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Solar Sail Department of Aerospace Engineering and Mechanics AEM 4332W – Spacecraft Design Spring 2007 2 Solar Sailing: 3 Project Overview – Motivation – Scope – Organization (tasks [%complete], groups, [who?]) – Present the scope of your design work. What are you setting out to do? – Explain how you have organized the work. What are the major tasks? What groups have you organized your team into, and who is in each group? 4 Team Members Orbit: Eric Blake, Daniel Kaseforth, Lucas Veverka Structure: Jon Braam, Kory Jenkins ADC: Brian Miller, Alex Ordway Power, Thermal and Communication: Raymond Haremza, Michael Hiti, Casey Shockman System Integration: Megan Williams 5 Design Strategy Not yet complete. Needs: – Describe all of the trade studies you are considering in this project – Describe the trade study conclusions and any other design decisions that you have already made – Discuss the unfinished trade studies and what effect they will have on your design – Summarize the key properties of the mission (orbit, anticipated lifetime, candidate launch vehicles) – Summarize the key properties of the spacecraft (mass, dimensions, peak and average power requirements, ADCS configuration, type of propulsion system, list of any moving parts, other important info as you see fit) – Show a 3D diagram of the spacecraft (use a CAD package, ie Solid Works or Pro-E) 6 Trade Study Results 7 Cost Estimate Delta II Launch: $42,000,000.00 Navigation System: Carbon fiber booms: $ 250,000.00 Aluminum Bus: $ 1,200.00 2 stepper motors (sail deployment): $ 80,000.00 Heater: Helium Tank: Star Tracker: $ 1,000,000.00 4 Step Motors (sliding masses): $ 160,000.00 Reaction Wheels: $ 600.00 Thrusters: Antenna Horn: Thermal Coating: Sail material: Solar Panels: Total: $43,491,800.00 Before Launch $ 1,491,800.00 Orbit Eric Blake Daniel Kaseforth Lucas Veverka Eric Blake Optimal Trajectory of a Solar Sail: Derivation of Feedback Control Laws 10 Recall Orbital Mechanics • The state of a spacecraft can be described by a vector of 6 orbital elements. – Semi-major axis, a – Eccentricity, e – Inclination, i – Right ascension of the ascending node, Ω – Argument of perihelion, ω – True anomaly, f • Equivalent to 6 Cartesian position and velocity components. 11 Orbital Elements 12 Equations of Motion ^ v 2 r 2 r n n rv 2 ^ r r ^ ^ ^ ^ n cos r sin cos p sin sin p r = Sail Lightness Number = Gravitational Parameter ^ p n sun line sail ^ ^ p r ^ r 13 Problem: Minimize Transfer Time ^ H ( x, , u ) r v 2 v r 2 r n v n 1 r 2 ^ r ^ ^ r 3 v 3 5 (r r )r 2 3 (r n)(v n) n 2(r n) r r r r ^ v r ^ p By Inspection: ^ ^ max{ n v } n v Transversality: ^ 2 ^ 2 ( r n ) p n ( r n ) p n v v r2 2 t t 0 r t t f n sun line sail ^ ^ p r ^ r 14 Solution • Iterative methods are needed to calculate costate boundary conditions. • Initial guess of the co-states must be close to the true value, otherwise the solution will not converge. • Difficult • Alternative: Parameter Optimization. – For given state boundary conditions, maximize each element of the orbital state by an appropriate feedback law. 15 Orbital Equations of Motion x g ( x, , ) d r3 sin( f )W df p sin i da 2 pr p Se sin f T df (1 e 2 ) 2 r 2 r de r r S sin f T 1 cos f T e df p p r d d r2 cos i S cos f T 1 sin df df e p di r 3 cos( f )W df p p df 2 dt r 2 S r r2 cos 3 p 1 e cos f T r2 cos 2 sin sin p a(1 e 2 ) f r r2 1 S cos f T 1 sin f p e W r 2 1 cos 2 sin cos = Sail Lightness Number = Gravitational Parameter 16 Maximizing solar force in an arbitrary direction ^ ^ ^ ^ n cos r sin cos p sin sin p r ^ Maximize: aq r n n q 2 ~ ^ ~ ~ ^ ~ ~ ^ ^ q cos r sin cos p sin sin p r 2 r Sail pointing for maximum acceleration in the q direction: ^ p sun line sail ~ n ^ ^ p r ^ r tan ~ 3 9 8 tan 2 ~ 4 tan 17 Locally Optimal Trajectories • Example: Use parameter optimization method to derive feedback controller for semi-major axis reduction. • Equations of motion for a: da 2 pr 2 df (1 e 2 ) 2 p Se sin f T r p r 1 e cos f p a(1 e ) 2 Feedback Law: ~ e sin f tan 1 e cos f 2 tan S T r2 cos 3 r 2 cos sin sin 2 ~ 3 9 8 tan 2 ~ 4 tan Use this procedure for all orbital elements 18 Method of patched local steering laws (LSL’s) • Initial Conditions: Earth Orbit a 1 e 0 i 0 0 0 t t0 0 • Final Conditions: semi-major axis: 0.48 AU inclination of 60 degrees a 0.48 AU e ~0 i 60 free free t tf free 19 Trajectory of SPI using LSL’s Time (years) 20 21 Global Optimal Solution – Although the method of patched LSL’s is not ideal, it is a solution that is close to the optimal solution. – Example: SPI Comparison of LSL’s and Optimal control. 22 Conclusion • Continuous thrust problems are common in spacecraft trajectory planning. • True global optimal solutions are difficult to calculate. • Local steering laws can be used effectively to provide a transfer time near that of the global solution. Lucas Veverka •Temperature •Orbit Implementation Optimal Trajectory of a Solar Sail: Orbit determination and Material properties. Lucas Veverka 25 Reflectivity Approximation • Reflectivity constant, r, negatively affects the solar radiation pressure force. f 2PArui n n 2 – – – – P is the solar pressure as a function of distance. A is the sail area being struck by the solar radiation. ui is the incident vector. n is the vector normal to the sail. • Emissivity and specular reflection neglected. • Assumed a Lambertian surface. 26 Sail Surface Temperature Fsolar Tsurface 2 4d sun 1 4 • Fsolar is the solar flux. • • • • α is the absorptance. ε is the emittance. σ is the Stefan-Boltzman constant. dsun is the distance from the sun. 27 Transfer Orbits • Objective: -Reach an orbit with semi-major axis of 0.48 AU and inclination of 60 degrees as quickly as possible. • Investigated four possible orbits -Cold transfer orbit -Hot transfer orbit -Inclination first transfer orbit -Simultaneous orbit 28 Cold Transfer Orbit • Advantages: – Very simple two-stage transfer. – Goes no closer to sun than necessary to avoid radiation damage. • Disadvantages: – Is not the quickest orbit available. • Order of operations: – Changes semi-major axis to 0.48 AU. – Cranks inclination to 60 degrees. • Time taken: – 10.1 years. 29 Cold Transfer Orbit 30 Hot Transfer Orbit • Advantages: – Still simple with three-stages. – Is a much quicker transfer. • Disadvantages: – Radiation is very intense at 0.3 AU. • Order of operations: – Changes semi-major axis to 0.3 AU. – Cranks inclination to 60 degrees. – Changes semi-major axis to 0.48 AU. • Time taken: – 7.45 years. 31 Hot Transfer Orbit 32 Inclination First Transfer Orbit • Advantages: – Very simple two-stage transfer. – Avoids as much radiation damage as possible. • Disadvantages: – Takes an extremely long time. • Order of operations: – Cranks inclination to 60 degrees. – Changes semi-major axis to 0.48 AU. • Time taken: – 20.15 years. 33 Inclination First Transfer Orbit 34 Conclusion • Simultaneous transfer is too complicated with little or no real benefit. • Inclination first transfer takes too long. • Hot transfer orbit is much quicker but submits materials to too much radiation. • Cold transfer orbit is slower than the hot but gets the equipment to the desired location safely. • Choice: Cold transfer orbit! Daniel Kaseforth Control Law Inputs and Navigation System 36 Structure Jon T Braam Kory Jenkins Jon T. Braam Structures Group: • Primary Structural Materials • Design Layout • 3-D Model • Graphics -72 Project Hours- 39 3-D Model • Blah blah blah (make something up) 40 Graphics • Kick ass picture 41 Launch Vehicle Weight and Volume Constraints • Delta II : 7400 Series • Launch into GEO – 3.0 m Ferring » Maximum payload mass: 1073 kg » Maximum payload volume: 22.65 m3 – 2.9 m Ferring » Maximum payload mass: 1110 kg » Maximum payload volume: 16.14 m3 42 Launch Vehicle 2.9 Meter Ferring In Flight Organization – Antenna Stowed – Solar panels folded to sides 43 Primary Structural Material Aluminum Alloy Unistrut® – 7075-T6 Aluminum Alloy • Density – 2700 kg/m3 – 168.55 lb/ft3 • Melting Point – 477 to 635oC http://www.matweb.com/SpecificMaterial.asp?bassnum=MA7075T6 44 Primary Structural Material • Density • Mechanical Properties – Allowing Unistrut design • Decreased volume • Factor of safety 15.0 – 0.06in thick (12 Ga) • Thermal Properties – Capable of taking thermal loads 45 Primary Structural Material • 7075-T6 Aluminum Alloy – Used in ISS structure – Useful Characteristics • Resistance to general corrosion • Resistance to pitting • Crack resistance – Inter-grainulairy – Stress corrosion 46 Design Layout • Constraints – Volume – Service task – Thermal consideration – Magnetic consideration – Vibration – G-loading (7.5) 47 Design Layout • Unistrut Design – Allowing all inside surfaces to be bonded to • Titanium hardware – Organization • Allowing all the pointing requirements to be met with minimal attitude adjustment 48 Design Layout 49 3-D Model • Sail Attachment – Single Tube • 0.09m Diameter – Allows Nitrogen to feel stabilizing thrusters – Supports Sail and Argon External Tank – Mounted to tube bus 50 Odds and Ends • Things to be improved upon – Safety Factor of 14 in Compression and Bending • Could be reduced to save weight Other Projects •Sliding mass movement design •Boom deployment methods •Moment of Inertia determination •Mass budget 51 Other Projects Trade Studies - Structural Materials used in ISS and other long term spacecraft. - Deployment methods and other autonomous movement used in space. - In space structural connections Kory Jenkins • Sail Support Structure • Anticipated Loading •Stress Analysis • Materials •Sail Deployment 53 Sail Sizing • Characteristic acceleration is a measure of sail performance. 2P ao s mp / A s ms / A • Characteristic acceleration increased with sail size. • Higher acceleration results in shorter transfer time. • Sail size is limited by launch vehicle size and deployment power requirements. 54 Sail Support Structure • Challenge: Design a robust, easy to deploy structure that will maintain sail shape. • A 150 x 150 meter sail covers the same area as 5 football fields. (22,500 square meters) • Solution: An inflatable boom structure based on the L’Garde design supports 4 triangular sail quadrants. • Booms are deployed in pairs to minimize power consumption. 55 Step 5 Step 1 Deployment cables retract to pull the sail quadrants out of their storage compartments. Heater: Raises boom temperature above glass transition temperature to 75 C. To sail quadrant Step 4 Once deployed, booms cool below glass transition temperature and rigidize. Step 2 Inflation gas inlet: booms are inflated to 120 KPa for deployment. Step 3 Cables attached to stepper motors maintain deployment rate of ~ 3 cm/s. To deployment motor 56 Estimate Worst Case Loading Solar Pressure P = 2/3 P_quadrant Assumptions: • Solar Pressure at 0.48 AU = 19.8 µN/m^2. • Thin wall tube. • Sail quadrant loading is evenly distributed between 3 attachment points. • Isotropic material properties. • Safety factor of 3. 57 Analysis of a Tapered Beam My Bending I 2 EI Buckling Pcr 4L2 Shear VQ max Iy Hoop stress (inflation pressure) Section Modulus Pmax t hoopt r x S ( x) dA (dB dA) 4 L 2 58 • • • Expected deployment loads of 20 N in compression dictate boom sizing. Booms sized to meet this requirement easily meet other criteria. Verified using laminate code that accounts for anisotropy of composite materials. 59 Boom Specifications • • • • • Cross-ply carbon fiber laminate. IM7 carbon fiber TP407 polyurethane matrix, Tg = 55 deg C Major Radius = 18 cm, minor radius = 10 cm. Length = 106 meters. Analysis of a Composite Laminate: EL V f E f Vm Em V f Vm ET E f Em 1 [Q]K [ o z T ] K 60 Conclusions and Future Work • Sail support structure can be reliably deployed and is adequately designed for all anticipated loading conditions. • Future Work – Reduce deployment power requirement. – Reduce weight of support structure. – Determine optimal sail tension. Attitude Determination and Control Brian Miller Alex Ordway Alex Ordway 60 hours worked Attitude Control Subsystem Component Selection and Analysis 63 Design Drivers • • • • • Meeting mission pointing requirements Meet power requirements Meet mass requirements Cost Miscellaneous Factors 64 Trade Study • Sliding Mass vs. Tip Thruster Configuration – Idea behind sliding mass 65 Trade Study • Sliding mass ACS offers – Low power consumption (24 W) – Reasonable mass (40 kg) – Low complexity – Limitations • Unknown torque provided until calculations are made • No roll capability • Initially decided to use combination of sliding mass and tip thrusters 66 ADCS System Overview • ADS – Goodrich HD1003 Star Tracker primary – Bradford Aerospace Sun Sensor secondary • ACS – Four 10 kg sliding masses primary • Driven by four Empire Magnetics CYVX-U21 motors – Three Honeywell HR14 reaction wheels secondary – Six Bradford Aero micro thrusters secondary • Dissipate residual momentum after sail release 67 ADS • Primary – Decision to use star tracker • Accuracy • Do not need slew rate afforded by other systems – Goodrich HD1003 star tracker • • • • • 2 arc-sec pitch/yaw accuracy 3.85 kg 10 W power draw -30°C - + 65 °C operational temp. range $1M – Not Chosen: Terma Space HE-5AS star tracker 68 ADS • Secondary – Two Bradford Aerospace sun sensors • • • • • Backup system; performance not as crucial Sensor located on opposite sides of craft 0.365 kg each 0.2 W each -80°C - +90°C 69 ACS • Sliding mass system – Why four masses? – Four Empire Magnetics CYVX-U21 Step Motors • • • • • • Cryo/space rated 1.5 kg each 28 W power draw each 200 °C $55 K each 42.4 N-cm torque 70 ACS • Gear matching- load inertia decreases by the gear ratio squared. Show that this system does not need to be geared. 1 2 70m a(600sec) 2 a 0.00389 sm2 F ma (10kg )(0.00389 sm2 ) F 0.0389 N 71 ACS • Three Honeywell HR14 reaction wheels – Mission application – Specifications • • • • • • 7.5 kg each 66 W power draw each (at full speed) -30ºC - +70ºC 0.2 N-m torque $200K each Not selected – Honeywell HR04 – Bradford Aerospace W18 72 ACS • Six Bradford micro thrusters – 0.4 kg each – 4.5 W power draw each – -30ºC - + 60ºC – 2000 N thrust – Supplied through N2 tank 73 Attitude Control • Conclusion – Robust ADCS • Meets and exceeds mission requirements • Marriage of simplicity and effectiveness • Redundancies against the unexpected Brian Miller •Tip Thrusters vs. Slidnig Mass •Attitude Control Simulation 75 Attitude Control • Conducted trade between tip thrusters and sliding mass as primary ACS • Considerations – Power required – Torque produced – Weight – Misc. Factors 76 Attitude Control • Tip Thrusters (spt-50) – Pros • High Torque Produced ~ 1.83 N-m • Low weight ~ 0.8 kg/thruster – Cons • Large Power Requirement ~ 310 Watts • Lifetime of 2000 hrs • Requires a fuel, either a solid or gas 77 Attitude Control • Attitude Control System Characteristics – Rotational Rate – Transfer Time – Required Torque – Accuracy – Disturbance compensation 78 Attitude Control • Requirements – Orbit • Make rotation rate as fast as possible • Roll spacecraft as inclination changes – Communications – Within Maximum Torque • Pitch and Yaw Axis ~ 0.34 N-m • Roll Axis ~ 0.2 N-m U mFz m – sliding mass M F – solar force z – distance from cg M – spacecraft mass 79 Attitude Control • Pitch and Yaw Axis • Rotation Rate = 0.144 rad/hr ~ 8.25 deg. • Transfer Time = 5300s ~ 1.47 hrs • Required Torque = 0.32 N-m ~ 98.8% of maximum produced • Converges to desired angle Torque Req. Transfer Time Slope = 0.00004 rad/s 80 Attitude Control • Roll Axis Torque Req. • Rotation Rate = 0.072 rad/hr ~ 4.12 deg • Transfer Time = 7000s ~ 1.94 hrs • Required Torque = 0.15 N-m ~ 75% of maximum produced • Converges to desired angle Transfer Time Slope = 0.00002 rad/s Power, Thermal and Communications Raymond Haremza Michael Hiti Casey Shockman Raymond Haremza Thermal Analysis •Solar Intensity and Thermal Environment •Film material •Thermal Properties of Spacecraft Parts •Analysis of Payload Module •Future Work Thermal Analysis and Design -Raymond Haremza 84 Design Approach Strategy 85 Decision to take “cold” orbit By taking longer to get to 0.48 AU, we in turn reduce the amount of design, analysis, production time and weight. Solar Sail Material and Thermal Analysis 86 87 Payload Panel Analysis The Carbon-Carbon Radiator has aluminum honeycomb sandwiched between it, and has thermal characteristics, Ky= Kx=230W/mK, and through the thickness Kz = 30W/mK which allows the craft to spread its heat to the cold side of the spacecraft, but also keeping the heat flux to the electric parts to a minimum. Material Properties 0.06 0.78 E 1.2e7 psi G 6.11e6 psi v 0.32 88 Spacecraft Heat Transfer Analysis 4 1026 W flux 2 2 4 d m 7.00E+03 6.00E+03 5.00E+03 4.00E+03 3.00E+03 2.00E+03 1.00E+03 0.00E+00 9.80E-01 8.80E-01 7.80E-01 6.80E-01 5.80E-01 4.80E-01 Qsun flux A Watts Distance from Sun (AU) Qsun Tsurface Atotal 1 4 Kelvin Solar Intensity (flux) (W/m^2) Solar Intensity vs Distance 89 Heat Transfer Analysis Qsun flux A 4 Qrad Atot Tsurf Tsurf Qsun Qrad 1 4 Setting the heat fluxes together yields the surface temperature of the object based on emmissivity, absorbitivity, size and geometry of the object. Atot A Thermal Analysis of Payload Module 90 Thermal Analysis of Payload Module 91 92 Temperature vs Distance (Side of Payload Module) 300 280 Temperature (K) 260 85 80 75 70 65 60 55 240 220 200 180 160 140 120 100 4.80E-01 5.80E-01 6.80E-01 7.80E-01 8.80E-01 Distance from Sun (AU) 9.80E-01 deg deg deg deg deg deg deg 93 Temperature vs Distance (Top of Payload Module) 450 400 Temperature (K) 350 0 incidence 5 deg 10 deg 15 deg 20 deg 25 deg 30 deg 35 deg 300 250 200 150 4.80E-01 5.80E-01 6.80E-01 7.80E-01 Distance from Sun (AU) 8.80E-01 9.80E-01 Spacecraft Component Thermal Management Notes: By using thermodynamics the amount of heat needed to be dissipated from the component taking into account its heat generation, shape, size, etcetera. If the component is found to be within its operating range, the analysis is done, if not a new thermal control must be added or changed. 94 95 Thermal Analysis of Antenna 96 Antennae Operating Temp (-373 to 373K) vs Distance With White Paint Reflector 390 Temperature (K) 370 350 330 310 290 270 250 4.80E-01 5.80E-01 6.80E-01 7.80E-01 Distance From Sun (AU) 8.80E-01 9.80E-01 97 Star Tracker Thermal Analysis Using the heat generated (10W), and using common coating material ( ); the required to maintain the star tracker’s temperature to 30 K can be found by. Qdiss Qtot T Atotal 4 s Knowing the heat needed to dissipate, a radiator size can be calculated, or other thermal control methods (MLI) can be used to maintain temperature. Arad Qsun Qgenerated Atotal 4 Ts 98 Heat Needed to Radiate Away From Star Tracker to Keep Temp 303K 1.60E+03 1.40E+03 1.20E+03 Heat (W) 1.00E+03 8.00E+02 6.00E+02 4.00E+02 2.00E+02 0.00E+00 4.80E-01 -2.00E+02 5.80E-01 6.80E-01 7.80E-01 Distance (AU) 8.80E-01 9.80E-01 99 Using the amount of heat needed to be radiated from star tracker, the additional area required to dissipate heat can be calculated and chosen. Area of Radiator Needed to Keep Star Tracker Surface Temp at 303K Area of Radiator (m^2) 2.50E+00 2.00E+00 1.50E+00 1.00E+00 5.00E-01 0.00E+00 4.50E-01 5.00E-01 5.50E-01 6.00E-01 Distance (AU) 6.50E-01 7.00E-01 100 Thermal Analysis of Microthruster Notes: Since Microthrusters need to be within 247 to 333 K, will have to add MLI to stay within thermal constraints. Analysis of Multilayer insulation… 101 Microthruster and Sun Senser Temperature vs Distance Temperature (K) 700 650 600 550 500 Microthruster Side Microthruster Top Sun Sensor 450 400 350 300 250 200 4.80E01 5.80E01 6.80E01 7.80E01 8.80E01 Distance (AU) 9.80E01 102 Thermal Analysis of Solar Panels Need to radiate heat away from solar sail, any ideas, stephanie, group? 103 Tempurature (K) Solar Panel Temp (Operating temp 123 to 400K) vs Distance from Sun 580 560 540 520 500 480 460 440 420 400 380 360 340 320 300 4.80E-01 5.80E-01 6.80E-01 7.80E-01 8.80E-01 Distance from sun (AU) 9.80E-01 104 Casey Shockman • Communications 105 Major Tasks • Trade Studies – Frequency – Antenna types – Power – Data transfer rates • Sizing the Antennas • Determine placement of antennas 106 Antenna Selection and Sizing • Initial Conditions – Payload stores data at a rate of 15.6 kbps. – Need to transmit data 1 or 2 times per week. • 1 week of storage is equal to around 9,500,000 kb. • We choose two 12,000,000 kb hard drives to store information. One hard drive will be used as backup. – Satellite needs to transmit data anywhere from .5 to 1.5 AU – All aspects of the DSN (size, SNR, noise temp.etc.) 107 Frequency • S-Band: 2 GHz – Used primarily for short distance. • X-Band: 8.4-8.5 GHz – This is the typical frequency used, so DSN is becoming overloaded at this frequency. • Ka-Band: 31.8-32.3 GHz – Due to overloaded X-Band frequency, the DSN is migrating to Ka-Band frequency. – Can transfer data much more quickly than X-Band. Solar Sail will use Ka-Band transmit with X-Band receive/transmit capabilities. 108 Process • This equation was then used with the following BER vs. SNR to solve for variables. 109 Bit Error Rate vs SNR 110 Process • A SNR is chosen to correspond to a BER of 10-6. • T is noise temperature which is based on the angle with the sun and earth, elevation angle of the earth antenna, weather conditions, distance between satellites. • From this, the gain and power transmitted was optimized for each frequency, antenna, distance and data transfer rate • The following chart was created for each antenna, frequency, and distance from the sun. Variables included power, noise temperature, and antenna size. 111 112 Antenna Types • Directional – – – – Parabolic Reflector Horn Array Helix • Omni-directional – Dipole – Conical 113 High Gain Directional Antennas 114 Directional Antennas • Parabolic Reflector – High data transfer rate with low power required. – Works with either X-Band receive/transit or KaBand receive/transit, not both. – Conventionally heavier than horn, but recent unused membrane dish antennas may be lighter in the future. – Can achieve high gain and a range of beamwidths. 115 Directional Antennas • Arrays – Gain is low for small areas. – Heavier than horn or parabolic reflector due to the large area needed to achieve desired level of gain. – Can attain any beamwidth. 116 Directional Antennas • Helix – Can attain any beamwidth necessary. – Antenna will have a low diameter but needs to be long to achieve high gain. – Length of antenna makes pointing and storage very difficult. – Length of antenna also adds resistance, so efficiency drops with length. 117 Directional Antennas • Horn – High data transfer rate with low power required. – Works directly with recently developed Small Deep Space Transponder. – New design works with X-Band and Ka-Band transmit as well as X-Band receive. – Smaller than conventional parabolic reflector and array. – High gain. – Ability to track using Delta Differenced One-Way Range (DDOR) because two tones can be sent at once (DSN stats.pdf 9). – Small beamwidth, suitable for long-range communications. – The Solar Sail will have two horn antennas. 118 119 120 Conclusions • The horn antenna was chosen because of its small size compared to the other choices. • The antenna cannot transmit at a Sun-EarthProbe angle smaller than .3 degrees or on a very stormy day at the ground station. • Different antennas would be used on the sun side and shade side of the antenna. • The sun side antenna would be .2 meters in diameter. The shade side antenna would be .075 meters. 121 More conclusions • The minimum transfer time for this setup is 1 hour using Ka-band transmission. • If the required signal to noise ratio is not met due to SEP angle or weather on earth, the transfer rate can be slowed to allow for more accurate data. • Power used for transfer is 30 watts. 122 Directivity • Horn directivity is estimated by the following equation: 225 HPBW *d 123 Beamwidths • Using this equation: – Sun-side antenna • X-Band HPBW=13.42 • Ka-Band HPBW=3.35. – Shade side antenna • X-Band HPBW=35.79 • Ka-Band HPBW=8.95 • These beamwidths are all much larger than the pointing accuracy so there will be very little pointing error. 124 Low Gain Omni-Directional Antennas 125 Low-Gain Antenna Selection • Omni-Directional Antenna – The goal is have a low data rate communications when not pointing at earth – There are many choices for low gain antennas. The solar sail will have two conical equiangular spiral antennas. – These two antenna will ensure the satellite will always be within contact with the DSN. Omni-directional Transfer Dsn stats 5 126 127 2-Arm Conical Equiangular Spiral Antenna Gain will be 0 dBi (isotropic) from -70 to +70 degrees. Gain will be -25 dBi from -90 to -70 and 70 to 90 degrees for each antenna. Using this configuration, at the worst case scenario, the low gain antenna can transmit 1 bps with an accuracy of 10-3. 128 Costs 129 DSN Cost Dsnstats.pdf This gives a cost of about $1100 per hour of transmission within the DSN network. 130 Antenna Costs • .2 m diameter horn antenna: • .075 m diameter antenna: • conical equiangular antenna: • hard drive: • Total cost: 131 Masses • .2 m diameter horn antenna: – 2.75 kg • .075 m diameter antenna: – .40 kg • conical equiangular antenna: – 2 x .25 kg • hard drive: – 2 x .79 kg • Miscellaneous – 1 kg • Total mass = 6.23 kg Michael Hiti Power 133 Objectives • Determine the amount of power required to support the payload instruments, and all other components of the spacecraft • Perform a trade study to determine whether to use a normal-pointing or conformal solar array • Determine appropriate solar array materials • Determine appropriate solar array size 134 Objectives (continued) • Determine appropriate battery type to be used in mission • Determine appropriate battery size 135 Power Requirements Peak Power (W) Remote Sensing Instruments Coronograph 4 All Sky Camera 3 EUV Imager 5 Magnetograph - Helioseismograph 5 Magnetometer 2 IN-SITU Instrument Package Solar Wind Ion Composition and Electron Spectrometer Energetic Particle (20keV - 2MeV) 3.5 2 Attitude Control Small Reaction Wheels 70 Large Reaction Wheel 70 Sliding Mass 40 Structure Heat Curing Elements 335 Communications Antenna Gimbal 8 Antenna 36 Thermal Management 50 Misc/Thermal TOTAL 633.5 • All power requirements for solar sail 136 Power Requirements (continued) Peak Power (W) Structure Heat Curing Elements 335 Communications Antenna 36 Large Reaction Wheel 70 Thermal Management 50 TOTAL 491 Attitude Control Misc/Thermal • Anticipated beginningof-life (BOL) power load 137 Power Requirements (continued) Remote Sensing Instruments Coronograph 4 All Sky Camera 3 EUV Imager 5 Magnetograph - Helioseismograph 5 Magnetometer 2 IN-SITU Instrument Package Solar Wind Ion Composition and Electron Spectrometer Energetic Particle (20keV - 2MeV) 3.5 2 Attitude Control Small Reaction Wheels 70 Communications Antenna Gimbal 8 Antenna 36 Thermal Management 50 Misc/Thermal TOTAL 188.5 • Anticipated endof-life (EOL) power load 138 Array Sizing • Key Equations Vchg = (1.2) * Vbus= 34.2 V Cchg = (PL* td ) / (Vbus* DOD) = 52.9 Ah Pchg = (Vchg* Cchg)/15h = 120.6 W PEOL = (PL + Pchg) = 310 W • • • • • • Vchg is the array voltage Cchg is the total charge capacity of the battery PL is the required power load at EOL td is the anticipated max load duration (2h) Pchg is the power required to charge the batteries DOD is the depth of discharge (0.25) 139 Array Sizing (continued) • The BOL power requirement is found by assessing the various efficiency factors that lead to the conditions at EOL Temperature efficiency = ηtemp = 1 - (0.005/K)*(Tmax – Tnom) Radiation efficiency = ηrad = 1- R Cosine loss = ηangle = cos(α) PEOL = ηtemp * ηrad * ηangle * PBOL • • • • Tmax is the maximum solar cell operating temperature Tnom is the nominal solar cell operating temperature R is the percent loss due to radiaiation damage α is the maximum angle off-normal to the sun 140 Array Sizing (continued) • Using a conformal solar array Assuming: ηtemp ≈ 0.51 ηrad ≈ 0.3 ηangle ≈ 0.81 PBOL = 1395 W 141 Array Sizing (continued) • Array area equations Acell = PBOL / ( ηGaAr* Is ) Aarray = Acell / ηpack • • • • • Acell is the area of the solar cells Aarray is the area of the array ηGaAr is the efficiancy of the solar cells Ηpack is the packing efficiency Is is the solar intensity 142 Array Sizing (continued) Acell = 0.8718 m^2 With a packing efficiency of 90% Aarray = 0.969 m^2 • These values reflect the sizes required to meet EOL power requirements at 0.48AU • We must check to make sure this array area will generate enough power to support the BOL requirements at 1AU 143 Array Sizing (continued) • Assuming that there is no radiation and cosine loss • Assuming a ηtemp ≈ 0.90 • Is = 1355W/m^2 at 1AUl The BOL load ≈ 546W This would require an Acell ≈ 1.413 m^2 and an Aarray ≈ 1.57 m^2 This means that the array sizing based on the EOL requirements will not support the BOL load requirments. • The BOL load requirements are the driving force behind the array sizing 144 Array Mass • Gallium Arsenide cells weigh 84mg/cm • Solar panels and coverslides weigh 2.06 kg/m^2 • Aluminum honeycomb panel backing weighs 0.9 kg/m^2 The total mass of a conformal array will be 5.963 kg 145 Solar Array • Solar cells and panels made by Spectrolab – Ultra Triple Junction GaAs cells – 28.5% efficiency – 84 mg/cm^2 (cells) – 2.06 kg/m^2 (panel) 146 Trade Study • Advantages to using of a normal-point solar array – Able to collect maximum possible solar energy – Requires smaller solar array – Array could be positioned to minimize thermal and radiation damage • Disadvantages to using of a normal-point solar array – Added mass of gimbal used for positional array – Added complexity to design – Creates problems regarding stowage in capsule 147 Trade Study (continued) • The BOL power requirements have caused our solar array to be nearly twice area required to meet the EOL power requirements • The reduction of mass is our highest priority • The smallest gimbal used for array positioning alone weighs approximately 5kg – This is nearly equal to the entire mass of our array • Since our array is already oversized for EOL requirements, an array with normal pointing capabilities will not be beneficial 148 Battery Sizing • Key Equations Cchg = (PL* td ) / (Vbus* DOD) = 52.9 Ah Ebat = (Vbus* Cchg) = 1508 W h mbat = Ebat / ebat • Ebat is the battery energy capacity • ebat is the energy density of the battery • mbat is the mass of the battery mbat = 8.6 kg 149 Battery • Batteries made by BST Systems – Silver-Zinc Battery – 1.5 V/cell – 175(W h) / kg 150 Demonstration of Success 151 Failure Modes and Effects Analysis • Boom fails to fully inflate due to problem with tank, heater, etc. – – Sail may still function, would apply different torques, difficult to control One or more of the booms could fail to extend fully. i.e. the heaters don't work, or the inflation gas tank ruptures or it gets caught on something. If that were to happen, it might be possible to run up the sail part way, although there would be a lot of slack in it, and therefore a loss of propulsion efficiency. And the attitude control system might not be able to compensate for the asymmetric torque...assuming the sliding mass on the malfunctioning boom worked at all...I mean, um...yeah, it'll work perfectly... • • Failure of navigation system... sail fails to know it's location and can no longer implement control laws; will not reach desired orbit. Failure Modes and Efffects: 1. Module structure fails at 7.5 g's and breaks it shit off on exit. Effect: It will spread debris throughout LEO. Something like the Chinese did about 9 months ago. Oops. My Bad. 2. The Sail gets kinked inside the Bus module and is unable to deploy or rips on deployment. Effect: Huge embarrassing failure for the UofM design team. 3. The solar array is not able to pivot downward from its storage/capsule setup to its working format. Effect: Same as #2. • FMEA Thermal can screw everything up. I don’t think I can narrow it down to one thing. If I have to I guess I will. Anyways, heres my FDR slides thus far, not done yet, but pretty much done calculating stuff. Now I have to explain things, add equations and graphics and explain what I would do if I had more time. I think stephanie will have plenty to say about what I have already. Thanks 152 Future Work 153 Acknowledgements • • • • • • Stephanie Thomas Professor Joseph Mueller Professor Jeff Hammer Dr. Williams Garrard Kit Ru…. ?? Who else??