1 Thermal Control
1.1 Preface
This system must maintain a thermal control over all the components of the satellite (especially electronic devices). This
objective can be achieved balancing incoming and outgoing thermal fluxes. Components operate most efficiently when
preferred operating temperatures are maintained. In addition, component life span will be degraded by either operating or
storing the component outside of its preferred temperature range.
1.2 Orbital Environment
The most critical global conditions that can occur during the mission are:


eclipse or ‘shadow’ condition (low thermal fluxes),
both Sun and Earth radiation incident (high thermal fluxes).
The temperatures are characterized by very different environmental conditions: these depend on solar power radiation and
operative modes of the spacecraft.
Satellite heat loads must be calculated for each orbit. Primary terms of the heat load will include incident solar radiation,
albedo, and Earth generated thermal radiation.
In this analysis, fundamental equations of heat transmission will be used and a few assumptions will be made in order to
simplify the calculus.
There are two different scenarios of mission:
Orbital parameters
Assembling Space Station
Assembling Spacecraft to Mars
h [km]
600
900
i [°]
5
51.8
1.3 Temperature Constraints on Components
The following temperature constraints are specified for each component:
Subsystem
Power
GNC
Telecom
CPU
Propulsion
Thermal
Equipment
Operating limits [°C]
Solar Array
Battery
Sun Sensor
Star Mapper
Momentum Wheel
RVS
Antenna
TM/TC
Network
CPU
Hydrazine
Xenon
Sensors
Heaters
-90
5
-80
-20
-20
0
-80
-40
-28
-50
2
0
-150
-32
Peak Power [W]
160
30
80
40
70
55
80
60
70
100
80
90
150
100
134
0.3
0.5
150
0.5
1
10
15
5
0
0
0.1
see details
From the table, it can be seen that the most critical components are the batteries and the docking sensor (RVS). Propellant
and pressurizing gas will be treated in order to keep a temperature around 17°C.
The structure of the satellite is made with Aluminum 2219 which thermal conductivity is 150 W/m/K.
1.4 Hot case and cold case
The satellite for the baseline mission will have a square base.
The worst situations in this case are:
Hot Case Cube:
1. Radiation from the Sun:
 Load assumed to strike at 45° on two faces,
 Effective area exposed to solar radiation is 24m2,
 Solar emitted heat flux is 1358 W/m2,
2. Radiation from the satellite to space:
 Radiation assumed to occur from all exposed surfaces,
 Deep space temperature is assumed to be 2.7K,
3. Radiation and albedo from the Earth:
 Infrared load is assumed to be 237 W/m2,
 Albedo coefficient is 30%,
 Effective area exposed to Earth radiation is 20m2,
4. Initial temperature of the satellite is 25 °C as expected upon ejection from the launcher
Cold Case Cube:
1. No radiation from the Sun
2. No albedo radiation
3. Radiation from the satellite to space:
 Radiation assumed to occur from all exposed surfaces ,
 Deep space temperature is assumed to be 2.7K.
In comparison, in the backup mission, the satellite is a prism with an hexagonal base.
Assumptions for this case are listed below:
Hot Case Hexagon:
1. Radiation From the Sun:
 Load assumed to strike on two faces, whose are under 30°,
 Effective area exposed to solar radiation is 7.80 m2,
 Solar emitted heat flux is 1358 W/m2,
2. Radiation from the satellite to space:
 Radiation assumed to occur from all exposed surfaces,
 Deep space temperature is assumed to be 2.7K,
3. Radiation and albedo from the Earth:
 Infrared load is assumed to be 237 W/m2,
 Albedo coefficient is 30%,
 Effective area exposed to Earth radiation is 7.80 m2,
4. Initial temperature of the satellite is 25 °C as expected upon ejection from the launcher
Cold case Hexagon:
1. No radiation from the Sun
2. No albedo radiation
3. Radiation from the satellite to space:
 Radiation assumed to occur from all exposed surfaces,
 Deep space temperature is assumed to be 2.7K.
The thermal subsystem has been designed considering that one face of the satellite will always be pointed towards the Earth.
This assumption is almost always true (during normal on station, orbit transfers and rendezvous), with the only exceptions
of spinning modes during final approaches and thrusting phases during BOB, when the satellite may have any attitude.
Anyway, it has been checked that if those maneuvers are accomplished in less than 2000 seconds (around 33 minutes, as
predicted by the Mission Analysis), temperatures will change only of almost 2 degrees.
As the module dimensions will be greater than satellite dimensions, it will be possible that the satellite will be shadowed for
several minutes, that is, the satellite will be in a kind of COLD CASE.
1.5 Single node thermal analysis
In first instance, the satellite is assumed to be a single spherical node, with area equal to the one of the entire satellite, that is
56 m2.
Average values for thermophysical properties are taken in account: ε = 0.8 and α = 0.3
An internal heat dissipation of 1000 W is also assumed.
Steady analysis gives the values listed below:
Altitude [km]
600
900
Tmin [°C]
-23.48
-28.14
Tmax [°C]
77.43
74.64
These temperatures approximately represent average values, but to know exactly each device’s temperatures a further
specific analysis will be mandatory.
1.6 Multi-node thermal analysis
In this model, it is necessary to consider all the specific geometrical and thermophysical characteristics of the satellite,
which is represented by a set of nodes. Incoming and outgoing heat fluxes are specified on each node. The rate of
conduction depends both on the length of the conductive path and on the choice of material. Radiation is neglected within
the satellite. The goal of the multinode analysis is to determine, with a good accuracy, the temperature mapping of any
parts. The multinode model has been implemented using MINITAN, a thermal analyzer software, distributed by ESA.
Further details will be given later.
1.7 Solar arrays
In both cases there are two solar arrays, fixed on opposite panels and two robotic arms on other two.
The solar arrays chosen are GaInP2/GaAs/Ge triple semiconductor junction, attached on aluminum panel.
Solar cells have good thermophysical properties (α = 0.92 and ε = 0.85) and high efficiency (η = 0.28).
The rear side of solar arrays is painted with a silver paint (α = 0.10 and ε = 0.85) in order to dissipate as mush as possible
towards the deep space.
In the SSM solar arrays area is around 18.85 m2, while in the S/C M it is around 15.72 m2.
In both missions, solar arrays will experience a temperature between -60 and 107 °C.
1.8 Baseline Mission
In order to meet requirements, different solutions have been adopted.
The configuration of the cubic satellite is shown on the figure below:
The whole satellite is covered with 2mm of Aclar, a thin Aluminum-based film, whose properties are: α=0.1 and ε=0.62.
The thermal control is provided with the following configuration:
1 – face towards the Earth
2 – face with main engine
5,6 – faces with solar panels
4 – face with docking port
The logical scheme for the active control of the satellite is shown in the following chart:
The battery heat (≈ 300W) is dissipated by louvers, whose emissivity is between 0.1 (cold case) and 0.9 (hot case), using a
heat-pipe. About half surface is covered with a film of MLI of Aluminum Polyamide, whose thermophysical properties are:
ε = 0.05 and α = 0.1.
In order to use louvers, it’s mandatory an active thermal control. This is provided using thermal-sensors, collocated on the
battery and changing the emissivity of the louver (up to 0.9) using a proportional control law. One louver is placed on face 5
and the other one on face 6. The surface of each louver is 0.4 m2.
Panel 2 and 3 are covered only with a thin film of Aluminum Polyamide (MLI). Two cases have been taken in account (600
Km and 900 Km). In both cases the Sun is in the orbital plane because this has been identified as the worst case. View
factors have been calculated to take in account the portion of the faces which radiates towards the deep space and which
receives radiation (infrared and albedo) from the Earth.
The mapping temperature (Hot Case and Cold Case) at 600 Km is shown below:
Node
D1
Antenna
D3
D5
Battery
D19
D21
Louver
D23
D13
CPU
D15
Star Mapper D17
Battery
D7
D9
Louver
D11
D2
Wheel
D4
D6
Hydrazine
D20
Wheel
D22
RVS
D24
D14
Wheel
D16
D18
D8
Wheel
D10
D12
T
48.86
48.39
48.65
24.66
14.96
10.05
44.31
44.31
36.12
24.64
14.92
9.96
38.00
40.22
34.72
33.52
30.66
34.61
36.56
36.55
34.70
33.51
30.62
34.07
600 km
Node
T
D1
20.50
Yes D3
14.22 Yes
D5
11.24
Yes D19
8.12 Yes
D21
1.29
Yes D23 -12.40 Yes
D13
2.90
Yes D15
3.46 Yes
Yes D17
-1.16 Yes
Yes D7
8.27 Yes
D9
1.54
Yes D11 -11.83 Yes
D2
12.71
Yes D4
11.20 Yes
D6
3.43
D20
9.77 Yes
Yes D22
9.33 Yes
Yes D24
3.33 Yes
D14
8.21
Yes D16
13.31 Yes
D18
3.34
D8
9.83
Yes D10
9.56 Yes
D12
5.93
Solar Array
Solar Array
Deep Space
D48
D49
B50
93.06 Yes
94.07 Yes
-270.44
D48
D49
B50
-35.91 Yes
-38.90 Yes
-270.44
All the components are in their respective ranges, namely batteries and RVS.
In the cold case it has been necessary to use heaters to keep hot batteries and propellant. These heaters give 140W globally.
Heaters, as louvers, need an active control thermal law which is based on measurements made by thermal-sensors.
The control law is proportional to the variation of the temperature.
In both cases, heat pipes create a path for the heat dissipated from the batteries until the louvers.
The mapping of the baseline mission at 900 Km is shown in the table below:
900 km
Node
T
Node
T
D1
31.74
D1
6.29
Antenna
D3
31.47 Yes D3
5.24 Yes
D5
31.75
D5
2.09
Battery
D19 22.24 Yes D19
5.12 Yes
D21 15.55
D21
-1.07
Louver
D23
3.47 Yes D23
-5.95 Yes
D13 41.25
D13
1.38
CPU
D15 41.37 Yes D15
-0.03 Yes
Star Mapper D17 41.76 Yes D17
-2.52 Yes
Battery
D7
22.25 Yes D7
5.18 Yes
D9
15.56
D9
-0.65
Louver
D11
3.50 Yes D11
-4.98 Yes
D2
29.28
D2
5.14
Wheel
D4
30.81 Yes D4
6.63 Yes
D6
26.61
D6
2.21
Hydrazine
D20 28.16
D20
4.49 Yes
Wheel
D22 26.93 Yes D22
4.19 Yes
RVS
D24 26.56 Yes D24
2.16 Yes
D14 31.92
D14
3.57
Wheel
D16 32.89 Yes D16
2.49 Yes
D18 26.69
D18
2.17
D8
28.16
D8
4.59
Wheel
D10 26.94 Yes D10
4.58 Yes
D12 26.72
D12
6.60
Solar Array D48 87.80 Yes D48 -60.44 Yes
Solar Array D49 88.78 Yes D49 -66.40 Yes
Deep Space B50 -270.44
B50 -270.44
The hot case for the satellite at 900 Km is treated in the same way.
In comparison, cold case has been controlled using heaters that give up to 145 W to batteries and propellant tanks.
The model for the baseline mission has been done with MINITAN using 27 nodes: this number has been chosen to have a
good accuracy in understanding heat transmission on the satellite.
1.9 Autonomous Robots
In order to accomplish assembling, two autonomous robots will be used, mounted respectively on face 3 and on face 1.
Robots are made with Aluminum and their internal power (when turned on) is around 200W that is, in the most part,
dissipated by moving parts of each robot.
The whole surface of the robots is white painted, except the boxes containing electronic devices, whose have panels
opposite to solar arrays painted of silver.
Thermal analysis shows that robots temperature is between -20°C and 25 °C.
Batteries have been insulated so that temperatures are kept over 5°C.
Thermal budget Baseline
Component
Area [m2]
MLI
11
Louvers
8*0.1
Heat Pipes
2*0.013
Heaters, Sensors and Sinks
-
Mass [Kg]
5.5
4.0
2.2
≈3
1.10 Baseline mission using electrochromic materials
In the future, the thermal control for this kind of mission should be done using Smart Radiator Devices (SRD).
Namely, SRD are thermochromic materials, thin film coatings that change their thermophysical properties in response to
external heat loads or temperature variations.
This technology has significant advantages:




Direct integration with spacecraft structures,
Minimal added mass ( their density is less than 20g/m2),
No mechanical moving parts,
Large dynamic variation in thermal conductance and emissivity.
The principal component of this kind of material is the VO2, doped in different ways to have the desired values of
emissivity.
Different types of thermochromic materials are listed below:
Material
AD-07-AL-01
MCL-02-AL-01
MCL-01-AL-01
MCL-04-AL-01
ε (Low Temperature)
0.45
0.32
0.29
0.23
ε (High Temperature)
0.72
0.58
0.54
0.48
Δε
0.27
0.26
0.25
0.25
The cubic shape has been analyzed at 600 Km.
Thermochromic materials are not enough to control the mapping temperature of the satellite; for this reason electrochromic
materials should be used. Electrochromic materials are similar to thermochromic materials but the ‘conductor to insulator’
transition is controlled applying a voltage. The promotions of electrons in the conduction band is due to the electric field
created by the applied voltage. MCL-04-AL-01 has been used. Besides it has been necessary to cover one face of the
satellite with MLI.
The mapping temperature for the case at 600 Km (Hot Case and Cold Case) is shown in the table below:
600 km
Node
T
D1
35.99
Antenna
D3
38.54
D5
41.67
Battery
D19
13.03
D21
3.59
Louver
D23
0.31
D13
37.32
CPU
D15
38.03
Star Mapper D17
41.13
Battery
D7
16.61
D9
11.26
Louver
D11
17.41
D2
27.73
Wheel
D4
31.76
D6
29.53
Hydrazine
D20
23.52
Wheel
D22
21.50
RVS
D24
29.40
D14
27.94
Wheel
D16
29.58
D18
29.52
D8
24.68
Wheel
D10
24.74
D12
32.16
Solar Array D48
92.91
Solar Array D49
94.07
Deep Space B50 -270.44
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Node
D1
D3
D5
D19
D21
D23
D13
D15
D17
D7
D9
D11
D2
D4
D6
D20
D22
D24
D14
D16
D18
D8
D10
D12
D48
D49
B50
T
12.65
7.05
4.75
7.41
-4.10
-10.72
1.73
1.49
-2.32
12.98
-1.60
-8.63
9.47
6.21
1.79
7.77
5.04
1.72
6.87
10.48
1.75
9.32
6.29
5.47
-35.79
-38.90
-270.44
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
In comparison to the conservative solution, using electrochromic materials temperatures are distributed in a different
manner within the satellite. Namely, batteries are kept at lower temperature in the hot case and at higher temperature in the
cold case. Devices operate more efficiently (namely hydrazine combustion will be better).
Using a 3μm thin film of electrochromic material, radiators,
heaters and heat pipes are not necessary. It means that the mass
budget decreases a lot.
The electrochromic materials need a proportional control law. The
electrical scheme is showed in the figure on right.
The Aluminum honeycomb structure is covered with a thin film of
electrochromic material. The external and internal interfaces are
connected to electrical cables. A thermal sensor sends
measurement at the controller which implements a proportional
control low in order to change the emissivity coefficient of the
material. The Δε allowable is 0.25.
The maximal tension required for the maximum emissivity
variation is about 3 mV.
Decreased mass budget and minimal tension required (low risk of
failure) encourage using these materials, this technology is still in
an experimental phase, so the conservative solution has been
chosen.
1.11 Backup Mission
The configuration of the backup mission is shown in the figure below:
Batteries are placed behind the panels supporting solar arrays (3 and 6).
The docking sensor is placed on the panel 9. Propellant tank is within the cylinder.
The whole of the satellite is covered with a 2mm of Aclar, a thin Aluminum-based film, whose properties are: α=0.1 and
ε=0.62.
The mapping temperature for the satellite at 600 Km is shown in the following table (the hot case is marked in red while the
cold case is marked in blue):
600 km
Node
D1
D2
T
40.37
21.30
Battery MLI
D3
29.31
MLI
MLI
Battery MLI
Propellant,Wheels
MLI
Star Mapper. RVS
D4
D5
D6
D7
D8
D9
D10
D11
D12
D13
D14
D15
D48
D49
D50
41.83
41.38
25.21
32.24
30.67
24.95
32.64
30.87
30.76
30.34
29.16
31.35
107.24
107.21
-270.44
CPU
Solar Array dx
Solar Array sx
Deep space
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Node
D1
D2
T
15.17
-0.26
D3
5.08
D4
D5
D6
D7
D8
D9
D10
D11
D12
D13
D14
D15
D48
D49
D50
2.16
2.45
7.70
7.48
4.70
3.27
5.12
5.71
7.06
6.22
5.14
5.40
-50.37
-50.28
-270.44
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Batteries and RVS are kept in their operative ranges using an extra Aluminum Polyamide Coating (MLI) layer, whose
thermophysical properties are: ε = 0.05 and α = 0.1 .
1– face towards the Earth
3,6 – faces with solar panels
8 – face with engine
9 – face with docking port
As it can be seen from the figures above, three sides of the hexagon are covered with MLI, in order to prevent batteries and
RVS from low temperatures.
It has been necessary to cover with MLI also the panel 8 of the hexagon.
No active devices have been used to reduce the mass dedicated to the thermal subsystem, except heaters giving up 140 W to
prevent low temperatures on batteries, RVS and propellant.
The model has been implemented with MINITAN using 18 nodes.
The case at 900 Km is shown in the table below:
900 km
Node
T
D1
36.70
D2
23.05
Battery MLI
D3
29.18
MLI
D4
41.76
MLI
D5
41.75
Battery MLI
D6
29.04
Propellant,Wheels D7
32.20
MLI
D8
30.84
Star Mapper. RVS D9
24.49
D10 32.60
D11 31.25
CPU
D12 30.71
D13 30.07
D14 29.27
D15 31.26
Solar array dx
D48 101.05
Solar Array sx
D49 101.04
Deep space
D50 -270.44
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Node
T
D1
18.77
D2
1.28
D3
5.44
D4
2.12
D5
2.10
D6
5.26
D7
7.52
D8
4.96
D9
2.87
D10
5.07
D11
5.40
D12
7.18
D13
6.76
D14
5.34
D15
5.43
D48 -56.65
D49 -56.66
D50 -270.44
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
Yes
This case is a little bit colder than the case at 600 Km, but devices temperatures are within their range.
Thermal budget Backup
Component
Area [m2]
MLI
12.66
Heat Pipes
7*0.013
Heaters,Sensors and Sinks
-
Mass [kg]
6.33
3.8
≈3
1.12 Trade off: Baseline vs. Backup
Thermal control subsystem operates efficiently in both cases keeping devices in their temperature ranges. MLI and thin
Aluminum films are used in both cases in order to prevent from cold. The mass budget is greater in the baseline mission (15
Kg in the baseline against 13 Kg in the backup). Active thermal control, used in the baseline mission, is more subjected to
failure risk because of both mechanical parts and electronic devices; its behavior depends also on other subsystems (power
subsystem, TT&DH subsystem...). Even if more expensive, active thermal system controls more efficiently temperature
ranges.
In the backup mission, there are only heaters in order to control the temperature of the batteries; heaters are less subjected to
failure because of the lower applied tensions.
1.13 Transient after launch phase
Following requirements about the launch, it can be found out that the satellite will be put on orbit with an initial temperature
of 25 °C. The transient has been calculated as shown in the figure below. Eclipse time is around 35 min for both orbits (600
Km and 900 Km). At t = 60 min the satellite enters the shadow and at t = 95 min the satellite exits the shadow.
Some nodes transient, namely batteries, hydrazine tank and RVS, are plotted in the figure:
Temperatures are always included between 17°C and 26°C. It means that hot case and cold case will not be reached while
satellite is waiting module. Heaters are not necessary during this phase.
Moreover, they will be kept onboard, because during docking manoeuvres the satellite could be shadowed for an amount of
time necessary to reach the cold case. So power is used by heaters only in this emergency case.
1.14 Thermal Budget
Temperature Sensors:
Dallas Semiconductor DS18B20 – Programmable Resolution 1-Wire Digital Thermometer
Mass: 0.15g
Voltage: 3.0V to 5.5V
Resolution: 9 to 12 bits
Current: 1.5 mA
Location: Batteries, RVS, Tanks, Optic Gyroscope, CPU, Star Mapper.
Purpose: To measure the temperature of each panel and send information to OBC&DH to make sure the temperatures are
within their operating range.
Heaters:
Minco Products, Inc. – Strip heater with Kapton Insulation
HK5365 R140 L12A
Mass: 3g
Resistance: 240 Ohms
Power: 2.4W to 2.6W
Location: Four heaters will be located on each battery. Heaters will be almost 60. In the baseline configuration, heaters are
used also for the RVS.
Purpose: To heat the batteries and RVS to within its required temperature limits.
Heaters
MLI:
Kapton tape
Conductivity: k=0.12 W/(m*°K)
Location: Around batteries pack, internal RVS and on external surfaces. This includes the sensors placed on the batteries.
Size: Insulation must be wrapped around component for a thickness of 0.5mm.
Purpose: To prevent heat loss to surrounding components.
MLI
Epoxy Filler:
Conductivity: h=15000 W/m2.
Location: this thin film is placed on all the interfaces between the devices and the structure.
Purpose: to prevent heat losses due to contact resistance. Calculating contact resistance is really hard and so filler have to be
used to reduce as much as possible heat losses in conductions paths.
Filler
Louvers:
Louvers provide thermal control by varying the emittance (from 0.1 to 0.9) of the underlying surface by varying the angle of
the low emittance blades that comprise our structure.
Louver
Heat pipes:
Heat pipes have different dimensions and use different fluids in order to transfer different quantity of heat from a
component to another one.
Heat Pipe
References
The Geostationary Applications Satellite (Peter Berlin) [ed. Cambridge University Press]
Elementi di trasmissione del calore (G.Guglielmini, C.Pisoni)
[ed. Masson]
Chimica – centralità di una scienza (T.L.Brown , H.E.LeMay,jr.) [ed. Zanichelli]
Spacecraft systems engineerin
MINITAN, “User manual”
Integrated Thin-Film Smart Coatings with Dynamically-Tunable Thermo-Optical Characteristics (Roman V. Kruzelecky
and Emile Haddad - MPB Communications Inc.) (Mohamed Soltani and Mohamed Chaker - INRS Énergie et Matériaux)
(Darius Nikanpour - Canadian Space Agency)
EMN_2002, chapter 03, “Environment”, ESA
EMN_2002, chapter 06, “Materials”, ESA
Smart Radiator Offering Autonomy in Satellite Thermal Control (Darius Nikanpour, Canadian Space Agency)
Elementi di controllo termico di satelliti (Davide Grassi)
Multilayer Insulation Material Guidelines
[M.M. Finckenor(Marshall Space Flight Center, Marshall Space Flight Center,
Alabama) e D. Dooling (D2 Associates, Huntsville, Alabama)
David Trujillo, “A general thermal contact resistance finite element”, march 1999, Boeing Satellite System
Fard and Rahumyan, “Design and Analysis of a Micro-Satellite Passive Thermal”, University of Teheran
Finckenor, “Multilayer Insulation Materials Guidelines”, Marshall Space Flight Center, April 1999
Haddad, Kruzelecky, Jamroz, “Dynamically Variable Thin-Film Smart Radiator Device”, 4th Round Table on Micro-Nano
Technologies for space
Klett, Conway, “Thermal management solutions utilizing high thermal conductivity graphite foams”, Carbon and Insulation
Materials Technology Group
Mark Fischer, “Thermal Control’, February 1995
NASA SP8105, “Spacecraft Thermal Control”, May 1973
Painter, “Finite element radiation transport in one dimension”, Lawrence Livermore National Laboratory
Web references
www.tpg.unige.it/skills
http://esapub.esrin.esa.it/publicat/publicat.htm
http://earth.esa.int/ers/
http://tycho.dm.unipi.it/~nobili/ggweb/phaseA/chapter4.html
http://linus.tesre.bo.cnr.it/Research/SAX/Mission/subsy_ttc.html
http://www.isis-innovation.com/licensing/further-info.html
http://www.etsi.org/technicalactiv/srd/srd.htm
http://www.nanothermel.org/
www.cordis.lu/nanotechnology/home.html
http://www.dlr.de
http://www.eng-tips.com/gthreadminder.cfm/lev2/15/lev3/57/pid/332
http://www.journals.tandf.co.uk/tf/
http://dau.ing.univaq.it/omhat/
http://www.hsssi.com/Applications/SpaceHabitat/
www.swales.com
www.minco.com
http://www.qrg.nwu.edu/projects/vss/docs/thermal/zoom-why.html
http://www.tak2000.com/data/planets/earth.htm
Special thanks
Thanks to Ing. Marco Molina (Gavazzi Space) for his advice.