2.1

Microsatellite size/mass/power thresholds

The key contributor to the uniqueness of this proposed TIR EO instrument are the severe mass and size constraints presented by the DMC micro-satellite platform. Although the trend for TIR EO instruments leverages recent advances in the miniaturization of infrared imaging technologies, very few proposed instruments have yet been taken to this extreme. The current limits taken by the TIR EO instrument include those set by DLR’s 30 kg BIRD instrument [3] and those proposed by NASA and ESA with the use of uncooled micro-bolometer arrays [9-11]. Although these designs make it possible to fly high-performing TIR (and mid-wave IR) instruments on micro-class satellites (< 100 kg), they would easily consume the entire payload budget for a DMC micro-satellite. Whilst the flight of these instruments in micro-satellite constellations have potential to fill an important TIR EO niche, it is the focus of Surrey in this study to push TIR EO instrument miniaturization to yet another level. The reasons for doing this are as follows. First, by developing a radically small TIR EO instrument, it can more easily be integrated into the already existing DMC imaging suite. This includes capabilities in medium (32 metre GSD) to medium-high (4 metre) resolution visible and NIR EO. Also, since the TIR is a relatively new development area for Surrey, technological risks can be significantly reduced by starting simple and small much like the earlier development of its current generation of visible and NIR EO capabilities. By slowly introducing TIR to the DMC constellation, the market and utility can be tested at a relatively low cost. Therefore, for this study, the following mass and size thresholds have been determined iteratively with the mission utility thresholds and the instrument design described in the next sections.

Tab. 1:

Baseline constraints for the TIR imager

2.2

Mission utility thresholds

The second boundary for the trade space are defined by the acceptable levels of GSD and NETD, in light of the increases in revisit time brought by the constellation. Traditionally, spatial (GSD) and radiometric (NETD) resolution requirements for new TIR EO instruments result from an active dialogue and eventual compromise between the user community, who always seeks greater spatial, radiometric, and temporal resolution, and instrument design engineers who determine what is physically and economically feasible. This compromise has solidified over the years and both the user community and instrument developers have since grounded their processes to the capabilities of heritage systems. As a result, somewhat of a departure must be made to validate the new class of data enable by micro-satellite constellations.

Therefore, when evaluating the mission utility of adding a TIR band to DMC, the existing field of TIR requirements has been used as a starting point but has been adapted to exploit the advantages brought by small satellites (reducing costs of existing data sets or providing niche services currently ill-served). The advantages or niche provided by this type of data has been illustrated in Fig. 2. In the upper half of the figure, representative course spatial but high temporal data is compared with more precise spatial but low temporal data (group 1 in Fig. 1). Although the latter data set is much better suited at geo-locating TIR phenomena, the former is better at monitoring the migration of thermal events across the scene.

Fig. 2:

Less precise TIR EO data (spatially and radiometrically) becomes more useful with decreased revisit time

In the lower half of Fig. 2, course radiometric but high temporal data is compared with precise radiometric but low temporal data (again group 1). In this case, the latter data set is better at determining the surface temperature/emissivity characteristics of the scene. However, the former is much better suited to detect a brief change in scene temperature/emissivity.

Therefore, in defining suitable mission areas for this new class of compact TIR EO instrument for flight in multi-satellite constellations, two general mission areas were the focus: 1) analysis of trends, and 2) detection of anomalous changes in the spatial or radiometric characteristics of the TIR signature of the Earth’s surface or atmosphere. Development of a sound set of mission applications for this new class of TIR EO data still requires a rigorous dialogue with the user community. However, using a 2003 ESA study (Thermal High-Resolution Earth Mapper - THEMA) reference [12], the following candidate mission applications have served as the starting point for this instrument development:

Tab. 2:

Sample Mission application for TIR EO micro-satellite constellations [12]

The THEMA study provides a good starting point because it represents a recent and thorough analysis of user requirements and the general acceptance by the TIR data users. Of special interest is forest fire detection, which, if used in conjunction with on-board detection algorithms and autonomous ground user data transfer, would fill an important niche in the global fire detection community.

2.3

Trade study (size/mass vs. performance)

With the size/weight and utility objectives defined, the iterative generation of the remainder of the instrument’s requirements was performed. During this process, trade-offs between performance (GSD, NETD, and revisit time) and estimated size, mass, and power requirements were made. For our instrument, NETD could be decreased by increasing the amount of scene energy available to the detector, i.e. lowering the f/#, increasing the spectral bandwidth, or employing time delayed integration (TDI). In the same way, spatial resolution and ground coverage could be increased or decreased by adjusting the imager focal length within the payload size limitations. The sub 8 - 12 µm wavebands in Tab. 3 have been chosen based on heritage TIR missions and the inevitable requirement for atmospheric correction, discriminating ground features, and customisation (e.g. utilizing a split window technique). Taking all these factors into account the following list of requirements was used for the baseline design.

Tab. 3:

Summary of requirements for TIR imager

2.4

Typical microsatellite mission

Given the fundamental limits encountered by a low-cost, compact imager, it’s performance as a complement to a larger imaging suite becomes important. In this section a potential mission concept resulting from the integration on a 90 kg DMC-type micro-satellite will be briefly presented. Given the varied specifications required by the multiple owners of DMC, there are several different versions of the DMC spacecraft. However, the orbital characteristics are all essentially the same; Sun-synchronous with a local time of the ascending node at 10:00 am, at near 700 km altitude. This has been chosen so that between 4 and 8 spacecraft can provide daily coverage at the Equator with 300 – 600 km ground swaths. A typical DMC imager is composed of six push-broom CCD sensors, configured in two banks of three (red, green and NIR), as part of an overall optical bench assembly. Two tri-axial vector magnetometers and four dual-axis sun-sensors are carried in order to determine the attitude of the spacecraft to better than ± 0.25 degrees, however control is relaxed at ±1 degree. For the addition of a TIR channel (also as an imager pair), the baseline spacecraft EO payload would most likely have the following characteristics:

Tab. 4:

Potential micro-satellite TIR imaging mission

The entire EO payload (3 imager pairs) would weigh less than 10 kg. While a wideband (8 - 12 µm), 325 m GSD instrument has been used in this example, the TIR channel for a real DMC mission would be customized to accommodate specific mission requirements.

3.1

Cooled vs. uncooled detector technology

The goal in detector selection has been to find candidate commercially available detector array that offer the best performance in terms of weight, size, power consumption, and cost. The TIR EO designer has a variety of cooled technologies to choose from, including ever-shrinking cryo-coolers that are increasingly more power efficient and relatively stable with regards to inducing micro-vibrations. However, in this study, we are looking for an order of magnitude decrease in mass–moving to approximately the 2-3 kg mass range–whilst still achieving reasonable mission utility. Cooled TIR technology may eventually become part of the DMC EO suite, but given reasons outlined in section 2.1, un-cooled detectors have become a practical first step to image in the TIR band.

3.2

Choices in un-cooled technology

In this survey the three main classes of un-cooled infrared technology (pyroelectric, thermoelectric, and resistive bolometer) were initially evaluated. The last class of un-cooled infrared technology, resistive bolometers, has recently been the most aggressively developed of the three for terrestrial applications. Resistive bolometers have been around for decades, but it has been the recent developments in MEMS and thin film technologies that has led to a surge in commercially available micro-bolometer arrays. These arrays are finding multiple markets in fire fighting, security, and manufacturing, which encourages streamlined production processes and increasingly less expensive and better performing devices.

After examining over 32 models from 11 different manufacturers, we have initially settled on a French COTS 320 × 240 amorphous silicon micro-bolometer array on which to base our preliminary design. The array has a pixel pitch of 45 µm with an 80 percent fill factor. The detector response is reported to be 4 mV/K and the thermal time constant is 4 ms. The NETD is given as 120 mK with f/1 optics observing a 300 K target. In addition, the array comes integrated with a thermally stabilized Peltier cooler, and read-out integrated circuit (described in section 3.4). The values for peak responsivity (ℜ) and rms noise (V_N) were not directly specified but were calculated to be approximately 7 × 10⁶ V/W and 480 µV respectively from the specified parameters. These values are consistent with data found in other previously published reports. [11]

The key rationale this array was chosen was that it is fairly representative of the state of the art and is readily available within the European academic community. However, the nature of rapid micro-satellite development times supports the insertion of any technology that is deemed appropriate during the initial design phase. For subsequent TIR imager designs, other COTS detector arrays may be selected using the best fit at the time.

3.3

Initial optics design

The final optical design will not be defined until after the laboratory test phase of our development cycle. However, a preliminary concept has been laid out given the required point spread function (PSF) and modulation transfer function (MTF) derived from the baseline requirements. For the bench-top prototype, the most liberal requirements were chosen to facilitate a low development cost and a short development timeline. The resultant baseline design is f/1.2 refractive lens triplet with a focal length of 100 mm.

3.4

Initial electronics design

At the micro-bolometer pixel level, absorbed incident radiation causes a change in pixel temperature, which, in the case of semiconducting bolometric materials, reduces its electrical resistance. Since the pixel is pulse biased with a set voltage, the resultant change in current signal can be measured by the detector read out integrated circuit (ROIC), which is part of the COTS detector package. After a thermal scene is ‘imaged’ on the array, the respective signal currents are reduced by a dark pixel current and then integrated into a voltage signal row by row at the bottom of each array column (ripple operation). In turn, all 76,800 pixels are read out as a 5 MHz analogue video signal resulting in a 50 Hz frame rate. One of the key advantages of using a COTS detector is the ease of electrical interfacing via the ROIC. The video output from the ROIC is similar to that found on some of our current imager systems, and so we had made use of our design heritage for image acquisition, storage and downloading.

After leaving the detector ROIC, the analogue video signal enters the SSTL engineered electronics. The video signal is first digitised to an 12-bit digital stream and then read into a solid-state data-recorder (SSDR). The data is then held until ready for transmission to Earth under the control of a field programmable gate array (FPGA), which links the SSDR to a dedicated downlink. The FPGA also contains all the sequencing and timing logic for system, as well as all the control signals needed to operate the imager under the various data collection scenarios. The command and control of the FPGA is performed by a microcontroller. The microcontroller sends and receives telecommands and telemetry though Control Area Network (CAN) packets from other on-board controllers. It can also provide on-board image processing such as automatic feature detection and data-compression. The downlink speed will depend on the particular host satellite. Typical rates for microsatellites range from 9.6 kbps to a few Mbps.

3.5

Calibration

One of the key challenges in flying TIR imagers on very small spacecraft is that of instrument calibration. In order to accurately measure ground feature temperatures, reference blackbodies are typically temporarily inserted within the imager field of view to calibrate the effects of inevitable drifts in the imager parameters. Standard practice with heritage systems is to utilize one cold and one hot blackbody to accurately reference the radiometer prior to taking each image. On larger systems, this has been done with the same steering mirror used for image scanning in a whiskbroom configuration. In smaller systems, mechanical arms, mirrors, or paddles are used. In most cases, cold space (3 K) is used as the required cold body and the hot body is on-board the spacecraft bus.

For our system, calibration schemes can complicate what has been so far a relatively simple and low-cost design. For some mission areas such as those focusing on thermal contrast instead of exact radiometric measurements, on-board calibration may not be required so a non-calibrated system is definitely a consideration for this design. However, one of the advantages of a small, nimble spacecraft is its ability to manoeuvre. The final calibration method is to be determined but heavy emphasis will be given to the use of manoeuvre (spacecraft pan towards cold space) or settling strictly on relative variation mapping with potential cross-calibration with other instruments (such as AATSR). However, consideration will be given to on-board calibration when mission requirements dictate.