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19.1 Attitude Determination and Control Systems

Scott R. Starin, NASA Goddard Space Flight Center

John Eterno, Southwest Research Institute

In the year 1900, Galveston, Texas, was a bustling direct hit as Ike came ashore. Almost 200 people in the
community of approximately 40,000 people. The Caribbean and the United States lost their lives; a
former capital of the Republic of Texas remained a tragedy to be sure, but far less deadly than the 1900
trade center for the state and was one of the largest storm. This time, people were prepared, having
cotton ports in the United States. On September 8 of received excellent warning from the GOES satellite
that year, however, a powerful hurricane struck network. The Geostationary Operational Environmental
Galveston island, tearing the Weather Bureau wind Satellites have been a continuous monitor of the
gauge away as the winds exceeded 100 mph and world’s weather since 1975, and they have since been
bringing a storm surge that flooded the entire city. The joined by other Earth-observing satellites. This weather
worst natural disaster in United States’ history—even surveillance to which so many now owe their lives is
today—the hurricane caused the deaths of between possible in part because of the ability to point
6000 and 8000 people. Critical in the events that led to accurately and steadily at the Earth below. The
such a terrible loss of life was the lack of precise importance of accurately pointing spacecraft to our
knowledge of the strength of the storm before it hit. daily lives is pervasive, yet somehow escapes the notice
of most people. But the example of the lives saved from
In 2008, Hurricane Ike, the third costliest hurricane ever Hurricane Ike as compared to the 1900 storm is
to hit the United States’ coast, traveled through the Gulf something no one should ignore. In this section, we will
of Mexico. Ike was gigantic, and the devastation in its summarize the processes and technologies used in
path included the Turk and Caicos Islands, Haiti, and designing and operating spacecraft pointing (i.e.
huge swaths of the coast of the Gulf of Mexico. Once attitude) systems.
again, Galveston, now a city of nearly 60,000, took the

Figure 19-1: Satellite image of Hurricane Ike (NASA image).

Attitude is the three-dimensional orientation of a can relate the angular velocity of the spacecraft, w, to H
vehicle with respect to a specified reference frame. by the equation
Attitude systems include the sensors, actuators,
avionics, algorithms, software, and ground support H=Iw+h
equipment used to determine and control the attitude of
a vehicle. Attitude systems can have a variety of names, where I is the moment of inertia and h is the angular
such as attitude determination and control system momentum stored by any rotating objects that are part
(ADCS), attitude ground system (AGS), attitude and of the spacecraft, such as momentum wheels or
orbit control system (AOCS), guidance, navigation and gyroscopes. So, by the product rule of calculus the
control (GNC), or whatever other term describes the Euler equations can be rewritten as a matrix equation:
designers’ focus in achieving the attitude needs of a
particular mission. When we use an acronym in this
Iw+Iw +h=T-wxH
section, we will use ADCS, but any given specialist
may be more familiar with other terms.
or, after moving some terms around:
Spacecraft attitude changes according to the
fundamental equations of motion for rotational Iw =T-h -I w-wxH
dynamics, the Euler equations, here expressed in vector
form in the spacecraft’s reference frame: The form of Equation 4 allows us to understand how
attitude can change due to a variety of causes. The first
H=T-wxH
term on the right-hand side represents external torque’s
direct contribution to attitude dynamics; this term
includes how some actuators can be used to control
This vector equation represents the conservation
spacecraft attitude by creating external torques. The
equations for the physical vector quantity of a body or
second term gives the relationship between changes in
collection of bodies called angular momentum, which is
onboard rotating objects’ speeds and changes in the
denoted by H. Recall that linear momentum is the
spacecraft’s rotational velocity; this term is where
translational motion of a body that will remain constant
certain other control actuators enter into the dynamics
unless a force acts to change it, and it is calculated (in
as so-called internal torques. The third term shows how
Newtonian physics anyway) as mass times velocity.
changes in the spacecraft moment of inertia
Analogously, angular momentum is the rotational
(representing how mass is distributed in the spacecraft),
motion of a body that will continue unless changed by a
such as by solar array articulation, can affect attitude
torque, and it is calculated as the body’s moment of
dynamics; in the absence of changes in mass properties,
inertia times its angular velocity. The moment of inertia
the third term disappears. The fourth term is called the
is a 3-by-3 matrix of values that describe the
gyroscopic torque, and it shows how the angular
distribution of mass in a body. There is always a
momentum appears to change direction, but not
coordinate frame, called the principal axis frame, for
magnitude, in the spacecraft’s frame of reference when
which the moment of inertia matrix is diagonal. The
the spacecraft is rotating. All these effects combine to
difference between the geometric and principal axis
determine the rate of change of the angular velocity on
frames is of great interest to ADCS designers, as we
the left-hand side.
will see later in this section.
Attitude determination is the process of combining
Note that in the form above, the Euler equation makes it
available sensor inputs with knowledge of the
clear that the magnitude of angular momentum in a
spacecraft dynamics to provide an accurate and unique
system can only be changed by applying external
solution for the attitude state as a function of time,
torques, T, because the change due to the term w x H
either onboard for immediate use, or after the fact (i.e.
can only change the direction of H, not the magnitude.
post-processing). With the powerful microprocessors
In fact, a body’s angular momentum is always
now available for spaceflight, most attitude algorithms
conserved in the absence of external torques upon it,
that formerly were performed as post-processing can
even when parts of the body can move with respect to
now be programmed as onboard calculations.
other parts, such as a gyroscope spinning in its housing.
Therefore, though there are still good engineering
If part of the body starts to spin in one direction in the
reasons for certain processes to be performed only by
absence of external torques, the rest of the body will
ground-based attitude systems, it will be sufficient to
have to spin in the opposite direction so that total
focus our attitude determination discussions in this
angular momentum is conserved. With this in mind, we
chapter on the design and implementation of onboard
systems.
 spacecraft orientation. These two general types of
The product of attitude determination, the attitude attitude control are not mutually exclusive. A spacecraft
estimate or solution, is attained by using sensors to may be mostly or usually passively controlled and yet
relate information about external references, such as the include actuators to adjust the attitude in small ways or
stars, the Sun, the Earth, or other celestial bodies, to the to make attitude maneuvers (i.e. slews) to meet other
orientation of the spacecraft. Frequently, any single objectives, such as targets of opportunity or
sensor has a noise level or other drawback that prevents communication needs.
it from providing a fully satisfactory attitude solution at
all times. Therefore, more than one sensor is often So, external torques change the total angular
required to meet all mission requirements for a given momentum, and internal torques exchange momentum
mission. between different rotating parts of the spacecraft. In this
way reaction wheels or control moment gyroscopes
The combination of information from multiple sensors may be used to change spacecraft pointing without
is a complex field of study. The possibilities for any affecting total angular momentum. Because
given mission range from simple logical combination of environmental disturbances create external torques on
sensors, depending on mode, to modern information the spacecraft, they also create angular momentum that
filtering methods, such as Kalman filtering. Many must be either stored or removed by the attitude system.
methods require some prediction of a future attitude Small external torques that vary over the course of an
from current conditions. Because all spacecraft sensors orbit but have a mean of zero may be managed just
must use the spacecraft’s reference frame as a basis for through storage, but those torques that have a non-zero
attitude determination, the development of angular mean (secular torques) will cause a gradual increase in
momentum according to the spacecraft’s frame of angular momentum, and this momentum build-up must
reference can be important for some attitude eventually be removed with actuators that create
determination algorithms. This is the reason for the external torques. Thrusters, magnetic torquers, or even
spacecraft-referenced aspect of the Euler equations. solar trim tabs can be used to create controlled external
torques on the spacecraft, thus controlling the total
Attitude control is the combination of the prediction of angular momentum.
and reaction to a vehicle’s rotational dynamics.
Because spacecraft exist in an environment of small and Attitude system design is an iterative process. Table 19-
often highly predictable disturbances, they may in 1 lists typical steps in a design process and what inputs
certain cases be passively controlled. That is, a and outputs would be expected for each step. Figure
spacecraft may be designed in such a way that the 19-2 presents all the processes involved in attitude
environmental disturbances cause the spacecraft systems. The FireSat II spacecraft, shown in Figure 19-
attitude to stabilize in the orientation needed to meet 3, and the Supplemental Communications System
mission goals. Alternately, a spacecraft may include (SCS) constellation of spacecraft, shown in Figure 19-
actuators that can be used to actively control the 4, will be used to illustrate this process.

Table 19-1: Steps in attitude system design. An iterative process is used for designing the ADCS.
Step Inputs Outputs
1 a) Define control modes Mission requirements, mission List of different control modes
1b) Define or derive system-level profile, type of insertion for launch during mission.
requirements by control mode vehicle Requirements and constraints.
2) Quantify disturbance environment Spacecraft geometry, orbit, Values for torques from external
solar/magnetic models, mission and internal sources
profile
3) Select type of spacecraft control by Payload, thermal & power needs Method for stabilization & control:
attitude control mode Orbit, pointing direction three-axis, spinning, gravity
Disturbance environment gradient, etc.
Accuracy requirements
4) Select and size ADCS hardware Spacecraft geometry and mass Sensor suite: Earth, Sun, inertial,
properties, required accuracy, orbit or other sensing devices.
geometry, mission lifetime, space Control actuators: reaction wheels,
environment, pointing direction, thrusters, magnetic torquers, etc.
slew rates. Data processing avionics, if any, or
processing requirements for other
 subsystems or ground computer.
5) Define determination and control Performance considerations Algorithms and parameters for
algorithms (stabilization method(s), attitude each determination and control
knowledge & control accuracy, mode, and logic for changing from
slew rates) balanced against one mode to another.
system-level limitations (power
and thermal needs, lifetime, jitter
sensitivity, spacecraft processor

6) Iterate and document All of above Refined mission and subsystem

requirements.
More detailed ADCS design.
Subsystem and component

Figure 19-2. Diagram of a Complete Attitude Determination and Control System. Definitive attitude
determination usually occurs in ground processing of telemetry, whereas onboard, real-time determination design
focuses on being extremely reliable and deterministic in its operation.
 Figure 19-3. Hypothetical FireSat II Spacecraft. We use this simplified example of a low-Earth orbiting satellite
to discuss key concepts throughout the section.

Figure 19-4. One Member of the Hypothetical Supplemental Communications System (SCS) Constellation. We
will also use this collection of three spacecraft in medium Earth orbit to illustrate attitude system design practices.

19.1.1 Control Modes and Requirements better understanding of the actual needs of the mission
The first step of the attitude system design process is often results from having these modes of controlling the
the definition of guiding requirements based on mission spacecraft well defined. This iteration takes place in a
goals. Since mission goals often require more than one trade space where a single set of ADCS hardware must
mode of operating a spacecraft, the guiding be used in different ways to meet different sets of
requirements generally begin with a description of the requirements. The ADCS will also be dependent on
control modes the ADCS is expected to execute to meet certain other subsystems, such as the power and
those goals. Tables 19-2 and 19-3 describe typical structural subsystems. Attitude needs will also impose
spacecraft control modes and requirements. requirements on other subsystems, such as propulsion,
thermal control, and structural stability. Figure 19-5
The final form of ADCS requirements and control shows many of the complex interactions needed to
modes will be the result of iteration; control modes are bring the ADCS design in line with the needs of the
designed to achieve certain sets of requirements, and whole mission.

Table 19-2: Typical attitude control modes. Performance requirements are frequently tailored to these different
control operation modes.
Mode Description
Orbit Insertion Period during and after boost while spacecraft is brought to final orbit. Options include no
spacecraft control, simple spin stabilization of solid rocket motor, and full spacecraft control
using liquid propulsion system. May drive certain aspects of ADCS design.
Acquisition Initial determination of attitude and stabilization of vehicle for communication with ground
and power generation. Also may be used to recover from power upsets or emergencies.
Normal Mission, Used for the vast majority of the mission. Requirements for this mode should drive system
On-Station design.
Slew Reorienting the vehicle when required.
Contingency or Used in emergencies if regular mode fails or is disabled. Will generally use less power or
Safe fewer components to meet minimal power and thermal needs.
Special Requirements may be different for special targets or time periods, such as when the satellite
passes through a celestial body’s shadow, or umbra.
Table 19-3: Typical attitude determination and control performance requirements. Requirements need to be
specified for each mode. The following lists the performance criteria frequently specified.
Criterion Definition* Examples/Comments
Accuracy Knowledge of and control over a vehicle’s 0.25 deg, 3 6 , often includes determination errors
attitude relative to a target attitude as defined along with control errors, or there may be separate
relative to an absolute reference requirements for determination & control, and even
for different axes
Range Range of angular motion over which Any attitude within 30 deg of nadir. Whenever
determination & control performance must rotational rates are less than 2 deg/sec.
be met
Jitter Specified bound on high-frequency angular 0.1 deg over 60 sec, 1 deg/s, 1 to 20 Hz; prevents
motion excessive blurring of sensor data
Drift Limit on slow, low-frequency angular 0.01 deg over 20 min, 0.05 deg max; used when
motion vehicle may drift off target with infrequent command
inputs
Transient Allowed settling time or max attitude 10% max overshoot, decaying to <0.1 deg in 1 min;
Response overshoot when acquiring new targets or may also limit excursions from a set path between
recover from upsets targets
*Definitions vary with procuring and designing agencies, especially in details (e.g. 1 6 or 3 6 , amount of averaging or
filtering allowed). It is always best to define exactly what is required.

Figure 19-5: The Impact of Mission Requirements and Other Subsystems on the ADCS. Direction of arrows
shows requirements flow from one subsystem to another.

For many spacecraft the ADCS must control vehicle Once the spacecraft is on station, the payload pointing
attitude during the firing of large liquid or solid rocket requirements usually dominate. These may require
motors for orbit insertion or management. Large motors planet-relative or inertial attitudes and fixed or spinning
can create large disturbance torques, which can drive fields of view. There is usually also a need for attitude
the design to larger actuators than may be needed for slew maneuvers, and the frequency and speed of those
the rest of the mission. maneuvers must be defined. Reasons for slews can
include:
 -Acquiring the desired spacecraft attitude driver for control mode and hardware selection, but the
initially or after a failure sensitivity of the ADCS designer must be to more than
-Repointing the payload’s sensing systems to just the external torque disturbances of the operational
targets of opportunity or for calibration purposes orbit. For example, some attitude sensors, such as star
-Tracking stationary or moving targets, cameras that use charge-coupled devices (CCDs) for
including communication stations imaging, can be highly sensitive to the intense radiation
-Directing the vehicle’s strongest motor(s) to in the Van Allen belts of the Earth’s magnetosphere.
the proper direction relative to orbital motion. Depending on the specific model, the star camera may
underperform or even provide no information at all
when the spacecraft occupies these regions.
19.1.2 Quantify the Disturbance Environment.
The environment in which the spacecraft will operate Only three or four sources of torque matter for the
constrains what types of control methods will be typical Earth-orbiting spacecraft: gravity-gradient
effective. For example, the relatively strong magnetic effects, magnetic field torques on a magnetized vehicle
fields that occur in low Earth orbit (LEO) can create (as most spacecraft will be), impingement by solar-
disturbance torques that need to be managed, but they radiation, and aerodynamic torques for LEO satellites.
also allow the use of magnetic torquers, a means of Figure 19-6 summarizes the relative effects of these
attitude control not available at much higher altitudes disturbances for different flight regimes. Chapter 7
like geosynchronous orbit (GEO). Here, we will focus describes the Earth environment in detail, and Hughes
on the torque disturbance environment as the primary [2004] provides a thorough treatment of disturbances.

Figure 19-6. Effects of major environmental disturbances on spacecraft attitude system design. The diagram
has a roughly logarithmic scale of altitude. The columns represent the four major disturbance sources, with the
intensity of color for each column indicating the strength of that disturbance in the various flight regimes.
 is created. Solar radiation pressure is more intense on
Other external disturbances to the spacecraft are either certain surfaces (reflective) than others (absorptive).
small relative to the four main external disturbances, The total pressure force over the Sun-pointing surface
such as infrared emission pressure, or they are limited of a spacecraft can be considered to act through a center
in time, such as outgassing. Occasionally, what is of pressure (cp) with an average reflectance, and the
normally negligible can become surprisingly large, offset of that point from the cm results in solar radiation
even exceeding the usual disturbance torque sources, pressure torque. The location of this cp is a function of
but this is one of the reasons for maintenance of healthy attitude as well as surface properties. Some modern
engineering margins and operational plans that are surfaces can have their reflectance change with a
adaptable to unforeseen events. change in applied voltage, usually for thermal reasons,
but which results in a controlled change in cp location.
Centroids. Estimation of environmental torques often So, in detailed modeling of spacecraft, the
requires the use of geometrical averaging. Anyone with determination of the weighted averages of various
a technical education will be familiar with the centroid forces is important to a good understanding of the
of an area, but it may have been some time since the torque environment.
reader encountered this concept. The centroid is the
point in an area through which any line drawn in any Modeling Major Disturbances. Now, we will present
direction will evenly divide moments about the line (or the equations used to model major disturbances with
any point along the line). To express it another way, the some explanation and demonstration of they can be
sum of all area elements multiplied by their distances used to design attitude systems. After the explanations,
from a line will be zero for any line passing through the Table 19-4 will show disturbance calculations for the
centroid. In a sense, it is the average point for the area. FireSat II and SCS examples.
If a source of pressure were applied evenly over the
area, the solar pressure force could be represented as Solar Radiation Pressure. Sunlight has momentum, and
being applied entirely at the centroid for the purposes of therefore it exerts pressure on those objects it strikes. If
determining moments, and therefore disturbance an object absorbs all the sunlight falling on it, then it
torques. A solid body can also have centroids. The absorbs all of its momentum and experiences a certain
center of mass (cm) is the point (usually inside) the pressure force because of it. If the sunlight is instead
body through which any plane will divide the mass reflected exactly back along its path, such as by a
moment evenly. By applying a force at or along the mirror, the pressure force felt is twice as much.
center of mass, no torques are created. This is why
freely rotating bodies rotate about their centers of mass. If a sunlit flat plate were mirrored on one half and
painted black on the other, the pressure distribution
As a practical example, the point that may be regarded across the plate would be uneven and a torque would
as the location of a body for purposes of gravitational result. Alternately, if the plate were all black, but a
forces is called the center of gravity (cg); i.e. all effects weight were attached to one end in the plate’s shadow,
of gravity on the body can be considered to act at the a torque would also result because the center of
cg. In the essentially uniform gravity that we humans pressure would be in the center of the plate, but the
occupy, the center of mass is usually indistinguishable center of mass would be closer to the weighted end.
from the center of gravity, but in the free-fall of a space These phenomena are called solar radiation pressure
orbit, the absence of direct gravitational forces and (SRP) torques.
torques means that the change, or gradient, of gravity
over the extent of a body can be important. For Now imagine a spacecraft like FireSat II in sunlight.
elongated or flattened objects in orbit, the cm may be Some parts of the spacecraft stick out further from the
offset from the cg, so that the gravitational force is center of mass than others. Some surfaces are more
effectively applied with an offset from the cm, creating reflective than others; solar arrays would absorb more
a torque—this is the gravity gradient torque. Note that light than reflective metallic surfaces would. Also,
the cg is a function of the current attitude of the surfaces that are angled with respect to the Sun would
spacecraft, not just its mass configuration, which is have less pressure on them than similar surfaces
critical in attitude analysis. directly facing the Sun. All this goes to demonstrate
that accurately predicting SRP torques is very tricky.
Other environmental effects can be understood in terms That said, a good starting estimate can be gleaned by
of offsets between centroids of different effects on a assuming a uniform reflectance and using the following
body. When the aerodynamic force centroid, which is at equation:
the centroid of the ram area (the area presented to the
velocity direction), is not aligned with the cm, a torque
 field, much like a compass needle. The Earth’s
Ts = Φ As ( 1 + q) (cps − cm)cos ϕ
magnetic field is complex, asymmetric, not aligned
with the Earth’s spin axis, and varies with both
where T s is the SRP torque, Φ is the solar constant geographical movement of the dipole and changes in
adjusted for actual distance from the Sun (average solar particle flux. However, for use in the ADCS
value: 1367 W/m 2 ), c is the speed of light (3 x 10 8 m/s), design process, it is usually sufficient to model the
As is the sunlit surface area in m2, q is the unitless Earth’s magnetic field as a dipole and to determine the
reflectance factor (ranging from 0 for perfect absorption maximum possible value of the magnetic torque for a
to 1 for perfect reflection), ϕ is the angle of incidence spacecraft’s altitude. The following equation yields this
of the Sun, and cp s and cm are the centers of solar maximum torque:
radiation pressure and mass.

Tm = DB D
Atmospheric Drag. In much the same way photons R
^ )
striking a spacecraft can exert pressure, so too can the
rarified atmosphere that clings to Earth (and certain where T m is the magnetic torque, D is the spacecraft’s
other planets) at the edge of space. The atmospheric residual dipole moment in A ⋅ m2 , and B is the magnetic
density is roughly an exponentially decaying function field strength in tesla. The magnetic field strength in
of altitude, so that generally only spacecraft in low turn is calculated from M, the magnetic moment of the
Earth orbit (LEO) encounter enough particles to cause Earth multiplied by the magnetic constant (M = 7.8 x
noticeable disturbances. Those that do experience a 3
1015 tesla⋅ m ); R, the distance between the spacecraft
pressure force known as atmospheric (or aerodynamic) and the Earth’s center in m, and λ, which is a unitless
drag. The atmospheric drag force itself is an important
function of the magnetic latitude that ranges from 1 at
consideration for orbit planning (Chapter 9) and orbit
the magnetic equator to 2 at the magnetic poles. So, a
prediction and tracking (Section 19.2). When the center
polar orbit will see roughly twice the maximum
of atmospheric pressure, determined by the spacecraft
magnetic torque of an equatorial orbit.
area exposed to the atmosphere in the direction of the
orbital velocity (i.e. ram direction), is not aligned with
Gravity Gradient. As described in the earlier subsection
the center of mass, a torque results. The atmospheric (or
on centroids, gravity gradient torques are caused when
aerodynamic) torque can be estimated as
a spacecraft’s center of gravity is not aligned with its
center of mass with respect to the local vertical.
Ta = 2 ρ Cd Ar V 2 (cp a − cm) Without getting into the math of the matter, the center
of gravity of a spacecraft in orbit is dependent on its
attitude relative to Earth (or whatever body the
where T a is the atmospheric drag torque, ρ is the spacecraft is orbiting), and that cg is not, in general, the
atmospheric density in kg/m3, Cd is the drag coefficient same as the center of mass. However, when one of the
(usually between 2.0 and 2.5 for spacecraft), A r is the spacecraft’s principal axes is aligned with the local
ram area in m2, V is the spacecraft’s orbital velocity in vertical, the cg is always on that principal axis, and
m/s, and cp a and cm are the centers of aerodynamic therefore there is no gravity gradient torque. The
pressure and mass in m. Average atmospheric density gravity gradient torque increases with the angle
and orbital velocity as functions of altitude are between the local vertical and the spacecraft’s principal
tabulated in the Appendices of this text. axes, always trying to align the minimum principal axis
with the local vertical.
Magnetic Field. The Earth’s liquid core is a dynamo
that generates a magnetic field powerful enough to have A simplified expression for the gravity gradient torque
important effects on the space surrounding the planet. for a spacecraft with the minimum principal axis in its
Most spacecraft have some level of residual magnetic Z direction is
moment, meaning they have a weak magnetic field of
their own. These residual moments can range anywhere
from 0.1-20 A⋅ m2 , or even more depending on the Tg = Iz Iy sin{ 2θ )
2R
spacecraft’s size and whether any onboard
compensation is provided. Where T g is the gravity gradient torque about the X
principal axis, µ is the Earth’s gravitational constant
When a spacecraft’s residual moment is not aligned (3.986 x 10 14 m3/s2) , R is the distance from the center
with a local magnetic field, it experiences a magnetic
of the Earth in m, θ is the angle between the local
torque that attempts to align the magnet to the local
vertical and the Z principal axis, and I y and Iz are the
moments of inertia about Y and Z in kg • m2 .

Table 19-4. Disturbance Torque Summary and Sample Calculations. See text for detailed discussion and
definition of symbols. FireSat II is mainly affected by magnetic torques, with the 30-degree offset attitude also
affected by gravity gradient torques. SCS satellites are mainly affected by solar radiation pressure torques.
Disturbance Type FireSat II SCS
Solar Cyclic for FireSat II is very small and Earth- SCS is small and Earth-pointing, so the
radiation Earth-oriented; oriented (though not Earth-pointing), surface area will be fairly small,
constant for and has body-mounted arrays. However, the need to balance the masses
Sun-oriented Therefore, its surface area is small and of three stowed satellites on the launch
its cp s is close to its cm. vehicle may reduce our control over
mass placement and may cause the
center of pressure to be considerably
offset from the cm with respect to the
Sun.

A s = 1.2 m x 1.1 m = 1.3 m2 ; q = 0.6 A s = 2 m x 1.5 m = 3 m 2 ; q = 0.6

cp = 0 deg; cp s –cm = 0.1 m cp = 0 deg; cp s–cm = 0.3 m

T s =(1 367)(1.3)(1+0.6)(0.1)/(3x1 0 8 ) T s =(1 367)(3.0)(1+0.6)(0.3)/(3x1 0 8 )

= 9.6x10-7 N• m = 6.6 x10-6 N • m
Atmospheric Constant for Similar to SRP. The ram face will Similar assumptions as for SRP, except
drag Earth-oriented; always be the same, so the cp and cm that being Earth-pointing, the same face
variable for locations can be designed a bit better in will be presented to the ram direction all
inertially this direction. the time, so we can expect a little more
oriented control over the cp and cm locations.

Ar = 1.3 m2; cp a–cm = 0.05 m; C d = 2.0 Ar = 3 m2; cp a–cm = 0.2 m; C d = 2.0

For 700 km orbit: For 21,000 km orbit:
p & 10-13 kg/m3; V = 7504 m/s p & 10-18 kg/m3; V = 3816 m/s

T a= (0.5)(10 -13 )(2.0)(1.3)(7504) 2 (0.05) T a=(0.5)(10 -18 )(2.0)(3)(3816) 2 (0.2)

= 3.7x10-7 N• m = 8.7x10-12 N• m
Magnetic Cyclic 55 deg inclination orbit; assume 0.5 Equatorial orbit; assume 1 A • m2 for a
field A• m2 for a very small, uncompensated small, uncompensated vehicle.
vehicle.
R = (6,378 + 21,000) km = 27,378 km
R = (6,378 + 700) km = 7,078 km D = 1 A• m2 ; X & 1.2 for equatorial orbit
D = 0.5 A• m2 ; X = 1.9 for 55 deg B =(1)(7.8x10 15 )(1.2)/(2.7378x10 7) 3
B = (7.8x10 15 )(1.9)/(7.078x10 6 ) 3 = 4.6x10-7 N • m
= 4.2x10-5 N • m T m = 1(B) = 4.6x10-7 N • m
T m = (0.5)B = 2.1x10 -5 N• m
Gravity Constant for With body-mounted arrays, FireSat II is Deployed solar arrays dominate moment
gradient Earth-oriented; fairly symmetric and small, so the of inertia, so I x = I z > I y . However, the
cyclic for moment of inertia can be balanced very ability to balance mass may be limited
inertially well: We’ll set 0 =1 deg. by the need to fit 3 satellites on the same
oriented. launch vehicle, so we’ll assume a large
R = 7,078 km difference between the geometric and
I = 25 kg• m2 ; Iy = 50 kg• m2 principal axes: 0 = 10 deg.

Normal mode: 0 = 1 deg R = 27,378 km

Tg = (3)(3.986x10 14 )|25-50|sin(2 deg) I z = 90 kg• m2 ; I y = 60 kg • m2
(2)(7.078x10 6) 3
= 1.5x10-6 N• m T g = (3)(3.986x10 14 )|90-60|sin(20 deg)
 (2)(2.7378x10 7 ) 3
Target-of-opportunity: 0 = 30 deg = 3.0x10-7 N • m
T g= 3.7x10-5 N• m

Remaining significant disturbances on the control Likewise, momentum wheel friction torques can be
system are internal to the spacecraft. Fortunately, we compensated in either a closed-loop or a compensatory
have some control over them. If we find that one is fashion; some reaction wheels are designed with
much larger than the rest, we can specify tighter values friction compensation included in some commanding
for that item. This change would reduce its significance modes. Liquid slosh and operating machinery torques
but most likely add to its cost or weight. Table 19-5 are of greater concern but depend on specific hardware.
summarizes the common internal disturbances. If a spacecraft component has fluid tanks or rotating
Misalignments in the center of gravity and in thrusters machinery, the system designer should investigate
will show up during thrusting only and are corrected in disturbance effects and ways to compensate for the
a closed-loop control system and through on-orbit disturbance, if required. Standard techniques include
calibration of the thrusters. propellant management devices (e.g. slosh baffles) or
counter-rotating elements.

Table 19-5 Principal internal disturbance torques. Spacecraft designers can minimize internal disturbances
through careful planning and precise manufacturing, which may increase cost.
Disturbances Effect on Vehicle Typical Values
Uncertainty in Center of Unbalanced torques during firing of 1-3 cm
Gravity (cg) couples thrusters
Unwanted torques during translation
thrusting
Thruster Misalignment Same as cg uncertainty 0.1-0.5 deg
Mismatch of Thruster Similar to cg uncertainty +/- 5%
Outputs
Reaction Wheel Friction Resistance that opposes control torque Roughly proportional to wheel speed,
and Electromotive Force effort. These torques are the limiting depending on model. At top speed, 100%
(i.e. back EMF) mechanism for wheels speed. of control torque (i.e. saturation)
Rotating Machinery Torques that perturb both stability and Dependent on spacecraft design; may be
(pumps, filter wheels) accuracy compensated by counter-rotating
elements
Liquid Slosh Torques due to liquid dynamic pressure Dependent on specific design; may be
on tank walls, as well as changes in cg mitigated by bladders or baffles
location.
Dynamics of Flexible Oscillatory resonance at Depends on spacecraft structure; flexible
Bodies bending/twisting frequencies, limiting frequencies within the control bandwidth
control bandwidth must be phase-stabilized, which may be
undesirable.
Thermal Shocks (“snap”) Attitude disturbances when Depends on spacecraft structure. Long
on Flexible Appendages entering/leaving umbra inertia booms and large solar arrays can
cause large disturbances.

19.1.3 Selection of Spacecraft Control Methods different modes of operating the spacecraft have
significantly different requirements or result in different
Now that we understand the requirements on the control disturbance profiles (as we will see in our FireSat II
system and the environment in which it will operate, we example). Table 19-6 lists several methods of control,
can select one or more methods of controlling the along with typical characteristics of each.
spacecraft. Multiple methods may be indicated when
Passive Control Techniques. Gravity-gradient control disturbances around the third axis (the spin axis) or else
uses the inertial properties of a vehicle to keep it have active means of correcting these disturbances.
pointed toward the Earth. This relies on the fact that an
elongated object in a gravity field tends to align its The vehicle is stable (in its minimum energy state) if it
longitudinal axis through the Earth’s center. The is spinning about the principal axis with the largest
torques that cause this alignment decrease with the cube moment of inertia. Energy dissipation mechanisms
of the orbit radius and are symmetric around the nadir onboard, such as propellant slosh, structural damping,
vector. Thus, the yaw of a spacecraft around the nadir or electrical harness movement, will cause any vehicle
vector is not controllable by this method. This to progress toward this state if uncontrolled. So, disk-
technique is used on simple spacecraft in near-Earth shaped vehicles are passively stable whereas pencil-
orbits without yaw orientation requirements, often with shaped vehicles are not. Spinners can be simple, survive
deployable booms to achieve the desired inertias. for long periods without attention, provide a thermally
benign environment for components (because of even
Frequently, we add dampers to gravity-gradient heating), and provide a scanning or sweeping motion
satellites to reduce libration—small oscillations off of for sensors. The principal disadvantages of spin
the nadir vector caused by other environmental stabilization are that the vehicle mass properties must
disturbances. For example, long deployed booms are be carefully managed during vehicle design and
particularly susceptible to thermal shocks when assembly to ensure the desired spin direction and
entering or leaving umbra. These spacecraft also need a stability, and that the angular momentum vector that
method of ensuring attitude capture with the correct end provides stability also limits maneuverability. More
pointed at nadir—the gravity-gradient torques stabilize fuel is required to reorient the vehicle than a vehicle
either end of the minimum inertia axis equally. with no net angular momentum, making this technique
less useful for payloads that must be repointed
In the simplest gravity-gradient spacecraft, only two frequently.
orientation axes are controlled; the orientation around
the nadir vector is unconstrained. To control this third A spinner requires extra fuel to reorient because of the
degree of freedom, a small constant-speed momentum gyroscopic stiffness, which also helps it resist
wheel is sometimes added along the intended pitch axis. disturbances. In reorienting a spinning body with
The momentum-biased wheel will be most stable when angular momentum, H, a constant torque, T, will
perpendicular to the nadir and velocity vectors, and produce an angular velocity, w, perpendicular to the
therefore parallel to the orbital momentum vector. The applied torque and angular momentum vector, of
stable state of the gravity-gradient plus momentum bias magnitude w=T/H. (This follows from the earlier-
wheel establishes the desired attitude through small introduced Euler equations.) Thus, the greater the
energy dissipations onboard without the need for active stored momentum, the more torque must be applied for
control. a given w. For a maneuver through an angle O, the
torque-time product—an indication of fuel required for
A third type of purely passive control uses permanent the maneuver—is a constant equal to H. O. Alternately,
magnets onboard to force alignment along the Earth’s for a vehicle with no initial angular momentum, a small
magnetic field. This is most effective in near-equatorial torque can be applied to start it rotating, with an
orbits where the North-South field orientation is opposite small torque to stop it when it has reached its
reasonably constant for an Earth-referenced satellite. new target. The fuel used for any angle maneuver can
be arbitrarily small if a slow maneuver is acceptable.
Spin Control Techniques. Spin stabilization is a (Note that the spinner can only be maneuvered
passive control technique in which the entire spacecraft relatively slowly; a fast slew is usually not an option.)
rotates so that its angular momentum vector remains
approximately fixed in inertial space. Spin-stabilized
spacecraft (or spinners), employ the gyroscopic
stability discussed earlier to passively resist disturbance
torques about two axes. Additionally, spinners are
generally designed to be either insensitive to

Table 19-6. Attitude control methods and their capabilities. As requirements become tighter, more complex
control systems become necessary.

Type Pointing Options Attitude Typical Accuracy Lifetime Limits

 Maneuverability
Gravity-Gradient Earth local vertical Very limited ±5 deg (2 axes) None
only
Gravity-Gradient Earth local vertical Very limited ±5 deg (3 axes) Life of wheel bearings
+ Momentum only
Bias
Passive Magnetic North/South only Very limited ±5 deg (2 axes) None
Rate-Damping + Usually Sun Generally used as ± 5-15 deg (2 axes) None
Target Vector (power) or Earth robust safe mode.
Acquisition (communication)
Pure Spin Inertially fixed Repoint with ± 0.1 deg to ± 1 deg Thruster propellant (if
Stabilization any direction precession in 2 axes applies)*
maneuvers; very (proportional to
slow with torquers, spin rate)
faster with thrusters
Dual-Spin Limited only by Same as above Same as above for Thruster propellant (if
Stabilization articulation on spun section. applies)*
despun platform Despun dictated Despun section bearings
by payload
reference and
pointing
Bias Momentum Local vertical Fast maneuvers ± 0.1 deg to ± 1 deg Propellant (if applies)*
(1 wheel) pointing or inertial possible around Life of sensor and wheel
targets momentum vector bearings
Repoint of
momentum vector as
with spin stabilized
Active Magnetic Any, but may drift Slow (several orbits ±1 deg to ± 5 deg Life of sensors
with Filtering over short periods to slew); faster at (depends on
lower altitudes sensors)
Zero Momentum No constraints No constraints ± 0.1 deg to 5 deg Propellant
(thruster only) High rates possible
Zero momentum No constraints No constraints ± 0.0001 deg to ± 1 Propellant (if applies)*
(3 wheels) deg (determined Life of sensors and wheel
by sensors and bearing
processor)
Zero Momentum No constraints No constraints ± 0.001 deg to ± 1
Propellant (if applies)*
(CMG) Short CMG life High rates possible deg Life sensors and CMG
may require high bearings
redundancy
• Thrusters may be used for slewing and momentum dumping at all altitudes, but propellant usage may be high.
Magnetic torquers may be used from LEO to GEO.

A useful variation on spin control is called dual-spin Also, by adding energy dissipation devices to the
stabilization, in which the spacecraft has two sections platform, a dual spinner can be passively stable
spinning at different rates about the same axis; this kind spinning about the axis with the smallest moment of
of spinner is also known as a gyrostat. Normally one inertia, as long as the rotor is spinning about its own
section, the rotor, spins rapidly to provide gyroscopic maximum moment of inertia. This permits more pencil-
stiffness, while the second section, the stator or shaped spacecraft, which fit better in launch vehicle
platform, is despun to keep one axis pointed toward the fairings and which would not normally be stable
Earth or Sun. By combining inertially fixed and rotating spinning about their long axes. The disadvantage of
sections on the same vehicle, dual spinners can dual-spin stabilization is the added complexity of the
accommodate a variety of payloads in a simple vehicle. platform bearing and slip rings between the sections.
(Slip rings permit power and electrical signals to flow adjust the spacecraft’s angular momentum vector to
between the two sections.) This complexity can counteract disturbance torques. In addition, we may
increase cost and reduce reliability compared to simple need to damp the nutation caused by disturbances or
spin stabilization. precession commands. Aggravating this nutation is the
effect of energy dissipating phenomena like structural
Spinning spacecraft, both simple and dual, exhibit flexure and flexible harness or fluid motion, which are
several distinct types of motion that are often confused. present in any spacecraft to some degree. Once the
Precession is the motion of the angular momentum excitation stops, nutation decreases as these same
vector caused by external torques, including thruster factors dissipate the kinetic energy added by the control
firings used to correct environmental disturbances. effort. However, this natural damping can take hours.
Coning (or wobbling) is the apparent motion of the We can neutralize this error source in minutes with
body when it is spinning about a principal axis of nutation dampers. We can also reduce the amount of
inertia that is not aligned with a body reference axis or nutation from these sources by increasing the spin rate
axis of symmetry—for example, the intended spin axis. and thus the stiffness of the spinning vehicle. If the spin
Coning looks like motion of the intended spin axis rate is 20 rpm and the nutation angle from a given
around the angular momentum vector at the spin rate. disturbance is 3 deg, then nutation from the same
Figure 19-7 shows various natural rotations. disturbance would be reduced to 1 deg if the spin rate
were 60 rpm. We seldom use spin rates above 90 rpm
because of the large centripetal forces demanded of the
structure and the consequent effects on design and
weight. In thrusting and pointing applications, spin
rates under 20 rpm are generally not used as they may
allow excessive nutation. However, applications
unrelated to attitude control, such as thermal control or
specialized payload requirements, are frequently less
Figure 19-7. Types of Rotational Motion. H = angular
sensitive to nutation and may employ very low spin
momentum vector; P = principal axis; w =
rates.
instantaneous rotation axis; Z = geometrical axis.
Three-axis Control Techniques. Spacecraft stabilized
Nutation is the torque-free motion of a simple rigid
in all three axes are much more common today than
body when the angular momentum vector is not
those using spin or gravity gradient stabilization. They
perfectly aligned with a principal axis of inertia. For
can maneuver relatively easily and can be more stable
rod-shaped objects, this motion is a slow rotation
and accurate, depending on their sensors and actuators,
(compared to spin rate) of the spin axis around the
than more passive stabilization techniques. They are,
angular momentum vector. For these objects spinning
however, more expensive. They are also often more
about a minimum inertia axis, additional energy
complex, but processor and reliability improvements
dissipation will cause increased nutation. For disk-
have allowed comparable or better total reliability than
shaped objects, spinning around a maximum inertia
some more passive systems. For critical space
axis, nutation appears as a tumbling rotation faster than
applications, there is no replacement for thorough risk
spin rate. Energy dissipation for these objects reduces
and reliability assessment (see Chapter 24).
nutation, resulting eventually in a clean spin. For these
reasons, minimum-axis (or minor-axis) spinners are
The control torques about the axes of three-axis systems
often concerned with minimizing energy dissipation,
come from combinations of momentum wheels,
whereas maximum-axis (or major-axis) may actually
reaction wheels, control moment gyros, thrusters, solar
include mechanisms, such as a passive nutation damper,
or aerodynamic control surfaces (e.g. tabs), or magnetic
to dissipate energy quickly.
torquers. Broadly, however, these systems take two
forms: one uses momentum bias by placing a
Nutation is caused by disturbances such as thruster
momentum wheel along the pitch axis; the other is
impulses, and can be seen as varying signals in body-
called zero momentum and does not use momentum
mounted inertial and external sensors. Wobble is caused
bias at all—any momentum bias effects are generally
by imbalances and appears as constant offsets in body-
regarded as disturbances. Either option usually needs
mounted sensors. Such constant offsets are rarely
some method of angular momentum management, such
discernable unless multiple sensors are available.
as thrusters or magnetic torquers, in addition to the
primary attitude actuators.
Spin stability normally requires active control, such as
cold-gas thrusters or magnetic torquers, to periodically
In a zero-momentum system, actuators such as reaction dumping or provision of small delta-Vs during other
wheels or thrusters respond to disturbances on the mission phases.
vehicle. For example, an attitude error in the vehicle
results in a control signal that torques the wheel, Momentum bias systems often have just one wheel with
creating a reaction torque in the vehicle that corrects the its spin axis mounted along the pitch axis, ideally
error. The torque on the wheel either speeds it up or normal to the orbit plane. The wheel is run at a nearly
slows it down; the aggregate effect is that all constant, high speed to provide gyroscopic stiffness to
disturbance torques are absorbed over time by the the vehicle, just as in spin stabilization, with similar
reaction wheels, sometimes requiring the collected nutation dynamics. Around the pitch axis, however, the
angular momentum to be removed. This momentum spacecraft can control attitude by torquing the wheel,
removal—called desaturation, momentum dumping, or slightly increasing or decreasing its speed. Periodically,
momentum management—can be accomplished by the momentum in the pitch wheel must be managed (i.e.
thrusters or magnetic torquers acting automatically or brought back to its nominal speed), as in zero-
by command from the ground. momentum systems, using thrusters, magnets, or other
means.
When high torque is required for large vehicles or fast
slews, a variation of three-axis control using control The dynamics of nadir-oriented momentum-bias
moment gyros, or CMGs, is available. These devices vehicles exhibit a phenomenon known as roll-yaw
work like momentum wheels on gimbals. The control coupling. To understand this coupling, consider an
of CMGs is complex and their lifespan is limited, but inertially fixed angular momentum vector at some error
their available torque for a given weight and power can angle with respect to the orbit plane. If the angle is
make them attractive. initially a positive roll error, then a quarter-orbit later it
appears as a negative yaw error with no roll component
A very specialized form of zero momentum control, remaining. As the vehicle continues around the orbit,
here called active magnetic control, can be attained the angle goes through negative roll and positive yaw
from a combination of a magnetometer, a Global before regaining its positive roll character. This
Positioning System (GPS) receiver, and coupling (or commutation) is due to the apparent
computationally intensive software filtering. The GPS motion of the Earth-fixed coordinate frame as seen
feeds the spacecraft location to the onboard processor, from the spacecraft, and it can be exploited to control
which then determines the local magnetic field based on roll and yaw over a quarter orbit using only a roll (or
onboard models. The magnetometer data is filtered only a yaw) sensor, instead of needing one sensor for
using the Euler equations to determine the attitude, and each of the roll and yaw axes.
magnetic torquers make corrections in the two available
directions at any given moment—corrections about the Effects of Requirements on Control Type. With the
magnetic field vector are not possible. Active magnetic above knowledge of control techniques, we can proceed
control can be an inexpensive backup control mode for to select a control type that will best meet mission
a LEO satellite, or it can be a primary control mode for requirements in the expected operational environment.
a satellite in a highly inclined orbit. (The highly Tables 19-7 and 19-8 describe the effects of orbit
inclined orbit has large changes in magnetic field insertion and payload slew requirements on the
direction, allowing the filtering algorithm to better selection process. It is also useful here to once again
determine a three-axis attitude solution.) This attitude reference Figure 19-6 for information on how altitude
knowledge can also be combined with other sensors, can affect the space environment. Certain control types,
such as Sun sensors, for more accuracy. While this is such as gravity-gradient stabilization or active magnetic
not a common control method, we include it here as an damping, are better in some orbits than in others.
example of how increased onboard computational
power and the presence of new resources, such as the A common control approach during orbit insertion is to
GPS constellation, can allow completely new methods use the short-term spin stability of the combination of
of attitude determination and control. spacecraft and orbit-insertion motor. Once on station
the motor may be jettisoned, the spacecraft despun
As a final demonstration of zero momentum three-axis using thrusters or a yo-yo device, and a different control
control, simple all-thruster systems are used for short technique used from that point on.
durations when high torque is needed, such as during
orbit insertion maneuvers or other orbit adjustments Payload pointing will influence the attitude control
(delta-V) from large motors. These thrusters then may method, the class of sensors, and the number and kind
be used for different purposes such as momentum of actuation devices. Occasionally, pointing accuracies
are so stringent that a separate, articulated platform is
necessary. An articulated platform can perform most stringent requirements will ultimately drive ADCS
scanning operations much more easily than the host component selection. Table 19-9 summarizes the
vehicle and with better accuracy and stability. Trade effects of accuracy requirements on the ADCS
studies on pointing requirements must consider approach for the spacecraft. Section 14.5 discusses how
accuracy in attitude determination and control, and the to develop pointing budgets.

Table 19-7. Orbit Transition Maneuvers and Their Effects. Using thrusters to change orbits creates special
challenges for the ADCS.
Requirement Effect on Spacecraft Effect on ADCS
Large impulse to complete orbit Solid motor or large bipropellant Inertial measurement unit for
insertion (thousands of m/s) stage. accurate reference and velocity
Large thrusters or a gimbaled engine measurement.
or spin stabilization for attitude Different actuators, sensors, and
control during burns. control laws for burn vs. coasting
phases.
Need for navigation or guidance.
On-orbit plane changes to meet Large thrusters needed, but these Separate control law for thrusting.
payload needs or vehicle operations thrusters may be needed for other Actuators sized for thrusting
(hundreds of m/s) reasons also, such as orbit insertion, disturbances (possibly two sizes of
coasting phase, or stationkeeping. thruster).
Onboard attitude reference for
thrusting phase.
Orbit maintenance/trim maneuvers One set of thrusters Thrusting control law.
(<100 m/s) Onboard attitude reference.

Table 19-8. Slewing Requirements That Affect Control Actuator Selection. Spacecraft slew agility can demand
larger actuators for intermittent use.
Slewing Effect on Spacecraft Effect on ADCS
None or Time- Spacecraft constrained to one - Reaction wheels, if planned, can be
Unconstrained attitude (highly improbable), or smaller
reorientations can take many hours. - If magnetic torquers can dump
momentum, reaction control thrusters may
not be needed
Low Rates Minimal - Depending on spacecraft size, reaction
From 0.05 deg/s wheels may be fully capable for slews
(orbital rate) to 0.5 deg/s - If reaction wheels not capable, thrusters
will be necessary
- Thrusters may be needed for other
reasons; i.e. stationkeeping
High Rates - Structural impact on appendages - Control moment gyros or thrusters needed.
>0.5 deg/s - Weight and cost increase If thrusters needed for other reasons, two
thrust levels may be needed.
Table 19-9. Effects of Control Accuracy Requirements on Sensor Selection and ADCS Design. More accurate
pointing requires better and more expensive sensors and actuators.
Required Effect on Spacecraft Effect on ADCS
Accuracy
(3σ)
>5 deg • Permits major cost savings Without attitude determination
• Permits gravity-gradient (GG) • No sensors required for GG stabilization
stabilization • Boom motor, GG damper, and a bias momentum wheel are
only required actuators
With attitude determination
• Sun sensors & magnetometer adequate for attitude
determination at >_ 2 deg
 • Higher accuracies may require star trackers or horizon
sensors
1 deg to • GG not feasible • Sun sensors and horizon sensors may be adequate for
5 deg • Spin stabilization feasible if stiff, sensors, especially a spinner
inertially fixed attitude is • Accuracy for 3-axis stabilization can be met with RCS
acceptable deadband control but reaction wheels will save propellant
• Payload needs may require despun for long missions
platform on spinner • Thrusters and damper adequate for spinner actuators
• 3-axis stabilization will work • Magnetic torquers (and magnetometer) useful
0.1 deg to • 3-axis and momentum-bias • Need for accurate attitude reference leads to star tracker or
1 deg stabilization feasible horizon sensors & possibly gyros
• Dual-spin stabilization also feasible • Reaction wheels typical with thrusters for momentum
unloading and coarse control
• Magnetic torquers feasible on light vehicles (magnetometer
also required)
< 0.1 deg • 3-axis stabilization is necessary • Same as above for 0.1 deg to 1 deg but needs star sensor
• May require articulated & and better class of gyros
vibration-isolated payload platform • Control laws and computational needs are more complex
with separate sensors • Flexible body performance very important

FireSat II Control Selection. For FireSat II, we consider For the optional off-nadir pointing requirement, three-
two options for orbit insertion control. First, the launch axis control with reaction wheels might be more
vehicle may directly inject the spacecraft into its appropriate. Mass and power will be especially precious
mission orbit. This common option simplifies the for this very small satellite, so carrying and running 3
spacecraft design since no special insertion mode is reaction wheels simultaneously may be more than the
needed. An alternate approach, useful for small system budgets can handle. Since the SCS will provide
spacecraft such as FireSat II, is to use a monopropellant an example of three-axis control with reaction wheels,
propulsion system onboard the spacecraft to fly itself up we will use the magnetic torque and momentum wheel
from a low parking orbit to its final altitude. For small combination for FireSat II, and for actuator sizing (table
insertion motors, reaction wheel torque or momentum 19-11) we will assume the 30 degree offset pointing is
bias stabilization may be sufficient to control the done about the pitch axis, so that the momentum wheel
vehicle during this burn. For larger motors, delta-V performs that slew also. If offset pointing is needed
thruster modulation or dedicated attitude control about roll, thrusters will have to be used; the thruster
thrusters become attractive. force sizing example in Table 19-12 shows how this
might work with the momentum wheel stopped.
Once on station the spacecraft must point its sensors at
nadir most of the time and slightly off-nadir for brief SCS Control Selection. For the Supplemental
periods. Since the payload needs to be despun and the Communication System, we will focus on taking
spacecraft frequently reoriented, spin stabilization is not advantage of the gentle disturbance environment and 1
the best choice. Gravity-gradient and passive magnetic deg accuracy requirement to design a light, inexpensive
control cannot meet the 0.1-deg pointing requirement or attitude system that can be installed in all 3 satellites.
the 30 deg slews. This leaves three-axis control and At the SCS altitude of 21,000 km, with the
momentum-bias stabilization as viable options for the configuration we’ve assumed, there are no good passive
on-station control. stabilization methods available to us. Reaction wheels
will be needed to reject the disturbances, which are
The periodic 180-deg yaw slews (i.e. slews around the dominated by the solar radiation pressure torques (see
nadir vector) would have to overcome any momentum Table 19-4). So, as long as the reaction wheels are there
bias in the Y axis (perpendicular to the orbit normal). anyway, we might as well use them for three-axis
When we do the example momentum bias sizing in stabilization. Also, three-axis control often can be
Table 19-11 later, we will see that a bias of exploited to simplify the solar array design by using
approximately 20 N ⋅ m⋅ s would be required to maintain one of the unconstrained payload axes (yaw, in this
the required 0.1 deg accuracy. The magnetic torquers case) to replace a solar array drive axis. Thus, the
could not perform this slew in 45 minutes because of reduced array size possible with 2 degrees of freedom
the bias to be overcome. Instead, we will use the can be achieved with one array axis drive and one
propulsion system to perform the yaw slew. spacecraft rotation.
While the greater part of the SRP torques will be cyclic, they are activated: magnetic torquers & thrusters (cold-
some small part will be secular. Therefore, the gas, hot-gas and electric) are the most commonly used
momentum stored in the reaction wheels will gradually in this category.
increase and will need to be removed periodically by
thrusters. The use of 100 A • m2 magnetic torquers for Wheel control provides smooth changes in torque,
momentum removal is feasible from the point of view allowing very accurate pointing of spacecraft. Some
of the available magnetic field. However, the wheels can cause vibrations, or jitter, at high speeds,
propulsion system has to be included anyway to but this can often be mitigated with vibration isolators
separate the 3 satellites and establish the constellation, or changes in structural design. Reaction wheels are
and constantly running torquers would be an additional essentially torque motors with high-inertia rotors. They
power drain. Since power is a major challenge for this can spin in either direction and provide one axis of
mission, we will use the thrusters for momentum control for each wheel. Momentum wheels are reaction
unloading. wheels with a nominal spin rate above zero to provide a
nearly constant angular momentum. This momentum
provides gyroscopic stiffness to two axes, and the
19.1.4 Selection and Sizing of ADCS Hardware motor torque may be controlled to change pointing
around the spin axis. In sizing wheels we must always
We are now ready to evaluate and select the individual consider two performance quantities: angular
ADCS components. For all ADCS hardware, we will momentum capacity, and torque authority.
determine the minimum performance level needed to
meet requirements. Then, standard components To determine the necessary momentum capacity, we
available from manufacturers will be selected if must distinguish between cyclic and secular
possible, sometimes resulting in better performance disturbances in the spacecraft’s environment. We
than the minimum required. If standard components are typically size reaction wheels to be able to store the full
not available, specialized components may be designed cyclic component of momentum without the need for
and built, but this can often be prohibitively costly for frequent momentum dumping. Therefore, the average
most agencies, and so a revision of the requirements is disturbance torque for 1/4 or 1/2 an orbit determines the
more likely to be more in line with available hardware. minimum capacity of the wheels. The secular
component of momentum will also need to be stored for
the amount of time the spacecraft must be operational
Actuators. Options for actuator selection are without a momentum dump being performed. This time
summarized in Table 19-10. First, we will discuss may be determined by requirements on payload
momentum-exchange devices, which conserve angular observation continuity, or it may be the amount of time
momentum in the spacecraft: reaction wheels, the spacecraft must survive without ground
momentum wheels, and control moment gyros. Then, intervention.
we will move on to external torque actuators, which
change the angular momentum of the spacecraft when

Table 19-10. Typical Attitude Actuators. Actuator weight and power usually scale with performance.

Actuator Typical Performance Range Mass (kg) Power (W)

Thrusters Thrusters produce force; multiply by moment arm for torque Mass and power vary
Hot Gas (Hydrazine) 0.5 to 9000 N (See Ch. 18 for details on
Cold Gas <5N propulsion systems)
Reaction and Maximum torques: 0.01 to 1 N • m Varies with speed:
2 to 20
Momentum Wheels Practical momentum storage capacity: 0.4 to 3000 N• m•s 10 to 100
Control Moment
Max. torques: 25 to 500 N • m > 10 90 to 150
Gyros (CMG)
Magnetic Torquers 1 to 4,000 A • m2 (See Table 19-4 for torque examples) 0.4 to 50 0.6 to 16
The necessary torque authority of the reaction wheels is For three-axis control, at least three wheels with non-
defined either by slew requirements or the need for coplanar spin axes are required. Often, a fourth wheel is
control authority to be well above the peak disturbance carried in case one of the primary wheels fails. If the
torque. If the peak disturbance is too close to the torque wheels are not orthogonal (and a redundant wheel never
authority available, pointing accuracy may suffer. is), additional torque authority or momentum capacity
may be necessary to compensate for non-optimal
geometry. However, the redundancy of the fourth wheel torquer produces a torque proportional and
can have additional benefits, such as being able to avoid perpendicular to the Earth’s varying magnetic field;
any wheel speed passing through zero (which can cause therefore, at any given moment, no torque can be
attitude error transients) or even as power storage, as provided about the Earth’s magnetic field vector.
driving a spinning wheel toward zero speed will However, as a spacecraft changes in latitude or altitude
provide power to the spacecraft. while following its orbit, different directions of the field
become available.
For spin-stabilized or momentum-bias systems, the
cyclic torques will cause cyclic variations of the Electromagnets have the advantage of no moving parts,
attitude, whereas the secular torques cause gradual requiring only a magnetometer for sensing and wire
divergence of the attitude away from the ideal target. coiled around a metallic rod in each axis. (Some
We typically design the stored angular momentum, torquers are even internally redundant, as they have two
determined by spin rate and inertia of the spinning coils around the same metallic core.) Because they use
body, to be large enough to keep the cyclic motion the Earth’s natural magnetic field, and this field reduces
within accuracy requirements without active control. in strength with the cube of distance from the Earth’s
Periodic torquing will still be needed to counteract center, magnetic torquers are less effective at higher
secular disturbances, and the cost of this torquing in orbits. We can specify the torquer’s dipole strength in
propellant, if performed by thrusters, will constrain the A• m2 and tailor it to any application. Table 19-11
maximum stored momentum. describes sizing rules of thumb for wheels and magnetic
torquers, and works through those rules for the FireSat
For high-torque applications in which fine control is II and SCS examples. SCS has no slewing
still needed, control moment gyros (CMG) may be used requirements, so we size its reaction wheels based only
instead of reaction wheels. CMGs are single- or double- on momentum storage needs.
gimbaled wheels spinning at constant speed (and
therefore providing momentum bias stiffness when not Thrusters (i.e. rocket engines) are possibly the most
actuating). By turning the gimbal axis, we can obtain a frequently flown attitude actuator because of their dual
high-output torque whose size depends on the speed of use in adjusting orbital parameters. Almost every
the rotor and the gimbal rate of rotation. Control spacecraft that needs to perform orbital maneuvers will
systems with two or more CMGs can produce large use thrusters to achieve that goal, and in many cases,
torques about all three of the spacecraft’s orthogonal some subset of the thrusters used is for attitude control.
axes, so we most often use them where agile (i.e. high Thrusters produce a force on the spacecraft by expelling
angular rate) maneuvers are required. The use of CMGs material, called propellant, at high velocity from their
requires complex control laws and careful momentum exit nozzles. Hot-gas propulsion systems include
management to avoid wheel saturation. Also, because thrusters that chemically alter the propellant to extract
the CMG torque is created by twisting what is the energy needed for rapid mass expulsion. These
essentially a stiff gyroscope perpendicular to its spin systems may be monopropellant, in which the
axis, the bearings of the wheel suffer a great deal of propellant is catalyzed to break down chemically, or
wear and tear, causing most CMGs to have shorter bipropellant, in which a fuel is mixed with an oxidizer
lifetimes than other actuators. Because CMGs combine to achieve combustion just prior to expulsion. Cold-gas
a short life with high cost, weight, and power needs, systems include thrusters whose propellant is not
they are generally used only on very large spacecraft altered chemically during propulsion. In a cold-gas
and only when absolutely necessary to achieve the system, the energy may come from phase change of the
mission goals. propellant, or simply from pre-pressurizing the
propellant in its tank. A third type of thruster is
Spacecraft also use magnetic torquers as actuation electrical, which, due to the usually small forces
devices. These torquers are magnetic coils involved, is only used for attitude control in special
(electromagnets) designed to generate magnetic dipole circumstances. Electrical propulsion is accomplished by
moments of commanded magnitude. When three using magnetic or electrostatic fields to eject plasma or
orthogonal torquers are mounted to a spacecraft, they magnetic fluid to achieve a reaction force on the
can create a magnetic dipole of any direction and spacecraft.
magnitude up to the strength of the torquers. Magnetic
torquers can compensate for a spacecraft’s residual Thrusters provide torque proportional to their moment
magnetic field or attitude drift from minor disturbance arm, which is the amount that their line of force is
torques. They also can be used to desaturate offset from the vehicle’s center of mass. When thrusters
momentum-exchange systems, though they usually apply their force along a line that intersects the center
require much more time than thrusters. A magnetic of mass of the vehicle, there is no torque. So, while the
amount of force available from a thruster may be large, degree slews that must be done periodically to keep the
the torque available is limited by the physical extent of Sun off the radiator. Because the slew must be about
the vehicle and how the thrusters are mounted on the the yaw axis, the momentum bias in the wheel would
vehicle. When mounted to maximize torque authority, resist the thruster torque, complicating the control
thrusters have the advantage of being able provide design. Instead, we will use the thrusters to remove the
large, instantaneous control torques at any time in the momentum bias, rotate the spacecraft, and then re-
orbit. The disadvantages of thrusters are that they use establish the momentum bias. One way to remove the
expendable propellant, can disrupt orbit determination wheel momentum is to use thrusters for three-axis
activities, and that the plumes of expelled matter can control while driving the wheel to zero speed. As the
impinge on the surface of the spacecraft, possibly wheel slows down, it causes the spacecraft to rotate,
heating or contaminating surfaces. and the thrusters are fired to counteract the rotation; the
same thing happens when the wheel is once again sped
Attitude functions to which thrusters can be applied up to re-establish the momentum bias. The thrusters
include controlling attitude, controlling nutation and must provide much more torque than the wheel, but
spin rate, performing large, rapid slews, and managing generally the attitude control accuracy requirements are
angular momentum. In all of these functions, large loosened for this kind of maneuver to avoid firing the
torque authority can be helpful, but also necessary to thrusters more than is necessary. Good accuracy is less
some extent is the ability to change angular momentum important here because science data would not be
by relatively small amounts. So, another consideration collected anyway during the slew. The total fuel mass
in selection of thrusters is how fine and how precise the needed for the momentum changes and the slew
delivered torque impulse needs to be. Torque applied depends on the specific impulse, I sp, of the selected
over a period time creates a change in angular thruster/propellant combination; we will assume the I sp
momentum, which has a corresponding change in the is 218 seconds.
rotational rate of the spacecraft. As attitude changes
with non-zero rate, attitude control accuracy is directly SCS Actuators. The SCS satellites will need to use
determined by the minimum thruster impulse available. thrusters for orbit maintenance and momentum
If a 50 N thruster with a 1 meter moment arm must dumping. We will want to minimize the number and
provide a change in angular momentum accurate to size of thrusters, since the spacecraft are so small.
within 1 N⋅ m⋅ s, then it will be necessary to fire the However, the lower limit will most likely come from
thruster for a period of time no greater than 20 msec. the orbital maneuvering needs, and not the attitude
The thruster valves must be capable of opening and control needs. We will suppose 4 1-lb thrusters are
closing that rapidly and precisely, and of propellant needed for orbital maneuvering and steering, and in
actually flowing into the thruster and out the nozzle, Table 19-12, we will present and work through
which takes a finite amount of time. This need for procedures and simplified equations for sizing thrusters
speed and precision in engine valve control also has using SCS as an example. SCS has no stated slew
implications in the design of the avionics that drive the requirement, but for the purposes of the example, a
thruster valves; often, special electronics cards must be slew requirement of 30 deg in 60 seconds will be
developed for a given mission to meet thruster-based supposed. A thorough discussion on the topic of
attitude control requirements. propulsion systems, including estimating propellant
needs, can be found in Chapter 18.
FireSat II Actuators. The baseline FireSat II
spacecraft will use magnetic torquers for momentum
dumping, but thrusters will be needed for the 180

TABLE 19-11. Simplified Equations for Sizing Reaction Wheels, Momentum Wheels, and Magnetic
Torquers. The FireSat II momentum wheel is sized for the baseline pointing requirements and for the optional
design with 30-deg slew requirements. SCS reaction wheels are sized for momentum storage capacity.

Parameter Simplified Equations Application to FireSat II and SCS Examples

Torque from Control torque, whether from reaction For the FireSat II spacecraft, magnetic torque dominates
Reaction Wheel wheels or magnetic torquers, must during regular observations:
for equal worst-case anticipated T D ≈ Tm = 2.1 × 10 –5 N·m (see Table 19.4).
Disturbance disturbance torque plus some margin:
Rejection T C = (TD)(Margin Factor) This torque is manageable by most reaction wheels and
magnetic torquers. Momentum requirements or slew
torque, not disturbance rejection, will be used for actuator
sizing.
Slew Torque for For max-acceleration slews with no FireSat II needs to make a 30-deg slew about the Y axis
Reaction resisting momentum (1/2 distance in (50 kg•m2 ) in 10 minutes using its momentum wheel:
Wheels 1/2 time):
0 /2 = 1/2 (T/I)(t/2) 2 T = (4)(30 deg)(n/180 deg)(50 kg• m2 )/(600 sec) 2
T = 4 0 I/t2 = 2.9 x 10-4 N• m (This is a small value.)

The total momentum change for the wheel during the

slew is T • t/2 = 0.087 N • m• s.
Momentum One approach to estimating wheel For FireSat II, the maximum gravity gradient torque
Storage in momentum, h, is to integrate the (during 30-deg off-pointing) is estimated at:
Reaction Wheel worst-case disturbance torque, T D , TD = 3.7 x 10 –5 N•m (Table 19-4)
over a full orbit, with period P. For and the 700-km orbital period is 5926 sec:
gravity gradient, the maximum h = (3.7 x 10-5 N• m)(5926 sec)(0.707/4)
momentum accumulates in 1/4 of an = 0.039 N• m• s
orbit. A simplified expression for such If the magnetic torquers were not capable of handling this
a sinusoidal disturbance is: momentum, a small reaction wheel with storage of 0.4
h = T D• P• (0.707)/4 N•m•s would be sufficient even if 30-deg off-pointing
where 0.707 is the rms average of a were held for an entire quarter-orbit.
sinusoidal function.
For SCS, T D=6.6x10-6 & P =45083 sec.
For the SCS solar pressure torques, the SCS storage: h=(6.6x10 -6 )(45083)(0.707/4)=0.06 N • m• s
torque buildup depends heavily on We will size 1.0 N • m• s wheels to minimize thruster burns
attitude and orbit geometry, but to and potential jitter from high wheel speeds.
keep this simple we will assume the
maximum also accumulates in 1/4 orbit.
Momentum For momentum-wheel stabilization, The value of h for the FireSat II yaw accuracy
Storage in roll and yaw accuracy depend on the requirement of 0.1 deg (0.0017 rad) would be
Momentum wheel’s momentum and the external h = (2.1x10-5 N• m/0.0017 rad)• (5926/4 sec)
Wheel disturbance torque. A simplified = 18.4 N • m• s
expression for the required momentum For 1 deg accuracy, we would need only 1.8 N•m•s.
storage is: h = (T/0 a) • (P/4) The same results would hold for a spinner.
where 0 a = allowable motion.
Note that this capacity would need to be added to the
momentum change needed for the 30 deg slews, which
for FireSat II is less than 0.1 N• m• s.
Torque from Magnetic torquers use electrical Table 19-4 estimates the maximum Earth field, B, for
Magnetic current through the torquer to create a FireSat II to be 4.2 x 10 –5 tesla. We calculate the torque
Torquers magnetic dipole (D). The Earth’s rod’s magnetic torquing ability (dipole) to counteract the
magnetic field, B, acts on D, resulting worst-case gravity gradient disturbance, T D , of 2.1 x 10 –5
in torque (T) on the vehicle: N•m•s as
D = T/B D = (2.1x10-5 N • m)/(4.2x10-5 tesla) = 0.5 A• m2
Magnets used for momentum dumping which is a small actuator. The Earth’s field is cyclic at
must equal the peak disturbance + twice orbital frequency; thus, maximum torque is
margin to compensate for the lack of available only twice per orbit. A torquer of 3 to 10 A•m2
complete directional control. (Note capability should provide sufficient margin, which also
that if magnetic torque is the dominant gives enough capacity to remove increased gravity
disturbance, the torquers need mainly gradient torques during offset pointing.
to cancel the spacecraft’s residual
 dipole.)

Note: For actuator sizing, the magnitude and direction of the disturbance torques must be considered. In
particular, momentum accumulation in inertial coordinates must be mapped to body-fixed actuator axes.

TABLE 19-12. Simplified Equations for Preliminary Sizing of Thruster Systems. SCS thruster requirements are
small for this low-disturbance, minimal slew application. It is likely that the thrusters needed for orbit maintenance
can also serve for momentum dumping. FireSat II thrusters are needed for 180 degree slews and for orbit insertion
and maintenance.

Simplified Equations Application to Examples

Thruster force level sizing for For the SCS worst case T D of 6.6 × 10 –6 N·m (Table 19-4) and a thruster
external disturbances: moment arm of 0.5 m
F=T D /L F=(6.6x10- 6 N • m)/(0.5m)=3.3x10- 6 N
This small value indicates orbit maintenance and momentum dumping
F is thruster force, T D is worst-case
requirements, not disturbance torques, will determine thruster size. Also,
disturbance torque, and L is the
using thrusters to fight cyclic disturbances uses precious propellant; it is
thruster’s moment arm
generally better to store the momentum in wheels.

Sizing force level to meet slew rates: Assume FireSat II does a 30-deg slew in less than 1 min (60 sec),
Determine fastest slew rate, w, accelerating for 5% of that time, coasting for 90%, and decelerating for 5%.
required in the mission profile. w = 30 deg / 60 sec = 0.5 deg/sec
Develop a slew profile that accelerates To reach 0.5 deg/sec in 5% of 1 min, which is 3 sec, requires an acceleration
the vehicle quickly, coasts at that rate,
and then decelerates quickly. The a = w /t = (0.5 deg/sec)/(3 sec) = 0.167 deg/sec2 = 0.003 rad/sec 2
acceleration required, a, comes from
equating these two torque definitions: F = Ia /L = (25 kg • m2 )(0.003 rad/sec2)/(0.5 m) = 0.15 N
This is very small. It is certainly within the 5 N of thrust that the propulsion
T = F • L = I• a design selected. However, it is so much smaller that specialized circuitry
would be needed to fire the thrusters for very brief amounts of time, which
may impact thruster efficiency.

Determining fuel needed to Changing momentum bias:

complete FireSat II yaw slew: The minimum thruster on-time needed to remove or create a given
Thrusters will need to remove the momentum, h, is t = h/T = h/(F • L). For FireSat II:
momentum bias from the wheel, rotate tbias = (18.4 N • m• s)/(5 N • 0.5 m) = 7.4 sec
the satellite 180 deg, and then re-
We will bring this up to 10 seconds to account for any additional attitude
establish the momentum. Fuel mass, correction firings needed during wheel torquing.
m, is estimated as:
m = F • t/(g• Isp), where t is the on-time Yaw slew:
of one thruster. We’ll plan on a slew rate of 1 deg/sec to finish the slew itself in under 5
minutes, leaving plenty of time for wheel stopping and starting. 1 deg/sec
amounts to an angular momentum of:
h = (40 kg• m2 )(1 deg/sec)( n rad/180 deg) = 0.7 N • m• s
and therefore an on-time for both starting and stopping the slew of:
 taccel = tdecel = (0.7 N. m. s)/(5 N . 0.5 m) = 0.3 sec

Total thruster on-time per slew: ttotal = 10 + 0.3 + 0.3 + 10 = 21 seconds

Fuel required per 180-degree slew:

m = (5 N)(21 s)/(9.8 m/s 2 . 218 s) = 49 g fuel

Sizing force level for momentum For SCS with 1.0 N•m•s wheels and 1-sec burns,
dumping:
F = (1.0 N.m.s)/(0.5 m * 1 sec)
F = h/(Lt) = 2.0 N
where
h = stored wheel momentum This is still well within the range of 1-lb thrusters, which are commonly used
L = thruster moment arm for orbit maintenance on small spacecraft. Reaction wheels with even larger
t = burn time capacity might be desirable if it would further reduce the number of times
the thrusters must be used.

Sensors. We complete this hardware unit by selecting periods, Sun-sensor-based attitude determination
the sensors needed for attitude determination. Consult systems must provide some way of tolerating the
Table 19-13 for a summary of typical devices as well as regular loss of this data without violating pointing
their performance and physical characteristics. Note, constraints.
however, that sensor technology is changing rapidly,
promising ever more accurate and lighter-weight Sun sensors can be quite accurate (<0.01 deg), but it is
sensors for future missions. not always possible to take advantage of that feature.
We usually mount Sun sensors near the ends of vehicles
Sun sensors are visible-light or infrared detectors that to obtain an unobstructed field view, so the Sun sensor
measure one or two angles between their mounting base accuracy can be limited by structural bending on large
and incident sunlight. They are popular, accurate, and spacecraft. Spinning satellites use specially designed
very reliable, but they require clear fields of view. They Sun sensors that measure the angle of the Sun with
can be used as part of the normal attitude determination respect to the spin axis of the vehicle, and they issue a
system, part of the initial acquisition or failure recovery pulse correlated to the time the Sun crosses the sensor
system, or part of an independent solar array orientation to provide spin-phase information. Also popular are
system. Since most low-Earth orbits include eclipse coarse Sun sensors, which are simply small solar cells
Table 19-13. Typical ADCS Sensors. Sensors have co ntinued to improve in performance while getting smaller and
sometimes less expensive.

Sensor Typical Performance Range Mass (kg) Power (W)

Gyroscopes Drift Rate = 0.003 deg/hr to 1 deg/hr < 0.1 to 15 < 1 to 200
Drift rate stability varies widely
Sun Sensors Accuracy = 0.005 deg to 3 deg 0.1 to 2 0 to 3
Star Sensors Accuracy = 1 arcsecond to 1 arcminute 2 to 5 5 to 20
(Scanners & Cameras) = 0.0003 deg to 0.01 deg
Horizon Sensors Accuracy:
- Scanner/Pipper 0.05 deg to 1 deg (0.1 deg is best for LEO) 1 to 4 5 to 10
- Fixed Head (Static) < 0.1 deg to 0.25 deg 0.5 to 3.5 0.3 to 5
Magnetometer Accuracy = 0.5 deg to 3 deg 0.3 to 1.2 <1

that issue a current roughly proportional to the cosine of many directions on a spacecraft, and then to estimate
the Sun angle. These sensors are so small and the Sun direction by solving the linear system equations
inexpensive that it is often feasible to put several in that results. Because coarse Sun sensors use no power
and almost never fail, they are often used in low-power staring sensors, which view the entire Earth disk (from
acquisition and fault recovery modes. GEO) or a portion of the limb (from LEO). The sensor
fields of view stay fixed with respect to the spacecraft.
Star sensors have improved rapidly in the past few This type works best for circular orbits, as they are
years and represent the most common sensor for high- often tuned for a tight range of altitudes.
accuracy missions. Star sensors can be scanners or
trackers. Scanners are used on spinning spacecraft. Horizon sensors provide Earth-relative information
Light from different stars passes through multiple slits directly for Earth-pointing spacecraft, which may
in the scanner’s field of view. After several star simplify onboard processing. The scanning types
crossings, we can derive the vehicle’s attitude. We use require clear fields of view for their scan cones
star trackers on three-axis stabilized spacecraft to track (typically 45, 60, or 90 deg half-angle). Typical
one or more stars to derive two- or three-axis attitude accuracies for systems using horizon sensors are 0.1 to
information. The majority of star trackers used today 0.25 deg, with some applications approaching 0.03 deg.
work much like digital cameras (and many of these are For the highest accuracy in low-Earth orbit, it is
increasingly called star cameras, rather than trackers), necessary to correct the data for Earth oblateness and
allowing starlight to fall on a CCD to create an image seasonal horizon variations.
of the star field. Then, internal processing determines a
three-axis attitude based on a star catalog. Many units Magnetometers are simple, reliable, lightweight sensors
are able to determine a very accurate attitude within that measure both the direction and magnitude of the
seconds of being turned on. Earth’s magnetic field. Magnetometer output helps us
establish the spacecraft’s attitude relative to the local
While star sensors excel in accuracy, care is required in magnetic field, which information can be combined
their specification and use. The most accurate star with magnetic field models and orbit information to
cameras are unable to determine attitude at all if the determine attitude relative to the Earth and inertial
spacecraft is rotating too fast, and other star sensors reference frames. However, their accuracy is not as
must know roughly where they are pointing to make good as that of star or horizon sensors. The Earth’s
their data useful. Therefore, the vehicle must be magnetic field can shift with time and is not known
stabilized to some extent before the trackers can operate precisely in the first place. To improve accuracy, we
effectively. This stabilization may require alternate often combine magnetometer data with data from Sun
sensors, which can increase total system cost. Also, star or horizon sensors. As a vehicle using magnetic
sensors are susceptible to being blinded by the Sun, torquers passes through changing magnetic fields
Moon, planets, or even high radiation levels, such as in during each orbit, we use a magnetometer to control the
the Van Allen belts, which is a disadvantage that must magnitude and direction of the torquers’ output relative
be accommodated in their application. Where the to the present magnetic field. In earlier spacecraft the
mission requires the highest accuracy and justifies a torquers usually needed to be inactive while the
high cost, we often use a combination of star trackers magnetometer was sampled to avoid corrupting the
and gyroscopes. The combination of these sensors is measurement. However, improvements in onboard
very effective: the gyros can be used for initial computing capability mean that coupling matrices can
stabilization and during periods of inference in the star be used to extract the torquer inputs from the field
trackers, while the trackers can be used to provide a measurement, allowing constant sampling even while
high-accuracy external reference unavailable to the torquing. Finally, good spacecraft ephemeris and
gyros. magnetic field models can be used in place of
magnetometers for some missions, but magnetometers
Horizon sensors (also known as Earth sensors) are will generally be more accurate.
infrared devices that detect the contrast between the
cold of deep space and the heat of the Earth’s GPS receivers are well known as high-accuracy
atmosphere (about 40 km above the surface in the navigation devices, but they can also be used for
sensed band). Simple narrow field-of-view fixed-head attitude determination. If a spacecraft is large enough to
types (called pippers or horizon crossing indicators) are place multiple antennas with sufficient separation,
used on spinning spacecraft to measure Earth phase and attitude can be determined by employing the
chord angles, which, together with orbit and mounting differential signals from the separate antennas. Such
geometry, define two angles to the Earth (nadir) vector. sensors offer the promise of low cost and weight for
Scanning horizon sensors use a rotating mirror or lens LEO missions. They can provide attitude knowledge
to replace (or augment) the spinning spacecraft body. accurate to 0.25 – 0.5 deg for antenna baselines on the
They are often used in pairs for improved performance order of 1 meter [Cohen 1996], and so are being used in
and redundancy. Some nadir-pointing spacecraft use low accuracy applications or as back-up sensors.
Development continues to improve their accuracy,
which is limited by the separation of the antennas, the Full three-axis knowledge requires at least two external,
ability to resolve small phase differences, the relatively non-parallel vector measurements, although we use
long wavelength of the GPS signal, and multipath IRUs or spacecraft angular momentum (in spinners or
effects due to reflections off spacecraft components. momentum-biased systems) to hold or propagate the
attitude between external measurements. In some cases,
Gyroscopes are inertial sensors that measure the speed if attitude knowledge can be held for a fraction of an
or angle of rotation from an initial reference, but orbit, the external vectors (e.g. Earth or magnetic) will
without any knowledge of an external, absolute have moved enough to provide the necessary
reference. We use gyros in spacecraft for precision information. In three-axis star trackers, each identified
attitude determination when combined with external star acts as a reference vector, which allows a single
references such as star or Sun sensors, or, for brief piece of hardware to generate a full three-axis attitude
periods, for nutation damping or attitude control during solution.
thruster firing. Manufacturers use a variety of physical
phenomena, from simple spinning rotors to ring lasers, For Earth-pointing spacecraft, horizon sensors provide
hemispherical resonating surfaces, and laser fiber optic a direct measurement of pitch and roll axes, but require
bundles. Gyros based on spinning rotors are called augmentation for yaw measurements. Depending on the
mechanical gyros, and they may be large iron gyros accuracy required, we use Sun sensors, magnetometers,
using ball or gas bearings, or may reach very small or momentum-bias control with its roll-yaw coupling
proportions in so-called MEMS gyros. (MEMS stands for the third degree of freedom. For inertially pointing
for microelectromechanical systems.) The gyro spacecraft, star and Sun sensors provide the most direct
manufacturers, driven largely by aircraft markets, measurements, and IRUs are ideally suited. Frequently,
steadily improve accuracy while reducing size and only one measurement is made in the ideal coordinate
mass. frame (Earth or inertial), and the spacecraft orbit
parameters are required to convert a second
Error models for gyroscopes vary with the technology, measurement or as an input to a magnetic field model.
but characterize the deterioration of attitude knowledge Either the orbit parameters are uplinked to the
with time. Some examples of model parameters are spacecraft from ground tracking and propagated by
drift bias, which is simply an additional, false rate the onboard processing, or they are obtained from onboard
sensor effectively adds to all rate measurements, and GPS antennas.
drift bias stability, which is a measure of how quickly
the drift bias changes. When used with an accurate FireSat II sensors. The external sensors for FireSat II
external reference, such as a star tracker, gyros can could consist of any of the types identified. For the 0.1
provide smoothing (filling in the gaps between tracker deg Earth pointing requirement, however, horizon
measurements) and higher frequency information (tens sensors are the most obvious choice since they directly
to hundreds of hertz), while the tracker provides lower measure the two axes we most need to control. The
frequency, absolutely referenced information whenever accuracy requirement makes a star sensor a strong
its field of view is clear. Individual gyros provide one candidate as well; its information would need to be
or two axes of information and are often grouped transformed, probably using an onboard orbit
together as an inertial reference unit (IRU) for three ephemeris calculation, to Earth-relative for our use. The
full axes and, sometimes, full redundancy. IRUs with 0.1 deg accuracy is at the low end of horizon sensors’
accelerometers added for position and velocity sensing typical performance, so we need to be careful to get the
are called inertial measurement units (IMU). most out of their data. We assume we also need a yaw
sensor capable of 0.1 deg, and this choice is less
Sensor Selection. Sensor selection is most directly obvious. Often, it is useful to question a tight yaw
influenced by the required orientation of the spacecraft requirement. Many payloads, e.g. antennas, some
(e.g. Earth-, Sun- or inertial-pointing) and its accuracy. cameras, and radars, are not sensitive to rotations
Other influences include redundancy, fault tolerance, around their pointing axis. For this discussion, we will
field of view requirements, and available data rates. assume this requirement is firm. We could use Sun
Typically, we identify candidate sensor suites and sensors, but their data needs to be replaced during
conduct a trade study to determine the most cost- eclipses. Magnetometers don’t have the necessary
effective approach that meets the needs of the mission. accuracy alone, but with our momentum-bias system,
In such studies the existence of off-the-shelf roll-yaw coupling, and some yaw data filtering, a
components and software can strongly affect the magnetometer-Sun sensor system could work for
outcome. In this section we will only briefly describe normal operations. The magnetometers would also
some selection guidelines.
improve the control effectiveness of the magnetic With rate information onboard, we only need an
torquers. occasional update from an attitude sensor. If the
magnetic field were stronger, we might be able to filter
At this point we consider the value of an inertial magnetometer data to get to 1 deg of accuracy, but it is
reference package. Such packages, although heavy and doubtful at this high altitude. Star cameras are small
expensive for high-accuracy equipment, provide a and very accurate, but they are expensive. One useful
short-term attitude reference that would permit the rule of thumb is: If at all possible, sense the thing you
Earth vector data to be used for full three-axis need to point at. Sun pointers should have Sun sensors
knowledge over an orbit. A gyro package would also and Earth pointers should have Earth sensors. Because
reduce the single measurement accuracy required of the the satellites will have the same direction pointing
horizon sensors, simplifying their selection and toward the Earth throughout their mission, the best
processing. Such packages are also useful to the control option appears to be an Earth sensor. We would need to
system if fast slews are required, and here is where select a sensor designed for high altitudes. However, we
FireSat II demands a gyro. We need to perform the 180 still have no yaw data, and since the satellites must
degree slews as quickly as we can, to avoid losing too point accurately at multiple targets simultaneously,
much data. So, we will include a MEMS-based inertial yaw accuracy is critical.
package for FireSat II. They will be arranged
perpendicular to one another, in a pyramid around the Z Because the satellites will be communicating with each
axis. This arrangement gives 3 axes of information with other, it is conceivable that the communication signal
maximal redundancy around the slewing axis. Now that strength could be used as an attitude determination data
we have decided to include gyros, a careful trade study source for yaw control. That is, a feedback loop would
should be done to determine whether an inexpensive close around the communication system’s own measure
MEMS gyro package combined with just the Earth of its link margin; maximizing the link margin would
sensors eliminates our need to include Sun sensors and provide the attitude goal we want. For this exercise we
magnetometers. Leaving out some of these other will not assume such an option is available. Instead, we
sensors could give better reliability or lower total cost. will choose the star camera after all; a simple onboard
We may also choose a slightly different arrangement of ephemeris calculation will tell the spacecraft where its
the sensors to improve accuracy in one direction at the target satellite is in inertial space. There may be clever
expense of accuracy in another direction. We would tricks that we could use with ground-based methods to
need to do these kinds of detailed trade studies in later avoid using star cameras, such as combining orbit
iterations of the design process. tracking and attitude data. However, the complexity of
operating three separate satellites that have to work
Finally, we will want a simple, coarse control mode to together will likely prove more expensive in software
initially point the arrays at the Sun and to protect the development and operating costs than just buying three
spacecraft in the event of an anomaly. By using 6 star cameras. At least we save money by not including
coarse Sun sensors pointing along the positive and horizon sensors. As for our rule of thumb of sensing
negative of each axis, the spacecraft can derive the what you’re pointing at, we now see there can be
location of the Sun from any attitude and use magnetic situations for which this rule cannot be followed. Still,
torquers to rotate the spacecraft so the arrays are lit. it’s always a good place to start.
Then, since the attitude relative to nadir will change as
FireSat II follows its polar orbit, we can be sure to get a We will propose the same plan for an initial acquisition
good communication signal at some point, so that we and safe control mode as for FireSat II. However, SCS
can receive telemetry and send commands. will use reaction wheels for control, and will have the
benefit of accurate rate sensors to improve
SCS sensors. For SCS attitude determination, low- performance. This is a good thing, since SCS satellites
power gyros can provide rate information. Accurate need more power than FireSat II, and so will probably
gyros can be heavy and often use a lot of power; we not have as much time to acquire the Sun (i.e. get the
have neither high accuracy needs nor an excess of solar arrays lit). Table 19-14 summarizes our FireSat II
power or mass in our budgets. Therefore, we will use and SCS hardware selections.
light and inexpensive MEMS gyros. We need a
minimum of 3 MEMS gyros—one for each axis—but
by employing 4-6 gyros we can cross-compare the gyro
data and remove the larger bias errors that MEMS
gyros normally have.


Table 19-14. FireSat II and SCS Spacecraft Control Component Selection.

Mission & Type Components Rationale
FireSat II Actuators Momentum Wheel (1) - Pitch axis torquing
- Roll and yaw axis passive stability
Magnetic Torquers (3) - Roll and yaw control
- Pitch wheel desaturation
Thrusters (4) - Rapid changes in pitch wheel momentum
- 180-deg yaw slew maneuvers
FireSat II Sensors Horizon Sensor (1) - Provide basic pitch and roll reference
- Can meet 0.1 deg accuracy
- Lower mass and cost than star sensors
Sun Sensors (6) - Initially acquire vehicle attitude from unknown
orientation
- Coarse attitude data
- Fine data for yaw
Magnetometer (1) - Coarse yaw data
- Needed for more precise magnetic torquing
MEMS Gyros (3) - Needed for rotational rate data during thruster
operations
SCS Actuators Reaction Wheels (3) - Three-axis stabilization
(for each satellite) - Storing momentum from solar radiation torques
Thrusters (4) - Thrusting and steering during orbit maneuvers
- Removing momentum from reaction wheels
SCS Sensors Star Camera (1) - Determining absolute attitude
(for each satellite) MEMS Gyros (4) - Determining rotational rates; having 4 gyros allows
large biases to be canceled out
- Propagating attitude solution when star sensor data
Sun Sensors (6) unavailable
- Initially acquire vehicle attitude relative to Sun to
get power quickly

Once the hardware selection is complete, it must be components satisfy all mission requirements, with
documented for use by other system and subsystem thrusters required for orbit injection and 180 degree
designers as follows. slews.
- Specify the power levels and weights required Components Type Weight Power Mounting
(kg) (W) Considerations
for each assembly Momentum Mid-size, < 5 total, 10 to Momentum
- Establish the electrical interface to the rest of Wheel 20 N•m•s with drive 20 vector on pitch
momentum electronics axis
the spacecraft Electromagnets 3, 10 A •m2 2, 5 to 10 Orthogonal
- Describe requirements for mounting, including configuration
alignment, or thermal control current best to reduce
drive cross-coupling
- Determine what telemetry data we must electronics
process Sun Sensors 6 wide- < 1 total 0.0 Free of
angle viewing
- Document how much software we need to coarse Sun obstructions
develop or purchase to support onboard sensors and reflections
providing
calculation of attitude solutions 4 7c
Specific numbers depend on the vendors selected. A steradian
typical list for FireSat II might look like Table 19-15, coverage;
5-10 deg
but the numbers could vary considerably with only accuracy
slight changes in subsystem accuracies or slewing Horizon Scanning 5 total 10 Unobstructed
Sensors type (2) view of Earth’s
requirements. plus horizon
electronics;
Table 19-15. FireSat II Spacecraft Control 0.1 deg
y
Subsystem Summary. The baseline ADCS Gyroscopes MEMS (J)
M 3 < 1 total 0.1 Mounted
 only low perpendicular Figure 19-8. Block diagram of a Typical Attitude
accuracy to each other,
needed in a pyramid Control System with Control along a Single Axis.
around Z axis Control algorithms are usually implemented in an
Thrusters Hydrazine; Propellant N/A Alignments
5 N thrust weight and moment
onboard processor and analyzed with detailed
depends on arm to center simulations.
mission of gravity are
critical
Magnetometer 3-axis <1 5 Need to isolate We typically apply linear theory only to preliminary
magnetometer analysis and design. We also maintain engineering
from
electromagnets, margin against performance targets when using linear
either theory because, as the design matures, so does our
physically or
by duty-cycling
understanding of the nonlinear effects in the system.
the magnets Nonlinear effects may be inherent or intentionally
introduced to improve the system’s performance.

19.1.5 Define the Determination and Control Another reason to maintain margin, especially in
Algorithms actuator sizing, is that while systems engineers always
Finally, we must tie all of the control components carefully budget mass, they cannot usually track the
together into a cohesive system. Generally, we begin moment of inertia matrix as well. Moment of inertia is
with a single-axis control system design (See Figure 19- the most important quantity to ADCS designers, and a
8). As we refine the design, we add or modify feedback given mass may have a wide range of moment of inertia
loops for rate and attitude data, define gains and values depending on how the mass is distributed.
constants, and enhance our representation of the system
to include three axes of motion (though we may still Feedback control systems are of two kinds based on the
treat these as decoupled for early design iterations). To flow of their control signals. They are continuous-data
confirm that our design will meet requirements, we systems when sensor data is electrically transformed
need good mathematical simulations of the entire directly into continuously flowing, uninterrupted
system, including sensor error models and internal and control signals to the actuators. By contrast, sampled-
external disturbances. Usually, linear differential data systems have sensor sampling at set intervals, and
equations with constant coefficients describe the control signals are issued or updated at those intervals.
dynamics of a control system well enough to allow us Most modern spacecraft process data through digital
to analyze its performance with the highly developed computers and therefore use sampled-data control
tools of linear servomechanism theory. With these same systems.
tools, we can easily design linear compensation to
satisfy specifications for performance. Although it is beyond the scope of this book to provide
detailed design guidance on feedback control systems,
the system designer should recognize the interacting
effects of attitude control system loop gain, capability
of the attitude control system to compensate for
disturbances, accuracy of attitude control, and control
system bandwidth.

Three-axis stabilization. Different types of active

control systems have different key parameters and
algorithms. For systems in which spacecraft rates will
be kept small, three-axis control can frequently be
decoupled into three independent axes. In this simplest
form, each axis of a spacecraft attitude control system
can be represented as a double-integrator plant (i.e.
1/Is2 in the s-domain) and may be controlled by a
proportional-derivative (PD) controller, which has
control torque T C = KPθ E + KDOE. Here, OE is the
attitude error angle, and ω E is the attitude rate error.
The most basic design parameter in each axis is its
proportional gain, KP . (KP is also often called a
position gain.) This gain represents the amount of
control torque we want to result from a unit of attitude
error, and so it has units of torque divided by angle, e.g. The value of KP also largely determines the attitude
newton-meters per radian. The proportional gain is control system bandwidth, which is a measure of its
selected by the designer and must be high enough to speed of response. The mathematical definition of
provide the required control accuracy in the presence of bandwidth requires a bit more explanation than we want
disturbances, which can be guaranteed by setting to go into here, but it can be approximated as (On =
KP>T D /θ E,max, where T D is the peak disturbance torque, (KP/I) 1/2 , where I is the spacecraft moment of inertia.
and θ E,max is the allowable attitude error. Note that this The bandwidth defines the frequency of disturbances
error will remain for as long as the disturbance torque and of commanded motions at which control authority
remains; the error that remains in the presence of begins to diminish. Attitude control and disturbance
disturbances is called steady-state error. rejection are effective from 0 frequency (i.e. constant or
d.c. inputs) up to the bandwidth. Speed of response is

Table 19-16 ADCS Vendors. Typical suppliers for ADCS components. An up-to-date version of this table can be
found at the SMAD web site.
Company Sun Earth Magnet- Star Gyro GPS Mom./ CMG Magnetic Thrusters
Sensors (horizon) ometers Sensors Reaction Torquers
Sensors Wheels
Adcole Corporation X
Aeroj et X
Ball Aerospace and
X
Technologies Corp.
Billingsley Aerospace &
X
Defense
Bradford Engineering X X X
Comtech AeroAstro X X
EADS Astrium X X X X
EADS SODERN X X
EMS Technologies, Inc. X
Finmeccanica (incl. SELEX
X X X X
Galileo
ITT Aerospace X
General Dynamics X
Goodrich (incl. Ithaco) X X X X
Honeywell Space Systems
X X
incl. Allied
Jena Optronik X X
L-3 Space & Navigation X X X
Kearfott Guidance &
X X
Navigation Corp.
Meda X
Micro Aerospace
X X
Solutions
Microcosm, Inc. X
NASA Goddard Space
X X
Flight Center
Northrop Grumman (incl
X X
Litton)
Øersted - DTU X
Optical Energy
X X
Technologies
Rockwell Collins
X
Deutschland (incl. Teldix
Servo Corp. of America X
StarVision Technologies X
Surrey Satellite
X X X X X X X X X
Technologies - US LLC
Systron Donner Inertial X
Terma X
Watson Industries, Inc. X X

disturbance torques to create it. The integrator causes

approximately the reciprocal of bandwidth. Note that the correcting control torque to gradually increase over
proportional gain is inversely proportional to allowable time to remove that steady-state error completely. If the
error, and bandwidth is proportional to the square root disturbance torque changes magnitude or direction, the
of the proportional gain. Therefore, high accuracy PID controller will take some time to adjust, but it will
implies high K P and high bandwidth. However, there track with no steady-state error. This improved
are practical limits on bandwidth. If the bandwidth is performance generally comes at the price of reduced
high enough that it includes bending frequencies for the stability margins, so the stability and performance
spacecraft structure, then control-structure interactions requirements must always be carefully balanced.
must be carefully considered. If the bandwidth is higher
than the maximum sampling rate of critical attitude In addressing control-structure interactions, the control
sensors, some care must be taken to avoid excessive designer must keep in mind that, if bending causes
attitude responses to the slow updates. Finally, high apparent attitude errors, the control system may
bandwidth control systems may be more likely to have aggravate the bending motion, eventually affecting
adverse reactions to inaccurate or lagging system control stability and perhaps damaging the bending
inputs; this condition is referred to as having structures. Though there are control design techniques
insufficient stability margin. to deal with these potential problems, the control
designer will often start by recommending a minimum
There are design techniques to address these concerns. frequency for structural bending modes that would
One such technique is to modify the system response to allow flexible effects to be neglected. This minimum
increase stability while keeping the same attitude frequency is generally an order of magnitude greater
accuracy by increasing attitude rate damping, which is than the bandwidth needed to achieve required
controlled by the derivative gain, or rate gain, KD . For accuracy. Then, structural engineers will attempt to
a given value of K P , increasing K D will slow down the honor this restriction in their design. For further
response of the control system, but it will not diminish discussion of how structural flexibility interacts with
the accuracy in the end, provided the disturbance feedback control systems, see section 3.12 of Agrawal
environment or slewing requirements do not require [1986].
very high speed of response. This technique is often
used to improve stability margins in systems where the Spin stabilization and momentum bias. The
final accuracy is more important than the amount of fundamental concept in spin stabilization is the nutation
time it takes to achieve that accuracy. Note that, though frequency of the vehicle. For a spinning body, the
there are multiple methods for selecting K P and KD , inertial nutation frequency ( wni) is equal to O)S I S/IT
such as linear-quadratic regulation (LQR) and other where ω S is the spin frequency, I S is the spin axis of
state-space methods, the system response generally inertia, and I T is the transverse axis inertia. For a
must still be understood in terms of bandwidth and momentum-bias vehicle with a non-rotating body and a
damping to achieve confidence that the control gains momentum wheel (or a dual-spin vehicle with a non-
will provide both stability and adequate performance. rotating platform), the nutation frequency is ω ni =h/I T
where h is the angular momentum of the spinning body
Sometimes the problem facing a designer is that (or bodies). (Note that this is really the same equation
stability margin requirements are not difficult to meet, as for a spinning body, for which h = ω S I S.) Thus,
but accuracy requirements are very difficult. A common spacecraft with large inertias and small wheels have
technique for improving on the performance of a PD is small nutation frequencies (i.e. long periods).
to add an integrator to the feedback loop; this is called a
proportional-integral-derivative controller (PID). Attempting to control the vehicle with a bandwidth
Recall that in a regular PD controller, a steady-state faster than the nutation frequency causes it to act more
error of some kind will remain as long as there are
like a three-axis stabilized vehicle, that is, the stiffness a dedicated reference such as Wertz [1978]. We will
of the spin is effectively reduced. In general, we highlight only some of the basic concepts.
attempt to control near the nutation frequency or
slower, with correspondingly small torques for given The basic algorithms for determination depend on the
attitude errors. Because we generally wish to avoid coordinate frames of interest and the geometry and
changing the spin rate with control torques, the torque parameterization of the measurements. Useful
is applied perpendicularly to the spin. We can then coordinate frames can include the individual sensor
relate the achieved angular rate, w, to the applied torque frames, the local vertical frame, and an inertial frame,
by simplifying Euler’s equations to only the gyroscopic such as Earth-centered inertial (ECI). The geometry of
torque component: T = w x H = wH where H is the the measurements is different for different sensors, as
system angular momentum. A lower limit on control discussed in Sec. 19.1.4, and they are generally
bandwidth is usually provided by the orbit rate wO, parameterized as sequential rotations about the axes of
a frame, called Euler angles (e.g. roll, pitch, yaw), or as
which for a circular orbit is wp = µ;/ r 3 where attitude quaternions, which are unit vectors with four
14m3 /s2 for Earth orbits and r is the orbit
µ =3.986 x 10 elements that define a rotational axis in a frame, called
radius. the eigenaxis, and the amount of rotation about the
eigenaxis. Inertial reference units (with their supporting
Example of Magnetic Momentum Management. software) and star sensors are well suited to
Managing the long-term buildup of angular momentum quaternions, whereas Earth-pointing vehicles often use
in the orbiting spacecraft often drives the ADCS a local-vertical set of Euler angles, much like aircraft.
requirements. As would be expected for an element so
central to ADCS design, there are many different Simple spacecraft may use the sensor output directly for
methods for keeping momentum to levels that do not control, whereas more complex vehicles or those with
degrade data or equipment. We present here a simple, higher accuracy requirements employ some form of
common, and relatively inexpensive algorithm for averaging, smoothing or Kalman filtering of the data.
commanding magnetic torquers using only The exact algorithms depend on the vehicle properties,
magnetometer data. It functions as an excellent active orbit, and sensors used.
rate damping controller, and after the spacecraft’s rate
are very low, it keeps the angular momentum at FireSat II algorithms. For our momentum-bias FireSat
manageable levels by canceling it out on an orbit- II example, control separates into pitch-axis control
averaged basis. Often called B-dot, this algorithm is using torque commands to the momentum wheel and
based on back-differencing magnetometer data to roll-yaw control using current commands to the
estimate the time derivative of the magnetic field (or magnetic torquers. The pitch-wheel desaturation
“Bdot”). For rate damping purposes, we assume the B- commands must also be fed (at a slower rate) to the
field is changing in the body due only to spacecraft torquers. The pitch-wheel control is straightforward,
rotation, which is true for all but rotation rates on the using PD control and, optionally, integral control, in
order of orbital rate. Specifically, Bdot = B k+1 – Bk = - which commands are augmented by a small torque
w x B. The spacecraft dipole vector, D, is set equal to a proportional to the integral of the attitude error. The
gain multiplied by the magnetic field: D = -k Bdot = roll-yaw control design starts by using the linearized
kw x B. Then, when the magnetic field acts on that nutation dynamics of the system, and is complicated by
vector, it produces a control torque T, = D x B that is the directional limitations of electromagnetic torque
always in a direction that slows down the spacecraft (the achievable torque is perpendicular to the
rotation except when w is aligned with B (in vector instantaneous Earth magnetic field).
terminology, w• T, ≤ 0). Once the spacecraft is slowed
down, the algorithm will attempt to follow the magnetic The nadir-oriented control system may use an Earth-
field vector, its success determined by the strength of referenced, aircraft-like Euler angle set (i.e. roll, pitch
the gain, k, that is chosen. However, whether it follows & yaw), although quaternions may simplify
the field doesn’t matter: The goal of slowing the calculations during off-nominal pointing; quaternion
spacecraft and then maintaining a slow momentum has calculations generally require fewer arithmetic steps,
been achieved. Flatley, Morgenstern, Reth and Bauer they simplify tracking commands, and they have no
[1997] did not invent the B-dot controller, but their singularities. The horizon sensors directly read two of
paper nicely explains in clear English and equations the angles of interest, pitch and roll. Yaw needs to be
how something so simple works so well. measured directly from Sun position (during orbit day)
or from the magnetometer readings (using a stored
Algorithms for attitude determination. A full model of the Earth’s field plus an orbit ephemeris), or
discussion of attitude determination algorithms requires inferred from the roll-yaw coupling described earlier.
MEMS gyro data can be used to enhance the attitude Agrawal, Brij N. 1986. Design of Geosynchronous
estimate, but they can have large biases that change Spacecraft. Englewood Cliffs, NF: Prentice-Hall, Inc.
with temperature, so it may be best not to have to rely
Chobotov, Vladimir A. 2002. Orbital Mechanics, 3rd
on them for our accuracy requirements. The magnetic
ed. Washington, DC: AIAA.
field and Sun information require an uplinked set of
orbit parameters, and increase the computational Cohen, Clark E. 1996. Global Positioning System:
requirements of the subsystem. Overall, meeting the 0.1 Theory and Applications, Vol. II, B. Parkinson, J.
deg yaw requirement when the Sun is not visible will be Spilker, P. Axelrad, and P. Enge (Eds.), American
the toughest challenge facing the ADCS designer, and Institute of Aeronautics and Astronautics.
planning to coast through the dark periods without
direct yaw control may be most appropriate. Flatley, Thomas W., W. Morgenstern, A. Reth, and
F.H. Bauer, 1997. “A B-Dot Controller for the
RADARSAT Spacecraft,” NASA GSFC Flight
SCS Algorithms. A 3-axis PD controller will be
Mechanics/Estimation Theory Symposium, Greenbelt,
sufficient for our control accuracy needs. In this control
MD.
algorithm, an error angle and error rate is calculated for
each axis, and a control torque calculated based on Fortescue, P. and J. Stark. 2003. Spacecraft Systems
rd
those. That calculation done for the 3 axes gives a Engineering, 3 ed. NY: John Wiley & Sons.
control torque vector, which is then distributed among
the reaction wheels. If one of the wheels is asked to Griffin, M. and J.R. French. 2004. Space Vehicle
issue a control torque that exceeds its capability, then Design. 2nd ed. Washington, DC: AIAA.
the spacecraft would have a total torque vector that is Hughes, Peter C. 2004. Spacecraft Attitude Dynamics,
not in the same direction as the requested vector. For Mineola, NY: Dover Publications, Inc.
our mission, that is not too much of a danger, since we
don’t have any elements that would be damaged by Sun Kaplan, Marshall H. 1976. Modern Spacecraft
exposure or the like. Still, it is a simple matter to scale Dynamics and Control. New York: John Wiley and
all 3 vector components by the same factor that would Sons.
make the largest of them equal to its limit while Thomson, William T. 1986. Introduction to Space
preserving the direction of the torque vector. Note that Dynamics, Mineola, NY: Dover Publications, Inc.
this is safe for a PD controller, but that stability
problems can result if the same tactic is used for a PID Wertz, James R. ed. 2009. Orbit & Constellation
controller without taking some care to prevent Design & Management. Hawthorne, CA: Microcosm
problems. Press and New York, NY: Springer.
Wertz, James R. ed. 1978. Spacecraft Attitude
The data from the star camera and the rate gyros in each Determination and Control. Dordrecht, The Nether-
SCS satellite can be combined in any of a number of lands: D. Reidel Publishing Company.
ways. We will simply trust the star camera attitude as
accurate whenever it gives us data it indicates is valid, Wie, Bong. 2008. Space Vehicle Dynamics and
and we will use the gyros to provide direct rate Control, 2nd ed., Reston, VA: AIAA Inc.
measurements and to propagate the attitude solution by
integrating the rate over the sampling time if the star
camera fails to provide a valid attitude. This method is
not the most accurate, but generally, more accuracy
comes at the expense of more complexity, which then
costs money and time in flight software development
and testing. This is a good lesson to wrap up this
section: Though it can be tempting to always reach for
your best, most expensive tool, it is better engineering
to try to get by with the cheapest system that will meet
your requirements with appropriate margin.

References
Keep all previous references and also include the
following (I have other texts and papers to add to this
list, but I wanted to get this draft out. I’ll send the
reference list later.):

1 Outer Planets
Solar& Interplanetary
(Heliocentric Trajectories)
Sun
Interplanetary spacecraft can get
very cold, have little power. May
want thrusters only as reaction
wheels may seize. Planetary
Mercury

Venus

Venus +

Mercury
)16 Orbits

Mars

SRP is only
40 Sun-Earth Lagrange
Points 1.5 million km 40
Gravity Gradient
fades to almost Trans Lunar
important
nothing far from Lunar Region
disturbance
400,000 km planets & moons. Orbit
far from
planets or
moons. fO .r,
Geosynchronous Orbit
3 (GEO) 35,786 km Cis Lunar
woo 4n
Magn c fieldtoo Region
weak at GEO for `
c magnetic torquers. - **I%b
W 20,000 km
E Gravity gradient too weak
V;N r_0
at high altitude for good
Y cv Medium Earth Orbit passive stabilization.
'^
r c (MEO) 2000 km -GEO
o

o`Cc 2,000 km
0V O
w
_T
r6 -
Low
Above 700-800 km, Earth Magnetic Field & O
aj atmosphere very thin; Orbit Gravity Gradient are
SRP>Orag. (LEO) both strong in LEO. Q
Cr ,w
LL,
Drag is T
. . est ISS 0,
W

Z f 200 km f`
350 km ^-

3 i C

02010 Microcosm
5ME-0172-01-C