The Orbit Parameter Message (OPM) is a standardized ASCII text format for exchanging spacecraft orbital state information at a single epoch. Unlike OEM which contains time-series ephemeris data, OPM represents a snapshot of spacecraft state and can include optional physical parameters and a single impulsive maneuver definition.
Key Components
An OPM file contains several distinct sections:
- Header: File-level metadata including format version, creation date, and originator
- Metadata: Orbital context such as object identification, reference frame, and time system
- State Vector: Spacecraft position and velocity at a specific epoch (Cartesian coordinates)
- Keplerian Elements (Optional): Alternative orbital element representation (semi-major axis, eccentricity, inclination, etc.)
- Spacecraft Parameters (Optional): Physical properties for orbit propagation (mass, drag coefficient, solar radiation pressure)
- Maneuver Parameters (Optional): Single impulsive maneuver definition with delta-V components
State Representation Options
OPM supports two ways to represent orbital state:
Cartesian State Vector
Position (X, Y, Z) and velocity (X_DOT, Y_DOT, Z_DOT) in a specified reference frame:
- X, Y, Z: Position components in kilometers
- X_DOT, Y_DOT, Z_DOT: Velocity components in km/s
Keplerian Elements (Optional Alternative)
Classical orbital elements describing the same state:
- SEMI_MAJOR_AXIS: Size of the orbit (km)
- ECCENTRICITY: Shape of the orbit (0 = circular, less than 1 = elliptical)
- INCLINATION: Angle between orbital plane and reference plane (degrees)
- RA_OF_ASC_NODE: Right ascension of ascending node (degrees)
- ARG_OF_PERICENTER: Argument of periapsis (degrees)
- TRUE_ANOMALY: Position along orbit (degrees)
Spacecraft Physical Parameters
OPM can include physical properties required for accurate orbit propagation:
- MASS: Spacecraft mass (kg)
- SOLAR_RAD_AREA: Effective area for solar radiation pressure (m²)
- SOLAR_RAD_COEFF: Solar radiation pressure coefficient
- DRAG_AREA: Effective area for atmospheric drag (m²)
- DRAG_COEFF: Atmospheric drag coefficient
Maneuver Definition
OPM supports a single impulsive maneuver with the following parameters:
- MAN_EPOCH_IGNITION: Maneuver execution time
- MAN_DURATION: Burn duration (seconds)
- MAN_DELTA_MASS: Propellant mass consumed (kg)
- MAN_REF_FRAME: Reference frame for delta-V components
- MAN_DV_1, MAN_DV_2, MAN_DV_3: Delta-V components (km/s)
OPM is designed for individual maneuvers only. For multiple maneuvers or complex burn sequences, use OCM (Orbit Comprehensive Message) instead.
Common Use Cases
- Orbit Determination Results: Sharing fitted state vectors after measurement processing
- State Vector Import: Loading initial conditions into mission planning systems
- Maneuver Planning: Defining planned burns with associated delta-V
- Data Exchange: Transferring orbital state between organizations
- Mission Analysis: Providing reference states for trajectory studies
Here is a sample OPM file in KVN format:
CCSDS_OPM_VERS = 3.0
CREATION_DATE = 2024-06-03T05:33:00.000
ORIGINATOR = VALAR
COMMENT Orbit determination solution from tracking pass
COMMENT State vector after maneuver execution
OBJECT_NAME = SPACECRAFT-ALPHA
OBJECT_ID = 2023-001A
CENTER_NAME = EARTH
REF_FRAME = EME2000
TIME_SYSTEM = UTC
EPOCH = 2024-06-03T00:00:00.000
X = 6655.9942 [km]
Y = -40218.5751 [km]
Z = -82.9177 [km]
X_DOT = 3.11548208 [km/s]
Y_DOT = 0.47042605 [km/s]
Z_DOT = -0.00101495 [km/s]
COMMENT Spacecraft physical parameters for propagation
MASS = 1913.000 [kg]
SOLAR_RAD_AREA = 10.000 [m**2]
SOLAR_RAD_COEFF = 1.300
DRAG_AREA = 10.000 [m**2]
DRAG_COEFF = 2.300
COMMENT Planned orbit-raising maneuver
MAN_EPOCH_IGNITION = 2024-06-03T09:00:34.1
MAN_DURATION = 132.60 [s]
MAN_DELTA_MASS = -18.418 [kg]
MAN_REF_FRAME = RTN
MAN_DV_1 = -0.02325700 [km/s]
MAN_DV_2 = 0.01683160 [km/s]
MAN_DV_3 = -0.00893444 [km/s]
OPM with Keplerian Elements Example
CCSDS_OPM_VERS = 3.0
CREATION_DATE = 2024-06-03T05:33:00.000
ORIGINATOR = VALAR
OBJECT_NAME = SPACECRAFT-BETA
OBJECT_ID = 2023-002B
CENTER_NAME = EARTH
REF_FRAME = EME2000
TIME_SYSTEM = UTC
EPOCH = 2024-06-03T12:00:00.000
COMMENT Keplerian elements representation
SEMI_MAJOR_AXIS = 42164.140 [km]
ECCENTRICITY = 0.0001234
INCLINATION = 0.05 [deg]
RA_OF_ASC_NODE = 75.123 [deg]
ARG_OF_PERICENTER = 180.456 [deg]
TRUE_ANOMALY = 90.789 [deg]
MASS = 2500.000 [kg]
SOLAR_RAD_AREA = 15.000 [m**2]
SOLAR_RAD_COEFF = 1.200
DRAG_AREA = 12.000 [m**2]
DRAG_COEFF = 2.200
| Feature | OPM | OEM | OCM |
|---|
| Temporal Scope | Single epoch | Time series | Single or multiple epochs |
| Maneuvers | One impulsive maneuver | Not supported | Multiple maneuvers |
| Keplerian Elements | Optional | Not supported | Optional |
| Covariances | Optional | Optional per epoch | Optional |
| Primary Use Case | State snapshots, OD output | Ephemeris sharing | Comprehensive mission plans |
The SpaceX OPM represents the state of the second stage, not the deployed spacecraft. Any position, velocity, attitude, or attitude-rate differences between the second stage and your spacecraft at separation must be accounted for by the recipient.
SpaceX OPM Fields
| Field | Description | Units |
|---|
| UTC time at liftoff | Launch time reference | DOY:HH:MM:SS.SS |
| UTC time of current state | Epoch of the state vector | DOY:HH:MM:SS.SS |
| Mission elapsed time | Time since liftoff | seconds |
| ECEF Position (X,Y,Z) | Position in WGS84 ECEF frame | meters |
| ECEF Velocity (X,Y,Z) | Earth-relative velocity in ECEF | m/s |
| LVLH to BODY quaternion | Attitude quaternion (scalar-first: S,X,Y,Z) | dimensionless |
| Inertial body rates (X,Y,Z) | Angular velocity | deg/s |
| Apogee Altitude | Maximum altitude (spherical Earth) | km |
| Perigee Altitude | Minimum altitude (spherical Earth) | km |
| Inclination | Orbital inclination | degrees |
| Argument of Perigee | Orientation of orbit ellipse | degrees |
| Longitude of Asc. Node | Ascending node referenced to Greenwich | degrees |
| True Anomaly | Position along orbit | degrees |
Key Differences from CCSDS OPM
| Aspect | CCSDS OPM | SpaceX OPM |
|---|
| Reference Frame | Inertial (EME2000, GCRF) | WGS84 ECEF, inertially frozen at state epoch |
| Position Units | kilometers | meters |
| Velocity Units | km/s | m/s |
| Velocity Reference | Inertial | Earth-relative |
| Time Format | ISO 8601 | Day-of-Year (DOY:HH:MM:SS.SS) |
| Ascending Node | Right Ascension (vernal equinox) | Longitude (Greenwich Meridian) |
| Altitude Reference | Typically WGS84 ellipsoid | Spherical Earth (6378.137 km radius) |
| Attitude Data | Not included | LVLH-to-body quaternion + body rates |
Sample SpaceX OPM
SpaceX OPM output (generated 2024-06-15-Sat-14-30-00):
All orbital elements are defined as osculating at the instant of the printed state.
Orbital elements are computed in an inertial frame realized by inertially freezing
the WGS84 ECEF frame at time of current state.
UTC time at liftoff: 166:14:00:00.00
UTC time of current state: 166:14:45:30.25
Mission elapsed time (s): +2730.25
ECEF (X,Y,Z) Position (m): +4523156.789, -3298765.432, +4012345.678
ECEF (X,Y,Z) Velocity* (m/s): +5234.567, +4123.456, -3456.789
LVLH to BODY quaternion (S,X,Y,Z): +0.7071068, +0.0000000, +0.7071068, +0.0000000
Inertial body rates (X,Y,Z) (deg/s): +0.0012345, -0.0023456, +0.0034567
Apogee Altitude** (km): +00850.123
Perigee Altitude** (km): +00320.456
Inclination (deg): +53.215
Argument of Perigee (deg): +090.123
Longitude of the Asc. Node*** (deg): +125.678
True Anomaly (deg): +45.890
Notes:
* ECEF velocity is Earth-relative
** Apogee/Perigee altitude assumes a spherical Earth, 6378.137 km radius
*** LAN is defined as the angle between Greenwich Meridian and the ascending node
Importing SpaceX OPM in VALAR
When importing SpaceX OPM files:
- Select SpaceX OPM as the format in the import dialog
- VALAR automatically converts the ECEF state to an inertial frame
- The Longitude of Ascending Node is converted to Right Ascension
- Position and velocity units are converted from meters to kilometers
SpaceX OPM format is documented in the Falcon User’s Guide. VALAR’s SpaceX OPM parser handles all necessary coordinate transformations automatically. 