FACULDADE DE E NGENHARIA DA U NIVERSIDADE DO P ORTO
Orbit Calculation and Re-Entry
Control of VORSat Satellite
Cristiana Monteiro Silva Ramos
Master in Electrical and Computers Engineering
Preparation for the MSc Dissertation Report
Supervisor: Sérgio Reis Cunha (PhD)
February 2011
Contents
1
Introduction
1.1 Motivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1.2 Objectives and Work Organization . . . . . . . . . . . . . . . . . . . . . . . . .
1.3 Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2
Case Study Presentation
2.1 CubeSat . . . . . . . . . . . . . . . .
2.1.1 QB50 . . . . . . . . . . . . .
2.2 Project . . . . . . . . . . . . . . . . .
2.2.1 Project’ Vision and Objectives
2.2.2 GAMA-Sat . . . . . . . . . .
2.2.3 VORSat . . . . . . . . . . . .
2.3 Summary . . . . . . . . . . . . . . .
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3
Literature Review
3.1 Satellites . . . . . . . . . . . . . .
3.1.1 Orbits . . . . . . . . . . .
3.2 Navigation System . . . . . . . .
3.2.1 Measuring from the Earth
3.2.2 Measuring from Space . .
3.2.3 GPS Antenna . . . . . . .
3.2.4 Communication . . . . . .
3.2.5 Microcontroller . . . . . .
3.3 Summary . . . . . . . . . . . . .
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References
1
1
2
3
27
i
ii
CONTENTS
List of Figures
1.1
Work Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.1
2.2
2.3
2.4
2.5
2.6
First draft of VORSat/GAMA-Sat architecture [1]
GAMA-Sat Structure [1] . . . . . . . . . . . . .
Capsule phases. Courtesy of João Gomes . . . .
Capsule Structure. Courtesy of João Gomes . . .
Interior of the Capsule. Adapted from [2] . . . .
Batery Charge [1] . . . . . . . . . . . . . . . .
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3.1
3.2
3.3
Evolution of satellites over the years . . . . . . . . . . . . . . . . . . . . . . . .
Ground Station. [3] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Equipment ON/OFF state depending on battery charge. [3] . . . . . . . . . . . .
14
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iii
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3
iv
LIST OF FIGURES
Chapter 1
Introduction
This chapter is a small introduction to a satellite program in which the work described in this
thesis is included. Additionally this work objectives and organization is presented, as well as a
brief description of the structure of this document.
1.1
Motivation
"The space age began when the cientific community began to accept the extraterrestrial origin
of meteorites. However, it took more one hundred and fifty years before technology could support
the engineering part of the space age: the design and construction of an artifact, its launch and
recovery from orbit" [4].
Despite the first man-made spacecraft was only launched in 1957, Sputnik-1, satellite orbits
had already been studied. Starting from Newton’s formulation of the law of gravity, scientists
sought continuously to develop and refine analytical theories describing the motion of the Earth’s
only natural satellite, the Moon. But, only at the 20th century was possible the physical (and not
only theoretical) exploration with the construction of first rockets. Common reasons for exploring
space include advancing scientific research, communications, uniting different nations, ensuring
the future survival of humanity and developing military and strategic advantages against other
countries.
In the past few years much of the attention of the space industry has shifted towards the development of small satellites. These satellites, often called picosats, nanosats, or microsats are
generally less than 200 kg and, in many cases, are as little as 1 - 5 kg. Such satellites, which range
in size from refrigerators to small soda cans, offer many potential benefits over traditional space
satellites. The development of small satellites and the use of commercial off-the-shelf components
has dropped the price of satellite launches to as low as a few million dollars for light satellites,
and a few tens of millions for heavy satellites. This fact led to a significant increase in the number
1
2
Introduction
of satellites developed and launched both by companies and universities . According to data provided by National Aeronautics and Space Administration (NASA), since the first artificial satellite,
Sputinik-1, thousands of satellites have already been launched. Currently there are approximately
3,000 satellites operating in Earth orbit, out of the roughly 8,000 man-made objects, together with
countless pieces of space debris[5]. When satellites reach the end of their mission, the majority of
these have fallen into unstable orbits and incinerated during reentry.
Even though Portugal had a successful operation of a satellite developed by a consortium
of mainly national entities (PoSAT-1), the fact is that the Portuguese aerospace industry is very
small and its activity is primarily focused on supplying high quality parts for the foreign market.
The Portuguese experience in fully designing and building a satellite and reentry capsule is rather
small[6].
In the present project, a literature review on Satellites and Navigation was carried out having
in mind a specific case study, concerning a small satellite program being developed at the Faculty
of Engineering of the University of Porto (FEUP). This project regards both the development of
a CubeSat in partnership with the Portuguese company TEKEVER (named GAMA-Sat) and an
Earth Reentry Capsule (ERC).
The cubesat has as main mission targets the determination of the satellite’s attitude from the
ground and from other near cubesats, the communication between satellites and the measurement
at the differential arm between them to obtain the differential drag affecting neighbour satellites.
The ERC main goals are to successfully achieve the survival of the reentry capsule, with a desired reentry path and in a desired landing site. The integration of a cork composite material in the
ablative Thermal Protection System (TPS) of a nanosatellite/picosatellite and the use of ARGOS
system inside ERC as the primary communication channel after reentry are being considered.
This MSc Thesis will focuses on the development of an onboard navigation and guidance system capable of predicting the satellite path and control the location of de-orbiting. Concerning the
capsule, the mission success may depend on using trajectory planning algorithms for the determination of the time and place of landing for its recovery. In this case, instead of having multiple
stations trying to track the satellite, they must continuously estimates its evolution from present
and past Global Position System (GPS) data and from model of the atmosphere and transmit it to
the mission control . Because of this, the algorithm must be able to estimate how many transmitions can be made, depending on the stage at which the orbit lies and battery status. The objective
is to maximize the number of useful transmissions while minimizing the battery usage.
1.2
Objectives and Work Organization
Based on the satellite operating architecture, the main objectives of this work are the following:
• To develop algorithms to calculate the evolution of both satellites in their Low Earth Orbit
(LEO), from their release until the time of reentry, by using present and past GPS data and a
the model of the atmosphere. It is particularly important to determine the time and the place
of reentry as well as the splashdown spot of the capsule for a successful recovery;
1.3 Structure
3
• To develop a methodology to update satellite ephemeris information obtained through sporadic GPS observations (omnidirectional antenna from array of patch antennas on different
faces).
•
To develop a control algorithm to adaptively adjust the point of re-acting on the drag
coefficient of the satellite. The actuation of the drag coefficient is yet to be defined, however
something like, opening and closing light spoilers on the surfaces of the satellite is foreseen.
The objectives mentioned above were undertaken during a period of approximately eight
months, and the work load was divided as shown in the Figure 1.1 :
Figure 1.1: Work Plan
1.3
Structure
This document is organized in three chapters. Each chapter starts with a small introductory text
describing the chapter’s intent and to the exception of the present one, ends with a brief summary.
The current introduction chapter presents the background motivation for Satellites in general and
the project. It also enumerates the proposed objectives as well as the work plan for the duration of
the project. Finally, it describes the document’s organization.
In the second chapter, the case study is presented and its architecture described. The present
development stage is also discussed while contextualizing the relevance of this work within the
project.
The third chapter offers a view on the state-of-art related with satellites and its navigation
system. The intent of this chapter is to introduce the different scientific themes and trends relevant
to this thesis and the developed work.
4
Introduction
Chapter 2
Case Study Presentation
This chapter aims to provide a general overview of the case in study, presenting and describing
its vision and objectives. The first section will provide a brief description of CubeSats and the
following presents the satellites that are being developed by FEUP CubeSat Team.
2.1
CubeSat
The CubeSat concept arose in 1999, through a collaboration between Dr. Bob Twiggs (Stanford University) and Dr. Jordi Puig-Suari (California Polytechnic State University San Luis
Obispo-Cal Poly)[7]. Cubesats were developed to help universities worldwide to perform space
science and exploration.
A CubeSat standard is a miniaturized satellite with the restricted dimensions and weight of, 10
cm cube and 1 kg, respectively. It offers the following standard functions of a normal satellite:
• Attitude determination and control;
• Uplink and downlink telecommunications;
• Power subsystem including a battery and body-mounted solar panels;
• On-board data handling and storage by a CPU ;
• Plus either a technology package or a small sensor or camera.
To keep it simple this is achieved by using commercial off-the-shelf components. Typically,
they are launched and deployed from a mechanism called a Poly-Picosatellite (P-POD). During
the launch phase the signal is turned off, the Cubesat are put into a Low Earth Orbit (LEO) and
deployed once the proper signal is received from the launch vehicle. Approximately 90% of
CubeSats were launched with P-POD [8].
5
6
Case Study Presentation
The bottleneck for small satellites remains in the physical space available, so the payloads
are very limited due to the extremely restricted size and mass of a CubeSat satellite. This seriously restricts payloads requiring large optics or bulky components. The limited surface area of a
CubeSat restricts the amount of solar power that may be generated, restricting power available for
computation, communications, and payloads.
Despite the limitations of CubeSats, in the past few years much of the attention of the space industry has shifted towards the development of small satellites because it represent a cost-effective
independent mean of getting a payload into orbit. They offer many potential benefits over traditional space satellites, such as, carry new scientific instruments, unproven technologies in low
prices but also greatly increase the available functionality.
Nevertheless, several companies but mostly universities have already developed CubeSats and
launched them with a mixed record of successfully orbited and failed missions. With almost thirty
CubeSats in orbit and over one hundred teams actively designing CubeSats, the CubeSat concept
has been greatly successful in providing access for space research.
2.1.1
QB50
A single CubeSat is too small to carry many or specialized sensors for significant scientific
research. However, when combining in a network a large number of CubeSats with identical sensors, besides the addition educational value, fundamental scientific questions can be addressed,
which otherwise would be inaccessible. The launch of a set of CubeSats have been under discussion in the CubeSat community for several years, but until now no university, company or space
agency took the initiative to create and coordinate a network that big. The problem of reliability
of the CubeSat is not an issue because the network can still achieve the objectives of his mission,
even if any CubeSats fail.
QB50 is an iniative of the Von Karman Institute, ESA and NASA and has the scientific objective of studying in-situ the temporal and spatial variations of a number of key constituents and
parameters in the lower thermosphere (90-320 km), using a network of 50 double CubeSats, separated by a few hundred kilometres and carrying identical sensors. QB50 will also study the re-entry
process by measuring a number of key parameters during re-entry and by comparing predicted and
actual CubeSat trajectories and orbital lifetimes [9].
Since the atmospheric network mission for in-situ measurements is a pioneering experience,
one of the low-cost solutions and an acceptable risk would be the use of CubeSats or very lowcost satellites. The cost of such a network for industry standards, would be extremely high and
not justified due to the limited lifetime in orbit and traditionally, universities do not have means of
funding and the development of CubeSats are for educational purposes.
Among the 50 satellites of the network, there are 38 CubeSats from 22 European countries.
Portugal have submitted a CubeSat proposal to be one of the satellites to be launched in QB50.
2.2 Project
2.2
7
Project
A small satellite program has started in the Faculty of Engineering of the University of Porto
(FEUP) regarding the development of a CubeSat in partnership with the Portuguese company
TEKEVER, GAMA-Sat, and an Earth Reentry Capsule (ERC). It is an academic project composed of graduate and PhD students and professors from different engineering fields (Electrical,
Industrial, Mechanical, etc.), none of which is directly related to aerospace sector.
The idea came with the challenge made by the Education Office of ESA, that worked previously with FEUP in the STRAtospheric PLatform EXperiment (STRAPLEX) program. The aim
was to carry European students’ experiments to the stratosphere using atmospheric balloons[10].
As state previously, this it can be considered an innovative project since, in the Portuguese
context the Portuguese aerospace industry is residual, and there has only been one successful
Portuguese satellite operation, PoSAT-1[6].
2.2.1
Project’ Vision and Objectives
The main goal is to show the capability of a Portuguese student team to perform the reentry of a
capsule trough the Earth atmosphere and the design and implementation of a CubeSat. This offers
the opportunity for students to apply their engineering qualifications to their common interest in
space.
Each of these missions and its goals will be described below in more detail.
2.2.2
GAMA-Sat
The name GAMA is a tribute to the famous Portuguese navigator Vasco da Gama who, during
the Fifteenth Century, discovered the maritime route from Europe to India.
GAMA-Sat will serve for both technology demonstration and scientific purposes. The technology demonstration will focus on the usage of Software Defined Radio (SDR) to establish inter
satellite links and support adhoc networking, range and attitude determination applications. These
capabilities will be used to serve the scientific purpose of calculating the differential evolution of
atmospheric drag between CubeSats [2].
The team will develop an innovative CubeSat transceiver module based on SDR technology,
capable of performing communications and navigation by GNSS (Global Navigation Satellite System). This device will be simultaneously a telecommunications transceiver and a GNSS receiver,
supporting multiple capabilities and applications, such as [2] :
• VHF, S-band and GNSS waveforms in a unique HW platform;
• Inter satellite adhoc networking capabilities, allowing each CubeSat to become a node in a
mobile adhoc network;
• GNSS, for receiving both GPS and GALILEO SIS (Signal in Space);
8
Case Study Presentation
• Range and attitude determination through the VHF Omni-directional radio Range (VOR)
principle (generating VHF waveforms that will be transmitted by patch antennas on each
face of the CubeSat).
As previously mentioned, the project represents the work of different people, integrated into
a single architecture. It has evolved from a simple control into a multilevel control architecture
depicted in Figure 2.1.
Figure 2.1: First draft of VORSat/GAMA-Sat architecture [1]
The satellite can be functionally broken down into the blocks described in 2.1. The four
main parts are the Electrical Power Subsystem, the Command Subsystem, the CW Beacon and
the communication system, one of the subjects of this thesis, more specifically the GPS. This
methodology was taken into consideration for the CubeSat to determine its position when were
necessary to adjust its trajectory due to the drag. The signal transmission depends on the available
energy and satellite location, and therefore the satellites will have a combination of solar panels
and antennas to maximize energy.
The Electrical Power Subsystem is composed by two power sources (the solar cells and the
secondary batteries), a MPPT unit and auxiliary circuits to perform power regulation and battery
charging (not included in the drawing).
The Command Subsystem is composed by a microcontroller, by its battery (primary battery)
and by a set of switches this microcontroller actuates to turn operation of the CW Beacon and
Communications Subsystems on and off.
2.2 Project
9
The CW Beacon subsystem is composed by a frequency synthesizer, a power amplifier and
the respective antennas. It will operate in the Amateur Radio band of the 1296.075 MHz reserved
for CW beacons.
The Communications Subsystem is the heart of the satellite. Its microcontroller is switched on
only when enough power is available to sustain its operation for a significant amount of time. Once
in operation, one of its tasks is to control de operation of the GPS receiver. The microcontroller
will seek position and velocity data from time to time in order to maintain and update the estimate
of its orbit (ephemeris). This ephemeris is stored in non-volatile memory and update periodically.
To measure the satellite attitude from the ground, the system is based on a set of RF signals transmitted from orbit. The idea is to combine multiple signals and antennas in such a way
that certain modulated information depends on the direction from which the signals are received.
Such information is coded in the form of signal phases, allowing to compute the satellite attitude
with one degree of expected accuracy. (A similar approach is used in the VOR - VHF the name
VORSat.)
The satellite will be travelling along its orbit and transmitting signals in all directions. Knowing the position of the satellite relatively to a ground station, by measuring the phase differences
of signals transmitted by different antennas on the faces of the satellite and their evolution as the
satellite passes by the ground station coverage area, it is possible to compute the satellite attitude
from the ground. Three orthogonal rings of antennas are necessary for this purpose, which implies
having 3 individual antennas per face,(see Figure 2.2 ). Other antennas are also required, for GPS
(which must be omnidireccional) and for a simple localization beacon.
Figure 2.2: GAMA-Sat Structure [1]
The transmitted signals are planned for the 2.4 GHz ISM band, using a bandwidth of just
15 KHz. The satellite will be completely solid-state, as the direction dependent signals will be
obtained by applying beam-forming techniques, through a combination of the several antennas.
10
Case Study Presentation
The ground-station requirements are particularly simple: a 2 DOF parabolic dish with a 2.4 GHz
feed and LNA,which is a combination of a tunable down-converter to base-band, a sound card and
a computer running software to demodulate the signals.
There will be a study for the adjustment of the drag coefficient, using the control loop with
feedback from GPS, in order to make the re-entry into the desired spot. In this particularly case
the re-entry will over Portugal, splashdown on the North Atlantic. The idea is to adjust the drag
coefficient by extending and retracting light pannels.
2.2.3
VORSat
This mission main goal is to successfully accomplish the reentry and recovery of an ERC.
Atmospheric reentries have been achieved since the mid Twentieth century, but it has not been
attempted at such a small scale.
Reentry is divided into three sub steps: adjusting friction in the higher atmosphere to adjust expected area of impact, surviving reentry during the deceleration phase and having a safety landing
and recovery. Figure 2.3 depicts the three steps described.
Figure 2.3: Capsule phases. Courtesy of João Gomes
To ensure the surviving reentry during the decelaration phase, it is needed to obtain a light
weight vehicle, Figure 2.4, therefore the structure will be mainly composed of Carbon Fiber Reinforced Polymer. Due to the high costs and use restrictions, the option for a non-ablative thermal
protection system (TPS), such as a Reinforced Carbon Carbon, has been discarded. It was decided
to use a light ablative material, preferably composed of cork, a considerably important product in
the Portuguese economy [2].
For the recovery of the ERC it is necessary to know the entry point and its flight path. The
prediction of the flight path will be achieved through a specially developed algorithm that uses
sporadic GPS observations. After the entry phase, the position of the capsule will be sent to the
ground stations (at least one stationary and several mobile) using, as the main communications
2.2 Project
11
channel, the ARGOS system, due to its low cost and high reliability. This system is commonly
used to track animals in their migrations[2].
The landing point is an essential part of the mission as it can determine the success or failure
of the recovery operation. Two main options are currently under analysis: splashing down in an
ocean or landing in a desert. These options are being considered due to the low population and
building density, reducing the possibilities of collateral damages. In order to minimize the force
acting upon landing, parachutes will be deployed by the capsule [2].
Due the small size of the capsule there are some constrains on the equipament placed onboard. This will, in turn, depend on the maximum dimensions allowed, the state of development
of the technology that is to be tested, space conflicts between different payloads, energy storage,
oscillation damping system, etc., Figure 2.5, [2].
Figure 2.4: Capsule Structure.
Courtesy of João Gomes
Figure 2.5: Interior of the Capsule. Adapted
from [2]
The battery also has some obstacles to the project due to its limited capacity. The algorithms
mentioned above running on the microcontroller, the GPS, the communications and the flash used
to facilitate the location of the capsule, consume energy. Therefore the algorithms must be able to
adjust its communications within the ground based on the values of the available battery. In the
following figure we can observe the level of consumption of each component:
Figure 2.6: Batery Charge [1]
12
Case Study Presentation
As in the CubeSat, the capsule will have a system of energy generation to maximize the signal
transmission.
2.3
Summary
The chapter’s main intent was to briefly present the project that is being developed in FEUP,
integrating it in the context of CubeSats and then introducing its objectives and its problems,
being this the motivational basis for the work described throughout this thesis. The project’s
current architecture was described in a brief manner as well as the different modules that make it
up. The project’s work description was complemented with figures and citations taken from the
project’s available literature. Moreover, some considerations on the project’s current development
status and its evolutional path were made. The following chapters will describe the research of the
literature and the process used to design, implement and integrate the navigation system.
Chapter 3
Literature Review
The intent of this chapter is to introduce the different scientific themes and trends relevant
to this thesis and the developed work that are related to the fields of artificial satellites and its
navigation system. The chapter is divided into 2 sections, each one presenting the state of art,
works and concepts related to a specific field.
3.1
Satellites
As already stated, the study of space began following World War II. Both the United States and
the Sovietic Union began researching the feasibility of attaching warheads to long-range rockets
capable of crossing half way round the world. These weapons were eventually called intercontinental ballistic missiles or ICBMs. They could be equipped with conventional or nuclear warheads. At the end of World War II, the U.S. had introduced the nuclear warfare age by dropping
atomic bombs on Japan [11]. However, it was in 1957 that the Sovietic Union launched the first
artifical satellite, the Sputnik 1 [12].
Sputnik was launched on October 4th and only with the issuance of a bip-bip, tuned by any
amateur radio in the frequencies between 20.005 and 40.002 MHz, began with the space age of
communications. It was at a height about 250 km (150 miles) of Earth and orbited for six months
before falling. Sputnik helped to identify the density of high atmospheric layers through measurement of its orbital change and provided data on radio-signal distribution in the ionosphere[12].
Since the first satellite was put into orbit, numerous countries launched their own artificial
satellites and several of them are currently under development in the last years. In Figure 3.1 it is
possible to see the development of satellites in the early stages [13].
Among several of the satellites presented in the Figure 3.1 , it is the satellite that carried for
the first time a living passenger into an orbit (Sputnik -2), the first human spaceflight, the first Portuguese satellite, PoSat-1 and the emergence of small satellites. About 75% of satellites launched
into space since 1957, had military purposes but nowadays satellites are, in generally, used for
13
14
Literature Review
Figure 3.1: Evolution of satellites over the years
radio communication and television signal. Beyond Earth observation and communications, satellites can be applied for scientific purposes, including observation of the weather, the exploration
of the universe and the data collection on Earth.
With Earth observation satellites in orbit, it is possible to analyse the environmental conditions, through the processes and interactions between land masses, oceans and global atmosphere,
in order to ensure our safety and quality of life. Among other things, they are also capable of
detecting environmental risks in a timely manner and monitor and manage the Earth’s natural resources. However, these types of satellites can be used for purposes of defense and security. The
main applications are the generation of maps, monitoring and tracking of targets and, if necessary,
the destruction of enemy warheads, satellites or other space assets. The first Earth observation
satellite was launched in 1959, Explorer VII, [14].
Communication satellites are used in transmission of digital information, specifically in the
civilian world. These are artificial satellites stationed in space with the purpose of telecommunications. The first and historically most important application of satellite communication was
intercontinental long-distance telephony. But when television became the main market, applied
to simultaneous delivery of relatively few signs broadband to many receivers with a more precise
match to the capabilities of geosynchronous comsats radio operators. After the 1990s, the technology of satellite communications has been used as a way to connect to the Internet via broadband
connections of data. This can be very useful for users who are located in very remote areas,
and can not access a broadband connection. Communication satellites are also used for military
communications applications, such as the Global Command and Control Systems[14].
The satellite navigation is a space-based radio positioning system, which includes one or more
satellite constellations, augmented as necessary to support the intended operation, and that provides 24 hours and three-dimensional position, velocity and time information to suitably equipped
3.1 Satellites
15
users anywhere on, or almost all the surface of the Earth. Users have sufficient accuracy and completeness of information to be usable for critical navigation applications. The GPS system is the
central element of the first satellite navigation system widely available to civilian users. They are
satellites that send out radio signals to mobile receivers on Earth, allowing the determination of
the accurate location. The direct reception of the signal from GPS satellites, combined with an
electronic increasingly developed, allows the GPS system to determine the position with an error
of a few meters in real time [14].
Weather forecast uses a variety of observations from which to analyses the current state of the
atmosphere. Since the launch of the first weather satellite in 1960, TIROS-1, global observations
have been possible, even in the remotest areas. [14].
Scientific research satellites provide applications such as earth science, marine science, and
atmospheric research [14].
To perform the functions mentioned above, satellites need to be semi computer-controlled
systems. They have to be able to realize many tasks, such as power generation, thermal control,
telemetry, attitude control and orbit control.
Depending on the purpose of the satellite and based on its height above the Earth, there are
several orbits.
• Lower earth orbit (LEO) satellites operate in orbits of around 500 km to 1,500 km above
the Earth’s surface – much lower than traditional communications satellites – which brings
them into frequent radio contact with ground stations. LEOs are used for a variety of civil,
scientific and military roles including Earth observation, radar, optical, telecoms and demonstrator. [4]
• Medium earth orbit (MEO) refers to an altitude below 35,786 km (geostationary orbit)
and above the altitude of Low Earth Orbit (LEO). Medium Earth Orbit enables a satellite
provider to cover the earth with fewer satellites than Low Earth Orbit, but requires more
satellites to do so than geostationary orbit. The most common use for satellites in this
region is for navigation, such as the GPS and Glonass constellations (Russian counterpart
to the United States GPS system). [4]
• Geostationary (GEO) Orbits is higher than an altitude of 35,786 km and they are so named
because they are placed in an orbit over the equator so that the satellite has a rotation period
equal to Earth’s, Objects in Geostationary orbit revolve around the earth at the same speed
as the earth rotates, means, 24 hours. Communications satellites and weather satellites are
often given geostationary orbits, so that the satellite antennas that communicate with them
do not have to move to track them, but can be pointed permanently at the position in the sky
where they stay. [4]
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Literature Review
• Highly Elliptical Orbits (HEO) is a geocentric orbit whose apogee lies above the geostationary orbit (35,786 km). HEO is mainly perturbed by the Earth’s oblateness and by gravitational attraction of the Sun and Moon. They are popular orbits for Earth magentospheric
measurements and astronomical observatories. [4]
3.1.1
Orbits
The applications mentioned above contains some restrictions on the range of launch, the need
to have or not a powerful amplifiers for a successful transmission, etc., which leads to a systematic
choice on the orbit where the satellite will be placed. The great majority of satellites are launched
into lower orbits at altitudes of 500-1500 km. Below that level, a satellite’s orbit would be quickly
decay due to the resistance of the Earth’s atmosphere, thus restricting extremely low-altitude orbits
to short-term ballistic missions or powered trajectories.
In any circular orbit, the centripetal force required to maintain the orbit is provided by the
gravitational force on the satellite. To calculate the geostationary orbit altitude, one begins with
this equivalence, and uses the fact that the orbital period is one sidereal day.
Fc = Fg
(3.1)
By Newton’s second law of motion, we can replace the forces F with the mass m of the object
multiplied by the acceleration felt by the object due to that force:
m × ac = m × g
(3.2)
We note that the mass of the satellite m appears on both sides — geostationary orbit is independent of the mass of the satellite. So calculating the altitude simplifies into calculating the
point where the magnitudes of the centripetal acceleration required for orbital motion and the
gravitational acceleration provided by Earth’s gravity are equal.
The centripetal acceleration’s magnitude is:
ac = w 2 × r
(3.3)
where w is the angular speed, and r is the orbital radius as measured from the Earth’s center of
mass.
The magnitude of the gravitational acceleration is:
v2 K × ME
K × ME
=
=
2
d
d
(RE + h)2
(3.4)
3.1 Satellites
17
From the expression 3.1.1 the speed at which the satellites travels in GEO orbit can easily be
obtained.
v=
!
K × ME
(RE + h)2
(3.5)
The angular speed w is found by dividing the angle travelled in one revolution by the orbital
period (the time it takes to make one full revolution: one sidereal day).
w =≈
w =≈
2×π
dayssolar
×
T
dayssideral
(3.6)
2×π
365.24
×
3600 × 24 366.24
(3.7)
The resulting orbital radius is the sum of the height of GEO satellite, 35786 km, and the mean
radius of the earth, 6371 km. The GEO orbital speed (how fast the satellite in GEO is moving
through space) is calculated by multiplying the angular speed by the orbital radius:
v = w × (RE + h) = 42213000 = 3.057km/s
(3.8)
Therefore, the velocity of a LEO satellite is:
"
√
RE + hGEO
42213
vLEO = vGEO × √
= 3061 ×
= 7.719km/s
6650
RE + hLEO
(3.9)
LEO satellites travel at 7.7 km/s while Earth travels at 3 km/s, making a complete revolution
around the Earth in about 90 minutes.
On the other hand, above 1500 km altitude, the satellite launchers must be more complex due to
the distance they want to achieve, that causes a higher cost than the launch of LEO. More powerful
amplifiers are required for a successful transmission and with the increasing of the distance, it
causes a delay of about 0.5 s in the signal. This delay brings more complex problems in the
protocols of verification and correction of incorrect data. The system has to stop transmitting data
communication while awaiting the response of the receiver that there is no error in the sent data.
[15]
However, the LEO systems have a few disadvantages. Once they are closer to Earth, LEO
satellites have to compensate for Doppler shifts cause by their relative movement and it is needed
a lot more satellites to cover the Earth’s surface. Due to the atmospheric drag, causing gradual
18
Literature Review
orbital deterioration, LEO have much shorter life span than MEO and GEO. They live for 5 to 8
years. [15]
3.2
Navigation System
Regardless of the level of autonomy that the satellites can reach, any man-made spacecraft
would have no value if it was unable to locate it and communicate with it. Furthermore, the
navigation system is concerned with the mathematical and physical description of artificial satellite
orbits, as well their control. Many of the satellites previously mentioned, need an active control
of their orbit in accordance with the specific mission requirements. Navigation is, therefore, an
essential part of satellite operations. It comprises the planning, the determination,the prediction
and the correction of satellite’s trajectory in line with the established mission goals.
To obtain measurements related to the instantaneous position of a satellite is necessary to
appeal to the using of the telemetry, tracking, command and communication (TTCC) system. It
provides meanings for the monitoring and the controlling of the satellite in orbit.
3.2.1
Measuring from the Earth
3.2.1.1
Tracking
The tracking subsystem provides facilities by the satellite orbit which can be determined. The
command subsystem provides the meanings from which the satellite is controlled and the communications subsystem provides links between the satellite and the earth station, for the transmission
of the necessary signals for these various purposes.
A variety of tracking systems may be used to obtain measurements related to the instantaneous
position of satellite or its rate of change.
1. Visually Tracking
One easy way to view a satellite in orbit is the unaided eye. That’s because, when the
satellite is near the Earth, it reflects the Sun’s light off their surfaces toward the observer and
becomes possible to follow it.
2. Radar Tracking
Most of these systems are based on radio signals transmitted to or from a ground antenna
3.2.
Common radio tracking systems are able to perform angle measurements by locating the
direction of a radio signal transmitted by a satellite. The resolution of these measurements
depends on the angular diameter of the antenna cone, which is determined by the ratio of
the carrier wavelength to the antenna diameter. Distance and velocity information can be
obtained by measuring the turn-around delay or Doppler-shift of a radio signal sent to the
spacecraft and returned via transponder.
3.2 Navigation System
19
Figure 3.2: Ground Station. [3]
3. Satellite Laser Ranging
Satellite laser ranging (SLR) systems provides highly accurate distance measurements by
determining the turn-around light time of laser pulses transmitted to a satellite and returned
by a retro-reflector. Depending on the distance and the resulting strength of the returned
signal, accuracies of several centimetres may be achieved. SLR is mainly used for scientific
and geodetic mission that requires an ultimate precision.
3.2.2
3.2.2.1
Measuring from Space
Telemetry
The telemetry subsystem supplies measurements of various parameters, such as pressures and
temperatures, on on-board equipment and converts them into electrical signals for transmission.
The data is send to the earth station, which is responsible for satellite management, and receives
the earth control station’s commands to perform equipment operation adjustments.
In our case, the 34 channels used for telemetry are being designed to operate in QPSK but
most of these measurements produce signals that vary slowly, the ideal it would be constant, and
requires only a few hertz of bandwidth so the signals can be multiplexed into a single carrier before transmission. Two methods of multiplexing are used, either frequency division multiplexing
(FMD) or time division multiplexing (TDM).
In FDM, the signal voltage produced by a sensor usually varies the output frequency of a
voltage-controlled subcarrier oscillator so that the oscillator frequency deviation represents the
amplitude of the output voltage of the sensor. For each sensor output signal, a different subcarrier
frequency is used. Subsequently, the outputs of all the subcarrier oscillators are combined and fed
into a single amplifier.
20
Literature Review
In TDM, the sensor outputs are fed into a commutator which samples in sequence the output of
each sensor several times per second, producing a series of pulses which represent the information
content. Time division multiplexing is only used with particularly low-frequency signals due to its
inherent limited frequency response.
In QPSK the phase of each one being maintained during each burst. Each frequency slot can
therefore encode 2 bits. If 32 of the slots carry information (and the remaining 2 are used for
checksum), then 64 bits of information can be sent with each burst. If the same information is
transmitted in the 6 faces of the satellite, then 20 different double words can be sent every second
(each face transmits different data through the X and Y channels). Therefore, a throughput of 1280
bps is achieved with this modulation method. As a detail, the phases of the two frequency slots
used for attitude are valuable as reference for the phases of each of the telemetry frequency slots
[3].
This throughput is more that enough for the satellite to continuously send the required telemetry data. This is composed by timing information (derived from the GPS receiver), by computed
ephemeris of the satellite and satellite status information.
1. Gobal Positing System (GPS)
The Global Positioning System (GPS), which was developed to meet military needs of the
Department of Defense but now is used in every day life, is a radio based navigation system
that gives three dimensional coverage of Earth 24 hours a day in any weather condition.
Going even further, GPS ranging signals offer the opportunity to obtain position measurements on-board a satellite, completely independently of a ground station.
It was decided that this system will be used for this mission because it is less expensive, it
requires less antennas, is less prone to orbit error, and is less labor intensive for scheduling,
collecting, and transferring data. GPS is the most accurate navigation device that functions within LEO, which is where the cubesat will be operating. A GPS receiver is a highaccuracy navigation device that obtains amplified signals from a GPS antenna (which obtains signals from GPS satellites) and outputs data in a coordinate format. The solution of
the GPS receiver includes the cubesat’s predicted position above the earth’s surface, velocity
vector, time, and date.
(a) GPS Limitations
i. Available Energy
Energy is the most critical factor for the use of GPS. The Electrical Power Subsystem is composed by two power sources (the solar cells and the secondary batteries), a MPPT unit and auxiliary circuits to perform power regulation and battery
charging (not included in the drawing). This subsystem is autonomous relatively
to the others since its operation state does not depend on any command or state
of other subsystems. The amount of energy available in the batteries can continuously be estimated by the Command Subsystem.
3.2 Navigation System
21
Using an area of 50 cm2 of solar panels with 25% efficiency per satellite face
and considering that the satellite is 50% under sunlight and 50% under darkness,
each face under random orientation, provides an average of 2.5 Watts of power
along time. This is clearly not enough to operate continuously the system, which
is especially true for the transmission of the radio signals [3].
Therefore, the satellite will selectively switch the most energy demanding devices
on and off according the energy level present in the batteries. 3.3 presents the
method provided for the energy management. All switching will have hysteresis
to avoid short duration cycles [3].
Figure 3.3: Equipment ON/OFF state depending on battery charge. [3]
A minimal energy level is required to switch the Communications Subsystem
microcontroller on. Until then, only the Command Subsystem microcontroller
is consuming a very limited amount of energy from its own long-life battery.
When operating, the main microcontroller decides to switch the GPS receiver
only when in need to obtain a position and velocity measurement to update the
satellite ephemeris and its clock. The energy consumption of this microcontroller
plus the GPS receiver can be designed to be significantly less than 2.5 Watts (less
than 0.5 Watts, in fact). Therefore, the satellite average energy accumulated in
the secondary batteries will have a global tendency to increase until reaching the
second level.
ii. Doppler Effect
The principle of position determination by GPS and the accuracy of the positions
strongly depends on the nature of the signals. When Sputnik was been monitoring through radio transmissions, a team of U.S. scientists led by Dr. Richard B.
Kershner discovered that,the frequency of the signal being transmitted by Sputnik was higher as the satellite approached, and lower as it continued away from
them. They realized that because they knew their exact location on the globe, they
could pinpoint where the satellite was along its orbit by measuring the Doppler
distortion [12] .
22
Literature Review
In other words, the Doppler effect causes a change in the frequency of emitted
waves produced by motion of an emitting source relative to an observer. Therefore, when we want to track the satellite it is necessary to take account the Doppler
effect. The change in frequency will be given by:
fd =
v
v
v
= c = f×
λ
c
f
(3.10)
where,
• v is the velocity of the satellite;
• c is the speed of wave;
• λ is the wavelength of the transmitted wave in the reference frame of the
source.
The satellite will transmit at a frequency of 1.575 GHz and will have a velocity
on Earth of 3 km/s 3.1.1 and approximately 9 km/s between the GPS and LEO
(moving in opposite directions).3.1.1.
The change in frequency on earth due the doppler effect it will be :
fd = 1.575 × 102 =
3 × 103
= 15KHz
3 × 108
(3.11)
fd = 1.575 × 102 =
9 × 103
≈ 50KHz
3 × 108
(3.12)
And on LEO:
As stated earlier, we can observe that the Doppler effect will be 3 times more on
LEO satellites than GEO. Therefore the maximum dynamic environment which
the carrier will be subject to, in LEO, is a Doppler shift of 100 Hz/s.
iii. COCOM restrictions
CoCom is an acronym for Coordinating Committee for Multilateral Export Controls. CoCom was established by Western bloc powers in the first five years after
the end of WWII, during the Cold War, to put an arms embargo on COMECON
(Warsaw Pact) countries [16].
Immediate access to satellite measurements and navigation results is disabled
when the receiver’s velocity is computed to be greater than 1000 knots (approximately 1,151 mph or 1,852 km/h), or its altitude is computed to be above 60
thousand feet (18,000 meters). The receiver continuously resets until the COCOM situation is cleared [16].
In GPS technology, the phrasing "COCOM Limits" is also used to refer to a limit
3.2 Navigation System
23
on how high and how fast a GPS will operate and a limit placed to GPS tracking devices that should disable tracking when the device realizes the restriction
mentioned above. This was intended to avoid the use of GPS in ICBM-like applications.
These restrictions make the prices of receivers become more competitive, which
means the price increases. Depending of the manufacturers they can apply this
limit literally (disable when both limits are reached), other manufacturers disable
tracking when a single limit is reached.
iv. Noise
When a GPS satellite passes across the sky, the signal that is sent out has to travel
through the atmosphere, and occasionally it bounces off the land before getting
to a GPS instrument. Both the atmosphere and the land create problems called
multi-path noise. This noise is suppressed as much as possible by the instruments,
but it is never completely eliminated, and it leads to error in high-precision GPS
positioning data. One solution is the use of Kalman Filter.
Since its introduction in 1960, the Kalman filter has become an integral component in thousands of military and civilian navigation systems. This deceptively
simple and recursive digital algorithm has been an early-on favorite for conveniently integrating (or fusing) navigation sensor data to achieve optimal overall
system performance [17] .
The Kalman filter is a set of mathematical equations that provides an efficient
computational (recursive) means to estimate the state of the process, in a way that
minimizes the squared error of the mean. The filter is very powerful in several
aspects: it supports estimations of the past, present, and even future states, and
when the precise nature of the modeled system is unknown.
Its purpose is to use measurements that are observed over time that contain noise
(random variations) and other inaccuracies, and produce values that tend to be
closer to the true values of the measurements and their associated calculated values.
Because of its optimum performance, versatility, and ease of implementation, the
Kalman filter has been especially popular in GPS/inertial and GPS stand-alone
devices.
3.2.3
GPS Antenna
To track a satellite in LEO orbit with 71 degrees of expected inclination, the antenna tracking
system must be capable of performing a full sky sweep. The existing antenna controller is able to
perform 400 degrees of azimuth rotation and 100 degrees of elevation with an average speed of
four degrees (TBC) per second with an accuracy of 0.5 degrees, which is compatible with LEO
satellite tracking. With a carrier frequency of 1.575 GHz, the maximum frequency shift caused by
24
Literature Review
Doppler Effect is 100 KHz. This frequency deviation requires an active control of the receiver in
order to keep a proper reception of the transmitted signals, which is also available [3].
3.2.4
Communication
Communication is an essential part of navigation. Without the communication it was not
possible to monitor the satellite in its LEO orbit.
Its microcontroller is switched on only when enough power is available to sustain its operation
for a significant amount of time. Once in operation, one of its tasks is to control de operation of
the GPS receiver. The microcontroller will seek position and velocity data from time to time in
order to maintain and update the estimate of its orbit (ephemeris). This ephemeris is stored in a
non-volatile memory and is updated periodically. This allows operating the GPS in low power
mode most of the time (or even switching it off, this detail to be defined later; nevertheless, the
GPS receiver is turned off together with the main microcontroller)[3].
Most CubeSats use the VHF band and S band (sometimes as secondary downlink frequency).
VHF practical limits data rates ( 9600 bps), while on the S-band high data rates, up to 256 kbps,
are used. The uplink frequencies show a similar distribution, but the S-band is scarce. Also the
data rates are limited, possibly due to the fact that uplinks are used mainly for short commands,
instead of large data packets. It can be stated that the communication skills between CubeSats and
ground are more limited, due to budget link instead of radio technology, from micro-electronics
for high data rates are widely available and are applied in the systems and the potential of ground
station tracking with high gain antennas is limited by the performance of dynamic attitude control
system.
3.2.5
Microcontroller
The Command Subsystem is composed by a microcontroller, by its battery (primary battery)
and by a set of switches this microcontroller acts to turn operation of the CW Beacon and Communications Subsystems on and off. This microcontroller, running failsafe software, consumes a
very limited amount of energy. The non-chargeable battery is designed to sustain the operation of
the microcontroller for at least two years after power up. The operation of microcontroller will
depend on the state of the main switch, which is activated when the satellite is released from the
P-POD. In the limit, this microcontroller can be replaced by comparators and logic systems (to be
defined in due time) [3].
The microcontroller is used as the brain of the satellite. The main tasks of this board are to
handle the commands received from the ground station on the earth, to perform tasks such as run
the algorithms, processing power, reading and storing the sensor values, and to send data back to
the ground station.
The specifications of the microcontroller board are largely determined from the tasks they
have to perform and the physical and electrical resources allocated to the board. Due to limited
space in CubeSats, the microcontroller should eliminates the need for many external components
3.3 Summary
25
required by other systems (analog-to-digital converters, serial communication interfaces, slave
micro-controllers, etc..), while remaining simple to program and implementable. It is also expected that the microcontroller provides a mechanism for issuing commands to instruments and
has the ability to store scientific data and system telemetry before downlink to ground station. The
consumption, the time, budget and besides being able to withstand the harsh orbital conditions,
the GPS receiver cannot be subject to COCOM constraints, which limits the choice of brand and
model.
Since the late nineties, low power microcontrollers have gradually become available. Their
processing performance is better than ever similar or ever better than most of the outdated but
space qualified processors used in satellites and their power consumption is only in the order of
milli-watts. The downside of those processors is their high susceptibility to particle radiation in
space. The risks are sometimes tackled by redundancy on different components and the distributed
command and data handling architecture with multiple microprocessor spread across the satellite
[18].
3.3
Summary
The chapter’s objective was to present the different fields relevant to this thesis. The currently
accepted view of an navigation main systems and functions was presented. Their motivations,
objectives and current implementations were also presented.
Moreover, the middleware technology currently used to enable communications between satellites and ground station was also presented. The description was focused in GPS Systems, although, other technologies that implement solutions for devices and software were also presented.
Throughout this document references will be made to this chapter and to the concepts and
systems it presented. This chapter can thus be perceived as the underlying basis for the developed
work linking it to the advancements in the described fields.
26
Literature Review
References
[1] Bruno Santos Carlos Capela Eduardo Sá João Gomes João Granjo João Silva Pedro Abreu
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presentation.
[2] G. Paredes C.Ramos, J. Gomes. Development of a cubesat and an earth reentry capsule at
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[3] Vorsat proposal.
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[5] National Aeronautics e Space Administration (NASA). Url: http://www.nasa.gov/.
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[13] The Satellite Encyclopedia.
[14] National Academy of Science. "earth observations from space".
[15] Michael Charles.
[16] M. Mastanduno. Economic containment: Cocom and the politics of east-west trade. Cornell
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[17] Eun-Jung Choi, Jae-Cheol Yoon, Byoung-Sun Lee, Sang-Young Park, e Kyu-Hong Choi.
Onboard orbit determination using gps observations based on the unscented kalman filter.
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