous meteorological imaging of polar regions from a two-satellite constellation. This orbit is characterized by three apogees (TAP) separated by 120°. The two satellites are 8 h apart, with repeatable ground track in the course of 2 days. Advantages are highlighted in comparison to the Molniya 12-h orbit described in detail in a previous study (Trishchenko and Garand). Orbital parameters (period, eccentricity, and inclination) are obtained as a result of an optimization process. The principles of orbit optimization are based on the following four key requirements: spatial resolution (apogee height), the altitude of crossing the trapped proton region at the equator (minimization of radiation doze caused by trapped protons), imaging time over the polar regions, and the stability of the orbit, which is mostly defined by the rotation of perigee. The interplay between these requirements points to a 16-h period with an eccentricity of 0.55 as the optimum solution. The practical range of orbit inclinations that could be maintained during the spacecraft lifetime can vary from a critical value of 63.435° to 70° (subject to the amount of propellant available for orbital maneuvers). In comparison to Molniya, this type of orbit reduces the radiation exposure to high-energy protons by factor of 103–104. On the other hand, the main advantage of 16 h versus longer orbital periods up to 24 h is better spatial resolution as a result of a lower apogee height. A two-satellite TAP constellation with an orbital inclination of 66° provides 100% temporal coverage above 60°N, >95% above 55°N, >85% above 50°N, and >75% above 45°N.

scholarly journals Orbital Perigee Deviation under Inclination Window for Sun Synchronized Low Earth Orbits

2019 ◽  
Vol 4 (10) ◽  
pp. 127-130
Author(s):  
Shkelzen Cakaj ◽  
Bexhet Kamo
Keyword(s):  
Orbital Plane ◽  
Earth Observation ◽  
Gas Content ◽  
The Sun ◽  
Orbital Inclination ◽  
The Earth ◽  
Low Earth Orbits ◽  
Earth Observation Satellites ◽  
Earth Orbits ◽  
Earth’S Oblateness

Data processing related to the Earth’s changes, gathered from different platforms and sensors implemented worldwide and monitoring the environment and structure represents Earth observation (EO). Environmental monitoring includes changes in Earth’s vegetation, atmospheric gas content, ocean state, melting level in the ice fields, etc. This process is mainly performed by satellites. The Earth observation satellites use Low Earth Orbits (LEO) for their missions. These missions are accomplished mainly based on photo imagery. Thus, the relative Sun’s position related to the observed area, it is very important for the photo imagery, in order the observed area from the satellite to be treated under the same lighting (illumination) conditions. This could be achieved by keeping a constant Sun position related to the orbital plane due to the Earth’s motion around the Sun. This is called Sun synchronization for low Earth orbits, the feature which is applied for satellites dedicated for the Earth observation. Nodal regression is the phenomenon which is utilized for low circular orbits providing to them the Sun synchronization. Nodal regression refers to the shift of the orbit’s line of nodes over time as Earth revolves around the Sun,  caused due to the Earth’s oblateness. Nodal regression depends on orbital altitude and orbital inclination angle. For the in advance defined range of altitudes stems the inclination window for the satellite low Earth orbits to be Sun synchronized. For analytical and simulation purposes, the altitudes from 600km to 1200km are considered. Further for the determined inclination window of the Sun synchronization it is simulated the orbital perigee deviation for the above considered altitudes and the eventual impact on the satellite’s mission.

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