I. Literature Review




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НазваниеI. Literature Review
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II. Rationale


The goal of this design project was to develop an integrated CubeSat deorbit device that when deployed could cause CubeSats to deorbit from a 900 km altitude in less than 25 years. Designs for deorbit mechanisms range from long tethers to inflatable assemblies. The current level of deorbit mechanism research provides many new areas to explore.


The method chosen to best meet the requirements for a CubeSat deorbit device is one that increases surface area, producing drag. This design relies on creating a drag force to slow the orbital velocity of the spacecraft (3). It can be shown that as the orbital velocity decreases, the orbital altitude must also decrease. The drag force is a function of the atmospheric density at the specific altitude. Atmospheric density varies between night and day, and is affected by solar activity. However, in general, atmospheric density decreases exponentially from the earth’s surface outward; therefore drag is higher at lower altitudes.


In spacecraft design, ballistic coefficient can be used to describe the drag efficiency. Ballistic coefficient is defined as where M is the mass of the orbiting object, CD is the drag coefficient, and A is the frontal area (6). In this case, a lower ballistic coefficient will have a larger surface area and be a better deorbit mechanism. A 1U CubeSat, with a mass of 1 kg, a cross sectional area of 0.01 m2, and a CD of 2, will have a ballistic coefficient of 50.


The method for increasing the frontal surface area of the spacecraft requires some sort of deployable mechanism. The two main ways to increase the surface area of a spacecraft are to deploy a rigid structure/array or an inflated device. Both of these methods require some sort of activation technique and movement.


The deployable array technique would be similar to deploying additional solar panels. The sides of the spacecraft could fold outward to increase the total surface area. This technique required a significant level of sophistication and power, and was discarded.


The second method for increasing surface area was to inflate a generic volume. From a drag perspective, the shape was irrelevant as long as the frontal surface area for producing the required spacecraft lifetime was achieved. An inflatable must use some sort of inflation gas. The thin-walled inflatable membrane and associated inflation gas and hardware can only occupy a small volume in order to provide sufficient volume for the actual payload. Consequently, the deorbit system, including the inflation gas and activation system must be capable of being packed into a very small volume. Material selection for the thin walled inflatable was also of importance since it must remain in space for approximately 10 years. Puncture resistance was a consideration, regardless of whether micrometeoroid impacts are a concern or not (6).


A possible modification to the inflatable system was considered utilizing a conductive ring inside the membrane material. The ring would allow a current to flow, which would interact with the Earth’s magnetic field. This interaction could be beneficial if the conductive ring was perpendicular to the Earth’s magnetic field lines, since then the ring would experience a torque, which could be used to orient the CubeSat. Alternatively, if the ring was oriented parallel to the magnetic field lines, the application of a suitable DC voltage could produce a current (clockwise versus counterclockwise), and the interaction could cause the ring to expand. This expansion could be useful. For example, if the inflatable device experienced a leak, the electromagnetic expansion could help keep the material from collapsing on itself, making it easier for the gas to inflate the structure.


Although a conductive ring was considered useful, the added complexity was a concern. One would need to ensure that dissipated heat from the ring did not degrade the adhesive in the inflatable structure. The folding and packaging of the material would need to be slightly altered. Also, the CubeSat would need to store more post-mission power, in order to supply the current to the ring. Lastly, and most importantly, a detailed numerical model would be necessary to predict the magnetic field of the Earth at the exact position of the CubeSat. The purpose of this computer program would be to ensure that the magnetic forces were directed appropriately ( according to the right-hand rule). Due to these complexities, a conductive ring was eliminated from further consideration.
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