DAVID: Lunar and Planetary Institute
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DAVID:
Diminutive Asteroid Visitor with Ion Drive
A Cubesat Asteroid Mission
Geoffrey A. Landis
NASA Glenn Research Center
COMPASS Team at NASA Glenn:
Steve Oleson, Melissa McGuire, Aloysius Hepp, James Stegeman, Mike Bur, Laura Burke,
Michael Martini, Jim Fittje, Lisa Kohout, James Fincannon, and Tom Packard
Collaborating Institutions:
Busek (propulsion); COSMIAC (spacecraft integration); Case Western (Ralph Harvey, Asteroid
Science lead); Planetary Science Institute (Asteroid Science)
1
DAVID visualization
2
Background:
• Purpose:
– a cubesat design for SLS
launch opportunity for EM-1
– EM-1 launch (2018) to a
lunar free return trajectory.
– A small ∆V before lunar fly-by
can adjust the escape
trajectory to C3 =0 (co-orbital
to Earth)
• Mission: visit an asteroid
3
Highly Constrained Project:
• Size constrained:
– 6U
– about the size of a large shoebox
• Mass constrained
– 12 kg total mass
– including 1.3 kg assumed growth
• Propulsion constrained
– no energetic components
• Cost constrained
– $5.6 million dollar cost cap
– about $5 million after subtracting
required cost reserves
– (does not include launch)
4
Choice of Target
• 12874 close approaches by asteroids were analyzed
• List narrowed down:
– nearest approach no earlier than 2019 (>1 year after launch date)
– Nearest approach no later than 2020.
• 2001 GP2 was chosen as a target.
– From visual magnitude, estimate* ~18-meter diameter.
• Fly-by and Rendezvous missions analyzed
• From escape, ∆V of ~400 m/s needed for May 2020 fly-by.
• October 2020 fly-by was just outside the window of the solicitation
• ∆V ~2000 m/s needed to achieve a rendezvous.
*depending on asteroid albedo
5
Asteroid 2001 GP2 Earth Fly-by
• Closest pass: Oct 3, 2020
• Closest approach at 0.5 to 4.3 time Lunar distance
• (~100,000 to 1 million miles)
– Further observations will decrease uncertainty
• V relative: 2.37 km/sec
Near Earth asteroid Eros (viewed by the NEAR spacecraft)
6
Itokawa: the smallest asteroid ever
visited by a spacecraft
viewed by the Hayabusa spacecraft
What makes Asteroid 2001 GP2
Interesting?
Asteroid Itokawa
• 10-20 meters in diameter
– Two orders of magnitude smaller than any other
asteroid ever visited
– Typical of “city killer” impact threats (much more
frequent than extinction-level threats)
8
What makes Asteroid 2001 GP2
Interesting?
• Asteroids this tiny have never been visited by Asteroid Itokawa
spacecraft
– Recent modeling suggests YORP-induced spin-ups
can disaggregate bodies smaller than 150 m in size.
– 2001–GP2 sits significantly below this model’s
threshold for stability:
• Will it be an Itokawa-like body with a blocks and finer
grained material, or a single cohesive block?
• Will it be tumbling chaotically or spinning?
• What history is evident or implied?
Representative of a whole class of objects
that are numerous and interesting, but
have never been observed up close
9
DAVID
(Diminutive Asteroid
Visitor using Ion
Drive)
• 6-U cubesat
• Mass limit 12 kg
• Design must include
margin on all systems
10Size
23.94 cm
36.59 cm Planetary Systems Corporation
Canisterized Satellite Dispenser
11.28 cmSLS 6U Cubesat Deployed Dimensions
32.69 cm
Arrays Rotated
90 Degrees
81.46 cm 13.86 cm
33.8 cm
52.29 cmSystem Schematic
13Propulsion Trade Off
Mission is mass and volume constrained
Electric propulsion systems:
• High Specific Impulse: Low propellant use, high power
requirement
– If Isp is too high, the power system mass dominates the system
• Low Specific Impulse: High propellant use, low power
requirement
• If Isp is too low, the propellant mass dominates the system
14Propulsion Choice
• Trade-off study included many propulsion systems
• The Hall thruster has the greatest base of experience in
operation in space, but the cases analyzed exceeded the initial
mass allowance of 12 kg, as did the Xe ion thruster.
• A case with two 10-W PUC (Propulsion Unit for Cubesats)
electrospray thrusters also exceeded the mass allowance, but a
revised case where this was reduced to lower power and a
single thruster was run, which met the requirements.
• Single PUC electrospray thruster chosen for baseline
design
• Rendezvous case requires higher Isp and larger solar array
– ion engine needed
15Electrospray Thruster
Propellant:
high density ionic liquid
16Busek Electrospray Thruster
• Mechanical Simplicity: No Moving Parts
• Small Volume, Mass, and Power
• Leverages $20M NASA ST7 Technology Flight
Development
• Leverages SBIR Work on Micro-Valves and Power
Management
• Non-Volatile Propellant
• Multi-Emitter Design Busek Electrospray Thruster
• Lisa Pathfinder Flight Heritage
• Propellant Stored in Low Pressure Stainless Steel
Bellows Tank
• Cold Ion Plume (No Hot Gas)
• Self Regulating Feed System
• Piezo-Actuated Isolation Micro-Valve
Integrated PPU/DCIU
(Engineering Model)
17LISA Pathfinder Thruster Integration
182001 GP2 Interplanetary Trajectory
2018 Launch
• Colloid (Electrospray) Thruster
Parameters:
– Power to thruster = 9W
– Isp = 800s
– Efficiency = 31%
– Duty Cycle = 90%
• Trajectory Assumptions:
– Fly-by of 2001 GP2
– Constant 9W to thruster
– SLS Launch Date: 12/17/2018
• 4 days, 10 m/s to correct for worst-
case SLS injection
– Spacecraft Wet mass = 12 kg
• Trajectory Details:
– Delta-V = 390 m/s
– Required Prop Mass = 0.58 kg
– TOF = 507 days
– Total Thrusting Time = 89 days2018 Thrust Profile
2001 GP2 Interplanetary Trajectory
mid-2018 Launch (Double Fly-by)
• Colloid (Electrospray) Thruster
Parameters:
– Power to thruster = 9W
– Isp = 800s
– Efficiency = 31%
– Duty Cycle = 90%
• Trajectory Assumptions:
– Double Fly-by of 2001 GP2
– Constant 9W to thruster
– SLS Launch Date: 7/31/2018
• 4 days, 10 m/s to correct for
worst-case SLS injection
– Spacecraft Wet mass = 12 kg
• Trajectory Details:
– Delta-V = 365 m/s
– Required Prop Mass = 0.546 kg
– TOF = 1037.8 days
– Total Thrusting Time = 75 days2001 GP2 Interplanetary Trajectory
Fly-by 2001 GP22001 GP2 Interplanetary Trajectory
Fly-by 2001 GP22001 GP2 Interplanetary Trajectory
2018 Launch (October Fly-by)
• Colloid (Electrospray) Thruster
Parameters:
– Power to thruster = 9W
– Isp = 800s
– Efficiency = 31%
– Duty Cycle = 90%
• Trajectory Assumptions:
– Fly-by of 2001 GP2
– Constant 9W to thruster
– SLS Launch Date: 12/17/2018
• 4 days, 10 m/s to correct for worst-
case SLS injection
– Spacecraft Wet mass = 12 kg
• Trajectory Details:
– Delta-V = 43 m/s
– Required Prop Mass = 0.065 kg
– TOF = 660 days
– Total Thrusting Time = 57 days2018 Thrust Profile (October Fly-by)
Instruments
26SLS 6U Cubesat External Components
Sun Sensor
Solar Array Wing
Sun Sensor
Solar Array
Gimbal Unit
Science Camera
Solar Array Wing Sun Sensor
Star
Tracker
Sun Sensor X-Band Omni
Antennas
Space Weather Neutralizer
Science
X-Band Patch
Antenna
Electrospray
Thruster
Sun Sensors
Solar Array Wing
X-Band Patch
AntennaStatus
• Proposal submitted to SIMPLEX solicitation, but not
selected for the EM-1 Mission Opportunity
• Proposal was rated highly, and selected for a one-year
technology development study
• Asked by program office to focus work on maturing the instrumentation
• Continuing to working on the engineering design
• Looking for a launch opportunity to C3=0 in 2019 or
early 2020
28Where do we go if we find a launch, but miss
the window for the Oct 2020 fly-by?
• Latest possible launch for 2020 fly-by is ~May 2020
• Asteroid 2011-CL50 has December 24 2020 fly-by
• not quite as good, but almost
• only slightly later
• If we miss that, 2010 UE51 has opportunity Dec. 2023
• Tiny asteroid (~10 m class)
• Farther away, but much slower fly-by speed
• Possibility to do a rendezvous mission
29Conclusions
• Asteroid mission is possible with a 6-U cubesat
Targeting near earth asteroids that fly close to Earth
minimizes the propulsion required for fly-by/rendezvous
• Upcoming Oct 2020 fly-by has VERY low ∆V
*assuming you can reach escape
• Rendezvous mission is possible as a stretch goal
– ∆V is very significant for a 6U cubesat:
– ~2000 m/s needed for rendezvous with 2001 GP2
30DAVID visualization
31Backup/Slides not used
32Mission Delta-V Summary • Total propellant in table represents usable propellant • Additional 5% of usable carried as margin in inert mass of vehicle • Isp = 800 s
Reacquiring 2001 GP2
• Asteroid hasn't been seen since its discovery year
• Uncertainty in orbit is a million miles!
– we need to refine orbit or else we miss it entirely
• The good news: it makes an Earth close approach in May 2019,
at a distance of 0.154 AU
• can reacquire asteroid and refine orbit one year before the May 2020
fly-by; 17 months before the Oct 2020 fly-by
• plenty of time to fine-tune trajectory
• phase angle to sun ~90° (directly ahead of Earth in orbit)
• The bad news: estimated magnitude at this approach is about
~24
– need a big telescope to see object that faint.
• Keck or similar capability ground-based telescope
• Hubble space telescope
• February 2020 close approach at 0.124 AU is near opposition
– 3 month lead time for May 2020 fly-by, 8 month for Oct 2020 fly-by
34Propulsion System Layout
Micro Valve Bellows Tank
NEXT
Style
Gimbal
PPU/DCIU
Propellantless Electrospray Thruster\l
Cathode
35SLS Cubesat Bus
Propulsion (EP Hardware) MEL
• SLS Cubesat Bus Propulsion (EP Hardware) Subsystem MEL
Unit Basic Total
WBS Description QTY Mass Mass Growth Growth Mass
Number Case #3.5 SLSCubesat 2013 CD-2013-101 (kg) (kg) (%) (kg) (kg)
12/3/2013
06 SLS Cubesat 7.80 16.8% 1.31 9.11
06.1 Cubesat Bus 7.80 16.8% 1.31 9.11
06.1.9 Propulsion (EP Hardware) 1.40 10.7% 0.15 1.55
06.1.9.a Primary EP System Hardware 0.72 10.8% 0.08 0.80
06.1.9.a.a EP System Hardware 0.72 10.8% 0.08 0.80
06.1.9.a.a.a Electrospray Thruster 1.00 0.52 0.52 8.0% 0.04 0.56
06.1.9.a.a.b Gimbal 1.00 0.20 0.20 18.0% 0.04 0.24
06.1.9.b Power Processing Unit (PPU) 0.50 8.0% 0.04 0.54
06.1.9.b.a Integrated PPU/DCIU 1.00 0.50 0.50 8.0% 0.04 0.54
06.1.9.c Propellant Management (EP) 0.18 18.0% 0.03 0.21
06.1.9.c.a EP Propellant Tank system 0.18 18.0% 0.03 0.21
06.1.9.c.a.a Custom Tank 1.00 0.13 0.13 18.0% 0.02 0.15
06.1.9.c.a.b Feed System 1.00 0.05 0.05 18.0% 0.01 0.06
36SLS Cubesat Bus
Propellant (EP) MEL
• SLS Cubesat Bus Propellant (EP) Subsystem MEL
Unit Basic Total
WBS Description QTY Mass Mass Growth Growth Mass
Number Case #3.5 SLSCubesat 2013 CD-2013-101 (kg) (kg) (%) (kg) (kg)
12/3/2013
06 SLS Cubesat 7.80 16.8% 1.31 9.11
06.1 Cubesat Bus 7.80 16.8% 1.31 9.11
06.1.10 Propellant (EP) 0.69 0.0% 0.00 0.69
06.1.10.a Main Propellant (EP) 0.69 0.0% 0.00 0.69
06.1.10.a.a Primary EP Propellant Used 1.00 0.63 0.63 0.0% 0.00 0.63
06.1.10.a.b Primary EP Propellant Residulals (Unused) 1.00 0.03 0.03 0.0% 0.00 0.03
06.1.10.a.c Primary EP Propellant Performance Margin (Unused) 1.00 0.03 0.03 0.0% 0.00 0.03
37Communication
For analysis purposes, a transmission antenna gain of 8 dBi was used for
both systems, not taking advantage of the ability of X-band to produce a
tighter beam spread. Both S-band and X-band transmission can
accomplish the required data rate of 174 bps with either receiving antenna.
The encounter data can thus be downlinked in two DSN sessions with the
34-m receiver, or in less than one session, if the 70-m dish can be used.
38CMAD Spectrometer
392001 GP2 Interplanetary Trajectory
Rendezvous
2017 Launch
• 3cm RF Ion Thruster Parameters:
– Power to thruster = 50W
– Isp = 2200s
– Efficiency = 27%
– Duty Cycle = 90%
• Trajectory Assumptions:
– Rendezvous of 2001 GP2
– Constant 50W to thruster
– SLS Launch Date: 12/17/2017
• 4 days, 10 m/s to correct for worst-
case SLS injection
– Spacecraft Wet mass = 12 kg
• Trajectory Details:
– Delta-V = 2088 m/s
– Required Prop Mass = 1.11 kg
– TOF = 611 days
– Total Thrusting Time = 456 days2001 GP2 Interplanetary Trajectory
Arrive at 2001 GP2
Requires 11,000 hours (456 days) of thrust
-ion thruster needs qualification for this duration of operationFly-by versus rendezvous summary
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