ExoMars 2016 CDR Presentation David Bartolo Shazib Elahi Aaron - - PowerPoint PPT Presentation

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ExoMars 2016 CDR Presentation David Bartolo Shazib Elahi Aaron - - PowerPoint PPT Presentation

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ExoMars 2016 CDR Presentation David Bartolo Shazib Elahi Aaron Tun Junyi Zhang Brief Recap - ExoMars 2016: ESA + Roscosmos - Joint operation aiming to investigate the Martian atmosphere and surface, and to gain more experience in

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ExoMars 2016

CDR Presentation

David Bartolo Shazib Elahi Aaron Tun Junyi Zhang

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Brief Recap

  • ExoMars 2016: ESA + Roscosmos
  • Joint operation aiming to investigate the Martian atmosphere and surface, and to gain more

experience in preparation for future missions

  • Trace Gas Orbiter (TGO)
  • Schiaparelli Entry, Descent, and Landing Demonstrator Module (EDM)
  • TGO: science operations in orbit, EDM: demonstrate ESA’s ability to land on surface of Mars,

some scientific objectives as well

  • Analyses focusing on EDM
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Subsystems Breakdown: After

Thermal Protection Systems External Structure Scientific Payloads Avionics Systems Landing Systems Propulsion System Mortar Deployed Parachute System 3 Clusters of 3 Hydrazine Pulse Engines Crushable Landing Structure Communication System UHF System Guidance, Navigation, and Control Power Systems Radar Doppler Inertial Measurement Unit (IMU) Back Shell Surface Payload Rechargeable Batt. Rechargeable EDL Battery Solar array Surface Operations Rechargeable Batt. Descent Camera (DECA) Accelerometers Gyroscopes Electronic Units (2x) Antennas (4x) Sun Sensors (2x) Aluminum with Carbon Fiber Reinforced Polymer Skin Back Shell Front Shell Ablative Material DREAMS SIS MicroARES MetHumi MetBaro MetWind MarsTem

Schiaparelli EDM

Software Sanity Check Predetermined Failsafe Calculations

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Subsystems Breakdown: After

Thermal Protection Systems External Structure Scientific Payloads Avionics Systems Landing Systems Propulsion System Mortar Deployed Parachute System 3 Clusters of 3 Hydrazine Pulse Engines Crushable Landing Structure Communication System UHF System Guidance, Navigation, and Control Power Systems Radar Doppler Inertial Measurement Unit (IMU) Back Shell Surface Payload Rechargeable Battery Rechargeable EDL Battery

Solar Array

Surface Operations Rechargeable Battery Descent Camera (DECA) Accelerometers Gyroscopes Electronic Units (2x) Antennas (4x) Sun Sensors (2x) Aluminum with Carbon Fiber Reinforced Polymer Skin Back Shell Front Shell Ablative Material DREAMS SIS MicroARES MetHumi MetBaro MetWind MarsTem

Schiaparelli EDM

Software Sanity Check Predetermined Failsafe Calculations

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Software Sanity Checks

  • Implement a Software Sanity Check that monitors the communication between the

Inertial Measurement Unit (IMU) and the Radar Doppler Altimeter (RDA).

  • The Sanity Check will focus on the attitude, altitude’s magnitude and change over time,

and the vertical acceleration (comparing with the Martian gravity).

  • The Sanity Check will remain on standby in parallel, analyzing the receiving data, for if

the readings of the IMU or the RDA become unreliable.

  • Exomars’ main flaw resided in the communication of the received information, and the

internal systems inability to respond to the inaccurate readings; however, despite the inertial measurement requiring a higher saturation limit, modeling the exact effects of the IMU reading when entering the Martian atmosphere and deploying the parachute is not feasible nor does the Avionics Test Bench simulator stimulate the IMU navigation function.

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Specific Quantitative Requirements Targeted for Analyses Analysis Quantitative Requirements Power Analysis

  • 25% Power Margin

Solar Array Design Analysis

  • Natural frequencies in the 5.90 - 7.24 Hz range (for modes 1-5)

Mass, Volume, Stability Analysis

  • Mass & Volume: less than 10% increase
  • Stability: longitudinal variation less than 12.5 degree
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Power Analyses

Goal of analysis

  • Determine the size of the solar array
  • Determine required battery capacity
  • Choose power regulation techniques and regulator
  • Analyze the dissipated heat from new power source
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Power Generation and Sources

Assumptions, Principles, & Methods

  • Assume required power consumption to be 77 Watts for both day and night cycles.
  • A 25 % power margin.
  • Method to determine the required solar array power output:
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Power Generation and Sources

Assumptions, Principles, & Methods

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Power Generation and Sources

Assumptions, Principles, & Methods

  • Determine the power that the solar array can produce
  • Determine the orbital solar irradiance:
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Power Generation and Sources

Assumptions, Principles, & Methods

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Power Generation and Sources

Assumptions, Principles, & Methods

  • Determine the Mars surface solar irradiance:
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Power Generation and Sources

Assumptions, Principles, & Methods

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Power Generation and Sources

Assumptions, Principles, & Methods

  • Determine the beginning-of-life (BOL) power production
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Power Generation and Sources

Assumptions, Principles, & Methods

  • Determine the Lifetime degradation due to dust

accumulation and radiation: D = % dust accumulation + % radiation

  • Assumed degradation of 0.028 % per sol for % dust

accumulation.

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Power Generation and Sources

Assumptions, Principles, & Methods

  • Calculate EOL power production:
  • Calculate the Solar Array Area:
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Power Generation and Sources

Results: Solar cell comparison :

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Power Energy Storage

Assumptions, Principles, & Methods

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Power Energy Storage

Assumptions, Principles, & Methods

  • Concluded with selecting Li-ion cell.
  • Size a secondary battery capacity, identify the parameters
  • Assumed a DOD of 50% because of small life cycles (~687

cycles)

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Power Energy Storage

Results

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Power Regulation and Control

Assumptions, Principles, & Methods

  • We will use DET
  • Disadvantages of PPT:

○ Additional weight ○ More complex system ○ Less efficient ○ Cost more

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Power Regulators

Assumptions, Principles, & Methods

  • We will use the Switch-Mode Regulators
  • Disadvantages of Linear Regulators:

○ Poor efficiencies

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Power Distribution And Heat Dissipation

P = i * V = i^2 * R,

  • Peak current: Ip = 2.5 [A]
  • Standard of Allowable Internal Resistance: R = 0.276

[Ohms]

  • Power Dissipated from heat: (2.5 [A])^2 * 0.276 [Ohms]

= 1.725 [Watt]

  • Power Capacity of total battery: 2569.32 [W* h]
  • From power dissipated, this results in an increases of

4.18 oC.

  • Required power consumption of the EDM: P = 77 [W*h]
  • Therefore, the actual power at the system’s disposal:

2569.23 - 1.725 = 2567.51 [W* h]; in comparison the EDM requires approximately 77* 12.3 = 924 [W* h] for a Martian night (12.3 h).

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Power

Discussion: Confidence, Weaknesses, Importance of Results

  • The estimated solar array size is reasonable, as well as the battery capacity.
  • The weakness of our analysis is that we based the solar array size on the

average solar irradiance throughout the year, which means it did not account for global and local dust storms.

  • The importance of the results suggest that we can move forward in the

process of adding the solar array, which can provide the required power extending the life of the EDM.

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Test Plan: Power

  • Electromagnetic Interference
  • Electromagnetic Compatibility
  • Thermal Cycling for the Battery
  • High-/-Low-Voltage Limits
  • Thermal Limits (Battery)
  • Performance
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Solar Array Design Analysis

Minor Analysis I

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Goal of Analysis

  • Ensure best possible solar array design is selected to meet power

requirements of Schiaparelli lander

  • Identify natural frequencies and mode shapes of deployed & stowed array
  • First step before assumptions, principles, and methods can be determined:

Trade Study

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Assumptions, Principles, and Methods

  • Simplifying assumptions:
  • Structural analysis done separately: for stowed and

deployed positions

  • Considerably simplifies modeling process in

Solidworks

  • Mass of solar cells is negligible and will be ignored
  • Array assembly components and materials will be

modeled as follows:

  • Substrate (gores): vectran (20% void)
  • Supporting structure (spars): aluminum

(6601-T6)

  • Assumptions are based on those made in a research

paper for finite element analysis on a larger variant of the UltraFlex design

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Assumptions, Principles, and Methods

  • Principles and Methods
  • Trade study to select best

design concept

  • UltraFlex
  • CAD modeling - Solidworks
  • Finite element analysis
  • Modal analysis -

Solidworks

  • Resonant and natural

frequencies

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Ultraflex Solar Array

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Math and Models

  • Ultraflex shape is decagonal
  • Area of decagon
  • Calculate side length a based on area

from power analysis

  • Then determine spar length and length of

centerline

  • Centerline = router + rinner
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Math and Models

  • To calculate rinner and router:
  • Referenced values from Semke et al.
  • Outer radius = 101.6
  • Inner Radius = 3.81
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Math and Models - Deployed Array

Models created in Solidworks Vectran gore Aluminum spar Perspective view Top-down view

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Math and Models - Stowed Array

Models created in Solidworks

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Math and Models - Finite Element Models

Models created in Solidworks

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Results: Deployed Array Analysis - Modes Shapes

Mode Shape 3 Mode Shape 2 Mode Shape 5 Mode Shape 4 Mode Shape 1

Frequency/Mode Number Rad/sec Hertz Seconds 1

26.197 4.1694 0.23984

2

29.48 4.6919 0.21314

3

32.173 5.1205 0.1953

4

35.456 5.643 0.17721

5

40.508 6.4471 0.15511

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Results: Deployed Array Analysis - Modes Shapes Animation

Fixed at one spar to simulate boundary conditions of deployed array

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Results: Stowed Array Analysis - Modes Shapes

Mode Shape 3 Mode Shape 2 Mode Shape 5 Mode Shape 4 Mode Shape 1

Frequency/Mode Number Rad/sec Hertz Seconds 1

57.973 9.2267 0.10838

2

58.235 9.2683 0.10789

3

58.521 9.3139 0.10737

4

58.523 9.3143 0.10736

5

58.528 9.3151 0.10735

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Discussion: Confidence, Weaknesses, Importance of Results

  • Results for deployed array modes were close to expected range
  • More information is required to determine the accuracy of each analysis,

especially the stowed configuration

  • Weaknesses: many simplifying assumptions that do not take into account many of the

components

  • These results are a rough estimate of what actual characteristics of the array would be
  • Importance of Results
  • The results for the deployed array are promising
  • Shows that this is a viable option for recharging the batteries onboard and powering the

Schiaparelli

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Test Plan to Check Analysis

1. Required data

a. Multiple deployed dynamics first mode frequency tests b. Measured natural frequency (ambient air) with off-loader to account for lower Martian gravity

2. Facilities, sensors, types of measurements

a. Johnson Space Center i. Vibration Laboratories for fixed-base modal testing ii. Active Response Gravity Offload System (ARGOS) b. Accelerometers needed to record data for determining modal frequencies

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Test Plan to Check Analysis

  • Shortcomings, Potential Impact of Errors/Oversight
  • It may be difficult or impossible to test the array with all environmental factors simultaneously,

so they must be done separately

  • Errors/oversight could cause the array to fail and end the scientific mission early or

prevent it from starting at all

  • Significance, Impact to Project if Requirements Not Met
  • If testing shows that requirements are not met, then trade study will have to be revisited
  • A new design concept will have to be picked
  • Project will be delayed and alternative design may impact other subsystems considerably,

causing even more delays

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Mass | Volume | Stability

Minor Analysis II

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Total Mass & Distribution Change Considerations

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Numbers: Total Mass Increase Solar Array

Data from Northrop Grumman: UltraFlex Solar Array System: 150 W/kg BOL Data from ABLE Engineering: 27% TJ cells: > 150 W/kg BOL Our choice: approx 30% ZTJ cells Consider industry data and a 1.15 fitting factor Specific Mass: 130 W/kg BOL

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Numbers: Total Mass Increase Solar Array

Total Power Needed: 282 W Power from one array: 141 W Mass of one array: 1.085 kg Total mass increase: 2.17 kg

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Numbers: Mass Increase Battery

From previous: 3.44 kg

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Numbers: Mass Increase Heat Shell

Data from Northrop Grumman and ABLE Engineering: Power per cube meter: 40 kW/m^3 Power needed of one array: 282/2 = 141 W Volume needed of one array: 141/40000 = 0.0035 m^3 Area of one array: 2.7/2 = 1.35 m^2 Area of folded one array: 1.35/20 = 0.0675 m^2 Thickness of one array folded: 0.0035/0.0675 = 5.19 cm

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Numbers: Mass Increase Heat Shell

Each solar array is placed in a direction align with the side of the heat shell The diameter of the heat shell will increase approx 10 cm Detailed specifications of heat shell is unavailable Diameter 2.4 m; Mass 80 kg; but thickness unknow Area increases: 0.55% Approximate heat shell mass increases: 0.4 kg

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System Mass Increase

Solar Array: 2.17 kg Battery: 3.44 kg Heat Shell: 0.4 kg Total: 6.01 kg

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Current System Capability Assessment

Consider really small mass increase Consider common industry practice of leaving a fitting factor Mass increase is negligible

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Key Points Check: Attitude

Attitude Control: EDM & TGO 4332 kg Solar arrays are attached on EDM, which is attached on one end of TGO. After installation of solar arrays, together with a size increase of heat shell, the center of mass will move towards EDM end. However, detailed mass distribution information is unavailable Estimation: 0.14% will not create an influence to the system

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Key Points Check: Landing Thrusters

Mars Gravity: 3.711 m/s² Mass: 6.01 kg F = 22.3031 N 9 hydrazine engines, 400N thrust each, a total of 3600N Initial EDM design: 600*3.711 = 2226.6 N Rough Estimate: Enough margin to accommodate increased mass Detailed calculations given in later slices

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Volume Increase

Launching Vehicle: Briz-M Stage Outer diameter more than 4 meters Consider 5 cm thickness of the solar array Consider Briz-M has variations and can be customized Payload space should not be a constraint

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Stability: Before Heat Shell Detachment

Two solar arrays arranged in symmetry Entry protection cover volume proportionally increase Center of mass does not change Aerodynamic center does not change Stability does not change

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Stability: After Heat Shell Detachment

Wind tunnel test to determine new stability conditions Exact composition of Mars atmosphere is still unclear Wind tunnel test will only give estimation Wind may come from any direction; model 10 directions; 36 degree each Evaluate new stability conditions of each of the 10 situations

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Test Plan: Mass, Volume, Stability

Mass: Measure the actual mass of each component Volume: Measure the actual volume after modification Stability: Wind tunnel test, velocity 250 km/h to 4 km/h (propulsion system ignition velocity to propulsion system off velocity) Check longitudinal variation (<12.5 degree)

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Discussion & Conclusion: Mass, Volume, Stability

Fitting factor of 1.15 was given for initial solar array mass calculation, so everything is an overestimation Unlike aircraft stability, in which one has control derivatives as well as static and dynamic stability, EDM stability is purely experimental determined From mass and volume perspectives, adding solar arrays is feasible Actual wind tunnel test is needed to give a conclusion on stability

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Additional Mass Effect on Propulsion System

  • 3 clusters of 3 hydrazine engines (400 N each), operated in pulse-modulation,

amounting to a total propellant mass: Mp = 46.2 [kg]

  • Using the ideal Change in Velocity for a Rocket equation (simplifying assumptions of

neglecting losses due to Drag and Gravity)

  • Also, the relation of total impulse over the weight
  • f propellant used:

Isp= Veq/ go

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Propellant Required due to Added Mass Calculation

  • Concluded from ESA’s guideline Entry, Descent & Landing:
  • MpTot = 46.2 [kg], MEDM = 280 [kg], Mo = 326.2 [kg],

Isp = 212 - 220 [s], go = 3.711 [m/s^2], ΔV = 270 - 4 [km/h] Mo/ [1 - 1/ e^(ΔV/Isp /go)] = MP

  • Mo = 326.2 [kg], MP = 29.24 [kg], MP/MPTot *100% = 63.3%
  • Mo2 = 326.2 + 6.01 [kg], MP2 = 29.78 [kg], MP2/MPTot*100% = 64.5%
  • Although simplifying assumptions were made to treat the rocket as ideal, conservative

values were used for this analysis such that it is reasonable there is enough excess propellant for the extra mass.

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Conclusion and Discussion

  • With the addition of a solar array, the Schiaparelli is now able to recharge its batteries and extend its mission life from a few days to a whole

Martian year

  • Strengths
  • Mission life extended significantly
  • Fitting factor 1.15 was given for initial solar array mass calculation; adding solar array is feasible from mass and volume perspective
  • Weaknesses
  • Structural analyses based on many simplifying assumptions
  • Detailed mass distribution unknown, mass distribution validity based on estimation
  • Solar array sizing is based on average solar irradiance throughout the year.
  • The idea to implement a solar array seems to be viable, though some open questions remain:
  • Stability testing results from wind tunnel test (after heat shell detachment)
  • Lessons learned:
  • Modal analysis can become very complicated
  • Rapid prototyping results can be inaccurate
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References

  • “UltraFlexTM Solar Array Systems: Breakthrough Performance, Competitive Cost, Space Flight Proven” Northrop Grumman Corporation,
  • Doc. FS007_15_1
  • “Exomars Media kit: EUROPE’S NEW ERA OF MARS EXPLORATION” ESA Doc. SCI-A-COEG-2016-001; March 2016
  • “UltraFlex Solar Array: Enabling Light Weight Low Volume Technology” ABLE Engineering,
  • “The Dynamics Characteristics of a Lightweight Deployable Solar Array” (research paper)
  • “Ultraflex-175 on Space Technology 8 (ST8) - validating the next-generation in lightweight solar arrays” (research paper)
  • “Advanced III-V Multijunction Cells for Space” King et al (research paper)
  • “Solar Power Technologies for Future Planetary Science Missions” NASA Research Paper
  • NASA Systems Engineering Handbook
  • Al 6061-T6: Department of Defense. MIL-HDBK-5J Metallic Materials and Elements for Aerospace Vehicle Structures. Department of

Defense, 2003.

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References

  • https://www.nasa.gov/centers/johnson/engineering/integrated_environments/active_response_gravity/
  • https://www.nasa.gov/centers/johnson/pdf/639713main_Vibration_Testing_FTI.pdf
  • https://www.jpl.nasa.gov/nmp/st8/tech/solar_array3.html#
  • http://rascal.nianet.org/wp-content/uploads/2015/07/Ultraflex-Solar-Array.pdf
  • https://www.northropgrumman.com/Capabilities/SolarArrays/Documents/UltraFlex_Factsheet.pdf
  • ESA EDM http://exploration.esa.int/mars/47852-entry-descent-and-landing-demonstrator-module/
  • Liquid Crystal Polymer Poissons’ ratio:

https://books.google.com/books?id=BfbiBwAAQBAJ&pg=PA170&lpg=PA170&dq=vectran%27s+poisson%27s+ratio&source=bl&ots=KCZCa _KTio&sig=ACfU3U1Z0ET25QxhjUjUrFz01dUCTgAc4A&hl=en&sa=X&ved=2ahUKEwi2t_uIgZPhAhVWqp4KHdMEAGoQ6AEwBXoECAkQ AQ#v=onepage&q=vectran's%20poisson's%20ratio&f=false

  • Thruster Isp: http://www.space-propulsion.com/spacecraft-propulsion/hydrazine-thrusters/400n-hydrazine-thruster.html
  • https://www.e3s-conferences.org/articles/e3sconf/pdf/2017/04/e3sconf_espc2017_13012.pdf
  • https://solaerotech.com/wp-content/uploads/2018/04/ZTJ-Datasheet-Updated-2018-v.1.pdf
  • https://ieeexplore.ieee.org/stamp/stamp.jsp?arnumber=4161575
  • https://trs.jpl.nasa.gov/bitstream/handle/2014/38400/05-3884.pdf?sequence=1&isAllowed=y
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References

ASI, Redazione. “Schiaparelli IMU Saturation.” A.S.I. - Agenzia Spaziale Italiana, 25 Nov. 2016, www.asi.it/en/news/schiaparelli-imu-saturation. Blau, Patrick. “Schiaparelli EDM.” Spaceflight101, spaceflight101.com/exomars/schiaparelli-edm/. Dunbar, Brian. “JSC Engineering - Active Response Gravity Offload System.” NASA, NASA, www.nasa.gov/centers/johnson/engineering/integrated_environments/active_response_gravity/. “ESA PR 07-2016: ExoMars on Its Way to Solve the Red Planet's Mysteries.” ESA - Robotic Exploration of Mars, exploration.esa.int/mars/57619-exomars-on-its-way-to-solve-the-red-planets-mysteries/. “ExoMars Trace Gas Orbiter and Schiaparelli Mission (2016).” ESA - Robotic Exploration of Mars, exploration.esa.int/mars/46124-mission-overview/. Jones, et al. “A High Specific Power Solar Array for Low to Mid-Power Spacecraft.” SAO/NASA ADS: ADS Home Page, 1 May 1993, adsabs.harvard.edu/abs/1993sprt.nasa..177J. “Schiaparelli: the ExoMars Entry, Descent and Landing Demonstrator Module.” ESA - Robotic Exploration of Mars, exploration.esa.int/mars/47852-entry-descent-and-landing-demonstrator-module/. “Timeline for ExoMars 2016.” ESA - Robotic Exploration of Mars, exploration.esa.int/mars/57607-timeline/. Tolker-Nielsen, Toni. EXOMARS 2016 - Schiaparelli Anomaly Inquiry. European Space Agency, 2017, EXOMARS 2016 - Schiaparelli Anomaly Inquiry. Wertz, James Richard, et al. Space Mission Engineering: the New SMAD. Microcosm Press, 2015.