The space radiation environment, driven by solar activity and galactic cosmic rays, provides a powerful and precise treatment to satellites in space near Earth. Understanding the atmospheric and ionospheric dynamic responses to solar powered particles, radiation and their air effects in space is critical and practical for satellite design and operation. The specific effects of the radiation environment as a satellite depend on the source, type and energy of the incident particle, as well as the satellite’s orbit or its location during solar energy events. Radiation reduction measures can increase the margin of safety between the part design boundaries and the applied stresses due to the impact of the particles and consequently increase the protection of the part. However, it is important to pay more attention to the solar maximum phase in all mitigation efforts because the rate of impact is higher in this range.
During peak solar storms, severe solar storms can cause a very short-term increase in radiation levels and high levels of single-event effects that existing mitigation measures may not be able to handle. In addition, as the dependence on satellite services increases, the economic and social risk associated with space weather increases and the potential impact may be unprecedented. In view of this, contingency plans have been proposed, including the possibility of switching to or utilizing other standalone satellite services. Upcoming multi-constellation GNSS receivers can play an important role in this context, such that individual GNSS receivers are naturally resistant to a satellite denial of service.
The increase in atmospheric drift on satellites due to space air and the resulting acceleration of orbital distortion is also an important disturbing force to be taken into account for low Earth orbit satellites. There is a brief review of the impact and mitigation of this phenomenon. It should be kept in mind that this review is concise and concise when compared to large-scale studies on the space radiation effects on satellites and their abatement methods.
Stopping Force, Particle Spacing, and Accumulated Dose in Spacecraft Materials
Particles, electrons and proton flows of various energies were analyzed from the NOAA database for 3 months (April – June 2010). The mass stopping power, range and possible accumulated proton dose are calculated and applied to the phenomenon of possible interaction of particles with the satellite surface and electrical, electronic and electrochemical components.
Stopping power is the average energy loss per unit length of a particle as it passes through the material. It is known that charged particles ionize the atom or molecule they encounter while passing through matter and lose energy in the process. The stopping power depends on the type of particle, its energy and the properties of the material it passes through. While the numerical values and units are the same for both quantities, the stopping power expresses the property of the material, while the energy loss per unit path length describes what happens to the particle. The ionization intensity along the particle path is proportional to the stopping power of the material because the production of an ion pair requires a fixed amount of energy. There is the Bethe-Bloch formula for stopping power derived from relative quantum mechanics, and it is as follows:
S= -d Ed x=4 πz2k20e4mv2[ ln2 metrev2ben- l n ( 1 -v2c2) -v2c2]S=-dEdx=4πz2k02e4mv2ln2mv2ben-ln1-v2c2-v2c2
Here z is the atomic number of the heavy particle, e is the magnitude of the electron charge, m is the electron rest mass, c is the speed of light, I is the average excitation energy of the medium, v is the velocity of the particle, and k is the Boltzmann constant (= 8.99 × 10 9 N m 2 C 2). The mass stopping power of the material is obtained by dividing the stopping power by the density of the material. This refers to the rate of charged particle energy loss per g / cm, since a useful quantity 2 is the medium through which it passes.
S= -d E d xS=-dEdx
The range R of a particle, for example the proton, its initial kinetic energy Ek and mass m is the average distance it travels before reaching a stop. R depends on the type of particle, its initial energy and the material through which it passes. A theoretical approach to determining the range of charged particles uses the term stopping power. For spacing calculations for the proton, Emin was taken as 1 mega-electron volt because there is a lot of data available at 1 mega-electron volt. R (Emax) is the range measured at maximum energy Emax. The previous work used the proposed experimental relationships to calculate the mass stopping power of particles in spacecraft materials. However, there are limitations to the equations as they were originally formulated for low energy particles. Values obtained using Bethe equations are higher and assumed to be more accurate at higher particle energies.
Dose Accumulation and Absorption
The total ionizing dose disclosed can be measured in terms of the absorbed dose and is a measure of the energy absorbed by that substance. Absorbed dose is measured using a unit called rad (radiation absorbed dose) or SI unit which is gray (Gy); 1 Gy = 100 cycles = 1 J / kg. The total dose accumulated in a satellite depends on orbit height, orientation, and time spent in orbit. To calculate the total ionizer dose, one must know the integrated particle energy spectrum, ø (E), or flux as a function of particle energy, and dose is a function of particle flow. Spacecraft become important as they spend more time in the space radiation environment. The stopping power is used to determine the dose from the charged particle with the relationship D = d E d xD = dEdx. Satellite and space probes typically encounter a total ionizing dose of between 10 krad (100 Gy) and 100 krad (Si) (1000 Gy (Si)). The time it takes for a satellite component to fail due to the total ionizing dose, t (in years), can be obtained by dividing the maximum absorbed dose or the total ionizing dose threshold by the total absorbed dose per year.
Theoretical calculations were made to estimate the average failure time of a model satellite due to the total ionizing dose. The assumption is that the body of the model satellite, the electrical, electronic and electrochemical components mainly contain materials of silicon (Si) and germanium (Ge), while calculations are based on E 78 mega electron volt particles. When particles of this energy range bombard and enter the satellite, some of their energy is lost due to the stopping power of the alloy, but the reminder generates a significant dose for the components. With continuous exposure, the dose continues to increase over time until the threshold is exceeded, resulting in complete failure of the affected satellite. Calculations have shown that the component of the model satellite can generate a dose of 10 krads in 3 years and 100 krats in 29 years.
Author: Ozlem Guvenc Agaoglu