This four-day course in space systems and space subsystems is for technical and management personnel who wish to gain an understanding of the important technical concepts in the development of space instrumentation, subsystems, and systems. The goal is to assist students to achieve their professional potential by endowing them with an understanding of the subsystems and supporting disciplines important to developing space instrumentation, space subsystems, and space systems. It designed for participants who expect to plan, design, build, integrate, test, launch, operate or manage subsystems, space systems, launch vehicles, spacecraft, payloads, or ground systems. The objective is to expose each participant to the fundamentals of each subsystem and their inter-relations, to not necessarily make each student a systems engineer, but to give aerospace engineers and managers a technically based space systems perspective. The fundamental concepts are introduced and illustrated by state-of-the-art examples. This course differs from the typical space systems course in that the technical aspects of each important subsystem are addressed.

1. SPACE SYSTEMS AND SPACE SUBSYSTEMS - FUNDAMENTALS
Instructor:
Dr. Vincent L. Pisacane
Course Schedule: http://www.ATIcourses.com/schedule.htm
Course Outline:
http://www.aticourses.com/Fundamentals_Of_Space_Systems_Space_Subsytems.htm

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3. CASSINI-HUYGENS
Interplanetary Mission to Saturn
• Saturn surrounded by Rings and 62 Moons
• Cassini launched in October 1997 arrived at Saturn June 2004
• The mission has been extended through September 2017
©Pisacane, 2013

4. CASSINI-HUYGENS
Trajectory
Planned 21 April
Planned 1 July
Planned 20 June
Planned 1 Dec
Planned 6 Oct
Planned 16 August
@ 1,170 km
Planned 30 Dec 2000
©Pisacane, 2013

5. NEAR
Configurations
©Pisacane, 2013

6. RISK MANAGEMENT
NASA’s Approach to Risk Management
 NASA identifies two activities critical to risk management
 Risk-Informed Decision Making (RIDM)

– Selection of alternatives based on assessment of requirements including risk
Continuous Risk Management (CRM)
– Systematic identification, assessment, and management of all risks
From: NASA Risk-Informed Decision Making
Handbook, NASA/SP-2010-576 Version 1.0 Apr 2010
©Pisacane, 2013

7. SYSTEM DEVELOPMENT
NASA Project Life Cycle Reviews
©Pisacane, 2013

8. SYSTEM TESTING
Sample NASA Payload Test Requirements
From: NASA-STD-7002A ─
Payload Test Requirements
©Pisacane, 2013

9. SPACECRAFT FAILURES
NOAA Spacecraft Radiation Induced Failures May 1998
 Data
from NOAA GOES (Geostationary Operational
Environmental Satellite) constellation
 Equator-S
failure attributed to latch-up in central
processor as result of a week or more of elevated
relativistic electron (top figure)
 POLAR processor loss of 6 hours of data attributed to
single-event upset (SEU) in processor from increased
proton flux (bottom figure)
 Galaxy
4 processor failure likely caused, by the
energetic electron environment most likely due to
deep dielectric, (or bulk) charging (top figure)
Space Environmental Conditions During April and May 1998: An Indicator
for the Upcoming Solar Maximum
D.N. Baker, J.H. Allen, S. G. Kanekal, and G.D. Reeves
©Pisacane, 2013

10. FAILURE ANALYSES
Burn-in Tests at Elevated Temperatures
 The standard life test for flight hardware parts is the dynamic (power on) burn-in test
o
o

for 1000 hours (41.7 d) at an ambient temperature of 125 C (257 F)
The Acceleration Factor (Af) is the test time multiplier derived from the Arrhenius
equation for operation at another temperature
E  1
1 
A f  1 if Tuse  Ttest
A f  exp a 


A f  1 if Tuse  Ttest
k  Tuse Ttest 

 
 Activation energy (Ea) is an empirical value of the minimum energy required to initiate a


specific type of failure mode that can occur within a technology type
– Failure modes include: oxide defects, bulk silicon defects, mask defects, electromigration, and contamination
Typical values of Ea for electronic devices
Acceleration Factors
are 0.5-1.0 eV, typically > 0.7
Equivalent Duration, y
Ea, For use temperatures
Table shows acceleration factors and
eV 25oC 35oC
45oC
25oC
35oC
45oC
equivalent durations
Parameters
Ea = Activation Energy of the failure
mode, eV
k = Boltzmann's Constant, 8.617 x
10-5 eV K-1
Tuse = Use Temperature, K
Ttest = Test Temperature, K
77oF
95oF
113oF
77oF
95oF
113oF
0.5
133
71
39
15
8
2
0.6
353
165
81
40
19
9
0.7
938
387
169
107
44
19
0.8
2,492
907
352
284
103
40
0.9
6,624 2,125
732
756
242
84
1,524
2,008
568
174
1.0 17,607 4,979
©Pisacane, 2013

11. FAILURE IDENTIFICATION
Sample FMEA Worksheet Failure Modes and Effects Analysis (FMEA)
 Typical FMEA worksheet is illustrated below for a spacecraft battery
Failure Modes, and Effects Analysis (FMEA)
System:
Part Name
Reference Drawing
Mission
Date
Sheet X of X
Compiled by: XXXX
Approved by: XXXX
Potential Effects of Failure Mode
Item
Function
or
Requirement
Potential
Failure
Modes
Potential
Causes
of
Failure
Mode
Battery
Provide
adequate
relay
voltage
Fails to
provide
adequate
power
Voltage
drops to
zero
Local
Effects
Battery
plates
shorted
Intermediate
Effects
Instrument
not
functional
End
Effects
Mission
Aborted
Detection
and
Mitigating
Factors
Test
battery
prior to
launch
O
c
c
u
r
r
e
n
c
e
4
D
e
t
e
c
t
i
o
n
4
S
e
v
e
r
i
t
y
0.5
+
0.3
X5
=
4
RPN
Actions
Recommendations
Responsibility
64
XXX
XXX
…
©Pisacane, 2013

12. RELIABILITY, AVAILABILITY, MAINTAINABILITY, and SAFETY
Derating Introduction
 Derating increases the margin of safety between operating stress level and actual
failure level for the part, providing added protection from unanticipated anomalies
 Derating is employed in electrical and electronic devices, wherein the device is
operated at lower than its rated maximum power dissipation, taking into account
– Case/body temperature
– Ambient temperature
– Type of cooling mechanism
 When derating, the application engineer applies a recommended derating factor
bases on the part specifications and operating environment
 For microcircuits, major derating factors are
–
–
–
–
–
Supply voltage
Power dissipation
Signal input voltages
Output voltages
Output currents
©Pisacane, 2013

13. RELIABILITY, AVAILABILITY, MAINTAINABILITY, and SAFETY
Calculating Reliabilities
 Series redundancy
– Reliability Rs of the series chain is given
by
n
Rs  Ri  R1R2R3 Rn
i1
– If all components have the same
reliability then Ri = R and
Rs  Rn
 Parallel redundancy
– The reliability of a parallel configuration
if only one device is needed is
Rs  1   1  Ri  1  1  R1 1  R2 1  Rn 
n
i1
– If all component s have the same
reliability then Ri = R and
Rs  1  1  R
n
©Pisacane, 2013

14. CELESTIAL MOTION
Principal Motion of the Celestial Ephemeris Pole
(more accurate number is 25,780 yrs)
(average of 50.26 sec of arc per year
or 0.1376 sec arc per day)
©Pisacane, 2013

15. COORDINATED UNIVERSAL TIME (UTC)
Variation in the Length of Day 2/2
25
From: http://www.ucolick.org/~sla/leapsecs/dutc.html
©Pisacane, 2013

16. REFERENCE SYSTEM
Geometrical Transformation Between GCRS and ITRS
 Figure shows transformation between terrestrial (ITRS) to celestial (GCRS) taking into
account (1) Pole Movement, (2) Earth Rotation , (3) Precession and Nutation
–
–
–
–
GCRS= Geocentric Celestial Reference System
ITRS = International Terrestrial Reference System
CIP = Celestial Intermediate Pole, instantaneous Earth spin axis
CTP = Conventional Terrestrial Pole, reference pole in ITRS (now average of pole positions from 1900 to 1905)
Modifiedfrom:ESA,http://navipedia.org/index.php/Transformation_bet
ween_Celestial_and_Terrestrial_Frames
©Pisacane, 2013

17. GRAVITATIONAL POTENTIAL
Geometrical Representation of Spherical Harmonics
 Pn,m(Cos q) Cos m(l – l n,m) has
− (n–m) sign changes or zeros 0  q  p (latitude of 180 degrees
− 2m zeros in interval 0  l < 2p (longitude of 180 degrees)
m=0
no longitudinal
variation
n = 2, m = 2
n ≠ m and m ≠ 0
Tessarae (Tiles)
n = 3, m = 3
n=m
no latitudinal
variation
n = 5, m = 0
n = 4, m =3
©Pisacane, 2013

18. TRAJECTORY PERTURBATIONS
Mars Global Surveyor Aerodynamic Braking
©Pisacane, 2013

19. ROCKET PROPULSION
Specific Impulse vs Thrust
Grayed area are
realized
characteristics
NH3 = Ammonia
N2H4 = Hydrazine
From: http://dawn.jpl.nasa.gov/mission/images/CR-1845.gif
©Pisacane, 2013

20. ROCKET PROPULSION
de Laval Nozzle
 The function of the nozzle is to convert the
chemical-thermal energy produced in the
combustion chamber into kinetic energy
 Thrust is the product of mass time velocity so a
very high gas velocity is desirable
 The nozzle converts slow moving, high pressure,
and high temperature gas in the combustion
chamber into high velocity gas of lower pressure
and temperature at the nozzle’s exit
 De Laval nozzles consist of a convergent and
divergent section
 The section with minimum area is the nozzle throat
 The nozzle is usually made long enough and the
exit area large enough to reduce the high pressure
in the combustion chamber to the ambient
pressure at the nozzle exit to create maximum
thrust
Typical DeLaval nozzle
T = temperature
p = pressures
v = speed
M = Mach number
From: http://en.wikipedia.org/wiki/Rocket_engine
©Pisacane, 2013

21. LAUNCH FLIGHT MECHANICS
Available Launch Inclinations in the United States
37
114
©Pisacane, 2013

22. COLD GAS PROPULSION SYSTEMS
Typical Cold Gas System Implementation
Service valve
Gas Regulator
P
GN2
T
Filter
F
Burst Valve
L
L
Access Port
NO Pyrovalve
normally open
F
NC
L
T
P
P
P
L
T
L
T
Latch Valve
T
Latch Valve
Check Valve, arrow
direction of flow
P
L
Pyrovalve
normally closed
L
P
Pressure Sensor
T
Temperature sensor
Typical cold gas thruster
Propellants
Air, Carbon Dioxide,
Helium, Hydrogen,
Methane, Nitrogen, Freon
©Pisacane, 2013

23. LIQUID PROPULSION SYSTEMS
Messenger Spacecraft Dual Mode Propulsion
 Illustrates the Messenger spacecraft
propulsion system with 17 thrusters
 Bipropellants ─ Hydrazine (N2H4)
and Dinitrogen Tetroxide (N2O4)
 Monopropellant ─ Hydrazine (N2H4)
S Wiley, K Dommer, L Mosher, Design and development of the
Messenger propulsion system, AIAA, PRA-053-03-14 July 2003
©Pisacane, 2013

24. TRANSFER TRAJECTORIES
Apollo 13 Circumlunar Free-Return Trajectory
CSM ─ Command Service Module,
DPS – Descent Propulsion System
EI – Entry Interface
GET ─Ground Elapse Time
LM – Lunar Module
MCC ─ Mid-Course Correction
PC – Pericynthion (closest point to moon)
S-IV4B – Saturn IVB
SM – Service Module
TLI –Trans Lunar Injection
JL Goodman , Apollo 13 Guidance, Navigation, and Control Challenges AIAA
SPACE 2009 Conference & Exposition, Sept 2009, Pasadena,, AIAA 2009-6455
©Pisacane, 2013

25. OVERVIEW
Attitude Control Schematic
©Pisacane, 2013

26. ATTITUDE KINEMATICS
Quaternion Mathematics 1/2
 Addition and subtraction
– Elements are added or subtracted
Q 1  Q 2  q1,1i  q1,2 j  q1,3k  q1,4   q2 ,1i  q2 ,2 j  q2 ,3k  q2 ,4 
 q1,1  q2 ,1 i  q1,2  q2 ,2  j  q1,3  q2 ,3 k  q1,4  q2 ,4 
 Multiplication
– Not communicative, Q1Q2 ≠ Q2Q1
– Multiple each component
Q 1Q 2  q1,1i  q1,2 j  q1,3k  q1,4 q2 ,1i  q2 ,2 j  q2 ,3k  q2 ,4 
where
time
s
1
i
j
k
1
1
i
j
k
i
i
─
1
k
─j
j
J
─
k
─1
i
k
k
j
─i
─1
 Equivalent quaternions
– Reversing signs on all 4 elements yields an equivalent quaternion
─Q = Q
©Pisacane, 2013

27. ATTITUDE SENSORS
ADCOL Two-Axis Digital Sun Sensor System
 Two-Axis Digital Sun Sensor System
– No of measurement axes:
• 2 each sensor)
– Number of sensors
• 5 typical per electronics
• 1 to 8 sensors can also be used
• Electronics selects sensor that has
sun in field of view
 Heritage
– Many systems flown with 1 to 8 sensor
heads per processing electronics
 Parameters
– Field of view: ±64° x ±64°
• Note: 4π steradians (full sphere)
coverage can be achieved with 5
sensors.
– Accuracy: ±0.25° (transition accuracy).
– Least Significant Bit Size: 0.5°
http://adcole.com/two-axis-dss.html
Sign bit
Most significant bit
Least significant bit
Interpolating bits
©Pisacane, 2013

28. INTRODUCTION
Function and Components of Spacecraft Power System
 Power system functions
–
–
–
–
–
–
Supply electrical power to spacecraft loads
Distribute and regulate electrical power
Satisfy average and peak power demands
Condition and convert voltages
Provide energy storage for eclipse and peak demands
Provide power for specific functions, e.g., firing ordinance for mechanism
deployment
– Ensure power to critical loads during critical phases and spacecraft anomalies
– Ensure power for mission duration
Primary Power
Source
Energy
Conversion
Power
Regulation
Power
Distribution
Power
Regulation
?
Critical
Loads
Non-Critical
Loads
Energy
Storage
Power
Regulation
©Pisacane, 2013

29. SECONDARY BATTERIES
Candidate Technologies
http://www.clyde-space.com/products/spacecraft_batteries/useful_info_about_batteries/secondary_batteries
©Pisacane, 2013

30. SOLAR ARRAYS
Solar Array Construction
 Cells connected in series to achieve


desired voltage
Cells connected in parallel to achieve
desired power
Arrays organized to minimize current
loops that result in dipole moment
©Pisacane, 2013

31. OVERVIEW
NEAR Spacecraft Spacecraft Communication System
From: RS Bokulic, MKE Flaherty, JR
Jensen, and TR McKnight, The NEAR
Spacecraft RF Telecommunications System,
Johns Hopkins APL Technical Digest, Vol
19, No 2 (1998)
Transponder unifies a number of communication functions - receiver,
command detector, telemetry modulator, exciters, beacon tone
generator, and control functions
Diplexer is a device that can split and combine audio and video
signals
©Pisacane, 2013

32. ANTENNAS
Typical Parabolic Antenna Pattern
©Pisacane, 2013

33. LINK ANALYSIS
Example Link Analysis
Transmitter power
20 W
+13.0 dBW
Spacecraft cable loss
1dB
─1 dB
Antenna boresight
gain
76.76
+18.9 dB
EIRP
Spacecraft antenna diameter = 1 m
Frequency = 1 GHz
Pointing error= 1/10 beamwidth
Receiver gain = 30 dB
Receiver system temperature = 400K
Bit rate = 106 bps
30.9 dBW
Antenna beamwidth
3 dB
─3.0 dB
Space loss at 10o
elevation @ 3000 km
1.58 x 1016
─162.0 dB
Pointing error, 0.1 BW
0.12 dB
─0.12 dB
Atmospheric loss
0.1 dB
─0.2 dB
K-1
Receiver G/T
1000/400
Boltzmann constant,
k
1.38x10-23
JK-1
+228.6 dB J1K
Bit rate
106 bps
2
2
6
9
 4 prf    4 p  3  10  1  10   1.58  1016
Ls  

 
3  108
 c  

2
l
 
12
12 0.1 q3dB 
Ll q  2  q2 dB 
 0.12 dB
i
2
i
q3dB
q3dB
2
─60 dB s
Receiver Eb/No
4.0
dbK-1
2
9
 πDf   0.70 π  1  1x10   76.76
Gboresight  



3x108 
 c 

38.2dB
1
a
a
Eb EIRP L s  L Other Losses  GRA  20  0.8  76.76   0.5  1.58  1016   0.97  0.63  1000 
 


  683  38.3dB
 T 
N0
kRb
1.38  1023  1  106
 400 
 s 
©Pisacane, 2013

34. THERMAL ANALYSES
Analysis Process
©Pisacane, 2013

35. MULTILAYER INSULATION
Gold and Black MLI
 Gold Thermal Blanket
− Outer layer is of a second surface mirror material with
high reflectivity and high emittance
− Consists of multiple layers of silver coated Kapton film
that gives it a gold color
− Except outer layers, all are perforated to allow entrapped
air to escape during launch and separated by a Dacron netting
− Edges are finished with a tape prior to sewing
− Individual blankets held together and to spacecraft by
dacron Velcro
 Black Thermal Blanket
− Black thermal blanket is used on the shade side of the
spacecraft
− Identical to the gold blanket except for the outer layer
generally Kapton filled with carbon powder
− Outer layer has a higher absorptance and lower
emittance than the gold Kapton
− This layer is also electrically conductive because of
carbon fill
− Grounding outer layer to the spacecraft frame dissipates
any charge build
Gold is multilayer insulation of
Cassini spacecraft; from
NASA
New Horizons spacecraft
http://www.boulder.swri.edu/pkb/ssr/ssrfountain.pdf
©Pisacane, 2013

36. DESIGN PROCESS
Overall Development Flow Chart
start
Launch Vehicle
Constraints
Conceptual and
Preliminary
Design
Critical Design
Fabrication
Integration
launch
Preliminary Launch
Loads
Launch Vehicle
Dynamic Model and
Forcing Functions
Spacecraft
Dynamic
Model
Coupled Launch
Vehicle and
Spacecraft Dynamic
Analysis
Spacecraft Dynamic
Response
Loads
Acceleration
Spacecraft Structural
Configuration
Preliminary
Spacecraft Structural
Design
Functional
Subsystem/Payloads
Requirements
Preliminary Natural
Frequency
Constraints
Thermal Analysis
Structural Analysis
Finite Element Model
Dynamic Analysis
Stress Analysis
Thermal Distortion
Assess Margins
Temperature
Distribution
Fabricate Spacecraft
Structure
Spin Balance and
Environmental
Testing
©Pisacane, 2013

37. STRUCTURAL CONFIGURATIONS
Structural Categories
 Structural
components are categorized by the different types of requirements,
environments, and methods of verification that drive their design
– Primary structures are usually designed to survive steady-state accelerations and
transient loading during launch and for stiffness
– Secondary and tertiary structures are usually designed for stiffness, positional
stability, and fatigue life
Secondary structures:
Primary structures:
• body structure
• launch vehicle adapter
•
•
•
•
•
•
appendage booms
support trusses
platforms
solar panels
Antenna
Extendibles
Tertiary structures:
• brackets
• electronics boxes
©Pisacane, 2013

38. INTRODUCTION
Space and Ground Based Systems
 Reliability, complexity, development costs, and operational costs are affected by
the partitioning of the computational load between the space and ground segment
From Wertz and Larson
©Pisacane, 2013

39. COMPUTER COMPONENTS
Typical Spacecraft Computer Schematic
 Figure is a simplified block diagram of a spacecraft computer system
 One or more processing units have access through bus structures to
–
–
–
–
–
Read only memory, random access memory, and special purpose memory
Mass storage
Input/output ports to spacecraft subsystems and payloads
Spacecraft communication system
Numerical coprocessor to carry out floating point arithmetic faster
From Pisacane, Fundamentals of space systems, Oxford University Press,
2005
©Pisacane, 2013

40. FAULT TOLERANCE
Summary Fault Tolerant Techniques
NMR = n-modular redundancy
ECC = Error Correction Coding
RESO = RE-computing with Shifted Operands; computation carried
out twice - once with usual input ─ once with shifted operands
Self-purging = each module has a capability to remove itself from
the system if faulty
Recovery blocks = Uses the concept of retrying the same
operation and expect the problem is resolved by the second
or later tries
©Pisacane, 2013

41. SPACECRAFT PROCESSORS
RAD6000 Processor
 Characteristics
−
–
–
–
–
–
–
–
–
35 Mbps at 33 MHz
Radiation Hardened 32-bit RISC
Super Scalar Single Chip CPU
8K Byte Internal Cache
Simplex or Dual Lock-step (compares CPU
operations)
Low Power 3.3 Volt Operation
72-bit (64 Data, 8 ECC) Memory Bus
Variable Power/Performance
Independent Fixed and Floating Point Units
 Radiation Hardness6 Levels
–
–
–
–
–
–
Total Dose: 2x10 rads(Si)
Prompt Dose Upset: 1x109 rads(Si)/sec
Survivability: 1x1012 rads(Si)/sec
Single Event Upset: 1x10-10 Upsets/Bit-Day
Neutron Fluence: 1x1014 N/cm
Device Latchup: Immune
From Lockheed Martin Federal Systems RAD6000 Radiation Hardened 32-Bit
Processor
atc2.aut.uah.es/~mprieto/asignaturas/satelites/pdf/rad6000.pdf
COP = Common on-chip processor interface
FPGA = Field Programmable Gate Array
HMC = Hardware Management Console
RS232 = Serial binary single ended data connector
VME bus = VersaModular Eurocard bus
Dual Lock Step
A technique that achieves high
reliability by adding a second
identical processor that monitors and
verifies the operation of the system
processor
©Pisacane, 2013

42. INTEGRATION AND TEST PROCEDURES
Integration and Test Procedure
From Spacecraft Computer Systems, JE Keesee
ocw.mit.edu/courses/aeronautics-and.../l19scraftcompsys.pdf
©Pisacane, 2013