Mastersat B: Mission and Analysis Design - Sapienza

Transcript

G. Baldesi, A. Califano, M. Di Marco, E. Di Litta,
G. Guarino, S. Mezzasoma, G. Orlando
Mastersat B: Mission and Analysis Design
Relazione di progetto
Master in Satelliti e Piattaforme Orbitanti
Anno Accademico 2002/2003
Tutors
Ing. Giorgio Perrotta
Ing. Guido Morelli
Il Direttore del Master
Prof. Paolo Gaudenzi
i
Mastersat B – Mission and Analysis Design
Study team
1. This report is the result of the work carried out by the following team:
Section
Team member(s)
Preliminary design
Andrea Califano
Mission objective
Marco Di Marco & Andrea Califano
Payload
Marco Di Marco & Giuseppe Orlando
Mission analysis
Andrea Califano
Design & Configuration
Gianluigi Baldesi & Elisa Di Litta
Propulsion
Andrea Califano
Thermal control
Giovanni Guarino
EPS
Silvia Mezzasoma
AOCS
Gianluigi Baldesi
DH, OBC, TT&C
Gianluigi Baldesi, Marco Di Marco & Giuseppe Orlando
Structure
Elisa Di Litta
Cost analysis, WBS
Andrea Califano
Team leader:
Andrea Califano
Università degli Studi di Roma “La Sapienza” - Master in Satelliti e piattaforme orbitanti
ii
Table of Contents
STUDY TEAM
I
TABLE OF CONTENTS
II
CHAPTER 1: INTRODUCTION
1
1.1
1
MISSION AND ANALYSIS DESIGN PROCESS
CHAPTER 2: MISSION OBJECTIVES
3
2.1
2.2
3
3
INTRODUCTION
SYSTEM CHARACTERISTICS
CHAPTER 3: PAYLOAD DEFINITION
5
3.1 PAYLOAD REQUIREMENTS AND DEFINITION
3.1.1 PAYLOAD ARCHITECTURE
3.1.2 ANTENNA DESIGN
3.1.3 LINK BUDGET FOR KU MISSION
3.1.4 LINK BUDGET FOR KA MISSION
3.1.5 MASS AND POWER BUDGETS
3.1.6 MORE OPTIONS
5
5
8
10
14
15
16
CHAPTER 4: PRELIMINARY DESIGN
17
4.1
4.2
17
18
INTRODUCTORY SPACECRAFT BUDGETS
MAIN SPACECRAFT PARAMETERS
CHAPTER 5: MISSION ANALYSIS
20
5.1
5.2
20
21
LAUNCHER
∆V BUDGET
CHAPTER 6: CONFIGURATION
24
6.1 REQUIREMENTS AND CONSTRAINTS
6.2 SPACECRAFT DESCRIPTION
6.2.1 BASELINE
6.2.2 LAUNCHER FAIRING STORAGE
6.3 OPTIONAL CONFIGURATION
24
24
24
27
28
CHAPTER 7: PROPULSION SUBSYSTEM
29
iii
7.1 DESIGN PROCESS
7.1.1 SPACECRAFT PROPULSION FUNCTIONS
7.1.2 ∆V BUDGET
7.1.3 TOTAL IMPULSE , THRUST LEVELS, DUTY CYCLES AND MISSION LIFE REQUIREMENTS
7.1.4 PROPULSION SYSTEM OPTIONS
7.1.5 ESTIMATE KEY PARAMETERS FOR EACH OPTION
7.1.6 ESTIMATE TOTAL MASS FOR EACH OPTION
7.2 OPTION 2
29
29
29
30
30
31
32
32
CHAPTER 8: THERMAL CONTROL SYSTEM
34
8.1
8.2
34
37
KU MISSION
KA MISSION
CHAPTER 9: POWER SUBSYSTEM
39
9.1 REQUIREMENTS
9.2 BASELINE DESIGN
9.3 DESIGN
9.3.1 SUBSYSTEM CONFIGURATION
9.3.2 SOLAR ARRAYS
9.3.3 DESIGN OF SOLAR ARRAYS ELECTRICAL NET
9.3.4 BATTERIES DESIGN
9.3.5 BATTERIES CHARGE
9.3.6 CHARGE POWER
9.3.7 MULTI-JUNCTION CELLS
39
39
40
40
40
41
43
44
44
45
CHAPTER 10: ATTITUDE ORBIT CONTROL SUBSYSTEM
46
10.1 ORBIT AND DESIGN DEFINITION
10.2 CONTROL MODES & REQUIREMENTS
10.3 DISTURBANCE TORQUE COMPUTATION
10.3.1 ENVIRONMENTAL
10.3.2 INTERNAL
10.3.3 RESULTS IN WORST-CASE
10.3.4 RESULTS BY ADS
10.4 ACTUATORS TRADE-OFF
10.4.1 PASSIVE STABILIZATION
10.4.2 ACTIVE STABILIZATION
10.4.3 RESULTS
10.5 SENSORS SELECTION
10.6 CONTROL MODE ARCHITECTURE
46
47
47
47
48
49
49
50
50
51
53
57
60
CHAPTER 11: TT&C, DH AND OBC
61
11.1 TT&C SUBSYSTEM
11.2 SPACECRAFT INTEGRATED CONTROL SUBSYSTEM
11.2.1 MASS BUDGET
61
63
65
CHAPTER 12: STRUCTURE
66
iv
12.1 INTRODUCTION
12.2 STRUCTURE DESCRIPTION (BASELINE)
12.3 SIMPLIFICATIONS AND ASSUMPTIONS
12.4 SOLUTIONS FOR THE SATELLITE
12.4.1 KU MISSION
12.4.2 KA MISSION
12.5 SUMMARY
66
66
66
67
67
69
70
CHAPTER 13: SYSTEM BUDGETS
71
13.1
13.2
71
72
MASS BUDGET
POWER BUDGET
CHAPTER 14: COST ANALYSIS
74
14.1
14.2
74
76
ELEMENTS OF ANALYSIS
COST ESTIMATE
CHAPTER 15: PLANNING
80
CONCLUDING REMARKS
81
REFERENCES
82
1
Chapter 1: Introduction
1.1 Mission and analysis design process
A practical approach to the space mission analysis and design process is summarized as follows:
Table 1.1 : The Space Mission Analysis and Design Process
Analysis and design are iterative, gradually refining both the requirements and methods of
achieving them. Successive iterations will usually lead to a more detailed, better-defined space
mission concept.
Once we have established alternative mission concepts, architectures, and system drivers, we
must further define the mission concepts in enough detail to allow meaningful evaluations of
effectiveness. This is done through an iterative process, whose steps are summarized in Table 1.2.
A
B
C
D
E
F
G
H
I
J
Define the preliminary mission concept
Define the subject characteristics
Determine the orbit or constellation characteristics
Determine payload size and performance
Select the mission operations approach
Design the spacecraft bus to meet payload, orbit and communications requirements
Select a launch and orbit transfer system
Determine deployment, logistics, and end-of-life strategies
Provide costing support
Document and iterate
Table 1.2 : Steps for the Concept Characterization Process
2
The previous table is generally followed not only from top to bottom, but also through several
interactions between the steps.
This is what happens automatically in the Concurrent Design Facility (CDF) of the Tech site of
the European Space Agency. Every change made by a subsystem would affect the others, in a
continuous exchange of information through the local net.
In the same manner, but in an old fashion man-driven iterative process, every member of the
team has worked giving his outputs as inputs to the other members, and vice versa while handling
external data, as the interactions in Fig. 1.1 explicitly point out.
Fig. 1.1: Characterization of the Mission Architecture Process
3
Chapter 2: Mission objectives
2.1 Introduction
The MASTERSAT mission aims to develop a communication satellite. It offers a good
opportunity to provide North America and Europe by broadcasting wide-band services such as
video, telephony and data.
The overall capacity consists of 2 Mbps data streams that can be either added together or divided
into smaller bands to fit user requirements. A broadcasting video service needs between 2 and 6
Mbps uncoded channels, whereas a high quality stereo quadraphonic channel requires about 256
kbps so that a good compromise is splitting the 2 Mbps into 8 streams of 256 kbps and each substream can be divided, on its turn, in 4 channels of 64 kbps.
We intend to develop a flexible and dynamic communication system, capable to vary the
parameter links according to the type of service and quality constraints. This system would allow
the service provider to define different type of users in terms of reliability and availability, and to
offer more performing channels at higher price. The objective is to obtain as much profit as possible
from these links, while providing remaining capacity to less demanding users. In fact, to prevent
more profitable links from outages, a possible choice could be either to put into operation additional
coding techniques, or to improve the link budget with damage to the other users.
Although many solutions may be found, our goal is to define one leading to, expectedly, good
cost/performance ratios.
For this purpose, two different technical solutions have been investigated: conventional
geostationary satellite with a capacity of more than 500 × 2 Mps channels in Ku band, and a second
one more challenging, able to offer more than 4000 × 2 Mbps channels in the Ka band.
The first solution takes advantage of reliable, available and low-cost technology. Yet, this solution
does not offer a good flexible system as required. Besides, cost analysis estimate shows that the cost
on a per channel/hour basis turns out to be not much competitive on the market.
The challenging solution fulfils technical requirements and, thanks to large number of channels,
it may pay off the higher cost, reaching the breakeven point after a few years. The drawback is the
difficulty to exploit the overall capacity resulting in a low fill factor. In addition, the Ku band is
much more weather-dependant with the consequence of many system outages.
2.2 System characteristics
In order to cover both continents, a geostationary orbit at the longitude of 30° West is a
reasonable choice. For both solutions, the antennas consist of multi-beams which fit the irregular
areas. However, the Ka solution provides smaller spot-beams that are supposed to handle a set of
channels each, improving the system capacity. On the other hand, the smaller spot-beam results in a
different pointing precision for the attitude control, as later described throughout this report.
Due to the technology innovation to be implemented in the second solution, the expected
operative life (> 10 years) is smaller than in the first solution (> 12 years), characterized by a
consolidated heritage.
4
In this preliminary study, we will consider a launch from the European base of Kourou, placing
the satellite in the lower part of the fairing of Ariane V.
Below, in Table 2.1, are summarized the system characteristics of the two possible solutions
described in this section, which will then be used as starting point of this preliminary study, and
represent at the same time the aimed target.
Orbit
Down-link frequency
Up-link frequency
Capacity
System Outages
Spacecraft Mass
Spacecraft Power
S/C antennas
Attitude Control
Spacecraft lifetime
Launcher
Option 1: Ku-band
GEO @ 30° West
12 GHz
14 GHz
> 500 × 2 Mbps channels
< 0.01 %
< 2000 kg
≅ 2000 W
Multi-beams
≤ 0.3 °
>12 years
Ariane 5
Option 2: Ka-band
GEO @ 30° West
20 GHz
30 GHz
> 4500 × 2 Mbps channels
< 0.1 %
< 3000 kg
≅ 2000 W
Spot-beams
≤ 0.05 °
> 10 years
Ariane 5
Table 2.1 : System characteristics.
5
Chapter 3: Payload Definition
3.1 Payload Requirements and definition
3.1.1 Payload Architecture
Both payloads consist of transparent transponders which receive incident signals from either
Europe or North America and then broadcast it to the area of concern.
As shown in Figure 3.1, the Ku solution has a band-width of 500 MHz in up-link and the same
amount in down-link; each band is divided into 6 smaller bands of 72 MHz accounting for the band
guards.
Up-link 14 GHz
72 MHz
VP
500 MHz
Fig 3.1 : Up–link signal band-width.
The single transponder capacity is therefore 72 MHz and it comprises 24×2 Mbps channels. In
order to improve the system capacity, frequency reuse has been adopted by making use of
orthogonal polarizations and space diversity between the two continents. Considering the frequency
reuse coefficient equals to 4 (2 continents × 2 polarizations) the satellite repeater can
simultaneously guarantee 4×24×6 = 576 × 2 Mbps channels.
The block diagram in Figure 3.2 displays a part of the receiver architecture for the conventional
communication satellite.
VP
1
LNA
Est beams
IFA
2
HP
LO
I
M
U
X
3
4
5
6
Fig. 3.2 : Receiver architecture
6
Next to the antenna, a polarization discriminator recovers the desired signal. Afterwards, the
LNA (Low Noise Amplifier) amplifies such weak signal which is then multiplied by the LO tone
(Local Oscillator) in the mixer, and converted to the intermediate frequency, that is the down-link
frequency. The IMUX (Input MUltipleXer), which is equivalent to 6 BPF, filters the 72 MHz
channels, each amplified by the following IFA (Intermediate Frequency Amplifier).
How the transmitting section works is illustrated in Figure 3.3. A variable gain amplifier
(CAMP, Controlled AMPlifier), optimizes the signal level for the next amplifying stages. A
linearizator block adjusts the signal to limit intermodulation products in TWTA while maintaining a
good efficiency. Finally, the 6 transponder signals are added together and filtered by an OMUX
(Output MUltipleXer) and sent to the antenna where the resultant signal is split into 6 elementary
beams.
VP
Lin
BPF
1
TWTA
Est Beams
2
3
4
O
M
U
X
HP
5
6
Fig 3.3 : Transmitter architecture
The second solution has a larger band-with of 800 MHz in both up-link and down-link, divided
into 6 smaller sub-bands, the same as in the previous solution. In this case the transponder capacity
is 120 MHz and provides 48×2 Mbps channels. The frequency reuse exploits other than the
techniques used for the first solution, space diversity inside the same area; of course, adjacent spots
don’t use the same sub-bands in order not to interfere each other. In fact, the spot-beam is designed
to handle only two of the 6 available sub-bands. Figure 3.4 points out band allocation.
120 MHz
1
VP
Up-link 30 GHz
2
3
4
5
6
800 MHz
Fig 3.4 : Up–link signal band-width
The receiver section associated with a spot-beam is shown in Figure 3.5:
7
Est Beam
LNA
1,4
Band Pass
Filter
IFA
Band Pass
Filter
IFA
LO
VP
HP
Fig. 3.5 : Receiver chain.
A frequency synthesizer supplies 6 frequencies of reference to the mixers so that the
intermediate frequency is set to a low and constant value.
By assuming 12 spot-beams to cover each continent, the frequency reuse coefficient R can be
calculated as follows:
R=
N ×S
M
where N is the number of sub-bands per spot-beam, S the number of spots in the service area and
M the number of available sub-bands; thus R = 2 × 12 / 6 = 4 . If we consider space diversity between
the two continents and polarization discrimination as well, the frequency reuse becomes 16. Since
the communication system manages 6 120 MHz transponders, the overall capacity is therefore 16 ×
6 ×48 2 Mbps channels, that are more than 4500 as required.
The flexibility of the system is performed by a controlled switch matrix, which addresses the
incoming 120 MHz signal to the transmitting section.
1
1
2
2
VP
VP
24
1
2
24
RF
Matrix
Switch
1
2
HP
HP
24
24
TT&C
Control
Fig. 3.6: RF Matrix Switch
8
Figure 3.7 gives an idea about the output section.
Note that a single TWTA is used to transmit the two transponder signals. It means that the Ka
communication payload, without redundancy, consists at least of 48 TWTA (12 beams × 2
polarizations × 2 continents).
2
IFA
VP
LO
Lin
CAMP
Est Beam
2,5
TWTA
HP
5
IFA
Fig. 3.7 : Transmitter chain.
3.1.2 Antenna design
In the antenna design, multi-beams architecture is the logical choice, due to the highly irregular
coverage. The paraboloidal antenna diameter required for a given beam-width (θ3dB) is computed by
the following approximate formula:
D=
h⋅λ
θ3dB
(3.1)
where h is a constant correlated to the type of illumination; it varies between 50 and 75 if θ3dB is
expressed as degrees. For the Ku mission, 6 elementary beams with θ3dB of 1.8 ° each are a good
compromise as shown in Figure 3.8. By assuming h=70 and the down-link wave-length, D = 97 cm.
The Ka solution provides 12 individual, narrow spot-beams per antenna reducing the 3dB beamwidth to 0.9°. The Earth coverage is displayed in Figure 3.9.
Fig. 3.8 : Earth coverage with elementary beams.
9
According to (3.1), the antenna diameter results D = 1 m, that is approximately the same as in the
Ku payload. However, the smaller beam-width, the more attitude control is needed. As rule of
thumb, minimum spot size is 10 times the p-p pointing error; for the spot-beams configuration the
pointing accuracy is very relevant and is about 0,05° along the three axes. In the spot-beams of the
Ku band instead, overlapping beams don’t interfere and a small gain shift is nearly irrelevant.
Therefore, an attitude control of 0,3° is enough adequate for the conventional satellite.
Fig. 3.9 : Earth coverage by spot-beams.
The nominal payload carries four antennas, a pair for transmitting (one for each continent) and a
pair for receiving. Besides, each antenna has dual polarization feed system and a gridded subreflector to exploit the polarization discrimination. In order to avoid electromagnetic interference,
for the both missions, 2 fixed RX antennas are mounted on Earth face, whereas TX antennas are
deployed along East-West panels.
The antenna gains are related to horizontal and vertical beam-width (θ3dBH, θ3dBV) as follows:
G=
K
θ 3dBH ⋅ θ 3dBV
where the constant K depends on antenna global efficiency. This value, typically in the range
[23000 ÷ 35000], is set to 28000 to take a bit of margin in the design.
Table 3.2 summarizes both transmitting and receiving antennas performance.
The ground antennas require at least 50 dB of gain. For a paraboloidal reflector uniformly
illuminated, the directive gain at the centre of the spot beam is given by:
Gideal
π ⋅ D 
=

 λ 
2
In the above relationship an efficient coefficient η has to be considered to take notice of nonuniform illumination, spill over and any kind of losses; thus the actual gain is G = ηGideal. The main
ground antenna features are listed in Table 3.3. Antenna diameters of 4 m and 2.4 m for Ku and Ka
band respectively, perform more than 50 dB of gain as required.
10
Ku Band
SatelliteAntennaTX
beam width
1,8
θx
θy
1,8
Gain
8641,98
39,37
K
28000
Diameter 0,97
h
70
[degree]
[degree]
[lin]
[dB]
[m]
SatelliteAntennaRX
beam width
1,8
θx
θy
1,8
Gain
8641,98
39,37
K
28000
Diameter 0,83
h
70
Tsys
400
[degree]
[degree]
[lin]
[dB]
[m]
[K]
Ka Band
SatelliteAntennaTX
beam-width
θx
0,9
θy
0,9
Gain
34567,90
45,39
K
28000
Diameter 1,00
h
70
[degree]
[degree]
[lin]
[dB]
[m]
SatelliteAntennaRX
beam-width
θx
0,9
θy
0,9
Gain
34567,90
45,39
K
28000
Diameter 0,67
h
70
Tsys
400
[degree]
[degree]
[lin]
[dB]
[m]
[K]
Table 3.2 : Satellite Antennas performance.
Ku Band
GroundAntennaTX
Gain
172188,60
52,36
Diameter
4
efficiency
0,5
[dB]
[m]
GroundAntennaRX
Gain
126505,91
51,02
[dB]
Diameter 4
[m]
efficiency 0,5
Tsys
300
[K]
Ka Band
GroundAntennaTX
Gain
284638,3
54,54
Diameter
2,4
efficiency
0,5
[dB]
[m]
GroundAntennaRX
Gain
126505,91
51,02
[dB]
Diameter 2,4
[m]
efficiency 0,5
Tsys
300
[K]
Table 3.3 : Ground Antennas performance.
3.1.3 Link Budget for Ku mission
The Ku communication payload is composed of 24 TWTA (6 area-beams × 2 continents × 2
polarizations); each of them is associated with a 72 MHz transponder and has the nominal output
power of 40 W.
In the ground station, the transmitting section is chosen to deliver an output power of 100 W per
72 MHz band.
11
It remains to compute the link budget and evaluate the Signal-to-Noise Ratio SNR or C/N0, by
the following relationship:
SNR =
C
Pr
EIRP ⋅ Gr
=
=
N 0 K ⋅ TSYS ⋅ B TSYS ⋅ L p ⋅ K ⋅ B
The parameters involved in the equation are explained below:
•
•
•
EIRP: a transmitter with output power Pt associated with an antenna of gain Gt can be
replaced, for the purpose of this calculation, by an isotropic radiator with output power PtGt.
This quantity, known as the Equivalent Isotropic Radiated Power (EIRP), characterizes the
transmitter. In the real case, additional losses have to be considered in the budget; these are
mainly end of coverage (e.o.c) losses, beam forming loss and steering losses.
Gr
: it corresponds to the ratio between the receiver antenna gain and the system noise
TSYS
temperature and gives a direct feeling of the technology implied in the receiver.
Lp : it stands for path loss and includes any kind of attenuation along the path. The main
contribution is due to the distance and it depends on operative frequency of the link by:
L p , freespace
 4π ⋅ d 
=

 λ 
2
the term Lp includes also additional atmospheric attenuation losses due to the ionosphere,
troposphere and hydrosphere. The latter contribute must be worked out through statistic
analysis. Table 3.4 shows typical rain attenuation Lr occurred over one year for the Ku band,
according to the Crane model.
% Time
Lr,up-link(dB)
12.3
6.9
5.5
0.01
0.05
0.1
Lr,down-link(dB)
10.8
5.3
3.8
Table 3.4 : Rain Attenuation vs outage percentage
The down-link budget is typically more critical than the up-link; as a result, the outage is
often assumed greater for the first link design.
If we set Pup = 0,0005 and Pdown = 0,001 from the table above (% Time can be espressed as
probability P by dividing it by 100), the system outage probability due to the large amount
of attenuation in one-way or both links is given then by:
(
)
Poutage ≅ 1 − 1 − ( Pup + Pdown − Pup ⋅ Pdown ) (1 − Psat ) ⋅ (1 − Pground ) 2
= 1 − (1 − 0, 0014995 ) ⋅ ( 0, 9999 ) ⋅ ( 0, 9999 ) ≈ 0.0018
2
(3.2)
where Psat= 0.00005 and Pground =0.00001 are the availability of the satellite and the ground
station (see Figure 3.10). In this case, Poutage corresponds to a possible outage of about
sixteen hours a year (0,18 % of the time).
12
Psat
Pdown
Pground
Pup
Pground
Fig. 3.10 : Probability of outage
•
B: it is the equivalent band occupied by the elementary channel. The larger B, the noisier is
the receiver chain. In fact, the input antenna referred equivalent noise contribution N0 is
given by KTSYS B, where K is the Boltzmann’s constant.
Table 3.5 shows the link parameters, including code gain which improves the signal-to-noise
ratio. The reference channel, as defined, is a 2 Mbps data stream. As a consequence, the output
power associated with the elementary channel is the power per transponder indicated in Table 3.5,
over 24 channels and 6 beams, that is 0,28 W.
UPLINK
Power per trasponder 100
20
Frequency
1,4E+10
Distance
38000000
Path Loss Lp
2,01E-21
-206,97
Bit rate Rp
2
Band-width B
2
KTB
1,10E-14
-139,57
Code Gain
4
Rain Attenuation
-6,9
e.o.c. losses
-3
pointing losses
-1
beam former loss
-0,8
DOWNLINK
40
16,02
1,2E+10
38000000
2,74E-21
-205,63
2
2
8,28E-15
-140,82
4
-3,8
-3
-1
-0,8
[W]
[dBW]
[Hz]
[m]
[lin]
[dB]
[Mbit/sec]
[MHz]
[W]
[dBW]
[dB]
[dB]
[dB]
[dB]
[dB]
Table 3.5 : Ku mission Link parameters.
The link-budget
C
in both links is illustrated in Table 3.6.
N0
13
Pt
Gt
Gr
EIRP
(Gr/Tsys)
Path Loss
Losses
KTB
Code gain
Received Signal
(C/No)
UPLINK
-2,08
52,36
39,37
50,28
13,35
-206,97
-11,7
-139,57
4
-125,02
28,48
14,54
DOWNLINK
-6,06
39,37
51,02
33,30
26,25
-205,63
-8,6
-140,82
4
-125,90
31,01
14,91
[dBW]
[dB]
[dB]
[dBW]
[dB K]
[dB]
[dB]
[dBW]
[dB]
[dB]
[lin]
[dB]
Table 3.6 : Link Budget.
The overall
C
can be evaluated as the harmonic average of the two links. Thus:
N0
−1
−1
−1
 C 
 C 
 C 
=
+




 N 0 overall  N 0 up −link  N 0  down −link
For an uncoded QPSK system, the probability of error, known as Bit Error Rate (BER), is:
 2 Eb
BER = Q
 N0




where Q is the complementary function of the monodimensional Gaussian probability
distribution and Eb is the energy of bit. By assuming BER ≤ 10-5, it implies from eq. (3.3) that Eb/N0
≥9.5 dB. In addition, from the following formula:
 Eb   B   C 

 =   ⋅ 

N
R
N
0
b
0
 overall
   

(3.3)
C/N0 ≥9.5 dB if the channel band-with B is equal to the bit-rate Rb.
The Margin in Table 3.7 refers to the worst case, that is, rain attenuation in both links. Indeed,
the system outage will be much lower than the percentage obtained by the eq. 3.2. By assuming a
possible outage only when both links are attenuated, the target of 0,01% listed in Table 3.1 is by far
fulfilled (Poutage< Pup⋅ Pdown = 0.00005 ≡ 0.005%).
Overall (C/N)
Overall (Eb/No)
Required
Margin
14,84
11,72
14,84
11,72
9,5
[lin]
[dB]
[lin]
[dB]
[dB]
2,2
Table 3.7 : Overall SNR
14
3.1.4 Link Budget for Ka mission
Table 3.8 and Table 3.9 recapitulates the main parameters link involved in the Ka mission,
including also the rain attenuations obtained with Pup = 0,0005 and Pdown = 0,001, as proposed in the
Ku solution; note how the attenuation increases with higher frequencies. To estimate the overall
(C/N0), climate circumstances are to be taken into account, as they deeply affect the link-budget.
First of all, rain in both links, implies likely system outage, even though the probability of
occurrence is practically negligible (Pup⋅Pdown= 0,00005 ≡ 0,005% << 0,1% as required). Table 3.10
examines the other more probable three cases; by assuming a system outage with rain in either
links, we get a little margin, proving that Poutage< Pup+ Pdown - Pup⋅ Pdown ≅ 0,0015 ≡ 0,15%, yet more
than the target of 0.1%. On the other hand, the up-link margin of 1,5 dB isn’t as critical as it seems.
In fact, it can be solved without increasing the satellite performance.
It shows the technical feasibility of the Ka mission, which may fulfill reliability and availability
requirements all over the lifecycle.
UPLINK
Power per trasponder 40
16,02
Frequency
3,0E+10
Distance
38000000
Path Loss
4,38E-22
-213,59
Bit rate/Channel
2
Bandwidth/Channel
2
KTB
1,10E-14
-139,57
Code Gain
4
Rain Attenuation
-9
e.o.c. losses
-3
steering losses
-1
beam former loss
-0,8
DOWNLINK
20
13,01
2,0E+10
38000000
9,85E-22
-210,06
2
2
8,28E-15
-140,82
4
-7
-3
-1
-0,8
[W]
[dBW]
[Hz]
[m]
[lin]
[dB]
[Mbit/sec]
[MHz]
[W]
[dBW]
[dB]
[dB]
[dB]
[dB]
[dB]
Table 3.8 : Ka mission Link parameters.
Pt
Gt
Gr
EIRP
(G/Tsys)
Path Loss
Losses w/o rain
KTB
Code gain
UPLINK DOWNLINK
-4,30
-7,31 [dBW]
54,54
45,39 [dB]
45,39
51,02 [dB]
50,24
38,07 [dBW]
19,37
26,25 [dB K]
-213,59
-210,06 [dB]
-4,8
-4,8 [dB]
-139,57
-140,82 [dBW]
4
4 [dB]
Table 3.9 : Link Budget.
15
Nice Weather
up-link
down-link
Received Signal -118,76 -121,769 [dB]
(C/N)
120,50 80,33406 [lin]
20,81
19,049 [dB]
Overall (Eb/No)
48,20 [lin]
16,83 [dB]
Required
9,5 [dB]
Margin
7,33 [dB]
up-link rain
up-link
-127,76
15,17
11,81
12,76
11,06
9,5
down-link rain
down-link
up-link
down-link
-121,77 [dB]
-118,76 -128,77 [dB]
80,33 [lin]
120,50
16,03 [lin]
19,05 [dB]
20,81
12,05 [dB]
[lin]
14,15 [lin]
[dB]
11,51 [dB]
[dB]
9,5 [dB]
1,56 [dB]
2,01 [dB]
Table 3.10 : Overall SNR.
3.1.5 Mass and Power budgets
The main payload characteristics in terms of mass and power are shown in Table 3.11 and Table
3.12:
Ku mission
Antenna
LNA
Synthesizer
Imux
Omux
IFAs blocks
TWTA blocks
RF harness
DC harness
miscellaneous
Unit
4
4
1
4
4
4
4
1
1
1
TOT
Mass [kg]
8
0,8
2
1,8
2,2
1,8
18
6
12
6
R. factor
1
2
1
1
1
1,33
1,33
1,5
1,5
1,5
Tot [kg]
32
6,4
2
7,2
8,8
9,6
96
9
18
9
Mass
Power/unit [W]
0
1
5
0
0
4,5
480
0
0
40
Tot. Power [W] remarks
0
Include dual beam former
4
wideband
5
single frequency redundant
0
six filters each unit
0
six filters each unit
18
six IFAs each unit
six TWTAs @ 3kg each & η=0,5
1920
0
rough estimate
0
rough estimate
40
hardware
198 Kg
DC Power
1987 W
Table 3.11 : Mass and Power budget for the Ku mission
The full redundancy is applied to LNA block, which is considered one of the most critical
elements for the mission, as a failure reduces the capacity by a factor 4, limiting most of the links.
The TWTAs are redounded 8+4 (two cold TWTAs every 6 TWTAs block); note that a TWTA
failure affects the system capacity by 24 2 Mbps channels.
Ka mission
Antenna
Receiver Chain
Transm. Chain w/o TWTA
Switch Matrix
Synthesizer
TWTA
miscellaneous (Harness)
TOT
Unit
4
48
48
1
1
48
1
Mass [kg]
8
0.7
0.55
4.27
3.1
2.5
15
R. factor
1
1.33
1.33
1
1
1.33
1.33
Mass
DC Power
Tot [kg]
32
44.69
35.11
4.27
3.1
159.6
19.95
Power/unit [W] Tot. Power [W]
0
0
0.55
26.4
0.5
24
12
12
20
20
20
1920
40
40
remarks
1 LNA+2 mix+2 BPF+2 IFA
1circ+2mix+2 IFA+BPF+lin
96 I/O + Controller
6 frequencies generator
@ η=0,5
hardware
298.72 Kg
2042.4 W
Table 2.12 : Mass and Power budget for the Ka mission
In the Ka mission a single LNA is associated with two transponders, and a failure isn’t as critical
as in the previous case, so that the whole receiver chain is redounded 4+3.
It can be noted how the two different satellite payloads consume the same amount of DC power,
though the Ka solution provides much more channels between the two continents.
16
3.1.6 More options
The good performance of the Ka payload pushes to design smaller satellites. One promising
solution is to halve the number of available links in order to keep low the satellite weight and DC
consumption. Such solution is analyzed in the following Table.
Antenna
Receiver Chain
Transm. Chain w/o TWTA
TWTA
Synthesizer
Switch Matrix
miscellaneous (Harness)
TOT
Unit
4
24
24
24
1
1
1
Mass [kg]
8
0,7
0,55
2,5
2,13
3,1
12
Mass
DC Power
R. factor
1
1,5
1,5
1,30
1
2
1,5
Tot [kg]
32
25,2
19,8
78
2,13
6,2
18
Power/unit [W] Tot. Power [W]
0
0
0,55
13,2
0,5
12
40
960
10
10
20
20
40
40
remarks
1 LNA+2 mix+2 BPF+2 IFA
1 circ.+ 2 mix+ 2 IFAs +BPF
@ η=0,5
6 frequencies generator
48 I/O + Controller
hardware
181,3 Kg
1055,2 W
Table 3.13 : Mass and power budget for an alternative mission.
As shown above, by halving the performance doesn’t imply cutting weight and mass in half. In
the first instance, the layout prevents the same technical solution from being reused. Moreover,
some electrical components such as the synthesizer, maintain their characteristics for both designs.
The only way to improve mass and weight features of the payload consists of exploiting new
technologies, by using lighter SSPAs instead of TWTAs for istance. However, every choice has its
drawback and it needs to be evaluated very carefully.
17
Chapter 4: Preliminary design
4.1 Introductory spacecraft budgets
The iterative process is ignited by the payload power and mass budget through a preliminary
design, which leads to a rough estimation of the spacecraft mass and power budget, represented in
Table 4.1 (a) to (d) for both solutions.
In addition to the total values calculated, a margin of 5% has been applied, in order to take into
account at this stage a generalized level of development for the various items.
Mass
[Kg]
Payload
200
Structure
128
TT&C
20
AOCS
75
Thermal
48
Propulsion
80
Power
201
Wiring
48
total dry
800
EOL 840
S/S
mission requirement
=
=
=
0.16 x M (sat. dry)
retrieved from analogue missions
0.06 x M (sat. dry)
0.1 x M (sat. dry)
batteries, solar arrays, electronics
=
=
0.06 x M (sat. dry)
4 x M (P/L)
=
1.05 x M (sat. dry)
Table 4.1 (a): Mass budget for Ku-band
S/S
Platform
Payload
TT&C
AOCS
Propulsion
BAPTA
Thermal
Charging
Regulation
Total Load
EOL
Power
[W]
2000
mission requirement
135
37
137
247
192
2749
2887
=
=
0.05 x P (sat.)
0.09 x P (sat.)
=
0.07 x P (sat.)
= P (Plt.) + P (sat.)
=
1.05 x P (sat.)
Table 4.1 (b): Power budget for Ku-band
18
Mass
[Kg]
Payload
300
Structure
192
TT&C
20
AOCS
75
Thermal
72
Propulsion
120
Power
201
Wiring
72
total dry
1200
EOL 1260
S/S
mission requirement
=
=
=
0.16 x M (sat. dry)
0.06 x M (sat. dry)
0.1 x M (sat. dry)
batteries, solar arrays, electronics
=
=
0.06 x M (sat. dry)
4 x M (P/L)
=
1.05 x M (sat. dry)
Table 4.1 (c): Mass budget for Ka-band
S/S
Platform
Payload
TT&C
AOCS
Propulsion
BAPTA
Thermal
Charging
Regulation
Total Load
EOL
Power
[W]
2000
mission requirement
135
37
137
247
192
2749
2887
=
0.05
=
0.09
=
0.07
= P (Plt.)
=
x
x
x
+
P (sat.)
P (sat.)
P (sat.)
P (sat.)
1.05 x P (sat.)
Table 4.1 (d): Power budget for Ka-band
4.2 Main spacecraft parameters
The result of this process, summarized in the following block diagrams (Fig. 4.2), will be used in
turn as inputs for a specialized analysis, developed in the following sections.
19
INPUTS
Ku mission
Satellite
Dry Mass
800 Kg
Payload
Mass
200 Kg
Payload
Power
2000 W
INPUTS
Cylinder
Height
2.3 m
Heat
Dissipation
1300 W
Radiating
Area
3.25 m2 ×2
Solar Array
Area
13.3 m2 ×2
Ka mission
OUTPUTS
Satellite
Dry Mass
1200 Kg
Payload
Mass
300 Kg
Payload
Power
2000 W
OUTPUTS
Cylinder
Height
2.6 m
Heat
Dissipation
1750 W
Radiating
Area
4.38 m2 ×2
Solar Array
Area
13.3 m2 ×2
Fig. 4.2 : Preliminary design block diagrams for the Ku (up) and the Ka (down) missions.
20
Chapter 5: Mission Analysis
The objective of the mission was already discussed in Chapter 2, whereas a definition of the
payload used to satisfy the given requirements was described in Chapter 3. This section analyzes the
various phases of the spacecraft, from its ascent with the launcher to the various steps leading to its
operative life.
5.1 Launcher
The first step in the launch system selection process is to establish the mission needs and
objectives, since they dictate the performance, trajectory, and the family of vehicles, which can
operate from suitable sites. With the mission requirements determined by the mission need, we
allocate them as functional requirements between the launch vehicle and payload. We must assess
each function required to achieve the mission objective through this process, and allocate functions
based on cost, reliability, and risk.
This iterative process should represent the baseline for a correct choice of the launch-system, but
is out of the scope of this study, and the launcher Ariane V was selected, as described in the section
of Propulsion Subsystem (Chapter 7).
As a general consideration, since launch-system reliability and cost are key factors to a
successful mission, it can be very cost effective to spend a bit more for a launch system with more
reliability. At this stage, Ariane V was selected, whereas a further study on alternative launchers
should be performed, aiming to achieve the same reliability at lower costs.
The main characteristics and performances of the chosen launcher are listed below in Table 5.1:
Ariane V characteristics
GTO parameters
Ra
Apogee radius 42166
Rp
Perigee radius
6898
i
Inclination
7
Dispersions (at 3σ)
Semimajor axis +/- 78
∆a
Inclination
0.3
∆i
km
km
°
km
°
Table 5.1 : Ariane V characteristics.
It follows a schematic diagram of the orbit staging, depicted in Fig. 5.1.
21
Fig. 5.1 : Orbit staging of the spacecraft system.
5.2 ∆v budget
After launching the satellite, its pre-operative life before reaching the final destination will be
divided into a series of different orbits, to each of them a cost in terms of propellant will be
combined.
In order to achieve the Geosynchronous operative orbit, an apogee “kick” will be performed.
Given the apogee transfer orbit velocity:
 1
1 
va = 2 µ E 
− 
 Ra 2 a 
and the geostationary orbit velocity:
vS =
µE
RS + RE
the ∆v needed to change the velocity from va to vS can be geometrically determined by means of
the Carnot theorem:
22
∆v = va2 + vs2 − 2va vs cos(i )
The corresponding geometry is depicted in Fig. 5.2:
Transfer
Orbit apogee
velocity va
∆v apogee engine
GEO Orbit
velocity va
Fig. 5.2 : ∆v apogee engine
The propellant used for this maneuver represents the main contribution to the total ∆v budget,
which is further composed of:
•
Dispersions due to the launcher;
o in the semi-major axis;
Differentiating the expression of ∆v in terms of the apogee velocity (as a
function of the semi-major axis), we obtain:
δ ∆v ( a ) =
o
va − v s cos(i )
δ va
∆v
in the inclination;
Differentiating the expression of ∆v in terms of the orbit inclination, we
obtain:
δ ∆v(i ) =
•
v a v s sin(i )
δi
∆v
Angle dispersions due to the apogee engine;
Being small the error, ε, in the angle of the apogee firing (we assume it equal to 1°
for the scope of this analysis), the corresponding δ∆v can be computed as:
δ ∆v = ε ⋅ ∆v
whose geometry is represented in Fig. 5.3.
23
correction ∆v
∆v
ε
Fig. 5.3 : Angle correction
•
Station keeping;
We will assume typical values of ∆v of 3 m/s/yr in the East/West direction and 48
m/s/yr in the North/South direction.
•
Attitude control;
0.35 m/s/yr, plus the 10% of the fuel consumption for the station keeping
•
De-orbiting;
The velocity change required to de-orbit a satellite with initial velocity vS in a
circular orbit with a higher semi-major axis ∆a can be computed for small values of
∆a (equal to 350 Km) as:
∆v =
1 v
∆a
2 2a
In addition, a margin has to be introduced, in order to take into account the inefficiency of the
thrusters and of the apogee motor.
Through these assumptions, the total budget for the ∆v can be written in a table.
maneuvre
delta v
m/s
apogee kick
launcher dispersions
apogee engine dispersion
station keeping E/W
station keeping N/S
attitude control
de-orbiting
1469.93
3.48
25.66
36.00
576.00
65.40
6.38
%
Margin
m/s
Total
m/s
2
2
2
5
5
5
5
1499.33
3.55
26.17
37.80
604.80
68.67
6.70
2247.02
29.40
0.07
0.51
1.80
28.80
3.27
0.32
TOTAL
Table 5.2 : ∆v budget
24
Chapter 6: Configuration
6.1 Requirements and constraints
The major drivers for the S/C configuration can be summarized as follow:
- Limited mass budget (the S/C should be as smaller and lighter as it is possible, with
same performances)
- I/F with the launcher
- Carrying large and heavy elements such as propellant tanks
- Providing direct load paths to the launcher
- Control type used
- Antennas accommodation
The spacecraft must provide accommodation to all the sub-systems and ensure compatibility
between them to throughout to the mission, i.e. avoiding electromagnetic interference among the
different antennas. Therefore each of the constraints as listed above must be fulfilled for every
operational mode and spacecraft attitude.
6.2 Spacecraft description
6.2.1 Baseline
Length and width of the spacecraft are mainly driven by the thermal control system (such as
radiators size), whereas height and the internal configuration by the propulsion system (such as the
accommodation of the fuel tanks). These dimensions are limited by the size of the launcher fairing.
Antennas, solar arrays and inner parts layout are selected in order to have the center of mass as
close as possible to the geometrical one. In fact a spacecraft has usually been modelled to be
symmetric, because the offset of the center of mass is directly connected to disturbing torques.
Configuration chosen:
Fig. 6.1 : Stowed configuration.
25
Ku mission
Fig. 6.2 : Spacecraft structure, main dimensions.
Fig. 6.3 : Spacecraft deployed configuration. Up view baseline.
26
Fig. 6.4 : Spacecraft deployed configuration. Down view baseline.
Ka mission
Fig. 6.5 : Spacecraft structure main dimensions.
27
Fig. 6.6 : Spacecraft deployed configuration. Baseline.
6.2.2 Launcher fairing storage
Defined the spacecraft structure main dimensions, it is
possible to choose where the satellite is stored in the
launcher faring. The stowed dimensions of both missions
allow using bottom position of Ariane 5 faring, which is
the cheapest one. In fact, nowadays there are many bigger
satellites that wait to be launched because they have no
partner to complete the minimum payload launcher
requirement.
Fig. 6.7 : Ariane 5 short faring main dimensions.
28
6.3 Optional configuration
Another possible configuration has been analyzed to increase rigidity of the spacecraft inversely
proportional to resonance frequency of deployed structure. This is very important for satellite
attitude control because if this frequency is too small the satellite could head towards instability
conditions. In fact, this new configuration of the solar arrays has the advantage to reduce their
lengths (factor ½), with the drawback of reducing the inertial characteristics (of a factor
approximately ½).
f ∝
I'
I
≈2 2 4
4
L'
L
L
2
I
I'≈
2
L' =
with
So the final result is positive: the resonance frequency is enhanced by a factor of 2 2 .
Fig. 6.8 : Spacecraft optional deployed configuration.
On the other side, the solar array deployment mechanism has to be more sophisticated,
increasing risks and costs of the mission. Although the development of more reliable deployment
mechanisms for this configuration is in progress, and probably mature for a more detailed study.
For this reason, it has not been taken into account in this phase. It should need more accurate
analysis because this new design could be a brilliant solution. In fact it should solve also others
problems connected to thermal control and structure subsystems that could allow to use another
cheaper launcher.
29
Chapter 7: Propulsion subsystem
7.1 Design Process
The process for selecting and sizing the elements of the propulsion subsystem requires the
following steps:
1. identify the functions of the satellite propulsion (e.g., orbit insertion, orbit
maintenance, attitude control, and controlled de-orbit or re-entry)
2. Determine ∆v budget and thrust level constraints for orbit insertion and maintenance
3. Determine total impulse for attitude control, thrust levels for control authority, duty
cycles (% on/off, total number of cycles) and mission life requirements
4. determine propulsion system options:
a. combined or separate propulsion systems for orbit and attitude control
b. high vs. low thrust
c. liquid vs. solid vs. Electric propulsion technology
5. estimate key parameters for each option
a. effective Isp for orbit and attitude control
b. propellant mass
c. propellant and pressurant volume
d. configure the subsystem and create equipment list
6. estimate total mass for each option
7.1.1 Spacecraft propulsion functions
After the satellite has been inserted in a Geo Transfer Orbit by the launcher, its propulsion
subsystem will provide the energy required to the transfer to operative orbit, to the station keeping,
to the attitude control (Sun/Earth acquisition, on orbit normal mode control, wheel desaturation, 3axis control during ∆v), and to the de-orbiting.
7.1.2 ∆v budget
The total ∆v budget was calculated in Chapter 5, while analyzing the spacecraft’s mission, and is
redrawn here for clarity’s sake.
maneuvre
delta v
m/s
apogee kick
station keeping E/W
station keeping N/S
attitude control
de-orbiting
1469.93
3.48
25.66
36.00
576.00
65.40
6.38
%
Margin
m/s
Total
m/s
2
2
2
5
5
5
5
1499.33
3.55
26.17
37.80
604.80
68.67
6.70
2247.02
29.40
0.07
0.51
1.80
28.80
3.27
0.32
TOTAL
Table 7.1 : ∆v budget
30
7.1.3 Total impulse , thrust levels, duty cycles and mission life requirements
This argument is not under the scope of this section, but will be analyzed while describing the
Attitude and Orbit Control System in Chapter 7.
7.1.4 Propulsion system options
a. Combined or separate propulsion systems for orbit and attitude control
A high ∆v, as the one needed for the orbit transfer, will require a high thrust, which can not
be achieved by a monopropellant system.
b. High vs. low thrust
Although a high thrust apogee engine is more efficient, it requires more powerful control
thrusters, due to higher angle dispersions. In order to fulfill these conflicting requirements, a
thrust for the apogee engine of 400 N has been chosen, which implies control thrusters of 10
N. The following figures show a possible selection:
Fig. 7.1 : 400 N Apogee engine (left) and 10 N control thruster (right)
c. Liquid vs. solid vs. Electric propulsion technology
A solid propellant system requires a huge effort to spend in the attitude control during the
apogee “kick”. Despite the very low fuel consumption (high Isp) of the electric propulsion
technology, the corresponding low thrust would not be adequate to operate the required
attitude control maneuvers.
The choice of a Unified Propulsion System (UPS) using bipropellant (N2O4/MMH) provide all
the three functions (orbit insertion and maneuvering, attitude control) to be performed with only one
higher performance system, having as a counterpart a more complex system. The use of
monomethyl hydrazine as fuel (MMH, ρ=0.88 Kg/dm3) and nitrogen tetroxide as oxidizer (N2O4,
ρ=1.47 Kg/dm3) would make possible to manufacture tanks of the same size, with a mixture ratio r
of 1.64, not far from their physical one, 1.67. A generalized UPS block diagram is represented in
Fig. 7.2.
31
Fig. 7.2 : Bipropellant system block diagram.
7.1.5 Estimate key parameters for each option
Having determined the Isp corresponding to the chosen propulsion system, we can estimate the
propellant mass as a fraction of the total launch mass:
− ∆v

∆M
g
⋅I
= 1.021 − e sp

M





where the propellant residual in the tank has been estimate at 2%.
Taking an average Isp of 300 s, we obtain that the propellant mass is 54.5% of the total mass.
Given a beginning-of-life (BOL) mass of 1600 Kg, the mass allocated to the propellant will be of
870 Kg, distributed in two spherical tanks with radius 0.44 m.
The propellant is fed to the thrust chamber simply by displacing it with a high pressure gas
(pressurant) contained in a tank, whose dimensions are easily computed using the perfect gas law,
with the assumption of an isothermal process, resulting in two tanks with radius 0.16 m.
Having in mind the Maximum Expected Operative Pressure (MEOP) experienced by the
propellant and pressurant tanks, and the ultimate stress of the selected material (aluminum in our
case), the corresponding wall thickness can be estimated using
t=
p⋅r
2σ
where a safety factor is taken into account in the allowable ultimate stress.
32
7.1.6 Estimate total mass for each option
A complete list of the propulsion subsystem is given below, where some of the data (not
computed in this section) were retrieved from a similar UPS system.
equipment
quantity
fill & drain valves
pressurant tank
liquid apogee engine
liquid filter
propellant tank
latching valve "A"
latching valve "B"
pyro valve NC
pyro valve NO
check valve
pressure regulator
pressure transducer "A"
pressure transducer "B"
pilot valve
reaction control thruster
UPS piping & fitting
mass
unit (Kg) total (Kg)
SET
2
1
2
2
2
2
5
2
4
1
1
2
1
16
SET
0.600
10.748
4.200
0.230
15.962
0.500
0.600
0.160
0.160
0.080
1.200
0.400
0.400
0.250
0.420
6.000
TOTAL
0.600
21.496
4.200
0.460
31.923
1.000
1.200
0.800
0.320
0.320
1.200
0.400
0.800
0.250
6.720
6.000
77.689
Table 7.2 : Mass budget
7.2 Option 2
The same type of considerations can be extended to Option 2 (Ka-band based payload), leading
to the following propellant and equipment mass budget (Table 7.3 and Table 7.4).
maneuvre
delta v
m/s
apogee kick
station keeping E/W
station keeping N/S
attitude control
de-orbiting
1469.93
3.48
25.66
30.00
480.00
54.50
6.38
%
Margin
m/s
Total
m/s
2
2
2
5
5
5
5
1499.33
3.55
26.17
31.50
504.00
57.23
6.70
2128.47
29.40
0.07
0.51
1.50
24.00
2.73
0.32
TOTAL
Table 7.3 : Mass budget for the second option.
33
equipment
fill & drain valves
pressurant tank
liquid apogee engine
liquid filter
propellant tank
latching valve "A"
latching valve "B"
pyro valve NC
pyro valve NO
check valve
pressure regulator
pressure transducer "A"
pressure transducer "B"
pilot valve
reaction control thruster
UPS piping & fitting
quantity
SET
2
1
2
2
2
2
5
2
4
1
1
2
1
16
SET
mass
unit (Kg) total (Kg)
0.600
16.191
4.200
0.230
24.045
0.500
0.600
0.160
0.160
0.080
1.200
0.400
0.400
0.250
0.420
6.000
TOTAL
0.600
32.382
4.200
0.460
48.091
1.000
1.200
0.800
0.320
0.320
1.200
0.400
0.800
0.250
6.720
6.000
104.743
Table 7.4 : ∆v budget for the second option.
34
Chapter 8: Thermal Control System
In this chapter we shall study the thermal control subsystem for two kind of satellites: one with a
communication payload which uses Ku band and another one with a payload in Ka band.
8.1 Ku mission
For the analysis of thermal control of a satellite, first of all, it is necessary to dimension the
radiating elements and afterwards the heaters.
Two radiator panels are considered. They are positioned on the planes, perpendicular to pitch
axis, one on the north side and another one on south side.
Radiator
Pannels
(±Y face)
Fig. 8.1
Using the formula for the radiator’s initial dimensioning, we obtain the value of its area:
A = Pw /(2*200) = 1300/400 = 3.25 m2
where Pw is the power which must be dissipated and 200 W/m is its capacity to dissipate power.
Radiator’s shape is a rectangle with sides equally to l1 = 1.4 m and l2 = 2.32 m.
We suppose that the specifications impose a radiator’s working temperature between 5°C and
45°C that is a typical range for GEO missions.
35
To level temperature’s value, heat pipes are provided inside the radiator; in this way it is justified
an analysis in isotherm and stationary conditions. It is clear that punctual solutions are foreseen for
possible hot or cold spots. On radiator’s surface OSR (Optical Solar Reflector) are settled. These
are characterized by a low value of solar absorptance and a high value of emittance in the infrared
band. They are suitable for direct solar radiation and for albedo, because they reflect most of the
incident energy and, at the same time, permit an effective dissipation of the on board generated
heat.
Fig 8.2
To verify that the working temperature’s conditions are respected, we use the fundamental
equation of thermal balance:
m Cp dT/dt = αs Csun 2A sinθ + Pw – εir 2A σ T4 η
where m = radiator mass
Cp = radiator specific heat
T = radiator temperature
αs = radiator assorbivity
Csun = solar constant
θ = sun incident angle on the radiator
εir = radiator infrared emittance
σ = Stefan-Boltzmann constant
η = radiator efficiency
It is clear that being the satellite on a geostationary orbit, in this equation the contribution of
albedo isn’t considered.
In stationary condition dT/dt = 0, then the fundamental equation of thermic balance becomes:
αs Csun 2A sinθ + Pw = εir 2A σ T4 η
The next step is to make verification in the case of maximum thermal load that occurs during
Solstice (called “hot” case) and of the minimum thermal load during Equinox (called “cold” case).
36
Fig 8.3 : solar illumination of GEO satellite
For the “hot” case:
T = [ ( αs Csun 2A sinθ + Pw)/ εir 2A σ η ]1/4
In the equation above, we suppose the values already used for other thermal analysis in GEO
missions. They are listed in the following table:
Pw
αs
εir
Csun
θ
η
1300 W
0.1 (BOL), 0.27 (EOL)
0.8
1399 W/m2
23°
0.95
Table 5.1
The obtained temperature is:
T = 292.35 K
as degree, T becomes:
T = 19.35 °C < 45 °C
Therefore, the working temperature is verified.
Heater’s power has been calculated assuming stationary conditions. Actually, during an eclipse,
satellite’s cooling happens to follow an exponential law so that it is difficult for the satellite to reach
the asymptotic temperature, as supposed in the thermal balance equation.
To characterize temperature’s exponential law it is essential to know satellite’s thermal capacity
that isn’t available in this analysis. So it is justified, in the equation of thermal balance, the use of
the lowest considered temperature of 0°C to calculate heater’s power.
Now we shall see what happens in the “cold” case. In this case θ = 0°, thus there is no
contribution from solar radiation in the equation of thermal balance. The temperature results equal
to:
T = 261 K → T = –12 °C < 5 °C
This condition is not acceptable; consequently it is necessary to use heaters. To know the power
Qheaters that the heaters should provide, we make use of the equation of thermal balance where we
require a working temperature of 5 °C.
37
Qheaters = – Pw + εir 2A σ T4 η
By replacing the values, we obtain a value of:
Qheaters = 373 W
In order to reduce the required power for the heaters, it is possible to utilize MLI (Multi Layer
Insulation) to replace a part of the radiator of 0.5 m2, ensuring that the temperature, in the “hot”
case, doesn’t exceed the limit of 45 °C:
Qheaters = 115.5 W
Now we make verification of the “hot” case:
T = 307.2 K → T = 34.2 °C < 45 °C
This condition fulfills the working requirements.
To complete the thermal design, it is expected that the satellite’s lateral panels are protected with
MLI, which is sufficient to screen the incident solar radiation, reminding that for these panels the
value of θ is very high (during equinox θ = 90°).
8.2 Ka mission
For this case the same criteria of the first satellite are used; the radiator panels are positioned in
the same way. By using the same formula for the initial dimensioning of radiator panels we obtain
an area of:
A = 4.375 m2
Radiator’s shape is a rectangle with sides equally to l1 = 1.69 m and l2 = 2.58 m.
By using the fundamental equation of thermal balance replacing the following values:
Pw
αs
εir
Csun
θ
η
1750 W
0.1 (BOL) , 0.27 (EOL)
0.8
1399 W/m2
23°
0.95
Table 8.2
we compute for the working temperature a value of:
T = 299.68 K → T = 26.68 °C < 45 °C
that verify the working requirements.
In the same way as the first satellite, the verification of the “cold” case is carried out. The
temperature is equal to:
T = 261.01 K → T = – 12.01 °C < 5 °C
Being this value not acceptable, heaters are necessary. The power that heaters have to supply is
given by:
Qheaters = – Pw + εir 2A σ T4 η = 502 W
38
To reduce the required power for the heaters, it is possible to use MLI (Multi Layer Insulation)
to replace a part of the radiator of 0.8 m2, making sure that the temperature, in the “hot” case,
doesn’t exceed the limit of 45 °C:
Qheaters = 90.27 W
The temperature in the “hot” case is:
T = 308.90 K
---›
T = 35.90 °C < 45 °C
To complete the thermal design, also for this satellite, the lateral panels are protected with MLI.
39
Chapter 9: Power subsystem
9.1 Requirements
The power subsystem was designed to supply the payload with the required power,
approximately 2 KW, and all the other bus subsystems, i.e., telemetry and telecommand, attitude
and orbit control, propulsion, thermal control. The power requirements for the satellite subsystems
are shown in Table 9.1.
SUBSYSTEM
Payload
Batteries
Power Supply
TCR, AOC, UPS
BAPTA
THC
POWER SUPPLY
(W)
2000
275
150
140
37
115
NOTES
Only in charge
Table 9.1 : Power requirements for the satellite subsystems.
The 150 W mentioned in the ‘power supply’ line in Table 9.1 is the power absorbed by the
control electronics supervising the power distribution, regulation and protections;
The power required by the battery is the one necessary to charge it, thus, it must be considered
only when the solar array is illuminated by solar flux.
The required power supply is the power absorbed by the control electronic that provides at the
power distribution.
TCR, AOC and UPS, represent the power absorbed by the Telemetry Command System, Attitude
Orbit Control and Unified Propulsion System.
BAPTA is the power required by SADMs (Solar Array Drive Mechanism) for the correct solar
array orientation.
Finally, THC is the thermal control system power consumption.
The power required by each subsystem is the same in both satellites; therefore only one EPS
design has been done. The power subsystem was designed for a 12 years mission.
9.2 Baseline design
During the solar array design, multi-junction cells were considered with respect to standard
silicon solar cells. Although multi-junction cells are more efficient than the silicon ones, resulting in
a smaller mass and size, they are more expensive and their technology is somewhat not mature for a
commercial mission, where reliability is a key factor.
The batteries selected use Ni-H2 cells rather than Li-Ion cells. Although Li-Ion cells allow to
halve the mass for the same power and do not need trickle charge, they require a different design for
the battery control electronic and different laws for the battery charge, with an increase of non
40
recurrent costs. Although developed very recently, this technology is nowadays commonly used on
modern spacecraft, suggesting a finer research for a further study.
A single bus architecture will be used in this configuration. Even though the double bus
simplifies the harness (that would not cross the satellite), it is more heavy than the single bus and
less easy to manage.
The power subsystem is totally regulated. For this reason a heavy and bulky discharge regulator
was included, assuring though constant current and voltage in every condition for the electronic
safety. As a rule of thumb, we use total regulation if the satellite power exceeds 2 KW.
The power supply voltage is fixed at 50 V. It derives from the fact that if the required power is
below 7 KW we have to use a voltage between 42,5 V and 50V.
9.3 Design
9.3.1 Subsystem configuration
The power subsystem contains:
•
2 solar array;
•
2 SADMs;
•
1 SADE, the SADM’s electronic;
•
1 MRU (Main Regulator Unit) that assures:
−
regulation with shunts in sunlight condition;
−
regulation in eclipse (BDR, Battery Discharge Regulator);
−
battery charge;
−
SGP (Single Ground Point);
−
external interface (EGSE, umbilical connection);
•
1 PPDU (Power Protection and Distribution Unit);
•
1 TCU (Thermal Control Unit);
•
1 PDU (Pyro Drive Unit);
•
2 batteries connected on the bus through the BDRs.
9.3.2 Solar arrays
We consider a GEO satellite for a 12 years mission, requiring
P ( sat ) = 2720W
of maximum power in the sun, during the equinox (including the power to charge batteries) at
the exit of MRU. The end-of-life power, P(EOL), can be determined as
P ( EOL) = 1,05
2720
= 2992 ≅ 3071W
0,93
where a margin of 5% has been consider for the power generated from solar arrays, and a
regulation efficiency of 0,93, given by
η=
P (out )
= 0,93
P (in)
so that there is a power loss of 7% in the shunts, in sunlight condition.
Furthermore, we can write
41
4
P ( EOL) = P ( BOL)∏ Ri
i =1
4
where P(BOL) is the beginning of life power, and
∏ R is the product of the following terms:
i
i =1
R(electrons/protons) = 0,82
loss for interaction between cells and protons and electrons;
R(micrometeorites/UV) = 0,97 loss for interaction between cells and micrometeorites and
ultraviolets rays;
R(failure) = 0,96
margin for failures;
R(calibration) = 0,98
margin for calibration error;
if we substitute:
P ( BOL) =
3071
= 4104W
0,82 ⋅ 0,97 ⋅ 0,96 ⋅ 0,98
with the relation
P ( BOL) = CAfη
where C is the sun constant, 1353 W/m2 in average, f is the filling factor, 0,85 (minimum value),
η is the solar cells efficiency at the beginning of life, 13% for silicon standard cells, we can
calculate the solar arrays area A:
A=
4104
= 27,45m 2
1353 ⋅ 0,85 ⋅ 0,13
9.3.3 Design of solar arrays electrical net
The diagram of the solar arrays electrical net is represented in Figure 9.1.
SLIP
RING
BLOCKAGE
DIODES
Vbus
Vstring
SOLAR
ARRAYS
BATTERIES
Fig. 9.1 : Solar array electrical net.
The diode interposed between the slip ring and the solar array prevents current from flowing in
the opposite way, in order to avoid current absorption by solar array thus power supply reduction.
According to this diagram, the following equation can be written
42
Vstring = Vbus + ∆Vdiodes + ∆Vslipring + ∆Vharness + tolerance
where ∆Vdioesi is 2,1V, ∆Vslip ring + ∆Vharness is 0,4V, the tolerance for Vbus is 0,5V, and imposing
the bus voltage at 50V, we have
Vstring = 50 + 2,1 + 0,4 + 0,5 = 53V
that is the voltage generated by the solar arrays.
The current generated by the solar arrays is:
Isa = Ibus =
Pbus 2720
=
= 54,4 A
Vbus
50
the solar arrays are constituted from frames producing a current of 54,4 A and an end of life
voltage of 53 V.
Before proceeding we set the working temperature of the solar array to 38°C. We get such value
by the thermal equation of the panel, as follows
αCS i = ηCS i + ε i S iσT 4 +ε bS bσT 4
where the emittance εi = 0,85 and the absorptance α = 0,74 are related to a solar cell with a
coverglass CMX, Si is the illuminated surface equals to 13,725 m2, Sb is the back surface of the
panel, C the solar constant set to 1353 Wm-2and η the solar cell efficiency (0,13).
The cells series number is thus:

 ∆V  
Ns Vmp ⋅ Rv − 
∆T  = 53V
 ∆T  

In the previous relation, Vmp is the maximum power voltage (0,454 V at 28 ˚C for silicon
standard cells), ∆T is the difference between working temperature and reference temperature, 28˚C,
∆V/∆T is the voltage variation with respect to the temperature difference of 1°C (2,2· 10-3 V/°C for
silicon standard cells), Rv is the loss coefficient, for the only voltage, for radiation
(electrons/protons) at maximum power point, 0,942. If we substitute, we have
Ns =
53
= 131
0,454 ⋅ 0,942 − 2,2 ⋅ 10 −3 ⋅ 10
The parallel cells number, Np, is

 ∆I  
Np Im p + 
∆T  ⋅ R = 54,4 A
 ∆T  

Imp is the maximum power current, if we use size cells of 4·6 cm2 we have Imp=36,6 mAcm-2 ·6·4
cm2; ∆I/∆T=0,0005·Imp, R is the total loss coefficient that is
R=
0,82
⋅ 0,98 ⋅ 0,96 ⋅ 0,97 = 0,80
0,942
leading to
Np =
54,4
= 78
[0,8784 + 0,0005 ⋅ 0,8784 ⋅ 10]⋅ 0,80
In conclusion, the solar arrays are constituted from 78 parallel frames, each of 131 cells, split in
43
2 arrays, each of 39 frames and with an area of 13,725 m2. The output voltage, at the end of life and
at the equinox sunlight condition, is 53 V, whereas the current is 54,4 A.
For the solar arrays with silicon standard cells, the generated power, at the end of life, is 40
W/Kg. Therefore, the resulting solar arrays weight, including solar cells, panels, coverglass and so
on, is 77 Kg.
9.3.4 Batteries design
In order to design the electric batteries we will make the following assumptions:
•
•
•
•
2 batteries, one for each radiator;
26 cells for each battery;
the battery capacity is calculated for the case of one broken cell in open circuit (worst case
of failure);
total regulated bus.
From the equation
P (bus ) = 2η [(n − 1) ⋅ V − 0,75] ⋅ DoD ⋅
C
t
where η is the discharge regulator efficiency, 0,90, n is the cells number, V is the voltage average
during battery discharge, 1,22 V for Ni-H2 cells, 0,75 is the loss voltage on the by-pass diode:
Vbattery
By-pass diode, for
each battery cell
Fig. 9.2 : Battery diagram.
44
DoD is the battery Depth of Discharge, that for Ni-H2 cells and for a 12 years missions can reach
today the 80%, P(bus) is the bus power, (2720 - 275) W, C is the battery capacity, t is the discharge
time, for GEO eclipse maximum is 1,2 hours, and if we are considering n-1, according to third
assumption, we have
C=
P (bus ) ⋅ t
= 69 Ahr
2η [(n − 1) ⋅ V − 0,75] ⋅ DoD
For Ni-H2 technology we can calculate the mass of batteries from the following relation:
energia
Whr
= 50
massa
Kg
then
M = 69 ⋅ 26 ⋅
1,22
= 43,8 Kg
50
In conclusion, the battery system is constituted from 2 Ni-H2 batteries, each of 26 cells with 69
Ahr capacity, and the total mass is 43,8 Kg. In the mass budget we also have to consider about 47
Kg, 30 Kg for the electronic regulation mass and 17 Kg for the electronic distribution mass, pyro
mass and BAPTA mass.
9.3.5 Batteries charge
For a DoD of 80% and a capacity of 69 Ahr, the batteries can discharge
0,80 ⋅ 69 = 55,2 Ahr
Since the charge efficiency is less than 100%, during the charge phase we have to release an
energy increased of 15 %, thus, we have an overload factor of 1,15.
We set the current for the battery charge, Ichg, to C/12 (it is included between C/10 and C/20)
1,15 ⋅ 55,2 = 63,48 Ahr
Ichg =
C
= 5,75 Ahr
12
then the charge time is
t=
63,48
= 11,04hr
5,75
If we carge the batteries in a sequence, we spend 22,08 hours. Given that
24hr − 1,2hr = 22.8
we have at most 22,8 hours of sunlight to charge the batteries, and we can complete the charge
before the eclipse
9.3.6 Charge power
We want to calculate the power absorbed by the batteries during the charging phase, Pchg;
assuming that the charge voltage of each cell is 1,5 V
Pchg = Vchg ⋅ Ichg = 1,5 ⋅ 26 ⋅ 5,75 = 224,25W
45
The batteries charge efficiency is 91%, resulting in a power coming from the bus
Pchg (bus ) =
224,25
= 246.5V
0,91
that is consistent with 275 W reserved to the batteries at the beginning of design, also including
about 20 W for the batteries trickle charge.
9.3.7 Multi-junction cells
If we use multi-junction cells instead of standard silicon cells, we may have advantages deriving
from arrays with lower size and mass. By using, in first approximation, the same R due to radiation,
from the relation
P ( BOL ) = CAfη
where η is 0,26 because the multi-junction efficiency is 26%, we obtain an area of 13,85 m2 with
respect to 27,7 m2.
For the mass array we have
M =
3100
= 56,4 Kg
55
instead of 77 Kg with the silicon standard cells.
46
Chapter 10: Attitude Orbit Control Subsystem
10.1 Orbit and Design Definition
To test and simulate the specific utilized AOCS it is used
ADS (AOCS Design Software) version 2.0 produced by
Alcatel for ESA. First of all it is necessary to define the design
and the orbit of the spacecraft:
1. Design:
see Configuration in Chapter 5.
It is important to remark that ADS has calculated the correct position of mass center into
body frame. The order of magnitude of this value is more or less 5 cm.
2. Orbit:
in order to respect the specifications, a GEO orbit is chosen with -30° (Westwards) for
longitude.
Fig. 10.1 : View of satellite on its orbit
47
10.2 Control modes & Requirements
In this preliminary study only nominal mode has been analyzed and modeled.
In order to respect the mission requirements directly connected to the payload, it is necessary
that the satellite be always pointed toward the Earth with an accuracy of:
I.
Ku: 0.2° for each axes
Payload requirements have fixed that spot beams have to be pointed with
an accuracy of 0.3° for each axes. This constraint has been applied to the
platform taking into account a margin in order to be source that it has been
respected also in not nominal situations such as station keeping maneuvers.
II.
Ka: 0.05° for each axes
Payload requirements have fixed that spot beams have to be pointed with
an accuracy of 0.05° for each axes. This constraint is much tightened and it
does not take into account the possible misalignments and off-loads of each
antennas, therefore an optional design has been introduced.
10.3 Disturbance Torque Computation
Before the selection of spacecraft control type, it is important to determine the magnitude of the
torques that the AOCS must tolerate. Once established the completed design, it is possible to
estimate disturbances, which are directly functioned to time and spacecraft attitude. To obtain an
order of magnitude of these perturbations they are calculated in worst-case.
10.3.1
Environmental
The principal sources of disturbing torques are:
• Gravity-gradient
(I zz − I yy )∆eϕ − I yz 
r
r
2 r
2
Tg = 3ω 0 e × Π e ≅ 3ω 0 (I zz − I xx )∆eϑ − I xz 


0
• Solar Radiation
ω 0 orbit angular velocity
with 
∆e attitude errors
r
r
r
Fsp = − pA (1 − cs ) s + 2 ( cs cos ϑ + 1 3 cd ) n  ≤ 2 pA
with
Tsp = Fsp ( c ps - cg )
 p solar pressure = 4.4 ⋅10−5 N 2
m

 A surface area
r
s versor of solar radiation
nr normal to A

cd diffusion coefficient
c scattering coefficient
 s
c ps center of solar pressure

cg center of gravity
• Magnetic Field
If we model Earth’s magnetic field with approximation by a dipolar field:
48
M
B (r , λ ) = 30 1 + 3 sin 2 λ
r
r
r r
Tm = D × B
with
with
M 0 magnetic moment = 8.056 ⋅1015 Tesla m3

r distance to Earth − center
λ latitude

 D residual dipole of S / C < 2 A·m 2

 B Earth' s magnetic field
• Aerodynamics
r
r
v
1
2
Fdrag = − C D ρ v A r
v with
2
Tdrag = Fdrag (c pa − cg )
 ρ atmopheric density
C = 2.2 drag coefficient
 D
 A surface area

v spacecraft velocity
c pa center of aerodynamics pressure

cg center of gravity
In GEO orbit the atmospheric density is negligible therefore it is not relevant.
10.3.2
Internal
Fortunately, we can respecify it to tighter values. This change would reduce its significance but
most likely add to its cost or weight.
The principal disturbances are due to emission power of antennas:
 Pa emission power of antenna i
r

Pai r
with c light velocity
FRF = − ∑
Z ai
working c
r r
antennas
 Z ai Z of antenna i
The following disturbances are not periodic but they produce their not negligible effects only at
particular instants:
•
•
•
•
•
•
•
•
Uncertainty in Center of Gravity
Thrust Misalignment and off-load
Liquid Sloshing
Gas Leakage
Rotating Machinery (pumps, tape recorders, gyros)
Dynamics of Flexible Bodies
Thermal Shocks on Flexible Appendages
Shocks (pyros, latching valves)
Therefore these impulsive disturbances are used to design the maximum torque that actuators
have to be provided.
Their order of magnitude can be assumed to be less than 10-2 Nm.
10.3.3
49
Results in worst-case
Disturbance
Gravity Gradient
Solar Radiation
Magnetic Field
Aerodynamics
Emission Power
TMAX
Order of magnitude [Nm]
Ku
Ka
10-7
10-7
6·10-5
6·10-5
2·10-7
2·10-7
negligible
negligible
1.5·10-7
1.5·10-7
6·10-5
6·10-5
Table 10.2 : Disturbances torque in worst-case
There is no big difference between the two cases, because their configurations and sizes are more
or less very close. In fact we have supposed to use the same solar array, and so on. The main
disturbance is due to solar radiation; the others are negligible.
10.3.4
Results by ADS
The same calculus is done by ADS. It is more precise because it takes account of the correct
design and the orbit of our satellite; therefore it is possible to obtain a profile for each disturbance.
It is possible to notice a correlation between the two results: they have the same order of magnitude.
Solar radiation
It is possible to notice that during this orbit the satellite is eclipsed by the Earth, in fact for few
minutes the torque due to solar radiation is zero.
Fig. 10.2 : Solar radiation disturbance torque calculated by ADS
50
Other disturbances
They are more then an order of magnitude below the solar radiation:
Fig 10.3 : Other disturbance torques calculated by ADS
The ADS profile has one bug; torque X is not plot, because it is out of scale. Its average value is
739·10-9 Nm.
10.4 Actuators Trade-off
Once we have defined the requirements and the order of magnitude of disturbing torques, we are
ready to select a method to control spacecraft attitude. First of all we have chosen if we want to use
a passive or active control technique.
10.4.1
Passive stabilization
It uses environmental proprieties in order to fix the satellite into the reference attitude even if
disturbing torques perturbs the vehicle. The principal types are:
•
Aerodynamic control
It uses the presence of the air by some mobile part like a flap to change the satellite
attitude. Therefore it is used only for near-Earth orbit where the atmospheric density
is considerable, it has no application for geostationary orbit.
•
Gravity-gradient control
It uses the inertial properties of the vehicle to keep it pointed toward the Earth. This
relies on the fact that an elongated object in a gravity field tends to align its
longitudinal axis through the Earth’s center. The torques which cause this alignment
decrease with the cube of the orbit radius, and are symmetric around the nadir vector,
51
M
thus not influencing the yaw of a spacecraft
around the nadir vector. To have a stabile
satellite we must have:
M L2 >
L
TMAX
3 ω0 2 ∆e
with
ω 0 orbit angular velocity

∆e attitude errors
X
Z
Fig 10.4 : Gravity gradient actuator
•
Magnetic control
It uses permanent magnets on board the spacecraft to force the alignment along the
Earth’s magnetic field. This is most effective in near-equatorial orbits where the field
orientation stays almost constant for an Earth-pointing vehicle, but it is not possible
to control torque around this axes. To have a stabile satellite we must have an on
board permanent magnet with:
D>
•
10.4.2
TMAX
B
D is magnetic dipole of S/C
Spin control
It employs the gyroscopic stiffness to reduce the effect of small, cyclic disturbance
torques. In fact if the body is initially spinning around an axis perpendicular to
applied torque, the body spin axis will precess, moving with a constant angular
velocity proportional to the torque. Thus, spinning bodies act like gyroscopes,
inherently resisting disturbance torques in 2 axes by responding with constant, rather
than increasing, angular velocity. The spinning motion is stable (in its minimum
energy state) if the vehicle is spinning about the axis having the largest moment of
inertia. Energy dissipation mechanisms on board (such as fuel slosh and structural
damping) will cause any vehicle to head towards this state if uncontrolled. This is
possible using the entire body spins or just a potion of it, such as a momentum wheel
or spinning rotor. For Earth pointing mission, specially for telecom where we have to
point each antenna at different part of the world, using spinners is not applicable, but
it is prefer to use or a dual-spin or a satellite with a momentum bias on board.
Active stabilization
It is more common today specially for missions that require high accuracies and a versatile
spacecraft. The disturbing torques are predicted and hindered by some actuators in order to
respect the expected attitude:
1. Momentum bias
These systems often have just one wheel with its spin axis mounted along the pitch
axis, normal to the orbit plane. The wheel is run at nearly constant, high speed to
provide gyroscopic stiffness to the vehicle, just as in spin stabilization, with similar
52
nutation dynamics. Around the pitch axis, however, the spacecraft can control
attitude by torquing the wheel, slightly increasing or decreasing its speed. This is the
main difference with passive stabilization. Periodically, the pitch wheel must be
desaturated (brought back to nominal speed) using thrusters or magnets.
One approach to estimating wheel momentum, h, is to integrate the worst-case
disturbance torque, TMAX, over a full orbit. Since the main disturbance is solar
radiation, the maximum disturbance accumulates in ¼ of an orbit. A simplified
expression for such a sinusoidal disturbance is:
H = (T⊥Wheel ) MAX
T⊥Wheel disturbing torque ⊥ wheel axes
P 1

with  P orbital period
4 ∆e
∆e allowable motion

2. Zero-momentum
The reaction wheels, which can be compared to momentum wheel with an initial
zero speed so do not provide a gyroscopic stiffness to the vehicle, respond to
disturbances changing their speed. If disturbance is cyclic during each orbit, the
wheel may not approach saturation speed for several orbits. Secular disturbances,
however, cause the wheel to drift toward saturation. To avoid it an external force has
to be applied usually by thrusters or magnetic torque in order to force the wheel
speed back to zero. This process, called desaturation, momentum unloading or
momentum dumping, can be done automatically or by command from the ground.
One approach to estimating wheel momentum, h, is similar to the previous one:
H = (TWheel )MAX
P
0,707
4
where P is the orbital period, TWheel disturbing torque along wheel axes and 0,707 is
the average value of an unitary sinusoidal function. In this expression does compare
angular accuracies because reaction wheels do not have a gyroscopic stiffness.
3. Gas jets or Thrusters
They produce toque by expelling mass, and are not governed by the same concerns
as momentum storage devices. They can provide large, instantaneous torques at any
point in the orbit, but, unfortunately, their plumes may impinge on the spacecraft,
contaminating surface, and they require expendable propellant.
So we have decided to use them only to provide station-keeping maneuvers and to
desaturate the wheels. See Propulsion subsystem part in order to have more details.
4. Solar sail
They provide torques by means of moving sails, by changing the satellite centre of
pressure. In fact the figure 10.5 shows how torques are obtained by orientation of the
solar arrays equipped with flaps.
53
Fig 10.5 : Solar sails actuators used by Matra Marconi Space
Only roll and yaw axes can be control by solar sails. These devices are new and not
completely analyzed; in fact only few “pioneer” companies use them. For this reason
they are not taken into account during this preliminary study.
10.4.3
Results
In order to respect attitude requirements defined before, it has calculated the size the size
of each control type
Control Type
Ku
Ka
Aerodynamic
Gravity Gradient
Magnetic Field
Momentum Wheel
Reaction Wheel
not applicable
M·L2 > 3.3·106 kg m2
D > 568 A m2
H > 123 N m
H>1Nm
not applicable
M·L2 > 7.2·106 kg m2
D > 568 A m2
H > 825 N m
H>1Nm
Table 10.3 : Control type results
For each actuator has to be able to produce a torque bigger than the maximum disturbing
torque.
Ku mission
Momentum bias, mounted along the pitch axis, has been selected in order to
respect the mission requirements and minimize the budget mass of AOCS. Also the
thrusters are used to desaturate the wheel and to execute station-keeping maneuvers.
54
• Momentum wheel:
In order to have a system that is able to reach more than 82 N m, it has decided
to uses two momentum wheels in V configuration to increase the angular
momentum on pitch axes:
α
α
α
α
Fig.10.5 : Momentum wheel configuration
H Y = 2 H wheel cos α
⇒
α = 34°
It has been chosen two wheels RDR 68 produced by TELDIX Bosch Telecom:
H
ω
TMAX
Pnom
PMAX
M
Parameter
angular momentum
rotation speed
max torque
nominal power
maximum power
weight
Value
51÷74 Nms
4500÷6600 rpm
± 0.085 Nm
9÷16 W
90 W
8 kg
Table 10.4 : Momentum wheel characteristics
• Thrusts:
A standard solution have been selected using 16 thrusters (8 for redundancy),
which are placed in the following configuration (Fig. 10.6). Other
configurations could be investigated during an advanced phase in order to find
more cost-effective ones. In fact, nowadays, thrusters are not more doubled for
55
redundancy, but other possible set of configuration are preferred to reduce the
number of thrusters increasing operative modes, which should allow to obtain
the same result of traditional configuration.
Z
X
Y
Fig.10.6 : RCT configuration
Thruster
Orientation
Set A
tilted in S/C X-Y plane
Set B
tilted in S/C X-Y plane
Set C
tilted in S/C X-Z plane
5C / 6C
tilted into +Z for minimization of
plume effects on antennas and
disturbance torque
Placement
on the corner of central cube
on the corner of central cube
on the edge of S/C side; the Ycomponent is chosen according
to the disturbance torque
compensation requirements of
apogee boost phase
S/C East and West side
Table 10.5 : RCT orientation and placement
Characteristics of selected thrusters:
Parameter
Nominal thrust
Thrust range
MIB nominal thrust
Value
10 N
7.4÷11.9 N
25 mNs
Table 10.6 : Thrusters characteristics
where MIB is the Minimum Impulse Bit.
56
Ka mission
The requirements of this mission are very narrow, in fact to keep satellite attitude
in a range of 0.05° only reaction wheel can be used. Again the thrusters are used to
desaturate the wheel and to execute station-keeping maneuvers.
• Thrusts:
Same reaction control thrusters are used like Ku mission.
• Reaction wheels
In order to completely control satellite attitude, three wheels are required and
an additional one is added for redundancy and also to reduce the mechanical
noises into AOCS. Reaction wheels configuration:
Fig.10.7 : Reaction wheels configuration
It has been chosen four wheels RSI 4-75 produced by TELDIX Bosch
Telecom:
H
ω
TMAX
PMAX
M
Parameter
angular momentum
rotation speed
max torque
maximum power
Weight
Value
4 Nms
6000 rpm
0.075 Nm
90 W
4.2 kg
Table 10.7 : Reaction wheel characteristics
57
Optional design for Ka mission
Accuracy requirements are so tightened that chosen AOCS subsystem would
be very expensive and complicated. Therefore it will be foreseen a new design of
the configuration: all antennas could be mounted on a two-axis controllable
support on Earth phase panel. This support would be completely isolated from
the main platform of the satellite, so it is possible to reduce global spacecraft
accuracy to, for example, 0.5 °. The final result would be: main attitude control
subsystem is less expensive and another more precise one for controlling Earth
facing platform. The first one could be like Ku mission using one or two
momentum wheels and the other one could be done by using 2 linear actuators
and 1 spherical hinge. This new configuration is more challenging but it allows
reducing a lot of AOCS mass and power budgets. This innovative idea has not
been analyzed in this phase of the project but it will be taken into account only in
an advanced phase.
spherical hinge
Z
Y
linear actuators
X
Fig.10.8 : Optional design for Ka mission
10.5 Sensors Selection
Sensors selection is most directly influenced by the required orientation of the spacecraft and its
accuracy. Other influences are included redundancy, fault tolerance, field of view requirements, and
available data rates. Typically, we identify candidate sensor suites and conduct a trade study to
determine the best, most cost-effective approach.
For medium-high accuracy missions, gyros have to be selected because they are able to measure
the speed or angle of the rotation from initial references, but without any knowledge of an external,
absolute reference. For this reasons they are used combined with external reference such as star or
sun sensors. Another possible use is a brief periods (such as during blinding condition), for nutation
damping or attitude control during thruster firing.
To provide basic pitch and roll reference a horizon sensor is selected.
58
Ku mission
In order to perform 0,3° sun sensors can be selected.
Sensors configuration:
Fig. 10.9 : Sensors configuration
• Sun Analog Sensors (SAS):
A sun sensor assembly of four 2-axis sun sensor heads of ±64° field of view
provides full coverage around the satellite y-axis.
Parameter
Module level linear FoV
Sensor level FoV
Measurement accuracy
Mass
Value
±64°
128° x 128°
0.3°
0.401 kg
Table 10.8 : SAS characteristics
• Infrared Earth Sensor (IRES):
A 2-axis infrared Earth sensor aligned along the spacecraft z-axis provides roll
and pitch attitude information.
Parameter
Linear FoV range
Earth-presence detection range
Noise (1σ)
Mass
Pitch axes
Roll axes
-8° ÷ 8°
-22° ÷ 22°
-8° ÷ 8°
-22° ÷ 22°
< 0.04°
< 0.04°
4.374 kg
Table 10.9 : IRES characteristics
• Ring Laser Gyros (RLG):
At this preliminary phase ring laser gyros have been selected to assure the top
performance of measurements, but it could be possible to change them with
any other less expensive gyros. This is possible only when simulations have
been done and investigated.
Two 2-axis gyros are provided: yaw axis and skew axis (on S/C X-Z plane).
Parameter
Rate linear field
Rate field with saturation
Measurement noise (1σ)
Yearly drift variation
Mass
59
Value
< 28.8°/s
< 375°/s
0.20 arcsec
< 0.02°/h/year
3.915 kg
Table 10.10 : RLG characteristics
Ka mission
To assure the requirements star sensors has been chosen. They can be quite accurate (<0.01°) but
its is not always possible to take advantage of that feature, because they are usually mounted near
the ends of the vehicle to obtain an unobstructed field of view, so their accuracy can be limited by
structural bending on large spacecraft.
IRES
Fig. 10.10 : Sensors configuration
• Star sensor
One 3-axis star sensor is required and an additional one is added for
redundancy and to reduce blinding period. Therefore one is mounted on + Y
face of the spacecraft and the other one on –Y, in order to avoid that sun rays
enter into star sensor optical head because it is very delicate. Its characteristics:
Parameter
Measurement field
Angular accuracy
Global bias error (3σ)
Mass
60
Value
21° x 31°
16.5 arcsec
< 11 arcsec
4 kg
Table 10.11 : Star sensor characteristics
• Infrared Earth Sensor (IRES)
It has been chosen the same sensor used in Ku mission.
• Ring Laser Gyros (RLG)
It has been chosen the same sensor used in Ku mission.
10.6 Control Mode Architecture
The standard representation of the AOCS control mode architecture can be modeled by ADS in
the following block diagram:
Fig. 10.11: AOCS control mode architecture
61
Chapter 11: TT&C, DH and OBC
11.1 TT&C subsystem
This communication subsystem provides the interface between the satellite and ground systems.
The main functions include the following:
-
Command reception and detection (receive the up-link signal and process it)
Telemetry modulation and transmission. (accept data from spacecraft systems, process them
and transmit them)
Ranging (receive, process and transmit ranging signals to determine the satellite’s position)
The main hardware involved in this subsystem consists of two transponders (one for
redundancy), RF front-ends and two omni-antennas.
The block diagram in Figure 11.1 shows the TT&C subsystems proposed for both missions:
Power
TM
Transmitter
Low-Pass
Filter
Transponder A
TC
Antenna A
Band
Reject
Filter
Diplexer
Gimbal/
Antenna
Control
Elect
Receiver
Transmit
RF switch
TM
Transmitter
Low-Pass
Filter
Transponder B
TC
Antenna B
Band
Reject
Filter
Diplexer
TC TM
Receiver
Low-Pass
Filter
Receive
RF switch
Power
GN&C
Gimbal/
Antenna
Control
Elect
GN&C
Low-Pass
Filter
TC TM
Fig 11.1: TT&C block diagram
From the left side the telemetry data streams enter the transponder where they are modulated
onto carrier output and then amplified. The output signal travels through a low-pass filter which
reduces second and higher-order harmonics, frequency spurs and intermodulation products. Next to
the filter, a transmit RF switch selects one of the two antennas and attenuates frequencies coming
62
from the transmitter and falling in the receiver’s pass band. Finally, a diplexer isolates the
transmitter from the receiver’s port, allowing transmitter and receiver to share the same antenna.
From the right side, the antenna receives the desired signal. The diplexer routs such signal to the
receive RF switch which then selects one the two antennas. A low-pass filter rejects unwanted
transmitter harmonics and frequency spurs; finally, the signal enters the transponder’s receiver
where it is demodulated and sent to the command and data handling subsystem.
In a typical 3-axis-stabilized satellite, omni-antennas are mounted to the top and bottom of the
satellite. All ground-link antennas are mounted to provide an unobstructed view of Earth and place
cross-link antennas to provide an unobstructed view of the relay satellite.
Table 11.1 summarizes the way we can apply a TT&C subsystem. For each application, the table
specifies frequency, modulation, and common antenna characteristics.
Application
Frequency
U/L
D/L
Modulation
U/L
D/L
Antenna
characteristics
Space-Ground
Link Subsystem
S-band
1,75 GHz
S-band
2,2 GHz
FSK
PCM
Earth
coverage
Ground-Space
Tracking and
Data Network
S-band
2.02 GHz
S-band
2,3 GHz
PSK
PSK
Hemispherical
coverage
Legend:
FSK = Frequency Shift Keying
PCM = Pulse Code Modulation
PSK = Phase Shift Keying
Tab 11.1: TT&C attributes
Table 11.2 contains detailed mass, power, and volume characteristics of a common S-band
TT&C subsystem (see Space Mission Analysis and Design – paragraph Telemetry Tracking and
Command).
Component
Transponder
Receiver
Transmitter
Filter/switch
diplexer
Antennas
Hemis
Parabola
Turnstile
Coax cables
TOTAL
Qty Mass [Kg]
(each)
2
6,87
Mass [Kg]
(total)
13,74
Power [W]
Dimensions [cm]
Remarks
14 × 33 × 7
17,5
40
12 W RF output
Solid state power amplifier
1
2
2
0
15 × 30 × 6
2
1
1
1
0,40
9,2
2,3
0,5
0,8
9,2
2,3
0,5
0
0
0
9,5 dia × 13
150 dia × 70
10 dia × 15
1,2 dia × 150
28,54
57,5
1 set
Cicular wave guide
4-dBi gain
Cavity Type
1 set
Tab 11.2: TT&C parameters
63
11.2 Spacecraft Integrated Control Subsystem
The on board computer is designed to handle telecommand and telemetry with the ground
station. The main functions of the on board computer are: to process on board information for
Attitude and Orbit Control Subsystems (AOCS), to do housekeeping of the satellite and its own
automatic unit reconfiguration acting functions redundancies. On board computer in the frame of
Spacecraft Integrated Control Subsystem (SICS) is named Spacecraft Control Unit (SCU).
The SCU is the core of the SICS of this commercial satellite. Through the 1553 bus, SCU
communicates with all the other units involved in the SICS and other subsystems which include:
- Remote Unit A (RU-A) which interface the Unified Propulsion Systems and AOC sensors
and actuators
- Platform Remote Unit B (RU-B) which interface the power and the thermal control units
- Hot redundant telecommand signal interface from TT&C transponder
- Cold redundant telemetry interface and transponder
Telemetry
TT&C
trasponder
SICS
1553 Bus
Sun & Earth
sensors
Telecommand
SCU
AOC remote
terminal
(RU-A)
Platf orm
Remote
Terminal
(RU-B)
Ground Segment
Fig. 11.2: SICS interfaces block diagram
1553 bus contains the information flows of the satellite and is passed through three types of
words. They are:
- Command word
- Data word
- Status word
The first class is transmitted only by the Bus Controller. This word directs a Remote Terminal to
either transmit or receive information across the data bus.
The second class is transmitted by the Bus Controller or a Remote Terminal. This word contains
the actual information that will be transferred from one avionic to another, across the bus.
The third class is transmitted only by a Remote Terminal. This word indicates the general status
of the Remote Terminal. It indicates whether any error conditions were detected in the information
received by the Remote Terminal other general Terminal status conditions.
64
SICS BLOCK DIAGRAM
Infrared
Earth
Sensor
Sun Analog
Sensor
Assembly
Gyro
Assembly
TM/CMD
Reconf.
CMD
H/W Alarm Signals
Spacecraft
Computer Unit
RF IN
(TC)
AOC
Remote
Terminal (RU-A)
Momentum
Wheel
Assembly
1553 Bus
Platform
Remote Terminal
(RU-B)
TM Video
Thrusters
UPS
Termal control
TT&C
Power
Configuration
Switch Drive
Units
TM/PL
TC/PL
To other 1553 Bus User
Fig. 11.3: SICS block diagram
11.2.1
65
Mass Budget
Ku mission
IT. Acron.
1
IRES
2
LDU
3
MIMU
4
MW
5
MWDE
6
RUA
7
RUB
8
SAS
9
SCU
TOTAL
Denomination
Infrared earth sensor
LAE driver Unit
Min. inertial measurement unit
Momentum wheel
Mom. Wheel drive electr.
Remote Unit “A” (AOC)
Remote Unit “B” (PTF)
Sun Analog Sensor
Spacecraft Control Unit
Q.ty
2
1
2
2
2
1
1
4
1
16
Unit [kg]
2,19
1,47
3,92
8,40
2,39
12,35
8,99
0,41
14,44
Total [kg]
4,37
1,47
7,83
16,80
4,78
12,35
8,99
1,64
14,44
72,67
Table 11.3: Ku mission SICS mass budgets
Ka mission
IT.
1
2
3
4
6
7
8
9
Acron.
IRES
LDU
MIMU
RW+RWDE
RUA
RUB
STS
SCU
TOTAL
Denomination
Infrared earth sensor
LAE driver Unit
Min. inertial measurement unit
Reaction wheel + drive electr.
Remote Unit “A” (AOC)
Remote Unit “B” (PTF)
Star Sensor
Spacecraft Control Unit
Q.ty
2
1
2
4
1
1
2
1
14
Unit [kg]
2,19
1,47
3,92
4,2
12,35
8,99
3,8
14,44
Total [kg]
4,37
1,47
7,83
16,80
12,35
8,99
7,6
14,44
73,85
Table 11.4: Ka mission SICS mass budgets
66
Chapter 12: Structure
12.1 Introduction
The analysis of structural feasibility has been performed with MSC_NASTRAN 2.0 for windows
with the main purpose to determine the eigen-values of the S/C. So, all the values reported in this
paragraph are referred to models implemented using MSC_NASTRAN.
The models have been built to evaluate the stress distribution. However they have not been postprocessed in correspondence of composite materials because it is out of scope for this work.
All results obtained by MSC-NASTRAN are consultable inside enclosed CD (See path: “filenastran”).
12.2 Structure Description (Baseline)
The S/C core is built up from a central cylinder that contains two propellant tanks.
The main structural parts are:
-
central cylinder stretching from the launcher interface to the top floor;
external panels connected to the central cylinder by 4 shear panels;
top floor panel and, bottom floor panel, directly connected to the central cylinder, and
connected to the shear panels;
P/L adapter cone connected to the central cylinder and to the bottom floor.
The launcher adapter cone and, the central cylinder, are made of solid aluminum. The shear
walls and the different panels are made of sandwich with carbon fiber skins and aluminum
honeycomb.
The main steps for to define all the nominal sizes are:
-
when we know the propellant mass, we shall define the central cylinder sizes ;
when we know the power that we have to dissipate we shall define the lateral panel sizes;
using the information above we can define all structural sizes.
12.3 Simplifications and assumptions
The distribution of the masses on the structure of the satellite has been made with a general
evaluation that has been aimed to simplify the FEM model of the satellite. In such optics the
internal units and components have been considered as non structural masses distributed on the
panels.
It has been considered that the S/C interfaces with the launcher via two dedicated adapters that
need a diameter of 960mm for satellite in frequency band Ku and diameter of 1200mm for satellite
in frequency band Ka. Ariane 4 user’s manual offers us a Ф937mm interface adapter and a
Ф1194mm interface adapter. So we must change our diameters in order to adapt to standard Ariane
adapters.
67
In order to simulate the launch conditions the satellite has been constrained in correspondence of
all the nodes of lower surface of the s/c adapter.
A static analysis has been performed in order to evaluate the stresses distribution and
displacements of the structure. The following gravity load has been applied in order to simulate the
quasi-static acceleration due to the launcher: X=5g; Y=5g; Z=10g.
The constrains that we have considered are:
-
budget of available mass, (should be around 130Kg structure for the satellite in Ku band and
190Kg structure for the satellite in Ka band);
the first lateral frequency (should be > 15 Hz).
12.4
12.4.1
Solutions for the satellite
Ku mission
• Central cylinder (diameter: 0.86m ; height: 2,32m)
• Box structure (1.4m x1.4m x 2.32m)
Due to the large dimensions of the cylinder, the first sizing has been done considering a
thickness of around 3 mm, (option1). This causes a value of the mass of the structure that would not
to be acceptable. If it is the case we will implement the second solution (option2). In the last case
the thickness of cylinder is 2 mm that means 14 kg less then first option.
Fig. 12.1 : first lateral frequency; on the right bottom XY view (option 1).
Option 1
Inside enclosed CD there is this option with path:
“file-nastran” → “file-nastran-ku” → “soluzione1”
68
The results of the modal analysis, have given a first eigen-value of 21Hz.(see Fig 12.1). This
result is acceptable if compared with our constraint. In a future detailed analysis such margin would
be used to optimize the solution.
The static analysis has been post-processed for the parts in aluminum, (see Fig. 12.2). In Table
12.1 the mass budget detail is reported.
STRUCTURE MASS BUDGET (Option 1 for frequency band: ku)
Item
Mass [kg]
Closure panels (4)
Shear walls (4)
Top Closing Panel
Central Cylinder
Propellant Tank
Support Brackets
Bottom floor
P/L adapter
Inserts and
Miscellaneous
Total
Non structural
Mass with
Material Surface [m2]
[kg/m2]
margin [kg]
mass [kg]
51,48
10,12
7,93
47,75
54,1
10,6
8,3
50,1
sandwich
sandwich
sandwich
aluminium
3
3,2
aluminium
7,92
6,31
8,3
6,6
sandwich
aluminium
5
5,3
139,51
146,5
13,000
2,510
1,962
6,265
460
120
50
800
35,38
47,81
25,48
127,69
1,962
40
20,39
1470
Table 12.1
Fig. 12.2 : Top VonMises Stress on the parts in aluminum
Option 2
“file-nastran” → “file-nastran-ku” → “soluzione2”
In this solution we have used a new structural parts: horizontal shelves. The results of the modal
analysis, have given a first eigen-value of 19Hz .
In Table 12.2 the mass budget detail is reported.
69
STRUCTURE MASS BUDGET (Option 2 for frequency band: ku)
Item
Mass [kg]
Mass with
Material
margin [kg]
Closure panels (4) 51,48
Shear walls (4)
10,12
Top Closing Panel 7,93
Central Cylinder
33,62
Propellant Tank
3
Support Brackets
Bottom floor
7,92
P/L adapter
6,31
Inserts and
5
Miscellaneous
Horizontal shelf
4,5
Total
129,88
54,1
10,6
8,3
35,3
sandwich
sandwich
sandwich
aluminium
3,2
aluminium
8,3
6,6
sandwich
aluminium
Surface [m2]
Non structural
[kg/m2]
mass [kg]
13,000
2,510
1,962
6,265
460
120
50
800
35,38
47,81
25,48
127,69
1,962
40
20,39
5,3
4,7
136,4
sandwich
1470
Table 12.2
12.4.2
Ka mission
• Central cylinder (diameter: 1m ; height: 2,58m)
• Box structure (1.69m x 1.69m x 2.58m)
Due to the large dimensions of the cylinder, the first sizing has been done considering a
thickness of around 3mm. Despite of this, we don’t get over the wanted structural mass budget.
In Table 8.3 the mass budget detail is reported.
STRUCTURE MASS BUDGET (Frequency band: ka)
Item
Mass [kg]
Closure panels (4) 70,48
Shear walls (4)
14,38
Top Closing Panel 11,56
Central Cylinder
63
Propellant Tank
6
Support Brackets
Bottom floor
11,54
P/L adapter
6,91
Inserts and
5
Miscellaneous
Total
188,87
Mass with
Non structural
Material Surface [m2]
[kg/m2]
margin [kg]
mass [kg]
74,0
15,1
12,1
66,2
sandwich
sandwich
sandwich
aluminium
6,3
aluminium
12,1
7,3
sandwich
aluminium
17,500
3,560
2,856
8,047
690
130
100
1200
39,43
36,52
35,01
149,13
2,856
90
31,51
5,3
2210
198,3
Table 12.3
70
The results of the modal analysis, have given a first eigen-value of 17.6 Hz .Some results of the
static analysis are reported in Fig.12.3
“file-nastran” → “file-nastran-ka” → “ka-finito-2”
Fig. 12.3 : Bot VonMises Stress on the parts in aluminium; total translation for the spacecraft.
12.5
Summary
The main results obtained in this chapter are reassumed in the following table.
S/C ku band "solution 2"
Contrain for
S/C ku band
S/C ka band
Contrain for
S/C ka band
130kg
130kg
189kg
190Kg
19Hz
1.69m x 1.69m x 2.58m
>15Hz
-
17,6Hz
1.4m x 1.4m x 2.32m
>15Hz
-
Top Von Mises Stress
98MPa
<120MPa
94MPa
<120MPa
Bot VonMises Stress
99MPa
<120MPa
100MPa
<120MPa
MAIN VALUES
Total structural mass
First eigen value
Box dimensions
Table 12.4
71
Chapter 13: System budgets
The power and mass budgets of the subsystem studied in the previous sections are summarized
in the tables below.
13.1 Mass budget
The mass identified in the system budget is based on the specified values of the individual units
and subsystems. Depending on the maturity status of the items, contingency is applied on unit/item
level. Different margins were already applied to each subsystem in the corresponding chapter:
Option 1
Ku
Subsystem
Payload
AOCS, DH
TT&C
Propulsion
Structure
Thermal
Power
Total Dry Mass
Propellant
Total Wet Mass
Option 2
Ka
% of S/C Tot
Dry Mass
Mass
[kg]
% of S/C Tot
Dry Mass
Mass
[kg]
27 %
10 %
4%
11 %
18 %
8%
23 %
198
73
29
78
130
61
168
32 %
8%
3%
11 %
20 %
7%
18 %
299
74
20
105
189
65
168
-
710
-
902
-
870
-
1100
-
1580
-
2002
Table 13.1: Spacecraft mass budget
In order to compare Mastersat mass characteristics among other large GEO telecommunications
satellites the following table is shown:
It is clear that both option 1 and option 2 are light satellites compared to those of the table 13.2
72
Table 13.2 : Mass distribution of some large GEO telecommunications satellites (Data from
MediaGlobe study, SpaceTech 1989-1999, TopTech studies, TU-Delft)
13.2 Power budget
Mastersat B
(Ku)
Mastersat B
(Ka)
Subsystem
Power [W]
Power [W]
Payload
AOCS, DH,UPS,TT&C
Batteries
BAPTA
Thermal
Power management &
distribution
1987
140
275
37
115
2042
140
275
37
90
150
150
Total Power
2704
2734
Table 13.3 : Spacecraft power budget (worst-case)
Also for the power budget a table with subsystem power loads of other GEO satellites, is
presented in order to compare Mastersat power values:
73
Satellite
Payload
TT&C
AOCS
Thermal
Propulsion
Power
Charging
Total Load
ANIK E
3000.0
42.0
28.0
100.0
?
25.0
287.0
3482.0
Arabsat (not 2)
990.5
38.3
125.1
90.5
?
18.2
99.2
1361.8
Astra 1B
2136.0
43.0
28.0
105.0
?
68.0
410.0
2790.0
DFS Kopernikus
896.0
28.0
39.0
235.0
?
46.0
168.0
1412.0
Fordsat
2461.0
51.3
130.1
92.0
?
41.0
335.8
3109.8
HS 601
2660.0
80.0
70.0
280.0
?
30.0
230.0
3350.0
Intelsat VII
2580.0
38.0
226.0
263.0
6.0
83.0
373.0
3569.0
Intelsat VIIA
3612.0
28.0
226.0
222.0
6.0
53.0
420.0
4567.0
OLYMPUS
2150.0
46.1
116.6
287.0
?
32.5
200.0
2832.2
SATCOM K3
2570.7
42.6
28.3
95.0
?
51.4
362.0
3150.0
TELSTAR 4
4816.5
98.0
76.0
137.0
?
38.0
507.4
5672.9
Table 13.4: Average power distribution (in Watt, EOL) for several large geostationary
telecommunications satellites (Data from MediaGlobe study, SpaceTech 1989-999, TopTech studies, TU-Delft)
Coherently with dry mass analysis, also for the power it is evident that our two options need a
low power budget. Like all the other satellites, this budget depends particularly on the payload that
represents about the 75% of the total load.
74
Chapter 14: Cost analysis
14.1 Elements of analysis
Applying an accurate cost analysis to the space mission under design is becoming more and
more a key factor for a successful operation. Whereas in the past the cost was one of the parameter
to be optimized among many others in order to achieve the performances required, now the trend is
toward a design-to-cost environment, where performance is maximized subject to cost constraints.
The cost analysis approach requires a preliminary development of the cost analysis requirements
description, which identifies the technical and operational parameters (cost drivers). These will be
used in turn as inputs feeding the chosen cost model. In order to categorize and normalize costs, an
organizational table must be drawn for all the phases of the program, defined as Work Breakdown
Structure (WBS). An extensive definition of a WBS is given by the U.S. Department of Defense:
“A work breakdown structure is a product-oriented family tree, composed of hardware, software,
services, data and facilities which results from system engineering efforts during the development
and production of a defense material item, and which completely defines the program. A work
breakdown structure displays and defines the products to be developed or produced and relates the
elements of work to be accomplished to each other and to the end product.”
In order to draw a WBS, it is necessary to describe the space mission in detail through a topdown process, as represented in the following chart (Fig. 14.1):
Work Breakdown Structure
RDT&E
Space Mission Architecture
Program
Level Costs
Management
SE&I
Space
Segment
Option A
Space System
Systems Level
Payload
Spacecraft Bus
Launch
Segment
Ground
Segment
Launch Vehicle
Launch Operations
Facilities
Equipment
Software
Logistics
Management
SE&I
Operations and
Maintenance
Personnel Training
Maintenance
Spares
Mission Operations
Command, Communications
and Control
RDT&E = Research Development Test and Evaluation
O&M = Operations and Maintanance
SE&I = Systems Engineering and Integration
Fig. 14.1: Representative Work Breakdown Structure.
Furthermore, it is possible to expand the single part in a more detailed description. This step was
done only for the space segment and is represented below in Fig. 14.2.
75
Product Tree
Ku Mission
GEO SAT
Space Segment
Structure
Cilinder
Shear panels
Side panels
Closing panels
Misc
Secondary structures
P\L Adapter
GSE Str.
AOCS /
Data Handl.
Thermal
Control
Sensor
Sun Sensor
IRES
Gyro
Actuators
MW
MT
PICS
GSE AOCS
OSR
MLI
Paintings
Heat Pipes
Heaters
Thermistors/
Thermostats
TCU
GSE Therm.
EPS
UPS
SAY
Batt.
MRU
PDU
GSE EPS
Prop. Tank
Press. Tank
RCT
LAE
Valves/Pipes
GSE UPS
PAYLOAD
Antennas
Transponders
WG
GSE P/L
TT & C
Antenna
Transponder
GSE TT&C
Fig. 14.2 : Representative Product tree.
In order to achieve a correct total cost estimate, costs must be evaluated in a detailed bottom-up
process, which requires a further expansion of the organizational chart in lower “branches”. Once
the desired detail is achieved, every “leaf” of the tree will be treated in a Work Package, and a Cost
Sheet will be attached. The described process is represented as an example in the next figures,
where the structure subsystem is first expanded (Fig. 14.3), then its function of Procurement is
analyzed in a Work Package (Table 14.1), and finally the corresponding Cost Sheet is computed
(Table 14.2).
Phase C
Program KU
WP n. 1B1-C
WP title: Procurement of the Structure S/S
Contractor: Master 1
WP responsible: Elisa Di Litta
Start event K.O. + 2
Proposal
Issue:1
Sheet:
End event K.O. + 10
INPUT:
- Structure S/S Specificationss
- Structure S/S Drawings
- Material Catalog
TASKS:
- Supplier Survey
- Preparing documents a supporto ordine
- Supplier Choice
OUTPUT:
- Structure S/S Procurement Specifications
- List of Materials
Table 14.1 : Work Package.
76
Detailed WBS
Ku Mission
Space Segment
Structure
S/S 1
1B1
S/S 2
1B2
S/S 3
1B3
programma
WP
title
proposal
econ. Cond.
company
Ku mission
1B1-C
Procurement S/S Strutture
21-feb-03
Mastersat
LABOUR COST
Procurement
1B1-C
MAIT
1B1-B
hours
management
engineering
manufacturing
Tests
PA
SW
1B1-L
Operations
1B1-D
Project Office
1B1-A
Management
1B1-AA
Engineering
1B1-AD
P.A.
1B1-AC
GSE
1B1-K
30
€
70
60
50
55
62
TOTAL
Tooling
1B1-BB
Testing
1B1-BD
hourly rate
150
250
10,500
15,000
0
0
1,860
27,360
Int. Spec. Fac.
Other Costs
raw materials
mech. Parts
semi lav
electrical components
electronic components
Hi-rel
external services
external major products
travels
transport& assurance
miscellaneous
0
ACQ
MH - 10%
tot
100,000
10,000 110,000
150,000
15,000 165,000
250,000
25,000 275,000
10,000
TOTAL OTHER COSTS
TOTAL COSTS OF WP
profit - 5%
TOTAL PRICE
0
10,000
560,000
587,360
28,868
616,228
Fig. 14.3 : Detailed WBS for the Structure S/S Table 14.2 : Cost Sheet corresponding to WP n. 1B1-C
14.2 Cost estimate
An estimate of the mission costs is summarized in Table 14.3 and 14.4 below. The following
data are intended to be only a rough order of magnitude retrieved from analogous missions and
from preliminary cost model software freely available at the NASA web site, but not the outputs of
a systematic cost analysis. Despite their inaccuracy, yet they underline important aspects of the
study carried out, and give a clear starting point for a further analysis.
In a first phase (before breakeven), the tariff will be such to reach the breakeven goal of 5 years,
given the available service to offer. After the breakeven, when the initial investment is paid back, it
will be possible to reduce the commercial tariff, increasing the number of users.
77
Ku MISSION
Spacecraft design lifetime
12 years
*5 years
COSTS BEFORE BREAKEVEN*
Satellite
Launch
Insurance
Non recurring
Stations
Total
50
40
13.5
10
15
128.5
Average cost of money
Maintenance
Running
Total yearly
Total cost at breakeven
Yearly cost at breakeven
3.86
0.75
0.5
5.11
M€
M€
M€
M€
M€
M€
M€/yr
M€/yr
M€/yr
M€/yr
including Program
15%
including LEOP
3% per year
154.03 M€
30.81 M€/yr
AVAILABLE SERVICE
Channels
24
Trasponders
24
Filling factor
0.7
Availability
0.99
Working Hours per year
8760
Reliability
0.95
Product
3,321,876
Commercial tariff to breakeven
9.27 €/hr/ch
AFTER BREAKEVEN UP TO EOL
Total operational costs
1.25 M€/yr
Assumed tariff
1.5 €/hr
Revenues
4.98 M€/yr
Net profit
reduced after breakeven
3.73 M€/yr
Table 14.3 : Mission cost and profit for the Ku case
To compare these tariffs with present cost range for transponders on board commercial
geostationary satellites, they have to be expressed in M€/year/trasponder:
Commercial tariff before breakeven
Commercial tariff after breakeven
1.28 M€/year/trasponder
Table 14.4 : Commercial tariffs for Mastersat Ku case
Present cost range for transponders on board commercial geostationary satellites:
o costs given in M€ per 40 MHz transponder
o data are typical and depend on specific contract conditions
o there are leasing cost differences between satellite owners, and between geographical areas
o partial transponder lease costs are 20 to 40% higher
78
To have an order of magnitude of present commercial tariffs for 72 Mhz transponder it has been
used a 1.5 factor of tariffs for 40 Mhz.
Band
Ku
[40 Mhz transponders]
Ku
[72 Mhz transponders]
U.S.
Latin
America
Asia
Europe
1.4-2
2.1 – 3.7
3.2 – 6.0
1.8 – 8.0
M€/year/trasponder
2.1-3.0
3.2 – 5.6
4.8 – 9.0 2.7 - 12.0
Table 14.5 : Present commercial tariffs for geostationary satellites
It is easy to notice that Mastersat tariffs are less expensive than present ones:
before breakeven
after breakeven
U.S.
39 %
75 %
Europe
29 %
81 %
Table 14.6 : Percent of Mastersat tariffs compared by present ones
Pushing the designer to find a more challenging solution, it has increased the available number
of channels, shifted the transmission to a higher band and had as a counterpart higher costs in terms
of mass and technology used, which would imply a bigger monetary effort for its development. The
second option leads to a net profit of 9.89 M€ per year, with a commercial cost of 1.93 €/hour,
assuming to allocate the 60% of all the available channels, whereas for the Ku-band a filling factor
has been considered.
79
Ka MISSION
Spacecraft design lifetime
10 years
*5 years
COSTS BEFORE BREAKEVEN*
Satellite
Launch
Insurance
Non recurring
Stations
Total
70
60
19.5
15
20
184.5
Average cost of money
Maintenance
Running
Total yearly
Total cost at breakeven
Yearly cost at breakeven
5.54
1
0.5
7.04
M€
M€
M€
M€
M€
M€
M€/yr
M€/yr
M€/yr
M€/yr
including Program
15%
including LEOP
3% per year
219.68 M€
43.94 M€/yr
AVAILABLE SERVICE
Channels
48
Trasponders
96
Filling factor
0.6
Availability
0.99
Working Hours per year
8760
Reliability
0.95
Product
22,778,579
1.93 €/hr/ch
AFTER BREAKEVEN UP TO EOL
Total operational costs
1.5 M€/yr
Assumed tariff
0.5 €/hr
Revenues
11.39 M€/yr
Net profit
reduced after breakeven
9.89 M€/yr
Table 14.7 : Mission cost and profit for the Ka case
Commercial tariff after breakeven
Table 14.8 : Commercial tariffs for Mastersat Ku case
To
Legend
PRR
To+13
Phase E: Launch
Phase D: Production
Phase C: Systemdesign and development
Phase A-B: Pre-feasibility studies and base-line
Hi-Rel QM
QMManufacturing & Test
Component procurement
Final design
Feasibilty studies
To+3
PDR
CDR
PDR
PRR
CDR
Bus & P/L A.I.T
To+28
Critical Design Review
Preliminary Design Review
Preliminary Requirements Review
FMManufacturing & Test
To+20
Satellite A.I.T
To+34
Launch
To+38
To+40
80
Chapter 15: Planning
The followed bar chart summaries Mastersat B hypothetic planning:
Fig. 14.3 : Mastersat B hypothetic planning
81
Concluding remarks
Finding new quotas in the market of telecommunications services is hard to achieve within the
current competency and perspectives. Yet, it is even harder when a provider is not able to offer a
service at lower price. Nowadays it would be impossible to conceive a mission without taking into
account costs and profits. Having this statement in mind, an effort to lower service costs has been
done, with the promising results obtained in the previous section. Both Mastersat options have
reached the main goal. Of course, a more accurate cost analysis should be performed to the
proposed solution, but the rough order of magnitude that has been given shows the economical
feasibility of the project and should raise the interest in investigating it in more detail.
Previous considerations should be completed by a market needs evaluation, which should be
instead the starting point anytime a new project is thought. In fact, the possibility that the Ka
mission has to allocate such a number of channels does not forecast whether those resources would
be used or not. Does the market need what we offer? Current trends show a certain tendency in this
direction, but we are not able to be sure at this stage, leaving the discussion open for a further study.
Mastersat B Working Team
82
References
[1] AA. VV., Space Mission Analysis and Design - 3rd edition, J. R. Wertz & W. J. Larson,
Dordrecht (NL) 1999.
[2] AA. VV., Spacecraft Systems Engineering - 2nd edition, P. Fortescue & J. Stark,
[3] AA. VV., Fundamentals of Space Systems – 1st edition, L. V. Pisacane & R. C. Moore,
[4] AA. VV., Comunicazioni Spaziali -A. Gilardini,
[5] AA. VV., Course Lectures, Master in Satelliti e Piattaforme Orbitanti, Roma 2002/2003.
[6] AA. VV., SUPAERO Course Lectures, Guidage et Pilotage des Satellites, Toulouse 2001/2002.
[7] AA. VV., SUPAERO Course Lectures, Ingénierie Satellite, Toulouse 2001/2002.
[8] AA. VV., SUPAERO Course Lectures, Dynamique et Stabilisation d’ Attitude des Satellites,
Toulouse 2001/2002.
[9] AA. VV., SUPAERO Course Lectures, Guidage et Pilotage des Satellites, Toulouse 2001/2002.
[10] AA. VV., “ADS Software User’s Manual”, ESA, ADS.MA.AER.001, 10/09/98.
[11] MSC/NASTRAN handbook for linear analysis. MSC-NASTRAN version 64 (The macnealschwendler corporation.)
[12] Ariane4 user’s manual. Arianespace 1999

Mastersat B: Mission and Analysis Design - Sapienza

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