Transiting Exoplanet Survey Satellite
August 22, 2018
Transiting Exoplanet Survey Satellite (TESS)Flight Dynamics CommissioningResults and Experiences
Joel J. K. Parker NASA Goddard Space Flight CenterRyan L. Lebois L3 Applied Defense SolutionsStephen Lutz L3 Applied Defense SolutionsCraig Nickel L3 Applied Defense SolutionsKevin Ferrant Omitron, Inc.Adam Michaels Omitron, Inc.
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https://ntrs.nasa.gov/search.jsp?R=20190000506 2019-04-29T20:09:35+00:00Z
Contents
Mission Overview
Flight Dynamics Ground System
TESS Commissioning Launch
Phasing Loops & Maneuver Execution
PAM & Extended Mission Design
Commissioning Results
Orbit Determination
Conclusions
2
Primary Goal: Discover Transiting Earths and Super-Earths Orbiting Bright, Nearby Stars Rocky planets & water worlds Habitable planets
Discover the “Best” ~1000 Small Exoplanets “Best” means “readily characterizable”
• Bright Host Stars• Measurable Mass & Atmospheric Properties
Unique lunar-resonant mission orbit provides long view periods without station-keeping
Mission Overview—Science Goals
Large-Area Survey of Bright Stars F, G, K dwarfs: +4 to +12 magnitude
M dwarfs known within ~60 parsecs
“All sky” observations in 2 years All stars observed >20 days Ecliptic poles observed ~1 year
(JWST Continuous Viewing Zone)
3
27 days
54 days
Mission Overview—Spacecraft
4
Northrop Grumman LEOStar-2/750 bus Propulsion:
1x 22N ΔV reaction engine assembly (REA) 4x 4.5N REA for attitude control
Communications: 2x S-band omnidirectional antennas 1x Ka-band high-gain antenna (HGA)
Attitude Control: 4x reaction wheel assemblies (RWAs), 2x star
tracker assemblies, 10x coarse sun sensors Keep-out zones associated with all sensors
Instrument: 4x camera assemblies 30° by ± 50° rectangular keep-out zone
s
THERMAL BLANKETS
REACTION WHEELS
SOLAR ARRAYS
MASTER COMPUTER
SUN SHADE
PROPULSION TANK
LENS HOOD
. LENSES
DETECTORS
ELECTRONICS
ANTENNA
STRUCTURE
Mission Overview—Trajectory Design
5
Overall mission design:3.5 phasing loops → lunar flyby → transfer orbit → mission orbit
Mission orbit features 2:1 lunar resonance (“P/2”), or 13.67 d mean orbit period
Earth-Moonrotating frame
Inertial frame
Mission Overview—Commissioning
6
Overall commissioning is ≤60d process
6 planned maneuvers + 5 optional/backup maneuvers
A2M },\.-----' I I I
ra = 270,000_..~i:!! __ ~ 1M Eng ! r,, Bum:
250 km all lnj
Phasing Loop 1 (6 d)
backup : : I
: I I
: : I
Phasing Loop 2 (9 d)
/,\-- ---
Phasing Loop 3 (9 d)
TCM3
r . = 478,000 km (75 Rel
r3 = 378,000 km (59 Re)
( Lunar flyby
PAM (Per Adjust)
Transfer Orbit (17 d)
Science Orbit 0 (14 cl)
Ascent and Commissioning (S60 days)
• IN Burn
• Optional Burn
• Perigee Passage
Science Orbit 1 Science Orbit2
Science Operations
Mission Overview—Navigation
7
Navigation support provided by NASA Deep Space Network (DSN) and Space Network (SN)
SN post-separation acquisition at +1 min Handover to DSN by +1.5 hr Near-continuous tracking through first phasing loop, then scheduled to cover maneuvers
and meet OD accuracy requirements Post-launch SN support scheduled around perigee maneuvers only
0
10
20
30
40
50
60
70
22 Sun
Apr 2018
1 Tue 8 Tue 15 Tue 22 Tue 1 Fri
Measurement Summary by Altitude Separation through PAM
Altitu
de
(R
e)
Time (UTCG)
Tess Tracking Tess No Tracking
lunar flyby
PAM
• =
Flight Dynamics Ground System
8
TESS Flight DynamicsGround System: NASA Flight Dynamics
Facility (FDF)
TESS Flight Dynamics System
FDF responsibilities: Facility/infrastructure
DSN/SN data interfaces
IOD external verification
FDS responsibilities: Maneuver planning
Orbit determination
Product generation
Maneuver reconstruction/calibration
Software utilized:Software UseL3 ADS Flight Dynamics System (FDS) Procedure Execution, Data ManagementNASA General Mission Analysis Tool (GMAT) Maneuver Planning, Ephemeris GenerationAGI Orbit Determination Tool Kit (ODTK) Primary Orbit DeterminationGoddard Trajectory Determination System (GTDS) Backup Orbit DeterminationSPICE Toolkit DSN Acquisition GenerationAGI Systems Tool Kit (STK) Analysis, QA, VisualizationsMATLAB Analysis, Plotting
NASA FDF
Common data
Tracking Data
.---------.,----------,,..--------,,--..... ---,,...--"!---, i DSN j MOC
Incoming Directory
soc CARA
CIL
Message Gateway
r----- -------------------------------- --------~ TESS FOS
! GTDS (100) ---------------'
Flight Dynamics Ground System
9
TESS Flight DynamicsGround System: NASA Flight Dynamics
Facility (FDF)
TESS Flight Dynamics System
FDF responsibilities: Facility/infrastructure
DSN/SN data interfaces
IOD external verification
FDS responsibilities: Maneuver planning
Orbit determination
Product generation
Maneuver reconstruction/calibration
Software utilized:Software UseL3 ADS Flight Dynamics System (FDS) Procedure Execution, Data ManagementNASA General Mission Analysis Tool (GMAT) Maneuver Planning, Ephemeris GenerationAGI Orbit Determination Tool Kit (ODTK) Primary Orbit DeterminationGoddard Trajectory Determination System (GTDS) Backup Orbit DeterminationSPICE Toolkit DSN Acquisition GenerationAGI Systems Tool Kit (STK) Analysis, QA, VisualizationsMATLAB Analysis, Plotting
NASA FDF
First end-to-end use of GMAT in primary maneuver planning role
Common data
Tracking Data
.---------.,----------,,..--------,,--..... ---,,...--"!---, i DSN j MOC
Incoming Directory
soc CARA
CIL
Message Gateway
r----- -------------------------------- --------~ TESS FOS
! GTDS (100) ---------------'
Launch Performance
Launch Date: April 18th 2018
Vehicle: SpaceX Falcon 9
Location: Cape CanaveralAir Force Station, SLC-40
10
Event Actual (UTC) Delta (s)
Liftoff 22:51:30.498 -0.502
Separation 23:41:03.177 +2.177
Element* Pre-Launch BET OD Delta 3σ Requirement Sigma
Apogee Altitude [km] 268,622.397 269,330.228 707.831 ± 20,000 0.11
Perigee Altitude [km] 248.456 248.755 0.299 ± 25 0.04
Inclination [deg] 29.563 29.579 0.016 ± 0.1 0.48
Argument of Perigee [deg]
228.111 228.088 -0.023 ± 0.3 -0.23
*All elements at epoch: 18 Apr 2018 23:45:30.666 UTC
Phasing Loops
11
Maneuver EpochDur. (sec)
Calibrated ΔV (m/s)
Performance Error (%)
Mean Pointing
Error (deg)
A1M 22 Apr 2018 01:59:06.628 50 3.915 -3.33 9.6
P1M 25 Apr 2018 05:36:42.053 449 32.265 -0.93 0.6
P2M 04 May 2018 08:05:46.650 7 0.430 -6.62 0.4
P3M 13 May 2018 11:37:48.648 29 1.862 +2.87 1.6
PAM 30 May 2018 01:20:23.149 923 53.409 -0.11 0.7
• All burns nominal
• A2M waived asunnecessary
• Performance error• <7% worst-
case• <1% for major
maneuvers
A1M
P1M P2M P3M
(flyby)
PAM
(A2M)
TESS Definitive Altitude Separation through PAM
/0
uo---=
50 -
2 10
~o
20 -
10 -
u 1
22Sun 1 l ue 8 l ue 1~ IU€ 2, l ue 1 1-n 1\p1 ?OIR (UTCG)
Maneuver Performance
12
Maneuver performance error trends with duration Shorter duration correlated with underperformance Trend explained as an artifact of thermal ramp-up of propulsion system
Analysis through P2M was used to predict likely performance of P3M flyby-targeting maneuver Maneuver thrust scaled by 95% during P3M planning to better target desired flyby
0
-1 'ii: a -2 L L
w -3 Q.J u C
E -4 L 0 't: -5
Q.J 0....
~ -6 +-' ro :e -7 ro u -8
-9
Maneuver Execution Error: Calibrated tiV vs. Reconstructed tiV
. • PlM, -0.93%
PAM, -0.11%
• P3M, -2.28%
• AlM, -3.33%
:.".":: P3M (predicted)
• P2M, -6.62%
0 100 200 300 400 500 600 700 800 900 1000
Burn Duration [sec]
PAM Design & Extended Mission Design
PAM: Period Adjust Maneuver Goal: Lower apogee to achieve 2:1 lunar resonance (approx. 13.67 day
orbit period)
Secondary objectives:• Improve long-term eclipse profile
• Maintain long-term orbit stability
Long-term extended-mission analysis performed to 18–25 years of mission life (to extent of prediction capability)
PAM was fine-tuned via parametric scanning process PAM start epoch fixes eclipse “trade space”
PAM duration chooses specific eclipse profile within trade space
Two major tools: Eclipse profile plots (a.k.a. “Napolean plots”)
Mean orbit period plots
13
Extended Mission Eclipse Profile
“Napoleon plot”: Each row shows eclipses over time for a trajectory associated with a given PAM duration (scale factor from nominal). Plot trade space is fixed with PAM start epoch (chosen by targeter) Selected profile (row) is fixed with choice of PAM duration (scale factor)
Selected PAM duration was chosen to avoid 3+ hr eclipses in 2021 and 4+ hreclipses in 2027 Selected duration = 97% of nominal (923 s, 53.6 m/s) Associated initial mission orbit period = 13.72 days
14
Gaps between long eclipses
Unscaled PAM
Selected PAM (97% dur.)
'-
I 1.2 -
1.15
1.1 -
0 t5 1.05 -«l u... <I)
tii ~ 0.95 -
~ 0.9 -a..
o.85 - I 0.8
0.75 -
I 0.7 ..__ _ ___., _ __._ ____________ __,__ __ ___.._ ___ _.__ __ ___., ___ __.__
2020 2022 2024 2026 2028 2030 2032 2034 2036 Eclipse Epoch
0-1 hr 1-2 hr 2- 3 hr 3-3.5 hr 3.5-4 hr 4+ hr
Extended Mission Stability
15
Each series = 1 PAM scaling value
25-year mean (above) is represented as one value (below)
Mean should be approx. 13.67 days to indicate stability
selected = 0.97
18
17
13
Nominal LRP Angle= 37.41 deg
--Resonance Orbit Period = 13.67 days
--72.75% 87.3%
--106.7% --121.25%
--Nominal (97%)
12 ~--~--~--~--~-~~-~--~--~--~--~-~~-~
l5 13.5 ., ~
13.4
13.3
13.2
2019 2021 2023 2025 2027 2029 2031 2033 2035 2037 2039 2041 2043
Time (UTC)
0.75 0.8 0.85 0.9 0.95 1.05 1. 1 1.15 1.2
Maneuver Scale Factor
Commissioning Results
PAM execution nominal;achieved long-term eclipseprofile shows no eclipses>3 hr duration
Achieved initialorbit period = 13.73 d
Long-term stability andperigee/apogee altitudepredictions meet expectations
All commissioning mission requirements were met:
16
Requirement Value Achieved
Orbit Period13.67 days (2:1 lunar resonance)
Achieved (orbit period oscillates about 13.67 days)
Maximum Perigee Radius ≤ 22 RE 18.70 RE
Maximum Apogee Radius < 90 RE 70.25 RE
Maximum Total ΔV ≤ 215 m/s 91.19 m/sMaximum Single Maneuver ΔV ≤ 95 m/s 53.41 m/sMaximum Commissioning Duration ≤ 2 months 54.87 days
Eclipses≤ 16 eclipses, ≤ 4 hours duration each (umbra + ½ penumbra)
10 eclipses in primary mission; longest = 2.5 hr
Orbit Determination Position Accuracy ≤ 6 km per axis Achieved throughout commissioningOrbit Determination Velocity Accuracy ≤ 7% of maneuver magnitude Achieved throughout commissioning
Eclipse Di.ration
2Ye rs 4 Ye rs 6Y ars 8Y rs 10 Y rs 12Y 14 Y ars 16Y ars 18Y
0 to 1 hr
• 1 to 2 hr
• 2 to 3 hr
Achieved Orbrt •• •• • • • • • 2020 2022 2024 2026 2028 2030 2032 2034 2036
Eclipse Epoch
Orbit Determination
17
Orbit determination was performed throughout commissioning
Software: AGI Orbit Determination Toolkit (ODTK)
DSN measurement types processed: TCP, Sequential Range
TDRS 5L Doppler measurements were available, but not used in final solution
Measurement types DSN TCP, DSN SeqRng, TDRS 5L DopplerDSN antennas DSS24, DSS26, DSS34, DSS36, DSS54, DSS65TDRS satellites TDRS-K, TDRS-L
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)SN.DSSG4 )Sf'\ -cp Meas f..e5iduals
)SN.0S536 )St'- -cp Mtct::i f..t::iiductl:s
)SN.0S565 )SI\ -cp Mtct::i Ft::iiduat:;
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Mec1surement Residuc1I / Sigmc1
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)Tue
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15 Tue 22 Tue
Time (UTCG)
)SN.DSS!:i4 )Sf\ Seq ,,g Weas Fesiduals DSf'-.OS324 D~N TCP f,,1ec;s P.esidi..al!
DSf'\.0S5.34 D~N TCP l\lt!G:S P.t!:silli..al:::
DSf'\.0S526 D~N TCP l\lt!c:s P.t:!:sillL.al:::
·"·Soma
I I I I I I I I I I
1 r·i c- rn
DSI\.D3324 C~N Seq Rn~ Mec:s P.esidi..al::
0S1\.05534 C~N St!tl Rn:.i Mt!c:s P.t::si t.11..al::: •
DSl\.05526 C~N St!t1 R11J Mt!c:s P.t:sitlL..al::: A ~
Orbit Determination
Minimal filter tuningwas required
Small injections ofprocess noise wereused sporadically toprevent collapse of covariance during perigee passes
Overall 3σ position uncertainty < 900 m < 450 m through phasing loops
Filter-smoother consistency well-behaved, remains w/in ±3σ bounds
18
Parameter Constant Bias Bias 1σ White Noise 1σ Bias Half-life [min]
DSN TCP -0.08 0.05 0.005 60DSN SeqRng [m] 0 5 0.25 60Spacecraft Cr 1.5 0.2 N/A 2880Spacecraft Cd 2.2 Not EstimatedSpacecraftTransponder Delay [ns]
5863.46 10 N/A 2880
Position Uncertainty (0.99P)
1000 Commi>iioning
900
800
- 700 E -V) 600 «l E C)
500
u5 400 Q.) Q.) ... .c
300
t- 200
100 l tol""tern,mJ:wmpDQY, 56
0 22 Sun 1 Tue 8Tue 15 Tue 22Tue 1 Fri 8 Fri
Apr 2018 Time (UTCG)
Tess 3-Sigma Radial Tess 3-Slgma lntrack Tess 3-Sigma Crosstrack
Conclusions
TESS will perform the first-everspaceborne all-sky survey of exoplanetstransiting bright stars.
TESS launched nominally on 18 Apr 2018,and successfully executed a 60-day flightdynamics commissioning phase.
All maneuvers executed nominally orwere waived as unnecessary.
All commissioning and primary missionrequirements were met or exceededand are expected to continue to be metfor 18+ years.
Mission ”firsts”: First mission designed for resonant orbit in the
primary mission Use of innovative techniques for fine-tuning
final maneuver for long-term characteristics First application of NASA’s open-source GMAT
in a primary role
TESS is now on-orbit and started science operations on July 25.
19
Lunar Flyby & Transfer Orbit
Lunar flyby: 17 May 2018
Performance asexpected:
Transfer orbit: Achieved apogee radius: 70.25 RE
Achieved perigee radius: 16.54 RE
21
Event Time (UTC) Lunar Altitude (km)
Expected 06:33:06 8183 km
Observed 06:34:36 8254 km
Delta 90 s 71 km
Element Pre-flyby Post-flyby
Perigee Radius (RE) 1.14 16.53
Apogee Radius (RE) 56.23 72.61
INC (deg) 29.3 36.6
RAAN (deg) 37 286
AOP (deg) 230 356
Orbit Determination
Filter-smoother consistency well-behaved, remains w/in ±3σ bounds
22
(Target-Reference) Position Consistency Statistics Commi>iioning
4--,-----------------------------------------------------------, 3
2
-2
-3
Apr 2018 22 Sun
Tess Radial
1 Tue 8 Tue
Tess In-Track
15 Tue 22 Tue 1 Fri 8 Fri
Time (UTCG)
Tess Cross-Track Upper Bound Lower Bound