This article intends to be a summary of the attitude actuators for spacecrafts either active or passive. It will introduce the reader to the basic attitude control hardware in use for the modern platforms.
The attitude control of a spacecraft can be considered being either actively controlled (meaning that a controller calculates necessary control torques and acting on the satellite to adjust its attitude to a desired position) or passively controlled (meaning that the satellite uses external torques that occurs due to its interaction with the environment and thus they cannot be avoided, in this case the disturbances being used for forcing the attitude of the satellite).
In general, the active control assures 3 axes stabilization, while the passive control gives the opportunity for 1 axis stabilization. Some disadvantageous points of the passive actuators are that the pointing accuracy is pretty bad and also that the natural damping is very small meaning that additional energy dissipation devices need to be installed on board of the spacecraft.
A summary of the active actuators include:
The torques generated by the thrusters is considered as external torques since the angular momentum of the entire satellite changes. The accuracy of the attitude control depends on the minimum impulse of the type of thruster used.
Depending on the size of the satellite and taking in consideration the complexity of this solution, different types of thrusters are normally being used: gas jets, ion jets or even nuclear propulsion.
Gas jets produce thrust by a collective acceleration of propellant molecules, with the energy coming from either a chemical reaction or thermodynamic expansion.
Gas jets are classified as hot gas when the energy is derived from a chemical reaction or cold gas when it is derived from the latent heat of a phase change, or from the work of compression if no phase change is involved. Hot gas jets generally produce a higher thrust level (>5 N) and a greater total impulse or time integral of the force. Cold gas systems operate more consistently, particularly when the system is operated in a pulsed mode, because there is no chemical reaction which must reach steady state. The lower thrust levels (<1N) of cold gas systems may facilitate more precise control than would be available with a high thrust system.
Hot gas systems may be either bipropellant or monopropellant. Fuel and oxidizer are stored separately in a bipropellant system; very high thrust levels (>500 N) can be obtained, but the complexity of a two component system is justified only when these thrust levels are required. Monopropellant systems use a catalyst or less frequently, high temperature to promote decomposition of a single component, which is commonly hydrazine (N2H4) or hydrogen peroxide (H2O2). Hydrazine with catalytic decomposition is the most frequently used hot gas monopropellant system on spacecraft. The problem of consistency manifests in two ways. First, the thrust is bellow nominal for the initial few seconds of firing because the reaction rate is bellow the steady state value until the catalyst bed reaches operating temperature. Second, the thrust profile, or time dependence of thrust, changes as a function of total thruster firing time; this is significant when a long series of short pulses is executed, because the thrust profile for the later pulses will differ from that for the earlier pulses.
The propellant supply required for jets is the major limitation on their use; a fuel budget is an important part of mission planning for any system using jets. Other considerations are the overall weight of the system and the need to position thrusters where the exhaust will not impinge on the spacecraft. The latter consideration is especially important when hydrazine is used, because the exhaust contains ammonia, which is corrosive.
In more distant orbits, jets are the only practical means of interchanging momentum with the environment. High thrust or total impulse requirements may indicate a hot gas system. Otherwise the cold gas system may be favored because hydrazine freezes at about 0 deg C and may require heaters if lower temperatures will be encountered during the mission. Specific components may affect the relative system reliability; for example hydrazine systems use tank diaphragms to separate the propellant from the pressurizing agent and also require a catalyst or heater to initiate decomposition; cold gas systems may have a pressure regulator between the tank and the thruster.
Ion jets accelerate individual ionized molecules electro-dynamically, with the energy ultimately coming from solar cells or self containing electric generators.
The first successful orbital test of the ion engine was made in 1968, and the engine was used to keep a satellite in geo-synchronous orbit. The test demonstrated that ion propulsion was possible in space, and that the electrical system did not interfere with any of the other systems on the spacecraft.
Ion jets are primarily used in applications that do not require large amounts of thrust, such as satellite control. The thrust produced by a typical ion engine is on the order of millinewtons, and thus cannot yet be used as a primary propulsion system for launching any spacecraft from the earth’s surface. However, the spacecraft utilizing the ion jet engine can be launched via chemical rocket, and then sufficient thrust can be developed in the frictionless atmosphere of space.
While chemical rocket engines produce thrust by burning large amounts of fuel and expelling hot gases, ion jets do so by expelling highly accelerated ions of inert gases. This greatly reduces the amount of required propellant, and removes limitations found in conventional rocket engines caused by the high thermal stresses of combustion.
- Angular momentum storage and exchange devices
These devices consist of a spinning wheel, either the orientation or the spin rate being changed. In the absence of external torques, the angular momentum of the entire satellite does not change, but whenever the rotation rate or the rotation axis of such a device is changed, the satellite will experience a torque in such a way that the angular momentum of the entire satellite system is constant.
On the same manner these devices can be used to compensate the external disturbance torques (which can cause a change in the angular momentum of the satellite) by producing their own internal torques.
The main advantage of these devices is that they assure very high accuracies, and their main disadvantage is that they have to be de-saturated using an actuator that generates external torques in the case when a maximum rotation rate is reached. One should also consider that such devices have big power consumptions and masses.
There are three types of such devices: momentum wheels, reaction wheels and fly wheels.
Most of the time magnetorquers are used as coils but more generally any conducting device can be used to perform this function. They are often used in combination with angular momentum storage and exchange devices, or they are used to de-saturate the wheels without using propellant.
The physical principle consists in generating a magnetic dipole moment in a desired direction, moment which will interact with the Earth’s magnetic field and thus external torques appear. However the use of such a device is limited to the low Earth orbits where the Earth’s magnetic field strength has usable values and should take in consideration that the generation of torques can be done just for the one perpendicular to the magnetic field vector. Considering this cosine dependency, should be very easily intuited the main disadvantage of this method- named that the absolute torque produced is very small.
A summary of the passive actuators include:
- Gravity gradient
This method is based on the fact that the gravitational force decreases with the square of the distance. By building the satellite in a certain configuration, parts of the satellites which are closer from the Earth’s center are subject to a bigger force than the ones which are further away and in consequence this effect can be used to produce torques that adjust the satellite to a certain position.
- Magnetic dipole
This method depends highly on the strength of the magnetic field in a specific point of spacecraft’s orbit and this is the reason why it can be used only for the low Earth orbits where the values of the magnetic field strength are big enough. The method is based on installation on board of the satellite of a strong constant magnetic dipole –or a permanent magnet, which will interact with the Earth’s magnetic field vector causing an adjustment of the two axes.
- Aerodynamic stabilization
The aerodynamics of the satellite (influenced by the geometry of the shape) can be used in flight to adjust the position with respect to the flight direction (based on the forces caused by the atmospheric drag).
However this method can only apply where the atmospheric drag is big enough (i.e. a low Earth orbit) and only for a short period of time, as this will lower the altitude and will shorten the decay period.
- Solar radiation
They can be used to generate torque on the spacecraft based on the solar radiation pressure created by the Sun. The orientation is then an axis that points to the Sun.