In order to increase the fidelity of hardware-in-the-loop ground-truth testing, it is desirable to create a dynamic scene of multiple, independently controlled IR point sources. ATK-Mission Research has developed and supplied the steering mirror systems for the 7V and 10V Space Simulation Test Chambers at the Arnold Engineering Development Center (AEDC), Air Force Materiel Command (AFMC). A portion of the 10V system incorporates multiple target sources beam-combined at the focal point of a 20K cryogenic collimator. Each IR source consists of a precision blackbody with cryogenic aperture and filter wheels mounted on a cryogenic two-axis translation stage. This point source target scene is steered by a high-speed steering mirror to produce further complex motion. The scene changes dynamically in order to simulate an actual operational scene as viewed by the System Under Test (SUT) as it executes various dynamic look-direction changes during its flight to a target. Synchronization and real-time hardware-in-the-loop control is accomplished using reflective memory for each subsystem control and feedback loop. This paper focuses on the steering mirror system and the required tradeoffs of optical performance, precision, repeatability and high-speed motion as well as the complications of encoder feedback calibration and operation at 20K.
RAMOS, the Russian American Observational Satellite program, is a cooperative space-based research and development program between the Russian Federation and the United States. The planned system configuration is a constellation of two satellites orbiting in approximately the same plane at an altitude of about 500 km, separated from one another by a variable distance centering on about 500 km. These satellites are equipped with passive electro-optical sensors, both US- and Russian-built, that operate over a range from infrared (IR) to ultraviolet (UV) and are designed for near-simultaneous stereo imaging capability. The sensor suite will include visible, IR and UV imaging radiometers, an IR spectrometer, and a short-wave infrared (SWIR) polarimeter. The projected launch date is 2008 with a planned minimum on-orbit lifetime of two years, and a five-year lifetime possible. This paper summarizes the program objectives, anticipated measurements and expected data, and presents the basic system design, expected performance characteristics, and the capabilities of each of the sensors.
The Space Dynamics Laboratory at Utah State University (SDL/USU) has built and flown an airborne hyperspectral imaging polarimeter (HIP)1,2 as a proof-of-principle sensor for a satellite-based polarimeter. This paper discusses measurement limitations and uncertainties associated with imaging polarimetric measurements in remote sensing applications, using experience and lessons learned from the HIP program and the design study for the proposed satellite demonstration sensor.
The Space Dynamics Laboratory at Utah State University has built and flown an airborne infrared Hyperspectral Imaging Polarimeter (HIP) as a proof-of-principle sensor for a satellite-based polarimeter. This paper briefly reviews the instrument design that was presented in previous SPIE papers1,2, details the changes and improvements made between the 1998 and 1999 measurements, and presents measurement data from the flights.
Measurement data from a series of flights in 1998 indicated the need for wider-band measurements than could be made with our ferroelectric liquid crystal polarimeter design. For this reason, the existing sensor was modified to use a rotating wire-grid polarization filter. The reasons for this choice, equipment design, and measurement equations will be given. A short description of the 1999 flights aboard FISTA3 (Flying Infrared Signatures Technology Aircraft), an Air Force KC-135 based at Edwards Air Force Base will be given, as well as a small sample of the four-dimensional data set will be presented.
To derive the polarization characteristics of a remotely sensed object, a time-sequential polarimeter must create multiple polarization response states during the course of each measurement set. A common method of creating these states is to rotate a polarizer element to a discrete location and hold that position while the detectors integrate and are sampled. The polarizer element is then rotated to the next position and the process is repeated. This time-sequential, advance-and-hold technique is widely used and easily understood because of its simplicity. However, it is not well suited for remote sensing applications where time delays caused by the advance-and-hold mechanism can limit measurement speed and reduce measurement accuracy. This paper introduces a continuously spinning polarizer (CSP) technique that eliminates the time delays and associated problems of an advance-and-hold polarimeter. A performance model for a linear Stokes polarimeter containing a CSP is derived, and a demonstration of the CSP technique based on the performance of the hyper-spectral imaging polarimeter (HIP) is presented.
The Space Dynamics Laboratory at Utah State University is building an infrared Hyperspectral Imaging Polarimeter (HIP). Designed for high spatial and spectral resolution polarimetry of backscattered sunlight from cloud tops in the 2.7 micrometer water band, it will fly aboard the Flying Infrared Signatures Technology Aircraft (FISTA), an Air Force KC-135. It is a proof-of-concept sensor, combining hyperspectral pushbroom imaging with high speed, solid state polarimetry, using as many off-the-shelf components as possible, and utilizing an optical breadboard design for rapid prototyping. It is based around a 256 X 320 window selectable InSb camera, a solid-state Ferro-electric Liquid Crystal (FLC) polarimeter, and a transmissive diffraction grating.
Satellite laser communication concepts have been under development for many years. The conventioanl approaches require sophisticated hardware and considerable spacecraft resources introducing concerns about cost, added weight, power consumption, and reliability. An optical tranceiver based on a modulating retroreflector is a relatively new concept which has not been explored for space communications. The majority of the hardware and complexity for such a communications link is located on the ground and only minimal spacecraft hardware is required. This technique can provide a modest telemetry link for spacecraft in low earth orbits while consuming negligible spacecraft resources when compared to a traditional RF system. A prototype for such a low power optical tranceiver has been constructed and tested over a 4 km ground path in preparation for a high altitude balloon demonstration. Presented here is an overview of the retromodulator communications concept, a link design, and results from prototype testing.
The Space Dynamics Laboratory at Utah State University built an infrared imaging radiometer with dual, large-format detector arrays and a passively cooled telescope for low earth orbit. The confocal detector arrays include a 128 X 128 HgCdTe array operating from 4.5 to 7.5 micrometers and a 256 X 256 InSb array operating from 2.0 to 4.5 micrometers . These arrays yield simultaneous dual-band images. A 13 cm aperture, passively cooled telescope with single- axis scan mirror gives high system sensitivity, excellent image quality, and precision tracking of targets and backgrounds without the usual complexity of cooled optics. High speed cryogenic filter wheels with 6 to 8 filters per detector provide for rapid band selection. A modular cooling system allows the detector arrays and filters to be cooled using either a mechanical cryocooler or a solid cryogen cryostat depending on mission requirements. An on-board calibration source performs pixel-to-pixel uniformity correction on- orbit.