China
Africa
Explore chapters and articles related to this topic
Spacecraft and Missions
Published in Julio Sanchez, Maria P. Canton, William Perrizo, Space Image Processing, 2018
Julio Sanchez, Maria P. Canton
The apogee is the point in the orbit in which the satellite is farthest away from the earth and the perigee the one in which the satellite is closest to the earth. The ascending node is the point where the satellite orbit crosses the equator when moving south to north, that is, on the ascending pass. The descending node is the point where the satellite orbit crosses the equator when moving north to south, on the descending pass. The points where the path crosses the equator are also called the equatorial crossing. The orbital inclination is the angle between the satellite track and the equator at the ascending node.
Defense Information, Communication, and Space Technology
Published in Anna M. Doro-on, Handbook of Systems Engineering and Risk Management in Control Systems, Communication, Space Technology, Missile, Security and Defense Operations, 2023
The ascending node is the point at which the northbound (ascending) satellite crosses the equator (Olsen 2007) as shown in Figure 5.4a. The right ascension of the ascending is the celestial longitude of this point, as another definition, it is the angle between the plane of the satellite orbit and the line connecting the earth and sun on the first day of spring, or vernal equinox as depicted in many illustrations of Figure 5.3 (Olsen 2007). It is characterized as an angle Ω computed from the vernal equinox towards the line of nodes in the direction of the earth’s rotation (Figure 5.2-C). The right ascension can also be described as being measured from the point of Aries (Olsen 2016). The inertial coordinate representation (of right ascension and declination) provides simple one angle (right ascension) for the orbital plane of the spacecraft which is inertial (Kramer 2002). Determining and utilizing of the appropriate angle of right ascension of the ascending node, Ω, makes sure that the satellite orbits in the designated plane. This can be done by selecting suitable injection time relative to the longitude. Angle Ω can be determined by measuring the difference between two angles: (1) the angle α between the direction of the vernal equinox and the longitude of the injection point and (2) the angle β between the line of nodes and the longitude of the injection point, as shown in Figure 5.4b. Angle β can be computed from: sinβ=cosisinlcoslsini
Force-System Resultants and Equilibrium
Published in Richard C. Dorf, The Engineering Handbook, 2018
Ω represents the right ascension (α) of the ascending node. The ascending node is the (projected) point where the satellite rises above the equator plane.
Direct trajectory optimization framework for vertical takeoff and vertical landing reusable rockets: case study of two-stage rockets
Published in Engineering Optimization, 2019
Lin Ma, Kexin Wang, Zhijiang Shao, Zhengyu Song, Lorenz T. Biegler
The first stage of the rocket returns to the Earth and finally lands at a specified site at time after stage separation; thus, the terminal constraint of the first stage is given as follows: The second stage of the rocket finally delivers the payload to the desired orbit at time , and the terminal position and velocity of the second stage can be transformed into orbital elements. The terminal constraint of the second stage is hence defined as follows: where the orbital elements represent the semi-major axis, eccentricity, inclination, right ascension of the ascending node and argument of perigee, respectively. The true anomaly is left undefined because the exact location within the orbit is not constrained.
Fully distributed finite-time adaptive robust time-varying formation-containment control for satellite formation
Published in International Journal of Control, 2023
Pingli Lu, Qing Jiang, Ye Tian, Haikuo Liu, Changkun Du
Consider a satellite formation system consisting of four leader satellites labelled as 1, 2, 3, 4, five follower satellites labelled as , respectively. The reference signal is labelled as 0 and only available for the 1th leader satellite. The relative motion dynamics of the ith satellite can be described by (2). Assume that the reference satellite moves along a near-circular orbit around the Earth with the following initial orbit elements, where a is the semi-major axis (in meters) of the orbit, e is the eccentricity, i is the inclination, Ω is the longitude of the ascending node, ϖ is the argument of perigee, and θ is the true anomaly.
Numerical Solving of Radiation Geometrical Inverse Problem
Published in Heat Transfer Engineering, 2023
Aleksey V. Nenarokomov, Evgeniy V. Chebakov, Dmitry L. Reviznikov, Irina V. Krainova
In the general case the orientation of an arbitrary surface element of a spacecraft can be determined by the following nine angles: Three angles determine the relative position of the equatorial XYZ and orbital nrb coordinate systems (Figure 1a): where Ω is the longitude of the ascending node, i is the inclination of orbit, u is the argument of latitude. The planetocentric equatorial coordinate system can be considered as an inertial coordinate system for most engineering problems. We assume that the corresponding angles are always known from the predetermined spaceflight control program.