Adcs orbit intro

Chapter1 ALEXSAT

CHAPTER (1)

1.1 Satellite Orbits

After a satellite is separated from launching vehicle, it moves in a path

around the Earth called an orbit. Satellite orbiting Earth due to the balance

between two forces, gravitational force which attracts the satellite towards the

Earth and centrifugal force (due to linear velocity of the satellite in orbit )

which causes repulsion of the satellite out from Earth,see Figure ( 1-1.) During

satellite mission design, the orbit is chosen which is appropriate to its mission.

So, a satellite that is in a very high orbit will not be able to see objects on Earth

as many details as orbits that are lower, and closer to the Earth's surface.

Similarly, the satellite velocity in orbit, the areas observed by the satellite, and

the frequency with which the satellite passes over the same portions of the

Earth are all important factors in satellite orbit selection. Essentially, there are

six orbital parameter called classical Keplerian orbital elements define the orbit

as shown in Figure ( 1-3).

Figure ( 1-1) Gravitational force and the centrifugal force acting on bodies

orbiting Earth

1. Semi-major axis. a This is a geometrical parameter of the elliptical

orbit. It can, however, be computed from known values of apogee and

perigee distances as for definition of apogee and perigee see Figure

( 1-2).

By Ahmad Farrag 2011 [email protected]

Chapter1 ALEXSAT

( 1.1) 2

perigeeapogeea +=

2. Eccentricity. e The orbit eccentricity is the ratio of the distance between

the centre of the ellipse and its focus to the semi-major axis of the

ellipse see Figure ( 1-2).

3. Right ascension of the ascending node Ω . it tells about the orientation of

the line of nodes, which is the line joining the ascending and descending

-nodes, with respect to the direction of the vernal equinox See Figure

( 1-3).

Vernal equinox is the line that intersects the Earth's equatorial plane and

the Earth's orbital plane, which passes through the centre of the Earth with

respect to the direction of the sun on 21 MarchError! Reference source not

found..

(a) (b)

4. Inclination i . is the angle that the normal to the orbital plane of the

satellite makes with the normal to the equatorial plane , Figure ( 1-4).

5. Argument of the perigee W. This parameter defines the location of the

major axis of the satellite orbit. It is measured as the angle ω between

the line joining the perigee and the focus of the ellipse and the line of

nodes in the same direction as that of the satellite orbit, see Figure ( 1-4).


Chapter1 ALEXSAT

Figure ( 1-2) apogee ,perigee of the orbit and semi-major axis

Figure ( 1-3) Right ascension of the ascending node

6. True anomaly of the satellite fo. This parameter is used to indicate the

position of the satellite in its orbit. It is defined as the angle, between the

line joining the perigee and the centre of the Earth with the line joining

the satellite and the centre of the Earth, see Figure ( 1-4)

Orbits can be classified according to different criteria, such as

1. According to orbit Altitude

o Low Earth Orbit (LEO): orbit altitude ranging in altitude from 200–1000 km

o Medium Earth Orbit (MEO): orbit altitude ranging from 1000 km to just below geosynchronous orbit at 35786 km.

o High Earth Orbit (HEO): orbit altitude above 35786 km.


Chapter1 ALEXSAT

Figure ( 1-4) Keplerian orbital elements

2. according to inclination

o Equatorial orbit : an orbit that co-planed with the equator i.e. orbit with zero inclination

o Polar orbit: An orbit that passes above or nearly above both poles of the Earth on each revolution. Therefore it has an inclination of about 90 degrees

o Inclined orbit: An orbit whose inclination between 0 and 90 degrees.

3. according to Eccentricity o Circular orbit: An orbit that has an eccentricity of 0 and whose

path traces a circle o Elliptic orbit: An orbit with an eccentricity greater than 0 and less

than 1 whose orbit traces the path of an ellipse


Chapter1 ALEXSAT

1.1.1 Special Orbits

An important consideration in space mission design is determining the

type of Earth Orbit that best suits the design goals and purpose of the mission.

A brief description for the special orbits which frequently used such as; low

Earth orbit, medium Earth orbit, geostationary orbit, polar orbit, Sun-

synchronous orbit and Molniya orbit, is presented.

1.1.1.1 Low Earth Orbit (LEO)

Orbiting the Earth at roughly 200-1000 Km altitude: Almost 90 percent

of all satellites in orbit are in LEO. LEO is often utilized because of the low

launch requirements that are needed to place a satellite into orbit. LEO

satellites orbit the Earth in roughly 90 minute periods. This means that they are

fast moving, and sophisticated ground equipment must be used to track the

satellite, LEO is used for such missions as flight tests, Earth observations,

astronomical observations, space stations and scientific.

Figure ( 1-5) LEO, MEO and GEO

1.1.1.2 Medium Earth Orbit (MEO)

MEO sometimes called Intermediate Circular Orbit (ICO), is the region

of space around the Earth above low Earth orbit (1,000 kilometers) and below

geostationary orbit (35,786 Km).The most common use for satellites in this

region is for navigation, such as the GPS (20,200 Km) and Galileo


http://en.wikipedia.org/wiki/Low_Earth_Orbit

http://en.wikipedia.org/wiki/Geostationary_orbit

http://en.wikipedia.org/wiki/Satellite

http://en.wikipedia.org/wiki/Global_Positioning_System

http://en.wikipedia.org/wiki/Galileo_positioning_system

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(23,222 Km) constellations. Communications satellites that cover the North and

South Pole are also put in MEO. The orbital periods of MEO satellites range

from about 2 o 12 hours. Telstar, one of the first and most famous experimental

satellites, orbited in

1.1.1.3 Geostationary/Geosynchronous Earth Orbit (GEO)

Satellite in geostationary orbit appears to remain in the same spot in the

sky all the time. Really, it is simply traveling at exactly the same speed as the

Earth is rotating below it, but it looks like it is staying still regardless of the

direction in which it travels, east or west. A satellite in geostationary orbit is

very high up, at 35,850 km above the Earth. Geostationary orbits, therefore, are

also known as high orbits; GEO is used for communications satellite

Figure ( 1-6) GEO satellites appear stationary with respect to a point on Earth

1.1.1.4 Polar Earth Orbit

For full global coverage of the Earth, a ground track would have to

cover latitudes up to ± 90o. The only orbit that satisfies this condition has an

inclination of 90°. These types of orbits are referred to as polar orbits. Polar

orbits are used extensively for the purpose of global observations.

1.1.1.5 Sun Synchronous Orbits (SSO)

A Sun-synchronous orbit (SSO) is a nearly polar orbit where the

ascending node precesses at 360 degrees per year or 0.9856 degrees per day.


http://en.wikipedia.org/wiki/Orbital_period

http://en.wikipedia.org/wiki/Telstar

Chapter1 ALEXSAT

Figure ( 1-7) Sun synchronous orbit

1.1.1.6 Molniya Orbit

Highly eccentric, inclined and elliptical orbits are used to cover higher

latitudes, which are otherwise not covered by geostationary orbits. A practical

example of this type of orbit is the Molniya orbit. It is a widely used satellite

orbit, used by Russia and other countries of the former Soviet Union to provide

communication services. Typical eccentricity and orbit inclination figures for

the Molniya orbit are 0.75 and 65° respectively. The apogee and perigee points

are about 40000 km and 400 km respectively from the surface of the Earth. It

has a 12-hour orbit and a satellite in this orbit remains near apogee for

approximately 11 hours per orbit before diving down to a low-level perigee.

Usually, three satellites at different phases of the same Molniya orbit are

capable of providing an uninterrupted service.


Chapter1 ALEXSAT

Figure ( 1-8) Molniya orbit

1.2 Reference Coordinate Systems

Several different reference coordinate systems or reference frames are

used to describe the attitude of a satellite in orbit. The most utilized coordinate

systems employed in attitude control problem are the inertial, Greenwich,

orbital, body, and device frames.

1.2.1 Geocentric Inertial Coordinate System

The Geocentric Inertial Coordinate System or Earth-Centered Inertial

(ECI)coordinate system has its origin in the Earth center The -axis points is

the axis of rotation of Earth. The -axis is in the direction of the vernal

equinox, and the -axis completes the right-hand rule for the coordinate

system. A demonstration for the geocentric inertial coordinate system is shown

in Figure ( 1-9).

IZ

IX

IY

1.2.2 Greenwich Coordinate System

The Greenwich Coordinate System or Earth-centered Earth-fixed

reference frame also has its origin at the center of the Earth, but it rotates

relative to inertial space, shown in Figure ( 1-10) The -axis direction is the GZ


Chapter1 ALEXSAT

GX

GY

Figure ( 1-9) Inertial coordinate system


Chapter1 ALEXSAT

Figure ( 1-10) Greenwich coordinate system

1.2.3 Orbital Coordinate System

The orbital coordinate system (OCS) is located at the mass center of the

satellite. This frame is non inertial because of orbital acceleration and the

rotation of the frame.

The motion of the frame depends on the orbit altitude. The -axis in the

direction from the satellite to the Earth , -axis in the direction opposite to

the orbit normal, and the -axis is perpendicular to the -axis and -axes

according to the right-hand rule . In circular orbits, is the direction of the

satellite velocity. The three directions , , and are also known as the roll,

pitch, and yaw axes, respectively. Figure ( 1-11) shows a comparison of the

inertial and orbital frames in an equatorial orbit.

OZ

OX OZ OY

OX

OX


Chapter1 ALEXSAT

Figure ( 1-11) orbital coordinate system

1.2.4 Body Coordinate System

Like the OCS frame, the body coordinate system has its origin at the

satellite’s mass center. This coordinate system is fixed in the body. The -axis in

the direction from the satellite to the Earth , -axis in the direction opposite to

the orbit normal, and the -axis is perpendicular to the -axis and -axes

according to the right-hand rule . In circular orbits, is the direction of the

satellite velocity. The relative orientation between the orbital and body frames

is the satellite attitude, when the satellite is nadir pointing OCS is co-onside

with BCS


Chapter1 ALEXSAT

Figure ( 1-12) Body coordinate system

1.2.5 Device Coordinate System

The device coordinate system is fixed at the device body (i.e. sensor or

actuator …). It define the orientation of the device with respect to satellite BCS

.As shown in Figure ( 1-13) the ZD- axis is Z-axis of the device 's body and XD-

axis is X-axis of the device 's body and YD-axis is perpendicular to ZD-axis and

XD-axis


Chapter1 ALEXSAT


Figure ( 1-13) device coordinate system

Chapter 2 ALEXSAT

CHAPTER (2)

ATTITUDE DETERMINATION AND CONTROL

SUBSYSTEM (ADCS)

In this chapter more detailed explanation about ADCS is introduced.

The impact of other subsystems requirements on ADCS and impact of ADCS

requirements on the other subsystems are presented. In addition, the tasks that

ADCS must perform all over the satellite lifetime and the ADCS operational

modes are describe. Then, an illustration for the physical concepts and

functions of ADCS devices such as sensors and actuators are exhibited.

Besides, different disturbances affecting rotational motion of the satellite are

demonstrated. Finally, the general control methods applied with ADCS are

presented. The control methods and

2.1 What is ADCS?

The attitude determination and control subsystem measures and controls

the satellite's angular orientation (pointing direction).The simplest satellite are

either uncontrolled or achieve control by passive methods such as spinning or

interacting with the Earth's magnetic or gravity fields. These may or may not

use sensors to measure the attitude or position. More complex systems employ

controllers to process the satellite attitude information obtained from sensors

and actuators torquers to control attitude, velocity, or angular momentum. SC

may have several bodies or appendages, such as solar array or communication

antennas, that required certain direction pointing. The complexity of the

attitude control subsystem depends on the number of body axes and appendage

to be controlled, control accuracy, and speed of response, maneuvering

requirements and the disturbance environment.


Chapter 2 ALEXSAT

2.2 Internal influence between satellite mission and other subsystems

upon ADCS

ADCS is very closely coupled with other subsystems; it is interactively

influences and being influenced by other satellite’s subsystems. In the

following section, a briefer description for interaction between ADCS and

other subsystem is presented.

2.2.1 Internal influence between ADCS and Mission requirement

Main mission of the satellite imposes the main requirements on ADCS.

Normally, the requirements associated with the mission are

Earth pointing or inertial pointing ( this will affect in ADCS control

techniques)

• Accuracy /stabilization requirements (this will affect in accuracy of

selected ADCS sensors).

• Slewing requirements (this will affect in selection of actuators types)

• Mission life time (this will affect in life time of selected ADCS devices)

• Orbit parameters (this will affect in the magnitude of environment

disturbance which will perturb ADCS)

2.2.2 Internal influence between ADCS and Structure Subsystem

The ADCS Subsystem directly interacts with the structure subsystem.

The structure of the satellite affects the space craft moment of inertia and

location of its center of mass, which is affecting the dynamics and stability of

the satellite. Also, the rigidity of the structure determines whether the model of

the satellite will be a rigid body or a flexible one. In addition, mounting

accuracies of ADCS devices are one of the main constrains upon the structural

design of the satellite.


Chapter 2 ALEXSAT

2.2.3 Internal influence between ADCS and Power Subsystem

The ADCS and the power subsystem are influencing each other. The

power budget of the satellite must take into account the requirements of the

ADCS sensors and actuators during different operational modes. For satellite

using solar panels, there are additional pointing requirements placed on the

ADCS, if solar panels must be kept aligned with the Sun for optimal

performance

2.2.4 Internal influence between ADCS and Communication Subsystem

If the satellite antenna is required to be pointed within a given accuracy

in order to communication with ground station, the Communication subsystem

will add pointing requirements on the ADCS Subsystem during communication

session.

2.2.5 Internal influence between ADCS and Command and Data

Handling Subsystem

Since the Command and data handling subsystem is the main brain that

organizes the data flow between satellite subsystems; so it imposes

requirements on the volume and rate of data transfer to ADCS or from ADCS

to other subsystems.

2.2.6 Internal influence between ADCS and thermal subsystem

In order to keep temperature of the satellite’s components within

specific range the thermal subsystem may impose maneuver requirements on

ADCS, by pointing the hot side to deep space and pointing the cold side

towards the sun

2.3 ADCS Tasks

According to the previous mutual impacts of ADCS with other

subsystems, ADCS has the following tasks must to be executed all over the

satellite life time. That is, ADCS executing the following tasks from the

moment of separation up to de-orbiting or discarding of the mission.


Chapter 2 ALEXSAT

1. Damping the satellite angular velocity, obtained from LV after satellite

separation.

2. Attitude acquisition of the satellite where the BCS is oriented to be

coincide with the assigned RCS (in Earth observation missions OCS will

be this RCS). In this attitude acquisition the satellite is initially oriented

towards the RCS supports the mission requirements.

3. The satellite three-axis stabilize in the RCS with the required accuracy

during the imaging sessions.

4. Three-axis stabilization in nadir pointing with low accuracy during non-

imaging periods

5. Attitude determination with the required accuracy during all ADCS

operational modes

2.4 Satellite operational modes

According to the above required tasks from ADCS, the ADCS

operational mode will be.

2.4.1 De-tumbling mode (DM)

This mode occurs after the satellite is released from the LV or after

loosing of orientation due to any failure. During this mode the ADCS suppers

the satellite angular velocity that received from the LV, Because of power

limitation this process should be completed within specified period.

2.4.2 Standby Mode (SM)

After DM satellite can have arbitrary attitude Automatically so after

finishing DM, ADCS transfers to SM in order to make attitude acquisition of

satellite (i.e. Orient the satellite BCS to be co-onside with OCS to get

stabilization at nadir pointing with low accuracy) and stay in this case

whenever there is no imaging tasks assigned to the satellite. In this mode the

satellite attitude should be kept even with a low accuracy to avoid loosing the

satellite’s attitude, it is a low accuracy mode. In this mode, the most important

thing is to save the system resources (i.e. lifetime of ADCS devices) and reduce


Chapter 2 ALEXSAT

the consumed power. ADCS stay in SM about 95% of the whole satellite

lifetime

2.4.3 High Accuracy Mode (HAM) or Imaging Mode (IM)

In this mode, ADCS should provide the required control to achieve the

pointing of the payload requirements. As an example, for imaging remote

sensing satellite using magnetic actuator the satellite must be stabilized at nadir

with high accuracy during imaging periods, so this mode called imaging mode

(IM)..

2.4.4 Emergency Mode (EM)

In case of any failure in ADCS (e.g. loosing satellite attitude or any

failure of ADCS devices ) ADCS automatically transfer to EM .In this mode

ADCS switch off all ADCS devices and make diagnostic for ADCS devices

according to command from ground and send TM to ground in order to take

the suitable decision.

2.4.5 Transferring from one operational mode to another

The organization of transfer from one mode to another is shown in

Figure ( 2-1).ADCS operational cyclogram and conditions for transferring

between modes are as follows:

1. After separation from LV and starting of satellite operation ADCS

enters DM.

2. When DM is finished, ADCS directly transfers the satellite to SM and

stay in SM.

3. Before imaging time, within specified period (i.e. Period sufficient to

stabilize the satellite at the required attitude with the required

accuracy),ADCS transfers the satellite to IM.

4. After finishing of imaging task, ADCS transfers the satellite again to

SM

5. In normal cases, the sequence of items 3-4 are repeated.


Chapter 2 ALEXSAT

6. In case of any failure (i.e. failure in ADCS devices or attitude

orientation ), ADCS directly transfers the satellite to EM.

DM

finishing

ADCS

failure

DM SM IM

EM

Imaging

command

Finishing imaging

session

ADCS

failure

ADCS

failure

Fixing of ADCS

failure

Figure ( 2-1) Organization of transferring from one operational mode to

another.

2.5 ADCS devices

A satellite in space must point to a given direction as assigned by the

mission requirements. Many satellites are Earth orientated while others are

inertial space object oriented such as sun or a star of interest. The orientation

of the satellite in space is known as its attitude. In order to achieve control and

stabilization of the satellite, attitude sensors are used to determine the current

attitude and actuators are used to generate required torque to maintain the

required attitude. This section gives brief description of the most common

used ADCS sensors and actuators.

2.5.1 ADCS Sensors

Sensors generally determine the attitude and pointing direction of

satellite with respect to reference objects, this object could be inertial space or a


Chapter 2 ALEXSAT

body of known position. The most commonly used reference objects, Earth,

Sun, stars, geomagnetic field and inertial space.

2.5.1.1 Earth’s Horizon sensor

For near-Earth satellites the Earth covers a large proportion of the sphere

of view and presents a large area for detection. The presence of the Earth alone

does not provide a satisfactory attitude reference hence the detection of the

Earth’s horizon is widely used.

Horizon sensor is infrared device that detect the contrast between the

cold of deep space and the heat of the Earth’s see Figure ( 2-2). Horizon sensors

can provide pitch and roll attitude knowledge for Earth-pointing satellite. For

the better accuracy in low Earth orbit (LEO), it is necessary to correct the data

for the Earth oblateness and seasonal changes in the apparent horizon .Earth’s

Horizon sensor is used in AEROS-I,-2, MAGSAT, SEASAT

Figure ( 2-2) principle of Earth horizon sensor

2.5.1.2 Sun sensor

Sun sensor is widely used with satellite mission due to the special

features of sun as a space object. One of these features is the brightness of the

sun, which makes it easy to be distinguished among other solar and stellar


Chapter 2 ALEXSAT

objects. also the Sun-Earth distance makes it appear as nearly a point source

(0.25 º). Those factors urge ADCS designer to rely upon sun sensors in high

pointing accuracy missions.

Sun sensor measures one or two angles between their mounting base

and incident sunlight. Categories of sensors are ranging from just sun presence

detector, which detects the existence of sun, rather accurate analogue sensor

measuring sun incidence angle, up to high accuracy digital instrument, which

measure the sun direction to accuracy down to one arc-minute. Typical digital

sun sensor is shown Figure ( 2-3).

Sun sensor is accurate and reliable, but require direct line of sight to the

sun. Since most low-Earth orbits include eclipse periods, the attitude

determination system should provide some way of handling the regular loss of

Sun vision. Sun sensor is used in AEROS-1,2 , GEOS-3, MAGSAT, SAGE,

SEASAT.

Figure ( 2-3) Sun sensors

2.5.1.3 Star mapper

Star mapper provides the most accurate absolute pointing information

possible for a satellite attitude. It contains Charged-Coupled Device (CCD)

sensors or Active Pixel Sensors (APS) which provides a relatively inexpensive


Chapter 2 ALEXSAT

Figure ( 2-4) Start sensor

The accuracy and autonomy provided by a star camera would be

impossible without high-speed microprocessors for image processing and star

identification. Star sensor is used in ATS-6, Egyptsat-1, LANDSAT-D·,

MAGSAT.

2.5.1.4 Magnetometers

Magnetometers are simple, lightweight sensors that measure both the

direction and magnitude of the Earth’s magnetic field. They are reliable but

require complex software for interpretation and provide relatively coarse

attitude determination as compared to horizon, sun, and star sensors.

Navigational information are used with a computer model of the Earth’s

magnetic field to approximate the field direction at the satellite’s current


Chapter 2 ALEXSAT

position. Comparison between measured and calculated earth magnetic field is

used to provide information about satellite orientation. Employing estimation

techniques such as Kalman filter, allows magnetometer to work as standalone

device for attitude determination. The Earth’s magnetic field also varies with

time and can't be calculated precisely, so a magnetometer is often used with

another sensor such as a sun, horizon or star sensor or a gyroscope in order to

improve the accuracy. Magnetometer is used in AEROS-1, Egyptsat1, GEOS-

3, SEASA.

Figure ( 2-5) flux-gate magnetometer

2.5.1.5 Inertial Sensor or Gyro

By definition, a gyroscope, is any instrument, which uses a rapidly

spinning mass to sense and respond to changes in the inertial orientation of its

spin axis. There are types of attitude sensing gyros: mechanical and optical

gyro. These sensors measure satellite orientation change.

• Mechanical Gyroscopes

The angular momentum of a gyro, in the absence of an external torque,

remains constant in magnitude and direction in space. Therefore, any rotation

of the satellite about the gyro's input axis results in a precession of the gimbal


Chapter 2 ALEXSAT

Figure ( 2-6) Three degree-of-freedom gyroscope construction geometry.

• Optical Gyroscopes

Optical gyros are gyroscopes that utilize a light ring instead of a

mechanical rotor as the main component to determine rotational changes. All

optical gyros work on the same principle, the Sagnac effect, This effect works

on relativistic principles but can be described in "normal" terms. Two light

beams are traveling through circular paths of the same length but in opposite

directions around in an optical coil. If the optical coil is rotating, one of the

light beams will take a longer period of time to travel the circumference of the

coil. This time lag is measured and converted into a rotational rate for the coil.

Thus, the rotation the gyro is feeling can be measured. The length changes

associated with the light beam are of nuclear dimensions and are difficult to

measure. However, great accuracy can be achieved through the use of this type

of gyroscope. The most common devices of this type is the Ring Laser Gyro


Chapter 2 ALEXSAT

(RLG) and Fiber Optic Gyros (FOG) .Gyros are used in ATS-6,

Egyptsat1,LANDSAT-D·, MAGSAT.

Figure ( 2-7) The QRS11Pro gyro used on Rømer

Typical values for accuracy of ADCS sensors are shown in the

following table

Table 2-1 Ranges of ADCS sensors accuracy

Sensor Accuracy

Earth’s Horizon sensor 0.05 deg. (GEO)

0.1 deg. (LEO)

Sun sensor 0.01 deg.

Star mapper 2 arc. sec.

Magnetometers 1.0 deg. (5,000 Km altitude)

5.0 deg. (200 Km altitude)

Gyro 0.001 deg./hr

2.5.2 ADCS Actuators

ADCS actuators are used to generate the required torque for correction

of satellite attitude. The generated torque is operated against the environmental

disturbance or to force the satellite to point to a cretin direction according to the


Chapter 2 ALEXSAT

control system requirement. A brief description of the commonly used

actuators is presented in this section.

2.5.2.1 Momentum and Reaction Wheel

Momentum wheels and reaction wheels are similar in construction; they

are simply motor with a flywheel mounted on the motor shaft, the difference in

terminology resulting primarily from the speed at which they operate. A

momentum wheel typically operates at constant speed, providing a means of

momentum storage, which in turn provides gyroscopic stabilization to the

satellite. Reaction wheels generally operate at varying speed, providing means

of reacting torque. According to Newton's third law, as a torque is electrically

applied on the motor shaft to cause the wheel to accelerate, an equal and

opposite torque is generated on the satellite, causing the attitude to change.

Momentum wheels are commonly used singly or in pairs to provide spin

stabilization. Normally, reaction wheel system consists of four wheels. Three

reaction wheels are aligned to the satellite pitch, yaw and roll control axes. The

fourth wheel is skewed symmetrically with respect to the orthogonal control

axes. This commonly used configuration provides full redundancy for roll or

pitch or yaw in case of wheel failure. An image of typical reaction wheel is

shown in Figure ( 2-8)


Chapter 2 ALEXSAT

Figure ( 2-8) The TELDIX Momentum and Reaction

Momentum and reaction wheels have the advantage of providing quick

and accurate attitude control. Also, they can be used at any altitude. Their

disadvantage is that they can be costly, massive, and require large amounts of

power. However, wheels may saturate since the RW is a motor that has

maximum speed, since the angular momentum that can be stored in the wheels

is limited, so a secondary control system is used to prevent the stored

momentum from reaching the maximum limit. The secondary control system

can be thrusters system or magnetorquers. Momentum and reaction wheels are

used in Egyptsat1, FLTSATCOM, MAGSAT and SEASAT Error! Reference

source not found..

2.5.2.2 Magnetic actuators

Magnetic actuators enforce a torque on the satellite by generating a

dipole moment, which interacts with the Earth's magnetic field. Generally,

there are two types of magnetic actuators, torque coils and magnetic rods or

magnetorqure.

1. Torque Coils


Chapter 2 ALEXSAT

The torque coil is simply a long copper wire, winded up into a coil.

Generally, three coils are used, one coil in each axis as shown in Figure ( 2-9

The generated dipole moment by each coil is calculated by L

ANiL coil ⋅⋅= ( 2.1)

Where, is the current in the coil, N is the number of windings in

the coil, and A is the area spanned by the coil.

coili

Figure ( 2-9) Torque Coils

2. Torque Rods

Torque rods operate on the same principle as torque coils, but instead of

a large area coil the windings is spun around a piece of ferromagnetic material

with very high permeability as shown in Figure ( 2-10). Ferromagnetic

materials, have a relative permeability, , of up to 106. the generated dipole

moment is calculated by the following formula

η

L

ANiL coil ⋅⋅⋅= η ( 2.2)

Hence, generating specified dipole moment from magnetic rod needs

current much lower than that needed to magnetic coil. However, the weight of


Chapter 2 ALEXSAT

magnetic rod increases drastically because of the metal core in the rods.

Another inconvenience of the torque rods is the hysteresis effect associated

with ferromagnetic core which add nonlinearity to the control loop. Advantages

and disadvantages of using magnetic actuator will be discussed in details in

Error! Reference source not found.. Magnetic actuators are used with

Egyptsat1, MAGSAT, TIROS-IX, LANDSAT-D and AEROS-1, 2Error!

Reference source not found. .

Figure ( 2-10) Torque rods

2.5.2.3 Thruster

Thruster works on the principle of Newton's third law, according to

which "for every action, there is an equal and opposite reaction". Referring to

this principle, if gas is propelled out of a nozzle, the satellite will accelerate in

opposite direction. However, if the nozzles are not pointed directly away from

the center of mass this will lead to cause rotational of satellite as well. In

addition, if two thrusters in opposite direction but not co-lined rotation only

will be generated. The source of the used gas defines the type of thruster .

Cold gass thrusters use high pressure storage tank. Hot gas thrusters use the

combustion of either monopropellant or bipropellant.

Six thrusters are needed to be mounted in pairs to generate the torque

needed for three-axis control. Thruster as actuator is highly accurate and

generate higher torque than RW and magnetic rods. On the other hand, the

structure used with the thrusters is large and heavy. Besides, run out of either


Chapter 2 ALEXSAT

gas or propellant will lead to stop functioning of thrusters. Thrusters are used in

ATS-3,6 , FLTSATCOM, GOES-I and SKYNETError! Reference source not

found. .

Figure ( 2-11) Torque generated thruster mounted to satellite

2.6 Disturbance Environment

In an Earth orbit, the space environment imposes several external

torques that the ADCS system must tolerate. According to orbit altitude, three

or four sources of disturbing torques are affecting the space craftError!

Reference source not found. . These torques are; gravity gradient, magnetic

field effect, solar radiation pressure, and aerodynamic forces. Those

disturbances are affected by the satellite’s geometry, orientation, and mass

properties in addition to satellite orbital altitude.

2.6.1 Gravity Gradient Disturbance

Any object with nonzero dimensions orbiting Earth will be subjected to

a “gravity-gradient” torque. In short, the portions of the satellite that are closer

to the Earth are subjected to a slightly larger force than those parts farther away

Error! Reference source not found. . This creates a force imbalance that has

a tendency to orient the satellite towards the center of Earth in order to

compensate this imbalance. According to [Error! Reference source not

found. the gravity gradient torque can be determined by equation ( 2.3) . The

worst case torque arises at o90=Θ


Chapter 2 ALEXSAT

)2sin(23

3 Θ−= iiZZgg JJR

T μ ( 2.3)

Where,

Tgg: is the resulting gravitational torque [Nm

μ: is the gravitational constant of the earth [m³/s² (μ = 3.896*1014m³/s²)

Jii :is the moment of inertia tensor for the satellite in i axis.(in body

coordinate system) [kgm² (i=x,y,z)

Θ Is the maximum deviation angel from the local vertical [rad

R: is the distance between satellite center of mass and earth center of

mass [km

The previous formula for calculation of gravity gradient is used to give

course estimation of gravity gradient disturbance torque but an accurate

formula given in Error! Reference source not found. is used in calculation of

satellite mathematical model

2.6.2 Magnetic Field Disturbance

Magnetic field torques are generated by interactions between the

satellite magnetic dipole and the Earth’s magnetic field. This satellite magnetic

dipole is the summation of two components; first component is the induced

magnetic dipole, which is caused by current running through the satellite

wiring harness and second component is the residual dipole moment, which is

caused due to magnetic properties of the satellite components. The satellite

magnetic dipole exhibits transient and periodic fluctuations due to power

switching between different subsystems. These effects can be minimized by

proper placement of the wiring harness. The magnetic torque is calculated by

following formula

BDTm ×= ( 2.4)

Where

D = the vector of total satellite magnetic dipole.

B = local geomagnetic field vector.


Chapter 2 ALEXSAT

In the worst case, the vectors are perpendicular to each other and the

cross product turns into a product of scalar values.

2.6.3 Solar Radiation Pressure Disturbance

Solar radiation pressure is a result of the transfer of momentum from

photons of light to the surface of the satellite. The result of this pressure across

the satellite surface is a force that acts through the center of pressure, , of the

satellite. In most cases, the center of pressure is not co-onside with the center of

mass of the satellite, thus a torque will be generated around the center of

mass see Figure ( 2-12). For Earth-orbiting satellite, where the distance from

the satellite to the Earth is small compared to the Earth-Sun distance, the mean

solar flux acting on the satellite is considered a constant (regardless of orbital

radius or position).

psc

cm

The solar radiation torque is calculated using the following equation

[Error! Reference source not found. .

( 2.5) )()cos()1( gpssSp cciqAc

SoT −⋅⋅+⋅⋅=

Where

So is solar constant [W/m² = 1428 W/m² (max)

c is speed of light [m/s = 3*108 m/s

A is the cross sectional area subjected to solar radiation pressure [m²

q is reflectance factor (0: perfectly absorbing, 1: perfectly reflecting)

si is the angle of sun light incidence [rad

cps is the center of pressure [m

cg is the center of gravity [m

Referring to the previous assumptions, the solar pressure disturbance

torque is the only one that is not dependent of the orbit altitude. However, it is

dependent of the sun incidence angle i. The worst case torque arises at i = 0°.


Chapter 2 ALEXSAT

2.6.4 Aerodynamic Disturbance

Aerodynamic torques are due to atmospheric drag acting on the satellite

as shown in Figure ( 2-12. Aerodynamic torques can be quite significant,

especially at low altitudes (less than 500). At higher altitudes the aerodynamic

torque is almost negligible. These torques is difficult to be calculate because

changing of some parameters, such as cross sectional area of satellite subjected

to the aerodynamic drag during tilting. In addition, atmospheric density varies

significantly with solar activity. The generated torque due to aerodynamic

effects is calculated by ( 2.6) .

( )gpaCDad ccvAcT −⋅⋅⋅⋅⋅= 2

21 ρ ( 2.6)

Where

ρ is the density [kg/m³

cD is the coefficient of drag

A is the cross sectional area subjected to atmospheric drag [m²

vc is the orbital velocity [m/s

cps is the center of pressure [m

cg is the center of gravity [m

Figure ( 2-12) Sunlight and drag effect

2.7 Attitude Control techniques

There are different techniques to apply control torque for disturbance

compensation and to maintain the required orientation . For these purposes, two

types of control techniques are often employed , passive and active control


Chapter 2 ALEXSAT

Error! Reference source not found. Error! Reference source not found. .

Since Attitude control system, is highly mission dependent, so the decision to

use a passive or an active control technique or a combination of them depends

on mission pointing and stabilization requirements.

2.7.1 Passive Control

For missions with rather coarse orientation requirements, passive control

techniques are used for attitude control. The main advantageous of these

techniques are saving resources concerning both mass and power and the

associated cost. In addition, they provide longer lifetime for the space mission.

However, a poor pointing accuracy is obtained. The most common passive

control techniques are passive magnetic system (i.e. Permanent magnate),

gravity gradient and spin stabilization Error! Reference source not found. .

2.7.1.1 Passive magnetic

In this method, the concept of magnetic compass is applied, that is, the

satellite is equipped with permanent magnet that will keep the alignment

between certain axis of the satellite with geomagnetic field vector .As a result,

the south pole of the magnet will be drawn towards the magnetic north pole of

the Earth, and vice versa. This will lead to a slight tumbling motion with two

revolutions per orbit and no possibilities of controlling spin around the magnets

axis as shown in Figure ( 2-13) so continues nadir pointing will not be possible.

Permanent magnet technique is used in AZUR-1 Error! Reference source not

found. .


Chapter 2 ALEXSAT

Figure ( 2-13) passive magnetic control orientation profile.

2.7.1.2 Gravity-gradient stability

Gravity-gradient stability uses the mass characteristics of the satellite to

maintain the nadir pointing towards Earth (as described in 2.6.1). The

magnitude of gravity-gradient torque decreases with the cube of the orbit

radius, and symmetric around the nadir vector, thus not influencing the yaw of

satellite. Therefore, the gravity gradient stability is used in simple satellite in

LEO without yaw orientation requirements Error! Reference source not

found. .

Yet, stability in the gravity gradient case depends upon the the

configuration of the mass characteristics of the space craft. The following

condition is necessary for gravity-gradient stability [Error! Reference source

not found. : JzzJxxJyy & Jzz Jxx Jyy +<>> ( 2.7)

Where Jii :is the moment of inertia tensor for the satellite in i axis.(in

body coordinate system) (i=x,y,z)


Chapter 2 ALEXSAT

As a result, the gravity gradient stability can be achieved by

manipulation of lay out of the satellite's components to grantee the above

mentioned condition ( 2.7). Other solution is to add a sufficient mass on a

deployed boom to reach the stability condition. This will increase the moment

of inertia in the directions transverse to the boom, and the satellite will be

stable with the mass pointed toward or away from the earth. Gravity gradient

stability is suffering from continuous oscillation about nadir due to lack of

damping. Hence, gravity-gradient stabilization should be supported with

damping system to reduce the small oscillation around the nadir vector.

Gravity-gradient stabilization technique is used in DODGE, GEOS-3, and

RAE-2 Error! Reference source not found. .

2.7.1.3 Spin stabilization

Spin stabilization technique applies the gyroscopic stability to passively

resist the effect of disturbance torques about the spinning axis. Spin-stabilized

satellites spins about their major or minor axes, so angular momentum vector

remains approximately fixed with respect to inertial space. [Error! Reference

source not found. . Spinning satellite is classified according to spinning object

to single or dual spin. The stability criteria and the corresponding spinning axis

is predicted according to the following analysis.

2.7.1.3.1. Single Spin

In single spin satellites, the whole satellite spins about the angular

momentum vector as shown in Figure ( 2-14) This method of stabilization is

simple and has a high reliability. The cost is generally low, and it has a long

system life. However, Spin-stabilized satellite are subject to nutation and

precession, but have a gyroscopic resistance which provides stability about the

transverse axis.

On the other side, spinning satellite will have poor maneuverability.

Beside, it will not be suitable for systems that need to be Earth pointing, such

as payload scanners and communication antennas. Single spin stabilization


Chapter 2 ALEXSAT

technique is used in AEROS-I,2, ALOUETIE-I,2and ARIEL-I Error!


Figure ( 2-14) spin stabilization

2.7.1.3.2. Dual Spin

In satellite with dual spin, a major portion of the satellite is spun, while

the payload section is despun. This technique is favorable because fixed inertial

orientation is possible on the despun portion. This method of stabilization has a

few disadvantages, however. This system is much more complex, which leads

to an increase in cost and a decrease in reliability. In addition, the stability is

sensitive to mass imbalances. Duel spin stabilization technique is used in ANS,

ATS-6, SEASAT and SMM Error! Reference source not found. .

2.7.2 Active control techniques

For complex mission requirements, satellite requires continues

autonomous control about the three axes during the mission. In general, active

control systems employ momentum exchange wheels, magnetic control

devices, and thrusters. Advantages of these systems are high pointing accuracy,

and a not constrained to inertial pointing like spin stabilization technique.

However, the hardware is often expensive, and complicated, leading to a higher

weight and power consumption.


Chapter 2 ALEXSAT


2.7.2.1 Momentum exchange Wheels

Three-axis stabilization through momentum exchange wheels applies

reaction wheels, momentum wheels, and control moment gyros. This is to

provide three axis stabilization. Advantages and disadvantages of this wheel

system are discussed in 2.5.2.1. Three-axis stabilization technique using wheels

is used in Egyptsat1, FLTSATCOM, MAGSAT and SEASAT Error!


2.7.2.2 Magnetic actuators

Magnetic actuators devices use the interaction of the satellite magnetic

dipole moment and the Earth’s magnetic field to provide a control torque.

Magnetic control torques work better in low Earth orbits than higher orbits,

such as geostationary, because as the distance from the Earth increases, the

geomagnetic strength decreases. Advantage and disadvantage of magnetic

actuators is discussed in 2.5.2.2 Three-axis stabilization technique using

magnetic actuators is used in Egyptsat1, MAGSAT, TIROS-IX, LANDSAT-D

and AEROS-1, 2Error! Reference source not found. .

2.7.2.3 Thrusters

Mass propulsive devices, such as thrusters, can be used for three-axis

stabilization. These often consist of six or more thrusters located on the satellite

body. The strength of the obtainable torque is dependent on the thrust level as

well as the torque-arm length about the axis of rotation. Advantage and

disadvantage of thrusters is discussed in 2.5.2.3 2.5.2.2. Three axis stabilization

technique using thrusters is used in ATS-3,6 , FLTSATCOM, GOES-I,

SKYNETError! Reference source not found. .

Adcs orbit intro

Engineering

Transcript of Adcs orbit intro