Gary H. Blackwood

StarLight mission: a formation-flying stellar interferometer

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.

Optical delay line nanometer-level pathlength control law design for space-based interferometry

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.

The New Millennium Formation Flying Optical Interferometer

Spaceborne opt i ca l inte~erometry has been identified as a critical technology for many of NASA’s 21* centu~ science visions. Included in this m“sion are interferometers that can probe the orighs of stars and can ultimately study Earthlike planets around nearby stars. To accomplish this feat, separation of an interferometer’s collecting apertures by large baselines are required from hundreds of meters up to thousands of kilometers. Thus the large separations require multiple spacecraft jlying in a formation. Furthermore, optical pathlengths over these distances must be controlled to the nanometer level. This level of control demands precision spacecraft controls, active optics, metrology, and starlight detection technologies, To date, some of these technologies have been demonstrated only in ground applications with baselines of order of a hundred meters; space operation will require a significant capability enhancement. This paper describes the New Millennium formation flying optical interferometer concept and associated technologz”es. The mission is designed to provide a technology demonstration for multiple spacecraft precision formation flying and very long baseline optical interferometry. The interferometer would be distributed over three spacecraft: two spacecraft would serve as collectors, directing ——— ——— ————-— ———---— ——— ——— ——--— —— —--— Copyright@ 1997 by the American Institute of Aeronautics and Astronautics, Inc. The U.S. Governmertt has a royalty-free license to exercise all rights under the copyright claimed herein for governmental purposes. All other rights are reserved by the copyright owner. starlight toward a third spacecraft which would combine the light and perform the in terferometric detection. The interferometer baselines would be variable, allowing baselines of 100 m to 1 km in an equilateral formation, providing angular resolutions from 1 to 0.1 milliarcsec.

Enabling design concepts for a flight-qualifiable optical delay line

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.

Interferometer instrument design for New Millennium Deep Space 3

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.

Separated-spacecraft interferometer concept for the New Millennium Program

A separated spacecraft optical interferometer mission concept proposed for NASAs 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.

Deep Space 3 metrology system

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.

System design and technology development for the Terrestrial Planet Finder infrared interferometer

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.

Formation-flying interferometry

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.

Combined fast steering and alignment mirror for space-based interferometers

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.