The Jupiter Magnetospheric Explorer (JMEX) is a UV observatory operating in an earth orbit proposed as part of NASA's Small Explorer (SMEX) class of missions. To meet mission requirements the residual jitter portion of the imaging error budget is set at 0.079 arcsec (3σ) over a 33.3 ms frame integration time and 0.01 arcsec (3σ) for all frequency content higher than 15 Hz. These requirements are challenging for a small, low cost mission and require some innovative system solutions to achieve these goals. The solution, discussed in the paper, was to combine several jitter rejection techniques fine-balanced reaction wheel mounted on an isolation assembly, post processing using science images and reaction wheel momentum control. This paper focuses primarily on meeting the high frequency portion of the requirements. To facilitate system performance verification, we leveraged an integrated model toolset, EOSyM (End-to-end Optical System Model), developed and used on various other advanced space-based missions over the last 9 years. Starting with individual subsystem models for the reaction wheel disturbances, the coupled payload/ spacecraft structural dynamics model, and the optical design, we were able to evaluate the end-to-end LOS performance under varying reaction wheel speeds. At the end we found that the requirements could be met by maintaining the reaction wheels operating range within a well-defined speed band. This paper describes the mission, the technical challenges, the integrated model, and system performance results.

The Terrestrial Planet Finder Coronagraph is a visible-light coronagraph to detect planets that are orbiting within the Habitable Zone of stars. The coronagraph instrument must achieve a contrast ratio stability of 2e-11 in order to achieve planet detection. This places stringent requirements on several spacecraft subsystems, such as pointing stability and structural vibration of the instrument in the presence of mechanical disturbance: for example, telescope pointing must be accurate to within 4 milli-arcseconds, and the jitter of optics must be less than 5 nm. This paper communicates the architecture and predicted performance of a precision pointing and vibration isolation approach for TPF-C called Disturbance Free Payload (DFP)* . In this architecture, the spacecraft and payload fly in close-proximity, and interact with forces and torques through a set of non-contact interface sensors and actuators. In contrast to other active vibration isolation approaches, this architecture allows for isolation down to zero frequency, and the performance of the isolation system is not limited by sensor characteristics. This paper describes the DFP architecture, interface hardware and technical maturity of the technology. In addition, an integrated model of TPF-C Flight Baseline 1 (FB1) is described that allows for explicit computation of performance metrics from system disturbance sources. Using this model, it is shown that the DFP pointing and isolation architecture meets all pointing and jitter stability requirements with substantial margin. This performance relative to requirements is presented, and several fruitful avenues for utilizing performance margin for system design simplification are identified.

SNAP is a proposed space-based experiment designed to study dark energy and alternate explanations of the acceleration of the universe's expansion by performing a series of complementary systematic-controlled astrophysical measurements. The principal mission activities are the construction of an accurate Type Ia supernova Hubble diagram (the supernova program), and conducting a wide-area weak gravitational lensing (WL) survey. WL measurements benefit from a highly constant point spread function (PSF). The goal of this study is to quantify the anticipated variations in PSF arising from on-orbit thermal variations and and shrinkage associated with dryout of the composite telescope metering structure. A combined thermo-mechanical-optical analysis tool was developed, and WL metrics whisker and effective anisotropy quantified for thermal and composite structure dryout effects. Stability limits necessary for WL are defined, and compared to stability tolerances defined for the supernova program. The mission is designed for operations at at the Earth-Sun L2 Lagrange point, where thermal disturbances from Earth are minimal. In this study, the effects of seasonal variations in solar flux, transients introduced when pointing the body-fixed Ka-band antenna toward Earth and 90° roll maneuvers (planned every three months of operations) are quantified, and introduced into the optical system. Whisker and effective anisotropy were computed, and found to be well below the WL requirement for stability. The effects of composite structure shrinkage due to on-orbit H₂O desorption are discussed, and estimated to be below WL limits for daily observations, at the beginning of the WL phase of the mission.

The Terrestrial Planet Finder Interferometer (TPF-I) mission requires a set of formation-flying collector telescopes that direct the incoming light to a beam combiner where the beams are combined and detected to identify habitable planets. A baseline TPF collector design, using a primary mirror of 4.2 meters in diameter, is used here to conduct a dynamic study. The objective is to investigate the effects of dynamic response of the spacecraft on the system optical performance at the presence of disturbances that arise from the reaction wheel assembly and thruster loading, respectively. Frequency responses where the frequency is associated with the flywheel speed are presented in the paper. The results focus on the surface oscillation of the primary mirror and the point at which the secondary mirror is located. Transient response simulations under the baseline four thruster-assembly configuration were conducted using various duty cycles and thrust levels determined by the TPF formation rotation requirements. This paper will also describe an investigation conducted using new IMOS (Integrated Modeling of Optical Systems), which is an open, multi-disciplinary, and Matlab-based dynamic/optical system simulation code. A pre-processor that is able to generate the sub-structure modal models required by ISYSD (Integrated System Dynamics) was developed in new IMOS. ISYSD is used to develop a high-fidelity system dynamic model by integrating the sub-structure modal models. Finally, the paper will summarize current and future work in order to meet the TPF dynamic requirements.

Two recent meetings sponsored by NASA have helped define the in-space capabilities (technology, operations and infrastructure) necessary to enable and enhance future space missions. The activities preceded NASA's roadmapping efforts that occurred from the fall of 2004 to spring of 2005. These Loya Jirga meetings (using a Pashto expression for "grand council') involved about 100 representatives from industry, academia and government. Three mission concepts were used to guide the products of the meetings: manned missions to Mars, large serviceable space telescopes, and unmanned nuclear-powered missions to the outer planets. The deliberations produced roadmaps for the timing and type of developments needed to support these missions, the interconnections of capabilities with missions and other details that can be used to guide investment planning.

The NASA Advanced Telescopes and Observatories (ATO) Capability Roadmap addresses technologies necessary for NASA to enable future space telescopes and observatories collecting all electromagnetic bands, ranging from x-rays to millimeter waves, and including gravity-waves. It has derived capability priorities from current and developing Space Missions Directorate (SMD) strategic roadmaps and, where appropriate, has ensured their consistency with other NASA Strategic and Capability Roadmaps. Technology topics include optics; wavefront sensing and control and interferometry; distributed and advanced spacecraft systems; cryogenic and thermal control systems; large precision structure for observatories; and the infrastructure essential to future space telescopes and observatories.

The President has identified the search for Earth-like planets around nearby stars as a critical part of NASA's long term Vision for Space Exploration. A suite of space-based missions will determine the incidence of Earth-like planets, detect and characterize the nearest planets, search for signs of life in their atmosphere, as well as make great advances in our understanding of how planetary systems form and evolve. A detailed roadmap lays out the required technology developments, the precursor scientific knowledge, and the capabilities of the relevant missions.

The NASA strategic roadmap on the Origin, Evolution, Structure and Destiny of the Universe is one of 13 roadmaps that outline NASA's approach to implement the vision for space exploration. The roadmap outlines a program to address the questions: What powered the Big Bang? What happens close to a Black Hole? What is Dark Energy? How did the infant universe grow into the galaxies, stars and planets, and set the stage for life? The roadmap builds upon the currently operating and successful missions such as HST, Chandra and Spitzer. The program contains two elements, Beyond Einstein and Pathways to Life, performed in three phases (2005-2015, 2015-2025 and >2025) with priorities set by inputs received from reviews undertaken by the National Academy of Sciences and technology readiness. The program includes the following missions: 2005-2015 GLAST, JWST and LISA; 2015-2025 Constellation-X and a series of Einstein Probes; and >2025 a number of ambitious vision missions which will be prioritized by results from the previous two phases.

This paper examines requirements trades involving areal density for large space telescope mirrors. A segmented mirror architecture is used to define a quantitative example that leads to relevant insight about the trades. In this architecture, the mirror consists of segments of non-structural optical elements held in place by a structural truss that rests behind the segments. An analysis is presented of the driving design requirements for typical on-orbit loads and ground-test loads. It is shown that the driving on-orbit load would be the resonance of the lowest mode of the mirror by a reaction wheel static unbalance. The driving ground-test load would be dynamics due to ground-induced random vibration. Two general conclusions are derived from these results. First, the areal density that can be allocated to the segments depends on the depth allocated to the structure. More depth in the structure allows the allocation of more mass to the segments. This, however, leads to large structural depth that might be a significant development challenge. Second, the requirement for ground-test-ability results in an order of magnitude or more depth in the structure than is required by the on-orbit loads. This leads to the proposition that avoiding ground test as a driving requirement should be a fundamental technology on par with the provision of deployable depth. Both are important structural challenges for these future systems.

The Hubble Space Telescope (HST) continues to provide spectacular views into the universe. Its findings have profoundly affected mankind's view of itself by revealing to scientist and layperson alike many previously unimaginable discoveries. These result from the technical capability of HST. This 2.4-meter aperture diameter telescope includes imaging, spectroscopic, as well as limited coronagraphic instrumentation. Current plans (as of 2/8/2005) are to operate HST until later this decade, without servicing, and then deorbit it in a controlled manner early next decade. Cost effective, 2.4 meter, near term replacements for HST are under study as part of the NASA sponsored Origin Probe studies. TPF-C, scheduled for launch in mid-next decade, will develop the large mirror technology that could enable a next generation UV/Optical/NIR facility. The next generation facility would be a very large aperture collector telescope with wide field of view (FOV) imagery, precise wavefront control, and high ultraviolet efficiency. The facility would provide spectroscopic capability in addition to imagery. This paper will explore design trades and configurations applicable to a future expanded HST.

As with ground based observatories, future space observatories will require increasingly large apertures in order to address the latest questions concerning the nature of our universe and the origin of the galaxies, stars, planets and life itself. Nearly a dozen 8-m to 10-m telescopes are currently in operation on the ground, and designs are being developed for telescopes with apertures 30, 50, and even 100 meters in diameter. Space-based telescopes will inevitably follow this trend in order to take advantage of their freedom from atmospheric effects, diurnal thermal cycling, and limits on their field of regard. The apertures of space telescopes are limited by the size of launch vehicle's payload fairings, however; so segmented deployable optics are currently required for telescopes with apertures larger than approximately 4-meters. This paper discusses the current state-of-the-art and future prospects for large deployable telescopes in space.

We view broadly the science and technology drivers for both space and ground optical telescopes, in order to identify the unique capabilities and limitations in each domain. This leads us to consider the potential for effective "divisions of labor" and synergies to enhance scientific value. We project the influence of new enabling technologies, human priorities, international collaboration issues, and funding expectations. Finally, we discuss current NASA and ESA optical astronomy mission goals, and speculate on long-term forecasts.

Recent work done in preparation for a potential robotic mission to the Hubble Space Telescope has verified that the capability now exists to assemble, upgrade, and service large space-based telescopes robotically. This paper recommends that future space-based telescopes explicitly take this capability into account and plan for periodic robotic upgrades and servicing, just as the Hubble Space Telescope planned for periodic human upgrades and servicing. A single robotic servicing spacecraft stored at the Earth-Moon Lagrangian point, EML1, can readily provide capabilities for assembly, instrument and component replacements, and routine and emergency servicing for all future space telescopes located in Earth orbit or at the Earth-Sun Lagrangian points, ESL1 and ESL2.

As optical systems grow in size, there becomes a point in which traditional system verification prior to launch will become impossible. This implies that observatory ground testing will not be completed. Our history does not support this premise and therefore results in an unacceptable programmatic risk. But, if the dream of building 20-30 meter systems is ever to become true, these realities must be accepted. To make this possible, new and better analytical tools and processes must be developed and certified on programs that can be tested on the ground. This change in paradigm does not eliminate critical testing; it just does it at different assembly levels and most likely adds alignment flexibility to correct optical errors after launch. This paper provides ideas on how the hardware, analysis tools, and testing may evolve to support these ambitious future programs.

As envisioned space-based telescopes, observatories, and constellations of sensors grow in size and complexity, the ability to perform complete ground verification becomes increasingly difficult or impossible. Integrated system modeling offers one bridge in analyzing the expected optical performance and metrology of extended platforms in space to an accuracy exceeding the optical testing that can be performed in 1-g. In addition, some aspects of the final integration and system performance testing will eventually progress to on-orbit operations in the not-to-distant future as the infrastructure for lunar and Mars manned exploration proceeds. Specifically, the possibility of an Earth-moon L1 Gateway or a similar "shipyard" in space opens up the potential for some final optical characterizations being performed in space while additional human or robotically assisted alignments and integrations can be performed prior to final deployment to distant operational destinations such as at the Earth-Sun L2. Programs like Laser Interferometer Space Antenna (LISA) and Terrestrial Planet Finder (TPF) are examples of missions where sole reliance on ground optical testing will be extremely difficult, impossible, or inconclusive. Spitzer is a recent example where modeling was a key component of predicting temperature environment and corresponding performance. The future will require a greater reliance on modeling and, where warranted, optical testing and final alignment utilizing on-orbit test facilities. In fact, the case can be made that system modeling will need to be embraced more strongly if space-based assembly and test are to be realized. The necessary analytical tools, verification ground testbeds, and confirming flight experiments are crucial along with the planning that will take full advantage of the flexibility of final system verification at a Gateway prior to a low energy transfer to the observatory's final deployed operating orbit.

Recent advances in technology and materials science have enabled a number of large telescopes on Earth. Adaptation of these methods to space systems (beyond those being demonstrated in the James Webb Space Telescope (JWST)) might provide an equivalent breakthrough in science productivity, but only if important problems related to integration and performance testing are resolved. This paper proposes a program of technology and process development that can lead to efficient and reliable methods for in-space performance testing, thereby overcoming the limitations imposed by testing in gravity, the limited size of test chambers, the challenges of creating synthetic starlight sources and other factors. Considerations are given to the in-space facilities that are required and the need for human presence or tele-presence during the test interval. Both deployed and assembled space systems are considered. The paper addresses the optimal allocation of various test activities including the role of modeling, and functional and performance testing. Risk issues are defined, along with the impacts that such an integration and test path imposes on the telescope designer. The principle goal of the paper is to define those parts of the test process that can (or must) be deferred until the system is in an operations-like environment and to define the processes and technologies that must be brought to maturity to assure that testing does not limit our ability to continue to upgrade our observational systems.

Future space telescope missions concepts have introduced new technologies such as precision formation flight, optical metrology, and segmented mirrors. These new technologies require demonstration and validation prior to deployment in final missions such as the James Webb Space Telescope, Terrestrial Planet Finder, and Darwin. Ground based demonstrations do not provide the precision necessary to obtain a high level of confidence in the technology; precursor free flyer space missions suffer from the same problems as the final missions. Therefore, this paper proposes the use of the International Space Station as an intermediate research environment where these technologies can be developed, demonstrated, and validated. The ISS provides special resources, such as human presence, communications, power, and a benign atmosphere which directly reduce the major challenges of space technology maturation: risk, complexity, cost, remote operations, and visibility. Successful design of experiments for use aboard the space station, by enabling iterative research and supporting multiple scientists, can further reduce the effects of these challenges of space technology maturation. This paper presents results of five previous MIT Space Systems Laboratory experiments aboard the Space Shuttle, MIR, and the ISS to illustrate successful technology maturation aboard these facilities.

The Single Aperture Far Infrared (SAFIR) observatory is a high priority mission for NASA and space astronomy. This ten-meter diameter telescope, operating at <10 Kelvin, will chart the formation of galaxies and elements in the early universe, map debris disks around stars to track hidden planets, and explore the chemistry of life in the universe. While baselined as an autonomously deployed telescope, we consider enabling factors that in-space operations would bring to this telescope - in particular, servicing opportunities that would dramatically increase the scientific lifetime and productivity of the observatory. The use of humans and robots to support and conduct servicing, at the operational site of Earth-Sun L2 and primarily at Earth-Moon L1, are considered, and the required capabilities are reviewed. SAFIR shares many characteristics of future large telescopes in space, and strategies developed for this strawman case are applicable for broader planning efforts.

This paper elaborates on theory and experiment of the formation flight control for the future space-borne tethered interferometers. The nonlinear equations of multi-vehicle tethered spacecraft system are derived by Lagrange equations and decoupling method. The preliminary analysis predicts unstable dynamics depending on the direction of the tether motor. The controllability analysis indicates that both array resizing and spin-up are fully controllable only by the reaction wheels and the tether motor, thereby eliminating the need for thrusters. Linear and nonlinear decentralized control techniques have been implemented into the tethered SPHERES testbed, and tested at the NASA MSFC's flat floor facility using two and three SPHERES configurations. The nonlinear control using feedback linearization technique performed successfully in both two SPHERES in-line configuration and three triangular configuration while varying the tether length. The relative metrology system, using the ultra sound metrology system and the inertial sensors as well as the decentralized nonlinear estimator, is developed to provide necessary state information.

Gen-X is a next generation concept x-ray telescope that would be approximately 1000-times more sensitive than current x-ray telescopes such as Chandra. Since Gen-X will require focal lengths greater than 50 meters, formation flying the detector module behind the primary mirror is a feasible option. This study investigates the viability of a novel approach referred to as Electromagnetic Formation Flight (EMFF). EMFF uses High Temperature Superconducting (HTS) coils to generate force and torque between the primary and detector modules. EMFF subsystems such as coils, thermal control and power are sized as a function of different system parameters such as slew rate, focal length and detector mass. To investigate the viability of EMFF, a comparison is made between three different techniques for keeping the detector at the focal length distance behind the primary mirror assembly. The first architecture uses a deployable truss to create a Structurally Connected X-ray telescope (SCX), the second uses propellant-based formation flying (PFF) and the third uses Electromagnetic Formation Flying (EMFF).

This paper describes a concept of a formation flyer for ASPICS (Association de Satellites Pour l'Imagerie et la Coronagraphie Solaire), a giant 100 m based, externally occulted coronagraph aimed at observing the inner corona (and the solar disk) in the visible and ultra-violet. The two-satellite formation approach, based on existing space systems, is composed of a Myriade micro-satellite supporting the occulter and a Proteus platform as the main system carrying the coronagraph and imager scientific instruments. Both spacecrafts are launched as a single composite and deployed once on orbit, preferably a 3-day orbit or at the L1 Lagrange point. The coronagraph satellite acts as the "master" and provides the main functions of the mission (data handling, communication, propulsion, Guidance Navigation and control) while the Myriade acts as the "slave". The control of the formation is performed in two steps: i) RF metrology for deployment and preliminary pointing, ii) classical optical attitude sensors and metrology based on diverging laser beams. This will insure the nominal requirement of a lateral positioning with an accuracy of 1 mm and a longitudinal positioning with an accuracy of 500 mm.

New space missions, such as the Terrestrial Planet Finder (TPF) and Darwin programs, call for the use of spacecraft which maintain precise formation to achieve the effective aperture of a much larger spacecraft. Achieving this requires the development of several new space technologies. The SPHERES program was specifically designed to develop a wide range of algorithms in support of formation flight systems. Specifically, SPHERES allows the incremental development of metrology, control, autonomy, artificial intelligence, and communications algorithms. To achieve this, SPHERES exhibits a wide array of features to 1) facilitate the iterative research process, 2) support experiments, 3) support multiple scientists, and 4) enable reconfiguration and modularity. The effectiveness of these aspects of the facility have been demonstrated by several programs including development of system identification routines, coarse formation flight control algorithms, and demonstration of tethered systems.

The JWST project at the GSFC is responsible for the development, launch, operations and science data processing for the James Webb Space Telescope. The JWST project is currently in phase B with its launch scheduled for August 2011. The project is a partnership between NASA, ESA and CSA. The U.S. JWST team is now fully in place with the selection of Northrop Grumman Space Technology (NGST) as the prime contractor for the telescope and the Space Telescope Science Institute (STScI) as the mission operations and science data processing lead. This paper will provide an overview of the current JWST architecture and mission status including technology developments and risks.

We have developed generic mission architecture with James Webb Space Telescope heritage that can accommodate a wide variety of future space observatories. This paper describes the optimization of this architecture for the Single Aperture Far InfraRed (SAFIR) mission. This mission calls for a 10-meter telescope in an L2 orbit that is actively cooled to 4 Kelvin, enabling background-limited observations of celestial objects in the 30 to 800 micron region of the spectrum. A key feature of our architecture is a boom that attaches the payload to the spacecraft, providing thermal and dynamic isolation and minimizing disturbances from the spacecraft bus. Precision mechanisms, hinges and latches enable folding the observatory into a 5-m diameter fairing for launch and a precision deployment once on orbit. Precision mechanisms also articulate the telescope to minimize solar torques and increase the field of regard. The details of our design and the trades considered during its development are also described

The Wide-field Infrared Survey Explorer (WISE), a NASA MIDEX mission, will survey the entire sky in four bands from 3.3 to 23 microns with a sensitivity 1000 times greater than the IRAS survey. The WISE survey will extend the Two Micron All Sky Survey into the thermal infrared and will provide an important catalog for the James Webb Space Telescope. Using 1024² HgCdTe and Si:As arrays at 3.3, 4.7, 12 and 23 microns, WISE will find the most luminous galaxies in the universe, the closest stars to the Sun, and it will detect most of the main belt asteroids larger than 3 km. The single WISE instrument consists of a 40 cm diamond-turned aluminum afocal telescope, a two-stage solid hydrogen cryostat, a scan mirror mechanism, and reimaging optics giving 5" resolution (full-width-half-maximum). The use of dichroics and beamsplitters allows four color images of a 47'x47' field of view to be taken every 8.8 seconds, synchronized with the orbital motion to provide total sky coverage with overlap between revolutions. WISE will be placed into a Sun-synchronous polar orbit on a Delta 7320-10 launch vehicle. The WISE survey approach is simple and efficient. The three-axis-stabilized spacecraft rotates at a constant rate while the scan mirror freezes the telescope line of sight during each exposure. WISE is currently in its Preliminary Design Phase, with the mission Preliminary Design Review scheduled for July, 2005. WISE is scheduled to launch in mid 2009; the project web site can be found at www.wise.ssl.berkeley.edu.

The New World Observer has the potential to discover and study planets around other stars without expensive and risky technical heroics. We describe the starshade, a large, deployable sheet on a separate spacecraft that is flown into position along the line of sight to a nearby star. We show how a starshade can be designed and built in a practical and affordable manner to fully remove starlight and leave only planet light entering a telescope. The simulations demonstrate That NWI can detect planetary system features as faint as comets, perform spectroscopy to look for water and life signs, and perform photometry to search for oceans, continents, clouds and polar caps.

The creation of next generation, ultra-large space telescopes (>20m diameter) will require novel technologies. Many current concepts involve curved membrane reflectors but the problem is creating a diffraction-limited, three-dimensional surface. Here we present the idea of using a flat diffractive element which requires no out-of-plane deformation and is thus much simpler to deploy. The primary is a photon sieve--a diffractive element consisting of a large number of precisely positioned holes distributed over a flat surface. Photon sieves can be simply designed to any conic, apodization and operating bandwidth. The photon sieve is easier to fabricate than the better known Fresnel zone plate as a single substrate can be used since there are no connected regions requiring support. Presented here are results of prototypes capable of diffraction-limited imaging over wide fields and useful bandwidths.

Fresnel lens technology gives the ability to make very large, 10s to 100s meter apertures and enables a whole new regime of science and exploration missions allowing a new look at currently planned missions. Examples of missions that would be enabled or modified by these extremely large aperture sizes will be discussed at a conceptual level. This paper will conclude with a roadmap of technical challenges to be solved that will enable this exciting technology to develop and move into the technological main stream.

Laboratory testing is undertaken for exploring the feasibility of "Laser Trapped Mirrors". These involve a soft membrane or two-dimensional array of nano-spheres having sub-micrometer thickness. It is constrained to an accurate optical figure by radiation pressure trapping in polychromatic standing waves formed by a pair of diverging and counter-propagating laser beams. The optical design and resulting properties are discussed as well as the mission aspects for several architectures of telescopes and hypertelescopes at scales from meters to hundreds of kilometers.

To obtain sky background limited viewing in the mid and far infrared, cooling the telescope itself is required. In the past the lowest temperatures have been achieved using stored cryogen systems. Due to weight constraints and the desire for long life observatories, stored cryogen coolers will be replaced by passive coolers, mechanical coolers and solid state coolers. Some of the challenges and opportunities presented by these large cold telescopes and cryogenic cooling systems are described with emphasis on telescopes that require temperatures below 6 K.

The next generation of larger space optics will need lightweight and deployed mirror systems in order to control costs and fit within current and planned launch vehicle fairings. These will require active control based on wavefront sensing to establish and maintain their optical quality. Such control has been the enabling factor for the current generation of 8 m class ground-based telescopes, whose mirrors are either single monoliths with detailed shape control or have multiple rigid segments with control of relative position. They use actuator densities of typically a few per square meter. For active space systems it will be highly desirable to test the full deployed spacecraft in a vacuum test with a scene simulator, to validate before launch the optical performance of the complete system with its closed loop control systems. To enable such testing, the space mirror system must be designed from the start to work in a 1g as well as zero g environment. The orientation we envisage has the spacecraft system pointed at the zenith, illuminated by a downward beam collimated with reference to a full aperture liquid flat. We consider here two space mirror systems. The first has rigid segments supported by position actuators to control only rigid body motions. Since the segments under test must hold their shape with an axial 1g load and no passive flotation supports, they must be smaller than for ground systems. If made of lightweighted silicon carbide or beryllium for diffraction limited imaging in the optical, they would have to be ~ 30 cm in diameter. A mirror systems made from such segments will require about 40 actuators and wavefront sensor sub-apertures per square meter. The second system is a lightweight 3.5x8 m monolith for very high contrast imaging, as is envisaged for NASA's Terrestrial Planet Finder. High accuracy control of Fourier components down to ~ 0.2 m period is required, requiring a deformable mirror with about 4000 actuators. If the primary itself is the deformable element, and has a 1 cm thick glass meniscus facesheet weighing 600 kg, the gravity-induced quilting during testing would be about 1 nm rms, low enough for ground testing of the complete system at the desired 10-10 contrast level.

Deployable optics comprised of an electroformed, replicated nickel optical surface supported by a reinforced shape memory resin composite substrate have the potential to meet the requirements for rapid fabrication of lightweight, monolithic, deployable, large optics. Evaluation has been completed for various composite constructions including shape memory resin, carbon fiber reinforcement and syntactic fillers bonded to the electroformed nickel surface. Results from optical and structural performance tests on the 0.5 meter aperture deployable test items are also applicable to non-deployable replicated composite optics.

The use of thin-film membranes is of considerable interest for lightweight mirror applications. The low areal density makes them ideal for large aperture imaging applications. One type of setup looked into in the past has been the lenticular design, which consists of a clear canopy attached to a reflective film that uses positive pressure to set the curvature of the mirror. One drawback to this concept has been the fact that too much error was introduced during the pass through the canopy due to material inhomogeneities and poor optical properties. This is no longer an issue thanks to developments over the past several years in the field of optical-quality polymer development. Thin-films (< 24 microns) can now be routinely made with surface roughness, thickness variation, and very good transmission properties well within specification for many visible and IR applications. The next step in this developmental process has been maintaining a prescribed figure in the mirror. This paper summarizes the current efforts in fabricating and testing a 1-meter class lenticular membrane mirror system utilizing active boundary control and stress-coating applications to form a usable aperture for visible imaging applications.

Mechanism interface mechanics play an important role in the static and dynamic dimensional stability of deployable optical instruments. Friction mechanics in deployment mechanisms has been found to be a source of kinematic indeterminacy allowing elastic energy to be stored throughout the structure. At submicron scales, microslip mechanics allow this behavior to persist well below the classical Coulomb friction limit. This paper presents the design of a cryogenic tribometer for measuring this behavior in candidate mechanism interfaces in both room temperature and cryogenic environments. Room temperature results are presented and compared to a proposed generalized microslip model form. This model form is intended to allow the parametric characterization of microslip behavior caused by smooth nonconforming contact as well as roughness-induced microslip. Spherical ball-on-flat interface geometries were used with two unlubricated material combinations: 440C stainless steel ball on a 440C stainless steel flat and a silicon nitride ball on a 440C stainless steel flat. Consistent parameters were identified for the generalized microslip model from steady cyclic shear responses for both of these interface cases. While these parameters exhibited a measurable sensitivity to normal preload levels, the model form appears to provide the necessary level of robustness. Non-ideal transient shear phenomena including rate dependence were also observed but should play only a secondary role in future modeling efforts.

The Jupiter Magnetosphere Explorer (JMEX) is a proposed earth-orbiting satellite which will image the planet Jupiter in the FUV with a 0.5 m telescope at 0.25 arcsec (") resolution. Because the satellite is small and lightweight, vibrations from the reaction wheels (even though isolated by dampers) produce random pointing errors with an amplitude as large as 5" at a frequency around 1 Hz. In order for the telescope to achieve a resolution of 0.25" FWHM during long exposures, we will use a novel post-processing scheme to correct the pointing error. The UV science camera is a photon-counting MCP detector which produces data as a time-stamped photon list with 0.08" spatial resolution and roughly 1 ms temporal resolution. Simultaneously, a 0.5" pixel video camera, fed by a pickoff mirror in the main beam, captures visible images of the planet's disk at 30Hz and, with onboard processing, the centroid of the planet is determined, frame-by-frame, with a resolution <0.02" (1/25 pixel). With inter-frame interpolation, each photon from the UV camera is position-corrected in ground post-processing to an accuracy of 0.02".To rigorously test this scheme, we have constructed a hardware mock-up consisting of a tip-tilt mirror, a beam-splitter, and two video cameras with controlled noise characteristics. The tip-tilt mirror produces controlled image motion over a range of amplitudes and frequencies. With all parameters at worst-case values, we have verified the specified performance of the system and achieved centroid correction close to the limit set by counting noise statistics.