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A Right qualified micro sun sensor is being developed and flight qualified for future Man missions. The micro sun sensor, which Is basically a small pinhole camera, consists of a small mask with pinholes, placed on top of an image detector. Images of the sun are formed on the image detector when the sun illuminates the mask. Image processing is performed in the sun sensor that outputs sun centroids.
A micro sun sensor is being developed for use on a Mars rover for the Mars Science Laboratory Mission. The micro sun sensor, which is basically a small pinhole camera, consists of a small mask with pinholes, placed on top of an image detector.
A prototype small, lightweight micro Sun sensor (MSS) has been flight qualified as part of the attitude-determination system of a spacecraft or for Mars surface operations. The MSS has previously been reported at a very early stage of development in NASA Tech Briefs, Vol. 28, No. 1 (January 2004). An MSS is essentially a miniature multiple-pinhole electronic camera combined with digital processing electronics that functions analogously to a sundial. A micromachined mask containing a number of microscopic pinholes is mounted in front of an active-pixel sensor (APS). Electronic circuits for controlling the operation of the APS, readout from the pixel photodetectors, and analog-to-digital conversion are all integrated onto the same chip along with the APS. The digital processing includes computation of the centroids of the pinhole Sun images on the APS. The spacecraft computer has the task of converting the Sun centroids into Sun angles utilizing a calibration polynomial. The micromachined mask comprises a 500-micron-thick silicon wafer, onto which is deposited a 57-nm-thick chromium adhesion- promotion layer followed by a 200-nm-thick gold light-absorption layer. The pinholes, 50 microns in diameter, are formed in the gold layer by photolithography. The chromium layer is thin enough to be penetrable by an amount of Sunlight adequate to form measurable pinhole images. A spacer frame between the mask and the APS maintains a gap of .1 mm between the pinhole plane and the photodetector plane of the APS. To minimize data volume, mass, and power consumption, the digital processing of the APS readouts takes place in a single field-programmable gate array (FPGA). The particular FPGA is a radiation- tolerant unit that contains .32,000 gates. No external memory is used so the FPGA calculates the centroids in real time as pixels are read off the APS with minimal internal memory. To enable the MSS to fit into a small package, the APS, the FPGA, and other components are mounted on a single two-sided board following chip-on-board design practices
We have developed a new technique to fabricate antireflection surface using silicon nano-tips for use on a micro sun sensor for Mars rovers. We have achieved randomly distributed nano-tips of radius spanning from 20 nm to 100 nm and aspect ratio of ~200 using a two-step dry etching process. The 30(deg) specular reflectance at the target wavelength of 1 (mu)m is only about 0.09 %, nearly three orders of magnitude lower than that of bare silicon, and the hemispherical reflectance is ~ 8%. By changing the density and aspect ratio of these nanotips, the change in reflectance is demonstrated. Using surfaces covered with these nano-tips, the critical problem of ghost images that are caused by multiple internal reflections in a micro sun sensor was solved.
A report describes the development of a compact micro Sun sensor for use as a part of the attitude determination subsystem aboard future miniature spacecraft and planetary robotic vehicles. The prototype unit has a mass of only 9 g, a volume of only 4.2 cm(sup 3), a power consumption of only 30 mW, and a 120 degree field of view. The unit has demonstrated an accuracy of 1 arcminute. The unit consists of a multiple pinhole camera: A micromachined mask containing a rectangular array of microscopic pinholes, machined utilizing the microectromechanical systems (MEMS), is mounted in front of an active-pixel sensor (APS) image detector. The APS consists of a 512 x 512-pixel array, on-chip 10-bit analog to digital converter (ADC), on-chip bias generation, and on-chip timing control for self-sequencing and easy programmability. The digitized output of the APS is processed to compute the centroids of the pinhole Sun images on the APS. The Sun angle, relative to a coordinate system fixed to the sensor unit, is then computed from the positions of the centroids.
A new generation of sun sensors is emerging. These sun sensors utilize an imaging detector and the sun sensor determines the sun angles based on an image of fringes or centroids on the detector plane. Typically determines the sun angle in two axes.
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Topics covered include: Miniature Intelligent Sensor Module; "Smart" Sensor Module; Portable Apparatus for Electrochemical Sensing of Ethylene; Increasing Linear Dynamic Range of a CMOS Image Sensor; Flight Qualified Micro Sun Sensor; Norbornene-Based Polymer Electrolytes for Lithium Cells; Making Single-Source Precursors of Ternary Semiconductors; Water-Free Proton-Conducting Membranes for Fuel Cells; Mo/Ti Diffusion Bonding for Making Thermoelectric Devices; Photodetectors on Coronagraph Mask for Pointing Control; High-Energy-Density, Low-Temperature Li/CFx Primary Cells; G4-FETs as Universal and Programmable Logic Gates; Fabrication of Buried Nanochannels From Nanowire Patterns; Diamond Smoothing Tools; Infrared Imaging System for Studying Brain Function; Rarefying Spectra of Whispering-Gallery-Mode Resonators; Large-Area Permanent-Magnet ECR Plasma Source; Slot-Antenna/Permanent-Magnet Device for Generating Plasma; Fiber-Optic Strain Gauge With High Resolution And Update Rate; Broadband Achromatic Telecentric Lens; Temperature-Corrected Model of Turbulence in Hot Jet Flows; Enhanced Elliptic Grid Generation; Automated Knowledge Discovery From Simulators; Electro-Optical Modulator Bias Control Using Bipolar Pulses; Generative Representations for Automated Design of Robots; Mars-Approach Navigation Using In Situ Orbiters; Efficient Optimization of Low-Thrust Spacecraft Trajectories; Cylindrical Asymmetrical Capacitors for Use in Outer Space; Protecting Against Faults in JPL Spacecraft; Algorithm Optimally Allocates Actuation of a Spacecraft; and Radar Interferometer for Topographic Mapping of Glaciers and Ice Sheets.
Described is a device having an anti-reflection surface. The device comprises a silicon substrate with a plurality of silicon spikes formed on the substrate. A first metallic layer is formed on the silicon spikes to form the anti-reflection surface. The device further includes an aperture that extends through the substrate. A second metallic layer is formed on the substrate. The second metallic layer includes a hole that is aligned with the aperture. A spacer is attached with the silicon substrate to provide a gap between an attached sensor apparatus. Therefore, operating as a Micro-sun sensor, light entering the hole passes through the aperture to be sensed by the sensor apparatus. Additionally, light reflected by the sensor apparatus toward the first side of the silicon substrate is absorbed by the first metallic layer and silicon spikes and is thereby prevented from being reflected back toward the sensor apparatus.
Described is a device having an anti-reflection surface. The device comprises a silicon substrate with a plurality of silicon spikes formed on the substrate. A first metallic layer is formed on the silicon spikes to form the anti-reflection surface. The device further includes an aperture that extends through the substrate. A second metallic layer is formed on the substrate. The second metallic layer includes a hole that is aligned with the aperture. A spacer is attached with the silicon substrate to provide a gap between an attached sensor apparatus. Therefore, operating as a Micro-sun sensor, light entering the hole passes through the aperture to be sensed by the sensor apparatus. Additionally, light reflected by the sensor apparatus toward the first side of the silicon substrate is absorbed by the first metallic layer and silicon spikes and is thereby prevented from being reflected back toward the sensor apparatus.
Topics covered include: Multisensor Instrument for Real-Time Biological Monitoring; Sensor for Monitoring Nanodevice-Fabrication Plasmas; Backed Bending Actuator; Compact Optoelectronic Compass; Micro Sun Sensor for Spacecraft; Passive IFF: Autonomous Nonintrusive Rapid Identification of Friendly Assets; Finned-Ladder Slow-Wave Circuit for a TWT; Directional Radio-Frequency Identification Tag Reader; Integrated Solar-Energy-Harvesting and -Storage Device; Event-Driven Random-Access-Windowing CCD Imaging System; Stroboscope Controller for Imaging Helicopter Rotors; Software for Checking State-charts; Program Predicts Broadband Noise from a Turbofan Engine; Protocol for a Delay-Tolerant Data-Communication Network; Software Implements a Space-Mission File-Transfer Protocol; Making Carbon-Nanotube Arrays Using Block Copolymers: Part 2; Modular Rake of Pitot Probes; Preloading To Accelerate Slow-Crack-Growth Testing; Miniature Blimps for Surveillance and Collection of Samples; Hybrid Automotive Engine Using Ethanol-Burning Miller Cycle; Fabricating Blazed Diffraction Gratings by X-Ray Lithography; Freeze-Tolerant Condensers; The StarLight Space Interferometer; Champagne Heat Pump; Controllable Sonar Lenses and Prisms Based on ERFs; Measuring Gravitation Using Polarization Spectroscopy; Serial-Turbo-Trellis-Coded Modulation with Rate-1 Inner Code; Enhanced Software for Scheduling Space-Shuttle Processing; Bayesian-Augmented Identification of Stars in a Narrow View; Spacecraft Orbits for Earth/Mars-Lander Radio Relay; and Self-Inflatable/Self-Rigidizable Reflectarray Antenna.
A compact optoelectronic sensor unit measures the apparent motion of the Sun across the sky. The data acquired by this chip are processed in an external processor to estimate the relative orientation of the axis of rotation of the Earth. Hence, the combination of this chip and the external processor finds the direction of true North relative to the chip: in other words, the combination acts as a solar compass. If the compass is further combined with a clock, then the combination can be used to establish a threeaxis inertial coordinate system. If, in addition, an auxiliary sensor measures the local vertical direction, then the resulting system can determine the geographic position. This chip and the software used in the processor are based mostly on the same design and operation as those of the unit described in Micro Sun Sensor for Spacecraft (NPO-30867) elsewhere in this issue of NASA Tech Briefs. Like the unit described in that article, this unit includes a small multiple-pinhole camera comprising a micromachined mask containing a rectangular array of microscopic pinholes mounted a short distance in front of an image detector of the active-pixel sensor (APS) type (see figure). Further as in the other unit, the digitized output of the APS in this chip is processed to compute the centroids of the pinhole Sun images on the APS. Then the direction to the Sun, relative to the compass chip, is computed from the positions of the centroids (just like a sundial). In the operation of this chip, one is interested not only in the instantaneous direction to the Sun but also in the apparent path traced out by the direction to the Sun as a result of rotation of the Earth during an observation interval (during which the Sun sensor must remain stationary with respect to the Earth). The apparent path of the Sun across the sky is projected on a sphere. The axis of rotation of the Earth lies at the center of the projected circle on the sphere surface. Hence, true North (not magnetic North), relative to the chip, can be estimated from paths of the Sun images across the APS. In a test, this solar compass has been found to yield a coarse estimate of the North (within tens of degrees) in an observation time of about ten minutes. As expected, the accuracy was found to increase with observation time: after a few hours, the estimated direction of the rotation axis becomes accurate to within a small fraction of a degree.
A concept study was undertaken to demonstrate how strategic valuation of a micro-sun sensor development could impact decision-making and planning during maturation of technology from proof-of-concept through validation and demonstration.
Miniature high-performance low-mass space avionics systems are desired for planned future outer planetary exploration missions (i.e. Europa Orbiter/Lander, Pluto-Kuiper Express). The spacecraft fuel and mass requirements enabling orbit insertion is the driving requirement. The Micro Navigator is an integrated autonomous Guidance, Navigation & Control (GN&C)micro-system that would provide the critical avionics function for navigation, pointing, and precision landing. The Micro Navigator hardware and software allow fusion of data from multiple sensors to provide a single integrated vehicle state vector necessary for six degrees of freedom GN&C. The benefits of this MicroNavigator include: 1) The Micro Navigator employs MEMS devices that promise orders of magnitude reductions in mass power and volume of inertial sensors (accelerometers and gyroscopes), celestial sensing devices (startracker, sun sensor), and computing element; 2) The highly integrated nature of the unit will reduce the cost of flight missions. a) The advanced miniaturization technologies employed by the Micro Navigator lend themselves to mass production, and therefore will reduce production cost of spacecraft. b) The integral approach simplifies interface issues associated with discrete components and reduces cost associated with integration and test of multiple components; and 3) The integration of sensors and processing elements into a single unit will allow the Micro Navigator to encapsulate attitude information and determination functions into a single object. This is particularly beneficial for object-oriented software architectures that are used in advanced spacecraft. Additional information is contained in the original extended abstract.
NASA is seeking to embark on a new set of human and robotic exploration missions back to the Moon, to Mars, and destinations beyond. Key strategic technical challenges will need to be addressed to realize this new vision for space exploration, including improvements in safety and reliability to improve robustness of space operations. Under sponsorship by NASA's Exploration Systems Mission, the Jet Propulsion Laboratory (JPL), together with its partners in government (NASA Johnson Space Center) and industry (Boeing, Vacco Industries, Ashwin-Ushas Inc.) is developing an ultra-low mass (<3.0 kg) free-flying micro-inspector spacecraft in an effort to enhance safety and reduce risk in future human and exploration missions. The micro-inspector will provide remote vehicle inspections to ensure safety and reliability, or to provide monitoring of in-space assembly. The micro-inspector spacecraft represents an inherently modular system addition that can improve safety and support multiple host vehicles in multiple applications. On human missions, it may help extend the reach of human explorers, decreasing human EVA time to reduce mission cost and risk. The micro-inspector development is the continuation of an effort begun under NASA's Office of Aerospace Technology Enabling Concepts and Technology (ECT) program. The micro-inspector uses miniaturized celestial sensors; relies on a combination of solar power and batteries (allowing for unlimited operation in the sun and up to 4 hours in the shade); utilizes a low-pressure, low-leakage liquid butane propellant system for added safety; and includes multi-functional structure for high system-level integration and miniaturization. Versions of this system to be designed and developed under the H&RT program will include additional capabilities for on-board, vision-based navigation, spacecraft inspection, and collision avoidance, and will be demonstrated in a ground-based, space-related environment. These features make the micro-inspector design unique in its ability to serve crewed as well as robotic spacecraft, well beyond Earth-orbit and into arenas such as robotic missions, where human teleoperation capability is not locally available.
The Space Technology 5 (ST-5) Project, currently in the implementation phase, is part of the National Aeronautics and Space Administration (NASA) s New Millennium Program (NMP). ST-5 will consist of a constellation of three miniature satellites, each with mass less than 25 kg and size approximately 60 cm by 30 cm. ST-5 addresses technology challenges, as well as fabrication, assembly, test and operations strategies for future micro-satellite missions. ST-5 will be deployed into a highly eccentric, geo-transfer orbit (GTO). This will expose the spacecraft to a high radiation environment as well as provide a low level magnetic background. A three-month flight demonstration phase is planned to validate the technologies and demonstrate concepts for future missions. Each ST-5 spacecraft incorporates NMP competitively-selected breakthrough technologies. These include Cold Gas Micro-Thrusters for propulsion and attitude control, miniature X-band transponder for space-ground communications, Variable Emittance Coatings for dynamic thermal control, and CULPRiT ultra low power logic chip used for Reed-Solomon encoding. The ST-5 spacecraft itself is a technology that can be infused into future missions. It is a fully functional micro-spacecraft built within tight volume and mass constraints. It is built to withstand a high radiation environment, large thermal variations, and high launch loads. The spacecraft power system is low-power and low-voltage, and is designed to turn on after separation &om the launch vehicle. Some of the innovations that are included in the ST-5 design are a custom spacecraft deployment structure, magnetometer deployment boom, nutation damper, X-band antenna, miniature spinning sun sensor, solar array with triple junction solar cells, integral card cage assembly containing single card Command and Data Handling and Power System Electronics, miniature magnetometer, and lithium ion battery. ST-5 will demonstrate the ability of a micro satellite to perform research-quality science. Each ST-5 spacecraft will deploy a precision magnetometer to be used both for attitude determination and as a representative science instrument. The spacecraft has been developed with a low magnetic signature to avoid interference with the magnetometer. The spacecraft will be able to detect and respond autonomously to science events, i.e. significant changes in the magnetic field measurements. The three spacecraft will be a pathfinder for future constellation missions. They will be deployed to demonstrate an appropriate geometry for scientific measurements as a constellation. They will be operationally managed as a constellation, demonstrating automation and communication strategies that will be useful for future missions. The technologies and future mission concepts will be validated both on the ground and in space. Technologies will be validated on the ground by a combination of component level and system level testing of the flight hardware in a thermal vacuum environment. In flight, specific validation runs are planned for each of the technologies. Each validation run consists of one or more orbits with a specific validation objective. This paper will describe the ST-5 mission, and the applicability of the NMP technologies, spacecraft, and mission concepts to future missions. It will also discuss the validation approach for the ST-5 technologies and mission concepts.
The sensory, control, and operation systems of the 'MITy' Mars micro-rover are discussed. It is shown that the customized sun tracker and laser rangefinder provide internal, autonomous dead reckoning and hazard detection in unstructured environments. The micro-rover consists of three articulated platforms with sensing, processing and payload subsystems connected by a dual spring suspension system. A reactive obstacle avoidance routine makes intelligent use of robot-centered laser information to maneuver through cluttered environments. The hazard sensors include a rangefinder, inclinometers, proximity sensors and collision sensors. A 486/66 laptop computer runs the graphical user interface and programming environment. A graphical window displays robot telemetry in real time and a small TV/VCR is used for real time supervisory control. Guidance, navigation, and control routines work in conjunction with the mapping and obstacle avoidance functions to provide heading and speed commands that maneuver the robot around obstacles and towards the target.