Satellite orbital instability generated by the perturbing factors.
Chiru, Anghel ; Pirna, Ion ; Candea, Ioan 等
1. INTRODUCTION
Due to the increased dependence of humankind on space technologies,
the impact of malfunctioning satellites would be great. In this
situation, research on the perturbing factors and the calculation of
their influence upon the orbit should be a prime step for the orbital
correction problems.
Until not long ago the shape of any orbit (as demonstrated by
Johannes Kepler in 1609 in "Astronomia nova") was thought to
be an ellipse with the central body in one of the foci. Analyzing the
real orbital trajectory using data collected during actual space
missions it became clear that there is a discrepancy between the
predicted and the real orbit.
The new problem was to find what exterior forces are acting upon
the spacecraft and then to try to calculate them in order to make better
predicaments for the trajectory.
A series of perturbing factors were presented as influencing the
movement on the orbit: gravity of other massive celestial bodies,
friction with the upper atmosphere, solar wind and solar pressure. Some
other insignificant factors as the flattening of the Earth or the
variable magnetic field are not taken into account in this paper. The
induced perturbations are very small, but taken a long period of time,
the impact upon the satellite orbit is significant
Each of these perturbing factors is at the moment well known and
can be calculated using specific formulas.
The novelty presented in this paper is the development of a method
that takes into account the added effect of all this factors. Based on
this model a recursive step-by-step computer program was created and
used to accurately predict the orbits.
2. GRAVITY OF CELESTIAL BODIES
The Sun, Moon and other planets can modify the trajectory of a
satellite by their gravity. The most significant bodies are the Sun
(huge mass) and the Moon (due to proximity). Equation 1 determines the
gravity forces that are acting on the satellite in a system with n
bodies (Schaub & Junkins, 2002).
[MATHEMATICAL EXPRESSION NOT REPRODUCIBLE IN ASCII] (1)
[FIGURE 1 OMITTED]
Where G = 6.67e-11 [m.sup.3]/[s.sup.2]kg is the universal constant
of attraction, [a.sub.g] is the gravity induced acceleration,
[m.sub.sat] is the mass of the satellite, [m.sub.j] is the mass of the
n'th body and [r.sub.sat j] is the distance between the satellite
and the n'th body.
It can be noticed form the figure 1 that the gravitational
acceleration is decreasing rapidly. This acceleration is directly
proportional with the attracting body mass, therefore the largest
celestial bodies will induce higher accelerations. But the acceleration
is inverse proportional with the distance, therefore the closest bodies
will induce larger accelerations and more distant planets will induce
only small accelerations.
Perturbations coming from these forces are relevant for satellites.
The acceleration induced by the Sun is almost constant (5.9e-3
m/[s.sup.2]), the acceleration induced by the Moon varies with the
satellite's position around the Earth (3e-5 m/[s.sup.2] .. 4e-5
m/[s.sup.2]). The accelerations induced by the other planets are very
small (1e-7 m/[s.sup.2]).
3. ATMOSPHERIC DRAG
The atmospheric drag is acting as a breaking force. The orbital
height of the satellite will decrease slightly.
The drag force is dependent on the air density (which decreases
with altitude), therefore it is inversely proportional with the
altitude. The atmospheric drag can be one of the main perturbations if
working on a satellite in LEO (Low Earth Orbit).
The acceleration from the atmospheric drag force can be expressed
as presented in equation 2 (Deliu, 2003):
[a.sub.d] = 1/2 [rho] x [C.sub.d] x A x [v.sup.2]/[m.sub.sat] (2)
where: p is the density of the air, [C.sub.d] is the drag
coefficient, A is the reference transversal area and v is the velocity
of the satellite relative to the air.
As a reference, the Space Shuttle (2008) with a mass of 92000 kg
and a sectional area of 362 [m.sup.2] at an altitude of 350 km has an
induced drag acceleration of 5e-3 m/[s.sup.2].
[FIGURE 2 OMITTED]
The variation of the induced drag acceleration upon the Space
Shuttle can be also deduced from figure 2.
4. SOLAR RADIATION PRESSURE
The Sun's radiation causes a small force on the spacecraft
that is exposed to it. This is because the Sun emits photons that are
either absorbed or reflected by the satellite. Therefore, the force
experienced by the satellite depends upon the surface area of the
satellite. The acceleration can be expressed as presented in equation 3
(Campbell & McCandless, 1996).
[a.sub.sp] = s x (1 - f) x k x P x A/c x 1/[m.sub.sat] (3)
where: k is the reflectivity factor and can be set between 1 and 2
(1 means that the radiation is totally absorbed, 2 means that it is
totally reflected; P is the solar power (=1400W/m2 at Earth's
location), A is the reference transversal area and c is the speed of
light.
The s factor is used to take into account the solar wind influence
(1< s < 1.2). It is dependent of solar activities. For maximum
solar activity, the coefficient can be taken as 1.2 and for low solar
activities the coefficient can be set to 1.
A normal satellite has a solar pressure induced acceleration of
1e-8m/[s.sup.2].
The f coefficient represents the time when the Earth is between the
Sun and the satellite and therefore the solar pressure and the solar
wind are no longer active on the craft. In figure 3 is presented the
light period for a family of orbits, from one can extract the f
coefficient.
In figure 3 can be observed the variation of the light function of
the orbital radius. Curve 1 represents an orbit with an inclination of 0
deg. (orbit in the ecliptic plane). Curve 2, 3 and 4 are inclined with
10, 30 and 50 deg. with respect at the ecliptic plane.
[FIGURE 3 OMITTED]
5. ONBOARD THRUSTERS SYSTEM
The satellite's own thrusters can change the orbit. During a
thrust orbital maneuver, the mass of the satellite will change as
propellant is consumed.
A simple, constant thrust model is however often sufficient to
describe the motion of a spacecraft during thrust arcs.
When a propulsion system ejects a mass m per time interval dt at a
velocity [v.sub.e], the spacecraft of mass [m.sub.sat] experiences a
thrust F which results in the acceleration presented in equation 4
(Frank, 1998).
[a.sub.t] = F/[m.sub.sat] = [absolute value of [??]]/[m.sub.sat]
[v.sub.e] (4)
6. IMPACT ON THE MOTION EQUATION
Starting from the gravitational force equation expression, the
acceleration of the satellite is expressed as in equation 5.
a = G x M/[r.sup.2] + [a.sub.g] + [a.sub.d] + [a.sub.sp] +
[a.sub.t] (5)
where M is the mass of the Earth.
Based on this equation, a computer program was written using
recursive functions and dividing time into small steps. Starting with
the known positions and speeds of all bodies, and giving values to
different parameters, the program computes the new position and speeds
over a small time step. Now this new positions and speeds are used as
the input for the next step.
Results were validated using as input data available online
information for the Hubble Space Telescope (***, 2008). Actual
calculated perturbations, for a time range of 1 year, are much accurate,
in the range of 1% difference from the real orbit compared with the
theoretical orbit. The previous calculations based on older model give
differences of 5% or more.
Future research could take into consideration some other smaller
perturbing factors and try to reduce even more the difference between
the predicted orbit and the real orbit. Also, the model has to be
validated using real data.
7. CONCLUSIONS
Using the equation presented above, one can calculate with an
iterative computer program the positions and acceleration for a
satellite. The iteration steps should be small in order to keep the
errors as low as possible. Then, the perturbation can be calculated
knowing both the real position and the theoretical position. With these
calculations, the correction of the orbit can be computed.
8. REFERENCES
Campbell, B. A.; McCandless W. (1996). Introduction to Space
Sciences and Spacecraft Applications, Gulf Publishing, ISBN 0-88415-411-4, Houston
Deliu, G. (2003). Aircraft Mechanics, Ed. Albastra, ISBN
973-650-029-2, Cluj-Napoca, Romania
Frank, W. B. (1998). Propulsion Flight Research at NASA, Dryden
Flight Research Center, ISBN 0-8027-1427-7, California
Schaub, H. & Junkins, J. L. (2002). Analytical Mechanics of
Aerospace Systems, Dover Publications, ISBN 1563475634, New York
*** (2008) http://hubble.nasa.gov/- National Aeronautics and Space
Administration, Hubble Space Telescope, Accessed on: 2008-04-11