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The Orbit Comprehensive Message (OCM) is the most feature-rich standardized format for exchanging spacecraft orbital information. Unlike OPM (single-epoch state) and OEM (ephemeris time series), OCM combines multiple data types into a single comprehensive message, supporting trajectory state history, spacecraft physical properties, covariance time history, multiple maneuvers, perturbation parameters, and user-defined data.

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Why OCM is “Comprehensive”

OCM aggregates and extends the capabilities of OPM, OMM, and OEM into one unified format:
  • Multiple Trajectory States: Time series of state vectors or Keplerian elements (like OEM)
  • Single-Epoch States: State snapshots at specific times (like OPM)
  • Multiple Maneuvers: Comprehensive maneuver planning with impulsive and continuous thrust
  • Physical Properties: Spacecraft mass, area, drag/SRP coefficients, attitude information
  • Covariance Time History: Uncertainty evolution over time
  • Perturbation Parameters: Force model configuration for propagation
  • User-Defined Data: Custom parameters for mission-specific needs
  • Parent-Child Scenarios: Support for deployment and separation events
OCM is the recommended format when you need to exchange comprehensive mission planning data, including orbital states, maneuver plans, spacecraft properties, and uncertainties in a single file.

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Key Components

An OCM file contains several distinct sections (all optional except Header and Metadata):
  • Header: File-level metadata (format version, creation date, originator)
  • Metadata: Mission context (object identification, time system, reference frame)
  • Trajectory State Time History: Cartesian state vectors or Keplerian elements over time
  • Physical Descriptor: Spacecraft physical properties (mass, area, drag coefficient, etc.)
  • Covariance Time History: Position/velocity uncertainty matrices over time
  • Maneuver Parameters: Multiple maneuver definitions (impulsive or continuous thrust)
  • Perturbation Parameters: Force model specification for propagation
  • Orbit Determination Parameters: OD configuration and results
  • User-Defined Parameters: Custom mission-specific data

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Structure Overview

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OCM Hierarchy

OCM File
├── Header (CCSDS_OCM_VERS, CREATION_DATE, ORIGINATOR)
├── Metadata Section (META_START to META_STOP)
│   ├── Time system, reference frame
│   ├── Object identification
│   ├── Epoch reference (EPOCH_TZERO)
│   └── Trajectory interpolation settings
├── Trajectory State Blocks (TRAJ_START to TRAJ_STOP) [Optional]
│   └── Multiple state vector or Keplerian element time series
├── Physical Descriptor Block (PHYS_START to PHYS_STOP) [Optional]
│   └── Mass, area, attitude, drag/SRP coefficients
├── Covariance Blocks (COV_START to COV_STOP) [Optional]
│   └── Time series of covariance matrices
├── Maneuver Blocks (MAN_START to MAN_STOP) [Optional]
│   └── Multiple maneuver definitions with unique IDs
├── Perturbation Block (PERT_START to PERT_STOP) [Optional]
│   └── Force model configuration
├── Orbit Determination Block (OD_START to OD_STOP) [Optional]
│   └── OD settings and results
└── User-Defined Block (USER_START to USER_STOP) [Optional]
    └── Custom parameters

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Metadata Fields

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Required Metadata

FieldDescriptionExample Values
TIME_SYSTEMTime system for all epochsUTC, TAI, GPS
EPOCH_TZEROReference epoch for relative times2024-01-01T00:00:00

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Object Identification

FieldDescriptionExample
OBJECT_NAMESpacecraft common nameSPACECRAFT-ALPHA
INTERNATIONAL_DESIGNATORCOSPAR ID2023-001A
CATALOG_NAMECatalog identifier sourceSATCAT
OBJECT_DESIGNATORObject ID in catalog12345

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Trajectory Configuration

FieldDescriptionValues
REF_FRAMEReference frame for statesGCRF, EME2000, ITRF
REF_FRAME_EPOCHEpoch for reference frameJ2000 epoch
TRAJ_TYPETrajectory representationCARTESIAN, KEPLERIAN, GEODETIC
TRAJ_UNITSUnits for trajectory data[km, km, km, km/s, km/s, km/s]
EPOCH_TZERO serves as the time origin for all relative time references in the OCM. Absolute times can also be used where appropriate.

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Trajectory State Time History

OCM supports time series of orbital states in multiple representations:

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Cartesian State Vectors

TRAJ_START
TRAJ_TYPE   = CARTESIAN
TRAJ_UNITS  = [km, km, km, km/s, km/s, km/s]

EPOCH       X           Y           Z           X_DOT       Y_DOT       Z_DOT
+0.000      6655.994    -40218.575  -82.918     3.115482    0.470426    -0.001015
+3600.000   6892.441    -39845.123  -76.234     3.087234    0.501287    -0.000923
+7200.000   7112.873    -39458.892  -69.445     3.058012    0.531456    -0.000831
TRAJ_STOP

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Keplerian Elements

TRAJ_START
TRAJ_TYPE   = KEPLERIAN
TRAJ_UNITS  = [km, n/a, deg, deg, deg, deg]

EPOCH       SEMI_MAJOR_AXIS  ECCENTRICITY  INCLINATION  RA_ASC_NODE  ARG_PERICENTER  TRUE_ANOMALY
+0.000      42164.140        0.0001234     0.05         75.123       180.456         90.789
+3600.000   42164.245        0.0001245     0.05         75.234       180.567         95.123
TRAJ_STOP
Time values can be:
  • Absolute: Full ISO 8601 timestamp
  • Relative: Offset in seconds from EPOCH_TZERO (indicated by + or - prefix)

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Physical Descriptor Block

Spacecraft physical properties for accurate propagation: 
PHYS_START
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
WET_MASS            = 2000.000 [kg]
DRY_MASS            = 1850.000 [kg]
OEB_PARENT_FRAME    = SPACECRAFT_BODY
OEB_Q1              = 0.707
OEB_Q2              = 0.000
OEB_Q3              = 0.000
OEB_QC              = 0.707
PHYS_STOP

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Key Physical Parameters

ParameterDescriptionUnits
MASSCurrent spacecraft masskg
WET_MASSFully fueled masskg
DRY_MASSMass without propellantkg
SOLAR_RAD_AREAEffective area for SRPm²
SOLAR_RAD_COEFFSRP coefficient (typically 1.0-1.5)dimensionless
DRAG_AREAEffective area for atmospheric dragm²
DRAG_COEFFDrag coefficient (typically 2.0-2.5)dimensionless
OEB_Q1, Q2, Q3, QCAttitude quaterniondimensionless

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Covariance Time History

OCM supports time-varying covariance matrices to represent uncertainty evolution: 
COV_START
COV_TYPE        = CARTESIAN
COV_ORDERING    = LTRI
COV_UNITS       = [km**2, km**2, km**2, km**2/s**2, km**2/s**2, km**2/s**2]

EPOCH           CX_X    CY_X    CY_Y    CZ_X    CZ_Y    CZ_Z
+0.000          4.142e+01  -8.579e+00  5.189e+01  -4.788e+01  -4.472e+01  5.960e+01

EPOCH           CX_DOT_X    CX_DOT_Y    CX_DOT_Z    CX_DOT_X_DOT
+0.000          3.523e-03   1.045e-03   -1.234e-03  8.945e-07

EPOCH           CY_DOT_X    CY_DOT_Y    CY_DOT_Z    CY_DOT_X_DOT    CY_DOT_Y_DOT
+0.000          -2.678e-03  -5.234e-04  1.876e-03   -3.456e-07      7.234e-07

EPOCH           CZ_DOT_X    CZ_DOT_Y    CZ_DOT_Z    CZ_DOT_X_DOT    CZ_DOT_Y_DOT    CZ_DOT_Z_DOT
+0.000          -1.234e-03  2.345e-03   -8.765e-04  2.134e-07       -4.567e-07      6.789e-07
COV_STOP
The covariance matrix is stored in lower-triangular form, representing the 6×6 position-velocity uncertainty.

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Maneuver Parameters

OCM’s most powerful feature is comprehensive maneuver specification supporting both impulsive and continuous thrust maneuvers.

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Impulsive Maneuver

A single instantaneous velocity change: 
MAN_START
MAN_ID              = MAN-001
MAN_BASIS           = PLANNED
MAN_DEVICE_ID       = THRUSTER-PRIMARY
MAN_PREV_ID         = N/A
MAN_NEXT_ID         = MAN-002
MAN_REF_FRAME       = RTN
MAN_FRAME_EPOCH     = 2024-06-03T09:00:00.000

DC_TYPE             = IMPULSIVE
MAN_COMPOSITION     = TIME_ABSOLUTE, MAN_DV_1, MAN_DV_2, MAN_DV_3
MAN_UNITS           = [s, km/s, km/s, km/s]

2024-06-03T09:00:34.100  -0.02325700  0.01683160  -0.00893444
MAN_STOP

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Continuous Thrust Maneuver

Extended burn with time-varying thrust profile: 
MAN_START
MAN_ID              = MAN-002
MAN_BASIS           = PLANNED
MAN_DEVICE_ID       = THRUSTER-ION
MAN_REF_FRAME       = RTN
DC_TYPE             = CONTINUOUS
MAN_COMPOSITION     = TIME_ABSOLUTE, MAN_DURA, THR_X, THR_Y, THR_Z, THR_ISP, DELTA_MASS
MAN_UNITS           = [s, s, N, N, N, s, kg]

2024-06-04T01:10:00.000  500.0  0.0000  0.0650  0.0000  3000.0  -0.78
2024-06-04T01:20:00.000  500.0  0.0000  0.0000  0.0650  3000.0  -0.78
2024-06-04T01:30:00.000  500.0  0.0650  0.0000  0.0000  3000.0  -0.78
MAN_STOP

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Maneuver Metadata Fields

FieldDescriptionValues
MAN_IDUnique maneuver identifierMAN-001, BURN-A, etc.
MAN_BASISManeuver statusPLANNED, ACCEPTED, ACTUAL
MAN_DEVICE_IDThruster or device IDThruster name or ALL
MAN_PREV_IDPrevious maneuver in sequenceMAN_ID or N/A
MAN_NEXT_IDNext maneuver in sequenceMAN_ID or N/A
MAN_REF_FRAMEReference frame for delta-VRTN, LVLH, TNW, inertial
DC_TYPEManeuver typeIMPULSIVE, CONTINUOUS

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MAN_COMPOSITION — Customizable Data Columns

The MAN_COMPOSITION field defines which data columns appear in the maneuver data rows. This allows flexible maneuver specification tailored to the mission needs. Common composition fields:
FieldDescriptionTypical Units
TIME_ABSOLUTEAbsolute time stampISO 8601 string
TIME_RELATIVETime relative to EPOCH_TZEROseconds
MAN_DURABurn durationseconds
MAN_DV_1, MAN_DV_2, MAN_DV_3Delta-V componentskm/s
THR_X, THR_Y, THR_ZThrust componentsNewtons
THR_MAGThrust magnitudeNewtons
THR_ISPSpecific impulseseconds
DELTA_MASSMass change (negative for consumption)kg
THR_EFFICIENCYThruster efficiency0.0 to 1.0
The order of fields in MAN_COMPOSITION must exactly match the order of data columns in subsequent maneuver data rows. Units must be specified in MAN_UNITS in the same order.

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Maneuver Chaining

OCM supports maneuver sequences through MAN_PREV_ID and MAN_NEXT_ID, enabling:
  • Maneuver dependency tracking
  • Sequence validation
  • Mission timeline reconstruction

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Perturbation Parameters

Define the force model configuration for orbit propagation: 
PERT_START
ATMOSPHERIC_MODEL       = NRLMSISE00
GRAVITY_MODEL           = EGM-96: 70D 70O
EQUATORIAL_RADIUS       = 6378.137 [km]
GM                      = 398600.4415 [km**3/s**2]
N_BODY_PERTURBATIONS    = SUN, MOON
SOLAR_RAD_PRESSURE      = YES
SOLID_EARTH_TIDES       = YES
OCEAN_TIDES             = YES
THIRD_BODY_GRAVITY      = YES
PERT_STOP

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Perturbation Parameters

ParameterDescriptionExample Values
ATMOSPHERIC_MODELAtmospheric density modelNRLMSISE00, JB2008, Harris-Priester
GRAVITY_MODELGeopotential model with degree/orderEGM-96: 70D 70O, WGS84: 8D 8O
GMGravitational parameter398600.4415 km³/s²
N_BODY_PERTURBATIONSThird-body perturbationsSUN, MOON, JUPITER
SOLAR_RAD_PRESSUREInclude SRPYES, NO
SOLID_EARTH_TIDESInclude solid Earth tidesYES, NO
OCEAN_TIDESInclude ocean tidesYES, NO
Perturbation parameters ensure the receiving system can accurately propagate the orbit using the same force model configuration as the originating system.

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User-Defined Parameters

OCM allows custom parameters for mission-specific data: 
USER_START
USER_PARAM_1 = MISSION_PHASE
USER_VALUE_1 = OPERATIONAL
USER_PARAM_2 = BATTERY_SOC
USER_VALUE_2 = 87.5 [%]
USER_PARAM_3 = CONTACT_FREQUENCY
USER_VALUE_3 = 2.2e9 [Hz]
USER_STOP
User-defined parameters should be used sparingly and documented in Interface Control Documents (ICDs). Their meaning must be clearly communicated to all parties exchanging the OCM file.

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Common Use Cases

  • Comprehensive Mission Planning: Complete mission state including trajectory, maneuvers, and spacecraft properties
  • Maneuver Sequencing: Planning and tracking multiple burn sequences
  • Conjunction Data Submission: Providing detailed orbital information to conjunction assessment services
  • Formation Flying: Parent-child deployment scenarios and relative state information
  • Mission Handover: Comprehensive state transfer between control centers
  • Long-Duration Missions: Time-varying properties (mass, drag coefficient) over mission lifetime
  • High-Fidelity Propagation: Providing complete force model and spacecraft parameters
Complete definition of the OCM standard in CCSDS 502.0-B-3 specification.

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Complete OCM Example

Here is a comprehensive OCM file demonstrating multiple features: 
CCSDS_OCM_VERS     = 3.0
CREATION_DATE      = 2024-06-15T08:30:00.000
ORIGINATOR         = VALAR

COMMENT Comprehensive mission planning file
COMMENT Includes trajectory, maneuvers, and physical properties

META_START
TIME_SYSTEM        = UTC
EPOCH_TZERO        = 2024-06-15T00:00:00.000
OBJECT_NAME        = SPACECRAFT-GAMMA
INTERNATIONAL_DESIGNATOR = 2024-001A
CATALOG_NAME       = SATCAT
OBJECT_DESIGNATOR  = 59234
REF_FRAME          = EME2000
TRAJ_TYPE          = CARTESIAN
TRAJ_UNITS         = [km, km, km, km/s, km/s, km/s]
META_STOP

COMMENT Initial trajectory state history

TRAJ_START
EPOCH       X           Y           Z           X_DOT       Y_DOT       Z_DOT
+0.000      6655.994    -40218.575  -82.918     3.115482    0.470426    -0.001015
+3600.000   6892.441    -39845.123  -76.234     3.087234    0.501287    -0.000923
+7200.000   7112.873    -39458.892  -69.445     3.058012    0.531456    -0.000831
TRAJ_STOP

COMMENT Spacecraft physical properties

PHYS_START
MASS                = 1913.000 [kg]
WET_MASS            = 2000.000 [kg]
DRY_MASS            = 1850.000 [kg]
SOLAR_RAD_AREA      = 10.000 [m**2]
SOLAR_RAD_COEFF     = 1.300
DRAG_AREA           = 10.000 [m**2]
DRAG_COEFF          = 2.300
PHYS_STOP

COMMENT First orbit-raising maneuver

MAN_START
MAN_ID              = RAISE-PERIGEE-1
MAN_BASIS           = PLANNED
MAN_DEVICE_ID       = THRUSTER-MAIN
MAN_PREV_ID         = N/A
MAN_NEXT_ID         = RAISE-APOGEE-1
MAN_REF_FRAME       = RTN
DC_TYPE             = IMPULSIVE
MAN_COMPOSITION     = TIME_ABSOLUTE, MAN_DV_1, MAN_DV_2, MAN_DV_3
MAN_UNITS           = [s, km/s, km/s, km/s]

2024-06-15T09:00:34.100  0.05234000  0.00124500  -0.00012300
MAN_STOP

COMMENT Second orbit-raising maneuver

MAN_START
MAN_ID              = RAISE-APOGEE-1
MAN_BASIS           = PLANNED
MAN_DEVICE_ID       = THRUSTER-MAIN
MAN_PREV_ID         = RAISE-PERIGEE-1
MAN_NEXT_ID         = CIRCULARIZE-1
MAN_REF_FRAME       = RTN
DC_TYPE             = IMPULSIVE
MAN_COMPOSITION     = TIME_ABSOLUTE, MAN_DV_1, MAN_DV_2, MAN_DV_3
MAN_UNITS           = [s, km/s, km/s, km/s]

2024-06-15T21:45:12.300  0.04876000  -0.00087600  0.00023400
MAN_STOP

COMMENT Circularization burn using low-thrust ion propulsion

MAN_START
MAN_ID              = CIRCULARIZE-1
MAN_BASIS           = PLANNED
MAN_DEVICE_ID       = ION-THRUSTER-1
MAN_PREV_ID         = RAISE-APOGEE-1
MAN_NEXT_ID         = N/A
MAN_REF_FRAME       = RTN
DC_TYPE             = CONTINUOUS
MAN_COMPOSITION     = TIME_ABSOLUTE, MAN_DURA, THR_X, THR_Y, THR_Z, THR_ISP, DELTA_MASS
MAN_UNITS           = [s, s, N, N, N, s, kg]

2024-06-16T10:00:00.000  600.0  0.0000  0.0850  0.0000  3200.0  -0.92
2024-06-16T10:10:00.000  600.0  0.0000  0.0850  0.0000  3200.0  -0.92
2024-06-16T10:20:00.000  600.0  0.0000  0.0850  0.0000  3200.0  -0.92
MAN_STOP

COMMENT Perturbation parameters for propagation

PERT_START
ATMOSPHERIC_MODEL       = NRLMSISE00
GRAVITY_MODEL           = EGM-96: 70D 70O
EQUATORIAL_RADIUS       = 6378.137 [km]
GM                      = 398600.4415 [km**3/s**2]
N_BODY_PERTURBATIONS    = SUN, MOON
SOLAR_RAD_PRESSURE      = YES
SOLID_EARTH_TIDES       = YES
OCEAN_TIDES             = NO
PERT_STOP

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Usage in VALAR

VALAR supports OCM for maneuver operations:

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Import of Maneuvers

  • OCM Import: Upload OCM files to import maneuver definitions into VALAR
  • Maneuver Parsing: VALAR extracts maneuver blocks (MAN_START to MAN_STOP) from OCM files
  • Both Types Supported: Import impulsive and continuous thrust maneuvers
  • Maneuver Sequencing: MAN_PREV_ID and MAN_NEXT_ID relationships are preserved

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Export of Maneuver Plans

  • OCM Export: Export maneuver plans from VALAR as OCM files
  • Multi-Maneuver Support: Multiple planned maneuvers are included in a single OCM file
  • Standard Compliance: Exported OCM files follow CCSDS 502.0-B-3 specification
  • Sharing with External Systems: OCM format enables maneuver plan exchange with other agencies
When importing OCM files, ensure all referenced MAN_DEVICE_ID entries match configured thrusters in VALAR. Unrecognized device IDs will cause import errors.
VALAR focuses on the maneuver portions of OCM. Other OCM sections (trajectory state history, physical properties, covariance, perturbations) are not currently imported or exported by VALAR.

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Best Practices

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Data Organization

  1. Use EPOCH_TZERO effectively: Set to mission start or a significant epoch for cleaner relative times
  2. Include physical properties: Always provide mass, area, and coefficient data for accurate propagation
  3. Document maneuver sequences: Use MAN_PREV_ID and MAN_NEXT_ID to chain related maneuvers

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Maneuver Specification

  1. Choose appropriate DC_TYPE: IMPULSIVE for chemical burns, CONTINUOUS for ion/low-thrust
  2. Specify reference frames clearly: Document whether RTN, LVLH, or other LOF is used
  3. Include all composition fields: Don’t omit important fields like DELTA_MASS or THR_ISP

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File Quality

  1. Add comments liberally: Explain mission phases, maneuver purposes, assumptions
  2. Validate before sharing: Ensure all referenced IDs exist and data is self-consistent
  3. Use appropriate units: Follow CCSDS conventions and specify units explicitly in MAN_UNITS

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Interoperability

  1. Minimize user-defined parameters: Use standard fields whenever possible
  2. Document extensions: If using USER blocks, provide clear documentation in ICDs
  3. Test round-trip: Verify OCM can be imported and re-exported without data loss

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Comparison with Other Formats

FeatureOCMOPMOEMTDM
Data TypeComprehensiveSingle stateEphemerisObservations
Temporal ScopeTime series + snapshotsSingle epochTime seriesObservation period
ManeuversMultiple (full specification)One impulsiveNoneNone
Physical PropertiesExtensiveBasicNoneNone
CovarianceTime-varyingOptional singleOptional per epochNone
PerturbationsForce model specificationNoneNoneNone
Use CaseMission planningState exchangeTrajectory sharingOD input
ComplexityHighLowMediumMedium

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When to Use OCM

Use OCM when you need:
  • Multiple maneuvers in a single file
  • Time-varying spacecraft properties (mass, drag coefficient)
  • Complete force model specification for propagation
  • Maneuver sequencing and dependencies
  • Parent-child deployment scenarios
  • Comprehensive mission state exchange
Use OPM instead when:
  • Sharing a simple single-epoch state vector
  • Only one maneuver needs to be specified
  • Simplicity and minimal file size are priorities
Use OEM instead when:
  • Providing pre-computed ephemeris for interpolation
  • No maneuvers or physical properties needed
  • Standard trajectory exchange format expected