5. The Rotation matrices

A rotation matrix for a newbie is a matrix which helps us determine a vector in a particular co-ordinates system when the vector or the co-ordinate frame is rotating. To add more clarity, let’s assume we have a particular point “P” in a 3-D space, represented by a vector \overrightarrow{P_{1}}  (say the three components of \overrightarrow{P_{1}} are P1X, P1Y, and P1Z) in a given co-ordinate frame C1. Further, if we rotate this co-ordinate frame by an angle \theta  about the Z axis and refer to this new orientation of C1 as C2, then the vector representation of P in C2 won’t be \overrightarrow{P_{1}}  but it will be given by the following equation

\left[ \begin{matrix} { { P }_{ 2X } } \\ { P }_{ 2Y } \\ { P }_{ 2Z } \end{matrix} \right] =\left[ \begin{matrix} cos\theta & -sin\theta & 0 \\ sin\theta & cos\theta & 0 \\ 0 & 0 & 1 \end{matrix} \right] *\left[ \begin{matrix} { P }_{ 1X } \\ { P }_{ 1X } \\ { P }_{ 1X } \end{matrix} \right]

Where P2X, P2Y, and P2Z are components of the vector \overrightarrow{P_{2}} , which is a vector representation of point P in the co-ordinate frame C2.

After this, say the co-ordinate frame C2 undergoes another rotation about the Y axis by an angle \phi and attains an orientation which is denoted by C3. So the vector representation of P in C3 will be given by

\left[ \begin{matrix} { { P }_{ 3X } } \\ { P }_{ 3Y } \\ { P }_{ 3Z } \end{matrix} \right] =\left[ \begin{matrix} cos\phi & 0 & sin\phi \\ 0 & 1 & 0 \\ -sin\phi & 0 & cos\phi \end{matrix} \right] *\left[ \begin{matrix} cos\theta & -sin\theta & 0 \\ sin\theta & cos\theta & 0 \\ 0 & 0 & 1 \end{matrix} \right] *\left[ \begin{matrix} { P }_{ 1X } \\ { P }_{ 2X } \\ { P }_{ 3X } \end{matrix} \right]

where P3X, P3Y, and P3Z are component of the vector \overrightarrow{P_{3}} , which is a vector representation of point P in the co-ordinate frame C3.

In case, we rotate the co-ordinate frame C 3 along X axis by an angle \psi , it will attain another orientation which we can refer by C4. The vector representation of P in C4 is given by

\left[ \begin{matrix} { P }_{ 4X } \\ { P }_{ 4Y } \\ { P }_{ 4Z } \end{matrix} \right] =\left[ \begin{matrix} 1 & 0 & 0 \\ 0 & cos\psi & -sin\psi \\ 0 & sin\psi & cos\psi \end{matrix} \right] *\left[ \begin{matrix} { P }_{ 3X } \\ { P }_{ 3Y } \\ { P }_{ 3Z } \end{matrix} \right]

Where P4X, P4Y, and P4Z are component of the vector \overrightarrow{P_{4}} , which is a vector representation of point P in the co-ordinate frame C4.

These rotation matrix can be used further in case the co-ordinate frame undergoes any more rotations. If the co-ordinate frame is rotated about the Z axis by an angle \alpha , then the vector in the newly oriented frame is obtained by multiplying the matrix \left[ \begin{matrix} cos\alpha & -sin\alpha & 0 \\ sin\alpha & cos\alpha & 0 \\ 0 & 0 & 1 \end{matrix} \right] with the vector observed in the co-ordinate frame before it got rotated about the Z axis by \alpha . In this case, the point P can be defined as a vector in the co-ordinate frame C5 (C4 after being rotated by angle \alpha about the Z axis) by the following equation

\left[ \begin{matrix} { P }_{ 5X } \\ { P }_{ 5Y } \\ { P }_{ 5Z } \end{matrix} \right] =\left[ \begin{matrix} cos\alpha & -sin\alpha & 0 \\ sin\alpha & cos\alpha & 0 \\ 0 & 0 & 1 \end{matrix} \right] *\left[ \begin{matrix} { P }_{ 4X } \\ { P }_{ 4Y } \\ { P }_{ 4Z } \end{matrix} \right]

Where P5X, P5Y, and P5Z are component of the vector \overrightarrow{P_{5}} , which is a vector representation of point P in the co-ordinate frame C5

Here, if C4 would have been rotated by an angle \alpha about Y axis instead of Z, then we would have multiplied with \left[ \begin{matrix} cos\alpha & 0 & sin\alpha \\ 0 & 1 & 0 \\ -sin\alpha & 0 & cos\alpha \end{matrix} \right] to \left[ \begin{matrix} { P }_{ 4X } \\ { P }_{ 4Y } \\ { P }_{ 4Z } \end{matrix} \right] get the vector in the newly oriented co-ordinate frame, on the other hand had it been rotated by X axis then we would have multiplied  \left[ \begin{matrix} 1 & 0 & 0 \\ 0 & cos\alpha & -sin\alpha \\ 0 & sin\alpha & cos\alpha \end{matrix} \right]  with \left[ \begin{matrix} { P }_{ 4X } \\ { P }_{ 4Y } \\ { P }_{ 4Z } \end{matrix} \right]  for the same purpose.

The rotation matrix we used in the above case are examples of passive rotation matrix, i.e., rotation matrix used to determine a vector when the frame is rotating.

When we need to make vector observation of a point which is changing its location, and if the co-ordinate frame of the observer is fixed in space, then we use active rotation matrix.

Imagine a vector \overrightarrow{V_{1}}  with vector components V1x , V1Y , and V1Z observed in a fixed coordinate frame C1. If this vector is rotated about the Z axis by an angle \theta , then the vector will change from \overrightarrow{V_{1}}   to \overrightarrow{V_{2}} , and the vector component of vector \overrightarrow{V_{2}}   in C1 co-ordinate frame is given by

\left[ \begin{matrix} { V }_{ 2X } \\ { V }_{ 2Y } \\ { V }_{ 2Z } \end{matrix} \right] =\left[ \begin{matrix} cos\theta & -sin\theta & 0 \\ sin\theta & cos\theta & 0 \\ 0 & 0 & 1 \end{matrix} \right] *\left[ \begin{matrix} { V }_{ 1X } \\ { V }_{ 1Y } \\ { V }_{ 1Z } \end{matrix} \right]

where V2X, V2Y and V2Z are components of vector \overrightarrow{V_{2}} after being rotated about the Z axis.  Further if vector \overrightarrow{V_{1}} after being rotated about the Z axis, is rotated about the Y axis by an angle \phi , then its new vector component V3X, V3Y and V3Z  are given by

\left[ \begin{matrix} { V }_{ 3X } \\ { V }_{ 3Y } \\ { V }_{ 3Z } \end{matrix} \right] =\left[ \begin{matrix} cos\phi & 0 & -sin\phi \\ 1 & 0 & 1 \\ sin\phi & 0 & cos\phi \end{matrix} \right] *\left[ \begin{matrix} cos\theta & -sin\theta & 0 \\ sin\theta & cos\theta & 0 \\ 0 & 0 & 1 \end{matrix} \right] *\left[ \begin{matrix} { V }_{ 1X } \\ { V }_{ 1Y } \\ { V }_{ 1Z } \end{matrix} \right]

Again, if this vector is given another rotation about X axis by an angle \psi , then the new vector components  V4x, V4Y, and V4Z can be obtained by

\left[ \begin{matrix} { V }_{ 4X } \\ { V }_{ 4Y } \\ { V }_{ 4Z } \end{matrix} \right] =\left[ \begin{matrix} 1 & 0 & 0 \\ 0 & cos\psi & -sin\psi \\ 0 & sin\psi & cos\psi \end{matrix} \right] *\left[ \begin{matrix} { V }_{ 3X } \\ { V }_{ 3Y } \\ { V }_{ 3Z } \end{matrix} \right]

These matrices can be further used to map a vector which is changing its orientation about any of the axis when observed from a fixed co-ordinate system. For example, if the above vector \overrightarrow{V_{4}} rotates about the Z axis by an angle \alpha , then the new components V5X, V5Y and V5Z can be obtained by

\left[ \begin{matrix} { V }_{ 5X } \\ { V }_{ 5Y } \\ { V }_{ 5Z } \end{matrix} \right] =\left[ \begin{matrix} cos\alpha & -sin\alpha & 0 \\ sin\alpha & cos\alpha & 0 \\ 0 & 0 & 1 \end{matrix} \right] *\left[ \begin{matrix} { V }_{ 4X } \\ { V }_{ 4Y } \\ { V }_{ 4Z } \end{matrix} \right]

On the other hand had \overrightarrow{V_{4}} been rotated by an angle \alpha , about the Y axis, then V5X, V5Y and V5Z would have been obtained by

\left[ \begin{matrix} { V }_{ 5X } \\ { V }_{ 5Y } \\ { V }_{ 5Z } \end{matrix} \right] =\left[ \begin{matrix} cos\alpha & 0 & -sin\alpha \\ 0 & 1 & 0 \\ sin\alpha & 0 & cos\alpha \end{matrix} \right] *\left[ \begin{matrix} { V }_{ 4X } \\ { V }_{ 4Y } \\ { V }_{ 4Z } \end{matrix} \right]

In the other case, had the rotation taken place along Z axis, by an angle \alpha , then the vector components V5X, V5Y and V5Z can be calculated by

\left[ \begin{matrix} { V }_{ 5X } \\ { V }_{ 5Y } \\ { V }_{ 5Z } \end{matrix} \right] =\left[ \begin{matrix} 1 & 0 & 0 \\ 0 & cos\alpha & sin\alpha \\ 0 & -sin\alpha & cos\alpha \end{matrix} \right] *\left[ \begin{matrix} { V }_{ 4X } \\ { V }_{ 4Y } \\ { V }_{ 4Z } \end{matrix} \right]

With this we complete our discussion on active and passive rotation matrices.  We can also observe that, the passive rotation matrices can be a way to define orientation of one co-ordinate frame with respect to the other.

With this, let’s get back to our Satellite Reference Frame. The passive rotation matrix to determine the orientation of orbit reference frame with respect to the ECI frame is given by

\left[ \begin{matrix} -sin(u)cos(\Omega )-cos(u)cos(i)sin(\Omega ) & -sin(u)sin(\Omega )+cos(u)cos(i)cos(\Omega ) & cos(u)sin(i) \\ -sin(i)sin(\Omega ) & sin(i)cos(\Omega ) & -cos(i) \\ -cos(u)cos(\Omega )+sin(u)cos(i)sin(\Omega ) & -cos(u)sin(\Omega )-sin(u)cos(i)cos(\Omega ) & -sin(u)sin(i) \end{matrix} \right]

Where u stands for the sum of argument of perigee and the true anomaly , i stands for the orbital inclination and \Omega  stands for the right ascension of the ascending node.

Since we know the orientation of the ECI frame, this relative orientation of orbit reference frame with respect to ECI frame can give us a great deal of information on how is the orbit reference frame oriented in space. So, in a similar way, if we can able to calculate the rotation matrix which can give us the orientation of the Satellite Body Frame with respect to the Satellite Reference Frame, with this we will not only know the orientation of Satellite Body Frame in space but we will also get to know how much is the satellite misaligned from its idea orientation. If we analyse these relative rotation parameters a bit more, we can calculate the rate at which satellite is deviating from its ideal orientation and this can help us to calculate the body rates of the satellite. With this we have answered the why’s, what‘s and how’s of the questions we raised in the beginning of the post.

4. How do we know where do we have to point ?

Hello!

In our previous post we discussed the need of the Attitude Determination and Control System (ADCS), shed some light on the different co-ordinate frames and skimmed through their applications.  With this we can dive further down into our study and understand ADCS in a better way.

As discussed earlier, ADCS is responsible for pointing the satellite in the desired direction in the space at any given time. However, it might not be possible for the satellite to attain hundred percent pointing accuracy, and it may happen that the satellite shakes a bit while trying to point towards a particular direction.  This shaky nature of the satellite can be quantified in terms of angular velocity, also known as the “body rates”. With this, the whole objective of ADCS boils down to pointing the satellite in a given direction with least possible body rates. Different satellites, depending on their payload requirements are tolerant to different body rates, and this tolerance determines the complexity of the ADCS on a satellite.

After reading all these, you must be wondering, how do we determine where the satellite has to point at any given time in space? How to calculate the pointing accuracy and how to measure the body rates of the satellite? The parameters, pointing accuracy and body rates can be measured with respect to some ideal values, so the next question which pops up is, how do we obtain those ideal values?

These why’s, what‘s and how’s will be answered in this and the upcoming posts.

So, let’s start with our first question, how do we determine where the satellite has to point at any given time in space? Different satellites need to point at different things while they are orbiting the earth. Some need to point at the earth all the time, while a few need to look at different heavenly bodies at different times, and a few of the satellites have to orient themselves along other satellites in space. The different axes of the Satellite Body Frame has to keep rotating in such a way that the satellite is able to point where it has to point. Hence, if we consider a hypothetical satellite whose body axis or the Satellite Body Frame keep on rotating with time in such a way, that the satellite is able to attain its desired orientation at any given time with hundered percent accuracy, then we will get something which we can call as the “reference satellite”. The Satellite Body Frame of this reference satellite is called the reference frame.

The reference frames for different satellites are different.  In a few cases where the satellite’s Z axis has to point continuously towards the earth (imagine a satellite which has to perform terrestrial imaging and has its camera coinciding with the Z axis) in such a way that the Y axis points towards the orbital angular momentum (Cross product of position vector and velocity vector of the satellite in ECI frame), then the reference frame of this satellite will be called the Orbit reference frame.

At any given time, the Satellite Reference Frame and Satellite Body Frame will be at the same location, i.e., the origin of the Satellite Body Frame will always coincide with that of the Satellite Reference Frame, however, the Satellite Reference Frame may not be aligned with Satellite Body Frame. The ADCS of a satellite works to align the Satellite Body Frame with the Satellite Reference Frame. To simplify our test cases and to understand ADCS in an easy way, we will be considering a satellite whose reference frame is the Orbit Reference Frame

With this, our first step in providing stability to a satellite is to trace the orientation and position of the reference satellite at any point in space.   Initially the position and velocity of the satellite can be acquired using an on-board GPS receiver or telemetry data containing the TLEs  (Something we will be discussing later on, for now you can read about this at https://en.wikipedia.org/wiki/Two-line_element_set), which can be sent to the satellite from the ground. Because of the power constraints on a satellite, it is not feasible to acquire position and velocity 24X7 using a GPS receiver or though telemetry data. This calls for the need of an on-board orbit propagator, which can predict the position and velocity of the satellite at any time instant, based on the initial data provided by the on-board GPS receiver or through the TLE containing telemetry data. Next, we find out the orientation of the reference frame at the satellite’s position.  Considering a satellite whose Satellite Reference Frame is the Orbit Reference Frame, we have a rotation matrix which gives us the orientation of the Orbit Reference Frame with respect to the ECI frame, something which we will be discussing in our next post.

 

 

2. The Different Frames and the Keplerian Elements

All Good Things Start With Frames

For a satellite to control its attitude, it needs to know its position along with that of some other stuff wandering in space. This can include a star, sun, moon or other heavenly bodies in our galaxy. Sensors on-board the satellites use these heavenly bodies as a reference to provide vector observations . Don’t worry if you could not get hold of the last line, things will be clearer as we progress through future posts. Apart from this, we can also determine the various physical disturbances acting on the satellite in space because of the different heavenly bodies, by accurately knowing their locations.

So, the obvious question which hits our curious mind is, how do we make vector observations of things which are zipping around the universe? To record a vector, we need a co-ordinate system, a frame of reference. Now, which frame should we be using to make some larger than life calculations?

As always, the key to simplify these vector calculations is to take the right frame of reference. And in this situation, where everything, from sun to moon, earth to satellite, is not at all stationary, we need to be extremely careful while selecting the frames and this selection will keep on varying with the observer and the subject.

Considering the non-inertial nature of satellites and the heavenly bodies, it will be a respite for us if we observe them from an inertial frame of reference. This stationary nature of observer can help us in easing out our calculations by getting rid of an additional velocity component that could have been imparted from the non-inertial frame.

Few of the important things which describe any co-ordinate system are, the fundamental plane(i.e. the X-Y place), the principal direction (i.e. the direction of X axis) and the direction of Z axis. Since the Z-axis must be perpendicular to the fundamental plane. The Y axis is chosen to form the right handed set of co-ordinate axes. Sounds a bit complicated? Don’t mind! You will develop a better idea about these things once we start reading specifically about the different frames.

Now, let’s talk about our first inertial frame of reference! The Earth Cantered Inertial (ECI) frame!

The Earth Cantered Inertial (ECI) frame

The Geocentric-Equatorial Coordinate System a.k.a. the Earth Centred Inertial frame has its origin right at the centre of the earth, however it is not fixed to the earth.

Although this frame has its origin at the centre of the earth, but it does not rotate with the earth.

this video can help you better visualize the ECI frame. The red lines in the video denote the three axes of the ECI frame. The fundamental plane contains the equator and the positive X-axis points in the vernal equinox direction. The Z-axis points in the direction of the geographical North Pole and the Y axis consequently completes the right hand set of co-ordinate axes. Here is a picture of this,

a1

Figure 1. The Earth Centred Inertial(ECI) frame

Any queries? Contact us. We will get back to you.

Now you might have been wondering, what do we mean by Vernal Equinox? Well, to make things a bit more interesting and to understand other inertial frames we need to throw some light on vernal equinox and something called the Keplerian Elements.

Vernal Equinox

You might have heard this strange term “vernal equinox” before, and in most cases we have been told that it is “the place in the sky where the sun rises on the first day of Spring”. This definition is not just vague and confusing, but it is terrible. Most of us don’t have any idea what is the first day of spring and why the sun should be in the same place in the sky on that date every year.

To get a better picture of vernal equinox, imagine the Sun’s orbit around the earth. Yes, right, the earth orbits the sun, but if we start observing from the earth, things will look the other way but the math is equally valid this way too (remember the concept of relative motion?). The plane formed by the hypothetical orbit of the sun around the earth is called the ecliptic. Similar to the ecliptic, we have the equatorial plane, the plane formed by the equator. Now, the sun will intersect the equatorial plane at two points in one orbit, one where the sun crosses the equator while ascending, the point where the sun pops out the equatorial plane and goes up, and the other one when the sun is descending, it punches the equator and dives down before coming up again. If you join these two points, you will get the nodal line for the sun’s orbit. Well, well, well, forgot to tell you something, the sun’s ascending node is called the Vernal Equinox!

Similar to the nodal line of the sun, we have nodal line for the satellites too. These lines are formed by the intersection of the plane formed by satellite’s orbit with the equatorial frame.

Since we have understood the concept of nodal line, let’s move forward and understand the Keplerian Elements.

Keplerian Elements, also referred as the orbital elements or simply the elements are a set of parameters used to define an elliptical orbit.

The first in the list is orbital eccentricity, sometimes referred as eccentricity

Eccentricity (e),

Eccentricity of an elliptical orbit lies between zero and one. In this domain of zero to one, higher the eccentricity, more elliptical is the orbit.

Semi major axis (a)

Imagine an ellipse traced by the satellite moving around the earth as its focus. The longest possible straight line in this ellipse is its major axis (2a, Figure 2), so half of this is the semi major axis. Cool?

Inclination (i),

The orbital inclination is the angle between the satellite’s orbital plane and the equatorial plane. Orbits having inclination lesser than 90 degrees are called prograde orbit and the orbits with inclination greater than 90 degrees are called retrograde orbit. Satellites in a prograde orbit rotate in the same direction as that of the earth, while those in retrograde, rotate in direction opposite to that of the earth.

Next comes the Right Ascension of Ascending node (RAAN)

 

Right Ascension of Ascending node (Ω)

Too weird a name, right? Never mind! This is the angle subtended between the nodal line of the sun (vernal equinox side) and that of the satellite (nodal line of the satellite from ascending node side).

Well, RAAN is not the only one with a horrible name, one another element with comparatively lesser weird name is the “Argument of Perigee”

Argument of Perigee

To visualise this we will have to dig a bit deeper into the satellite’s orbit. A satellite in an elliptical orbit will have the earth at one of its focus. So the satellite following the elliptical path around the earth will be closest to the earth at one point of time and farthest at the other. The point where the satellite and our planet maintains the minimum distance is known as perigee, on the other hand, the point where satellite is at the highest possible distance from the earth is called the apogee.

a2Figure 2. Elliptical Orbit

The vector in the direction from the earth’s centre to the perigee defines the eccentricity vector, whose magnitude equals the eccentricity of the orbit.   The angle subtended between the eccentricity vector and the nodal line of the satellite marks the “Argument of Perigee”.

True anomaly

Yet another orbital element with another weird name!

Consider a vector from earth’s origin to the satellite, let’s call this vector as the position vector. Now imagine an angle subtended by this vector and the eccentricity vector. This angle is what we call the “True anomaly “

Mean motion

Number of revolution the satellite completes per day. As simple as that!

Mean Anomaly

For a satellite moving in an elliptical orbit, if we draw a circle which passes through the apogee and perigee of the ellipse and has its centre at the centre of the ellipse, and consider a hypothetical satellite to be moving along that circle with angular velocity equal to the average angular velocity of the satellite moving in the elliptical orbit, then the angle subtended between the position vector of the satellite moving on the newly made circle and the eccentricity vector of the elliptical orbit is called the mean anomaly.

a3Figure 3. Orbital Elements

Now, since we are done with the orbital elements, we can peacefully move on and understand the other reference frames

Perifocal Frame

This is popularly known as the “natural frame” for an orbit. This frame is centred at the centre of the earth, the orbital plane is the fundamental plane (XY plane) itself, X axis is directed to the eccentricity vector, Z axis is in the direction of the satellite’s angular momentum which lies perpendicular to the orbital plane, and the Y axis completes the right hand set of co-ordinate axis.

a4Figure 4. Perifocal Frame

Let’s now move towards understanding the non-inertial frame. We previously said that an inertial frame helps us in simplifying the calculations, so why are we now bringing the non-inertial frames into picture? One way of answering this question is by looking at the Earth’s magnetism. The magnetic field produced by earth differs with the location. So if we have a frame which rotates with the earth, then it will be very convenient to locate different points on earth relative to this moving frame as all points on the earth will be stationary to this frame. This fact is crucial for a satellite having a magnetic field sensor (magnetometer) which needs to know its position with respect to different points of the earth. This non inertial frame is not just important for predicting magnetic field at different points in the orbit but it is also important for any satellite which is bothered about the different points on earth, be it for taking pictures of different parts of the world or for locating the different ground station on the home planet. And last but not the least, the non-inertial frame is the key for analysing the dynamics of a satellite!

Earth Centred Earth Fixed (ECEF) Frame

This frame keeps on rotating with the Earth. Centred at the equator, it has the equatorial plane as its fundamental plane (XY plane), the X axis can be traced by joining a line starting from the centre of the earth, to the point of intersection of the prime meridian and the equator, the Z axis points towards the geographical north pole and the Y axis completes the right hand set of co-ordinate axes.

The green axes in the video represent the ECEF frame.

a5Figure 5. Earth Centred Earth Fixed (ECEF) frame

Since the magnetic field produced by the satellite at different points in space in known, the ECEF frame can be used to locate the position of the satellite with respect to the moving earth, and consequently find the magnetic field at that point.

The frames we have talked about so far are the ones used to observe the satellite and other things in space from the earth, but then we cannot observe everything from our planet, to stabilise the satellite, we need to change our perspective, we need to look at the things from the satellite.

Orbit Reference Frame

The orbit reference frame has its origin at the centre of mass of the satellite, the Z axis points towards the centre of mass of the earth, the X axis which is perpendicular to the Z axis, and is in the direction of the velocity of the satellite, and the Y axis completes the right hand set of co-ordinate axis. In this article, the axes of orbit reference frame are denoted with a subscript R (XR, YR and ZR). Irrespective of the satellite’s orientation in the space, the orbit reference frame will always have its Z axis pointing towards the centre of the Earth, X axis pointing towards the velocity and the Y axis will be completing the right hand set of co-ordinate axes.

a6Figure 6. Inertial frame, Orbit Reference Frame and the Satellite Body frame

The inertial frame has been denoted by the axes set XI, YI, and ZI.

Satellite Body frame

Satellite body frame is fixed to the satellite’s body, with its origin at the centre of mass of the satellite. This frame is used to represent the actual satellite in space. The X, Y and Z axis need to be perpendicular to each other and should be popping out of the different faces of the satellite. An example of Satellite Body frame has been given in figure 7.

a7Figure 7. Satellite Body Frame

 

1. And So It Begins

Greetings,

All things require introduction and so we begin with ours,

Most of us are undergraduate students enrolled in a college located on a small coastal town of south India. We have been working on developing the Attitude Determination and Control Sub-System (ADCS) for our university funded nano-satellite.

It’s been the most wonderful experience, filled with all the emotions in any spectrum, having learned and gained a great deal we would like to share from our endeavors.

The main aim of the blog is to help novices to develop an understanding of the Attitude Determination and Control Sub-System as well as help model one if need be.

Explanations pertaining to the same will be as concise and elaborate as possible, we will also explain the system in the required flow with the required amount of detail. Pre-requisite knowledge is kept at a minimum.

And so we begin,

A nano-satellite, for the unacquainted is just a small satellite of low mass and size (For more information, click on this https://en.wikipedia.org/wiki/Miniaturized_satellite).

Since the work involved in building a nano-satellite is both widespread and (very) multi-disciplinary, the system has been divided into the following sub-systems (alphabetically listed, or priority wise, depending on your sub-system)

(1) Attitude Determination and Control Sub-system

This sub-system is primarily responsible for providing (either fine or coarse) pointing accuracy. What this means is being able to point the satellite in a certain direction with a certain degree of accuracy.

(2) Communication and Ground Station Sub-system

This sub-system is concerned with the transfer of data from the nano-satellite to the ground station and vice versa.

(3)Electrical Power Sub-system

Since the only source of power in space is solar energy, this sub-system is involved with the harnessing, storing and distribution of power.

(4) On board data handling Sub-system

Mostly anything and everything pertaining to the hardware architecture of the on-board computer comes under the umbrella of this sub-system.

(5) Payload

Payload is the end-goal or the purpose with which the satellite is sent into orbit to begin with.

(6) Structures, Thermals and Mechanical Sub-system

As the name suggest it is to do with the structure of the satellite, and is responsible for protection of the sub-system both during launch and from the space environment generally.

These are the basic sub-systems of our currently in progress nano-satellite, Parikshit (for more information, click on this à http://parikshit.org/). Variations to the above system exist considering no nano-satellites are built with exactly the same purpose. Let us begin with the obvious question,

Why do we need an ADCS sub-system to begin with?

It is both complex to understand and a mouthful to swallow. Well, the ADCS sub-system has been designed to provide both rate and attitude control, this requirement comes from the payload, power and communication sub-systems. The ADCS subsystem can also provide us with the position and velocity of our satellite while in orbit, this is useful for both payload and antenna pointing.

As to the question of how is this established? The answer to that question for now is that suitable hardware works in synchronization with complex algorithm. Both aspects should be made clear with the subsequent posts.