Asteroid Sample Return AEM 4332 Final Design Review 5/7/2008 Becky Wacker Carla Bodensteiner Ashley Chipman John Edquist Paul Krueger Jessica Lattimer Nick Meinhardt Derek Steffes Sam Zarovy 1

Becky Wacker Team Overview Attachment Method 2

Mission Goal Drill into the asteroid and return 100 g core sample from a depth of 1 meter for study on Earth. 3

Our Mission Objectives 1.Leave orbit about the asteroid and position for landing 2.Land on the asteroid with drill and return vehicle intact 3.Attach the spacecraft to the asteroid for 24 hours withstanding a 10 N reaction force and a 0.1 N/m torque from the drill 4

Requirements Spacecraft must weigh no more than 750kg including: – 50kg payload package (drill and stereo cameras) – 200kg return vehicle The spacecraft must land and support drilling operations for 24 hours Landing must not exceed 15 g’s Spacecraft must support drilling operations: – 10 N reaction force – 0.1 N/m torque 5

Expectations Explain the results of all system and hardware trade studies including rationale for final selections Provide spacecraft layout drawing(s) Demonstrate compliance with all requirements Provide high-level block diagram of the spacecraft subsystems Define system and hardware requirements sufficiently to allow future subsystem architecture definition and hardware trade studies Demonstrate that the landing loads do not exceed requirements and can be supported by the structural/mechanical configuration 6

Assumptions Orbiter will supply pictures of asteroid to give targeted landing area Asteroid is only gravitational pull in area Orbiter can place lander in preferred initial placement Major Tasks 1 and 2 are most important Key asteroid physical properties a. Rotation period is 5.0 hours b. Semimajor axis: 1.5 AU c. Radius: 10 km (assume spherical shape) d. Mass is 1x10 16 kg e. Surface rock distribution is Gaussian with 0.25% probability of a 0.5 m rock in a 5 m 2 area. f. Surface gravity: m/s 2 7

Major Tasks 1) Perform a trade study of approach, landing and anchoring options which considers: a. Control system architecture b. Structural impact loads c. Anchoring mechanism d. Propulsion system options 2) Develop a spacecraft layout including a. Mechanical/structural configuration b. GN&C sensor locations and field-of-view (FOV) clearances c. Location and orientations of reaction control system thrusters d. Telecommunication component locations including antenna FOV clearances 8

Major Tasks Continued 3) Identify major hardware components of each spacecraft subsystem a. Guidance, Navigation, and Control b. Command and Data Handling c. Electrical Power d. Telecommunications e. Propulsion f. Structures and mechanisms g. Thermal control system 4) Define the key system and hardware requirements 5) Determine landing loads and associated structure/absorber sizing 9

Team Organization 10

Structures Group Responsibilities Design Structure to hold all components and protect payload package and return vehicle Perform ANSYS analysis on structural components to ensure structural stability Provide solution for thermal control needs Determine attachment method and prove effectiveness Determine Center of Mass of spacecraft 11

GN&C and Propulsion Determine proper trajectory for mission Complete trade studies for GN&C components Develop high level block diagrams for GN&C systems Provide necessary propulsion system to satisfy mission requirements Provide solution for telecommunication needs 12

Project Summary Approach asteroid from orbiter using a Hohmmann transfer and descend from 500m altitude Use pressure regulated tank with monoprop clusters Use spikes post-impact to attach to asteroid surface 13

Layout 14

Attachment Method 15

Initial Trade Study IdeaHow worksProsCons Spikes on Impact Drive stakes in when land from legs Don’t need separate propulsion system (less weight) Highest probability for success HarpoonShoot something out and reel ourselves in Uses more fuel Many mechanisms involved Possibly get tangled to roll back up. Thrusters Propulsion Only Use thrusters to hold in place Extra fuel Torque when drilling Cork ScrewScrews into ground like a cork screw into a cork Not a high resulting forceComplicated mechanisms and possibility of stirring soil instead of screwing in Spikes Post- Impact Two arms with small explosion that would rotate and come back while nailing into the ground Would work…kind of Able to retry Even more complicated than harpoon 16

Initial Trade Study Cont. ThrustersHarpoonSpikes on Impact Spikes Post- Impact Corkscrew WeightBad GoodOkayBad Mechanical Complexity GoodBadGoodOkayBad Fuel Consumption BadGood OkayGood Power Requirements GoodBadGood Bad DependabilityGoodBadOkayGoodBad Applied Loads Good BadOkayGood ReusabilityGoodBad Okay 17

Attachment Method Results 18 Spikes Post-ImpactThruster Attachment Mechanical Mass8 kg (approx)0kg Fuel Mass1.198 kg590.4 kg Total Mass9.198 kg590.4 kg

Attachment Method Overview Assume built weight for guns to be same as hilti gun (2kg ea.) Use F=m(v/t) to determine force from anchor deployment t is assumed to be 0.01 s Deploy two anchors simultaneously at 45° to reduce thruster requirements Determine required Thrust and Fuel based on Force Assume 30s to deploy all 4 anchors Use Anchor based on future research 19

Gun Use Mechanical fastening system based on Hilti Gun model DX E72.22 caliber powder actuated fastening tool Drives up to 7.1cm nails into concrete or steel 2 kg each 20

Reaction Force Ballistics information found at: Used: Remington.22 Short CB Cap Bullet Weight = 1.87 g Velocity = m/s Force from Attachment = (213.36/.01)*(1.87/1000)=39.898N per anchor 21

Spike deployment requires a 21 N force from the 4 thrusters pointing in the +Z direction using a safety factor of 1.5 (56/4=14*1.5(safety factor)=21N/per thruster) Reaction Force Continued Spacecraft +28N -28N +28N Gun #1 Gun #2 22

Future Work Make gun space worthy -space worthy construction material Test force required to insert nails -soil testing experiments Determine spike specifications -force experiments 23

Structures and Mechanisms Paul Krueger John Edquist Derek Steffes Carla Bodensteiner 24

Paul Krueger Spike Experiment Packaging 25

Force Required to hold Spacecraft 26

Shear Force on Spikes 27

Spike Force Experiment Force spike can withstand anchored in ground? Driven in at 45 degree angle 45 L x 3 D mm spike – 3.6 N normal force – 4.3 N shear force – Within Hilti gun driving capacity 28

Spacecraft Frame Basic truss properties Opening for return vehicle and drill Octagonal Shape Dimensions based on interior components – Each side 1m x 2m Constructed of aluminum and aluminum honeycomb panels 29

Packaging Trade Study IdeaHow worksProsCons Dual Shear PlateMounts electronics on flat plates that are mounted to shear plates that are then bolted to frame Strong structure, Efficient packaging, Good heat dissipation Needs custom electronics, Expensive ShelfShelf inside spacecraft where electronics are mounted Uses standard electronics, Inefficient packaging, Complicated heat dissipation Skin Panel/FrameElectronics mounted to panels that are mounted to frame Uses standard electronics, Good heat dissipation, Easy access Less rigid than Dual Shear Plate configuration 30

Packaging Trade Study Dual Shear PlateShelfSkin Panel/Frame Structure Strength HighLowMedium Heat-transferGoodOkayGood Volume neededSmallMediumSmall Needs custom electronics YesNo CostHighLow 31

Future Work Build basic mock up of feet Add mounting brackets to frame – Return vehicle, drill, fuel tank, etc Lighten spacecraft frame 32

John Edquist Landing Gear 33

Landing Gear Basic Truss – Solid bar, D=20mm Low gravity allows smaller structure Multiple designs – ANSYS prevented further testing

Landing Gear Ratcheting System Gear is spring loaded Must prevent it from bouncing off asteroid Stopper prevents craft from bottoming out

Ratchet Spring Calculations Total mass of craft: 519 kg Force from gravity = Mass*Gravity = 1.5*519 * m/s 2 = 5 N Force on impact: F=SF*Mass*Velocity/Δt =1.5*(519 kg)*(2.45 m/s)/(1 sec) = 1912 N Force per pad = (1912+5)/4 = N Force from weight negligible due to very low gravity Spring constant k=Force*Distance/Δθ = (479.3 N)*(0.385 m)/(π/6 Rad) = 352 N*m/rad (per pad) = 6.2 N*m/degree (Δθ=30⁰) Divide by 4 because 4 springs in parallel per pad: Individual spring constant: 1.5 N*m/degree Source:

Recommendations Optimize landing gear to reduce weight and areas of high stress concentration Continue design of ratchet system and work with Derek to perform ANSYS testing Perform more detailed analysis of internal components and structure for the SolidWorks model

Derek Steffes Structural Impact Loads Mass Budget 38

Structural Impact Loads 39

Structural Shape 40

Mass Budget From Design – GN&C: 21 kg – Propulsion: 26.8 kg – Telecommunication: 10 kg – Structure: kg – Thermal protection: 18.4 kg – Attachment guns/spikes: 8 kg – Propellant: 4.4 kg Given – Drill: 50 kg – Return vehicle: 200 kg Allocated – Electrical: 50 kg – Data handling: 20 kg Total mass: 519 kg Mass margin: 231 kg 41

Center of Mass 42

Structural Analysis Material Type – Aluminum Alloy 6061-T6 Source: Alcoa Engineered Products – Aluminum Honeycomb Paneling 3003 Source: Portafab FEA element type – 10-node tetragon 43

Structural Analysis Impact Load – Linear momentum-Impulse mv i +(F Imp /SF)Δt=mv f – Gravitational Force F g = SF(mg) – Total impact load F = F imp + F g Values m = 519 kg v i = 2.45 m/s v f = 0 m/s SF = 1.5 Δt = 1.0 s g = m/s 2 F Imp = N F g = -5 N F = N 44

Landing Assembly (Stress) 45

Landing Assembly (Stress) 46

Landing Assembly (Strain) 47

Landing Assembly (Strain) 48

Octagonal Structure (Stress) 49

Octagonal Structure (Stress) 50

Octagonal Structure (Strain) 51

Octagonal Structure (Strain) 52

Hilti Gun Recoil (Stress) 53

Hilti Gun Recoil (Stress) 54

Hilti Gun Recoil (Strain) 55

Hilti Gun Recoil (Strain) 56

Future Work Single-structure analysis – Requires 256,000+ elements Dynamic load analysis Spring-Ratchet system analysis Non-ideal landing conditions analysis – Single landing assembly impact 57

Carla Bodensteiner Thermal Control 58

Requirements Define high level diagram of thermal control Identify major components required Define system and hardware requirements sufficiently to allow future thermal control architecture definition and hardware trade studies 59

Passive Thermal Control – Multilayer Insulation (MLI) – Surface Finishes Active Thermal Control – Panel (strip) heaters Closed switch controlled by PRTs (Platinum Resistance Thermometers) monitoring optimal temperature for spacecraft components Used to heat hydrazine thrusters before initial burns Basic Hardware Required 60

Active Thermal Control Electronics and Thrusters PRT monitor equipment Command and Data Handling Panel Heating ON/OFF 61

Plume Protection 62

Plume Protection Trade Study MaterialBasic Information ProsCons Molybdenum Alloys Used in turbine blades Meets temperature requirements Unsure of space worthiness PICAUsed as heat shield for entry into Earth/Mars atmospheres Light weight, used in previous space missions Expensive materials 63

Plume Protection Trade Study Cont. MolybdenumPICA DensityOkayGood Space worthinessUnknownGood Temperature Appropriate Good CostsOkayBad 64

Basic Material Properties PICA (ablative material) (Phenolic Impregnated Carbon Ablators) Used on heat shield of Stardust Sample Return Capsule Designed by Lockheed Martin Density of g/cm 3 Withstand at least 2700 degrees C 65

Propulsion Nick Meinhardt 66

Propulsion Trade Study for Translational Motion Main ThrustersIsp (s)FuelOxidizer Power RequiredThrust Flight HistoryComplexity Controlled shutoff/ restart Monoprop Rocketmediumstorable N/A middle rangeextensivesimpleyes Storable Bipropmediumstorable N/A middle rangeextensivemoderateyes Solid Rocketmediumstorable N/A highextensivesimpleno Cryogenic Biprophigh not storable N/A extremely highlimitedcomplexyes Ion Engine very highstorable N/A high extremely smalllimitedcomplexyes Hall Thrusters very highstorable N/A highsmalllimitedcomplexyes Electrothermal (Arcjet)highstorable N/A highsmallextensivecomplexyes Electrothermal (Resistojet)highstorable N/A extremely highsmallextensivecomplexyes

Propulsion Options for Attitude Control Attitude ControlAdjustmentTranslation Power Required 3 Direction Rotation Rapid Change in Direction Combined Rot/Trans Professional's Opinion (Steve Lee) Monoprop Clustersfine-roughyesnoyes yes, off modulationgreat idea! Momentum Wheelvery finenoneyes no yes, with main thrust probably would use a reaction wheel instead Reaction Wheelvery finenoneyes no yes, with main thrust overkill for this application, usually used for photos 1 68

Thruster Requirements Attachment Method F required = 21N Provide attitude control capability Provide translational capability Exit mass flow rate 69

Flow Component Requirements Design Drivers: Minimal pressure drop for rated flow Monitor fuel conditions Isolate systems: multi-fault tolerance against thruster fire Provide parallel & series valve redundancy for reliability Additional affecting factors: Low power consumption (affects sizing of power supply) Low response time (affects minimum turn angle) 70

Fuel Tank Requirements - assume incompressible flow - assume conservation of mass BurnΔV or FBurn Time (s) Fuel (kg) Fuel (in 3 ) ΔV1ΔV m/s ΔV2ΔV m/s F attach 22.5 N Total

Pressure Regulated vs. Blowdown Tank Pressurization ComparisonPressure RegulatedBlowdown Component ComplexityComplexSimple Controls ComplexitySimpleComplex Pressure lossNoneFunction of propellant consumed Loss in thruster forceNoneFunction of tank pressure MassAdditional mass from pressurant monitoring components Additional mass from added tank thickness to withstand high pressures Pressurization RequiredLow  299 psiHigh  721 psi to complete mission #tanks neededOnePotentially two if not enough fuel volume available in high pressure tank Change in performance between ΔV and Attitude maneuvers NoneIsentropic expansion of pressurant results in temperature drop within the tank  temporary lull in pressure. Fits mission profileYesNo 72

Propulsion Isolation Assembly & Thruster Orientation Parallel Series 73

Fuel Lines Same length and number of 90 degree bends in lines going to each thruster. L total = 3.78m Bends = 9 74

Propulsion System Losses 20 N HYDRAZINE THRUSTER Model CHT 20 Max Tank Operating Pressure: psia Filter Pressure Drop: 5 psid at rated flow Latch Valve Pressure Drop: <1 psid at rated flow Solenoid Valve Pressure Drop: 25 psid (max) at rated flow Total System Pressure Drop: 31.6 psid Thruster Operating Pressure for Required Force: 267 psia Regulate Tank Pressure Near: 299 psia In – line pressure drop: psid at 283 K, fully developed laminar flow 75

Component Properties ComponentProductMax Operating Pressure ∆P, Flow ratePower Consumption Max Response Time Fuel Tank385.3 psia N/A Fuel Filter585.3 psia5 psid kg/s N/A Torque-Driven Latch Valve psia1 psid kg/s 11.1 watts50 ms Dual-Seat Solenoid Valve psia25 psid (max) kg/s 27 watts15 ms Service Valve650 psia30 psid kg/s N/A 20 N Thruster320 psia kg/s 27.5 watts (heater: 10 Solenoid: 17.5) - 76

Propulsion Mass Budget Component Type Manufacturer/Model NumberQtyUnit Mass (kg)Total Mass (kg)Heritage Propellant TankAKT-PSI/ I.U.E. FilterVACCO/F1D Not Listed Pressure TransducerGulton-Statham/PA DS1 Service Valve (gas)VACCO/V1E FDV DS1 Service Valve (liq)MOOG/Ref Model MOOG IR & D, PMA, NSTAR Latch ValveMOOG/Ref Model Koreasat, LM 700, Globalstar, Iridium®, ETS-8, Smart-1, MT- SAT2, Cosmo Sky Med Solenoid ValveMOOG/Ref Model GPS, ACE, CSII, Centaur, Skyret 4, COBE, Topex, SFV,DMSP, USERS, SERVIS Engines (15N req)EADS Astrium CH EURECA, HAPS, XMM, Integral, METOP, Herschel, Plank Tubing1/4" 0.028" wall tubing N/A DS1 PressurantGHe N/A DS1/MER Total DryN/A 25.24N/A Propellant loadedN2H4N2H4 N/A Requirements Total WetN/A 29.67N/A 77

Off-Modulation System Impact: Operating up to 4 thrusters at a given time to produce translational motion Max required flow rate through the latch & solenoid valves = exit flow rate = F/ve = kg/s Max required flow rate through the filter = 4 * exit flow rate = kg/s Max power consumed at once = 262 Watts Def: turn off thrusters during a ∆V maneuver to compensate for unwanted rotations that are caused by an offset in C M. 78

Future Work Model pressure control assembly & find components. Complete off-modulation dynamic model to find more accurate fuel consumption information. CFD analysis using Monte Carlo methods to determine plume impingement effects on thrust. 79

GN&C Jessica Lattimer Sam Zarovy 80

Jessica Lattimer Approach GN&C Hardware 81

GN&C Assumptions There exists space qualified versions of all devices No redundancy in system devices – single string Probability of landing on large rocks is so small that hazard avoidance control system is not needed Do not need pin point landing Asteroid is the only gravity force acting on the spacecraft Only disturbance torque acting on spacecraft is gravity Cruise stage has cameras and GN&C system to navigate and determine landing spot Cruise stage will put spacecraft in any initial orbit we desire Current project will use radar, next team will do trade study between LIDAR and radar Flight computer has enough processing power for all calculation and command execution The impulse force during landing only lasts t = 0.1 sec 82

Approach Hohmman Transfer Non-Hohmman Transfer Time OkayGood Fuel Efficiency GoodOkay Burn Complexity GoodBad 83

Free Fall Distance Farther AwayCloser Time OkayGood Crash Probability GoodBad Reposition Chance GoodBad Used a MatLab script to quantitatively analyze free fall distance 84

Approach 85

Attitude Determination Star trackerSun SensorsMagnometersGyroscopes Earth-Horizon Scanners Range of UseGoodOkayBadGoodBad ReliabilityGoodOkayBadOkayBad ErrorsGood WeightGood SizeGood PowerGood OkayGood 86

Star Tracker Galileo Avionica A-STR Autonomous Star Tracker FOV: 16.4 X 16.4˚ Power Consumption: 8.9 W at 20˚C 13.5 W at 60˚C Size: 195 (L) X 175 (W) X 288 (H) mm Mass: 3 kg 87

IMU Honeywell Miniature IMU Mass: 5 kg Size: 25 (D) X 20 (H) cm Power Consumption: 30 W 88

Radar 3 Antennas – Mass : 1 kg each – Size : 20 cm diameter x 1 cm high each Central control box – Mass: 10 kg – Size: 30 cm x 20 cm x 15 cm Based off of MSL radar design 89

Radar Field of View Antennas only objects on bottom of spacecraft Landing gear is at 45˚ angle form vertical in descent 45˚ 90

Sam Zarovy Control System 91

Thruster Logic Vehicle Model Sensor Update Guidance Algorithm + - Error Vehicle State Control Inputs Estimated Vehicle State Control Block Control Update Control System Architecture Desired Vehicle State 92

Trajectory Input 93

Control Block Attitude Error Kp_att + Attitude Update Ki_att Kd_att 1/S S + + Angular Rate Error Kp_AR + Angular Rate Update Ki_AR Kd_AR 1/S S + + Velocity Error Kp_vel + Velocity Update Ki_vel Kd_vel 1/S S + + Altitude Error Kp_alt + Altitude Update Ki_alt Kd_alt 1/S S

Vehicle Model Use thruster firing as input to force and moments equations Estimate how thrusters will change vehicle state 95

Thruster Logic Torque produced by one thruster couple firing = 48 N*m Minimum angle achievable: X-Axis = deg Y-Axis = deg Z-Axis = deg 96

Control System Architecture: Sensor Update Block Star tracker Attitude Measurement IMU - Gyros Angular Rate Measurement Filter: blend slow update rate of star tracker with fast update rate of gyros. Use highly accurate star tracker to remove gyro bias. Estimated State Filter: Use measurement to update estimated attitude and angular rate. Estimated Attitude IMU - Accels Acceleration Measurement Estimated Velocity Filter: Use measurement to update estimated velocity. Radar Altitude Measurement Estimated Altitude Filter: Use measurement to update estimated altitude. Attitude State Velocity State Altitude State Vehicle State Filter: blend Radar ground speed measurements with acceleration measurements to calculate velocity measurement. Ground Speed Measurement Angular Rate State Estimated Angular Rate 97

Flight Computer Guidance and Navigation Telecommunications Propulsion Thermal Control 98

Future Work LIDAR trade study Further refine control blocks 99

Telecommunications Ashley Chipman 100

Telecommunications Requirement: Report back to orbiter Antenna:  Conical Log Spiral (from Spacecraft Mission Analysis and Design)  7.2 GHz receive band (X-band)  Transmit 8.5 GHz Solid State Amp vs TWTA (traveling wave tube amp):  Requires more power input  SSA has a lower mass (wt limitations)  more reliable (require lower voltages) 101

Telecommunications 102

Telecommunications 103 ComponentMass (kg)Power (W) Dimensions (m) Placement Med Gain Antenna (Conical Log Spiral) Cone Diam 0.14 Tube Diam 0.03 Cone Depth 0.16 Ideally on top 15 degree FOV Transponder -- receiver -- transmitter x 0.33 x 0.07Where fit ** includes solid state amp Comm. Detector Unit N/A Same as main command unit RF switches/cables (est. 10% of total) 0.908N/A TBD Where needed Telemetry Conditioner Unit x x Where fit Total Mass: kgTotal Power: 66.3W

Field of View Star Tracker Antenna 1.82 meters 104

Thank you! Steve Lee – JPL Professor Garrard Professor Flaten Professor Ketema 105

Questions? 106