parseTLElines: Parse the lines of a TLE

In Rafael-Ayala/asteRisk: Computation of Satellite Position

Parse the lines of a TLE

Description

TLE (Two-/Three- Line Element) is the standard format for representing orbital state vectors. This function parses a character vector where each element represents a line of the TLE. The supplied character vector can have either 2 (for Two Line Elements) or 3 (for Three Line Elements) elements. The two lines of a Two Line Element contain all the information. The additional line in a Three Line Element is optional, and contains just the satellite name. For a detailed description of the TLE format, see https://celestrak.org/columns/v04n03/#FAQ01.

Usage

parseTLElines(lines)

Arguments

lines

Character vector where each element is a string corresponding to a line of the TLE. The character vector must have either 2 or 3 elements.

Value

A list with the following elements that define the orbital state vector of the satellite:

NORADcatalogNumber	NORAD Catalog Number, also known as Satellite Catalog Number, assigned by United States Space Command to each artificial object orbiting Earth
classificationLevel	Classification level of the information for the orbiting object. Can be unclassified, classified, secret or unknown
internationalDesignator	International Designator, also known as COSPAR ID, of the object. It consists of the launch year, separated by a hyphen from a three-digit number indicating the launch number for that year and a set of one to three letters indicating the piece for a launch with multiple pieces.
launchYear	The launch year of the object
launchNumber	The launch number of the object during its launch year
launchPiece	The piece for the launch of the object, if it was a launch with multiple pieces
dateTime	Date time string to which the orbital state vector corresponds
elementNumber	Element number for the object. In principle, every time a new TLE is generated for an object, the element number is incremented, and therefore element numbers could be used to assess if all the TLEs for a certain object are available. However, in practice it is observed that this is not always the case, with some numbers skipped and some numbers repeated.
inclination	Mean orbital inclination of the satellite in degrees. This is the angle between the orbital plane of the satellite and the equatorial plane
ascension	Mean longitude of the ascending node of the satellite at epoch, also known as right ascension of the ascending node, in degrees. This is the angle between the direction of the ascending node (the point where the satellite crosses the equatorial plane moving north) and the direction of the First Point of Aries (which indicates the location of the vernal equinox)
eccentricity	Mean eccentricity of the orbit of the object. Eccentricity is a measurement of how much the orbit deviates from a circular shape, with 0 indicating a perfectly circular orbit and 1 indicating an extreme case of parabolic trajectory
perigeeArgument	Mean argument of the perigee of the object in degrees. This is the angle between the direction of the ascending node and the direction of the perigee (the point of the orbit at which the object is closest to the Earth)
meanAnomaly	Mean anomaly of the orbit of the object in degrees. This indicates where the satellite is along its orbital path. It is provided as the angle between the direction of the perigee and the hypothetical point where the object would be if it was moving in a circular orbit with the same period as its true orbit after the same amount of time since it last crossed the perigee had ellapsed. Therefore, 0 denotes that the object is at the perigee
meanMotion	Mean motion of the satellite at epoch in revolutions/day
meanMotionDerivative	First time derivative of the mean motion of the satellite in revolutions/day^2^
meanMotionSecondDerivative	Second time derivative of the mean motion of the satellite in revolutions/day^3^.
Bstar	Drag coefficient of the satellite in units of (earth radii)^-1^. Bstar is an adjusted value of the ballistic coefficient of the satellite, and it indicates how susceptible it is to atmospheric drag.
ephemerisType	Source for the ephemeris (orbital state vector). Most commonly, it is distributed data obtained by combaining multiple observations with the SGP4/SDP4 models
epochRevolutionNumber	Number of full orbital revolutions completed by the object
objectName	Name of the object, retrieved from the first line of the TLE if a Three Line Element was provided

References

https://celestrak.org/columns/v04n03/#FAQ01

Examples

# The following lines correspond to a TLE for Italsat 2 the 26th of June, 2006
# at 00:58:29.34 UTC.

italsat2_lines <- c("ITALSAT 2",
"1 24208U 96044A   06177.04061740 -.00000094  00000-0  10000-3 0  1600",
"2 24208   3.8536  80.0121 0026640 311.0977  48.3000  1.00778054 36119")

italsat2_TLE <- parseTLElines(italsat2_lines)
italsat2_TLE

Rafael-Ayala/asteRisk documentation built on May 16, 2024, 5:24 p.m.