. . . that mean no need to do adjustments for an Earth’s inclination to the ecliptic. Everything was ready to add engine's impulses to a simulation. Impulses was imported into TRA application as a data from plot on engine’s firing tests . . . For a convenience it is possible to use parameter PropCoeff to adjust a total impulse. For sure real test’s data of a real engine should be use – but for an approximation it is ok for now. Also that PropCoeff can be used to simulate a real engine firing – satellite can be rotated and vector thrust can be controlled by a rotation, in that case an impulse can be less that 100% - and PropCoeff with a corresponded value will be a good approximation for that firing. The first study for a flight trajectory was done – it uses Kepler’s elements for imaginary satellite and engine from http://www.canadianrocketry.org/motor_files/37148-O4900-BS-P.pdf. In all test’s runs was used the direction opposite of a vector velocity of the satellite. Debug mode of a compiler found to be useful for checking results. First tests targeted to find a firing time on an orbit to achieve a maximum possible apogee. TRA were set to try all points (engine’s firing time) with interval of 1 sec. It was found that the Sun’s position can do impact on the orbit. Then was confirmed that (approximately) a same mean anomaly (place on an orbit) gives a similar apogee (difference depends on rotation of the Earth around the Sun). Then were adjusted scale coefficient (PropCoeff 1.0~3.0) to see how it can impact the orbit. Next in test’s runs were discovered that proper firing of an engine can give a difference as a twice high apogee (i.e. 135,000 km compare to 70,000km), possibility that this is just is a bug in TRA application is not ruled out. Then it was found that experiments better to be done with satellite’s total weight variation then with impulse variation. At one test satellite actually reached one of Lagrange points. This case was investigated and it was found that actually difference in going to a Lagrange point or staying on high apogee Earth orbit is 0.02kg (20 grams) at total satellite weight. Was found proportion of an engine weight (~80kg), PropCoeff (3.0 == 51kg of a solid rocket propellant) to achieve distance 390,000 km (which is not good!). When experiments was finished (it was basically done to understand ballistics of a satellite orbiting the Earth) was made an attempt to see how to flyby the Moon. Was used previous step’s mean anomaly (position on an orbit) to achieve minimum distance to the Moon on apogee (may be apogee is not a good point?). As it was expected it shows that it will require some time (up to 30 days == one revolution of the Moon around the Earth) to achieve an optimized impulse.
(see more on http://www.adobri.com/ProjectTra.aspx)