0
SPACE SYSTEMS AND SPACE SUBSYSTEMS - FUNDAMENTALS

Instructor:
Dr. Vincent L. Pisacane

Course Schedule: http://www.ATIcou...
www.ATIcourses.com
Boost Your Skills
with On-Site Courses
Tailored to Your Needs

349 Berkshire Drive
Riva, Maryland 21140...
CASSINI-HUYGENS
Interplanetary Mission to Saturn

• Saturn surrounded by Rings and 62 Moons
• Cassini launched in October ...
CASSINI-HUYGENS
Trajectory

Planned 21 April
Planned 1 July
Planned 20 June

Planned 1 Dec

Planned 6 Oct
Planned 16 Augus...
NEAR
Configurations

©Pisacane, 2013
RISK MANAGEMENT
NASA’s Approach to Risk Management

 NASA identifies two activities critical to risk management
 Risk-In...
SYSTEM DEVELOPMENT

NASA Project Life Cycle Reviews

©Pisacane, 2013
SYSTEM TESTING
Sample NASA Payload Test Requirements

From: NASA-STD-7002A ─
Payload Test Requirements

©Pisacane, 2013
SPACECRAFT FAILURES
NOAA Spacecraft Radiation Induced Failures May 1998

 Data

from NOAA GOES (Geostationary Operational...
FAILURE ANALYSES
Burn-in Tests at Elevated Temperatures

 The standard life test for flight hardware parts is the dynamic...
FAILURE IDENTIFICATION
Sample FMEA Worksheet Failure Modes and Effects Analysis (FMEA)

 Typical FMEA worksheet is illust...
RELIABILITY, AVAILABILITY, MAINTAINABILITY, and SAFETY
Derating Introduction

 Derating increases the margin of safety be...
RELIABILITY, AVAILABILITY, MAINTAINABILITY, and SAFETY
Calculating Reliabilities

 Series redundancy

– Reliability Rs of...
CELESTIAL MOTION
Principal Motion of the Celestial Ephemeris Pole

(more accurate number is 25,780 yrs)

(average of 50.26...
COORDINATED UNIVERSAL TIME (UTC)
Variation in the Length of Day 2/2

25

From: http://www.ucolick.org/~sla/leapsecs/dutc.h...
REFERENCE SYSTEM
Geometrical Transformation Between GCRS and ITRS

 Figure shows transformation between terrestrial (ITRS...
GRAVITATIONAL POTENTIAL
Geometrical Representation of Spherical Harmonics
 Pn,m(Cos q) Cos m(l – l n,m) has
− (n–m) sign ...
TRAJECTORY PERTURBATIONS
Mars Global Surveyor Aerodynamic Braking

©Pisacane, 2013
ROCKET PROPULSION
Specific Impulse vs Thrust

Grayed area are
realized
characteristics

NH3 = Ammonia
N2H4 = Hydrazine

Fr...
ROCKET PROPULSION
de Laval Nozzle

 The function of the nozzle is to convert the
chemical-thermal energy produced in the
...
LAUNCH FLIGHT MECHANICS
Available Launch Inclinations in the United States

37

114

©Pisacane, 2013
COLD GAS PROPULSION SYSTEMS
Typical Cold Gas System Implementation
Service valve

Gas Regulator
P

GN2

T

Filter

F

Burs...
LIQUID PROPULSION SYSTEMS
Messenger Spacecraft Dual Mode Propulsion
 Illustrates the Messenger spacecraft
propulsion syst...
TRANSFER TRAJECTORIES
Apollo 13 Circumlunar Free-Return Trajectory

CSM ─ Command Service Module,
DPS – Descent Propulsion...
OVERVIEW
Attitude Control Schematic

©Pisacane, 2013
ATTITUDE KINEMATICS
Quaternion Mathematics 1/2

 Addition and subtraction

– Elements are added or subtracted

Q 1  Q 2 ...
ATTITUDE SENSORS
ADCOL Two-Axis Digital Sun Sensor System

 Two-Axis Digital Sun Sensor System

– No of measurement axes:...
INTRODUCTION
Function and Components of Spacecraft Power System

 Power system functions
–
–
–
–
–
–

Supply electrical p...
SECONDARY BATTERIES
Candidate Technologies

http://www.clyde-space.com/products/spacecraft_batteries/useful_info_about_bat...
SOLAR ARRAYS
Solar Array Construction

 Cells connected in series to achieve



desired voltage
Cells connected in para...
OVERVIEW
NEAR Spacecraft Spacecraft Communication System

From: RS Bokulic, MKE Flaherty, JR
Jensen, and TR McKnight, The ...
ANTENNAS
Typical Parabolic Antenna Pattern

©Pisacane, 2013
LINK ANALYSIS
Example Link Analysis
Transmitter power

20 W

+13.0 dBW

Spacecraft cable loss

1dB

─1 dB

Antenna boresig...
THERMAL ANALYSES
Analysis Process

©Pisacane, 2013
MULTILAYER INSULATION
Gold and Black MLI

 Gold Thermal Blanket

− Outer layer is of a second surface mirror material wit...
DESIGN PROCESS
Overall Development Flow Chart
start
Launch Vehicle
Constraints

Conceptual and
Preliminary
Design

Critica...
STRUCTURAL CONFIGURATIONS
Structural Categories

 Structural

components are categorized by the different types of requir...
INTRODUCTION
Space and Ground Based Systems

 Reliability, complexity, development costs, and operational costs are affec...
COMPUTER COMPONENTS
Typical Spacecraft Computer Schematic

 Figure is a simplified block diagram of a spacecraft computer...
FAULT TOLERANCE
Summary Fault Tolerant Techniques

NMR = n-modular redundancy
ECC = Error Correction Coding
RESO = RE-comp...
SPACECRAFT PROCESSORS
RAD6000 Processor

 Characteristics
−
–
–
–
–
–
–
–
–

35 Mbps at 33 MHz
Radiation Hardened 32-bit ...
INTEGRATION AND TEST PROCEDURES
Integration and Test Procedure

From Spacecraft Computer Systems, JE Keesee
ocw.mit.edu/co...
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Space Systems & Space Subsystems Fundamentals Technical Training Course Sampler

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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.

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  1. 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
  2. 2. www.ATIcourses.com Boost Your Skills with On-Site Courses Tailored to Your Needs 349 Berkshire Drive Riva, Maryland 21140 Telephone 1-888-501-2100 / (410) 965-8805 Fax (410) 956-5785 Email: ATI@ATIcourses.com The Applied Technology Institute specializes in training programs for technical professionals. Our courses keep you current in the state-of-the-art technology that is essential to keep your company on the cutting edge in today’s highly competitive marketplace. Since 1984, ATI has earned the trust of training departments nationwide, and has presented on-site training at the major Navy, Air Force and NASA centers, and for a large number of contractors. Our training increases effectiveness and productivity. Learn from the proven best. For a Free On-Site Quote Visit Us At: http://www.ATIcourses.com/free_onsite_quote.asp For Our Current Public Course Schedule Go To: http://www.ATIcourses.com/schedule.htm
  3. 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. 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. 5. NEAR Configurations ©Pisacane, 2013
  6. 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. 7. SYSTEM DEVELOPMENT NASA Project Life Cycle Reviews ©Pisacane, 2013
  8. 8. SYSTEM TESTING Sample NASA Payload Test Requirements From: NASA-STD-7002A ─ Payload Test Requirements ©Pisacane, 2013
  9. 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. 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. 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. 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. 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. 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. 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. 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. 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. 18. TRAJECTORY PERTURBATIONS Mars Global Surveyor Aerodynamic Braking ©Pisacane, 2013
  19. 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. 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. 21. LAUNCH FLIGHT MECHANICS Available Launch Inclinations in the United States 37 114 ©Pisacane, 2013
  22. 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. 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. 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. 25. OVERVIEW Attitude Control Schematic ©Pisacane, 2013
  26. 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. 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. 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. 29. SECONDARY BATTERIES Candidate Technologies http://www.clyde-space.com/products/spacecraft_batteries/useful_info_about_batteries/secondary_batteries ©Pisacane, 2013
  30. 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. 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. 32. ANTENNAS Typical Parabolic Antenna Pattern ©Pisacane, 2013
  33. 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. 34. THERMAL ANALYSES Analysis Process ©Pisacane, 2013
  35. 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. 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. 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. 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. 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. 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. 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. 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
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