ESTCube-1 Space System and Mission Description
Author: ESTCube Team
Edited by Jouni Envall
17.4.2009
CONTENTS
Contents 1
2
3
Introduction
8
1.1
Mission objectives . . . . . . . . . . . . . . . . . . . . . . . . .
9
1.2
Spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
9
Mission plan 2.1
Sequence of events
2.2
Electric solar wind sail
2.3
Tether experiment
. . . . . . . . . . . . . . . . . . . . . . . .
11
. . . . . . . . . . . . . . . . . . . . . .
13
. . . . . . . . . . . . . . . . . . . . . . . .
13
Orbit selection and determination
15
3.1
Keplerian elements
. . . . . . . . . . . . . . . . . . . . . . . .
15
3.2
Formats
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15
3.2.1
Two-line elements . . . . . . . . . . . . . . . . . . . . .
15
3.2.2
AMSAT format . . . . . . . . . . . . . . . . . . . . . .
16
3.2.3
One-Line Charlie Elements . . . . . . . . . . . . . . .
17
3.3
Prediction models . . . . . . . . . . . . . . . . . . . . . . . . .
18
3.4
GPS
18
3.5
Orbit for ESTCube-1 mission
3.6
4
11
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20
3.5.1
Equatorial LEO . . . . . . . . . . . . . . . . . . . . . .
21
3.5.2
Polar LEO . . . . . . . . . . . . . . . . . . . . . . . . .
21
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22
210 Structure (STR)
24
4.1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24
4.2
System members
. . . . . . . . . . . . . . . . . . . . . . . . .
24
4.3
System description
. . . . . . . . . . . . . . . . . . . . . . . .
24
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CONTENTS 4.3.1
Material . . . . . . . . . . . . . . . . . . . . . . . . . .
24
4.3.2
Mechanical parts
. . . . . . . . . . . . . . . . . . . . .
26
Deduced requirements from the Cubesat standard . . . . . . .
27
4.4.1
Dimensional requirements for the satellite structure . .
27
4.4.2
Electrical requirements . . . . . . . . . . . . . . . . . .
28
4.5
Layout proposal . . . . . . . . . . . . . . . . . . . . . . . . . .
28
4.6
Technical results
. . . . . . . . . . . . . . . . . . . . . . . . .
31
4.7
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32
4.4
5
220 Attitude Determination and Control System (ADCS) 33 5.1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33
5.2
System members
. . . . . . . . . . . . . . . . . . . . . . . . .
33
5.3
System description
. . . . . . . . . . . . . . . . . . . . . . . .
33
5.3.1
Deduced requirements from the Cubesat standard . . .
34
5.3.2
Deduced requirements from the payload
. . . . . . . .
34
5.3.3
Considerations about the spinning axis . . . . . . . . .
34
Layout proposal . . . . . . . . . . . . . . . . . . . . . . . . . .
39
5.4.1
System components . . . . . . . . . . . . . . . . . . . .
39
5.4.2
Subsystem operation phases
. . . . . . . . . . . . . . .
39
5.5
Comparison of the three spin-up methods . . . . . . . . . . . .
42
5.6
Momentum wheel power calculation . . . . . . . . . . . . . . .
42
5.6.1
Starting
43
5.6.2
Maintaining the speed
5.4
5.7
Budgets 5.7.1
5.8
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
43
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
43
Average ADCS power consumption during phases
. . .
44
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
44
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CONTENTS 6
230 Electrical Power System (EPS)
47
6.1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . .
47
6.2
System members
. . . . . . . . . . . . . . . . . . . . . . . . .
47
6.3
System description
. . . . . . . . . . . . . . . . . . . . . . . .
48
6.3.1
Deduced requirements from the Cubesat standard . . .
48
6.3.2
Layout proposal . . . . . . . . . . . . . . . . . . . . . .
49
6.3.3
Mass and cost budbet
. . . . . . . . . . . . . . . . . .
49
6.3.4
Power generation during dierent phases . . . . . . . .
50
6.3.5
Discussion . . . . . . . . . . . . . . . . . . . . . . . . .
51
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51
6.4
7
240 Thermal Control System (TCS)
52
7.1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . .
52
7.2
System members
52
7.3
Examples of typical temperature ranges for satellite components 52
7.4
Thermal environment for satellite in space (LEO)
7.5
Mechanisms of thermal energy transfer
7.6
Temperature equilibrium and thermal balance calculations
7.7
Thermal balance calculation of a raw satellite
7.8
7.9
. . . . . . . . . . . . . . . . . . . . . . . . .
. . . . . . .
53
. . . . . . . . . . . . .
54
. .
55
. . . . . . . .
56
7.7.1
Temperature equilibrium of the spacecraft at the sunlight 57
7.7.2
Temperature equilibrium of the spacecraft at the eclipse 58
7.7.3
Summary
. . . . . . . . . . . . . . . . . . . . . . . . .
58
Thermal control mechanisms . . . . . . . . . . . . . . . . . . .
58
7.8.1
Passive control
. . . . . . . . . . . . . . . . . . . . . .
58
7.8.2
Active control . . . . . . . . . . . . . . . . . . . . . . .
59
Conclusions of the preliminary status of TCS . . . . . . . . . .
59
7.10 Future steps and plans
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59
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CONTENTS 8
250 Communications system (COM)
61
8.1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61
8.2
System Members
. . . . . . . . . . . . . . . . . . . . . . . . .
61
8.3
System Description . . . . . . . . . . . . . . . . . . . . . . . .
61
8.3.1
Deduced requirements from the Cubesat standard . . .
61
8.3.2
Deduced requirements from the Payload
. . . . . . . .
62
8.3.3
Layout Proposals . . . . . . . . . . . . . . . . . . . . .
62
8.3.4
Possible problems . . . . . . . . . . . . . . . . . . . . .
63
8.4
9
Link budgets
. . . . . . . . . . . . . . . . . . . . . . . . . . .
63
8.4.1
Downlink budget
. . . . . . . . . . . . . . . . . . . . .
63
8.4.2
Uplink budget . . . . . . . . . . . . . . . . . . . . . . .
63
8.4.3
Beacon . . . . . . . . . . . . . . . . . . . . . . . . . . .
63
8.5
Mass and power budgets . . . . . . . . . . . . . . . . . . . . .
67
8.6
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
67
260 Command and data handling system (CDHS)
69
9.1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . .
69
9.2
System Members
. . . . . . . . . . . . . . . . . . . . . . . . .
69
9.3
Functionality
. . . . . . . . . . . . . . . . . . . . . . . . . . .
70
9.3.1
Deduced requirements from the Cubesat standard . . .
70
9.3.2
Layout proposal . . . . . . . . . . . . . . . . . . . . . .
71
9.3.3
Technical requirements . . . . . . . . . . . . . . . . . .
72
Components . . . . . . . . . . . . . . . . . . . . . . . . . . . .
72
9.4.1
Microcontroller selection . . . . . . . . . . . . . . . . .
72
9.4.2
8-bit Atmel AVR Family . . . . . . . . . . . . . . . . .
73
9.4.3
Memory device selection
. . . . . . . . . . . . . . . . .
74
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
74
9.4
9.5
Budgets
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CONTENTS
9.6
9.5.1
List of Components . . . . . . . . . . . . . . . . . . . .
74
9.5.2
Power usage characteristics . . . . . . . . . . . . . . . .
74
9.5.3
Mission Phases During Orbiting . . . . . . . . . . . . .
75
9.5.4
Mass budget . . . . . . . . . . . . . . . . . . . . . . . .
77
9.5.5
Data budget . . . . . . . . . . . . . . . . . . . . . . . .
78
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
79
10 270 Payload (PL)
80
10.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . .
80
10.2 System Members
. . . . . . . . . . . . . . . . . . . . . . . . .
80
. . . . . . . . . . . . . . . . . . . . . . . .
80
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
81
10.3 System description 10.4 Hardware
10.5 Mass and power budget
. . . . . . . . . . . . . . . . . . . . .
83
10.6 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
84
11 Conclusion
85
References
87
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1 INTRODUCTION
1 Introduction
The Estonian Student Satellite project started in the summer of 2008 at Tartu University with the objectives for promoting space, being an invaluable educational tool for science, technology, engineering, mathematics (STEM) subjects and giving students hands-on experience on developing space technologies. In addition, our further goal is to foster the development of Estonian space and high-tech industry by training experts and disseminating knowledge about space technologies. [1] In addition the satellite project is expected to have a signi cant role in educating and inspiring the general public and improve their awareness of space research. In the mean time the project has grown into a full-scale international collaboration with participating students from Tartu University, Tallinn University of Technology and Estonian Flight Academy. By now students from University of Surrey, UK and International Space University, FR are also taking part in the development and our project has had the honour of hosting a visitor from Aachen University of Applied Sciences, DE. It has been a challenging exercise to start work in such a geographically disperse environment, but we have overcome all the di culties and are fully committed on reaching our goal to launch the satellite. The international dimension is also added to the project as the development of the payload is carried out in international collaboration, with Finnish Meteorological Institute (FMI) being the prime payload development coordinator. Currently, at the end of phase 0, seven bachelors and masters theses are being written on the subject of ESTCube-1. ESTCube-1 is based on the Cubesat standard.[2] Cubesat is a satellite standard developed by the California Polytechnic State University (Cal Poly) and the Stanford University.
The standard is primarily destined to serve
the needs of the student satellite teams.
The basic Cubesat, or the single
Cubesat, is a cube whose volume is 1 litre and whose mass must not exceed 1 kg. The Estonian Student Satellite team is very thankful towards its supporters: Skype, Estonian Information Technology Foundation and collaborators Estonian Physical Society, Vangelis Space LLC, Tartu Observatory. This document summarizes the work performed during ESTCube-1 phase 0 study and includes status, main results of preliminary analysis and the
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1 INTRODUCTION conceptual baseline design for space system con guration, mission and operations and payload design.
1.1
Mission objectives
The mission objectives of the ESTCube-1 mission are:
To test the deployment of a 10 meter Hoytether as a part of the development work of the Electric solar wind sail.
1.2
To measure the electric sail force, interacting with the tether.
Spacecraft
ESTCube-1 will be a single Cubesat. Its size will be 100×100×113,5 mm
3
and it's mass must not exceed 1 kg, as stated in the standard. The satellite is divided into several subsystems, each responsible of a speci ed task. The subsystems are
210 Structure (STR)
220 Attitude Determination and Control System (ADCS)
230 Electrical Power System (EPS)
240 Thermal Control System (TCS)
250 Communications System (COM)
260 Command and Data Handling System (CDHS)
270 Payload (PL)
The tether experiment requires ESTCube-1 to spin around its axis, the tether will then be deployed with the help of the centrifugal force. This sets great demands for the ADCS. Once the satellite has been ejected from the launch vehicle into orbit, it has to be detumbled rst. Once this has been achieved, the ADCS should commence the spinning in a controlled way.
Once the
desired angular velocity (approximately 1 revolution per second) has been achieved, the tether is deployed with the motorized spool.
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The successful
J. Envall (editor)
1 INTRODUCTION deployment of the tether will be veri ed with the measured decrease in the angular velocity. In addition, an onboard camera is used to take a picture of the deployed tether. Finally, ADCS will repeat the detumbling process, so that the satellite will be nadir pointed, and the camera can take pictures of Earth.
The gathered data will be stored in the CDHS and transmitted
to Earth. Also the COM system has to face challenges due to the spinning. The antenna system must be designed in such a way that both uplink and downlink connections are working at all times. This is also in uenced by the choice of the spinning axis. Each subsystem is designed by a designated group of students. The following sections give more detailed information about the subsystems.
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2 MISSION PLAN
2 Mission plan
As stated in section 1.1, the objectives of the ESTCube-1 mission include the successful deployment of a conductive space tether, and measuring the electric sail eect. The following sections give a more detailed description of the events that take place, starting from the moment the satellite is deployed from the carrier rocket, and reaching to the end of the mission life.
2.1
Sequence of events
Once the satellite has been ejected from the launch pod of the carrier rocket, the following sequence is commenced.
1. Detumbling 2. Nadir pointing 3. Taking photographs 4. Commence spinning 5. Tether deployment 6. High voltage to tether 7. Measure electric sail force
If the payload system will include the option of retracting the tether, then the following events will follow the ones listed above.
8. Retract tether 9. Detumbling 10. Nadir pointing
Once the experiments have been performed (regardless of whether the phases 810 took place), there is still on optional phase, which will end the mission.
11. Electric sail deorbiting (optional)
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2 MISSION PLAN Detumbling
Stoppping the initial spinning, right after the deployment
from the carrier rocket, is called detumbling. See section 5.4.2 for details.
Nadir pointing
In this phase the attitude determination and control sys-
tem (ADCS) of the satellite will orient the satellitte in such a way that the camera is constantly viewing the Earth.
The orientation needs to be con-
stantly checked and corrected throughout the orbit period.
Photographs
Once the camera is viewing the Earth, pictures are taken
and submitted to the ground station.
The photographs should be taken
before the tether experiment, since the possible problems with the experiment (e.g. in retracting the tether) may otherwise interfere with this objective.
Spinning
Once the orientation of the satellite is suitable (see section 2.3
for further discussion), the ADCS starts to spin the satellite. The purpose of the spinning is to create the centrifugal force needed to deploy the tether.
Tether deployment
Once the satellite has reached the desired angular
velocity, the tether is deployed. The deployment is done with the use of the centrifugal force. The succesful deployment is observed from the telemetry data, recorded by the onboard sensors. Also, a photograph is taken of the tether. It should be noted that when this phase has been reached, the rst mission objective has been met.
High voltage to tether
A high voltage is applied to the tether.
The
purpose of the voltage is to cause the tether to interact with the surrounding plasma.
Measure electric sail force
Once the voltage has been applied, the tether
is being in uenced by the Lorentz force, initiating from the Earth magnetic eld, and the electric sail force, initiating from the plasma surroundings. The eect of these forces can be observed as perturbations in the satellite spinning. Section 2.3 gives a more detailed description of the experiment.
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2 MISSION PLAN Retract tether
The tether is retracted by reeling it in, using the dedicated
motor of the payload subsystem. This step is optional.
Detumbling
Detumbling is performed as in the beginning.
This step is
optional.
Nadir pointing
Nadir pointing is performed as in the beginning.
This
step is optional. After the nal nadir pointing, photographs can be taken, or a new sequence of tether experiment can be commenced.
Electric sail deorbiting (optional)
We might also continue operating
in the electric sail mode for a time long enough to observe the measurable lowering of the satellite orbit. This would demonstrate the electric sail deorbiting.
2.2
Electric solar wind sail
The electric solar wind sail (the electric sail) is a propulsion innovation, made at the Finnish Meteorological Institute in 2006. The system will get its thrust from the solar wind plasma. The sail itself consists of long, thin, conductive tethers, which are held at a high potential. [4, 5, 6] The electric sail is predicted to have a wide range of applications, including travels to the far regions of the solar system, traveling to the interstellar space, and maneuvering in the solar system. In the proposed electric sail vessel the sail would open using the centrifugal force to deploy the tethers.
This very same approach is attempted in the
ESTCube-1 mission.
2.3
Tether experiment
When the satellite has reached the desired angular velocity (approximately 1 revolution per second), the tether is deployed. The tether itself is extremely light. The proposed measures are 0,2 grams and 10 meters, see section 10 for further details. Because of the lightness of the tether, an end mass will
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2 MISSION PLAN probably be used at the tip of the tether. The end mass would also act as a re ector, making it easier to photograph the deployment. When the spin plane is perpendicular to the Earth magnetic eld, the voltage is applied to the tether with a polarity selected according to the spin, to increase or decrease spin rate by the electric sail and Lorentz forces, thus measuring both forces from the observed change of the spin rate (using accelerometer, gyrosensor and/or Earth limb -seeing camera). This procedure is repeated under dierent conditions until enough measurements have been made. As noted above, the experiment can and will be performed under several conditions in order to receive various information about the measured phenomena.
The list below contains some considerations about the possible
actions that may be taken.
If and when the electric sail and Lorentz forces work as expected, it is possible to modify the satellite spin at will when the tether is open and under voltage.
When the tether is charged positive, it experiences both Lorentz and electric sail forces. The electric sail force is larger, but not by much.
When the tether is charged negative, it experiences only the electric sail force and almost no Lorentz force, since a negative tether is quite ine cient current gatherer. This is due to the low speed of the ambient ions relative to the electrons.
If the initial spin is clockwise, the Lorentz force, which appears when the tether is positively charged, tends to accelerate the spin when the satellite is at the equator and to decelerate it when it is at one of the poles (and vice versa for the initially counterclockwise spin).
The electric sail force can be used to accelerate and decelerate the spin (or not to do anything about it) by judicious choice of the voltage phasing with the rotation. For a polar orbit this can be done at the poles, but not at the equator since at the equator the spacecraft velocity vector is perpendicular to the spin plane so the electric sail force tends only to bend the tether instead of modifying the spin. For an equatorial orbit this can be done at any time. In both cases it can be done both for positive and negative tether voltage separately, so the positive and negative electric sail eects can both be measured independently in this way.
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3 ORBIT SELECTION AND DETERMINATION
3 Orbit selection and determination 3.1
Keplerian elements
Elements of orbit are the parameters needed to uniquely specify an orbit. Given an inertial frame of reference and a speci ed point in time, six parameters are necessary to completely describe an orbit and motion of a satellite. Traditional orbital elements are the six Keplerian elements:
Eccentricity - de nes shape of orbit.
Semi-major axis - de nes the size of the orbit.
Inclination - de nes the angle between the equator and the orbit plane.
Right Ascension of the Ascending Node - de nes the angle between vernal equinox (the direction of the constellation Aries) and the point where the orbit crosses the equatorial plane (from Southern Hemisphere to Northern Hemisphere).
Argument of Perigee - de nes the point where the satellite is closest to Earth.
True/Mean Anomaly - de nes where the satellite is on its orbital path. The mean anomaly ranges from 0 to 360 degrees and is referenced to perigee. (At perigee the mean anomaly would be 0).
3.2 3.2.1
Formats Two-line elements
Keplerian elements can be encoded as text in a number of formats. The most common of them is the NASA/NORAD Two-Line Elements (TLE) format. Data for each satellite consists of three lines in the following format:
NOAA 14 1 23455U 94089A 97320.90946019 .00000140 00000-0 10191-3 0 2621 2 23455 99.0090 272.6745 0008546 223.1686 136.8816 14.11711747148495
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3 ORBIT SELECTION AND DETERMINATION Table 1: TLE; Line 1 explanation Column
Description
01
Line Number of Element Data
0307
Satellite Number
08
Classi cation(U=Unclassi ed)
1011
International Designator (Last two digits of launch year)
1214
International Designator (Launch number of the year)
1517
International Designator (Piece of the launch)
1920
Epoch Year (Last two digits of year)
2132
Epoch (Day of the year and fractional portion of the day)
3443
First Time Derivative of the Mean Motion,
4552
Second Time Derivative of Mean Motion (decimal point assumed)
5461
BSTAR drag term (decimal point assumed) for GP4 peturbations (otherwise radiation pressure coe cent)
63
Ephemeris type
6568
Element number
69
Checksum (Modulo 10) (Letters, blanks, periods, plus signs = 0; minus signs = 1)
Line 0 is a twenty-four character name of the satellite (to be consistent with the name length in the NORAD SATCAT). Lines 1 and 2 are the standard Two-Line Orbital Element Set Format identical to that used by NORAD and NASA. The format is explained in Tables 1 and 2. Each number is speci ed in a column. Spaces are signi cant. The last digit on each line is a mod-10 check digit, which is checked by program.
The
program also checks the sequence numbers ( rst column), and checks each orbital element for reasonable range. That is why this format is relatively safe and robust.
3.2.2
AMSAT format
The Radio Amateur Satellite Corporation (AMSAT) developed its own Keplerian element format. Table 3 shows an example. This format is designed to be read and edited by humans, and for this reason, each eld has a descriptive tag, and the organization of data is not as rigid as in the TLE format. It is said that there are problems with converting from the AMSAT format into the two-line format due to considerations of the accuracy of the predic-
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3 ORBIT SELECTION AND DETERMINATION Table 2: TLE; Line 2 explanation Column
Description
01
Line Number of Element Data
0307
Satellite Number
09-16
Inclination (degrees)
1825
Right Ascension of the Ascending Node (degrees)
2733
Eccentricity (decimal point assumed at beginning)
3542
Argument of Perigee (degrees)
4451
Mean Anomaly (degrees)
5363
Mean Motion (revs per day)
6468
Revolution number at epoch (revs)
69
Checksum (modulo 10)
tion models and it is not recommended to convert between the AMSAT and TLE formats if accuracy is a concern. Spaces are not signi cant. Each orbital element must appear on a separate line.
The order in which orbital elements appear is not signi cant, except
that each element set should begin with a line containing the word satellite. A blank line is usually interpreted as ending the element set. This le format does not contain any check digits, but an overall checksum is sometimes used.
3.2.3
One-Line Charlie Elements
The One Line Element (OLE) format is a format to express the Keplerian elements, which is rarely used nowadays. The only virtue to this format is its brevity. The format looks like the following:
123456789012345678901234567890123456789012345678901234567890 nnnnnyydddffffffddddddiiiiiinnnnnneeeeeeaaaaaammmmmmxxxxxxxx Table 4 gives the description of the format.
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3 ORBIT SELECTION AND DETERMINATION Table 3: Example of AMSAT format.
3.3
Description
Example
Satellite
AO-13
Catalog number
19216
Epoch time
94311.77313192
Element set
994
Inclination
57.6728 deg
RA of node
221.5174 deg
Eccentricity
0.7242728
Arg of perigee
354.2960 deg
Mean anomaly
0.7033 deg
Mean motion Decay rate
rev 2.09727084 day rev −5.78 × 10−06 day 2
Epoch rev
4902
Checksum
312
Prediction models
From Spacetrack Report NO.3: The most important point to be noted is that not just any prediction model will su ce. The NORAD element sets are mean values obtained by removing periodic variantions in a particular way. In order to obtain good predictions, these periodic variations must be reconstructed. [7] Orbital positions can be calculated from TLEs through algorithms. of them is SGP4.
One
SGP4 (Simpli ed General Perturbations Satellite Orbit
Model 4) is the most common algorithm used to calculate the orbit positions for satellites with orbit periods less than 225 minutes. This model is believed to be currently used to generate the NORAD Two-Line Elements and that is the reason why it usually is the best choice in practice. The accuracy of SGP4 is typically about 1 km in position. Downside is that TLE data is less accurate with each day.
3.4
GPS
When precise orbit determination is necessary, a Global Positioning System (GPS) receiver is needed. This paragraph will make an overview of available GPS receivers and some antennas that could be used on Cubesats. It will also
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3 ORBIT SELECTION AND DETERMINATION Table 4: Description of OLE format. Columns
Description
Format
1 - 5
NORAD catalog number
nnnnn
6 - 7
Year
yy
years
8 - 10
Day number
ddd
days
11 - 16
Fraction of a day
0.
17 - 22
Drag
0.dddddd
23 - 28
Inclination
iii.iii
days rev day 2 degrees
29 - 34
R.A.A.N.
nnn.nnn
degrees
35 - 40
Eccentricity
0.eeeeee
dimensionless
41 - 46
Argument of Perigee
aaa.aaa
degrees
47 - 52
Mean Anomaly
mmm.mmm
53 - 60
Mean Motion
xx.xxxxxx
degrees rev day
Note:
Units
R.A.A.N - Right Ascension of the Ascending Node
discuss if such high accuracy is reasonable compared to mission requirements and especially the budgets needed for the GPS system. One might think that if even mobile phones contain GPS systems then it would be relatively easy to incorporate one into a picosatellite but there is more. The most problematic are the COCOM trade limitations which limit the GPS receivers' maximum altitude to 18 km and velocity to 515 m/s. Because of that most manufactures refuse to make GPS receivers without these limitations and it is a big challenge to nd some without them.
To
complicate things even more, only a handful of Cubesats have had a GPS system on board and most of them (if not all) have been specially meant for use in space. Two interesting possible products come from Timble: Lassen iQ and Copernicus II. Both are reported working over 18 km by amateur high-altitude balloonists. The rst one is also assumed to work in a Cubesat.[8] One uncertain aspect is the speed because Lassen iQ data sheet lists 515m/s as the maximum speed because of the COCOM requirements. measures 26 mm W
×
26 mm L
×
The module itself
6 mm H and weighs 6,5 g. Typical power
consumption is 86 mW. The study estimates the full GPS system to weigh 57,4 g including antenna and mechanics. The Copernicus II is a newer model which according to the manual has operating limits of 515 m/s and 50 km. It is also smaller, measuring 19 mm W
×
19 mm L
×
2,54 mm H and weighing 2 grams. Continous tracking would
comsume 120 mW. With embedded 3,3 V antenna which measures 22 mm L
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3 ORBIT SELECTION AND DETERMINATION ×
21 mm W
×
8 mm H and having a mass of 20 g it might be possible to
include one of them into EstCube-1. As stated earlier, there have been some reports from amateur high altitude balloonists over GPS receivers that can operate above the altitude 18 km. These limits can vary between models or even software versions. It is not very important because the manuals often state a little bit higher altitude limits too.
Unfortunately one cannot test the velocity limit with high altitude
balloons.
Manufactures forgetting to include these limits on some models
is doubtful and there is no certain way knowing if they work at the most probable EstCube-1 orbit with an altitude of 600800 km and a velocity of approximately 7500 m/s. Despite that, there is always also a possibility to sort out the legal issues, and if then the companies were interested to modify their software for use in space, it would greatly bene t the Cubesat community. The 100 mW power consumption would be ne for constant use in most situations. Now an overview of GPS systems used by picosatellites to see if any of them has used commercial, o-the-shelf (COTS) GPS receivers:
Compass-1: Used Phoenix GPS receiver from DLR having consumer GP4020 chip from Zarlink. The chip includes the same COCOM limitations but the receiver additionally has necessary software for such conditions.
The receiver measures 70 mm
×
50 mm
×
12 mm and
has a power consumtion of 800 mW which makes it di cult to use it continuously.
Can-X series:
Used Superstar OEM board from CMC Electronics.
Company was bought up by Novatel and the Superstar GPS receiver lineup has now been discontinued. 71 mm
×
Superstar II measures 46 mm
13 mm and weighs 22 g.
×
Typical power consumption is
0,50,8 W. Product datasheet lists the COCOM limitations so one can speculate that they had a special version from the company.
AtmoCube: Uses the GPS receiver SSTL SGR-05 from Surrey Satellite Technologies, which is meant for use in space. It measures 70 mm 45 mm
3.5
×
×
10 mm, weighs 20 g and consumes 800 mW of power.
Orbit for ESTCube-1 mission
The choice of the orbit for ESTCube-1 mission is a tradeo between the mission requirements, technical limitations and the availability of launch op-
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3 ORBIT SELECTION AND DETERMINATION portunities. Considering only the requirements of the tether experiment, it would be desirable to have an orbit with an apogee outside the Earth's magnetosphere. This would maximise the available solar wind plasma density, and therefore the electric solar sail force.
Such an orbit would, however,
increase the demand for protection against radiation. Also the communication with the satellite would be demanding. Given the limited resources of a single unit Cubesat mission, a low Earth orbit (LEO) is preferred. LEO is an orbit, whose apogee remains within 2000 km over the Earth's surface. Also the good availability of launch opportunities can be considered a factor supporting the LEO. Continuing the orbit speculation with the assumption of LEO, we are basically down to deciding the actual parameters of the orbit.
Mainly, the
altitude and the inclination should be chosen. Due to the technical requirements for the tether mission, we are currently considering to use either a (near) equatorial LEO or a (near) polar LEO.
3.5.1
Equatorial LEO
An equatorial orbit is an orbit with an inclination close to zero.
In other
words, the plane of the orbit coincides, or nearly coincides, with the plane of the equator. For ESTCube-1 mission the equatorial orbit would be an ideal choice, since the tether experiment requires the Earth magnetic eld lines to be perpendicular against the spin plane of the satellite, see section 2.3 for further details. On an equatorial orbit this condition could be satis ed throughout the orbit period. On the other hand, the communication with the satellite could not be handled with the Tartu ground station. Also, nding a suitable launch could be challenging.
3.5.2
Polar LEO
◦ A polar orbit is an orbit with an inclination close to 90 .
Such an orbit
passes over the poles, and its plane rotates, causing the satellite to pass the equatorial plane at a dierent longitude during each period.
This makes
the polar LEO an appealing choice for remote sensing applications.
For
ESTCube-1 mission a polar orbit would oer an opportunity to have frequent overpasses over Tartu, as well as other radio amateur stations all over the world. Also the launch opportunities are more frequent, compared to those for an equatorial LEO. The drawback of using a polar orbit is the orientation
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3 ORBIT SELECTION AND DETERMINATION of the magnetic eld lines. The experiment could not be run constantly, since the eld lines are not properly aligned, with respect to the satellite spin plane, throughout the orbit period.
3.6
Conclusion
The orbit for the ESTCube-1 mission has not been selected yet. In practice, the choice will be made between the equatorial and polar LEO. Even though the technical requirements of the tether experiment would favor the equatorial orbit, all other factors seem to be in favor of the polar orbit. Firstly, the availability of the launch opportunities is considerably better for the polar orbits. This is due to the suitability of the polar orbit for several mapping, remote sensing and Earth observation applications. Equatorial LEO is sometimes preferred in remote sensing, when the equatorial region is of interest. Still, launch opportunities for polar orbits are far more frequent. In addition, the use of an equatorial orbit would require accessibility to a ground station located near the equator. The factors listed above seem to favor the polar orbit.
However, we must
remember that the one issue supporting the equatorial orbit is a strong one. The alignment of the Earth magnetic eld lines, and therefore the possibility to conduct the tether experiment at all times, is a powerful argument. Therefore, the question of the orbit remains open for now. The satellite will be designed to be applicable for either orbit. This can, of course, change in the future, once nal selection of the orbit has been made. As the last remark for the orbit speculation, one must bear in mind that if an oer for a free launch arises (e.g. the Vega rocket of ESA), letting such an opportunity go pass would be foolish. In student satellite projects, such chances may sometimes overrun technical considerations. The objectives of the mission would not, however, be compromised. The determination of the orbit and the position of ESTCube-1, during the mission, will rely on the TLE's, provided by NORAD. With the best knowledge available to us at the time of writing this report, we see no need to include a GPS receiver in the satellite. While some GPS receivers work over the altitude of 18 km, they still have an altitude limitation. Also, no consumer GPS receiver is known not to have the velocity limit. GPS receivers, which are meant for space, have quite a high power consumption and a noticeable mass. Taking into account the relatively tight power and mass budgets of ESTCube-1, these factors do not favor the use of GPS.
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3 ORBIT SELECTION AND DETERMINATION The NORAD TLE's are widely used for locating the satellites and predicting overpasses for Cubesat missions.
They have proved to be well suited for
this task, given the required accuracy of positioning of such missions. Also, the accuracy requirements for the ESTCube-1 mission are not considered to exceed the accuracy oered by the NORAD TLE's.
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4 210 Structure (STR) 4.1
4 210 STRUCTURE (STR)
Introduction
The STR system's task is to build a strong structure and mechanical systems for the satellite.
All other systems and the payload are integrated to the
structure. The weight of a spacecraft structure should be as small as possible, but the structure should be still strong enough against forces to the satellite in the launch phase and nally in the zero gravity environment. In the building of our subsystem we will follow the Cubesat standard. The standard gives us dimensions of the satellite and its weight, which we have to follow. Material selection is based also on the Cubesat standard and the materials used should be meant for space use. Our main task will be to keep the mechanical systems as simple as possible, so that they would later work in space and not fail.
4.2
System members
Members of the STR team are listed in Table 5.
Table 5: Members of the STR team. Name University Role
4.3 4.3.1
Paul Liias
TUT
Subsystem coordinator
Endel Soolo
UT
Subsystem member
System description Material
In the Cubesat Design Speci cation there are two types of aluminum suggested for the main structure.
The types are AW 6061-T6 and AW 7075.
If other materials are used, the thermal expansion must be similar to that of Aluminum 7075-T73 (P-POD material) and approved by Cal Poly launch personnel.
For our rst satellite we decided to use one of the suggested
aluminums. Selection criteria for the satellite material are:
Speci c strength
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4 210 STRUCTURE (STR) Table 6: Comparison of materials. Parameter
AW 6061-T6
AW 7075
Unit
Density
2,7
2,81
23,6
23,6
3 g/cm µm/m◦ C
Modulus of Elasticity
69
72
CTE, linear
20◦ C
Heat Capacity
0,896
0,96
GPa ◦ J/g C
Tensile Strength, yield
275
505
GPa
Price
Cheap
Expensive
Speci c stiness
Stress corrosion resistance
Fracture and fatigue resistance
Thermal expansion coe cient and conductivity
Ease of manufacture
As we can see from Table 6, these two materials have just some minor dierences.[9] The main advantage of the 7075 is its very high strength, optimized for highly stressed structural parts. For our satellite, however, the 6061-T6 is the excellent choice.
6061 has a relatively high strength, good
workability, and high resistance to corrosion. In addition, AW 6061 is more easily available than AW 7075. The following parts will be constructed of aluminum (the structure of the frame is further discussed in Section 4.5):
Rails
Beams
Battery box
Rails and beams together will be the frame of the satellite. As said above, the frame has to be made of aluminum 6061-T6. Its weight should be as small as possible, but it has to be strong enough to survive in dierent conditions. All other components of the spacecraft are attached to the frame. The battery box is used as protection for the battery itself against the pressures caused by vacuum and temperature dierences and against a battery
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4 210 STRUCTURE (STR) explosion, which could destroy some of the satellite components. The box is also made of aluminum, but it can be a dierent type of aluminum than the material of the structure. In case the weight of the satellite needs to be reduced, we will need to use some other material instead of aluminum.
Carbon ber reinforced plastic
(CFRP) oers a suitable solution for this. CFRP is often used in aerospace and automotive elds, as well as in sport equipment. This material is very 3 strong and light. Density of CFRP is just 1,5 g/cm . Its coe cient of expansion is low, as are the thermal and electrical conduction. CFRP would be a poor choice for the main structure but for the side plates it would be applicable.
The CFRP panels need to be designed to have a coe cient of
thermal expansion similar to that of aluminum 6061-T6 between the frame and panels.
On the panels a thin aluminum foil is mounted during the
curing process in order to make the panels more resistant to free oxygen and radiation, and most important of all to decrease the temperature range of the satellite during operations.
With this material we can reduce ca 45%
weight of each panel. As a drawback, it will be more expensive and harder to develop.
4.3.2
Mechanical parts
Kill switch
Kill switch also called deployment switch is used to
activate the power system of the satellite after deployment. It is necessary as when the satellite is in the P-POD no electronics shall be active during launch to prevent any electrical or RF interference with the launch vehicle and primary payloads. There will be two switches and they are connected parallel to decrease the probability of a failure.
Remove before Flight Pin
Along with the kill switches there is also the
requirement for a remove before ight pin to disable the satellite before and during integration with the deployer. Once the satellites are loaded into the deployer, the remove-before- ight pin is removed.
Antenna deployment system ployment of the antennae.
A speci c system is designed for the de-
This system contains the antenna itself and a
nylon ber which is melted to deploy one end of the antenna. The material for the antenna will be beryllium copper. Beryllium copper is used in springs
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4 210 STRUCTURE (STR) and other parts that must retain their shapes during periods in which they are subjected to repeated strain.
Separation Spring
The top of the rails are to be equipped with separation
springs to enable the deployment from the P-POD.
4.4 4.4.1
Deduced requirements from the Cubesat standard Dimensional requirements for the satellite structure
The structure of the Cubesat must be strong enough to survive maximum loading de ned in the testing requirements and cumulative loading of all required tests and the launch. The Cubesat structure must be compatible with the P-POD. Speci c requirements from the Cubesat Design Speci cation:[2]
The Cubesat shall be 100.0+0.1 mm wide (X and Y) and 113.5+0.1 mm tall (Z).
Rails must be smooth and edges must be rounded to a minimum radius of 1 mm.
No external components other then the rails shall be in contact with the internals of the P-POD.
Each rail shall be a minimum of 8.5 mm wide.
At least 75% (85.125 mm of a possible 113.5mm) of the rail must be in contact with the P-POD rails. 25% of the rails may be recessed and NO part of the rails may exceed the speci cation.
All rails must be hard anodized to prevent cold-welding, reduce wear, and provide electrical isolation between the Cubesats and the P-POD.
Separation springs must be included at designated contact points (Appendix A of the standard).
The use of Aluminum 7075 or 6061-T6 is suggested for the main structure.
Deployables must be constrained by the Cubesat.
The P-POD rails
and walls are NOT to be used to constrain deployables.
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4 210 STRUCTURE (STR)
Figure 1: Main frame.
4.4.2
Electrical requirements
Speci c requirements from the Cubesat Design Speci cation:
No electronics shall be active during launch to prevent any electrical or RF interference with the launch vehicle and primary payloads.
One deployment switch is required (two are recommended) for each Cubesat.
The deployment switch should be located at designated
points (Appendix A of the standard).
Deployment switch shall be compatible with +Z contact point(s).
A remove before ight (RBF) pin is required to deactivate the Cubesats during integration outside the P-POD.
4.5
Layout proposal
The satellite body consists of the main frame, shown in Figure 1, and skin plates that are attached to it.
For the structure of the main frame there
are two possible solutions. One solution is that the frame contains six parts
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4 210 STRUCTURE (STR)
Figure 2: Rails.
(rails and beams, see explanation below), the other solution would be to manufacture the whole frame out of a solid block of aluminum.
The rst
option would be cheaper, but the one piece frame would be stronger and a little bit lighter, because there would not be any connections.
Rails
The rails (two pieces) are the only parts of the satellite that will have
contact to the P-POD. The rails are depicted in Figure 2.
Beams
The beams (four pieces) hold the rails together. The structure of
a beam is shown in Figure 3.
Plates
The plates (four side plates and two top plates) will protect the
inside of the satellite, mainly the the PCB's, against radiation and temperature changes. The solar panels and the antenna deployment system will be installed on the plates. In some of the panels there will be holes, for example for the atennas, the tether, electron gun and the camera. The structure of the plates is shown separately for the side plates and the topp plates in Figure 4. These plates are often referred to as skin plates or skin panels. Figure 5 illustrates the assembled satellite frame. Figure 6 shows the layout of the satellite with the subsystems and the payload inside the frame.
The payload will be placed into the middle of the
satellite and it will be xed with the PCB-board to the structure, like all
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4 210 STRUCTURE (STR)
Figure 3: Beams.
Figure 4: On the left: side plate. On the right: top plate.
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4 210 STRUCTURE (STR)
Figure 5: Assembled satellite frame.
Table 7: Power and mass budget of the STR system.
The smaller values
refer to the use of the CFRP skin panels. The mass of the battery box has not been included in the sum, since it will probably not be used. Component
Qty
Power, mW
Mass, g
Frame
1
-
110
Skin panels
6
-
132 / 84
Antenna dpl sys
1
?
10
Battery box
1
-
(30)
Kill switch
2
-
10
Separation Spring
2
-
6
?
268 / 220
P
other subsystems. The thickness of the payload compartment can be 15 mm high and max 90x96 mm wide. And it has three openings in the side panels for the tether, the camera and the electron gun.
4.6
Technical results
Cubesat standard sets the weight of a single Cubesat to be less than 1 kg. Therefore we have to calculate the estimated weight of ESTCube in the phase 0 to see if this mission is possible. The heaviest components in this subsystem are the frame and the skin panels. If we need to further cut down the mass of the structure, it can be done by using the CFRP parts as skin plates. Also the use of the battery box can be discarded. Power is needed
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4 210 STRUCTURE (STR)
Figure 6: Layout of the satellite.
just at the beginning for the antenna deployment system. The exact amount of power needed for the antenna deployment is not known at the moment. The estimate of the mass and power budget is shown in Table 7. The mass of several components is likely to be slightly over estimated in this preliminary estimate.
4.7
Conclusion
We have decided to use aluminum 6061-T6 as the frame material and we were able to do the rst mass calculations for the satellite structure.
In
the phase 0 the mass calculation is the most important one as we need to verify that the mass of the satellite will stay within the limits of the Cubesat standard. The estimated weight of the STR subsystem is currently 268 g. If the CFRP skin panels are used, the mass will reduce to 220 g and after the more detailed calculations are nished, it is possible that the mass will be even further reduced. This weight includes the weight of the frame, as well as the weight of the other components of the STR system.
The decisions
concerning the structure of the main frame (single block vs rails and beams) and the skin panel material will be made in the future.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS)
5 220 Attitude Determination and Control System (ADCS) 5.1
Introduction
The purpose of the attitude determination and control system is to stabilize the satellite in orbit, start and maintain the rotation during the tether experiment, stop the rotation after the experiment has concluded and maintain the required side towards Earth. After the satellite has been released into space, the detumbling mode will stabilize the movement by stopping the rotation caused by the ejection from the P-POD. During the experiment phase the system will start and maintain the needed rotation around one axis and will stop the rotation again after the experiment has reached its end. Thereafter the main objective is to keep one of the sides pointed towards Earth (nadir pointing) so that the secondary mission objectives can take place.
5.2
System members
Members of the ADCS team are listed in Table 8.
Table 8: Members of the ADCS team. Name University Role
5.3
Erik Kulu
UT
Subsystem coordinator
Endel Soolo
UT
Subsystem member
Katrin Tuude
UT
Subsystem member
System description
Description of the ADCS:
Determine the position of the spacecraft in respect to Earth.
Determine the rotational movement of the spacecraft.
Stop the initial tumbling.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS)
Align the satellite as required.
Start and stop the rotation needed for the experiment.
5.3.1
Deduced requirements from the Cubesat standard Size: Smaller than (equal to) 10cm
×
10cm
×
10cm together with all
the parts of every subsystem.
Total weight: Smaller than (equal to) 1 kg together with all the parts of every subsystem.
5.3.2
No reactive propulsion.
Deduced requirements from the payload
The necessary instruments:
Sensors to determine the tumbling
Actuators to stop the tumbling
Sensors to measure the alignment of the spacecraft
Actuators to correct the alignment
Actuator for achieving the high velocity rotation
Gyroscope to measure the changes in angular velocity
5.3.3
Considerations about the spinning axis The tether has to be in same direction as the centrifugal force, perpendicular with the spinning axis.
The axis of the momentum wheel must be parallel with the spinning axis.
The axis of the tether spool should be parallel with the satellite's spinning axis, otherwise deploying the tether would change the satellite's spinning axis.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS)
The tether has to be in the visual eld of the camera, so camera has to be on the same side and in the same direction as the tether.
The antennas have maximum gain in the direction perpendicular with the direction of the antennas. To have good reception all the time while spinning, the antennas should be parallel with the spinning axis if the direction to base station is perpendicular with the spinning axis.
If
then the base station can be held always in the direction of spinning axis, then the antenna can be perpendicular with the spinning axis.
If the spinning axis is vertical in some part of the orbit, then it will ◦ be horizontal in the 90 shifted position on the orbit. The axis cannot be constantly pointed toward earth without constantly using energy to rotate it.
If the antenna is parallel with the spinning axis, then satellite's spinning axis has to be perpendicular with the orbital plane, otherwise the antennas would be pointed towards earth and lose their reception at some part of the orbit.
The antennas can be made thinner and lighter if they are perpendicular with the spinning axis, so the centrifugal force would keep them straight. However, that would increase the momentum of inertia and need more speed and power for the momentum wheel.
Before experiment, camera is pointed to earth, so starting to rotate around horizontal axis needs less maneuvering than spinning around vertical axis.
The electron cannon should be pointed away from the tether to avoid the electrons stopping on the tether.
Maybe the electron cannon should be pointed away from the plane that the tether is going through during one revolution, so it should not be perpendicular with the spinning axis.
Maybe the electron cannon should be pointed away from the antennas, then it should not be parallel with the spinning axis.
If the previous two concerns are justi ed, then the electron cannon ◦ could be pointed at e.g. 45 angle from the spinning axis.
The antennas should be directed so that the tether can not tangle around them if the rotation is stopped while the tether is still out.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS)
Magnetic torquer can be used to start spinning only around axes perpendicular to the magnetic eld lines.
Possible antenna directions
There are two possible con gurations for
the antennas in respect of the spinning axis that allow communication during the phase of nadir pointing.
Antenna parallel with the spinning axis (Figure 7)
⊕
Least momentum of inertia, least power needed to start spinning.
Spinning around an axis that has neither maximum nor minimum momentum of inertia for a given body is inherently unstable.
No communication while over poles.
Antenna perpendicular with the spinning axis (Figure 8)
⊕
Lighter antennas
⊕
More stable spinning axis
Larger momentum of inertia, needs more power to start spinning.
Discontinuous communication while over equator.
The variant 1 of antenna placement is preferred because it requires least power for the momentum wheel to start spinning and it leaves open the possibility of using equatorial orbit.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS)
Figure 7: Antenna direction parallel with the spinning axis (variant 1).
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS)
Figure 8: Antenna direction perpendicular with the spinning axis (variant 2).
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) Conclusion
Without considering the Lorenzt force experinent, the best
con guration would be if the spinning axis was perpendicular with the orbital plane. Tether, view direction of the camera and the maximum gain directions of the antennas would be in the orbital plane.
Antennas and all axes of
rotating satellite components would be parallel with the spinning axis of the satellite body. However, for the Lorentz force experiment, the only solution on polar orbit would have the spinning axis in the orbital plane. The spinning axis has to be parallel with the axis of Earth. Such spinning axis makes it impossible to keep the antennas in a direction suitable for communication with Earth during the whole orbit and also makes it di cult to start rotation of the satellite without a momentum wheel.
5.4 5.4.1
Layout proposal System components
Sensors:
Digital magnetometer
Sun sensor (fast CMOS image sensor)
MEMS gyroscope
MEMS accelerometer
Actuators:
Magnetic torquer
Momentum wheel
5.4.2
Subsystem operation phases
1. Detumbling
After launch and separation from the carrier rocket, the
satellite will rst go into the detumbling mode where the objective is to stop the random rotation of the satellite.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) The magnetometer readings are taken at sub-second intervals.
From the
time-line of the readings the rotation period and rotation axis are determined. A desirable magnetic moment vector is calculated based on the angular velocity vector and the current magnetic eld vector reading, so as to slow down the rotation in the fastest possible time. The coils can be only switched on or o, but the magnet moment vectors in any direction can be simulated by independently varying the activation pulse widths of separate coils.
2. Nadir pointing
The objective of this phase is to have the maximum
gain direction of the antennas and the view direction of the camera continually pointed towards Earth, so the communication can be started and some pictures can be taken. The magnetometer and sun sensor readings at regular intervals are used to locate the satellite and determine its attitude and rotation speed based on the orbit data and a map of the earth magnetic eld. When the attitude is unsuitable, i.e., if the satellite side with camera and antennas is more than ◦ 10 o the nadir direction or if some special camera orientation is required, then a magnetic coil is powered for a short period to start rotation towards the correct orientation.
When that direction has been attained, another
short magnetic pulse is used to stop the rotation.
When the satellite has
been stabilised at the preferred attitude, the ADCS system goes to power saving mode and is only activated for short time periods once a minute to check the orientation and correct it if necessary.
3. Experiment
The objective of this phase is to have the satellite rotating
around the correct axis at the rate of 1 revolution per second when the deployment of the tether starts, then maintain its spinning axis, but not to do any sudden attempts to correct in order to avoid tangling the tether. The deployed tether slows down the satellite's rotation. That eect has to be monitored using the gyroscopic sensors and reported back to the base station. Three possible methods have been proposed for starting the rotation, each having dierent equipment and/or power requirements.
Variant A
The satellite is checked to be in the nadir pointing.
Then
the momentum wheel is quickly started in the direction against the desired
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) rotation direction of the satellite. The momentum wheel is kept running at constant angular velocity until the end of the experiment. spinning is slowed down by braking the momentum wheel.
Thereafter the There can be
some momentum left due to external torques after the momentum wheel has been stopped. Such momentum has to be killed with the coils in phase 4.
Pros: Fast starting and stopping, therefore well determined axis of rotation. Cons: High power requirement during the experiment. In case of power failure
during the experiment phase, the tether will tangle itself.
Variant B
There is no reaction wheel. Spinning is started using the mag-
nettorquer coils during an extended period of time while constantly monitoring the mode of rotation.
Pros: The lightest method. Less moving parts. No power needed during the
experiment to maintain spinning. Cons: The rotation axis may wander away from the optimal direction during
the slow starting of rotation. Control algorithm is more di cult and more error-prone than in variant A. Needs faster sun sensors to check the orientation while spinning.
Variant C
The satellite is veri ed to be in the nadir pointing. Then the
momentum wheel is started in the same direction as the desired rotation direction of the satellite. The rotation of the satellite body is slowly stopped using the magnettorquer coils. The satellite body is turned back to the initial attitude of nadir pointing, taking into consideration the precession caused by the powered-up momentum wheel. When the orientation has been veri ed, the momentum wheel is stopped by electrically braking its motor. The satellite quickly starts to spin. The momentum wheel motor and magnettorquers stay switched o during the experiment.
Pros: Satellite starts to rotate quikly around a well-de ned axis. No power
needed during the experiment to maintain spinning. Cons: The most di cult control algorithm.
Requires precise magnetometer
and good timing to distinguish the earth magnetic eld from the stray magnetic eld of the working reaction wheel motor.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) Table 9: Wheel vs coils in the tether experiment. A: Wheel
B: Coils
C: W & C
Peak power ( rst 3 sec)
6000mW
820mW
6820mW
Power during spin-up
430mW
820mW
1200mW
Power during the experiment
430mW
40mW
40mW
Mass
70g+70g
70g
70g+70g
Time to start
3 sec
20 + 36
Volume
cm
3
2 hour 2 cm
2 hour
20
20 + 36
cm
Axis stability
Good
Not so good
Good
Control
Simple
Di cult
Di cult
4. Detumbling
3
After experiment the spinning of the satellite has to be
stopped. The procedure is same as in the initial detumbling phase.
5. Nadir pointing
The antennas are oriented to the earth by the same
procedure as in phase 2. A repeated experiment may be requested by the operator. In that case, the execution of phase 3 follows, but by default the nadir pointing is held until the end of the satellite lifetime.
5.5
Comparison of the three spin-up methods
The choice of the actuator used for spinning the satellite during the experiment is a trade o between the performance, size, total mass and the electrical power required. Table 9 shows the comparison of these properties.
5.6
Momentum wheel power calculation
For the following calculation an assumption was made that the antennas would be parallel with the spinning axis. That way their momentum of inertia is negligible. Motor type: Portescap nuvoDisc32 BF[10] Motor mass: 26g Wheel mass:
20g + spokes
Mass of the whole momentum wheel assembly: Wheel radius:
≈ 70g
r = 40mm
i = m · r2 = 3.2 · 10−5 kg · m2 M ·a2 Inertial momentum of the satellite as a solid cube: I = = 1.7·10−3 kg · m2 6
Inertial momentum of the wheel as a thin ring
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) Ω = 1rps = 6.28 rad s speed of the wheel: ω = Ω ·
Rotation speed of the satellite: Required counterrotation
5.6.1
= 54rps = 334 rad s
I i
Starting
Kinetic energy in the rotation of the wheel:
2
E2 = I·Ω = 0.034J 2 i·ω 2 E1 = 2 = 1.8J
Kinetic energy in the rotation of the satellite:
Possible starting regime for the motor: current limited to 0.5A, constant −3 torque output at 3.9 · 10 Nm, start-up time 3 seconds. Conclusion (with safety factors): The start-up needs up to 6 watts of power for up to 10 seconds.
5.6.2
Maintaining the speed
If ree = 65mA Urated = 12V
Motor no-load current: Motor rated voltage:
Motor back-EMF at the rated free-running speed:
EM F1 = 0.82 · 10−3 ·
12800 = 10.5V ω : EM F2 = 0.82 · 10−3 · 3.2 = 2.7V Motor voltage at ω : U = Urated − EM F1 + EM F2 = 4.2V Motor power at ω : Pcalc = 280mW E cency of motor commutation: η = 80%
Motor back-EMF at
Conclusion: the continuous power estimate for running the momentum wheel P motor during the experiment is P = calc = 350mW η
Note: if the antennas are not parallel with the spinning axis, then the moment of inertia of the satellite is considerably larger and motor speed and power have to exceed this calculation.
5.7
Budgets
Table 10 shows the mass and power requirements of the ADCS subsystem components.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) Table 10: Mass and power of the ADCS components. Subsystem part
Qty
Mass (each), g
Power (each), mW
Gyro-sensor
3
1
5
Digital Magnetometer
1
28
35
Magnetic torquer
3
23
350
Sun sensor
6
7
150
Accelerometer
1
5
10
DC motor+wheel
1
70
6000/350
PCB: controller
1
99
10
P
16
316
Table 11: Detumbling
5.7.1
Component
Power (mW)
Operation period
Controller
10
always on
Magnetometer
35
always on
One coil + switching losses
350
on 90% of time
Second coil + losses
350
on 50% of time
Third coil + losses
350
on 20% of time
Safety factor
+ 10 %
Average
650
2 hours
Average ADCS power consumption during phases
The average power consumption of the ADCS during dierent phases of operation are shown in Tables 11 through 17.
5.8
Conclusion
The proposed Lorentz force experiment requires a mode of rotation that is somewhat incompatible with other requirements for communication. Despite that, the required orientation and rotation modes can still be achieved and from this standpoint the proposed mission is feasible.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) Table 12: Variant A: During experiment
Note:
Component
Power (mW)
Operation period
Gyro-sensor
15
always on
Accelerometer
10
always on
Controller
10
always on
Magnetometer
35
on 1/60 of time
Sun sensors
150
on 1/60 of time
Wheel + losses
350
always on
Safety factor
+ 10 %
Average
430
can be months
Starting the momentum wheel would additionally need 6W for 3 seconds.
Table 13: Variant B: Spinning started with coils Component
Power (mW)
Operation period
Gyro-sensor
15
always on
Accelerometer
10
always on
Controller
10
always on
Magnetometer
35
always on
Sun sensors
150
one always on
One coil + switching losses
350
almost always on
Second coil + losses
350
on 50% of time
Safety factor
+ 10 %
Average
820
Note:
2 hours
The total time of 2 hours may be spread over certain parts of several consecutive orbits. Table 14: Variant C: Stopping the momentum wheel. Component
Power (mW)
Operation period
Gyro-sensor
15
always on
Controller
10
always on
Magnetometer
35
always on
Sun sensors
150
one always on
One coil + switching losses
350
almost always on
Second coil + losses
350
on 50% of time
Wheel + losses
350
always on
Safety factor
+ 10 %
Average
1200
Note:
2 hours
The total time of 2 hours may be spread over certain parts of several consecutive orbits.
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5 220 ATTITUDE DETERMINATION AND CONTROL SYSTEM (ADCS) Table 15: Changing orientation Component
Power (mW)
Operation period
Gyro-sensor
15
always on
Accelerometer
10
always on
Controller
10
always on
Magnetometer
35
always on
Sun sensors
150
one always on
One coil + switching losses
350
almost always on
Second coil + losses
350
on 50% of time
Safety factor
+ 10 %
Average
820
2 minutes
Table 16: Variant B or C: During experiment
Spinning started with coils or stopping the momentum wheel Component
Power (mW)
Operation period
Gyro-sensor
15
always on
Accelerometer
10
always on
Controller
10
always on
Magnetometer
35
on 1/60 of time on 1/60 of time
Sun sensors
150
Safety factor
+ 10 %
Average
40
can be months
Table 17: Stand-by and orientation keeping Component
Power (mW)
Operation period
Controller
10
always on
Magnetometer
35
on 1/60 of time
Sun sensors
150
on 1/60 of time
One coil + switching losses
350
when needed
Safety factor
+ 10 %
Average
20
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6 230 ELECTRICAL POWER SYSTEM (EPS)
6 230 Electrical Power System (EPS) 6.1
Introduction
The Electrical Power System (EPS) is responsible for gathering power from solar panels and storing it into batteries for later use by total system. ESTCube will probably have sun synchronous polar orbit in such a way that it will be in eclipse during about 30solar energy must be kept in a battery in order to return it to the subsystems during the eclipse (during this time the solar cells are of course quasi ine cient). The payload will take photographs during the eclipse if tether is deployed but this work does not consume a lot of energy. EPS supplies a continuous source of electrical power to spacecraft payload during the mission. It has to control, distribute and regulate power for subsystems and to balance power requirements and generation during eclipse and sunphase. Three main ESTCube electrical consumers are:
ADCS (magnetic torquers, motors)
PL (Electron gun)
EPS (converters and regulators)
EPS provides command and telemetry capability of its health and status. EPS protects against transient bus voltages and Single Event Latchup. This phenomenon happens when a high energy particle hits the device.
If the
impact on the device is of a serious nature, the high energy particle can directly cause damage to the device. This phenomenon happens very quickly and must be detected and corrected in hardware. In case of the power not being turned o at a latch-up, a burn-out can occur and destroy the chip.
6.2
System members
Members of the EPS team are listed in Table 18.
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6 230 ELECTRICAL POWER SYSTEM (EPS) Table 18: Members of the EPS team. Name University Role
6.3 6.3.1
Veigo Evard
TUT
Subsystem coordinator
Ramon Rantsus
UT
Subsystem member
Mirko Mustonen
TUT
Subsystem member
System description Deduced requirements from the Cubesat standard Nominal length of 100 mm per side. Solar cell standard size is 40 x 80 mm and we put two cells per side.
Each single Cubesat may not exceed 1 kg mass.
Estimated battery
weight is 62 g.
Center of mass must be within 20 mm of its geometric center. Batteries have probably highest mass density so these have to be near the center of cube.
No electrics shall be active during launch to prevent any electrical or RF interference with the launch vehicle and primary payloads. We have two deployment switches to cut the power from electronics.
One deployment switch is required (two are recommended) for each Cubesat.
Developers who wish to perform testing and battery charging after integration must provide ground support equipment (GSE) that connects to the Cubesat through designated access ports.
A remove-before- ight (RBF) pin is required to deactivate the Cubesats during integration outside the P-POD.
Cubesats with rechargeable batteries shall have the capability to receive a transmitter shutdown command, as per Federal Communications Commission (FCC) regulation.
To allow adequate separation of Cubesats, antennas may be deployed 15 minutes after ejection from the P-POD (as detected by Cubesat deployment switches).
Larger deployables such as booms and solar
panels may be deployed 30 minutes after ejection from the P-POD.
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6 230 ELECTRICAL POWER SYSTEM (EPS)
Figure 9: EPS block diagram.
Random vibration testing at a level higher than the published launch vehicle envelope outlined in the Mission Test Plan (MTP).
Thermal vacuum bakeout to ensure proper outgassing of components. The test cycle and duration will be outlined in the MTP.
6.3.2
Layout proposal
Figure 9 shows the proposed block diagram of the EPS. There are two cells per side that are connected serial and dierent sides are connected parallel to each other. Two opposite sides have one dedicated active maximum power point tracking for higher eciency and boost-converter for battery charging. We considered to use 4,1 V Li-Po batteries from Clyde Space. Charge system and battery will be doubled for backup. Power Distribution Unit (PDU) can switch subsystems o in case of low power. Only one MPPT might be used for all solar panels because electronics may take too much volume.
6.3.3
Mass and cost budbet
Table 19 shows the estimated mass and cost budget for the EPS.
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6 230 ELECTRICAL POWER SYSTEM (EPS) Table 19: EPS mass and cost budget. Component
Qnt.
Battery
2
Electronics Solar cells
P
6.3.4
10
Total Mass
Price
Temp. limits
grams
EEK
°C
Min
Max
Min
Max
Min
Max
70
130
500
11000
0
30
50
100
500
5000
-40
+85
26
40
TBD
TBD
TBD
TBD
146
270
TBD
TBD
Power generation during dierent phases
Table 20 shows the estimated power consumption and generation during the dierent phases of operation. The two bottom rows give the power available for the payload. Calculations of the output power levels have been done for a temperature of ◦ 28 C, a solar cell eciency of 28%, and assuming to have cells on ve sides of the satellite, two cells on each side. Worst case orbit has 41% of time in eclipse. Temperature variations have about
±3%
eect on nal power gen-
eration. Following losses have been included: re ection from cell, shadowing losses, ultraviolet degradation, radiation degradation, fatigue (thermal cycling), micrometeoroid loss, maximum power point tracker eciency, diode power loss. The maximum continuous power available from the solar panels depends on satellite attitude. For short periods higher power levels can be achieved from batteries (starting up momentum wheel for example). The Li-Po batteries we intend to use will have a capacity of at least 10 Whr. Power levels up to 100 W can be achieved for a limited time when the battery is used as the power source. This feature can be utilized when the power need of the system exceeds power generation by solar panels, as is the case for example in the beginning of the spinning phase. Payload power consumption is assumingly about 1,5 W and it will be powered for 1/3 of orbit time. Worst case scenario shows that if we use momentum wheel to spin the satellite (variant A) we can power payload every other orbit before batteries have to be recharged. If we use coils or momentum wheel togeteher with coils (variants B and C) to spin, payload can be powered 9/10 of polar passes. It is also possible to skip these recharge cycles but if batteries run empty it will take more time to recharge and batteries lifetime will shorten due to higher depth of discharge.
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6 230 ELECTRICAL POWER SYSTEM (EPS) Table 20: Power generation during dierent phases. Avg Cons (W)
6.3.5
Detumb
Orient
Exp, var A
Exp, var B C
Sun phase
3,14
2,008
1,576
1,576
Period
1,85
1,2
0,936
0,936
ADCS
-0,65
-0,048
-0,43
-0,12
CDHS
-0,061
-0,061
-0,14
-0,14
COM
-0,1
-0,1
-0,1
-0,1
EPS
-0,1
-0,1
-0,1
-0,1
Available, sun
2,229
1,699
0,806
1,116
Available, per
0,939
0,891
0,166
0,476
Discussion
Batteries depend on power consumption. For reference, see [11] or [12]. The ◦ temperature of the electronics is usually between -20 and 65 C, but if we buy industry speci ed components the temperature range will be -40 to ◦ +85 C. Solar cells are specially made for space so we assume that temperature changes are not a problem. Cells temperature must be kept as low as possible to maximize the power output.
6.4
Conclusion
At the moment calculations show that EPS will provide enough power for all subsystems but it has to be use eeciently during experiment. The payload power consumption needs to be updated to get more accurate information for calculations.
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7 240 THERMAL CONTROL SYSTEM (TCS)
7 240 Thermal Control System (TCS) 7.1
Introduction
A Thermal Control System (TCS) is a subsystem that is responsible for the maintenance of temperature constraints of all the hardware components of the Cubesat.
TCS is dependent of all the subsystems (including itself ) as
they all contribute to the overall heat equilibrium of the satellite. Most of the equipment on the spacecraft is usually designed to operate on or near room temperature (RT). The main task of the TCS is therefore to provide requisite temperatures for Cubesat components in space environment.
As
the conditions in space are harsh and uctuatinghardly ever near suitable RTspecial means have to be utilized in order to assure a properly working Cubesat.
7.2
System members
Members of the TCS team are listed in Table 21.
Table 21: Members of the TCS team. Name University Role
7.3
Ott Scheler
UT
Subsystem coordinator
Marko Lõoke
UT
Subsystem member
Examples of typical temperature ranges for satellite components
Table 22 shows a ballpark estimate for temperature ranges of dierent satellite components, the values have been taken from literature [13] and other
Table 22: Typical temperature ranges. ◦ Component Temperature Range, C
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Electronics
-15. . . +50
Batteries
0. . . +20
Solar cells
-100. . . +100
Structures
-45. . . +65
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7 240 THERMAL CONTROL SYSTEM (TCS)
Figure 10: Sources of thermal radiation
Table 23: Sources of heat in orbit. Source Value Unit Solar radiation
1371
Albedo radiation
480
Planetary radiation
237
Internal (battery)
∼2 ∼3
Relic radiation
Cubesat projects.
2 W/m 2 W/m 2 W/m W K
All the temperatures are for reference only and exact
values have to be gathered from the manufacturers.
7.4
Thermal environment for satellite in space (LEO)
Table 23 and Figure 10 depict the most signi cant sources of thermal energy for a satellite in orbit. The value given for solar radiation refers to the average radiation, called solar constant, it varies seasonally due to the elliptical orbit of the sun: at least.
±5%
Albedo radiation stands for the re ection of solar radiation from
the Earth surface, dependent on the surface characteristics (∼80% re ection from clouds, 5% from water and forest). The planetary radiation is originated from heat from the Earth at infrared wavelengths (thermal radiation). Relic
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7 240 THERMAL CONTROL SYSTEM (TCS) radiation, or cosmic microwave background radiation, can be neglected due to its weakness. For an orbiting spacecraft we have to consider the exact duration of eclipse periods when there is no solar and albedo radiation. Cubesat itself also emits infrared radiation into the space, and the resulting equilibrium temperature needs to be calculated.
7.5
Mechanisms of thermal energy transfer
Convection
Convection can be neglected in space as there is no liquid nor
gaseous environment.
Conduction
Spontaneous transfer of heat inside Cubesat through contact
from component of higher temperature to another component of lower temperature. Conductive heat ow follows the equation
Qc = where
Qc
λA ∆T, l λ
is the conductive heat ow,
(1)
is the thermal conductivity,
the cross sectional area between the components, duction path, and
∆T
l
A
is
is the length of the con-
is the temperature dierence in Kelvins. ABC of the
conduction: metals are good conductors, plastics are poor conductors = good insulators.
Radiation
Heat transfer between dierent bodies via radiation is depen-
dent on the surface temperature, radiative view factors and surface properties. Radiative heat ow follows the Stefan-Boltzmann law of radiation
Qr = σAT 4 , where
Qr
is the radiative heat ow,
Boltzmann constant,
1
A
(2)
is the emissivity,
is the surface area, and
T
σ
is the Stefan-
is the absolute tempera-
ture.
1σ
≈ 5, 67 · 10−8 Wm−2 K−4
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7 240 THERMAL CONTROL SYSTEM (TCS) 7.6
Temperature equilibrium and thermal balance calculations
Temperature of the spacecraft (all the components inside and on the surface of the satellite) depends on the balance between the heat received from external and internal resources, and the heat radiated into space by the spacecraft itself. Delicate balance between the heat absorbed and the heat radiated determines the nal temperature of the Cubesat and its components. Satellite will achieve its thermal equilibrium some time after the beginning of the solar irradiation. Equilibrium is obtained when the absorbed power power emitted by radiation
Q .
Qα
is equal to
The aforementioned powers can be expressed
as
where
S0
is the irradiance,
illuminated area,
Te
α
Qα = S0 αAi n
(3)
Q = σAout T4 ,
(4)
is the absorptivity of the material,
Aout
is the emissivity,
Ain
is the
is the radiating surface area, and
is the equilibrium temperature.
Equilibrium condition is
Qα = Q
(5)
S0 αAin = σAout T4
(6)
s
T =
4
S0 αAin σAout
(7)
Because the side panels of a Cubesat are facing the sources of radiation at dierent angles, the change in the projectional area account, i.e. the angle of incidence
θ
Ain
must be taken into
must be included in the equation:
Qα = S0 αAin cos θ.
(8)
Combined properties of absorptivity and emissivity (α/ ratio) characterizes the surface material:
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7 240 THERMAL CONTROL SYSTEM (TCS) Table 24: Surface characteristics of some satellite materials. Material α α/
High
Low
Gold
0,25
0,04
6,25
Aluminum (polished)
0,24
0,08
3
Black paint (epoxy)
0,95
0,85
1,12
Black paint (polyurethane)
0,95
0,9
1,06
Solar cells (GaAs)
0,88
0,8
1,10
White paint (silicone)
0,26
0,83
0,31
White paint (silicate)
0,12
0,9
0,13
α/ α/
ratio: surface is a good absorber, but a bad radiator ratio: surface is a poor absorber, but a good radiator
Table 24 shows the surface characteristics for a few materials commonly used in satellites.
7.7
Thermal balance calculation of a raw satellite
For the preliminary calculations a raw (without any thermal control mechanism applied) satellite thermal equilibrium is calculated as shown in Equation (7). Hence
Aα Jabsorbed = A Jradiated
(9)
(Asun Js + Aalbedo Ja )α + AEarth Je + Q = Asurf ace σT 4 , where
αsun
and
αEarth
(10)
are the absorptivities of the Sun and the Earth.
Since we do not know the exact positioning of the spaceship in time, an assumption is made that energy from sun is adsorbed so that the tetrahedron of the cube is facing the sun. This should be the average simpli ed case. The angle between a cube face and the farthest corner is calculated as θ = √ arctan(1/ 2) = 35, 26◦ . The incidence angle is then θ = 90◦ − 35.26◦ = 54.74◦ . Three faces are illuminated at the same incidence angle. Eective area is then the area of the three sides facing the sun, multiplied by
cos θ.
Table 25 gives estimated values needed for the temperature calculations for a 1-liter Cubesat.
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7 240 THERMAL CONTROL SYSTEM (TCS) Table 25: Material usage and thermal parameters. Material ASun , m2 AEarth , m2 α
7.7.1
Solar cells
0,0110
0
0,88
0,8
Aluminum (polished)
0,0062
0,001
0,24
0,08
Temperature equilibrium of the spacecraft at the sunlight
Using the equations and material parameters introduced above, we can now derive an estimate for the equilibrium temperature for the satellite at Sun light. First we calculate the absorbed power from the three sides facing the Sun. We obtain
αSun · ASun · Gs = 0, 88 · 0, 0110 · 1371 = 13, 27 W
(11)
for the solar cells and
αSun · ASun · Gs = 0, 24 · 0, 0062 · 1371 = 2, 04 W
(12)
for the aluminum frame. The absorbed power from the Earth infrared radiation is
αEarth · AEarth · q1 = 0, 08 · 0, 01 · 230 = 0, 184 W.
(13)
The absorbed power from the Earth albedo radiation is
a · αSun · ASun · Gs = 0, 24 · 0, 01 · 1371 · 0, 3 = 0, 98 W.
(14)
When we sum up the powers of Equations (11)(14) and add to the sum the internally dissipated power, which is approximately 2W, we obtain the total absorbed power 18,59W. This value corresponds to the numerator inside the fourth root in Equation (7). The denominator is obtained as
σ · · A = 0, 0000000567 · (0, 032 · 0, 8 + 0, 028 · 0, 08) = 1, 57853 · 10−9 W K −4 . (15)
◦ Now, using Equation (7), we get the equilibrium temperature 50,11 C. ESTCube-1 SSMD
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7 240 THERMAL CONTROL SYSTEM (TCS) Table 26: Summary of equilibrium temperatures. ◦ Phase Equilibrium Temp, C
7.7.2
Sun
50,11
Eclipse
−80, 14
Temperature equilibrium of the spacecraft at the eclipse
The calculation is performed with the same principle as in section 7.7.1. In the eclipse phase the absorbed power is originated solely from the Earth infrared radiation.
When using the numbers given above, we obtain the ◦ equilibrium temperature -80,14 C.
7.7.3
Summary
Table 26 summarizes the equilibrium temperatures. Assuming that the Sunlight time is two thirds of the time and the eclipse time is one third then ◦ the equilibrium temperature of the spaceship should be about 5 C. Calculated average temperature estimate is close to the RT that is the goal of TCS. Using dierent passive thermal control methods, permanent sub RT environment is very likely achievable.
7.8
Thermal control mechanisms
As it is necessary to maintain temperature in the satellite constantly and steadily near RT, several methods have been developed in order to achieve this. Methods of thermal control can be divided into two categories: passive and active methods. Due to the small size of a Cubesats and relatively small and simple payload, passive methods are ordinarily used as they contribute very little to the overall weight, space and energy budget; also the possibility of a system failure during the mission is minimal.
7.8.1
Passive control
Passive control means using dierent kind of surface nishes (with dierent
α/
ratios), paints, coatings or Multilayer Insulation (MLI) to provide
necessary temperature ranges for Cubesat and its components. Preliminary intentions for the ESTCube are to use only passive methods for temperature
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7 240 THERMAL CONTROL SYSTEM (TCS) control. In addition to surface coatings and paints (on the surface and inside of the satellite) various insulation materials will be considered for maintaining components inside the cube near RT (Te on tapes, Kapton tapes, Kapton lms etc.). As there is no knowledge about the size and thermal properties of the components of the spaceship so far, precise requisites, materials and amounts are yet to be determined.
7.8.2
Active control
When it has been proved being impossible to meet the requirements for the temperature control by passive methods, active methods have to be used. Active control methods use heaters, heat pipes, radiators and refrigerators etc.
in order to achieve necessary temperatures for satellite components.
At the moment none of the active methods are being actively considered for ESTCube. When such a necessity comes up (in case of too low temperature in some compartment or node of a spacecraft) small heater-thermostat system will probably be introduced to the region of interest.
7.9
Conclusions of the preliminary status of TCS First calculations show that raw body of our Cubesat is on average already near suitable RT.
Equilibrium temperature maximum and minimum values for sunlight and eclipse respectively, exceed probable working temperature ranges for Cubesat components as predicted.
Steady sub RT environment is very likely achievable using only passive methods (surface coating and insulation).
7.10
Active thermal control should and probably can be avoided.
Future steps and plans Accurate attitude determination and orbital analysis is needed for precise calculations of dierent radiation eects on satellite surface. Orbital model of our Cubesat will be made using STK software.
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7 240 THERMAL CONTROL SYSTEM (TCS)
Precise 3D model of Cubesat and all of its components has to be made using CAD software. Without 3D modeling exact temperature uctuations on and inside satellite cannot be predicted. Requirements for the other subsystems: all the components, their ma-
terial, size, shape and probable location and orientation inside satellite have to be known (tight collaboration with 210 STR). Positioning of the components is extremely important as the more temperature sensitive components can be shielded from regions with harmful temperature by using insulation materials as well as other components of Cubesat.
Thermal model of the Cubesat will be made out of the 3D model. Thermal capacity and thermal conduction values of all the components and TCS elements have to be considered in order to make a precise thermal model. All possible passive thermal control materials and their characteristics have to be gathered. Engineering software ANSYS Workbench will probably be used for making a thermal model.
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8 250 COMMUNICATIONS SYSTEM (COM)
8 250 Communications system (COM) 8.1
Introduction
The communication subsystem is responsible for the communication between the ground station (GS) and the spacecraft.
It can receive telecommands
from the GS for setting dierent operating modes and requests to transmit data. There are two dierent types of downlink transmission modes:
LPTM - Low Power Transmission Mode (Beacon)
HPTM - High Power Transmission Mode (Data)
The beacon data contains the most important system information, like battery power and system status.
The HPTM is used for transmitting large
amounts of mission data, for example a picture taken by the camera. Because COM subsystem needs to be fault tolerant and reliable, it has autonomous modes for low-energy situations and dierent recovery operations in case errors occur.
8.2
System Members
Members of the COM team are listed in Table 27. Table 27: Members of the COM team. Name University Role
8.3 8.3.1
Urmas Kvell
UT
Subsystem coordinator
Siim Meerits
UT
Subsystem member
Toomas Vahter
TUT
Subsystem member
System Description Deduced requirements from the Cubesat standard
Based on the Cubesat standard, the COM can start its work for sending beacon after 15 minutes, and data after 30 minutes after ejection. Naturally the restrictions for the size and mass, dictated by the standard, also in uence the COM system.
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8 250 COMMUNICATIONS SYSTEM (COM)
Figure 11: Block diagram of the COM subsystem.
8.3.2
Deduced requirements from the Payload
During the experiment the satellite will start spinning in one axis and communication needs to be working during that time. The experiment will include an electron gun which will give the satellite a positive charge.
8.3.3
Layout Proposals
Figure 11 illustrates the proposed block diagram of the COM subsystem. The system consists of two antennas, a switch, a beacon, a transmitter, a receiver and a micro controller. The downlink antenna is intended for sending the housekeeping data and the data generated by the payload. The uplink antenna is for receiving the telecommands sent from the GS. There is a switch which controls the usage of the downlink antenna so that we can use the same antenna for sending housekeeping and payload data. COM micro controller is used for controlling the data ow from the receiver and to the transmitter. There is a latch for preventing short circuit.
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8 250 COMMUNICATIONS SYSTEM (COM) 8.3.4
Possible problems Due to the experiment phase spin axis, two antennas might not enable working communications during the whole experiment. The issues related to antenna positioning are further discussed in Section 5.3.3.
8.4
The electron gun might disturb the communications.
Link budgets
Downlink
Antenna chosen for the downlink is a half-wave dipole with
antenna gain 2,15dBi. E ciency of the dipole antenna is 99,6%. The chosen modulation scheme is FSK and the required Eb/N0 for BER 10-5 is 13,8 dB. FSK is the best choice as we can use it non-coherently (demodulator does not need a backing signal with the exact frequency and phase on the receiver side).
Uplink
Antenna chosen for the uplink is a quarter-wave monopole with
antenna gain 3dBi. A monopole will have a lower input resistance, resulting in overall lower e ciency compared to dipole antenna. Chosen modulation scheme is non-coherent FSK and the required Eb/N0 for BER 10-5 is 13,8 dB.
8.4.1
Downlink budget
Table 28 shows the downlink budget.
8.4.2
Uplink budget
Table 29 shows the uplink budget.
8.4.3
Beacon
Table 30 shows the beacon link budget.
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8 250 COMMUNICATIONS SYSTEM (COM)
Table 28: Link budget: Downlink
Parameter
Value
Unit
0,5
W
-3,01029996
dBW
26,9897
dBm
Frequency
0,44
GHz
Transmit Antenna Gain
2,15
dBi
Transmission Line Losses
2
dB
EIRP
-0,86029996
dBW
Range
800
km
Free Space Loss
-143,368221
dB
Antenna Pointing Loss
-3
dB
Atmospheric/Ionospheric Loss
-2
dB
Satellite to Ground Polarization Loss
-3
dB
Total loss:
-151,368221
dB
Antenna Gain
17
dBi
System Noise Temperature
550
deg K
Transmission Line Losses
2
dB
Antenna Pointing Loss
1
dB
Figure of Merit (G/T)
-10,4036269
dB/K
1200
bit
30,7918125
dBHz
k
-228,6
dB
Final Eb/No=EIRP + G/T - L - 10log(rb) - k
35,1760397
dB
Required Eb/No for given modulation and coding
13,8
dB
Modulation implementation loss
1
dB
Link margin=Final Eb/No-Required Eb/No
20,3760397
dB
Transmitter (Satellite) Transmit Power
Channel
Receiver (Ground Station)
General Bit Rate
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8 250 COMMUNICATIONS SYSTEM (COM)
Table 29: Link budget: Uplink
Parameter
Value
Unit
20
W
13,0103
dBW
43,0103
dBm
Frequency
0,14
GHz
Transmit Antenna Gain
17
dBi
Transmission Line Losses
5
dB
EIRP
30,0103
dBW
Range
800
km
Free Space Loss
-133,421728
dB
Antenna Pointing Loss
-3
dB
Atmospheric/Ionospheric Loss
-2
dB
Satellite to Ground Polarization Loss
-3
dB
Total loss:
-141,421728
dB
Antenna Gain
2,15
dBi
System Noise Temperature
270
deg K
Transmission Line Losses
2
dB
Antenna Pointing Loss
3
dB
Figure of Merit (G/T)
-22,1636376
dB/K
1200
bit
30,7918125
dBHz
k
-228,6
dB
Final Eb/No=EIRP + G/T - L - 10log(rb) - k
64,2331217
dB
Required Eb/No for given modulation and coding
13,8
dB
Modulation implementation loss
1
dB
Link margin=Final Eb/No-Required Eb/No
49,4331217
dB
Transmitter (Ground Station) Transmit Power
Channel
Receiver (Satellite)
General Bit Rate
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8 250 COMMUNICATIONS SYSTEM (COM)
Table 30: Link budget: Beacon
Parameter
Value
Unit
0,1
W
-10
dBW
20
dBm
Frequency
0,44
GHz
Transmit Antenna Gain
3
dBi
Transmission Line Losses
2
dB
EIRP
-7
dBW
Range
800
km
Free Space Loss
-143,368221
dB
Antenna Pointing Loss
-3
dB
Atmospheric/Ionospheric Loss
-2
dB
Satellite to Ground Polarization Loss
-3
dB
Total loss:
-151,368221
dB
Antenna Gain
17
dBi
System Noise Temperature
550
deg K
Transmission Line Losses
2
dB
Antenna Pointing Loss
1
dB
Figure of Merit (G/T)
-10,4036269
dB/K
1200
bit
30,7918125
dBHz
k
-228,6
dB
Final Eb/No=EIRP + G/T - L - 10log(rb) - k
29,0363397
dB
Required Eb/No for given modulation and coding
14
dB
Modulation implementation loss
1
dB
Link margin=Final Eb/No-Required Eb/No
14,0363397
dB
Transmitter (Satellite) Transmit Power
Channel
Receiver (Ground Station)
General Bit Rate
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8 250 COMMUNICATIONS SYSTEM (COM) Table 31: Weight and power budget of the COM system; best case estimates.
Component
Qty
Mass (each)
Power (each)
Monopole antenna
1
11 g
Dipole antenna
1
7 g
0,5 W
Receiver
1
15 g
16 mW
Transmitter
1
15 g
10 mW
Controller
1
8 g
20 mW
Memory
2
8 g
25 mW
PCB
1
32 g
Coating lack
1
10 g
LPTM
0,1 W
HPTM
0,5 W
P (LP T M ) P
114 g
196 mW
114 g
596 mW
(HP T M )
8.5
Mass and power budgets
Tables 31 and 32 show the best and worst case estimates for the weight and power budget.
8.6
Conclusion
The current con guration of the COM subsystem is able to support communication without the need for nadir pointing during the experiment when the satellite is spinning around one axis. Current mass estimations are quite rough and need to be improved. To increase the resistance to failure we need to use a separate receiver and transmitter. Also separate memory chips are needed to provide our own memory buer, separated from the CDHS subsystem. It is best to consider the sending power of 1W for HPTM but we can estimate the mass about 114 g (with one transceiver) + 15 g if we want to include two transceivers.
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8 250 COMMUNICATIONS SYSTEM (COM)
Table 32: Weight and power budget of the COM system; worst case estimates.
Component
Qty
Mass (each)
Power (each)
Monopole antenna
1
18 g
Dipole antenna
1
11 g
1 W
Receiver
1
15 g
16 mW
Transmitter
2
15 g
10 mW
Controller
1
8 g
20 mW
Memory
2
8 g
25 mW
PCB
1
32 g
Coating lack
1
10 g
LPTM
0,1 W
HPTM
1,0 W
P (LP T M ) P
140 g
206 mW
140 g
1106 mW
(HP T M )
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS)
9 260 Command and data handling system (CDHS) 9.1
Introduction
The mission objectives of a satellite mission are spread in a wide eld of applications. From scienti c missions over Earth monitoring to communication applications, but the combined objectives are to collect data and send these down to Earth. For this purpose an on-board data system is required to collect, administrate and store the generated data. The type of mission, orbit, payload and the selected ground stations have a signi cant eect on this on-board data system.
The collected data can be transmitted from a
geostationary satellite constantly, because the spacecraft is visible permanently from one ground segment. A circling satellite in the low-Earth-orbit requires a reliable data storing during the non-contact periods between the spacecraft and the ground station(s). Due to the short LEO/ground station link time the internal generated data of a satellite do not have to exceed the sendable data volume during one day. The CDHS is the data- and telecommand administration system of ESTCube1. The CDHS memories store the generated internal data of the satellite like housekeeping data, experiment data and the transmitted telecommands from the ground station. In a regular time interval the system is requesting data from each system and stores this data in the memory devices. The actual design schedules one memory device for the housekeeping data, one for the experiment data and one for the telecommands. The incoming telecommands will be forwarded to the according systems of the spacecraft.
9.2
System Members
Members of the CDHS team are listed in Table 33. Johannes Piepenbrock is a visitor from the Aachen University of Applied Sciences (FHAC). He was a team member in the Compass-1 project, the rst student satellite project of FHAC, and is currently the project manager of the Compass-2 project. Piepenbrock has acted as an advisor for the members of the ESTCube team.
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) Name
Table 33: Members of the CDHS team. University Role
Priit Laes
UT
Subsystem coordinator
Indrek Sünter
UT
Subsystem member
FHAC
Subsystem member
Johannes Piepenbrock
9.3
Functionality
Internal functions of an onboard data handling system contain the following:
Enabling the ow of housekeeping and mission data
Receiving and distributing commands
Performing telemetry and telecommands protocols
Time distribution around the spacecraft
Providing data storage
Executing commands and scheduled events
Controlling payloads and subsystems
Monitoring spacecraft health
Making autonomous decisions
Performing data compression
9.3.1
Deduced requirements from the Cubesat standard
The main requirements for cubesats are in respect to the total weight of maximum 1kg (single-unit cubesat) and the dimensions of
100mm.
100mm ∗ 100mm ∗
In respect to the CDHS there are only requirements to the size of
the PCB-board. The size is scheduled to be about
90mm ∗ 90mm ∗ 2.5mm.
Also the power input is limited concerning to the de ned surface area of the spacecraft. Otherwise, there are no further requirements deducible from the Cubesat standard concerning the CDHS.
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS)
Figure 12: Proposed sytem layout of CDHS.
9.3.2
Layout proposal
Figure 12 shows the proposed system layout of the CDHS. All electrical devices like the MCU, Memory devices and the other electrical devices (resistors, capacitors) are surface mounted devices (SMD). The advantages are:
Weight reduction, because no additional connection wires are required
Low power consumption of the SMD devices
Space reduction, because the PCB size and the space inside the satellite are limited
Improvement of HF-characteristic (shorter circuit paths, lower resistance, lower impendance)
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) 9.3.3
Technical requirements
The technical requirement consist of the total weight, the power consumption for CDHS and data storage needs for the whole satellite.
9.4 9.4.1
Components Microcontroller selection
Microcontroller is a functional computer system-on-a-chip. It contains a processor core (CPU), memory (RAM/ROM), and programmable input/output peripherals. The microcontroller selection was mostly aected by hard requirements the controller needs to meet and soft requirements that generally simplify the development of required functionality.
Hard requirements for a microcontroller:
availability of discrete input and output bits, allowing control/detection of the logic state of an individual package pin.
weight and size
low power mode availability
integrated clock generator
Soft requirements for a controller:
serial input/output (UART)
2 I C or CAN controller
in-circuit programming/debugging support
low pin count
watchdog timer
brownout detection
open source toolchain
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) 9.4.2
8-bit Atmel AVR Family
After listing all the dierent controllers from various hardware families and various CPU architectures, the nal choice became Atmel 8-bit AVR family.
Pros
Integrated UART/SPI support Integrated TWI (I2 C) support The AT90CAN32/64/128 series also have an integrated CANcontroller (supporting 2.0A and 2.0B protocol)
5 dierent sleep modes Low power consumption Open source toolchain available
Cons
8-bit instruction set (no data crunching needed, but memory access becomes a bit more complicated)
We also considered following microcontroller families extensively:
16-bit PIC24F family
lack of convenient open source toolchain
16-bit MSP430x
Although open-source toolchain exists (http://mspgcc.sourceforge.
net/),
it is outdated
STxx series from STMicroElectronics
No open source toolchain. We chose the AT90CAN128-16AU microcontroller from the 8-bit Atmel AVR 2 Family because it has integrated UART, I C (TWI) and CAN controllers. Atmel also has an automotive product line and AT90CAN128-16AU is one of those products. The AT90CAN128 can bear temperatures ranging from ◦ ◦ -40 C to 85 C, which should be enough.
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) 9.4.3
Memory device selection
The memory devices were chosen from Atmel as well. Three AT25DF321A chips were chosen because they have low power requirements, fast program and erase times and can be interfaced over SPI bus (Serial Peripheral Interface).
This also means that we do not need extra multiplexer for commu-
nicating with these devices. The memory chip with a temperature range of ◦ ◦ -55 C to 125 C is even more tolerant than the MCU.
9.5 9.5.1
Budgets List of Components
Table 34: List of components. Component
Product code
Vcc Range
Main controller
Atmel AT90CAN128-AU
2.7-5.5V
Memory chip (3x)
Atmel AT25DF321A
2.3-3.6V
Other electr. comp.
-
-
9.5.2
Notes
Not chosen
Power usage characteristics
Main Controller (Atmel AT90CAN128-AU)
Main controller in our
case is the central hub of the whole system that collects data coming from various subsystems and acts based on the analyzes of incoming data. It also manages various payload systems during experiment phases. See Table 35.
Memory chips (Atmel AT25DF321A)
32-Mbit/8-MiB memory chips
are meant for storing housekeeping and payload data and use SPI protocol for communicating with master microcontroller. For operation they require 2.7-3.6V. Their power requirements are listed in Table 36.
Other electrical components
An estimated 20mW of power should be
reserved for components that have not been chosen yet.
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) Table 35: Main controller power requirements. Power usage (Vcc=5V/3V) 125
Operating mode (8 Mhz)
1
Active Idle
14mA/7mA
1
Power down mode
14mA/7mA
4
2 3
0.12/0.08mA
5
1
Using internal RC oscillator 2 AT90CAN128-AU Datasheet: Figure 29-3. Active Supply Current vs. Vcc (Internal RC Oscillator 8 MHz) 3 AT90CAN128-AU Datasheet: Figure 29-9. Idle Supply Current vs. Vcc (Internal RC Oscillator 8 MHz) 4 Watchdog enabled 5 AT90CAN128-AU Datasheet: Figure 29-15. Power-down Supply Current vs. Vcc (Watchdog Timer Enabled) - Temp.= 125
Table 36: Power requirements for memory chip Operation
Power requirement
Operation speed (worst case)
Typical
Max
Write (Program)
10 mA
15 mA
5ms (for 256 Bytes block)
Read
10 mA
14 mA
-
Erase
12 mA
18 mA
200/600/950ms (4/32/64 Kbytes)
Sleep/Standby
25
Deep powerdown
10
1
µA µA
50 10
µA µA
1
10ms (wakeup time) 10ms (wakeup time)
1 - Full chip erase takes between 32-56 sec
9.5.3
Mission Phases During Orbiting
Power requirements on orbit can be divided into three main phases during which all have dierent power usage patterns and therefore dierent power needs.
Pre-experiment phase
Only collection and storage of housekeeping data
and optional communication with communication subsystem take place. Once every minute, main controller wakes up, requests sensor data from other subsystems, stores it in ash memory 1 and assembles and forwards beacon data to the communication subsystem. In addition to waking up once every minute in order to request sensor data from other subsystems, the main controller can also receive exceptional wakeups for taking pictures.
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) When taking pictures, we need to fetch data from camera, compress it and store it in memory 2. The other two memory devices will be in low-power suspend state during that time.
The current that is taken up by controller ports is still to be calculated.
Main Controller: Flash1:
active
erase operation
Flash2, Flash3:
standby mode
Table 37: Power consumption in the pre-experiment phase. Calculations Maximum peak Standby
Power
20 ∗ 5 + (14 + 2 ∗ (0.035)) ∗ 3 ≈ 145 mW 7 ∗ 5 + (3 ∗ 0.035) ∗ 3 ≈ 36 mW
The average power usage during this phase depends on the frequency of incoming picture-taking requests.
Assuming that one picture is 400KB of
data, the power consumption of the storage, reading and erasing processes can be calculated. See Tables 38 and 39.
Table 38: Worst case power consumption during picture handling. Process
Calculations
erase
14 ∗ 5 + (20 + 2 ∗ (0.035)) ∗ 3 ≈ 131 14 ∗ 5 + (18 + 2 ∗ (0.035)) ∗ 3 ≈ 125 14 ∗ 5 + (14 + 2 ∗ (0.035)) ∗ 3 ≈ 113
write read
Power
Duration
mW
5 seconds
mW
8 seconds
mW
-
Table 39: Typical power consumption during picture handling. Process erase write read
Calculations
Power
Duration
0.12 ∗ 5 + (20 + 2 ∗ (0.035)) ∗ 3 ≈ 622mW 14 ∗ 5 + (12 + 2 ∗ (0.035)) ∗ 3 ≈ 107 mW 14 ∗ 5 + (10 + 2 ∗ (0.035)) ∗ 3 ≈ 101 mW
1
5 seconds 8 seconds -
1 Worst case
2 Main controller can be in power-down mode when the memory device is erasing a 64 KBytes block
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) Experiment phase
Power consumption during the experiment mode is
shown in Table 40. During this phase we possibly need to monitor the engine and camera, so main controller has to be active during the whole sequence. During experiment we need to fetch data from payload which is stored on ash memory 2. Due to design considerations, it is possible to access only one memory chip at a time, which means that when operating with one chip, other two will be in low-power suspend state. Table 40: Power consumption in the experiment phase. Calculations Maximum peak
1
Standby
Power
20 ∗ 5 + (14 + 2 ∗ (0.035)) ∗ 3 ≈ 145 18 ∗ 5 + (14 + 2 ∗ (0.035)) ∗ 3 ≈ 135
mW mW
1 For up to 56 seconds in case we want to erase the whole chip
Post-experiment phase
Power consumption during the post-experiment
mode is shown in Table 41 From the perspective of power consumption, the post-experiment phase is equal to the pre-experimnet phase.
See the pre-
experiment phase section for more information. Table 41: Power consumption in the post-experiment phase. Calculations Maximum peak Standby
1
Power
20 ∗ 5 + (14 + 2 ∗ (0.035)) ∗ 3 ≈ 145 mW 7 ∗ 5 + (3 ∗ 0.035) ∗ 3 ≈ 36 mW
1 For up to 56 seconds in case we need to erase the whole chip The average power usage during this phase depends on the frequency of incoming picture-taking requests.
Power Budget
Table 42 summarises the power need of the CDHS during
the dierent phases of operatrion.
9.5.4
Mass budget
The mass budget of CDHS stays constant throughout all experiment stages and only depends on the mass of the electrical circuit. The mass budget is shown in Table 43.
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) Table 42: Power consumption of the CDHS. Phase
Average
Maximum requirement
1
≈ 50mW ≈ 135mW ≈ 50mW 1
Pre-experiment Experiment phase Post-experiment
145mW (up ≈ 145mW ≈ 145mW 2
to 5 seconds)
1 Controller is either active for 5 seconds once every minute (assuming that no pictures are being taken) 2 During picture taking, packing Table 43: Mass budget of the CDHS. Number
1
Component
Mass [g ]
1
AT90CAN128
0.5
3
AT25DF321A
0.3
x
Extra components
1
4
Sum weight [g ]
1
0.5
2
0.9
1 2
3
-
5
PCB
-
32
1
Coating lack
-
10
1
Solder
-
15
x
Wires
-
30
Σ
-
-
≈ 95
AT90CAN128
is
available
in
the
following
3 3 3 3
packages:
TQFP64 (0.35 - 0.4 grams), QFN64 (0.2 grams). AT90CAN128-16AU is probably TQFP64. 2 AT25DF321A is available in the following packages: 8MA1, 8S2. Masses are still to be determined. 3 Those are pretty rough estimations 4 Passive components like resistors, capacitors
9.5.5
Data budget
The data budget for the CDHS is shown in Table 44. During 24 hours the systems are generating the following amount of housekeeping data: bit (370 + X) min × 1440min∗Byte∗KByte ≈ 65 + XKByte 8bit∗1024Byte The payload data memory size is estimated to be
≈
4000 KByte in total.
With this memory it is possible to store 10 images with a size of 400 KByte each image.
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9 260 COMMAND AND DATA HANDLING SYSTEM (CDHS) Table 44: Data budget of the CDHS. System
Number
Data
Data size [bit ]
Sum [bit ]
210 STR
2
Antenna deployment
1
2
210 STR
1
System status -TBD-
2
2
220 ADCS
1
System status
2
2
220 ADCS
5
Sun vector
20
100
220 ADCS
1
Power consumption
10
10
220 ADCS
1
Calculated vector
20
20
230 EPS
1
System status
2
2
230 EPS
8
INPUT-Power
12
96
230 EPS
1
OUTPUT-Power
12
12
240 TCS
6
Temperature
4
24
240 TCS
-
-
-
-TBD-
250 COM
1
System status
2
2
250 COM
-
-
-
-TBD-
260 CDHS
1
System status
2
2
260 CDHS
3
WRITING-Pointer
16
48
260 CDHS
3
READING-Pointer
16
48
Σ
-
-
-
370 + X
270 Camera
1
System status
16
16
270 Camera
10
Image
3276800
32768000
-
-
3276816
32768016
Σ 9.6
Conclusion
There are no unresolvable problems foreseen in the CDHS subsystem, although mass and power budgets are still rough estimates that are needed to be con rmed on real hardware.
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10 270 Payload (PL) 10.1
10 270 PAYLOAD (PL)
Introduction
The payload (PL) subsystem includes the essential hardware and instruments needed for the tether experiment, the most essential one being the tether itself. The following sections give a closer view of the system.
10.2
System Members
The payload is being designed and manufactured by an international team, coordinated by FMI. The team includes members from FMI, University of Jyväskylä Finland (JU), German Aerospace Center (DLR) and University of Helsinki (HU). Table 45 gives a list of the people most involved in the PL design. The list is not inclusive.
Name
Table 45: Members of the PL team. Institute Role
Pekka Janhunen
FMI
Scienti c manager (overall)
Walter Schmidt
FMI
Technical manager (overall)
Jouni Polkko
FMI
Overall design
Harri Haukka
FMI
Overall design
Markku Mäkelä
FMI
Overall design
Pekka Riihelä
FMI
Overall design
Hannu Koivisto
JU
Electron gun
Olli Tarvainen
JU
Electron gun
Lutz Richter
DLR Bremen
Tether reel
Olaf Krömer
DLR Bremen
Tether reel
Edward Haeggström
HU
Tether manufacture
Henri Seppänen
HU
Tether manufacture
Risto Kurppa
HU
Tether manufacture
10.3
System description
The space for the payload is in the centre of the upper panel of the cube. The spin axis of the satellite should be perpendicular to the tether deployment plane. The payload compartment is 15 mm high and (excluding the satellite
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10 270 PAYLOAD (PL)
Figure 13: Proposed layout for the payload.
walls) 95 mm
×
95 mm wide.
It has three openings for the tether, the
camera and the electron gun. Figure 13 presents one possible arrangement of the payload components. The purpose of this gure is only to show that there is enough space to t all components. In any placement, it is sensible to put the reel motor next to the capstan where its torque is needed, and if the reel drum is also motorised then it must be close to the motor as well. The camera and the electron gun must be mounted at the surface. The electron gun should not be very close to the tether.
10.4
Hardware
Tether reel
The proposed reel is a 6-cm diameter, 1-cm wide drum, which
is capable of holding a 9 mm wide Hoytether. The mass of the reel is 11,3 g, estimated by assuming that the drum is made of 1 mm thick plastic (density 1,5).
Tether reel bearings and xing
3 Estimated 2 cm of Al equivalent needed,
mass 5.4 g.
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10 270 PAYLOAD (PL) Tether reel motor
Estimated mass is 10 g. The baseline approach is that
the capstan is motorised but the reel drum is only braked with passive friction. This design is not capable of retracting the tether. Whether retraction ability is wanted or not in this mission is not yet clear and should be decided. At the rst phase we will design a non-retracting version and perhaps later consider going to a retractable design, if it looks technically feasible, risk-free and ts to the available mass and power budget at that stage.
Tether reel launch lock
This mechanism will prohibit the reel from ro-
tating during launch. Estimated mass is 5.65 g (50% of reel mass).
Tether capstan mechanism
The capstan is a rotating motorised reel
made of elastic rubber-type material, which presses towards a passively rotating metal axis, a design mimicked from cassette tape decks. If the capstan reel is 1 cm in diameter and 1,5 cm in length, its mass is 1,8 g if made of plastic (density 1,5). The metal axis mass is 1 g if made of steel, assuming a 0,3-cm diameter. If one reserves 1 g each for reel and axis bearings, the total mass is 5 g after rounding upward.
Tether itself
The proposed tether is a 50
µm
diameter Al, 4,3-times
Hoytether, with 10 m length, mass 0,2 g.
Tether end mass
0.5 g is close to optimal. At this end mass, if the initial
spin is 1 revolution per second and the tether length is 10 m, at the end of deployment the spin period is 35 s and the tether tension at the tip is 17 mg (equivalent to 73 cm of tether hanging at earth). If the end mass is heavier, the nal spin period will be longer and the nal tension lower, which may not be enough to straighten the tether. On the other hand if the end mass is lighter, the tension at the start of deployment may be dangerously low. Making the end mass much lighter than 0,5 g does not appreciably increase the nal tension since the tether's own mass 0,2 g also contributes.
Tether end mass xation
The end mass must be xed during the launch
at the tether opening and released at the start of the deployment. Since we do not know yet how the xation will be implemented, let us reserve 2 g for it (four times larger than the end mass itself ), plus 1 g for cabling and connector since the release signal must be transmitted from the controller.
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10 270 PAYLOAD (PL) High voltage (HV) source
A 1 W, 200 V voltage source is needed in the
tether experiment. Estimated mass is 25 g. Assuming a DC-DC converter with AC intermediate stage, the devise needs a transformer with coils. Therefore it cannot be extremely lightweight. This HV source is also responsible for producing the cathode heater voltage (∼1.5 V) for the electron gun.
Wiring
Insulated single wire, 2,75 g/m. Total of 3 g reserved.
Electron gun
A 5 mA, 200 V, 1 W device, with a 0,2 W heater has been
proposed. The planning of the electron gun has begun. We plan to have two redundant guns, each of them ts in about 5 mm The electrons exit from a 0,6 mm
×
×
5 mm
×
15 mm box.
13 mm slit. The beams are inclined to
shoot away from the satellite and the tether spin plane. We reserve 3 g for each gun.
In Figure 13, the internals of the gun have been magni ed and
drawn in vertical instead of the actual horizontal placement.
Tether current measurement
Could be almost free depending on how
HV source is constructed, but 2 g is reserved for it.
Camera with electronic adapter
Estimated mass for the camera is 15 g.
Most of the mass is for the electronics.
Electric connector between payload and spacecraft
Single Micro-9
pin connector with screws is 1 g. We want to eliminate connectors as much as possible and replace them with soldered joints, but this one is good to have to avoid hassle in the nal integration phase.
Mechanical xation of payload
Some screws etc., total 5 g.
This in-
cludes all extra mass needed for attaching the components of the payload (reel, motor, bearings, HV source, camera, electron gun) to the satellite body.
10.5
Mass and power budget
The mass and power budget of the payload subsystem are given in Table 46.
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10 270 PAYLOAD (PL) Table 46: Mass and power budget of payload.
Tether reel
Power cons, W
11,3
Bearings, xings
5,4
Reel motor
10
TBD
Launch lock
5,7
TBD
Capstan
5
Tether
0,2
End mass
0,5
End mass xing
3
TBD
25
1,1
Wiring
3
E-guns (2)
6
0,2 (heater)
Current meas.
2
TBD
15
TBD
Connector
1
Mechanical x.
5
98,1
TBD
HV source
Camera
Σ 10.6
Mass, g
Conclusion
It seems that one can implement the payload with 100 g. Some mass could be saved by reducing the diameter of the reel drum from 6 cm to 45 cm, but maybe at some risk of curling up the tether in long-term storage. It should also be noted that the overall satellite mass must be symmetrical around the rotation axis. This may require some additional mass to compensate for the non-symmetric mass distribution of the payload components (heaviest individual component is the HV source).
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11 CONCLUSION
11 Conclusion
Table 47 gives the mass budget of ESTCube-1 for three scenarios, which relate to the design choices made for the STR and ADCS. Very small reductions, if any, could be made for the masses of other subsystems.
The
scenarios are:
Scenario 1
The skin panels of the satellite are made of aluminum, the
spinning is executed with the reaction wheel (variant A or C, see Section 5.4.2).
Scenario 2
The skin panels of the satellite are made of CFRP (see Section
4.3.1), the spinning is executed with the reaction wheel (variant A or C).
Scenario 3
The skin panels of the satellite are made of CFRP, the spinning
is executed with the magnetic torquer coils (variant B). It is unlikely that Scenario 1 could be utilized, even if the weight could be further cut down. Scenario 3 would oer the most safe margin, but it would force us to use the magnetic torquers to actuate the spinning. The use of the reaction wheel would be preferable due to the better stability and control, making Scenario 2 the most appealing one with respect to the operational viewpoints. In Scenarios 1 and 2 there is still the open question of whether to use variant A or C for the experiment mode.
The use of either variant would ensure
Table 47: Mass budget of ESTCube-1. Subsystem
Mass, g
Mass, g
Mass, g
Scenario 1
Scenario 2
Scenario 3
STR
268
220
220
ADCS
316
316
246
EPS
146
146
146
TCS
-
-
-
COM
129
129
129
CDHS
95
95
95
PL
98
98
98
P
1052
1004
934
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11 CONCLUSION a rapid buildup of angular velocity around a well de ned axis.
Variant A
would be simpler of the two. On the other hand, variant C would require considerably less electrical power during the tether experiment. This would provide us with the opportunity to conduct the experiment more frequently. The power budget, given in Table 20, indicates that the mission operations, including the tether experiment, can be carried out with the available electrical power. It is the opinion of the Estonian Student Satellite Team that the mission described in this document is viable and that the development of ESTCube-1 should move onto phase A.
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REFERENCES
References
[1] Towards an Estonian space policy and strategy, 2008, Enterprise Estonia [2] http://www.cubesat.org/ [3] R. Hoyt and R.L. Forward, Alternate interconnection Hoytether failure resistant multiline tether, US Pat. 6286788 B1, 2001. [4] http://www.electric-sailing.com/ [5] P. Janhunen and A. Sandroos, Simulation study of solar wind push on a charged wire: basis of solar wind electric sail propulsion, Ann. Geophys.
25, 755767 (2007); available online as http://www.electric-
sailing.com/paper2.pdf [6] P. Janhunen, On the feasibility of a negative polarity electric sail, Ann. Geophys.
27, 14391447 (2009); available online as http://www.electric-
sailing.com/paper3.pdf [7] http://www.celestrak.com/NORAD/documentation/spacetrk.pdf [8] M. Polaschegg, Study of a Cube-Sat Mission, Submitted in partial ful llment of the requirements for the degree of Master of Sciences, Karl Franzens University, Graz, Austria (2005). [9] http://www.matweb.com/ [10] http://www.portescap.com/ [11] http://www.clyde-space.com/documents/35 [12] http://swisscube.epf l.ch/webdav/site/cubesat/shared/technical_documentation/s3a-eps-1-0-eps.pdf [13] P. Fortescue, J. Stark, and G. Swinerd, Spacecraft Systems Engineering, John Wiley & Sons, UK (2003). [14] http://www.raumfahrt.fh-aachen.de/
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