Lynx, one of four strategic mission concepts under study for the 2020 Astrophysics Decadal Survey, will provide leaps in capability over previous and planned X-ray missions, and will provide synergistic observations in the 2030s to a multitude of space- and ground-based observatories across all wavelengths. Lynx will have orders of magnitude improvement in sensitivity, on-axis sub-arcsecond imaging with arcsecond angular resolution over a large field of view, and high-resolution spectroscopy for point-like and extended sources. The Lynx architecture enables a broad range of unique and compelling science, to be carried out mainly through a General Observer Program. This Program is envisioned to include detecting the very first supermassive black holes, revealing the high-energy drivers of galaxy and structure formation, characterizing the mechanisms that govern stellar activity - including effects on planet habitability, and exploring the highest redshift galaxy clusters. An overview and status of the Lynx concept are summarized.
This overview paper is a progress report about the system design and technology development of two interferometer concepts studied for the Terrestrial Planet Finder (TPF) project. The two concepts are a structurally-connected interferometer (SCI) intended to fulfill minimum TPF science goals and a formation-flying interferometer (FFI) intended to fulfill full science goals. Described are major trades, analyses, and technology experiments completed. Near term plans are also described. This paper covers progress since August 2003 and serves as an update to a paper presented at that month's SPIE conference, "Techniques and Instrumentation for Detection of Exoplanets."
This paper describes the technical program that will demonstrate the viability of two mid-infrared nulling interferometer architectures for the Terrestrial Planet Finder (TPF) to support a mission concept downselect in 2006 between a nulling interferometer and a visible coronagraph. The TPF science objectives are to survey a statistically significant number of nearby solar-type stars for radiation from terrestrial planets, to characterize these planets and to perform spectroscopy for detection of biomarkers. A 4-telescope, 36-m Structurally-Connected Interferometer using a dual-chopped Bracewell nuller will meet the minimum science requirement to completely survey at least 30 nearby stars and partially survey 120 others. A Formation-Flying Interferometer is being designed to meet the full science requirement to completely survey at least 150 stars, and involves a trade between dual-chopped Bracewell, degenerate Angel Cross, and the Darwin bow-tie configuration. The system engineering trades for the connected structure and formation-flying architectures are described. The top technical concerns for these architectures are mapped to technology developments that will retire these concerns prior to the project downselect.
There are many advantages to space-based interferometry, but monolithic, single-spacecraft platforms set limits on the collecting area and baseline length. These constraints can be overcome by distributing the optical elements of the interferometer over a system of multiple spacecraft flying in precise formation, opening up new realms of angular resolution and sensitivity. While the principles of interferometry are the same as for structurally-connected systems, formation-flying interferometers must integrate a wide range of technologies to provide an optically stable platform capable of finding, tracking and measuring fringes. This paper discusses some of the key differences between formation-flying and structurally-connected interferometers, including formation configurations, controlling beam shear, station-keeping, and the importance of delay and delay rate estimation in determining the instrument sensitivity.
Proposed future formation-flying interferometer missions include the Terrestrial Planet Finder (TPF), Darwin, the Submillimeter Probe of the Evolution of Cosmic Structure (SPECS), the Stellar Imager, the Micro-Arcsecond Xray Imaging Mission (MAXIM), and its precursor, MAXIM Pathfinder. In addition, Life Finder and Planet Imager have been identified as two formation-flying missions capable of detailed characterization of habitable exo-planets. The parameters for these missions are compared and described briefly.
The StarLight mission is designed to validate the technologies of formation flying and stellar interferometry in space. The mission consists of two spacecraft in an earth-trailing orbit that formation-fly over relative ranges of 40 to 600m to an accuracy of 10 cm. The relative range and bearing of the spacecraft is sensed by a novel RF sensor, the Autonomous Formation Flyer sensor, which provides 2cm and 1mrad range and bearing knowledge between the spacecraft. The spacecraft each host instrument payloads for a Michelson interferometer that exploit the moving spacecraft to generate variable observing baselines between 30 and 125m. The StarLight preliminary design has shown that a formation-flying interferometer involves significant coupling between the major system elements - spacecraft, formation-flying control, formation-flying sensor, and the interferometer instrument. Mission requirements drive innovative approaches for long-range heterodyne metrology, optical design, glint suppression, formation estimation and control, spacecraft design, and mission operation. Experimental results are described for new technology development areas.
A metrology subsystem on board the Deep Space 3, a separated spacecraft interferometer mission, is used to determine stellar fringe delay jitter, delay rate, and initial delay. The subsystem implements two capabilities: linear metrology for optical pathlength determination and angular metrology needed to determine the configuration and orientation of the spacecraft constellation. Frequency modulated metrology concept is used to implement high-precision (5nm) interferometric linear measurements over large target ranges (1km). System is made angle sensitive by using an articulated flat mirror at the target.
In an interferometer, an Optical Delay Line (ODL) must be able to inject a commanded pathlength change in incoming starlight as it proceeds from a collecting aperture to the beam combiner. Fringe visibility requirements for space interferometry prescribe that the optical path length difference between the two arms must be equal and stable to less than 5 nm RMS to a bandwidth of 1 kHz. For a space mission, an ODL must also operate in a vacuum for years, survive temperature extremes, and survive the launch environment. As part of the interferometer technology program (ITP) at JPL, a prototype ODL was designed and built to meet typical space mission requirements. It has survived environmental testing at flight qualification levels, and control design studies indicate the 5 nm RMS pathlength stability requirements can be met. The design philosophy for this ODL was to crete as many design concepts as possible which would allow a priori attainment of requirements, in order to minimize analysis, testing, and reliance on workmanship. Many of these concepts proved to be synergistic, and many attacked more than one requirement. This paper reviews the science and flight qualification requirements for the ITP ODL and details design concepts used to meet these requirements. Examples of hardware implementations are given, and general applicability to the field of optomechanics will be noted.
This article is concerned with the discussion of a control law design for a brassboard optical delay line (ODL) developed for the interferometry technology program at the JPL to support the space-based optical interferometry missions. Variations on the ODL brassboard design will be flown on the space interferometry mission and new millennium separated spacecraft interferometer. The brassboard ODL was designed to meet both the performance and environmental requirements for space interferometry. A control experiment was contrived to evaluate how well the brassboard optical delay line can control optical pathlength jitter. Fringe visibility resolution requirements for space interferometry prescribe that the optical pathlength from the two collecting telescope apertures must be equal and stable to within a few nanometers RMS. This paper describes the classical frequency domain lop shaping techniques that were used to design a control law for the experiment. Included is a description of a methodology for managing the control authority for the three actuation stages of the ODL, as well as, an input shaping technique for handling the large dynamic range issues. Experimental performance results characterizing closed loop control of residual optical jitter in an ambient laboratory environment are reported.
Deep Space 3 will fly a stellar optical interferometer on three separate spacecraft in heliocentric orbits: one spacecraft for the Michelson beam combining optics, and two spacecraft for each of the starlight apertures. The spacecraft will formation fly to relative spacecraft distances from 100 meters to 1 kilometer, enabling an instrument resolution of 1 to 0.1 milliarcsecond. At each baseline length and orientation - up to 100 points in the synthetic aperture plane for a given astrophysical target - the instrument will measure source visibility amplitude form which the source brightness distribution can be determined. An infrared metrology system performs both linear and angular metrology between spacecraft and is sued to estimate delay jitter, interferometer delay and delay rate. Pointing and control mechanisms use the metrology error signals to stabilize delay jitter and to null delay and delay rate to enable detection and tracking of a white light fringe on a photon-counting detector. Once stabilized, fringes can be dispersed on a CCD in up to 80 spectral channels to attain high-accuracy measurements of visibility amplitude as a function of wavelength.
A separated spacecraft optical interferometer mission concept proposed for NASA's New Millennium Program is described. The interferometer instrument is distributed over three small spacecraft: two spacecraft serve as collectors, directing starlight toward a third spacecraft which combines the light and performs the interferometric detection. As the primary objective is technology demonstration, the optics are modest size, with a 12-cm aperture. The interferometer baseline is variable from 100 m to 1 km, providing angular resolutions from 1 to 0.1 milliarcseconds. Laser metrology is used to measure relative motions of the three spacecraft. High-bandwidth corrections for stationkeeping errors are accomplished by feedforward to an optical delay line in the combiner spacecraft; low-bandwidth corrections are accomplished by spacecraft control with an electric propulsion or cold-gas system. Determination of rotation of the constellation as a whole uses a Kilometric Optical Gyro, which employs counter-propagating laser beams among the three spacecraft to measure rotation with high accuracy. The mission is deployed in a low-disturbance solar orbit to minimize the stationkeeping burden. As it is well beyond the coverage of the GPS constellation, deployment and coarse stationkeeping are monitored with a GPS-like system, with each spacecraft providing both transmit and receive ranging and attitude functions.
This advanced steering mirror design combines large angular travel with high bandwidth dynamic response and high accuracy. The benefits for space-based interferometry include more commonality between mechanisms, reduced spares inventory, lower procurement costs, and reduced risk. These devices are used for alignment and fine-steering functions in the coherent combination of light from several collectors to independent combiner optics. Since this design can be used for alignment and fine-steering functions, a reduced number of component designs are required for interferometric missions. In some cases functions can be combined into a reduced number of mechanisms. The steering mirror design achieves this with a simplified electromagnetic actuator configuration having no iron other than the magnets in the magnetic path. Other benefits of the simplified design include: a compact steering mirror envelope that is only slightly larger than the mirror itself, simplified fabrication and assembly, and reduced power consumption. This paper includes the application, requirements and configuration along with performance analyses and verification test data. Analytical models for force, power, thermal, magnetic, dynamic and mass properties as well as various figures of merit are described.
The Stellar Interferometer Technology Experiment (SITE) is a near-term precursor mission for spaceborne optical interferometry. Proposed by the MIT Space Engineering Research Center and NASA's Jet Propulsion Laboratory, SITE is a two-aperture stellar interferometer located in the payload bay of the Space Shuttle. It has a baseline of four meters, operates with a detection bandwidth of 300 nanometers in the visible spectrum, and consists of three optical benches kinematically mounted inside a precision truss structure. The objective of SITE is to demonstrate system-level functionality of a space-based stellar interferometer through the use of enabling and enhancing Controlled Structures Technologies such as vibration isolation and suppression. Moreover, SITE will validate, in the space environment, technologies such as optical delay lines, laser metrology systems, fringe detectors, active fringe trackers, and high- bandwidth pointing control systems which are critical for realizing future space-based astrometric and imaging interferometers.
Lead magnesium niobate (PMN) has many attractive features for precision submicron control. At room temperature hysteresis is less than 1%, thermal expansion is less than 1 ppm/ degree(s)C, and the sensitivity is 375 ppm strain at 600 V/mm. There has been recent interest in using PMN actuators in applications near 0 degree(s)C, which is near the Curie temperature of the PMN material. An investigation was conducted to obtain data on PMN:BA strain response and hysteresis at lower temperatures. Results of these experiments which were conducted to characterize the longitudinal and transverse field-induced strains at temperatures between -7 degree(s)C to 24 degree(s)C are provided for SELECT multilayer actuators and electroceramic plates. Measurements were made at 1 Hz both at constant maximum field (600 V/mm) as well as constant maximum strain (300 ppm for longitudinal; 140 ppm for transverse). Data shows that hysteresis and strain sensitivity to field increase monotonically as the temperature is decreased throughout the test range. Transverse strain is shown to track the longitudinal strain closely, within a simple scale factor. A comparison is made between constant field and constant strain hysteresis for both the longitudinal and transverse cases. Finally, data is presented which shows a factor of four reduction in hysteresis using passive charge control.
A class of proposed space-based astronomical missions requiring large baselines and precision alignment can benefit from the application of Controlled Structures Technology. One candidate mission, that of a 35 meter baseline orbiting optical interferometer, is studied as a focus mission for a testbed for controlled structures research. Interferometry science requirements are investigated and used to design a laboratory testbed which captures the essential architecture, physics and performance requirements of a full scale instrument. Testbed hardware used for identification and control is presented, including an on-board six-axis laser metrology system using state of the art cat's eye retroreflectors. The testbed and research program are discussed in terms of controlled structures design and in terms of the expected benefits to the optical engineering and science communities.