The Junocam instrument consists of two subsystems: the camera head (CH), which uses a build-to-print copy of the camera head electronics developed for the Mars Science Laboratory (MSL) mission with slightly modified logic and Juno-specific optics and housings, and the Juno Digital Electronics Assembly (JDEA), which contains an image buffer, power conversion circuitry, and the interface to the spacecraft. The camera head is mounted on the spacecraft’s upper deck, while the JDEA is mounted to the side of the main avionics vault on the spacecraft. The two subsystems communicate via a spacecraft-provided wiring harness. Table 2 summarizes the Junocam physical characteristics.
Like previous MSSS cameras (e.g., Mars Reconnaissance Orbiter’s Mars Color Imager) Junocam is a “pushframe” imager. The detector has multiple filter strips, each with a different bandpass, bonded directly to its photoactive surface. Each strip extends the entire width of the detector, but only a fraction of its height; Junocam’s filter strips are 1600 pixels wide (spanning 58o) and about 155 rows high. The filter strips are scanned across the target by spacecraft rotation. At the nominal spin rate of 2 RPM, frames are acquired about every 400 ms. This process is illustrated in Fig. 9. At a perijove altitude of ∼5000 km the nadir pixel, with an angular size of 0.6727 mrad, has a spatial scale of ∼3 km. Over the pole the altitude is ∼one jovian radius and the nadir spatial scale is ∼50 km.
The spacecraft spin rate would cause more than a pixel’s worth of image blurring for exposures longer than about 3.2 ms. For the illumination conditions at Jupiter such short exposures would result in unacceptably low SNR, so the camera provides Time-Delayed-Integration (TDI). TDI vertically shifts the image one row each 3.2 ms over the course of the exposure, canceling the scene motion induced by rotation. Up to about 100 TDI steps can be used for the orbital timing case while still maintaining the needed frame rate for frame-to-frame overlap.
The pushframe imaging mode requires additional processing for image reconstruction. First, each exposed frame is read out to the spacecraft and the desired bands are extracted into 128-pixel-high “framelets”, editing out the unused lines between filters which may suffer from spectral crosstalk. After optional summing and compression, the framelets from all of the frames in an image are transmitted to Earth. The MSSS Ground Data System then treats each framelet as an individual image, using spacecraft attitude telemetry to map-project it onto a planetary shape model. Finally, each map-projected framelet is composited into an overall mosaic by spatial location and bandpass to form an output map.
Camera Head (CH)
Electronics and Detector
The CH electronics are designed around the Kodak KAI-2020 Charge-Coupled Device (CCD) image sensor. This detector has 1640×1214 7.4-micron pixels (1600×1200 photoactive), and uses interline transfer to implement electronic shuttering. The sensor incorporates microlenses to improve its quantum efficiency, which peaks at about 55 %. The CCD quantum efficiency as a function of wavelength is shown in Fig. 10. The “fast-dump” capability of the sensor is used to clear residual charge prior to integration and also allows vertical subframing of the final image.
The output signal from the CCD is AC-coupled and then amplified. The amplified signal is digitized to 12 bits at a maximum rate of 5 Mpixels/s. For each pixel, both reset and video levels are digitized and then subtracted in the digital domain to perform correlated double sampling (CDS), resulting in a typical 11 bits of dynamic range.
All CH functions are supervised by a single Actel RTSX field-programmable gate array (FPGA). In response to commands from the JDEA, the FPGA generates the CCD clocks, reads samples from the analog-to-digital converter (ADC) and performs digital CDS, and transmits the pixels to the JDEA.
The CH operates using regulated 5 V and ±15 V power provided by the JDEA. A platinum resistance thermometer (PRT) on the camera focal plane is read by the spacecraft to provide temperature knowledge for radiometric calibration. An additional pair of PRTs and redundant etched-foil heaters are attached to the outside of the camera head and thermostatically controlled by the spacecraft.
The CH electronics are laid out as a single rigid-flex printed circuit board with three rigid sections. The sections are sandwiched between housing sections that provide mechanical support and radiation shielding, and the flexible interconnects are enclosed in metal covers. For Junocam, additional radiation shielding was required and was incorporated into the housings, which are made of titanium. An additional copper-tungsten enclosure surrounds the image sensor. The total mass of the CH is about 2.6 kg.
A functional block diagram of the Junocam electronics is shown in Fig. 11.
A color filter array with four spectral bands is bonded directly to the CCD, as shown in Fig. 12. The four bands are red (600–800 nm), green (500–600 nm), blue (420–520 nm), selected to meet SNR requirements over the pole and for ease of color reconstruction by the public, and methane (880–900 nm). The Junocam filters were fabricated by Barr Associates. The filter transmissions as measured by Barr from witness samples are shown in Fig. 13 (note that this plot also includes spectral variation of the Junocam optics). The measured center wavelengths and bandwidths are given in Table 3.
The purpose of the methane filter is to image within the narrow methane absorption band centered at 889 nm, to enhance the contrast of higher altitude clouds, as described in Sect. 2. Jupiter is quite dark in this band (<5 % albedo) but fairly bright immediately outside it (>30 %). This is significant because the methane filter bandpass varies from the center to the edges of the field due to the variation in the incidence angle of the lens. An estimate of the impact of this variation is shown in Fig. 14. The dotted line shows the spectrum of Jupiter (Karkoschka 1994), with the methane absorption feature at the center. The red and blue lines show the signal as a function of wavelength for the middle of the field (red) and edges (blue).The blue line shows a 2× higher peak than the red at the shorter side of the bandpass, and a 30 % shorter peak relative to the red on the long side. This difference yields of order 10 % more leakage (i.e., signal not from the bottom of the absorption band) at the edge of the field than at the center. It should be noted that this difference is an estimate based on the measured filter normal incidence transmission and modeled values for the off-normal transmission, and applies to both edges of the field of view.
The Junocam optics are comprised of a 14-element all-refractive lens with a nominal focal length of 11 mm and a field of view of about 58 degrees (horizontal.) T/number varies somewhat across the field and with wavelength, but the nominal on-axis T/number is 3.2. The first five front elements are made of radiation-hard glasses to provide shielding for the remaining elements, and the optics are additionally shielded by a thick titanium housing. An alignment cube is mounted to the optics to facilitate precision mounting on the spacecraft. The Junocam optics, shown in Fig. 15, were fabricated by Rockwell-Collins Optronics.
Junocam Digital Electronics Assembly (JDEA)
As originally proposed, Junocam was to have used a copy of the MSL Digital Electronics Assembly (DEA), which takes raw digital image data from the camera head, compresses it in real time, and stores it in a non-volatile memory buffer for later transmission. However, it soon became apparent that the digital electronics used in the DEA (particularly its Xilinx FPGA) would likely suffer too many radiation-induced upsets from the energetic protons trapped in the jovian radiation belts. Since most of the capabilities of the DEA were unneeded for Juno, we designed a new, more radiation-resistant version, called the Junocam DEA or JDEA.
The JDEA provides regulated power to the camera head, implements a minimal command sequencing capability to manage camera head pushframe operation, receives the raw digital image data from the camera head, applies 12-to-8-bit non-linear companding, and stores the image data in a 128 MB internal DRAM buffer. The CH command/data interface is a three-signal Low Voltage Differential Signaling (LVDS) synchronous serial link transmitting commands from JDEA to CH at 2 Mbit/s and a four-signal synchronous 3-bit parallel interface from CH to JDEA at a rate of 30 Mbit/s. The JDEA also contains a command/data interface with the spacecraft, receiving higher-level imaging commands and returning image data. The command interface is a bidirectional asynchronous RS-422 interface running at 57.6 Kbaud; the data interface is a unidirectional three-signal RS-422 synchronous interface running at 20 Mbits/s.
The JDEA uses an Actel RTSX FPGA. Most of the logic design is inherited from the previously built MSSS context imager on the Mars Reconnaissance Orbiter (MRO CTX) and Lunar Reconnaissance Orbiter Camera (LROC). The power subsystem uses Interpoint components and is derived from the MSL design.
The JDEA electronics are laid out as a single rectangular printed circuit board, sandwiched between housing sections that provide mechanical support and radiation shielding. The JDEA housings are aluminum, since considerable radiation shielding is provided by the spacecraft avionics vault.
As indicated above, there is no software resident in the instrument. All additional processing is performed by software provided by Junocam and running in the spacecraft computer. This software has significant commonality with that previously developed by MSSS for the Mars Odyssey and MRO missions. It is written in ANSI C and uses the VxWorks multitasking facility so that processing can occur when the spacecraft computer is otherwise unoccupied.
The software receives commands to acquire images from the spacecraft’s command sequence engine. Each image command contains parameters such as exposure time, number of TDI stages, number of frames, interframe time, summing, and compression. Optionally, each image can be commanded relative to the spin phase (based on information provided by the spacecraft’s attitude control system) so that only frames that are pointed at the planet need be acquired. The software instructs the JDEA to begin imaging at the appropriate time and then delays until the entire multi-frame image is acquired. It then reads out the JDEA DRAM. The raw image data are stored in spacecraft DRAM and then read out, processed, and formatted for downlinking. Processing consists of frame editing, optional summing, optional median filtering to remove radiation-induced pixel transients, and optional lossy transform-based or lossless predictive image compression.
There are three radiation effects likely to be observable in Junocam images. The first will be persistent hot pixels caused by displacement damage from energetic particle hits to the sensor, primarily trapped protons. Ground testing indicates that at the end of orbit 8, Junocam will have accumulated less than a hundred hot pixels, a number that can be dealt with by onboard median filter processing. There will also be a global increase in dark current, leading to a slow degradation of image quality. We estimate the dark current increase will be less than 2× through the end of orbit 8.
There will also be transient effects caused by high particle flux, effects that will vary from image to image. These effects will be localized to portions of the orbit near perijove. Because of the relatively short exposure and readout time for the visible bands (<0.20 s) and the relatively high signal levels, the impact of these transient effects are limited, and are well within the ability of the instruments median filter to remove (Fig. 16). The integration and readout time for the methane band is about twice as long as for the visible and the signal levels are more than 10× lower, making the transient radiation impact much worse (Fig. 17). Based on this analysis, the methane band will not be usable for a window less than an hour long centered at 0.4 h after perijove.