More Related Content Similar to Space Systems & Space Subsystems Fundamentals Technical Training Course Sampler (20) More from Jim Jenkins (20) Space Systems & Space Subsystems Fundamentals Technical Training Course Sampler1. 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
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
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
i1
– 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
i1
– 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
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
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
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
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
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 1023 1 106
400
s
©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