Senior Design - Europa Lander Mission Presentation

PDR for The Richter Program’s “Europa

CT Scanning” RFP

May 27th, 2015 – Northrop Grumman Corp.

Kronus Inc. Team MembersMichael Corpuz (Team Lead)

Randall C. Acosta (Deputy)

Omar Alhassen

Matt Bergman

Brendan J. Clarke

Frank Garcia

Kasbar Gulbenli

Jeremiah Kho

Sean Matthews

Juan Sanchez

Aerospace Department of California State Polytechnic University, Pomona5/27/15 1

Topics Covered

• Spacecraft Con Ops and Design

• Scientific Payload

• Spacecraft Subsystems

• Mass Tables

• End of Mission

• Preliminary Cost

• Compliance Matrix

5/27/15 Michael Corpuz 2

Europa

Earth

Mission Design Summary

Time

Launch Vehicle: SpaceX Falcon HeavyLaunch Payload Mass: 13,500 kg (estimated)Launch Window: Feb. 27, 2020 to Mar. 26, 2020

Venus Earth Jupiter

Ganymede Callisto

Deployment of 8 Landers: Mar. 2026

June 2020

April 2021

June2024

Venus-Earth Gravity Assist Jovian Satellite Tour

Europa

2026

2020

Spacecraft Deployment

5/27/15 3Matt Bergman

Earth to Jupiter: Trajectory Options

VEGA & VEEGA Trajectory:• VEGA:

– High ΔV trajectory

– Short time of flight

– Satisfies the RFP 2026 arrival date.

• VEEGA: 2nd Earth flyby reduces Jupiter orbit insertion maneuver by 320 m/s.

– Reduction of total ΔV by 542 m/s.

– Increase in time of flight by 1.35 years.

– Does not satisfy RFP requirement of 2026 landing date.

Launch Vehicle Selection:• VEGA: due to high spacecraft wet mass, SpaceX

Falcon Heavy [1].

• VEEGA: Delta IV Heavy with the reduction of 8 landers to 7.

Earth Gravity Assists

Ve

nu

s Gravity A

ssists

Mars Gravity Assists

Trajectory to Jupiter

Earth Launch

Date

Maximum Launch C3 (km2/s2)

Launch Vehicle Choice

Launch Vehicle Payload Mass

(kg)

Spacecraft Wet Mass

(kg)

Numberof Europa Landers

Europa Arrival Date

Total Time of Flight (years)

Total Mission ΔV (m/s)

VEGA 3/18/2020 13 Falcon Heavy

13,500 [1] 9,780 8 2/1/2026 5.9 2952

VEEGA 3/11/2020 13 Delta IV Heavy

7,920 7,600 7 5/29/2027 7.25 2410

5/27/2015Matt Bergman 4

Earth to Jupiter: Cruise NavigationGround-Based Navigation:• Goal: correct delivery errors & re-

optimize interplanetary trajectory after each planetary gravity-assist to minimize ΔV.

• X-band coherent 2-way Doppler and range data.

– Radiometric data to measure position and velocity of spacecraft.

– Before 2021 Earth flyby: primarily with LGA, while HGA and solar arrays remain sun-orientated.

– After 2021 Earth flyby: primarily with HGA.

• Planetary flybys: optical navigation (with Narrow Angle Camera) to improve delivery accuracy.

Event Planet Date Details

Earliest Launch

Earth 2/27/2020C3 = 13 km2/s2

Falcon Heavy Payload Mass = 13,500 kg

Optimal Launch

Earth 3/18/2020C3 = 11.9 km2/s2

Declination = 2.67°Departure V∞= 3.45 km/s

Flyby 1 Venus 7/1/2020Transfer Time = 105 daysAltitude = 21,935 kmDeparture V∞= 6.41 km/s

Flyby 2 Earth 4/28/2021Transfer Time = 301 daysAltitude = 1,292 kmDeparture V∞= 9.72 km/s

Arrival Jupiter 6/11/2024Transfer Time = 1,140 daysAltitude = 15 RJ

Arrival V∞= 6.40 km/s

Time of Flight to Jupiter 4.3 years

Critical Dates for VEGA Trajectory:

Matt Bergman 5

Jupiter to Europa: Trajectory OptionsJovian Tours: Banzai Pipeline [2] 12-L1 [3]

Total ΔV (m/s) 2,390 2,100

Total Radiation Dose (krad) 89 124

Total Time of Flight (months) 20 16

Orbital Trim Maneuvers 7 4

Number of Satellite Flybys 17 12

Lowest Flyby Altitude (km) 100 (Ganymede) 120 (Ganymede)

Last Flyby to EOI Δt (days) 3 2

Mission Tour: Banzai Pipeline• 10 Calisto and 7 Ganymede flybys.

• Low tour radiation dose (no Europa or Io flybys).

• On average, higher flyby altitudes than 12-L1.

• 3 critical flybys (below 500 km altitude), compared to 5 for the 12-L1.

• Larger flyby separation time.

• 500 km Ganymede flyby before Jupiter Orbit Insertion.

5/27/2015 Matt Bergman 6

Jupiter to Europa: Tour NavigationGround-Based Navigation:• Flybys (G0-C12) with a last-encounter time ≥ 14 days will

be ground-based.

• Based on Cassini’s 3-maneuver navigation strategy:

– 1) Cleanup maneuver: 3 days after encounter.

– 2) Targeting maneuver: near apoapsis.

– 3) Approach maneuver: 3 days before encounter.

Autonomous Navigation (AutoNav):• Reliant on the optical navigation system (with the

Narrow Angle Camera).

• Based on EPOXI, STARDUST, and Deep Space 1 heritage.

• Last 4 flybys (C13-G16): flight separation time between each encounter is less than 14 days.

– Can be preformed by ground utilizing only one maneuver.

– Or, on-board autonomous navigation for multiple-maneuver processing and state estimation.

Description Event Date Δt (days)Altitude

(km)Initial

Ganymede Flyby

G0 6/11/2024 0 500

Jupiter Orbit Insertion

JOI 6/11/2024 0

Perijove Raise PJR 9/4/2024 85

First Tour Flyby G1 12/8/2024 95 629

First Orbital Trim Maneuver

OTM1 1/10/2025 33

Lowest Altitude Flyby

G2 2/16/2025 37 100

Double FlybyG4 4/22/2025 65 643C5 4/26/2025 4 221

Last Ground-Based Flyby

C12 12/31/2025 249 1185

Autonomous Navigation

C13 1/12/2026 12.5 6681C14 1/20/2026 8 6563G15 1/26/2026 5.5 5182G16 1/29/2026 3 1751

Europa Orbit Insertion

EOI 2/1/2026 3 100

Critical Dates for Jovian Tour:

3σ radial uncertainty for the 100 km orbit insertion[4] ≈ 54 km.5/27/2015 Matt Bergman 7

Mission Design Summary

Maneuvers: ΔV (m/s)

Trajectory Correction Maneuvers 90

Deep Space Maneuver 100

Jupiter Orbit Insertion 1125

Perijove Raise Manuever 110

Tour Orbital Trim Maneuvers 223

Tour Statistical Maneuvers 32

Europa Orbit Insertion (100 km Orbit) 1202

Europa Orbit Maintenance (120 Days) 70

Total 2952

Earth to Jupiter Trajectory:• Venus-Earth Gravity Assist (VEGA).• One month launch window: 2/27/2020 to

3/26/2020.• Launch vehicle: Falcon Heavy.• Jupiter arrival date: 6/11/2024.• Jupiter orbit insertion into large elliptic 15

by 242.5 RJ orbit.

Jupiter to Europa Tour:• Based on the Banzai Pipeline [3] tour.• Autonomous navigation required.• Europa arrival date: 2/1/2026.• Europa orbit insertion into 100 km polar

orbit.• At Europa: 120 day mission.

• 30 day mapping phase.• 90 day science phase.

Total Mission ΔV Budget:


Europa Radiation Shielding


Orbiter Radiation Design Factor: 3120 kradLander Radiation Design Factor: 1134 krad

Radiation Vaults:• Electronics in vault accumulate 50 krad.

• 100% design margin• Orbiter & lander electronics rated

for 100 krad total dose.• Lander payload (camera &

seismometer) accumulates 15 krad.

Orbiter and Lander Radiation Vaults & Coverglass Thickness

Orbiter SpectroLab UTJ Solar Cells Lander ATK TJ Solar CellsCoverglass

(mils)120 Day Fluence

(rad)Power

Attenuation42 m2 Mass

(kg)120 Day Fluence

(rad)Power

Attenuation17.1 m2 Mass

(kg)3 mils 9.44E+15 28% 8.2 9.33E+15 28% 3.36 mils 2.01E+15 18% 16.3 1.91E+15 18% 6.6

12 mils 4.86E+14 12% 32.6 3.91E+14 11% 13.320 mils 2.09E+14 9% 54.4 1.26E+14 8% 22.230 mils 1.40E+14 8% 81.6 6.72E+13 6% 33.2

Material:Orbiter Vault 1

Mass (kg)Orbiter Vault 2

Mass (kg)Lander Vault

Mass (kg)Lander Camera

Shielding Mass (kg)Lander SeismometerShielding Mass (kg)

Aluminum 100.2 20.7 12.3 7.3 17.1

Copper-Tunsten(CW80)

91.5 18.9 15.6 10.9 25.6

Aliminum-Tantulumbi-layer

107.7 25.0 18.5 11.7 27.5

Titanium 95.7 19.8 13.8 10.9 25.6

Tantulum 108.3 22.4 17.5 10.9 25.5


First Month at Europa: Mapping (1/2)Orbital Characteristics:• 100-km near-polar, near-circular orbit.

• Global coverage of Europa, including poles.• Near-dawn/dusk (Full Sun).

• Optimal sunlight for solar arrays.

Mapping Phase: Overview• In 8 days: the WAC will provide a global coverage mosaic

(150 m/pixel) of Europa.• Would ensure global characterization of landforms for

general landing site area evaluation.• In 22 days: The NAC will provide 8 high-resolution (1

m/pixel) landing site mosaics.

100 km

Wide Angle CameraFOV: 61.3°Narrow Angle CameraFOV: 2.85°

Europa

Orbit geometry as seen from Earth

NAC Landing Site Coverage

5/27/15 11Brendan Clarke

First Month at Europa: Mapping (2/2)

Mapping Phase: Purpose• Critical for landing mission success.• Global search for safe landing sites

• Areas with most solar illumination.• Areas with least amount of sloped

terrain.• Update Europa’s characteristics: shape

(volume), albedo, rotation rate, mass, and gravity values.

• Update landing trajectory based on landing site selection.

• Post-landing images provide lander geolocation for seismic data and lander imaging correlation.

Landing Site Selection:• Hazard topology best detected at high

incidence angles.• Solar incidence angle ≈ 90°.

• Seismologically interesting landing sites: scientific desirability vs. technical feasibility. Number of coverage passes in 22 days: 16

Mean coverage duration: 2 minutesTotal coverage duration: 43 minutes(For polar lander: 90 coverage passes)

r = 137 km

Orbiter Ground Track Path

Narrow Angle CameraCoverage Area: 250 km2

• NAC’s average coverage area per landing site (in 22 days): 7620 km2.

5/27/1512

Brendan Clarke

Landing Sequence (1/4)


Deorbit Burn• 102 sec duration• ΔV = 800 m/s• -15.2° FPA at end




Coast• 84 sec duration




Coast• 84 sec duration

Gravity Turn Burn• 338 sec duration• ΔV = 660.3 m/s


5/27/15 Brendan Clarke 16

10 MeV1 MeV

100 keV

10 MeV

1 MeV

100 keV

Orbiter

Lander

Design Configuration

5/27/15 17Kasbar Gulbenli

Made in Solidworks

Deployed Configuration2.5m High Gain Antenna

8x Landers2x Hydrazine + NTOPropellant tanks

4x Solar Array

Helium Pressurant Tank

2x TR-308 Main Engine

(4x) 3 ACS Thrusters

4 x Low Gain Antenna

5/27/15 18Kasbar Gulbenli Made in Solidworks

2x Orbiter Camera

No Spin

Spin

Top View

Stowed Configuration

Falcon Heavy Payload Fairing

Front View

5/27/15 19Kasbar Gulbenli Made in Solidworks

13.9 m

5.6m

4.6m

4.3m

Static Envelope 0.3mDynamic Envelope 0.2m

Orbiter Stowed Configuration 3-View


Spin

No Spin

Made in Solidworks

Ix = 634,000kg-m2

Iy = 1,776,000kg-m2

Iz = 1,776,000kg-m2

C.G:X = 0.00mY = 1.72mZ = 0.00m

Spin

Orbiter Un-Stowed Configuration 3-View


Made in Solidworks

Ix = 640,000kg-m2

Iy = 1,783,000kg-m2

Iz = 1,783,000kg-m2

C.G:X = 0.00mY = 1.70mZ = 0.00m

Lander Stowed Configuration

Seismometer + Camera Payload2x Folding Solar Arrays

4x Folding Legs

4x NTO + Hydrazine Propellant Tanks

(2x) 5 ACS Thrusters

4x R-1E Main Engine

2x Helium Pressurant Tank

2 ACS Thrusters


Made in Solidworks

2x Altimeter

2x Lidar

Doppler

Lander Stowed Configuration 3-View


Made in Solidworks

C.G:X = 0.00mY = 0.29mZ = 0.00m

Ix = 41.19kg-m2

Iy = 48.91kg-m2

Iz = 72.62kg-m2

Lander Un-Stowed Configuration 3-View


Made in Solidworks

C.G:X = 0.00mY = 0.29mZ = 0.00m

Ix = 41.19kg-m2

Iy = 48.91kg-m2

Iz = 72.62kg-m2

FEM/FEA

Buckling Analysis on Orbiter top busFEM/FEA done in Siemens Femap with NX Nastran


Structural AnalysisStructure: Max

Bending Stress (MPa)

Margin of Safety

Max Compressive Stress (Mpa)

Margin of Safety Max Deflection (mm)

Orbiter Top Bus 232 0.39 479 0.70 3

Orbiter Bottom Bus

149 1.17 706 0.15 5

Lander Structure

161 1.01 383 1.13 2

Structure: Dimensions Thickness(mm)

Axial Force (N) Lateral Force (N)

Orbiter Top Bus R= 0.06mTubing

10 9800 14,400

Orbiter Bottom Bus R= 0.08mTubing

12 252,000 10,600

Lander Structure 4mm L-Plate and Flat Plate

2 30,219 19,000

Factor of Safety 1.55/27/15 26Kasbar Gulbenli

Modal Analysis (1/2)Mode Upper Bus f (Hz) Max Displacement (m)

(no damping)Max Displacement (m)(with damping)

1 1463 1.55 0.062

2 1465 1.55 0.062

3 3358 1.26 0.050

4 3361 1.26 0.050

5 4757 1.20 0.048

6 9757 1.19 0.047

7 9761 1.20 0.048

8 10562 1.21 0.048

9 12469 1.21 0.048

10 12475 1.21 0.048

5/27/15 27Kasbar GulbenliCritical Damping 0.02

Modal Analysis (2/2)Mode Lower Bus f (Hz) Max Displacement (m)

(no damping)Max Displacement (m)(with damping)

1 1255 1.58 0.063

2 1257 1.58 0.063

3 2147 1.24 0.049

4 2149 1.24 0.049

5 2812 1.22 0.048

6 10505 1.31 0.052

7 10508 1.31 0.052

8 11715 1.23 0.049

9 12233 1.55 0.062

10 12236 1.55 0.062

5/27/15 28Kasbar GulbenliCritical Damping 0.02

Science Payload

Frank Garcia

5/27/15 29Frank Garcia

Payload Activity at Europa’s Surface

• Temperature at surface: -220 °C to -160 °C

• 90 Day Radiation Dosage at surface: 400 krad

• Seismic Activity: Unknown

• 8 Seismometer stations placed across the surface of Europa

– Each station placed with a camera on a boom for 360° FOV

Image Courtesy of U. of Texas at Austin

Determine ice shell thicknessDetermine

presence of subsurface ocean

Take pictures of Europa’s surface


Seismometer and Camera Payload

• Seismometer – Insight SEIS Inheritance

• Vacuum Sphere

• Mass: 2.9 kg

• Power: 1.5 W

• Data Collected: 12.5 kbps

• SP and VBB Sensors (3 Each)

– 0.001 Hz to 50 Hz Frequency

• Thermal Heating – RHU (3 Units)

• Camera - Beagle 2 Stereo System Inheritance (2 used)

• Each Camera– Low Mass: 175 g

– Low Power: 0.9 W

– Size: 79 x 63 x 75 mm

– Temp Range: -150 to 100°C

– FOV: 48 deg.

– Focus: 1.2 m to infinity

– Data Size: 1.31 Mb per picture

(Compressed)

Image Courtesy of Micro-Cameras & Space Exploration


Payload Mission Design (1/2)

90° Tilt Capability

360° ActuatorRotation

Seismometer: Continuously gathering data

Extendable Boom

Orbiter has large FOV so that it can constantly

receive data from Payload Package

48° FOV

149 kbps Transfer Rate

• 2 Cameras In Use

• Payload Mass: 5.42 kg

• Typical Power Usage: 8.3 W

– Linear and Rotary Actuators

• Aluminum Shielding (2000 mil): 24.45 kg

– Estimated Mission Lifetime: 120 days

– Radiation Dosage: 15 krad

• Total Payload Mass (w/ shielding): 29.9 kg


Payload Mission Design (2/2)

Deployed Seismometer PackageSolar Arrays Deployed

Deployed Lander

Cameras take 1 image for every 4° of solar

horizontal azimuth

8 Landers Across the Surface of Europa for 90

day mission

Cameras at the Poles will also take aesthetic photos of Jupiter as seen from Europa


Payload Data at the Surface of Europa

• Camera – Takes various 2 π sr. coverage mosaics

– 90 pictures (118 Mb) per mosaic

– Between 1 – 7 Mosaics per week, per lander

– Aesthetic pictures of Jupiter

• Seismometer constantly collecting data – 12.5 kbps

• Approx. Data received by orbiter (90 day Surface mission)

• No. of Mosaics = 394

• No. of Pictures = 35,460

• Total Camera Data Size = 46.5 Gb

• Total Seismometer Data Size = 502 Gb (Compressed)

• Data Acquisition limited by available orbiter to lander transfer time


Payload Data Transferred (90 days)

Lander to Orbiter Transfer Data (90 days)

Lander 1

Lander 2

Lander 3

Lander 4

Lander 5

Lander 6

Lander 7

Lander 8

Total

Time Available to Transfer Data (hr)

221.40 196.64 70.52 43.26 42.68 71.27 197.70 222.29 1065.75

Data Transfer Rate (kbps) 149 149 149 149 149 149 149 149 N/A

No. of 2 pi Mosaics 90 70 15 15 21 19 74 90 394

Total No. of Pics taken 8100 6300 1350 1350 1890 1710 6660 8100 35460

Total Camera Data Transferred (Gb)

10.6 8.3 1.8 1.8 2.5 2.2 8.7 10.62 46.5

Seismic Data Collected (per day) (Gb)

1.08 1.08 1.08 1.08 1.08 1.08 1.08 1.08 8.64

Seismic Data Collected (Gb) 97.2 97.2 97.2 97.2 97.2 97.2 97.2 97.2 777.6

Seismic Data Transferred (Gb) 97.2 97.2 35.96 21.38 20.41 35.96 97.2 97.2 502.52

Total Data Transferred (Gb) 107.82 105.46 37.73 23.15 22.89 38.21 105.93 107.82 549.00


TelecommunicationsCommand and Data Handling

5/27/15 Omar Alhassen 36

Equipment List

Orbiter• X-Band Transponder • X-Band TWTA • X-Band Diplexor• X-Band Switching Network • X-Band Cables • X-Band Low Gain Antennas • X-Band High Gain Antenna• Ultrastable oscillator • Rad750 3u Compact PCI• Rad750• S990 3U Compact PCI Radiation

Tolerant Memory Module• SRR (1 Tb)

Lander

• Ultrastable oscillator

• S-Band Transponder

• S-Band Diplexor

• S-Band Cables

• S-Band Low Gain Antennas

• Rad750 3u Compact PCI

• Rad750

• S990 3U Compact PCI Radiation Tolerant Memory Module

• SRR (1 Gb per lander)


Flight Modes for Mission

• Safe (Emergency) Mode– All non-essential systems are shut down, only minimal

essential functions are active

• Cruise (Normal) Mode– Scientific instruments powered, while spacecraft on

its planned trajectory

• Sleep mode– Low power mode

• Orbiter Science mode– X Band amplifiers turned on continuously to support

gravity science


Safe Mode Requirements

• Detection of condition or event that may cause loss of control or damage

• Triggered event System Failure

• Power removed from non-essential subsystems

• Active functions during safe mode to be:– Thermal Management

– Communication (using LG)

– Attitude Control


Method for Full System Reset

• Forces spacecraft to engage in safety and self-protection

• Used for computer rebooting

• Spacecraft enters safe mode

• In safe mode– Science instruments are shut down– Spacecraft attempts to maintain orientation with Sun– Spacecraft awaits commands sent by the mission control center – These commands are received by the low gain antenna


Recovery (from safe mode)

• Re-establishing communication between spacecraft and mission control

• Diagnostic data is downloaded

• Sequencing power back on to the subsystems shut down prior are resumed


Scheduled Contact with Spacecraft

• Spacecraft will be contacted from ground through DSN

• Data sent to DSN from spacecraft will later be transmitted from DSN ground stations on Earth

• From ground station, commands uplinked via DSN to orbiter

• Will have contact with spacecraft every day through DSN– Request more time if emergency or required

• Signal transfer from Jupiter to earth is roughly 50 minutes


Low and High Gain Antennas

Lander to Orbiter Communication

• Use of Low Gain (LG) Antenna

• Data Transfer Rate using LG provides 149 kbps

Orbiter to Earth Communication

• Use of High Gain (HG) Antenna

• Low Gain at times

• Data transfer rate using HG provides 360 kbps


Communicating with Earth

• During mapping– Wide angle camera will take 46 pictures per orbit

• 2576 Mb of data

– Narrow angle camera will take 13 pictures per orbit• 2600 Mb of data

• Both high and low gain used to communicate

• An orbit of 7414.3 seconds, or 2.06 hours

• Worst case (use of LG) data rate of 149 kbps


Communication with Earth (cont.)

• With 149 kbps (from Low Gain):

– Data transfer of 1104.76 Mb per orbit

• With 360 kbps (from High Gain)

– Data transfer of 2669.22 Mb per orbit

• Data Storage if necessary

– 1 Tb storage on orbiter


Power System

5/27/15 46Jeremiah Kho

Spacecraft Power Overview

• Solar Power system is simpler

– System can be integrated easier than RTGs

– Cheaper system

• Solar Power is more available

– At 90.6 W at EOM, System would need 12 MMRTGs• 4 for orbiter

• 8 for landers

• Total of 38.4 kg of plutonium


Europa Orbit Sun DataOrbiter will be in a full sun orbit

– Spacecraft will only be eclipsed by Jupiter every 69.5 hours

– Eclipse time = 2.86 hours

– Solar flux at Jupiter is about 51 W/m2

STK Data:1 week periodTime

Eclip

se

5/27/1548Jeremiah Kho

Europa Surface Sun DataLander will be subjected to eclipses from Europa and Jupiter

– Total sunlight duration per period: 39.44 hours

– Total eclipse duration per period: 45.78 hours

– Solar flux = 51 W/m2

STK Data:1 week period

Time

Eclipse

Sunlight


Power SourcesOrbiter Lander

Four 10.5 m2 Rigid Panel Solar Arrays• Total Area = 42 m2

• Power Generated = 411.66 W• Mass = 168 kg

Two 3.3m Ø Ultraflex Solar Arrays• Total Area = 17.1 m2

• Power Generated = 43.89 W • Mass = 32.02 kg

100 Ah Li-ion Battery• Energy Capacity = 2880 Wh• Mass = 17.76 kg

121 Ah Battery• Energy Capacity = 3485 Wh• Mass = 21.12 kg


Average Power EstimatesOrbiter Power W % Lander Power W %

Thermal 82.00 25.68 Thermal 7.2 20.48

Attitude Control 56.74 17.77 Attitude Control 0 0.00

Power 41.40 12.96 Power 2.65 7.54

Command & Data 20.00 6.26 Command & Data 10.8 30.73

Telecom 40.00 12.53 Telecom 6.5 18.49

Propulsion 0.00 0.00 Propulsion 0 0.00

Mechanisms 10.50 3.29 Mechanisms 0 0.00

Payload 68.70 21.51 Payload 8.3 22.76

Total 319.34 100.00 Total 35.45 100.00

Power Generated 411.66 W Power Generated 43.89 W

Excess Power 92.33 W Excess Power 8.44 W

Power Margin 28.91 % Power Margin 23.81 %

Battery Excess 1170.61 Wh Battery Excess 291.02 Wh5/27/15 51Jeremiah Kho

Power System Block Diagram


Power System Mass SummaryOrbiter

Solar Array 168 Kg

Battery 17.76 kg

Power Conditioning Unit 16.6 kg

Power Distribution Unit 13.2 kg

Total 215.56 kg

Lander

Solar Array 32.02 kg

Battery 21.12 kg

Power Control System 1.5 kg

Total 54.64 kg5/27/15 53Jeremiah Kho

Thermal Systems

Randall Acosta

5/27/15 54Randall Acosta

Thermal Control Overview

Topics

• Key Thermal Drivers

• Thermal System Elements

- Active

- Passive

• Worst Case Heating/Cooling

- Orbiter

- Lander

• Mass Estimates

• Power EstimatesSpacecraft Image showing MLI

Source: CAD model fromTeam Member: Kasbar G.


• TPS must keep SC elements within their respective operating temperatures

• Battery is significant key driver due to losses

• Prefers a 273K to 283K thermal environment or approximately 0°C to 10°C

• Will function as low at 233K, but will rapidly lose efficiency below approx. 260k

Instrument Operating Temperatures

Component Min. Temp. [K] Max. Temp. [K]

Seismometer 123 423

Camera (w/boom) 123 373

Additional Operating Temperatures

Component Min. Temp. [K] Max. Temp. [K]

RHU 300 -

Lithium Battery 233 323

HG Antenna 216 334

Propellant 263 313

Star Tracker 243 323

Key Drivers - Thermal


Thermal System Elements

Active Thermal Control

• RHUs

• Resistance Heaters

Passive Thermal Control

• Optical Solar Reflectors

• Black & White Coatings

• Multi-Layer Insulation

Notes:

• Radiator and louvers not used for thermal environment during Venus flyby

• Spherical approximations and shape factors currently used


Sun

Spacecraft – during Venus flyby (lowest altitude)

Maximum Temperatures(without thermal control)Spacecraft: Tmax = 363 KSolar arrays: Tmax = 307 K

Note: Flyby altitude between 20,000km and 23,000km

Approx. 22000 km

Note: bodies and distances not to scale

Worst Case Heating - Spacecraft

Spacecraft – closest direct exposure to sunlight results in hottest thermal environment

Venus


Alleviation for Worst Case Heating• Turn down power consumption during Venus flyby (319W to 194W)• Use HGA as thermal shield (chart below produced assuming no alleviation)

• Reduces spacecraft temp. from 363K to 291K

5/27/15 Randall Acosta 59

Jupiter

Spacecraft – eclipse period during Europa orbit results in coldest thermal environment

Minimum Temperatures(without thermal control)Spacecraft: Tmin = 116 KSolar arrays: Tmin = 74 K

100 km

Europa

Worst Case Cooling - Spacecraft

Sun

Note: bodies and distances not to scale


Important Notes:• Total eclipse time is approx.

2.86hrs.• Min. S/c temp. occurs at approx.

1.9hrs into eclipse

Alleviation for Worst Case Cooling

• Internal heat generation of electronics also utilized

• Increases spacecraft temp. from 116K to 272K

5/27/15 Randall Acosta 61

PropertyMLI MLI Percent

(New) (Degraded) Change

Absorptance, α 0.13 0.18 27.78

Emissivity, ε 0.79 0.75 -5.33

Melting Temp, T 272.59 [°C] 250 [°C] -9.036

• MLI, black thermal coating, resistance heaters, and RHUs

• Note: MLI will experience degradation mostly due to particle radiation. After 90days = 1500krad, the following properties will change: (see Table below)

Not a concern. Thermal environment at Europa is well below the melting temp. shown above.

Mass for Elements of TPS [Units in Kg]

MLI Coatings (Wht & Blk) OSR Heating Units Total

Orbiter 88.8 5.5 ** 7.1* 101.4

Lander 4.21 1.2 ** 2.42* 7.82

Total (8 landers) 122.48 15.0 ** 26.46 164

Mass Estimates - Thermal


** OSR mass included in solar panel mass

* Resistance heaters and RHU masses included together

Power for Elements of TPS [Units in W]

ResistanceHeaters RHUs

Orbiter 33.6* 0.0 33.6

Lander 4.2* 0.0 4.2

Total (8 landers) 67.2 0.0 67.2

Power Estimates - Thermal

Note: The following also do not require power:

- Multi-Layer Insulation

- Thermal Coatings

- Optical Solar Reflectors

Power saved on lander heating units, for landers in high sunlight


* Highest power value provided here.

• Wattage for heater units is variable (0-4.2W) since thermal needs are determined in situ. Adjusted via solid state controller

• Heating requirements will vary on a lander to lander basis due to geological location and differing solar exposure

In case of single heater failure:

• Two redundant patch heaters per lander.

• Increase lander power output by 9.9W

Propulsion

Kronus Inc

Propulsion System- Key Drivers• Orbiter Architecture:• Variable thrust• Maximize Isp• Thermal equilibrium• Storability

• Lander Architecture:• Short nozzle length• Maximize Isp• Variable thrust preferred• but not necessary

Courtesy: Permanent.com

655/27/15 Juan Sanchez

Propulsion System-Orbiter• NTO and Hydrazine:

• Balance between thermal equilibrium and specific impulse

• Storable at spacecraft operating temperature

• (272 K)• Less heating power

required• Handles multiple start

ups wellKronus Inc


Propulsion System- OrbiterEngine Selection• TR-308 Liquid Apogee

Engine by Northrop Grumman

• Two engines are required

• Rated for 50 minute maximum firing duration

• Can handle mission’s longest burn (40 minutes,20% margin)

27.8 in

11.8 inCourtesy: Northrop

Grumman675/27/15

Juan Sanchez

Propulsion System-Orbiter

Monopropellant (HYD) ACS

Flowmeter

Latch Valve

Pyrotechnic Valve

Pressure Transducer

Solenoid Valve

System Filter

He 1 He 2

2 X 471 N Engine

NTO 1 NTO 2 HYD 1 HYD 2Centrifugal Tanks


Propulsion System-Orbiter

Propulsion System Mass

2026 Arrival (ΔV~2.95 km/s)

Falcon Heavy Delta IV Heavy

Propellant

Mass (kg)5530 5530

Prop System Wet Mass (kg)

6203 6203

LV Margin (kg) 3722 -1859

2027 Arrival (ΔV~2.41 km/s)

Delta IV Heavy

(8 Landers)

Delta IV Heavy

(7 Landers)

Propellant

Mass (kg)4080 3695

Prop System Wet Mass (kg)

4869 4410

Margin (kg) -406 320

*Note: All propellant masses account for an extra 10% ΔV, losses due to startup, and 3% residual.


Propulsion System-Liquid BiProp Lander

• NTO and MMH:

• Engine nozzle size best

suits structural design

• Specific impulse meets

landing scheme criteria

• Variable thrust

• Handles multiple start ups

well

Kronus Inc



Engine Selection:• Made by Aerojet

Rocketdyne• Four engines are

required• Short nozzle length• 330,000 pulses

available to vary thrust

• Maximum firing duration exceeds mission needs

• Lightweight ~ 2 kg Courtesy: Aerojet Rocketdyne



Flowmeter

Latch Valve

Pyrotechnic Valve

Pressure Transducer

Solenoid Valve

System Filter

He 1 He 2

4 X 110 N EngineR-1E by Rocketjet Aerodyne

NTO 1 NTO 2 MMH 1 MMH 2

Kronus Inc


Propulsion System-Lander Comparison

Propulsion System Mass Per Lander

Benefits:• Engine pulses can help vary

thrust• High specific impulse (280 s)• In case of engine failure, ACS can

couple with the thrusters balance the thrust

Liquid BiProp Engines (ΔV~1.48 km/s)

Propellant Mass (kg) 144

Total Wet Mass (kg) 169.5

Total Dry (kg) 25

Kronus Inc

*Note: Propellant mass accounts for an extra 10% ΔV, losses due to startup, and 3% residual


Attitude Control System

Sean Matthews

5/27/15 74Sean Matthews

Orbiter Attitude Overview

Dual Spin Stabilized Orbiter

Dual Spin Line


•Adapts characteristics from both 3-axis and spin stabilization

• Spin Mechanism Assembly (SMA)• Stabilized along minor axis

• 3 RPM• Major Burns

• Spin Up-Burn-Spin Down• 10 RPM

Despun Section

Spun Section

Orbiter Attitude Sensors

Courtesy of BATC

STAR TRACKER (x2)• Orbiter will be equipped with 2 CT-602 Star Trackers• Tracks up to 5 stars

Inertial Measurement Unit (x2)

• Small, Lightweight, highly reliable, fiber optic

• Three solid-state fiber optic gyros• 1 degree/hr gyroscope

Sun Sensor

• Adcole Spinning Sun Sensor

• Adds redundancy for Star Trackers

CT-602 Star Tracker

Adcole Spinning Sun Sensor

Courtesy of Adcole


Courtesy of NGC

LN-200 Core IMU

Orbiter Attitude Actuators• MR-106E 22N Thruster• 4 Rocket Engine Modules (REM)

• 3 thrusters each module• 2 tangential• 1 radial

Courtesy of Aerojet RocketdyneRocket Engine Module

Inertial Measurement Unit

Lander Attitude Sensors3-Axis Stabilized

LN-200 Core IMU

Courtesy of Northrop GrummanCourtesy of Honeywell

HG-8500 Radar Altimeter

Radar Altimeter

Courtesy of Aerojet Rocketdyne

Lander Actuators3-Axis Stabilized

• Two LN-200 Core IMUs per lander• ALHAT Sensors

• Two Peregrine 3D Flash LIDARS• Two HG-8500 Laser Altimeters per lander• 1 Doppler LIDAR

• MR-111C 4N Thruster• 2 Rocket Engine Modules (REM) per lander

• 5 thrusters each module• 2 thrusters on top of each lander


Thrusters

Rocket Engine Module

Lander Deployment

• Pyrotechnic Bolts (ΔV = .05 - .5 m/s)

• Two landers will be deployed simultaneously to avoid CG shift

• Using the star tracker and sun sensor, the deployment will occur as the chosen landers are directly on top and bottom of the dual spin orbiter (in reference to Europa)

• To the right is the direction vector of each of the 3 bodies after the deployment: 2 landers, 1 orbiter

• Only 3 RPM so the tangential velocity will be miniscule (<0.48 m/s)

1

5

12

3

45

6

7

8

• The CG shift only shifted along the spin axis (y-axis)


Landing and Hazard Detection

•Difficulties to overcome • Unpredictable terrain• Need to be fully autonomous with great precision

• Project Morpheus (NASA): ALHAT – Autonomous Landing Hazard Avoidance Technology

• 3D Flash LIDAR• Doppler LIDAR• High-Altitude Laser Altimeter

Description Size

Lighting Conditions Any

Global landing precision ± 90 m

Local landing precision ± 3 m

Detects hazard elevation > 30 cm

Detects hazard slopes > 15°

Courtesy of ASC

Courtesy of Honeywell

HG-8500 Altimeter Peregrine Flash LIDAR


ALHAT Sequence

> 10 km

Phases

TRN HDA HRN

Laser Altimeter

Navigation Doppler LIDAR

HDS

0.5 km

Sensors

DEM

Safe Landing Site

Safe Site Determination

TRN- Terrain Relative NavigationHDA- Hazard Detection and AvoidanceHRN- Hazard Relative NavigationDEM- Digital Elevation Model/MapHDS- Hazard Detection System

IMU

End HD (20 m)


ALHAT Redundancy

Instrument

Altimeter Flash LIDAR Doppler LIDAR IMU

Amount 2 2 1 2

Redundant? YES YES NO YES

Action in case of FAILURE

Use Doppler LIDAR to calculate distance

traveled from orbiter to calculate altitude

Hazard Avoidance will be surpassed. Lander

will land without hazard avoidance

IMUs will be sensor for trajectory and

orientation

Doppler LIDAR will be the primary sensor for

trajectory and orientation

• ALHAT has 3 sensors, but to ensure redundancy landers will be equipped with 2 3D flash LIDARS, and 2 Altimeters•A minimum of 3 instruments must fail for the lander to fail

•Various signals•Accelerometer- 3g•Damping Mechanism- kinetic energy -> electrical energy

•Voltage and current to trigger a touch-down signal•Position sensor- damping mechanism detects soft landing by evaluating small movements with the on-board computer

Touchdown Confirmation


Mass Table


Orbiter (kg) % Mass 1 Lander (kg) % Mass

Structure 157.5 4.4 26.8 13.6

Thermal 115.4 3.2 12.7 6.4

ACS 65.0 1.8 10.9 5.5

Power 237.1 6.6 60.1 30.4

Radiation 154.6 4.3 46.2 23.4

Cabling 56.1 1.6 6.1 3.1

Propulsion 675.0 18.9 25.5 12.9

Telecom 27.7 0.8 2.6 1.3

CDS 21.6 0.6 1.2 0.6

Payload 2738.2 76.6 5.4 2.7

Total Dry 3573.2 100.0 197.5 100.0

Total Wet 9777.1 342.3

Margin 3722.9

% Margin 38%Falcon Heavy Launch Mass ~ 13500 kg

Mission is possible on the Falcon Heavy

End of Mission (EOM)

• Orbit will naturally decay (periapsis decrease) and crash into Europa after 80-200 days after mission

• Europa-escape burn not feasible: ΔV ≈ 650 m/s

– Escape burn from Europa alone adds ~ 1500 kg of mass to entire system (50% of the launch margin)

– Most efficient method of disposal is controlled crash onto Europa

– Mapping will aid in determining suitable disposal site

– S/C will be adhere to Planetary Protection Protocols


EOM Disposal

• From NATIONAL ACADEMY PRESS Report Preventing the Forward Contamination of Europa (2000):

• Another conclusion reached from this particular sample analysis is that to meet the requirement that Pc ≤ 10-4 , it will be necessary to do at least one of the following:– 1. Demonstrate that no Type C or Type D organisms are on the

spacecraft; or– 2. Demonstrate that the probability of impacting the surface is less

than 10-4 for the entire time the spacecraft is in the vicinity of Europa(regardless of whether the spacecraft is operational or not); or

– 3. Show by probabilistic calculations that the 10-4 standard can be met through a combination of spacecraft cleaning, selective and/or whole-spacecraft sterilization, and exposure of the spacecraft to the radiation environment at Europa for a long enough period of time to reduce the bioload to the required level (“near sterilization”)


End of Mission (EOM)

Must follow NASA Planetary Protection Provisions• Category IV: Interested in the evolution and/or the origin of

life.• NPR 5.4.1: …reduce the probability of inadvertent

contamination of … liquid water body to less than 1 x 10-4

per mission• S/C payload and components shall be assembled and

maintained and payloads in Class 100,00• Methods will include (but not limited to) dry heat microbial

reduction (DHMR) and vapor phase hydrogen peroxide (VHP).

• Will use enhanced Viking methods at discretion of NASA PPO.


Preliminary Cost Estimates

• 2 Cost Methods for S/C estimate Used:– Unmanned Space Vehicle Cost Model, Eighth Edition

(USCM8)

– Project Cost Estimating Capability (PCEC) based of NAFCOM

• FY2020 Estimates for 1 Orbiter and 8 Landers:– USCM8: 2608 $M

– PCEC: 3727 $M

• PCEC seems more reasonable– Includes System Integration Costs


Preliminary Cost Estimates

• Estimated LV & Launch Ops Cost: ~ 200 $M• Mission Ops Cost ~ 128 $M• Software Cost ~ 184 $M (assuming 107 lines of

code)

• Total S/C Cost : ~ $4.2 billion

• Compared to other programs:– Juno Program Cost ~ $1.1 billion– Europa Clipper Projected Program Cost ~ $2.1 billion


Compliance MatrixReq # Requirement Compliance Justification

Tech 1.1

The array of seismometers shall be able to determine the interior structure and composition of Europa.

C 3-axis seismometers able to detect all 3 waves

Tech 1.2

There shall be sufficient geographic dispersion of the seismometers to determine the interior organization of Europa (7 minimum).

C Logarithmic Landing Sites in Longitude and Latitude.

Tech 1.3

A delivery system shall place the seismometers at their desired locations on the Europa surface.

C Soft Lander Design w/ Orbiter chosen.

Tech 1.4

The seismometer array shall last for a minimum of 90 days.

C Radiation Provides adequate protection for 90 days



Tech 1.5

There shall be a minimum of 7 seismometers on the Europa surface.

C Seismometer package chosen, one for each of the 8 landers

Tech 1.6

The system shall be able to return seismometer and imaging data (to orbiter relay → Earth).

C Acceptable data rates have been determined to relay data from orbiter to earth once per week, with a 1TB storage to hold data.

Tech 1.7

The landers shall be placed on Europa with a Legendre-Gauss-Lobatto point distribution. (changed Req.)

C The lander placement design for logarithmic scale.

Tech 1.8

The latitude pattern shall have minimum six calculated latitudes from the starting latitude.

C Latitude pattern created for lander insertion.

Tech1.9

The seismometers shall be able to detect P, S, and L waves.

C 3-axis seismometers able to detect all 3 waves.



Tech 1.10

The latitudes shall span 180°from pole to pole. (changed Req.)

C Landing trajectory set to start at north pole and end at south pole.

Tech 1.11

The longitudes shall span 360°. (changed Req.)

C Landing trajectory set to have spacing at logarithmic integrals spanning integralmultiples of 360°longitude.

Tech 1.12

Seismometers at sites 1-6+ shall be placed within 5° radius of the planned target point.

C Extra fuel margin to compensate for trajectory error.

Tech 1.13

Knowledge of the landing site must be determined by the end of mission to 0.1°.

C Accurate telecom instruments for the orbiter to determine location.



Tech 1.14

Each seismometer station shall carry a wide angle, color camera on a boom capable of 360°panning in azimuth and tilting in elevation

C Camera w/ boom has been modeled in Solidworks

Tech 1.15

Steradian coverage mosaickingshall permit >2π and Image sets shall be taken automatically

C Camera w/ boom has been modeled in Solidworks(Field of Views available for >2π steradian coverage)

Tech 1.16

There shall be at least one image set for every 4° of solar elevation angle above and below the local horizon

C Taken into account with boom design as the boom is capable of 180 degree horizontal/vertical turn

Tech 1.17

Cameras shall haveautoexposure, commandablepointing, and exposure controls and focus from 10 m to infinity

C Auto exposure available, as well as focus to infinity



Tech 1.18

Data shall be returned to Earth with a frequency set by the landers’ system production rate, the seismometer data production rate, and the camera production rate, and available recording system volumes, unless otherwise commanded.

C Lander Data Collected every 2 hours. Orbiter will transmit at 7.153 GHz.

Tech 1.19

Design a spacecraft to safely transport the payload from Earth to the surface of Europa

C Spacecraft model and trajectory has been designed.

Tech 1.20

Spacecraft shall arrive at Europa by 2026

C Spacecraft arrives February 1st 2026.


Conclusion

• Design Summary: 1 Orbiter and 8 Landers

– Solar Powered and Liquid Biprop

– Data is relayed from Landers to Orbiter to Earth

• Design is compliant with all RFP Requirements


Thank You For Listening!Questions?

Source: NASA

Documents

Senior Design - Europa Lander Mission Presentation