DSPSE will be launched from Vandenberg Air Force Base in California via a refurbished Titan IIG booster, which is a Titan IIB with a modified nose cone. The Titan IIB was originally used to launch ICBM's. It will be launched at an inclination of 67o and placed in a nominal 140 x 160 nmi Low-Earth Orbit (LEO). The LEO orbit may be circularized at 160 nmi to reduce the effect of drag on the orbit. The vehicle will remain in LEO orbit for two days.
After two days in LEO, the vehicle will spin up to 60 rpm and a solid rocket motor will fire to place the spacecraft on a transfer trajectory to the Moon. (See Figure 1.) This burn is called the Transfer Trajectory Insertion (TTI) burn. Originally the solid motor was sized so that the total Delta-V obtained from it would place the radius of apogee of the spacecraft orbit's at lunar distance (distance from Earth to Moon); however, it was determined that the Isp of the onboard liquid motor was greater than the Isp of the solid motor and the total payload to orbit could be increased if the solid motor was staged with the liquid motor. Therefore, the spacecraft is put into an orbit that has a radius of apogee less than lunar distance.
The technique used for the lunar transfer is a phasing loop approach. The original phasing loop approach allows the s/c to remain in 2.5 transfer orbits before reaching the moon. The period of each orbit originally equaled that of the period of a direct injection orbit, similar to an Apollo period. Therefore, the period of a phasing loop was approximately 10 days, but the stunted phasing loops replaced the original non-stunted loops. Regardless, the spacecraft is inserted into a trans-lunar orbit 18 days prior to the required arrival date at the Moon. This allows the capability to correct for any off-nominal burns during the TTI sequence. These errors could include total magnitude Delta-V error, either an overburn or an underburn due to a different Isp than modeled or due to precession of the s/c during spinup, misalignment errors due attitude misalignment which would result in a line of apsides rotation, inclination differential, or a combination of both. Also, the phasing loops would allow for a greater TTI burn window since there is more time to correct errors. The bottom line is the phasing loop approach allows for lower Delta-V corrections since many corrections can be performed at apogee.
The phasing loops also allow multiple passes through the Van Allen radiation belts. Normally, this would not a good result; however, an objective of this mission is to test the radiation hardness of the sensors.
Finally, the s/c will perform autonomous navigation experiments during this phase by measuring the limb of the Earth and Moon and measuring the direction to a star simultaneously. A Kalman Filter uses measurements to estimate the position and velocity of the s/c.
Another experiment uses the laser ranging sensor to actively range and track the departing solid rocket motor as it is separated from the s/c. This will require equations of the Hill's/Euler type for relative motion in conjunction with the attitude control system. This also allows calibration of the sensors.
Following the phasing loops, the s/c will be placed in polar orbit about the Moon via a two burn sequence. The first burn will place the s/c in a high polar orbit with a large orbital period. The second burn, performed one day later, will place the s/c in the mapping orbit. The advantage of the two burn method is it allows time for accurate orbit determination about the Moon before the actual burn to the mapping orbit is performed. The mapping orbit has the argument of periselene 30o below the lunar equator for the first month of mapping. The mapping orbit for the second month has the argument of periselene at 30o above the lunar equator in the Northern hemisphere. This transfer to the Northern hemisphere will be acheived through a sequence of two burns with a transfer orbit resulting from the first burn and then the second month's mapping orbit resulting from the second burn.
The mapping orbit is designed such that the orbital period is an integral number of orbits of one sidereal lunar month (27.322 days) such that every other mapping strip is recorded during the first month and then the second month fills in those gaps. The width of the strip and the subsequent orbit depends on the sensor field of view and the stability of the lunar orbit. The orbit is polar with an eccentricity of 0.375 - 0.360, a semi-major axis of 3443 km, a period of 4 hours 59 minutes, a periselene altitude of 425± 25 km, and a longitude of the node of 179 degrees. The node depends on the desire to have the line joining the Sun-Moon-and-Earth to line up after the first month of mapping for scientific mapping purposes and it depends on the proper geometry so that the next phase of the mission to the asteriod can be accomplished.
The s/c will also perform the autonomous estimation experiment while in lunar orbit. This experiment will see limited time since much time will be spent mapping the Moon.
The lunar departure occurs much the same way as the lunar insertion with two burns, the first of which places the s/c in an orbit with a higher orbital period and then the second allows the s/c to escape from the Moon.
The next phase of the mission occurs with the s/c returning to Earth, performing a flyby of Earth, travelling to apogee which extends far beyond the lunar distance, then performing another flyby of Earth at perigee #2, and then performing a swingby of the Moon to gain additional energy to allow the s/c to achieve the correct orbit to the asteroid. (See Figure 2.) Neither of the flybys nor the swingby is power-assisted, which means changing the magnitude or direction of the velocity vector at perigee by burning the thrusters.
Transfer to Geographos:
The s/c then enters the transfer to the asteroid Geographos phase. The s/c will flyby the asteroid when the asteroid crosses the ecliptic plane. The position of the s/c and that of the asteroid must be well known to achieve a 100 km flyby. The s/c's position is determined through range and range-rate orbit determination. The asteroid is tracked optically during opposition and within the last few weeks prior to encounter. During the last few days, it is expected that radar measurements will be taken of the asteroid to improve its orbit.
The terminal encounter with the asteroid will involve identifying Geographos through use of the high- resolution visible sensor from the star background approximately one day prior to encounter. Once identified, the on-board autonomous estimation algorithms will estimate the relative state of the spacecraft and asteroid. Final trajectory burns will be performed in the last day. The estimate of the relative state does not change much nor improve until within minutes of the encounter. Due to the nature of the 100 km flyby and fact that the s/c will be 8 million km away from earth which limits command time, the flyby will have to be done autonomously. The algorithms will estimate the relative position and perform a B-plane targeting to estimate the necessary rotation rates and accelerations to be sent to the ACS. The flyby will include numerous images through the laser ranger, visible, near infrared, and long infrared sensors.
After the flyby, if enough propellant is left, a burn could be performed which would allow the s/c to return to Earth in 13 years.
Saturday, 28-Aug-1999 13:15:59 CDT
CSR/TSGC Team Web