Stellar Day Orbits

Launch loop vehicle mass is limited; however, with high precision launch it is "relatively" easy to rendezvous vehicles to connect into larger assemblies. For example, launching rocket/heatshield packages, followed later by cargo packages to complete a vehicle for apogee plane change and subsequent aerobraking-assisted orbit changes.

The launch loop will rotate under the perigee of a previous launch once per stellar day, every 86164.0989 seconds, about 4 minutes shorter than a solar day. Launch times should be scheduled for exact integer multiples of these orbit times. Otherwise, the loop will not be exactly positioned at the perigee time.

The actual rendezvous time will be modified by the nonlinear gravity field of the Earth due to the equatorial bulge, and the gravity of the Sun, Moon, and Jupiter. These perturbations are not included here, but for a real system they can be predicted and computed to millimeter accuracy. There will also be less-predictable perturbations due to residual atmospheric gas drag (a function of solar UV activity), spacecraft out-gassing, light pressure, and other small effects. The following is for rough estimates only.

The Earth orbital period T (in seconds) is (to first order) a function of the standard gravitational parameter μ = 398600.44 km³/s² and the semimajor axis a . ~~~ T =~ \sqrt{ a^3 / \mu } ~ . ~ ~ The semimajor axis is half the sum of the perigee and apogee radius: 2 ~ a ~=~ r_a + r_p . So, we can rearrange to compute the sum, r_a + r_p = \left( 2 \mu T^2 / \pi^2 \right)^{1/3} . If D is the number of sidereal days, t = 86141.09 \times D . ombining terms, r_a + r_p = 84313.33 \times D^{2/3} km .

 r_a + r_p in kilometers versus D in stellar days D r_a + r_p D r_a + r_p D r_a + r_p D r_a + r_p D r_a + r_p S. days kilometers S. days kilometers S. days kilometers S. days kilometers S. days kilometers 1 84328.34 6 278446.06 12 442005.57 25 720997.21 50 1144511.72 2 133862.90 7 308583.19 14 489845.29 30 814181.22 55 1219594.25 3 175410.03 8 337313.38 16 535451.61 35 902302.73 60 1292432.12 4 212494.11 9 364867.56 18 579191.14 40 986310.30 65 1363271.89 5 246577.58 10 391417.50 20 621336.56 45 1066879.21 70 1432316.30

Choose a perigee r_p to avoid existing satellite orbits. Subtract that from the sum to compute the apogee radius.

The average radius of the Moon is around 385,000 km, and orbits to those large distances will be strongly influenced by lunar gravity. Beyond 1.5 million kilometers, solar tidal forces are stronger than the Earth's gravity. These very slow orbits cannot be estimated simply.

While difficult to compute, and requiring continual accurate measurement and tiny correction thrusts, the multi-body tidal effects can be exploited to pump energy into or out of long period orbits; a subject for more adept orbit calculators than I am. Belbruno's Fly Me to the Moon describes these "chaotic" effects for lay audiences.

StellarDayOrbits (last edited 2018-03-29 07:10:33 by KeithLofstrom)