Orbit Parameter Message (OPM)

For the complete documentation index, see llms.txt.

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 one or more maneuver definitions.

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): Supplementary orbital element representation (semi-major axis, eccentricity, inclination, etc.) that accompanies the state vector
Spacecraft Parameters (Optional): Physical properties for orbit propagation (mass, drag coefficient, solar radiation pressure)
Maneuver Parameters (Optional): One or more maneuvers (impulsive or finite), each with delta-V components

State Representation

Every OPM carries a mandatory Cartesian state vector. Osculating Keplerian elements may be added as an optional supplementary block describing the same state — they do not replace the state vector.

Cartesian State Vector (Mandatory)

Position and velocity at the epoch, in the reference frame given by REF_FRAME:

Element	Description	Unit
X, Y, Z	Position components	km
X_DOT, Y_DOT, Z_DOT	Velocity components	km/s

Osculating Keplerian Elements (Optional)

A supplementary block. If included, all of its elements must be present — including GM:

Element	Description	Unit
SEMI_MAJOR_AXIS	Size of the orbit	km
ECCENTRICITY	Shape of the orbit (0 = circular, < 1 = elliptical)	—
INCLINATION	Angle between the orbital plane and the reference plane	deg
RA_OF_ASC_NODE	Right ascension of ascending node	deg
ARG_OF_PERICENTER	Argument of pericenter	deg
TRUE_ANOMALY or MEAN_ANOMALY	Position along the orbit	deg
GM	Gravitational coefficient (G × central mass)	km³/s²

Spacecraft Physical Parameters

OPM can include physical properties required for accurate orbit propagation:

MASS: Spacecraft mass (kg)
SOLAR_RAD_AREA: Cross-sectional area for solar radiation pressure (m²)
SOLAR_RAD_COEFF: Solar radiation pressure coefficient
DRAG_AREA: Cross-sectional area for atmospheric drag (m²)
DRAG_COEFF: Atmospheric drag coefficient

Common Use Cases

Orbit Determination Results: Sharing fitted state vectors after measurement processing
State Vector Import: Loading initial conditions into mission planning systems
Data Exchange: Transferring orbital state between organizations
Mission Analysis: Providing reference states for trajectory studies

Complete definition of the OPM standard on CCSDS 502.0-B-3 guidelines.

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 Optional Keplerian Elements

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
X                    = 36127.350 [km]
Y                    = 21794.520 [km]
Z                    = 36.800 [km]
X_DOT                = -1.588400 [km/s]
Y_DOT                = 2.632800 [km/s]
Z_DOT                = 0.000100 [km/s]

COMMENT Osculating Keplerian elements for the state above

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]
GM                   = 398600.4418 [km**3/s**2]

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

Importing OPM in VALAR

Import an OPM from the State Vectors page: click Import, choose OPM (CCSDS Standard) as the format, and upload a .txt or .opm file. VALAR resolves the target spacecraft automatically from the file’s OBJECT_ID (COSPAR ID) or OBJECT_NAME, so you do not select a spacecraft for CCSDS OPM imports. A successful import adds a state vector to that spacecraft.

VALAR imports the Cartesian state vector — position (X, Y, Z) in km and velocity (X_DOT, Y_DOT, Z_DOT) in km/s. An OPM that provides only Keplerian elements (SEMI_MAJOR_AXIS, ECCENTRICITY, …) and no Cartesian state vector is not a conformant OPM and cannot be imported. Re-export the OPM with the Cartesian state vector before importing.

Every imported OPM must also include a MASS value (kg), which VALAR uses to propagate the state.

Import Error Modes

If an import is rejected, the dialog shows a red alert with an error code and message. The most common OPM import errors:

Code	Cause	Remediation
OS-4004	The OPM has no Cartesian state vector — for example, it provides only Keplerian elements.	Re-export the OPM with the Cartesian state-vector block (`X`, `Y`, `Z`, `X_DOT`, `Y_DOT`, `Z_DOT`). Keplerian-only OPMs are not supported.
OS-4003	The file is not valid CCSDS OPM — unrecognized keywords, malformed values, or stray non-CCSDS text.	Check the file against the CCSDS OPM format and remove any non-CCSDS labels or stray text.
OS-4001	The OPM has no `MASS` keyword. Mass is required to propagate the imported state.	Add a `MASS` line (kg) to the OPM and re-import.
OS-4014	The OPM `MASS` is less than the spacecraft’s configured dry mass.	Correct the OPM mass, or update the spacecraft’s dry mass on the Spacecraft page.
OS-4016	The OPM has neither `OBJECT_ID` nor `OBJECT_NAME`, so VALAR cannot match it to a spacecraft.	Add an `OBJECT_ID` (COSPAR ID) or `OBJECT_NAME` that matches a spacecraft in your fleet.
OS-4015	The `OBJECT_ID` or `OBJECT_NAME` does not match any spacecraft in your fleet.	Create the spacecraft first, or correct the identifier in the OPM.
OS-4012	(SpaceX OPM) The OPM’s `OBJECT_ID` does not match the spacecraft you selected.	Select the matching spacecraft, or correct the OPM’s `OBJECT_ID`.
SV-4003	The state vector `EPOCH` is in the future.	Use an OPM whose `EPOCH` is at or before the current time (UTC).

Differences Between OPM and Other Formats

Feature	OPM	OEM	OCM
Temporal Scope	Single epoch	Time series	Single or multiple epochs
Maneuvers	One or more (impulsive or finite)	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

SpaceX OPM Format

SpaceX provides state vectors to customers via a proprietary OPM variant derived from the CCSDS standard. This format is generated from Falcon second stage flight telemetry and has several key differences from standard CCSDS OPM.

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.

Core Concepts

File Formats

Key Components

State Representation

Cartesian State Vector (Mandatory)

Osculating Keplerian Elements (Optional)

Spacecraft Physical Parameters

Common Use Cases

OPM with Optional Keplerian Elements

Importing OPM in VALAR

Import Error Modes

Differences Between OPM and Other Formats

SpaceX OPM Format

SpaceX OPM Fields

Key Differences from CCSDS OPM

Sample SpaceX OPM

Importing SpaceX OPM in VALAR

​Key Components

​State Representation

​Cartesian State Vector (Mandatory)

​Osculating Keplerian Elements (Optional)

​Spacecraft Physical Parameters

​Common Use Cases

​OPM with Optional Keplerian Elements

​Importing OPM in VALAR

​Import Error Modes

​Differences Between OPM and Other Formats

​SpaceX OPM Format

​SpaceX OPM Fields

​Key Differences from CCSDS OPM

​Sample SpaceX OPM

​Importing SpaceX OPM in VALAR

Key Components

State Representation

Cartesian State Vector (Mandatory)

Osculating Keplerian Elements (Optional)

Spacecraft Physical Parameters

Common Use Cases

OPM with Optional Keplerian Elements

Importing OPM in VALAR

Import Error Modes

Differences Between OPM and Other Formats

SpaceX OPM Format

SpaceX OPM Fields

Key Differences from CCSDS OPM

Sample SpaceX OPM

Importing SpaceX OPM in VALAR