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

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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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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:::

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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

Questions?

Courtesy of MIT TESS Science OfficeTESS first public image – Centaurus star field

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