ORBVIEW-2 Orbit Raising Operations Report

1.0 Summary

Space Exploration Engineering, Inc. planned and executed the orbit-raising task for the SeaStar (ORBVIEW-2) spacecraft launched August 1, 1997. During the previous three years, SEE worked with Orbital Sciences Corporation (OSC) to design the final target orbit, develop the Rapid Response Mission Analysis Tools (RR-MAT) to allow real-time orbit raising decisions to be made, and develop the operational scenario and constraints for the orbit raising task.

The main mission requirements to be met by the orbit were as follows:

1. The final orbit shall be sun-synchronous with a descending node local time of day equal to 12:05 PM. This means that the spacecraft will cross the equator of the earth on its southbound path when the local time of day below the spacecraft is 12:05 PM.

2. The final orbit shall maintain a descending node local time of day to within 15 minutes of Noon for the lifetime of the spacecraft (5 years).

3. The final orbit shall be approximately circular at 705 km altitude. (This forces the inclination to be 98.2175 deg. In order for the orbit to be sun-synchronous).

4. The final orbit shall be slightly elliptical, and the perigee of the orbit (closest approach point) shall be 'frozen', i.e. it shall stay within 10 degrees of the northernmost latitude of the orbit for the duration of the mission (this produces a constant altitude with latitude profile).

1.1 Estimated Orbview-2 Spacecraft Parameters

The Orbview-2 spacecraft was estimated to have the following parameters at launch, as specified by the OSC Ground Operations team:


Spacecraft Dry Weight:

Fuel:

Total Thrust:

Tank Pressure:

Thruster Duty Cycle

Maximum Allowed Yaw.

715 lbs.

150 lbs.

4.1 lbs. (4, 1 lbs Hydrazine Thrusters)

350 lbs

47% (Av. Thruster on-time during burn)

10 deg.


A pre-launch nominal burn plan, based upon the estimated insertion orbit and 3-sigma errors obtained from the Pegasus program office, was developed. A nominal burn plan was developed prior to launch that used 27 separate motor burns to achieve the target orbit without violating any of the 13 mission operations constraints. This burn plan raised the spacecraft to the final target orbit in a period of 30 days while maintaining ample margins in all parameters of interest. The nominal burn plan was used as a baseline for staffing and operations planning during the orbit-raising portion of the mission

Space Exploration Engineering developed the RR-MAT software to create the nominal burn plan, and used this same software to create the actual burn plan once the spacecraft was on-orbit. RR-MAT was used extensively before the launch of SeaStar to develop various contingency plans, as well as to refine the software and familiarize the operators with its use. RR-MAT contains an integrator with heritage from the JPL standard POHOP and ASAP code which was re-coded into the C programming language, along with various other orbital mechanics routines. It is implemented with a user interface that adds flexibility, simplicity and versatility. This interface, based on the commercial Microsoft Excel spreadsheet software, provids an extremely useful tool that has all the accuracy and heritage of the original code, and adds the ability to monitor a virtually unlimited number of mission constraints during the burn plan. Not only can numerous mission constraints be monitored, additional constraints can be added and monitored at any time with ease. SEE's RR-MAT software provided much more than just the ability to generate the Orbview-2 orbit-raising plan. In addition, it provided a general set of mission analysis tools that could be used to verify and debug problems with other subsystems, and track the progress of the mission.

RR-MAT was used in conjunction with the OASYS orbit determination software provided by Integral Systems Incorporated (ISI) to verify the performance of burns and then to plan subsequent burns.

As Table 1.1 illustrates, the actual insertion orbit varied considerably in apogee altitude from the predicted insertion orbit. This variance was outside the 3-sigma worst case, as estimated by the Pegasus program office. A post-launch nominal burn plan was created in real-time, and included an additional four motor burns to accommodate the lower apogee.


Mean Orbital Parameters

Predicted Insertion Orbit

Actual Insertion

Orbit

Error

Semi Major Axis

6733.4 km

6679.4 km

-54 km

Eccentricity

0.0061

0.0001

-0.006

True Anomaly

unconstrained

200.0 deg.

N/A

Arg. Of Perigee

unconstrained

241.9 deg.

N/A

Inclination

98.2175 deg.

98.274 deg.

+0.0565 deg.

Longitude of Node

307.675 deg.

307.8 deg

+0.125 deg.

Altitude at Perigee

310.0 km

299.9 km

-10 km

Altitude at Apogee

400.0 km

302.5 km

-97.5 km

Mean Time of Node

11:47:00 AM

11:47:29 AM

+29 sec

Epoch

 

8/2/1997 6:44:00 UTC

 


Table 1.1

Table 1.2 contains a brief description of the real-time anomalies encountered as the post-launch burn plan was being executed.

Anomaly

Response

Unreliable GPS data used to update ACS propagator resulting in loss of attitude knowledge. Spacecraft enters safe-mode aborting further burns.

Assist GPS and ACS subsystem engineers to generate state vectors following each burn. Replan groupings of burns to allow additional ground station contact and minimize possibility of ACS-induced aborts. Perform quick check real-time GPS state vector and confirm each burn execution.

Slight underperformance of propulsion system relative to predictions

Alter predictive model to accommodate. Target slightly greater performance than required to achieve desired results.

Uncommanded yaw resulting in inclination error accumulation. Errors occurred at random intervals but always resulted in a decrease in inclination

Replan specific burns to null accumulated error. Move final burns off of the equator to minimize inclination effect of uncommanded yaw.

Orbit determination performed incorrectly resulting in the targeting of the spacecraft to an incorrect orbit.

Replan final burn set to include inclination correction, as well as argument of perigee correction.


Table 1.2

The combined effect of these anomalies altered the list of the mission operations constraints and necessitated the addition of another burn near the end of the orbit-raising maneuvers.

The actual burn plan used 32 separate motor burns, met all of the 13 mission operations constraints, and raised the spacecraft to the final target orbit in a period of 31 days while maintaining ample margins in all parameters of interest. Table 1.3 contains the targeted final orbit, allowable errors, actual final orbit, and actual errors.

Parameter

Targeted Final Orbit

Allowable Error

Actual Final Orbit

Actual Error

Semi-major Axis

7083 km

5 km

7083 km

0.0 km

Eccentricity

0.001

0.001

0.001

0.0

Arg. Of Perigee

90.0

5.0 deg

93.64 deg

+ 3.64 deg

Inclination

98.2175 deg

0.02 deg

98.2187 deg

+0.0012 deg

Perigee Alt.

697 km

3 km

712 km

0.0 km

Apogee Alt.

712 km

3 km

712 km

0.0 km

Mean Time of D Node.

12:05 PM

5 minutes

12:05:14 PM

+ 15 sec


Table 1.3

2.0 Actual Performance vs. Predicted.

The following figures show the actual orbital and spacecraft parameters during the orbit raising vs. the predicted values from the pre-launch nominal burn plan. Figure 2.1 shows the planned apogee and perigee history vs. the actual history. Various unforeseen real-time constraints and anomalies caused the final altitude profile to deviate slightly from that of the nominal burn plan, but these deviations had no effect on the ability the spacecraft to reach the final required orbit.

Figure 2.1 Predicted Apogee and Perigee vs. Actual

Figure 2.2 shows the response of the orbit-raising team to anomalies in inclination during the course of orbit raising. The nominal plan naturally assumed that no errors in inclination corrections would occur, but these anomalies did occur, and SEE was able to respond to them and ultimately still target the inclination to its desired point.

Figure 2.2 Predicted Inclination vs.Actual

Figure 2.3 shows how the Descending Node Time of day tracked during orbit raising. The progression of this parameter was precisely planned by SEE prior to launch, and margin was built into the nominal orbit-raising plan to accommodate unforeseen delays and complications that might effect this parameter. As the figure shows, the planning was accurate and the margin was sufficient so that this mission requirement was met exactly.

Figure 2.3 Predicted vs. Actual Mean Descending Node Local Time

3.0 Conclusion

The SeaStar Orbit-Raising procedure was a complete success. The nominal orbit was achieved to an accuracy that was well within the error tolerances. Multiple, unforeseen anomalies including: low launch altitude, inclination errors, orbit determination errors, GPS errors, ACS pointing errors, and propulsion underperformance, were accommodated. The RR-MAT developed by Space Exploration Engineering, Inc. for this mission proved to be exceptionally flexible not only for its designed purpose, but also in the support of other subsystems. The RR-MAT software allowed for quick turn-around time when re-planning of a burn sequence was required, in addition to providing a 'quick-and-dirty' tool for doing 'sanity checks' on other subsystems.

Overall, the design and simplifying assumptions made when implementing the RR-MAT software were validated, as was the approach taken by SEE in meeting the requirements for the SeaStar mission.