LightSail-A: estimated post-sail deployment elements

From: Ted Molczan via Seesat-l <seesat-l_at_satobs.org>
Date: Sat, 6 Jun 2015 23:47:34 -0400
Communication with LightSail-A (15025L / 40661) resumed Saturday, and its solar sail may be deployed on Sunday 2015 Jun
07, not long after 18:02 UTC:

http://www.planetary.org/blogs/jason-davis/2015/20150506-lightsail-wakes-second-time-1.html

"If battery levels continue to trend stably during Sunday's early morning ground station passes, sail deployment will be
scheduled for 2:02 p.m. EDT (18:02 UTC)."

The current orbital elements are:

1 90726U          15156.84281789  .00025322  00000-0  67301-3 0   253
2 90726  55.0121 269.9399 0247556 221.6417 288.5573 15.12666037  2455

Observers between about 20 N and 56 N will have morning visibility of the orbit. Those south of about 43 S will have
evening visibility.

I estimate that with its sail deployed, LightSail-A's standard visual magnitude will be about 4.4 (1000 km range, 90 deg
phase angle), resulting in mag 2 to 3 on high-elevation, well illuminated passes. 

Judging by the much smaller NanoSail-D (10062L / 37361), LightSail-A's brightness may vary considerably from one pass to
another. It could be much fainter than expected, or flare to negative magnitudes. I suspect it will begin tumbling
during its first pass through perigee. Marco Langbroek analyzed the flash period of NanoSail-D: 

http://sattrackcam.blogspot.ca/2011/06/nanosail-d-evolution-of-flash-pattern.html

With its sail deployed, the rate of decay of LightSail-A will be enormous and difficult to predict with precision. I
estimate final descent late on 2015 Jun 9 UTC. I have estimated the following TLEs to assist in visual, optical and
radio tracking during the first 18 h post-sail deployment.

1 70001U          15158.79734543  .08393463  00000-0  21359 0 0    20
2 70001  55.0101 261.4702 0243810 226.4735 132.4935 15.13949689    03
1 70002U          15158.86337413  .08613548  00000-0  21534 0 0    62
2 70002  55.0093 261.1832 0240018 226.8670 132.2125 15.15058108    02
1 70003U          15158.92902983  .08770975  00000-0  21528 0 0    49
2 70003  55.0099 260.8978 0236172 227.2742 130.1542 15.16189165    06
1 70004U          15158.99551514  .09386835  00000-0  22594 0 0    28
2 70004  55.0104 260.6091 0232270 227.6713 132.9136 15.17355447    03
1 70005U          15159.06032682  .09589525  00000-0  22592 0 0    27
2 70005  55.0116 260.3268 0228342 228.0528 126.8258 15.18572200    00
1 70006U          15159.12662708  .09808663  00000-0  22573 0 0    85
2 70006  55.0113 260.0374 0224364 228.4612 129.1443 15.19843776    03
1 70007U          15159.19203951  .10166942  00000-0  22884 0 0    77
2 70007  55.0090 259.7499 0220091 228.8694 126.9040 15.21126993    04
1 70008U          15159.25842929  .10793040  00000-0  23694 0 0    96
2 70008  55.0065 259.4571 0215818 229.2351 130.3703 15.22476955    03
1 70009U          15159.32430804  .11150602  00000-0  23840 0 0    39
2 70009  55.0067 259.1661 0211305 229.5642 131.3984 15.23899019    09
1 70010U          15159.38976868  .11468796  00000-0  23854 0 0    66
2 70010  55.0080 258.8773 0206687 229.9966 130.3720 15.25358870    00
1 70011U          15159.45515580  .11800000  00000-0  23862 0 0    51
2 70011  55.0090 258.5890 0201864 230.4677 129.2531 15.26858693    03

For planning observations based on the above TLEs, I recommend comparing predictions against the pre-sail deployment
TLE, and taking 50% of the difference in prediction time as the uncertainty, e.g. if a 700XX TLE predicts a pass 5 min.,
earlier than the 90726 TLE, then allow 2.5 min. prediction time uncertainty.

I estimated the TLEs using the following procedure. I used SGP4 to propagate the 90726 TLE to the time of deployment,
which I took to be Jun 07 at 18:10 UTC. TLE Analyzer 2.12 converted the result to a state vector type compatible with
GMAT 2014a (General Mission Analysis Tool). 

I used GMAT to numerically propagate the orbit, based on the mass of the spacecraft (4.5 kg), the dimensions of its sail
(~5.66 x 5.66 m^2), my guess that it will be tumbling, and the forecast space weather.

GMAT produced Cartesian state vectors at close time intervals. I selected state vectors near successive ascending nodes
and converted them to TLEs. I computed the mean ndot/2 and corresponding B* decay terms between successive TLEs, and
inserted them into the TLEs.

I intend to revise these elements if the deployment occurs at a significantly different time, or subsequent radio
tracking reveals a significantly different rate of decay.

I also intend to produce TLEs past the period covered above, if I have sufficient data. 

Ted Molczan



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Received on Sat Jun 06 2015 - 22:49:01 UTC

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