Mission Concept for a Satellite Mission to Test Special Relativity

Kurzfassung

In 1913 Albert Einstein developed the theory of Special Relativity. This theory describes the relation between space and time and revolutionized the understanding of the universe. While the concept is generally accepted new experimental setups are constantly being developed to challenge the theory, but so far no contradictions have been found. One of the postulates Einsteins theory of Relativity is based on states that the speed of light in vacuum is the highest possible velocity. Furthermore, it is demanded that the speed of light is independent of any chosen frame of reference. If an experiment would find a contradiction of these demands, the theory as such would have to be revised. To challenge the constancy of the speed of light the socalled Kennedy Thorndike experiment has been developed. A possible setup to conduct a Kennedy Thorndike experiment consists of comparing two independent clocks. Likewise experiments have been executed in laboratory environments. Within the scope of this work, the orbital requirements for the first space-based Kennedy Thorndike experiment called BOOST will be investigated. BOOST consists of an iodine clock, which serves as a time reference, and an optical cavity, which serves as a length reference. The mechanisms of the two clocks are different and can therefore be employed to investigate possible deviations in the speed of light. While similar experiments have been performed on Earth, space offers many advantages for the setup. First, one orbit takes roughly 90 min for a satellite based experiment. In comparison with the 24 h duration on Earth it is obvious that a space-based experiment offers higher statistics. Additionally the optical clock stability has to be kept for shorter periods, increasing the sensitivity. Third, the velocity of the experimental setup is larger. This results in an increased experiment accuracy since any deviation in the speed of light would increase with increasing experiment velocity. A satellite planted in a Low Earth Orbit (LEO) travels with a velocity of roughly 7km=s. Establishing an Earth-bound experiment that travels with a constant velocity of that order is impossible. Finally, space offers a very quiet environment where no disturbances, such as vibrations, act upon the experiment, which is practically unavoidable in a laboratory environment. This thesis is contrived of two main chapters. The chapter titled "Mission Level" exploits orbital candidates. Here, possible orbits are explained in detail and the associated advantages and problems are investigated. It also contains a discussion about ground visibility and downlink feasibility for each option. Finally, a nominal mission scenario is sketched. The other chapter is called "Sub- Systems". Within this chapter the subsystems of the spacecraft are examined. To examine the possible orbits it is necessary to define criteria according to which the quality of the orbits can be determined. The first criteria reflects upon the scientific outcome of the mission. This is mainly governed by the achievable velocity and the orbital geometry. The second criteria discriminates according to the mission costs. These include the launch, orbital injection, de-orbiting, satellite development, and orbital maintenance. The final criteria defines the requirements in terms of mission feasibility and risks, e.g. radiation. The criteria definition is followed by explaining the mission objectives and requirements. Each requirement is then discussed in terms of feasibility. The most important parameters, such as altitude, inclination, and the right ascension of the ascending node (RAAN), are discussed for each orbital option and an optimal range is picked. The optimal altitude depends on several factors, such as the decay rate, radiation concerns, experimental contributions, and eclipse duration. For the presented mission an altitude of 600 km seems to be the best fit. Alongside the optimal altitude possible de-orbiting scenarios are investigated. It is concluded that de-orbiting of the satellite is possible without any further external influence. Thus, no additional thrusters are required to de-orbit the satellite. The de-orbiting scenario has been simulated with systems tool kit (STK). From the simulation it can be concluded, that the satellite can be deorbited within 25 years. This estimation meets the requirements set for the mission. Another very important parameter is the accumulative eclipse duration per year for a given orbit. For this calculation it is necessary to know the relative positions and motion of the Earth and the Sun. From this the eclipse duration per orbit for different altitudes is gained. Ground visibilities for orbital options are examined for two possible ground stations. The theory is based on the geometrical relation between the satellite and the ground stations. The results are also shown an agreement with the related STK simulations. Finally, the both ground stations are found adequate to maintain the necessary contact between the satellite and the ground station. In the trade-off section, orbit candidates are examined with a more detail. Results from the previous sections with some additional issues such as the experiment sensitivities, radiation concern and thermal stability are discussed to conclude which candidate is the best for the mission. As a result of the trade-off, two scenarios are explained in the "Nominal Mission Scenario" section which covers a baseline scenario and a secondary scenario. After selecting a baseline orbit, two sub-systems of the satellite are examined. In the section of "Attitude Control System (ACS)" where the question of "Which attitude control method is more suitable for the mission?" is tried to be answered. A trade-off among two common control methods those are 3-axes stabilization and spin stabilization is made. For making the trade-off possible external disturbances in space are estimated for two imaginary satellite bodies. Then, it is concluded that by a spin stabilization method maintaining the attitude is not feasible. Thus, the ACS should be built on the method of 3-axes stabilization. As the second sub-system the possible power system of the satellite is examined. The total size and the weight of the solar arrays are estimated for two different power loads. Then, the battery capacity which will be sufficient for the power system budget is estimated together with the total mass of the batteries. In the last section, a conclusion of the thesis work is made and the possible future works for the BOOST mission are stated.