This question appears in Mumbai University > Satellite Communication and Network subject

Marks: 10 M

Year: Dec 2013, May 2013

Satellite Structural Design

• The cost of launching a satellite is a function of its mass. As a result of this, the cost of launching one is very high, more so in the case of as geostationary satellite. One of the most basic requirements is therefore lightness of its mechanical structure.
• All efforts are therefore made to reduce the structural mass of the satellite to the minimum. This is achieved by using materials that are light yet strong. Some of the materials used in the structure include aluminium alloys, magnesium, titanium, beryllium, Kevlar fibres and more commonly the composite materials.
• The design of the structural subsystem relies heavily on the results of a large number of computer simulations where the structural design is subjected to stresses and strains similar to those likely to be encountered by the satellite during the mission.
• The structural subsystem design should be such that it can withstand mechanical accelerations and vibrations, which are particularly severe during the launch phase. Therefore the material should be such that it can dampen vibrations. Kevlar has these properties.
• The satellite structure is subjected to thermal cycles throughout its lifetime. It is subjected to large differences in temperature as the sun is periodically eclipsed by earth. The temperatures are typically several hundred degrees Celsius in the side facing the sun and several tens of degrees below zero degrees Celsius on the shaded side. Designers keep this in mind while choosing material for the structural subsystem.
• The space environment generates many other potentially dangerous effects. The satellite must be protected from collision with micrometeorites, space junk and charged particles floating in space. The material used to cover the outside of a satellite should also be resistant to puncture by these fast travelling particles.
• The structural subsystem also plays an important role in ensuring reliable operation in space of certain processes such as separation of the satellite from the launcher, deployment and orientation of solar panels, precise pointing of satellite antenna, operation of rotating parts and so on.

Satellite Structural Design

Spacecraft mass and power estimate:

• During the early stages of as satellite communication system simplified spacecraft mass and power estimation techniques are useful in assessing the sensitivity of these parameters to changes in system parameters such as the satellite EIRP, eclipse operation and redundancy configurations.
• Here we have considered Pritchard model which provides adequate accuracy for early system planning. The model uses broad operational requirements together with the prevailing technological factors to estimate mass and power of a communication satellite.
• The required RF transmitter power, number of transponders, eclipse service capability and satellite lifetime are used as the initial input to the model.
• These parameters are a function of the desired traffic capacity, channel quality objectives, coverage area and space segment cost.
• Power estimate: The RF transmitter power levels together with the number of transponders are used to estimate the payload DC power through simple relationships.
• The total primary DC power Pt for a satellite transmitter is given by

• Other parameters such as satellite lifetime, eclipse operation conditions and housekeeping requirements are then introduced to estimate the total primary power and array size.
• Mass Estimate: The mass estimates of the various sub-systems are based on the state of the prevailing technology and the use of empirical relationships obtained from available database on communication satellites.
• The mass of the primary power system is given by

• The payload mass (MPI) consist of the sum of the transponder and antenna masses.

$where R_u= redundancy in the i^{th} transmitterr$

$b=factor to account for mass of receivers, switches, up/down converters, filters and other$

$components (b=1.1-1.5 depending upon complexity)$

• Mass of the platform $M_f$: It is related to the dry mass

where MD is the dry mass (mass of the satellite without propellant) given as:

• Wet mass: Wet mass of spacecraft is the sum of the dry mass and the propellants. For the beginning of life mass it is given as:

• Mass in the transfer orbit is obtained as

where

$I= value of specific impulse of fuel$ used by the apogee or perigee kick motors

$∆v=velocity$ required for injection into transfer orbit to circularize the orbit and to correct the inclination. Transfer orbit weights generally range between 1.7 and 2.3 times the in-orbit weight of a communication satellite.

• Cost estimate: The total space segment cost (launch plus payload) for a single satellite in a circular orbit upto geostationary orbit is given by