parseTLElines: Parse the lines of a TLE

View source: R/fileParsers.R

parseTLElinesR Documentation

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.