Comet Sample Return CSR Design Review March 20, 2003
Transcription
Comet Sample Return CSR Design Review March 20, 2003
Comet Sample Return CSR Design Review March 20, 2003 University of California, Santa Cruz Electrical Engineering 127 & 128 Acknowledgements Alec Gallimore Ken Lande Greg Laughlin Douglas Lin Claire Max Graham Smith Martin Yabroff Mission Statement The Objective of the Comet Sample Return project is to develop a viable, robust space system for returning pristine comet nucleus samples for analysis on Earth, thus enabling a better understanding of the formation and evolution of the Solar System and the origins of life on Earth. Why Sample a Comet? • Comets are thought to be the oldest remains of the early Solar System and are composed of the same matter as the early solar nebula. • Comets are also thought to be the main source for water and organic molecules on earth Why Sample a Comet? • National Research Council – “Only a returned sample will allow the necessary elemental, isotopic, organic, and mineralogical measurements to be preformed.” • Better understanding of how the Solar System formed and life on Earth formed. Ways to Understand the Early Solar System Study Pros Cons Geology Comets Asteroids Computer Modeling Meteorites •Human on site analysis •Earth based •Pristine •Easier to samples from land on than the eerily comets solar system •Easier to drill than asteroids •Easy to do •Earth based •Extraterres trial Sample •Earth based •The earth is constantly changing over time •Difficult to land on •Run present state backwards •Numerous outcomes •No initial conditions •Sample can change chemically while entering the atmosphere •Difficult to drill into Scientific Objectives of CSR • To study: – Internal Structure – Ion and Dust tail – Organic Parent Molecules – Rotation – Activity while Approaching the Sun • To Return: – Pristine Samples of a Comets Nucleus – Pictorial Atlas of the Surface of a comet – Sensor Data • • • • • Rotational Spectrometer Permittivity Magnetometer Dust Analyzer Comet Exploration • Giotto, Suisei, Sakigake, Vega 1 & 2 (1985) Halley flyby • ICE (1985) Comet GiacobiniZinner flyby • Stardust (2004) Dust sample collection and return • Rosetta (Planned 2003 launch) Orbiter and Landers • Deep Impact (2004 launch) Sensor probe to impact comets surface • CSR the next step Mission Overview Mission Overview • Rendezvous and trail a comet, at a distance of approximately 1-1.7 AU from the sun, entering into orbit around the comet around 1.7 AU. • Provide ~30 days mapping of comet, with the resolution needed to select an appropriate landing site. – Full surface mapping 10x – Potential landing site mapping 50x • 5-10 sites each mapped 5-10 times Encke Orbit Mission Overview (Cont) • Land on the comet at approximately 2 AU from the Sun and provide real time data on comet environment during the landing process • Deploy Smart Dust seismic motes on the surface of the comet for seismic sounding • Drill at least 0.3m deep to obtain a sample of the comet • Perform a preliminary in-situ compositional analysis of the sample Mission Overview (Cont) • Deploy coma and tail probes to orbit the comet while it approaches the sun • Maintain the sample at the comets ambient temperature (50-200° K) for chemical preservation during voyage back to Earth. • Final result: Return at least 100g of preserved comet material to a Lower Earth Orbit to be recovered by the Space Shuttle. Candidate Comets Wild 2 Encke Encke Orbiter and Lander Configuration Orbiter & Lander Configuration Return Configuration Lander after separation Lander Configuration during Drilling Ops Innovations • Wireless networked miniature sensors (Smart Dust, Micro Motes) • Seismic imaging of the interior of the comet • Portable drilling and sample collection technology • On site sample analysis devices Project Overview: Teams UCSC and University of Michigan Joint Effort UCSC U of M • Mission Science & Architecture • Structures & Materials • Nucleus Sample Return & Analysis • Power & Thermal • Remote Sensing & Analysis • Smart Dust • Communications • Payload Computers Complete System Design • Trajectory & Propulsion • Guidance, Navigation, Command, Control, & Communications • Test & Quality Control • Budget & Scheduling Joint Tasks: Mission Science Communications Michigan Preliminary Study Overview • • • • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications Spacecraft Computer System and Networking Space System Integration Summary and Conclusions Mission Science and Architecture Deutron Kebebew Jason Canoy Outline • • • • • Main Objective Background Mission Architecture Mission Science Mission Conclusion www. comet-cartoons. com Main Objective • To provide humanity with a better explanation and understanding of the source of organic and non-organic molecules and their relation to Comets and life on Earth. • To give insight on the evolution of comets from their origins to their destiny, and to understand the source and evolution of the solar system, the planets, asteroids, and space debris. Why Comets? Determine source space debris Start Determine comets composition Understand how nebulas were formed Understand how life appeared on Earth Understand how the Solar System was formed What Are Comets? • Comets are small, fragile, irregularly shaped bodies composed of mixture of grains and frozen gases. • Comet structures are diverse and very dynamic, but they all develop a surrounding cloud of diffused material. • As comets approach the Sun they develop enormous tails of luminous material that extends for millions of kilometers from the head. Life Cycle of a Comet 6AU, comet inactive 4AU, comet becomes active Tail can be 1-2AU long Comet has welldeveloped tail Earth’s Orbit Comet at perihelion 6AU, comet inactive again Jupiter’s Orbit Anatomy of a Comet Nucleus 1-50 km Bluish ion tail 106-107 km Coma 104-105 km Yellowish dust tail 106-107 km Target Trade Study Best = Comet Orbital Period Perihelion Distance Semi-Major Axis Orbital Inclination Encke 3.3 yrs 0.340 AU 2.21 AU 11.8 deg Wild2 6.39 yrs 1.583 AU 3.44 AU 3.2 deg Subject Profile • Comet P2/Encke Advantages: – It provides clear understanding for the source for dust particles in the inner solar system – It provides insight to the evolution of comets, the solar system and asteroids – Short term space mission and cost effective due to Encke’s short orbital period – It provides a layer highly organic rich surface due to the many encounters with solar wind and solar light – Possibly rock like structure for the nucleus easy for landing and less gases activity Disadvantage: – It provide less of pristine volatile material that is found on comets Mission Architecture • Spacecraft will consist of: – – – – Orbiter Lander Coma Probe Tail Probe • Spacecraft will rendezvous with comet at 1.7-1.9AU. • Orbiter will hold Lander and probes until deployment. Orbital Relation of Comet Encke Deploy tail and coma probes receive and analysis data Continue monitoring the activities of the comet 0.34 AU Earth orbit 1AU Mars orbit 1.5AU Lander live the comet surface Encke orbit 2.2 AU Jupiter orbit 5.2 AU Start following the comet Rendezvous with comet and start imaging, mapping and remote sensing Deploy the Lander and start extraction of sample The Orbiter • Acts as the super node – Communicates with Earth, Lander, and probe • Orbiter Probe measurements – – – – – Surface and Nucleus imaging IR and UV Spectra Ion mass measurements Surface mapping Magnetic field spaceprojects.arc.nasa.gov/ .../pioneer/ PNhist.html The Lander • Sample extraction • Onboard sample analysis – Analysis of chemical, mineral, biological, and isotopic compositions. – Near surface strength – Density and texture – Thermal measurement • Surface imaging • Seismic signal analysis • Return sample to Earth www.mssl.ucl.ac.uk/www_plasma/ missions/rosetta.html Coma Probe • Coma Probe measurements – Measure the material production rate – Measure the different isotope CO and CO2 – Monitor the evolution and production CO, CO2, and H2O – Rational energy of the partials – Photolytic process – Physicochemical process (formation) – Measure neutral gas • Hydrogen Envelope – Measure hydrogen isotopes www.mssl.ucl.ac.uk/www_plasma/ missions/rosetta.html The Tail Probe • Tail Probe measurements – Dust tail analysis • Composition dust particles • Partial size, density, and velocity • Optical prosperities of dust particles – Ionic tail analysis • • • • Ion pressure Magnetic field Ionic particles Plasma http://www.boeing.com/defense-space/space/bss/factsheets/scientific/pioneer/pioneer.html Orbiter Coma probe Lander Tail probe Mission Architecture Architecture Cost Weight Scientific Benefits Mission ~180W ≈ 100kg Sample surface and core of nucleus Mars observer Analysis of coma and nucleus Rosetta Power Lander ≈ 100150M Orbiter/Lander ≈ 170200M ~270W ≈ 250290kg Orbiter/Probe ≈ 150175M ~120W ≈ 200270kg Analysis of coma and tail Star Dust Lander/Probe ≈ 100150M ~180W ≈ 200210kg Analysis of nucleus and tail Pioneer Lander/Orbiter/ Probes ≈ 170250M ~300W ≈ 420470kg Analysis of entire comet CSR Mission Science Objectives • Determine characteristics of nucleus – Mass, size, composition • Study comet environment • Analyze organic compounds and isotopes – Are these molecules connected with the origin of life? Mission Science • Remote Measurements – – – – – – – – – Imaging Mapping Spectrometer readings Rotation Magnetic fields Pressure Gas expulsion rates Particle velocities Photolytic physicochemical process • In Situ Measurements – – – – – – – Compound abundance Particle velocities Density of nucleus Nucleus composition Near surface strength Porosity Seismic Important Measurements Primary = Secondary = What How Why Mass/Size Radio Science Investigation (Rosetta) Obtain better understanding Near Surface Strength Sonic Drill of comet nucleus. Isotopes UV Spectrometer Determine connection with C, N, H, O IR Spectrometer comets to origin of life. EM Fields Magnetometer Radio Noise Heterodyne Radio Receiver Temperature Smart Dust Determine characteristics of comet Total Pressure Smart Dust Determine characteristics of comet Determine characteristics of comet Determine gas production rate of comet Conclusion • Encke is the best target for our mission. • Design an Orbiter/Lander/Probes to carry out the mission. • Desire measurements that describe the nucleus and indicate origins of life. Overview • • • • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications Spacecraft Computer System and Networking Space System Integration Summary and Conclusions Lander Sensors and Sample Collection Kaeton Lee Adam Seger Austin Roman Outline • Drilling • Seismic sources • Sample encapsulation and cooling • In situ analysis Requirements • Drilling – – – – 100g samples Depth greater than 0.3m Pristine sample Power/Mass/Volume • Sample Storage – Find sample temperature – Keep sample at temp. for entire trip back. – Power/Mass/Volume • In-situ Measurements – Obtain scientific analysis incase returned samples are not useable or pristine. – Power/Mass/Volume • Seismic source – Produce seismic source for smart dust analysis – Power/Mass/Volume What Drill to Use? SATM (sample acquisition and transfer mechanism) vs. USDC (ultrasonic/sonic drill corer) SATM Features -Acquire samples to 1.2 meters below surface -Features in-situ instruments for analysis and return containers -Developed for low temperature operation -Ability to prevent cross contamination USDC features - Small light-weight mechanism of drilling and coring with integrated sensors - Can be used as a seismic source for mapping comet - No drill bit sharpening - Drilling technique prevents clogging during core sampling - Ability to take in-situ analysis and core samples for return USDC vs. SATM DRILL POWER (W) axial PreLoad (N) MASS (Kg) Sample Size/ depth of sample Drilling Speed Maintenance Integrated Sensors Autonomous Sample Control Seismic Source USDC 510 < 10 .4 Small/ Short depth slow little yes In Development (date unknown) yes SATM 70 80 32.5 100 g/ slow 112m required no Developed & Operational no Gopher - - - >100g/ slow 1-3m little yes In Development (date unknown) yes Our Choice: USDC with Integrated Sensors Drill Conclusions Why we chose the Ultrasonic Gopher: - Low axial force - Small, light weight, power efficient solution - Ability to take core sample and in-situ analysis - Seismic source: mapping of comet - Integrated Sensor ability: *accelerometers, fiber optic probes, temperature sensors and electrodes - Exciting New Technology with amazing potential Seismic Sources • Seismic sources – Accelerated weight drop system – Vibrators – Drill Seismic Sources Frquency (Hz) Accelerated Weight Drop System Vibrators 7-250 Drill 60-100; 20,000 Best Worst Weight (kg) 45-900 20,000 0.3 Power (kW) 15 130-175 .003-.01 Cooling System Requirements • Must cool to different temperatures – Be able to cool the sample to temperatures between 50-200K, depending on temperature upon acquisition • Must provide adequate cooling at all temperatures • Must operate within mass, power and volume constraints Cooling System Trade Study Cooling System Ball E100 Ball E200 Sunpower M77 Sunpower M87 radiant cooling Mass (kg) Cooling Power Power Consumption Radiation Flight at 70K (W) at 70K (W) Hardness (krad) Tested 4.1 2 60 Yes 6.5 3 70 250 Yes 3 3 60 Yes 2.7 4 100 No 8 2 0 No Best Worst In-situ Analysis • Scientific Objective – The instruments chosen need to determine the elemental, molecular, mineralogical, and isotopic composition of the cometary surface and subsurface. – They need to measure properties like nearsurface strength, density, texture, porosity, ice phases and thermal properties of the comet. Instrument Description • COSAC – (Cometary Sampling and Composition experiment) – Evolved gas analyzer that detects and identifies complex organic molecules from their elemental and molecular composition. Instrument Description • MUPUS – (Multi-Purpose Sensors for Surface and Subsurface Science) – Uses sensors on the Lander’s anchor, a probe and sensors on the exterior of the Lander to measure the density, thermal and mechanical properties of the surface. Instrument Description • MODULUS Ptolemy – (Methods Of Determining and Understanding Light elements from Unequivocal Stable isotope compositions.) – Evolved gas analyzer, which obtains accurate measurements of isotopic ratios of light elements. Instrument Description • ROMAP – (Rosetta Lander Magnetometer and Plasma Monitor) – Studies the local magnetic field and the comet/solar-wind interaction. • Permittivity Probe – Performs an electrical examination of the cometary surface layer. Conclusions • A sonic drill with encapsulation capability • The drill will be used as a seismic source • The Sun power, Inc. M77 cryogenic cooler attached to an encapsulating drill • Five instruments for in-situ analysis • • • • • COSAC ROMAP MUPUS MODULUS Ptolemy Permittivity Probe Overview • • • • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications Spacecraft Computer System and Networking Space System Integration Summary and Conclusions Orbiter Sensors and Sample Collection Bautista Fernandez Edison Estacio Thomas Jun Orbiter Sensors & Data Collection Objective • Map P2/Encke to acquire images in 3D to help determine landing site. • Sense where the space craft is relative to comet nucleus. • Sense molecules and other isotopic compositions of volatile components from both the cometary coma and nucleus. Requirements • Map comet surface for landing – 1-10 m resolution from safe distance (est. 30-100 km) – Determine surface roughness from 1m to 10m • Monitor distance to comet – Measure distance within 1 to 10 m accuracy • Sense molecules and other science information – IR , UV, and Ion observations Remote Sensors • Imaging systems • Spectrometers – Thermal Infrared – Ultraviolet – Ion Mass • Heterodyne radio receiver • Radar altimeter Imaging System - OSIRIS Optical, Spectroscopic, and Infrared Remote Imaging Systems (OSIRIS) • Used on the Rosetta orbiter • Consists of two CCD cameras – Narrow angle camera (NAC) – Wide angle camera (WAC) Imaging System - OSIRIS • NAC is designed to have high spatial resolution to look at comet’s nucleus. • WAC has a wide field of view (FOV) and a high straylight rejection to look at the dust and gas above the surface of the comet’s nucleus. OSIRIS - Distance Requirement NAC WAC Ground Resolutions (m) Height (km) Height (km) 1 50 10 2 100 20 3 150 30 4 200 40 5 250 50 6 300 60 7 350 70 8 400 80 9 450 90 10 500 100 Imaging System - MSI Multi Spectral Imager (MSI) • Used on NEAR Shoemaker • Provides visible and infrared images • At 12-bits per pixel, an uncompressed image is 1.6 Mbits Imaging Systems MSI OSIRIS Total mass 9.55 kg Total Power: 13.04 W Total mass 23kg. NAC WAC 14 microns 14 microns Pixel Size - Array 244 X 537 Focal Length (mm) 168 700 140 Resolution 95 X 161 mrad 20 µrad/px 100 µrad/px FOV 2.25° x 2.9° 2.35° x 2.35° 12.1° x 12.1° Wavelength 400-1100 nm 250-1000 nm 250-1000 nm 2048x2048 pixel x pixel 2048x2048 pixel x pixel Problem with Requirements • Resolution of the MSI instrument is too coarse to satisfy requirements. • For MSI to obtain a surface resolution of 1-10m, it has to be a distance of 10.5-105.3 m from the comet. Preliminary Molecule List Electromagnetic Spectrum Wavelengths Ultraviolet Infrared 400 nm-10 nm 700 nm-1000 um Spectrometer - TES Thermal Emission Spectrometer (TES) • Used on Mars Global Surveyor • Measures incoming infrared and visible energy. • Measures broadband solar reflectance and thermal emittance Thermal Emission Spectrometer (TES) FOV 8.3 mrad per 3x2 array Spatial Resolution 3 km from MGS orbit Wavelength Range 6 - 50 microns Broadband solar reflectance and thermal emittance 0.3 - 2.7 µm and 4.5 - 100 µm Mass 14.4 Kg Average Power 14.5 W Problem with Requirements • The wavelength of TES must be modify to meet Infrared requirement (2.5-50 microns). Spectrometer - Ultraviolet Ultraviolet imaging spectrometer (ALICE) • Used on the Rosetta orbiter • Obtains far UV-spectra, FUV, multispectral or monochromatic images of the comet • Allows analyses of FUV properties of the nucleus and solid grains • Is suited to determine abundances of He, Ne, Ar, Kr, H2O, CO, and CO2 Spectrometer - Ultraviolet F/3 Primary Mirror 0.1x6 deg2 Bandpass 700-2050Å Spectra Resolution 5-6Å pt. source 12.5Å extended source at 700Å 9.8Å extended source at 2050Å Spatial Resolution 0.1x0.6 deg2 Active FOV 0.1x6 deg2 Normal Efficiency Area 0.03-0.53cm2 Mass 2.2 kg Power 2.9 W Spectrometer -IMS Ion Mass Spectrometer (IMS) • Used on Giotto • IMS contains two detectors – High-Energy-Range Spectrometer (HERS) – High-Intensity Spectrometer (HIS). Ion Mass Spectrometer (IMS) Energy Range 20eV to 16keV Mass Resolution ≈ 20 Mass 9 kg Average Power 6.3 W Heterodyne Radio Receiver Heterodyne radio receiver (HIFI) • Has continuous coverage in the frequency of 492-1113ghz • Provides high resolution spectroscopy (R=104-107) over the frequency interval 480-1170ghz (250-625microns) • Provides the out-gassing rate of the comet’s H20 rotational lines Heterodyne Radio Receiver Radar Altimeter TOPEX/Poseidon • Measures topographical inclinations and altitudes of surfaces • Dual frequency (C- and Ku-Band) • Poseidon can measure the surface height within 4.3 centimeters. Poseidon Frequencies Mass Power 13.6GHz and 5.6GHz 23kg 49W Summary & Conclusions • Candidate imaging system MSI does not have sufficient resolution that meets the 1 to 10m requirement • Candidate spectrometers (TES) must be modify to meet Infrared requirement (2.5-50 microns) • Heterodyne Radio Receiver does not meet mass and power requirements • Imaging System - OSIRIS • Spectrometers - Ion Mass, TES, ALICE • Altimeter - TOPEX/Poseidon Outstanding Issues • A light Heterodyne Radio Receiver that consumes minimal amounts of power needs to be researched Overview • • • • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications Spacecraft Computer System and Networking Space System Integration Summary and Conclusions Smart Dust David Ibañez Vincent Lin Smart Dust Objective: • To investigate applications of miniature sensors in a wireless network (Smart Dust) useful for our CSR mission. • Plan out possible ways to complete proposed applications. Smart Dust Outline of Presentation • • • • • • • • • Introduction on Micro Motes Seismic Application Accelerometer Application Landing Site Application Wireless Mote Trade Study Power, Mass, and Volume Deployment Anchoring Conclusion Introduction to Smart Dust Introduction: • Currently under research at Berkeley research labs • Millimeter-scale wireless sensor devices capable of collecting various chemical and physical stimuli. • Micro Motes built from off-the-shelf components RF Micro Mote Seismic Application Illustration Seismic Application • Develop an 2D image Comet’s internal structure • Seismic Tomography – Seismic Source for sending waves into the comet – Collect data with Seismic sensors • Requirements – Good sensor coverage on comet surface. – Appropriate seismic source and sensors(low frequency <500Hz). Seismic Tomography • Earth's structure at 200 km depth below Southeast Asia • Blue color area represent high seismic velocity • Red color area represent low seismic velocity Seismic Sensor Geophones Freq. (Hz) Sample Rate (ms) Mass (g) Volume (cm3) SM-15 14 1 74 16.214 SM-24 10 2 74 16.214 Seismic Formation Accelerometer Application Illustration Accelerometers Trade Study of Accelerometers Accelerometers Power Range Dimensions Axis Consumption Bandwidth Analog Devices ADXL202 2.7~5.25 V ±2g 5x5x2mm dual 0.5uA 10hz Analog Devices ADXL311 3V ±2g 5x5x2mm dual 0.4mA 10hz Analog Devices ADXL210E 3~5v ±10g 5x5x2mm dual 0.6mA 10hz Reiker B3 3~6V ±3g 25 dia.x11mm dual ~1mA 0~160hz Fujikura 3 axis 5V ±2g 5x5x2.6mm tri <50mA 0~150hz Sislicon VSG CRS03-11 5V ±10g 27x13x27mm tri <50mA 10hz Landing Application Illustration Landing Guidance Application • Deploy Optical Motes to potential landing site • Attempt optical communication between motes – If level land site, optical communication are expected to be functional – If there are obstacles, optical communication fails because of direct-line-of-sight requirement • Optical communication may fail due to debris. (Need to use Radar Altimeter) Mote Comparison Trade Study of COTs Dust Mote Type Transmission Range Data Transfer Rate Transmission Properties Dimensions Mass (g) RF 3 - 200 m RFM Transceivers 1 - 20 kbps Laser ~10's km CCR IrDA Sleep/Wake Current@3volt 300 - 916.5 MHz 75x25x12.5mm ~150 1 + 0.75uA/ 7 + 8mA 4 bps 650 nm 25x25x51mm ~160 1uA/25mA ~150 m 30 bps Passive Laser 18x18x2.5mm ~160 50uA/10mA ~60 cm 115 kbps 800-300 nm 50x11x11mm N/A RF Mote battery life (3V Lithium ion battery): 66 hours continuous operation, 1.5 year at 1% duty cycle. Transceiver • RF Monolithics TR1000 – Greatest transmission range (3~200m) – Lowest current consumption – Data transfer rate of 20Kbps • Requires 7.79 cm vertical antenna Power, Mass, and Volume Subject RF Mote Power (mW) Mass (g) Volume (cm3) 2160 7200 1180.8 Geophone SM15 Seismic App. 0 5920 1297.12 2160 13120 2477.92 Accelerometer 2217.6 7200 1183.2 Landing N/A N/A N/A Deployment Canister 32 motes 22.6 cm dia. x 7.62cm 16 motes 11.3 cm dia. x 7.62cm Anchors • Multi-leg • Best stability • Mounted on sides – Mote – Geophone • Compatibility with mode of deployment Conclusion • Three applications for the Smart Dust team – Seismic, Accelerometer, and Landing Guidance • Cannot do Landing Guidance • Depending on Power, Mass, and Volume constraints set for the wireless mote applications, we may not be able to do both the Seismic and Accelerometer applications. • We need to determine the trajectory of the Canister deployment methods. Overview • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications • Spacecraft Computer System and Networking • Space System Integration • Summary and Conclusions Data Collection, Compression, and Communications Rinesh Patel Parisa Toorani Data Collection and Communications Objectives • Design communication links to transmit data – Between the sensors, the probes, the Lander and the orbiter – Between earth and the orbiter and between earth and the Lander • Theoretical frequency and data rate analysis Link Considerations • High gain antenna for space-to-earth link for the Lander and the orbiter • Low gain antenna for communication between the probes, the orbiter, and the Lander • Ka-band links between the orbiter and earth and between the Lander and earth • UHF links between the space modules • NASA Deep Space Network (DSN) Communication Phases Phases of mission: • Cruise – high gain antenna communication (4 hour access time/day) • Reach Encke – orbiter maps the comet and transmits the images via a high gain antenna (10 hours continuous access time/day) Communication Phases (cont.) • Landing - Lander communicates with the probes and orbiter via low gain antennas - orbiter and Lander communicate with earth (4 hour access time/day for each) • Return trip – Lander and orbiter continue transmitting data (3 hour access time/day for orbiter, 5 hour/day for Lander) Link Budget Signal to Noise Ratio: Eb/No = (P • Ll • Gt • Ls • La • Gr)/(k • Ts • R) • • • • • • • • • P – Power (W) = 12 W La – Atmospheric Loss (dB) ≈ 0.9 Ls – Space Loss (dB) = (λ/(4π(4.5x1011 m ))2 Ll – Link Loss (dB) ≈ 0.9 Gt – Transmitter Gain (dB) = 4(π•r)2/λ2 Gr – Receiver Gain (dB) = 4(π•17.5 m)2/λ2 k – Boltzmann’s constant = 1.38x10-23 J/K Ts – System Temperature = 100 K R – Data Rate (bps) Data Rate Estimation Eb/No = 2.7, Receiver Ant - 34 m, 25 GHz, BPSK Reed-Solomon w/ Viterbi coding Data Transmission Time • • • • • Image file 2048 x 2048 pixels (24-bit) ~ 12.3 MB Time of propagation ~ 30 min Data Rate ~ 23 kbps (60 cm diameter Antenna) Send one image (uncompressed) ~ 2 hr 4:1 compressed image (3 MB) ~ 48 min Image Resolution Wide Angle Camera Narrow Angle Camera Distance (km) FOV (m) Resolution (m/pixel) Distance (m) FOV (m) Resolution (m/pixel) 10 km 2119.73 1.03 10 km 410.2 .2 20 km 4239.46 2.07 20 km 820.42 .4 30 km 6359.2 3.11 30 km 1230.6 .6 40 km 8478.9 4.14 40 km 1640.8 .8 50 km 10598.7 5.17 50 km 2051.05 1 # Frames vs. Data Size Narrow Angle Camera Wide Angle Camera FOV # of Data FOV # of (m) Frames Size (m) Frames 410.2 75 225 MB 2119.73 3 820.42 19 57 MB 4239.46 1 1230.6 8 24 MB 6359.2 1 1640.8 5 15 MB 8478.9 1 2051.1 3 12 MB 10599 1 Data Size 9 MB 3 MB 3 MB 3 MB 3 MB Total Estimated Data • Each camera mapping Encke 10 times 40 Uncompressed Images (2048 x 2048) ~ 492 MB • Housekeeping ~ 49.2 MB • Sensors ~ 49.2 MB • Total Data ~ 591 MB • Sufficient storage space X- and Ka-Band Frequencies Frequency Ranges* Frequency Band Transmit f Receive f DSN Operating Frequencies DSN Uplink f DSN Downlink f x (MHz) 7145 – 7190 8400 – 8450 7900 – 8400 7250 – 7750 ka (GHz) 34.20 – 34.70 31.80 – 32.30 27.50 – 31.00 17.70 – 19.70 * Regulated by International Telecommunications Union (ITU) and the World Administrative Radio Conference (WARC) Atmospheric Attenuation This plot shows the attenuation of electromagnetic waves as they enter from different heights above the atmosphere. NASA Deep Space Network (DSN) • Three stations each separated by 120° longitudinally - California - Australia - Spain • 26 m, 34 m, and 70 m antennas • Estimated 10 hour time slot per day Picture obtained from JPL Data Compression Lossless Compression • Original data can be recovered almost entirely • Does not work if original data is analog • Compression ratio varies from 2:1 to 4:1 Lossy compression • Original pixel intensities can not be recovered • Can reconstruct most of the original image • Compression ratio varies from 4:1 to 40:1 Data Compression Trade Studies + indicates rating Information obtained from EV-3M project Antenna Trade Studies Manufacturer Performance COM DEV Ltd. 23 GHz (180 MHz bandwidth) 36 dBi gain COM DEV Ltd. 19 GHz (190 MHz bandwidth) 27 dBi gain Power Size Mass 5.5 kg 0.254 m 5.44 kg diameter, 0.305 m high Transponder Trade Studies Manufacturer Performance Motorola Ka-band (31.8-32.3 GHz) downlink X-band Uplink (7.1457.235 GHz) Output power: (x) – 12 dBm, (ka) – 4 dBm 12.9 W Frequency Range 17.721.2 GHz Output Power: 26 dBm max < 11 W Alcatel Power Size Mass 3 kg 201 x 66 x < 1.25 kg 152 mm Amplifier Trade Studies Manufacturer Motorola SSSD Performance 3.4 W output 23 GHz Power Size 19 W 10.2 x 22.9 x 2.54 cm AIL Systems Inc. 13.8 dBm output per phase 27.5 GHz 18.3 dB nominal gain 13.5 W 3.76 x 2.49 x 0.305 cm Alcatel 2 W output 17.7-20.2 GHz (500 MHz BW) 30 dB nominal gain 12 W 60 x 60 x 15 mm (for 2) Mass 75 g Conclusion The following are being considered: • • • Ka-band frequency 20–25 GHz link UHF links JPEG 2000 data compression Outstanding Issues • Find equipment to closely match our expectations • Decide how much data to send/store • Time scheduling between orbiter-to-earth and Lander-to-earth communication links Overview • • • • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications Spacecraft Computer System and Networking Space System Integration Summary and Conclusions Spacecraft Computer Sys. Turtle Kalus Tarik Elsorady Jeffry Gosal Why do we need a computer? • • • • Communications Instrument and Sensor Control Signal and Data Processing Data Compression • Navigation and Attitude Control • Power and Resource Management Objectives • Design a Computer Control System which can handle our mission requirements • We want to optimize: – Capability – Flexibility – Reliability Computing Requirements • • • • • Speed / MIPS Power Consumption Networkability Memory Capacity Built-in Error Correction • Radiation Hardened • • • • • Latch up Immunity Reliability Cost OS Compatibility Smallest Possible Package • Redundancy Central Processors • Processors – – – – – Aitech S220 Aitech S320 Honeywell Space Computer BAE Systems RAD750 Seakr PPC 603e Central Processor Comparison • Computer Systems Aitech S220 Mass N/A MIPS N/A Power (W) nom./max. 15 / 17.5 Radiation Hardness (kRad) 20 Redundancy single Reliability / years .995 / 25 Boot/User ROM 2MB / 64MB Speed (MHz) 133 SRAM 256MB Temperature Range (°C) -40 to +71 Better Best Aitech S320 N/A N/A 10 / 12.5 20 single .995 / 25 2MB / 64MB 133 256MB -40 to +71 BAE RAD750 549 g 21 <10.2 > 100 single .975 / ~44.5 256kB / 1MB 133 128MB -55 to +80 Honeywell HSC 4545 g 2.2 <5 > 100 dual .995 / ~12 N/A 12 512kB N/A Seakr PPC603e < 1000 g 22 <~10 > 100 single N/A 512kB / 2MB 133 128MB N/A Seakr Engineering PPC 603e Processor Card Radiation Characteristics • Radiation and Temperature Tolerant • Reliable, maintenance-free operation in space • Built using fully controlled and documented processes, approved for use in Deep Space • All components are immune to latch up • Low Single Event Upset (SEU) rate Computing Specs • • • • • • • PowerPC Compatible RISC µ-processor Standard VMEbus Interface 222 MIPS @ 133 MHz Power Consumption < ~10 Watts 128MB DRAM and 2MB EEPROM Built-in 10Base-T Ethernet Interface Can be replaced with a PPC750 Memory System • Magnetic based (ordinary hard drive) – Highly susceptible to corruption in space – Bad choice • Solid State – Solid state memory better withstands the radiation environment of space. Memory • Memory Systems – – – – Aitech S290 16GB High Density Flash Memory BAE Systems 1GB Solid State Recorder Chrislin CI-VME64 – 4GB Seakr 24GB Solid State Recorder Memory Comparison • Memory Systems Aitech 290 Power (W) nom/max 4/5 Radiation Hardness (kRad) N/A Reliability / years N/A Temperature Range (°C) -40 to +71 Memory Capacity 128MB to 16GB Better Best BAE SSR N/A > 60 NA / ~44 -30 to +65 1GB Chrislin CI-VME64 10 / 12.5 N/A N/A 0 to +50 32MB to 1GB Seakr 24-GB VME N/A N/A N/A -50 to +70 24GB Seakr Engineering 24GB VME Memory Card Seakr Engineering Memory Board • Seakr Engineering Solid State Memory Card – – – – – – – – Standard 6U VME Form Factor Total Storage: 24GB Power Consumption: 4 to 6W Operating Temperature Range: -50°C to +70°C 30 MB/sec transfer rate Non-Volatile Flash Ram Proprietary Seakr EDAC software available > 70,000 hours between failures * Seakr Engineering Command and Data Handling System Seakr Engineering C&DH Sys • Command and Data Handling System – Standard VME backplane • • • • Widely used form factor Many COTS applications Saves money in development Expands selection for other peripherals – Readily Available from Seakr • Assures compatibility Seakr Engineering Hardware Image Compressor Seakr Engineering Image Compressor • JPEG2000 Hardware Image Compressor and Data Buffer – – – – – High-speed real time compression Compression ratios greater than 80:1 Flexible quality, resolution and color Solid state memory storage Frees up main processor(s) Seakr Engineering Re-Configurable Computer (FPGA Board) Seakr Re-Configurable Computer • FPGA Board – – – – – Four Reconfigurable Co-Processors 1 GB High-Speed SDRAM Local Memory Processor Controlled Reconfiguration SEU Fault Tolerant Automatic SEU Detection and Scrubbing Equipment Overview • Orbiter – – – – 2 - PPC 603e boards 1 - 24GB SSR 1 – FPGA board 1 – Image Compressor Board – 1 – C&DH System • Lander – – – – 2 - PPC 603e boards 1 - 24GB SSR 1 – FPGA board 1 – C&DH System If Image Compression is needed aboard the Lander, use FPGAs Computing Redundancy • Orbiter – Hardware • Cold Standby Spare during Cruise Phase • Hot Standby Spare and using Seakr’s FPGA board for processor intensive calculations during Intercept and Landing Phase – Software • Use Seakr’s built-in EDAC system • Lander – Hardware • Complete Shutdown during Cruise Phase • Same configuration as Orbiter for Intercept, Landing Phase and Return Phase. – Software • Use Seakr’s built-in EDAC system Memory and Data Redundancy • Orbiter and Lander – Two duplicate sets of data on both the Orbiter and the Lander. – Encode data using Hamming code. Computer Systems Block Diagram 24GB VME Memory FPGA Board PowerPC 603e card 1 PowerPC 603e card 2 Image Compressor & Data Buffer Only in Orbiter Comp. Communications link Embedded Systems Mission Data-Link Phases Comet Orbiter Lander Separation Landing Phase Intercept and Mapping Phase Cruise Phase Sample Collection Phase Lander Return Phase Probe Data Collection Phase Real Time Operating System • • • • Small Memory Footprint Short Interrupt Latency Predictable Behavior Designed for Reliability and Critical Applications Real Time Operating Systems • • • • • Integrity LynxOS 4.0 QNX 6.2 VxWorks 5.x VxWorks AE RTOS Comparison VxWorks AE Support for PowerPC family Yes Support for Protection Domains Yes Number of Priority Levels 256 Maximum number of Tasks Limt'd by mem. Kernel ROM (min, max) N/A Kernel RAM (min, max) N/A Minimum RAM per process N/A Minimum RAM per thread N/A Minimum RAM per queue N/A Typical thread switch latency N/A Guaranteed max. inerrupt latency N/A System Clock Resolution N/A Priority Inversion Avoidance Mech. Yes Multiprocess Support Yes Multiprocessor Support Yes VxWorks 5.5 Yes No 256 Limt'd by mem. N/A N/A N/A N/A N/A N/A N/A N/A Yes Yes Yes Inegrity Yes No 255 N/A 70K 20K 1 Page 1 Page 128 bytes 50-100clks N/A N/A Yes Yes Yes LynxOS 4.0 Yes No 256 Limt'd by mem. 280K, 4M 500K, 4G 1073 bytes 1073 bytes 80 bytes 4us to 19us 14us 20us Yes Yes Yes QNX 6.2 Yes Yes 64 Finite 64K, 64K N/A N/A N/A N/A depends depends depends Yes Yes Yes VxWorks AE • Derived from the widely adopted VxWorks 5.X RTOS • PowerPC support • Many network protocols supported • Implements “protection domain” technology which effectively separates resources and helps prevent conflicts Protection Implemented in VxWorks AE • Protection Domain Technology – Single, flat physical address space is extended to multiple partitions – Enables developers to create logical “containers” to isolate and protect applications Selection • Computer Systems – – – – 4 Seakr PPC 603e 2 Seakr SSR - 24GB 2 Seakr C&DH System 1 Seakr JPEG2000 Image Compressor (only aboard Orbiter) – 2 Seakr RCC – FPGA Board – VxWorks AE Outstanding Issues • Networking • Cost Analysis • Other teams computing needs Overview • • • • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications Spacecraft Computer System and Networking Space System Integration Summary and Conclusions Space System Integration Paul Mellentine Richard Harris Space System Design Integration • Objective • UCSC integration with University of Michigan • U of Michigan PDR Results – Power System – Propulsion • Launch Vehicle • Spacecraft • Mass Itemization Objective – To integrate the UCSC payload designs with the University of Michigan spacecraft design – To keep the UCSC design within power, mass, and volume constraints UCSC and University of Michigan Joint Effort UCSC U of M • Mission Science & Architecture • Structures & Materials • Nucleus Sample Return & Analysis • Power & Thermal • Remote Sensing & Analysis Complete System Design • Trajectory & Propulsion • Smart Dust • Guidance, Navigation, Command, Control, & Communications • Communications • Test & Quality Control • Payload Computers • Budget & Scheduling Joint Tasks: Mission Science Communications Teaming with U of Michigan • Prelimary Study Teaming with U of Michigan • Aerospace Engineering 483 – Cometary Science and Sampling Endeavor (CSSE) • Preliminary Design Review Differences with Michigan • Michigan Mission Architecture – Comet Wild 2 5.2 AU – Orbiter & Lander Return • UCSC Mission Architecture – – – – Comet Encke Lander Return Smart Dust Dust Probes 2.2 AU Orbiter Payload Power Mission Phases Orbiter initial trip mapping landing drilling & take-off return trip analysis Command and Data Handling 30 30 30 30 30 30 Telecommunications 4* 4* 4* 4* 4* 4* Cameras 7* 17* 17* 7* 7* 7* Spectrometers 10* 10* 10* 17* 10* 10* Radar Altimeter 10* 49 49 49 49 TOTALS (W): 51 110 110 107 100 51 * Estimate from U of M CORSAIR Power requirement ~ 110 W Lander Payload Power Mission Phases Lander initial trip mapping landing drilling & take-off return trip analysis Command and Data Handling 30 30 30 30 30 30 Telecommunications 4* 4* 4* 4* 4* 4* Narrow Angle Camera 7* 15* 15* 7* 7* Instruments 10* 10* 10* 17 10* 10* Drilling Device 10 Sample Containment 10 10 10 100 100 100 TOTALS (W): 61 69 69 168 151 149 * Estimate from U of M CORSAIR Power requirement ~ 170 W Propulsion NASA 457-M Hall Thruster Advanced Concept Ion Thruster • • • • • • • Use: Earth escape/capture, comet rendezvous Operating conditions: - 58.5 kW nominal power - Max thrust = 2.5 N - Discharge Isp = 2995 s Will provide a total ∆V of 4.55 km/s Total burn time: 0.54 yrs = 198 days Required Xe Propellant: 1454 kg Total system dry mass: 344 kg • • • • • • • • • Use: Deep space cruise between Earth and Wild-2 System will employ 4 thrusters with 3 thrusters in operation at all times Operating conditions: - 3*20 = 60 kW nominal power - Max thrust = 3*0.400=1.200 N - Discharge Isp = 7500 s Will provide a total ∆V of 14.89 km/s Total burn time: 2.44 yrs = 892 days Required Xe Propellant: 1258 kg Total system dry mass: 556 kg Source: University of Michigan Cometary Science and Sampling Endeavor PDR Power Generation Systems Advantages Solar Radioisotope • Unlimited Power Source • Low fuel consumption • Not reliant on the sun • Unlimited life • High output to • Long discharge fuel ration life • Produces radiation • Heat radiator required • Produces radiation • Requires storage of water and gases 25 – 200 5 – 20 2 – 40 275 0.2 – 300 0.2 – 10 5 – 300 0.2 – 50 • Poor at large distances from the sun Disadvantages • Cells can be damaged by comet dust tail Specific Power* (W/kg) Power Range* (kW) Nuclear Fuel Cell Source:Wertz and Larson, ed. Space Mission Analysis and Design. Third Edition Power Systems Orbiter Power • Radioisotope thermoelectric generators (RTGs) • Power output for two: 20 kW • Specific Power: 65.09 W/kg • T/E conversion: 7% efficiency • Power system mass: 307.22 kg Lander Power • Heat-pipe cooled nuclear reactor • Power output: 61 kW • Specific Power: 17 W/kg • T/E conversion: Brayton cycle (27% efficiency) • Power system mass: 3,588 kg (10% margin) Launch Vehicle • Phase One – Liftoff from Earth – Escape from Earth’s sphere of Influence • What Launch Vehicles will work for our mission? – Mass – Volume Spacecraft Mass Itemization Lander Command and Data Handling Telecommunications High Gain Antenna Low Gain Antenna Transponders Amplifier Wide angle camera *CORSAIR Laser Range Finder *CORSAIR Payload Instruments COSAC ROMAP Modulus Ptolemy MUPUS Permittivity Probe Drilling Device (USDC) Crytogenic Cooler Samples (3) Guidance, Navigation, & Control * Electronic Propulsion System * Thermal Control ** Structure ** Intial Estimate 30% Design Margin Total Mass (kg) 12.000 5.500 5.440 1.250 0.075 13.300 2.160 4.850 0.700 4.500 4.500 0.500 0.400 3.000 0.300 30.108 3158.500 231.93 1159.67 4638.690 1391.607 6030.297 Orbiter Command and Data Handling Telecommunications High Gain Antenna Low Gain Antenna Transponders Amplifier Cameras (OSIRIS) Spectrometers Ion Mass Thermal Ir & Visual UV Radar Altimeter Probes Smartdust Guidance, Navigation, & Control * Electronic Propulsion System * Thermal Control ** Structure ** Intial Estimate 30% Design Margin Mass (kg) 12.000 5.500 5.440 1.250 0.075 23.000 2.200 14.400 2.200 23.000 150.000 11.200 25.583 453.500 36.467 182.337 729.348 218.804 Total 948.152 Orbiter + Lander 6978.45 307.940 3588.000 RTG Power System Nuclear Power System (10% margin)* Spacecraft Total Delta IV-H Payload 10874.389 10889.000 Mission Launch Vehicles Mission Total Mass Launch Vehicle CSR 10875 kg Delta IV-H Cassini 5712 kg Titan IV Rosetta 2900 kg Ariane V Deep Impact 1020 kg Delta II Giotto 960 kg Ariane I Stardust 385 kg Delta II Gravitational Assist Rosetta Mission Example Integration Team Conclusions • Power Requirements – Orbiter Payload – Orbiter Ion Engine 0.110 kW 20 kW – Lander Payload – Lander Hall Thruster 0.170 kW 60 kW • Power Generation Systems – Orbiter • Radioisotope thermoelectric generators 20 kW – Lander • Heat-pipe cooled nuclear reactor 61 kW Integration Team Conclusions • Mass – Total Spacecraft 10875 kg (30% Margin) • Launch Vehicle – Delta IV-H • 10899 kg Payload with Earth Escape • 4 meter Fairing Outstanding Issues • Total Mass – Too large? – Explore Alternate Deep Space Travel Methods • Gravitational Assist Overview • • • • • • • • • Introduction Mission Science and Architecture Lander Sensors and Sample Collection Orbiter Sensors and Sample Collection Smart Dust Applications Data Collection, Compression, and Communications Spacecraft Computer System and Networking Space System Integration Summary and Conclusions CSR Objective • To provide humanity with a better explanation and understanding of the source of organic and non-organic molecules and their relation to Comets and life on Earth. • To give insight on the evolution of comets from their origins to their destiny, and to understand the source and evolution of the solar system, planets, asteroids, and space debris. Scientific Objectives of CSR • To study: – Internal Structure – Ion and Dust tail – Organic Parent Molecules – Rotation – Activity while Approaching the Sun • To Return: – Pristine Samples of a Comets Nucleus – Pictorial Atlas of the Surface of a comet – Sensor Data • • • • • Rotational Spectrometer Permittivity Magnetometer Dust Analyzer Mission Components Orbiter Coma probe Lander Tail probe Launch • Delta IV-H – Spacecraft mass 10875 (30% margin) http://www.boeing.com/companyoffices/gallery/images/space/delta_iv/d4_1st_flight_28.htm Cruise • Hall Thruster – 60 kW power – 1 on Lander • Use: Earth escape/return • Ion Engines – 20 kW power per engine – 3 on Lander • Use: Deep Space Cruise – 1 on orbiter • Use: Following comet • Approach Comet Power • Orbiter – 2 Radioisotope Thermoelectric Generators • 20 kW • Lander – Heat-pipe cooled nuclear reactor • 61 kW Communications • Data Rate ~ 10 kbps • Frequency – ka band (~ 20–34 GHz) – Antenna • Com Dev Ltd. 19 GHz 0.254m • JPEG 2000 Data Compression • Alcatel Transponder 26 dBm • Alcatel Amplifier 30 dB gain DSN Access Time Estimates Phase Use (hours/day) Cruise 2 Orbit 10 Landing 7 Return 5 Computer System • Seakr PPC603e CPU – 133 MHz, 128 MB SRAM • Seakr 24 GB VME Solid State Memory Card • Redundancy – 4 Processors: 2 Orbiter, 2 Lander – Two 24GB Solid State Memory: 1 Lander, 1 Orbiter • Mirrored data storage until Lander returns to Earth • VxWorks AE Comet Rendezvous • Follow Comet at Lagrange Point ~50 km • Orbiter – Surface and Nucleus imaging and mapping • OSIRIS 1m resolution • Determine landing – Molecular Composition Analysis Spectrometers • TES Thermal Emission • ALICE Ultraviolet imaging • Ion Mass – TOPEX Radar Altimeter • Topographical map • Orbiter altitude Landing • Deploy Lander • Drill Nucleus Sample – Drill at least 1ft deep for pristine sample using USDC – Three 100g samples – Sunpower M77 Cooler • In situ Measurements – – – – – COSAC elemental, molecular gas analyzer ROMAP magnetometer and plasma monitor Modulus Ptolemy isotopic ratio MUPUS density, thermal properties Permittivity Probe Smart Dust • Small COTS Wireless RF Network Sensors • Analysis – Seismic Tomography using SM-15 Geophone • 48 Motes for half hemisphere coverage & communication to Lander – Rotational using ADXL311 Accelerometer • Canister Deployment Lander Earth Return • Return to Earth – Ion engines – LEO orbit • Hall thruster – Space Shuttle pickup Probes • Orbiter deploy probes – – – – – Surface and Nucleus imaging IR and UV Spectra Ion mass measurements Surface mapping Magnetic field • Orbiter relays probe data Why Comet Sample Return • “Fossils” in revealing primitive composition of Solar System – Comet Composition: elemental, isotopic, organic, and mineralogical measurements – Understand formation and evolution of Solar System • Chemistry in extreme physiochemical conditions Outstanding Issues • • • • Lightweight Heterodyne sensor Computer Networking Smart Dust Canister Deployment Dynamics Total Spacecraft Mass – Gravitational Assist • Michigan Integration – Mission Architecture Differences Final Comments Comets are sources of valuable scientific data for clues to understanding the origin of solar system and origin of life on Earth. Acknowledgements Alec Gallimore Greg Laughlin Douglas Lin Claire Max Graham Smith Martin Yabroff Ken Lande