4.0 Mission Details

4.1 Mission Timeline and Trajectory Details

Below is the general launch timeline for each P-STAR transponder delivery.   The actual launch windows of each launch date will definitely introduce some variance in these times, but they are qualitatively accurate:

Final Design - TK Model ms_sts.tk

Mission Time Event Mass
T+0:00 Shuttle Launch
2,040,000 kg
T+45:00 min LEO Delivery of Upper Stage
23,640 kg
T+2:00:00 hr Interplanetary Injection DV = 6.31 km/s
2,050 kg
T+2.73 year Powered Flyby (Semi-circularizing) Burn DV = 0.82 km/s
1,530 kg
T+2.73 year Heliocentric Z Component Addition Burn DV = 1.14 km/s
860 kg

With the help of the data from hohangle.f and position.f the following launch dates were chosen based primarily to provide significant displacements between each orbit.   Also, the first two launches are temporally spread out to provide ample time for any redesign or upgrading in the final launches that may be found critical for the operation of P-STAR.

Launch Date Long. Asc. Node 1 2 3 4 5 6
02 - 06 - 2000 129º - -162º -225º 103º 69º 32º
07 - 19 - 2005 291º 162º - -62º 229º 231º 194º
09 - 22 - 2007 353º 225º 62º - 327º 293º 256º
10 - 26 - 2008 26º -103º -229º -327º - -34º -71º
11 - 30 - 2009 60º -69º -231º -293º 34º - -37º
01 - 04 - 2011 97º -32º -194º -256º 71º 37º -

Note that the table lists the differences in each transponder's longitude of ascending node.   The closest two orbits are Transponders 1 and 6 with only a 32º separation in the X-Y plane, but they have an almost 60º (calculated from a 12.7 year period and 11 year period between delivery) difference in argument of perihelion, providing ample displacement.

4.2 Power Requirements

This is perhaps the trickiest part of this report.   As an aerospace engineer, trying to calculate the mass of the delivered payload without knowing the necessary components is the equivalent of telling the electrical engineers who would design the comm system:

This is the volume and mass you can play with...get the job done within these limitations.

This is not the best of situations for the comm design team to be in.   At 5-6 AU (and farther considering the orbits are slightly eccentric) Solar power is not what it is in the Earth orbit environs.   To make matters worse, there has been a worldwide outspoken resistance to Radio Thermal Generators (nuclear power supplies based upon a decaying isotopes thermal energy) not to mention full fledged fission reactors.   It is not clear if a fission reactor could even be designed to fit within the mass constraints of the current P-STAR launch and delivery regime.

So one must assume rather large Solar arrays which will siphon vital mass from the actual comm and chronometer subsystems of each transponder, yet there may be a silver lining to this cloud of power limitations.   It is conceivable that a large array could:

Both of these ideas are qualitative and are potential ways to offset the mass taken up by an effective solar array at Jovian distances.

4.3 Operations

The real meat of the P-STAR system is how it would change the business of spacecraft tracking and spacecraft operations.   The concept is that a Standard Operations Team would actively monitor the P-STAR system year round and spot check the onboard tracking algorithms of any spacecraft in interplanetary coast mode.   A Special Operations Team would be formed to focus on launches, trajectory correction maneuvers, and solar body encounters.   The special operations team would remain active in a rotation of projects:

Launch Mission A
Launch Mission B
Correct Trajectory A
Launch Mission C
Correct Trajectory B
Encounter with A
Correct Trajectory C
Correct Trajectory B
Encounter with C
Encounter with B

Where mission A and C would be mid-length tracking missions and mission C would be a longer tracking mission.   With standard team maintenance and automated tracking during cruise periods and focused attention where needed by the special operations team, a large number of missions could be tracked using two very skilled, very experienced, but small and utilitarian tracking groups.

4.4 Contingencies

Fuel contingencies of 2% at Hohmann insertion and powered flyby and of 4% at Z component addition were included in all mass calculations and estimates.   Two transponders were in place for redundancy in the event of mechanical failure or solar eclipse of a transponder.   In the event that two transponders should fail, a seventh transponder could be set op on Earth that would sweep signals outward from 1 AU.

Contents | Texas Space Grant | CSR Homepage | NASA | UT Aerospace

Tim Crain
Graduate Aerospace Engineering
The University of Texas at Austin
Last updated: