The Two-Line Element Set (TLE) is a standardized data format used to describe the orbital elements of an Earth-orbiting object. It provides the necessary parameters to compute the satellite’s position and velocity at a given time using specialized orbital propagation models such as SGP4 (Simplified General Perturbations) or SDP4 (Simplified Deep Space Perturbations).
TLEs are distributed by organizations such as CelesTrak and Space-Track, and are widely used in both professional and amateur orbital analysis systems for tracking satellites, debris, and spacecraft.
Structure Overview
A TLE consists of two lines of 69 characters each, sometimes preceded by an optional title line containing the satellite’s name. Each field in the lines has a fixed position and meaning, making it a rigid but efficient format for data exchange.
Key Components
Optional Title Line (Line 0)
- Satellite common name (not part of the formal TLE specification)
- Catalog identification and classification
- Launch designator
- Epoch (date and time of element set)
- Mean motion derivatives (orbital decay parameters)
- BSTAR drag term (atmospheric drag coefficient)
Line 2: Keplerian Orbital Elements
- Classical orbital elements (inclination, RAAN, eccentricity, argument of perigee, mean anomaly)
- Mean motion (orbital period)
- Revolution number at epoch
Line 1 Field Breakdown
| Columns | Field | Description |
|---|
| 1 | Line number | Always 1 |
| 3–7 | Satellite number | Unique 5-digit NORAD catalog number |
| 8 | Classification | U = Unclassified, C = Classified, S = Secret |
| 10–11 | Launch year | Last two digits of the launch year |
| 12–14 | Launch number | Launch count of that year |
| 15–17 | Launch piece | Piece of the launch (e.g., A, B, C) |
| 19–32 | Epoch | Format YYDDD.DDDDDDDD, where DDD is day of year and decimal is fractional day |
| 34–43 | First derivative of mean motion | Mean motion rate (rev/day²), orbital decay rate |
| 45–52 | Second derivative of mean motion | Usually 0 for SGP4 model |
| 54–61 | BSTAR drag term | Atmospheric drag coefficient (Earth radii⁻¹) |
| 63 | Ephemeris type | Usually 0 |
| 65–68 | Element set number | Version number of the element set |
| 69 | Checksum | Sum of all digits in the line modulo 10 |
Line 2 Field Breakdown
| Columns | Field | Description |
|---|
| 1 | Line number | Always 2 |
| 3–7 | Satellite number | Must match Line 1 |
| 9–16 | Inclination (deg) | Angle between orbital plane and Earth’s equator |
| 18–25 | Right Ascension of Ascending Node (deg) | Longitude of the ascending node |
| 27–33 | Eccentricity | Decimal point assumed (e.g., 0005742 = 0.0005742) |
| 35–42 | Argument of Perigee (deg) | Angle from ascending node to perigee |
| 44–51 | Mean Anomaly (deg) | Satellite’s position within its orbit at epoch |
| 53–63 | Mean Motion (rev/day) | Number of revolutions per day |
| 64–68 | Revolution number at epoch | Total number of orbits completed since launch |
| 69 | Checksum | Sum of all digits in the line modulo 10 |
Critical Fields
Epoch (Line 1, Columns 19-32)
Defines the precise date and time when the orbital elements are valid. Propagators use this reference time to compute position and velocity at other times. Accuracy degrades as you propagate further from the epoch.
BSTAR Drag Term (Line 1, Columns 54-61)
Critical for low-Earth orbit propagation. Models atmospheric decay effects and is expressed in units of Earth radii⁻¹. Larger values indicate stronger atmospheric drag.
Mean Motion (Line 2, Columns 53-63)
Determines the orbital period using the formula:
Orbital Period (minutes) = 1440 / Mean Motion (rev/day)
For example, a mean motion of 15.5 rev/day corresponds to an orbital period of approximately 93 minutes.
Checksums (Line 1 & 2, Column 69)
Each line ends with a checksum digit calculated as the sum of all numeric digits modulo 10. The minus sign (-) counts as 1. This ensures data integrity during transmission.
Common Use Cases
- Satellite Tracking: Computing real-time positions for orbit visualization
- Collision Avoidance: Propagating orbits for conjunction assessment
- Space Situational Awareness: Monitoring debris and active satellites
- Mission Planning: Quick orbital analysis and visualization
- Catalog Correlation: Matching observations to cataloged objects
TLE format documentation available at CelesTrak. For orbit propagation theory, see Vallado’s Fundamentals of Astrodynamics and Applications. ISS (ZARYA)
1 25544U 98067A 24315.54791547 .00003100 00000-0 61805-4 0 9991
2 25544 51.6414 66.0545 0005742 30.4418 95.9135 15.50049823389163
Parsing the Example
Line 0 (Title): ISS (ZARYA), common name of the spacecraft
Line 1 Breakdown:
- Satellite Number:
25544 (ISS NORAD catalog ID)
- Classification:
U (Unclassified)
- Launch Designator:
98067A (1998, launch 67, piece A)
- Epoch:
24315.54791547 (November 10, 2024 at 13:09:00 UTC)
- Mean Motion Derivative:
0.00003100 rev/day² (slight orbital decay)
- BSTAR:
0.000061805 (atmospheric drag coefficient)
- Element Set Number:
999 (version)
Line 2 Breakdown:
- Satellite Number:
25544 (matches Line 1)
- Inclination:
51.6414° (typical ISS orbit)
- RAAN:
66.0545°
- Eccentricity:
0.0005742 (nearly circular)
- Argument of Perigee:
30.4418°
- Mean Anomaly:
95.9135°
- Mean Motion:
15.50049823 rev/day (≈93 min orbital period)
- Revolution Number:
38916 (orbits since launch)
Working with TLE Data
TLEs are a widely-used format for sharing orbital information, particularly from public catalogs like Space-Track and CelesTrak. While TLEs provide a convenient orbital snapshot, they have specific characteristics to understand:
SGP4/SDP4 Propagation Models
TLEs are designed to work with specialized propagation models:
- SGP4: Simplified General Perturbations for near-Earth orbits (period less than 225 minutes)
- SDP4: Simplified Deep Space Perturbations for higher orbits (period greater than 225 minutes)
These models use mean orbital elements and simplified perturbation theory, making them computationally efficient but less accurate than high-fidelity numerical propagators.
Converting TLE to State Vectors
To use TLE data with high-precision orbit propagation tools, you’ll need to:
- Parse the TLE using an SGP4/SDP4 library
- Propagate to the desired epoch using the SGP4/SDP4 model
- Convert the resulting position and velocity from TEME frame to your working frame (typically GCRF or ITRF)
- Import the state vector in your preferred format (e.g., OPM)
TLE accuracy degrades over time due to atmospheric drag, solar pressure, and gravitational perturbations. For high-precision applications, use TLEs with recent epochs or switch to higher-fidelity state vectors (OPM/OEM).
Limitations and Best Practices
Accuracy Degradation
- Low-Earth Orbit: TLEs remain accurate for 1-7 days
- Medium-Earth Orbit: Accuracy extends to several weeks
- Geostationary Orbit: Can remain accurate for months
Best Practices
- Use Recent TLEs: Always use the most recent element set available
- Check Epoch Date: Verify TLE epoch is close to your analysis period
- Update Regularly: Refresh TLEs frequently for operational missions
- Validate Checksums: Ensure data integrity by verifying checksums
- Understand Limitations: TLEs are mean elements, not osculating (instantaneous) elements
| Feature | TLE | OPM | OEM |
|---|
| Format Type | Fixed-width ASCII | Keyword-value (KVN) | Keyword-value (KVN) |
| Temporal Scope | Single epoch | Single epoch | Time series |
| Propagation Model | SGP4/SDP4 required | High-fidelity models | Interpolation |
| Element Type | Mean elements | Osculating elements | Osculating state |
| Accuracy | Moderate (days) | High (hours-days) | Very high (minutes) |
| Use Case | Tracking, awareness | Mission operations | High-precision ops |
