MAXIM Pathfinder IMDC Study 13 May 2002

Science Team Keith Gendreau Code 662 GSFC Webster CashUniversity of Colorado Ann ShipleyUniversity of Colorado

Objectives –Demonstrate X-ray interferometry in space as pathfinder to full up MAXIM –Image with 100 micro-arc second resolution using a 1-2 m baseline –1000 times improvement on Chandra Coronae of nearby stars Jets from black holes Accretion disks Two spacecraft flying in formation: –Telescope spacecraft with all the optics –300 micro arc sec pointing control –30 micro arc sec knowledge –“Detector spacecraft” positioned km  10 m and laterally aligned  2 mm from Telescope spacecraft to make fringes well matched to detector pixels –Detector and optics fit within medium class launch vehicle (e.g., Delta IV H) MAXIM Pathfinder Overview Optic Spacecraft Detector Spacecraft L= km!

But, to resolve the event horizon of a black hole, you need more… A Star like Capella is A few mas across.. But the event horizon for the black hole in M87 is only a few  as across….

Full Maxim Design 10 KM 5000 KM CONSTELLATION BORESIGHT Hub Spacecraft SPACECRAFT DELAY LINE 200 M baseline Optics divided between multiple spacecraft. 0.1  as Angular Resolution “Extreme” Formation Flying Detector flown 1000s of km from optics to make fringes comparable to detector pixel sizes

The “New” MAXIM Pathfinder 2 mission modes –mode 1: 100  as Science Very similar to original MP concept, but some looser tolerances 2 formation flying s/c Studies Stars, AGN, Black hole Jets and Accretion Disks –mode 2: 1  as Science Adds “N” s/c to extend angular resolution to a few  as Tougher Formation Flying tolerances Tougher Line-Of-Sight Requirements Get a Glimpse of a Black Hole Event Horizon! Optimize Full MAXIM mission Design to accomplish all mode 1 science with capability to explore mode 2 science N=2

Mode 1 Detector S/C ~320 kg (inst+bus est.) ~400 watts (inst+bus est.) Pitch, Yaw, Roll control to ~ arcminutes Pitch, Yaw, Roll Knowledge to arcseconds Line-of-Sight Control to 5 cm Optics Hub S/C ~600 kg (inst+bus est.) ~200 watts (inst+bus est.) Pitch, Yaw, control to ~ 1 arcsecond (TBR) Roll control to arcminutes Pitch, Yaw, Roll Knowledge to +/- 1 arcsecond 200 km +/- 5 m Formation Flying Challenges: Range control to 10 meters, knowledge to cm LOS Control to “inches” (detector size) LOS to target knowledge to ~30  as (~ km) ~10 degree slews every 1-2 weeks. (~100 targets in 2 years) Target Science: Stars Separate AGN Jets from Disks Neutron Stars SNR Other 100  as science targets 5cm control +/-15  m Knowledge

Mode 2 Detector S/C Optics Hub S/C Activate metrology system to freeflyer spacecrafts 20,000 km +/- 5 m New Formation Flying Challenges: LOS to target knowledge to ~0.1  as (~15 20,000 km) ~10 degree slews every 1-2 months. (~5 targets in 1 year) Target Science: Event Horizons Separate AGN Jets from Disks Neutron Stars Other 1  as science targets 5cm control +/-15  m Knowledge FreeFlyer S/C ~150 kg (inst+bus est.) ~100 watts (inst+bus est.) Pitch, Yaw control to ~1 arcsecond Pitch, Yaw Knowledge to arcseconds Roll Control to ~1 arcsec m Control to ~10 microns 2 arcsec

Mode 2 Detector S/C Optics Hub S/C Activate metrology system to freeflyer spacecrafts 20,000 km +/- 5 m The array of optics spacecrafts: These fly in a virtual plane that is normal to the Line-of-Sight to within 2 arcseconds How many can we fly without needing a second launch vehicle? 5cm control +/-15  m Knowledge 2 arcsec

Some Typical Numbers “MODE 1” 100  as Science (5x radians)  Angstroms S=10  m  F=200 km  2B= 2 m  FOV=250 arcseconds  Long. Control=7 cm “MODE 2” 1  as Science (5x radians)  Angstroms S=10  m  F=20000 km  2B= 200 m  FOV=2.5 arcseconds  Long. Control=700  m

Optics Hub Spacecraft 71 Modules in a ~4m Diameter (shrink as necessary) Each Module has 2 30cm long optics 0.7 cm 2 /Module ~5.4 kg/Module Total Module Mass: 383 kg Total Effective Area: 50 cm 2 13 cm 26 cm Thickness of assembly: 1+ m Additional mass terms: structure, Beacon, star tracker, s/c bus. +~180 kg??? Optics s/c comes to about 600 kg now…

What would one of these modules look like? m msin(g) m/6 Gap~msin(g) Pitch Control msin(g) 3/2m+d m/3 + msin(g) 2(w+gap)+msin(g) By 2(w+gap)+msin(g)+m/3+actuator+encoder ASSUME: w+gap~5 cm Encoder+encoder~5cm Sin(g)~1/30 -->(10cm+m/30)x(15cm+m/3+m/30) -->m=30cm-> 13cmx26cm

Optics Hub Spacecraft Total Mass: ~600 kg ACS Control: Pitch, Yaw ~1 arcsec Roll ~ 1 arcminute ACS Knowledge: Pitch, Raw, Roll ~ 1 arcsecond Payload Power Requirements (* to be refined): Mirror Heaters: ~2 watts/module-> 160 Watts Beacons & Metrology: ~100 Watts

Detector Spacecraft Prime Instrument: X-ray CCD ~100 kg ~50 Watts Wants to be in temperature range of -130 to -90 C Precision Line-Of-Sight Alignment Instrument: “Super Star Tracker” (ISAL) -120 kg 0.5m diam x 2 meter tall -250 watts -cryocoolers need ways to dump power at room temp and 100 K Science Telemetry: ~ 5 kbits/s TOTAL Science Payload Mass: 220 kg TOTAL Science Payload Power: 300 watts ACS Requirements: pitch,yaw,roll control to 1 arcminute (knowldege~ 1 arcsecond) Line-Of-Sight control to ~5 cm.

CCD Camera 50 kg 30cm 40cm ~1 m 50 kg 30cm To Radiator to unload heat from CCD. (900 cm 2 to -130) Electronics Box for CCD This side looks to optics hub s/c

“Line of Sight (LOS)” Instrument Beacon Tracker and Superconducting Gyroscope Insert. (0.5 m diam x 2 m long) Will need radiators (Mike Dipirro) CCD Camera And Electronics. Maybe share a radiator With the LOS instrument? S/C bus Detector S/C

FreeFlyer Spacecraft In Mode 2, we will use extra optical spacecraft to extend baselines to 100s of meters.. ~ 1m diam x 1 m long pill box ~45 kg payload How many can we add before needing a new launch vehicle? ~ 10 watts power (for heater and beacon system) ACS: –Pitch, Yaw, Roll control to ~arcseconds –Translation control to ~10  m –Integrated with BALL “swarm” sensor.

FreeFlyer Spacecraft This Side points toward Optics Hub S/C. 8 Mirror Modules (~5 cm 2 ) 1m How much can we exploit miniature component technologies for “nano” and “micro” sats?

What we want…. This is a ~>2015 mission. Are we right? Cost estimate Refinement of mission profile (eg orbit analysis, …) Identification of required “Miracles” and performance drivers Nice figures… We will be back- but hopefully starting at a better position.

Extensions My guestimate of the total masses suggest that all three s/c total to ~1100 kg. I suspect that there is more capacity in the Delta III (maybe >2500 kg?). If so, how many more free flyers can we put on? As with many of these distant missions, I wonder why we design a mission going 15+ years from now to use present, off-the-shelf technology… Eg. Solar concentrators….

Backup Slides

Designing a Mission 2: What angular resolution at what wavelength do you want?  res  What is the smallest X-ray pixel size(  m) you can imagine? s The baseline (2B) needs to be: 2B =  res The Focal Length (F) needs to be: F =  s/  res The FOV will be: FOV = 80  (s/ ) 2  res Tightest Formation Flying Tolerance between optics s/c = s. “Lateral” Longitudal Formation Flying Tolerance between optics s/c = 80  (s/ ) 2  res B The Difference Here is That we will have fringes 10x bigger than the CCD pixel Size.

Mirror Module Dimensions The Mirror modules are pairs of flat (better than /100) mirrors. One mirror is fixed, the other has pitch (~mas) and yaw (arcminute) control. The module also has the ability to adjust the spacing of the mirrors at the nm level to introduce ~ angstrom pathlength control. Thermal control consistent to maintain optical figure (~0.1 degrees). There is structure to hold the module together.

Actuator Requirements: The pitch control should be to the some fraction of the diffraction spot size.  (m*sin(g))~30 /m ~ 6 mas for m=100 cm, 62 mas for m=10cm ~ 30 nm of control for any size mirror The range of pitch control should be able to accommodate the range of baselines over the range of focal lengths.  max ~ B/F =  s) ~ 1 arcsecond of range ~ 5x10 -6 m of linear range for a mirror of length m. where s=CCD pixel size

dX dd In 1-D we have: 3 unknowns:  d,  o, dx But laser interferometers Only give us 2 measurements: Dred, and Dgreen Interference pattern from optics space craft laser interferometer  Dred Interference Pattern from Detector Space craft laser interferometer  Dgreen oo Two Laser interferometers can make the two spacecrafts virtually rigid- but we still need a tie-in to the celestial sphere-> we still need a star tracker.

dX dd Using a “Super Startracker” to align two spacecraft to a target. In the simplest concept, a Super Star Tracker Sees both Reference stars and a beacon on the other space craft. It should be able to track relative drift between the reference and the beacon to 30 microarcseconds- in the case of MAXIM Pathfinder. The basic procedure here, is to align three points (the detector, the optics, and the target) so they form a straight line with “kinks” less than the angular resolution. The detector and the optics behave as thin lenses- and we are basically insensitive to their rotations. We are sensitive to a displacement from the Line-of-Sight (eg dX). oo For a number of reasons (proper motion, aberration of light, faintness of stars,…) an inertial reference may be more appropriate than guiding on stars. The inertial reference has to be stable at a fraction of the angular resolution for hours to a day. This would require an extremely stable gyroscope (eg GP-B, superfluid gyroscopes, atomic interferometer gyroscopes).

Sizing the Laser Beacon for LOS 1 Watt = 6.25x10 18 eV/s 620 nm = 0.5 eV 620 nm ~ 12.5x10 18 photons/sec Aperture of SST Beacon Tracker= D = 125 mm Laser Beam Expander Diameter = L <~ 1 m Laser wavelength = ~ 620 nm or so Distance from Laser To Tracker = F = 2x10 7 m MAX Required accuracy of LOS=  ~10 -7 arcseconds Telescope is good to ~ 1 arcsecond Required number of Photons to get to accuracy: N=(1 arcsec/  ) 2 ~10 14 OPTICS HUB Size of Footprint of laser=  ( /L*F) 2 Fractional coverage of tracker area: {  D 2 /4}/  ( /L*F) 2 =(DL) 2 /(2F ) 2 Number of Photons/sec into tracker: =(DL) 2 /(2F ) 2 *W*12. 5x10 18 photons/s ~3x10 14 xL 2 xW photons/sec L in meters, W in output laser power. LISA Laser power efficiency=10% NOTE: The LASER DIVERGANCE MIGHT DRIVE THE PITCH AND YAW CONTROL ON THE HUB to A FEW 1/10s of an arc second!!!

28 Option 2: Gravity Probe-B-Like Telescope only centroids on beacon Gyro provides inertial reference frame Advantages –Gyros exist!! 1/3 micro arc sec/day –No need to find stars –Just a beacon tracker telescope –Cryo-cooler: TRL 5 by 2005 ($2-5M for cryo-cooler (flight model only), FM + EM $3 to $7 million, mass about 20kg Launches on Con-X ~ 2010 –Gyro was to launch this year (Oct 2002) $10-100M for both cooler & gyro Disadvantages –Must know aberration Delta-V to 3cm/sec (Landis checks) –GP-B = expensive –Cryogen or coolers/vibrations

29 Option 2: Gravity Probe-B-Like (cont.) Deltas on GP-B –Since 1/3 micro arc sec /day is not required, then we may be able to back off on this capability –Cryo-coolers mean normal conductivity launching –If negligible magnetic L2, simplifies magnetic shielding design –Requirements on magnets in s/c –Neutralize cosmic-ray charging –Respinning up gyro? (dynamic range) –Cryo getters less important? –Proof mass? Yes-maybe –Squids will be better –Venting vs. cooler mechanism –Mass, power,size,cost,other req’ts (mag,jitter,thermal) –Need to integrate over 10 sec you get 10E-13radians???? –ConX, NGST Cooler specs: 150W BOL, 250W EOL, 10 yr lifetime, 20-30kg include electronics, heat syncs temp, produces 7.5mWatts 6K Selection in March 2002 –Cryo cooler mating to Adiabatic Demagnetization Refrigerator (ADR) starts in Temperature 2 K or less. Estimate 30 watts.

Super Star Tracker Beacon System

Thermal Meeting a pointing requirement of 30 micro arc sec requires extreme structural stability between the attitude sensor and the instrument. Mounting the attitude sensor and the instrument to the same platform would help accomplish this; in any case, the interfaces must be extremely rigid to prevent drift. In addition, temperatures must be tightly controlled to maintain dimensional stability, even with structural materials having a very low coefficient of thermal expansion (CTE). Allowable temperature difference was calculated as follows: 1. Assume a perfectly rigid structure. 2. Assume the best structural material currently available (M55J composite, with a CTE = per o C) Since (sin 30 micro arc sec) is 1.5 x , this represents the distortion limit. The temperature change producing this error is (1.5 x ) / (10 -7 ) = 1.5mK Developing a lower CTE structural material should be a high research priority! Otherwise, temperature control must be to about 0.1mK, which has been done only in labs on very small sample volumes. Control to 1mK has been achieved on space instruments, but again in small volumes. This is another area requiring a technology upgrade.

Detectors Note: Mass is 4 kg for 2 detectors