1.1

CLICK Mission

The CubeSat Laser Infrared CrosslinK (CLICK) mission will demonstrate low SWaP laser communication terminals, capable of conducting full-duplex high data rate downlinks and crosslinks, as well as high precision ranging and time transfer. A joint project between the Massachusetts Institute of Technology (MIT), the University of Florida (UF), and NASA Ames Research Center, CLICK consists of two separate demonstration flights: the initial CLICK-A, which will demonstrate a space-to-ground downlink and serve as a risk-reduction mission, and CLICK-B/C, a crosslink demonstration mission. The CLICK technology demonstration will ultimately enable mission concepts including space radio interferometry, GPS-denied navigation, time synchronization for synthetic aperture telescopes, and improved bandwidth for science data telemetry on SWAP-limited platforms like nanosatellites.

The CLICK mission has its origins in MIT’s Nanosatellite Optical Downlink Experiment (NODE), previously described in Refs. 1 and 2, and leverages the design of UF’s Miniature Optical Communications Transmitter (MOCT).^{3, 4} The CLICK payloads, consisting of laser transceivers and pointing, acquisition, and tracking (PAT) systems, will each fly on a 3U CubeSat bus from Blue Canyon Technologies. The optical downlink and crosslink experiments will be performed in low Earth orbit (LEO), at an altitude of roughly 400 km-600 km. In a concurrent effort, a portable optical ground station (OGS) is being developed for the CLICK mission. The 30 cm aperture Portable Telescope for Lasercom (PorTeL), which features a beacon laser uplink, receiver system, and coarse and fine pointing capabilities, will serve as the ground receiver for the downlink demonstration.⁵

This paper provides an overview of the CLICK-A flight demonstration by describing the mission concept of operations (ConOps) and payload specifications. In addition, the current status of the mission is described. At the time of writing, the final assembly and testing of the CLICK-A payload has been completed and the payload has been delivered for integration with the spacecraft bus. This paper covers the final stages of assembly, integration, and testing for CLICK-A and discusses the procedures and results of the tests, including the functional testing of the transmitter and the pointing, acquisition, and tracking (PAT) system. We discuss implications of these results for on-orbit operations. Finally, the CLICK-BC flight demonstration is described and an update on its current status is presented.

2.1

CLICK-A ConOps

The ConOps of the demonstration are shown in Fig. 1. Following the initial deployment from the International Space Station (ISS), RF communication will be established with the spacecraft using the KSATLite ground station network. KSATLite ground stations will transmit commands to and receive telemetry and mission data from the spacecraft, and the network will relay this information to the mission operations center (MOC), which is located at MIT. The MOC will conduct passes with the KSATLite ground stations and coordinate all spacecraft operations, beginning with a functional check-out of the payload subsystem and including the optical downlink experiments.

Figure 1:

CLICK-A Mission Concept of Operations

The optical downlink experiment is conducted in several stages. Prior to the start of a pass, the spacecraft bus initiates its attitude control sequence to point the payload aperture at the optical ground station. The bus system also powers on the payload to enable heating and initial boot-up. At the beginning of the pass itself, the uplink beacon laser is transmitted from the OGS. Specifications of this laser and the payload transmitter are given in Table 1. On ground, the OGS operates in coarse stage tracking in open loop (CSTOL) mode, in which it will track the spacecraft based on ephemeris data generated by the on-board GPS receiver. The ephemeris data is obtained prior to the start of the demonstration, having been downlinked during a previous RF pass with a KSATLite ground station.

Table 1:

CLICK-A Laser Specifications

Using the body pointing of the spacecraft, the beacon laser signal is detected and acquired using the beacon tracking camera. This initiates the closed-loop fine stage tracking (FST) mode, which uses the payload’s fine steering mirror (FSM). The FSM is controlled using the error offset on the camera sensor between the beacon spot and a spot generated by an on-board calibration laser that has been multiplexed with the payload transmit laser. For more detail, see Ref. 6. With the payload conducting fine pointing, the payload downlink laser is detected by the OGS using its own near-infrared (NIR) tracking camera. The OGS will subsequently operate for the duration of the pass in its coarse stage tracking in closed loop (CSCTL) mode, in which it will track the downlink laser from the spacecraft instead of relying on the ephemeris data. In the final stage of the experiment, the OGS will enter FST mode as well, using its own internal FSM which is similarly controlled using the tracking camera error signal. In this mode, the downlink laser from the spacecraft will be centered on the receiver avalanche photodiode (APD). The signal from the APD is sampled using an oscilloscope throughout the pass for offline analysis.⁷

After the pass has ended, the payload will perform any necessary shut-down activities, such as data logging and transferring, before being switched off by the spacecraft bus.

2.2

CLICK-A Payload Design

The CLICK-A payload consists of a laser transmitter and fine pointing PAT system in a 1.2U form-factor. As shown in Fig. 2, the maximum envelope of the payload is roughly 96mm x 96mm x 138mm in size, and it has a mass of approximately 1.17 kg. The upper half of the payload hosts the optical bench for the fine pointing system, while the lower half of the payload contains the transmitter optoelectronics. Both of these halves sit above the fiber raceway. The mechanical assembly is further detailed in Fig. 3. The optical bench includes the beacon tracking camera, a Matrix Vision mvBlueFOX, and its lens assembly, a Schneider Xenoplan lens, as well as the FSM (from Mirrorcle Technologies) used to conduct beam steering. The payload optomechanical design, which supports the ConOps described in the Section 2.1, is summarized in Fig. 4.

Figure 2:

CLICK-A Payload Structure

Figure 3:

CLICK-A Payload Mechanical Design

Figure 4:

CLICK-A Payload Optical Design

The 1550nm transmit laser and 980nm calibration laser, coupled through wavelength division multiplexing (WDM), are collimated and directed to the FSM. The FSM conducts fine pointing by steering the beam to the dichroic beam splitter (DBS). The transmit laser is reflected and exits via the payload aperture, while the calibration laser is directed using a mirror to the lens assembly and the camera’s focal plane array. The 976nm uplink beacon signal follows a similar path through the DBS to the camera and lens assembly. A double bandpass filter is used at the aperture for stray light isolation.

The lower half of the payload, as illustrated in Fig. 3, contains the transmitter optoelectronics, including a COTS Erbium-Doped Fiber Amplifier (EDFA) and modulation and control electronics board stack. The primary controller of the payload is the Raspberry Pi-based CPU board, which provides command and data handling and interfaces with the spacecraft bus. An additional microcontroller is utilized to enable the Raspberry Pi to be reprogrammed on orbit. The CPU board also interfaces with the beacon camera and runs the software for the PAT sequence. A field programmable gate array (FPGA) is utilized for the control and modulation of the transmitter seed laser. The transmitted signal is modulated using M-ary pulse position modulation (PPM), which is implemented with a pulse duration of 5 ns and a variable PPM order ranging from 4 to 128. The FPGA, a Xilinx Spartan 6, is located on its own board with additional circuitry for current sensing and interfacing with the daughterboard. The daughterboard is primarily an interface board for the signals from the FPGA to the various other components. It includes interfaces and driver electronics for the payload heaters, resistance temperature detectors (RTDs), FSM, photodiodes, calibration laser, and transmitter seed laser. This system architecture is shown in Fig. 5.

Figure 5:

CLICK-A Payload System Architecture

The 1550nm downlink transmitter uses a master-oscillator power-amplifier (MOPA) design based on a COTS transmitter optical sub-assembly (TOSA) for the seed laser and an EDFA for the optical amplifier. The transmitter design is based on that of NODE, as previously described in Refs. 8-1. The FPGA controls the bias current and temperature (using a thermo-electric cooler, or TEC) of the seed laser diode in order to shift the output signal wavelength in and out of the passband of a Fiber Bragg Grating (FBG), thereby achieving a high extinction ratio. A modulated signal from the FPGA, along with the TEC control, is applied to the TOSA to induce a signal of a certain wavelength which is then directed via optical fiber to the FBG through a circulator. Wavelengths within the passband of the FBG, signifying a PPM pulse, are reflected back into the circulator and directed to the EDFA. Wavelengths outside the passband, indicating an empty PPM time slot, are transmitted through the FBG to an attenuator and photodetector. The EDFA amplifies the signal and directs it to a collimator to achieve the specified beam divergence. This design is illustrated in Fig. 6.

Figure 6:

CLICK-A Transmitter Architecture

Three PIN photodiodes are used to measure the optical power at various points to provide feedback and allow closed-loop control of the transmitter. The FPGA utilizes these measured values to adjust the bias current and temperature of the laser diode to finely tune the transmitter wavelength. This also enables the built-in self-test (BIST) capability of the transmitter to verify the output signal.

A comprehensive discussion of the transmitter design can be found in Ref. 8. The performance of this system has previously been evaluated and an analysis of the communication link has been completed in Refs. 6-7.

2.3

CLICK-A Status

Recently, the final assembly and testing of the CLICK-A payload has been completed and the payload has been delivered for integration with the spacecraft bus. The fully assembled flight model is shown in Fig. 7. The final testing of the payload focused on the validation of the optoelectronic and mechanical design of the transmitter and PAT system. In parallel, the payload flight software was developed, tested, and finalized as well. For environmental tests, the assembled payload was subject to thermal vacuum testing in various thermal conditions. Ultimately, the testing campaign conducted at MIT sought to demonstrate the functional operation of the payload and characterize its performance. Select results from this testing are presented and discussed here.

Figure 7:

CLICK-A Payload Flight Model

To enable the functional testing of the payload on ground, an all-in-one instrument for the ground support equipment (GSE) was developed. The design of the GSE is illustrated in Fig. 8. Equipped with its own beacon laser, FSM, APD, and IR camera, as well as an M² measurement system from Thorlabs for beam profiling, the GSE enables the measurement and characterization of the transmitter beam quality and divergence, received signal pulse shape and amplitude, and the accuracy of the PAT system. In order to streamline testing further, automated scripts were developed to run with the payload controller and GSE. These self-test scripts will also be run via command on orbit to check the functionality of the payload.

Figure 8:

CLICK Payload Testing Ground Support Equipment (GSE)

The testing of the transmitter included confirming the seed laser power and modulation, characterizing the pulse shape, measuring the peak power of the EDFA, and examining the beam quality in a variety of environmental conditions. Figure 9 shows the power of the seed laser for the CLICK-A payload over its output wavelength ranges at various PPM orders as well as the insertion loss of the FBG used for modulation. Due to chirp, the wavelength of the seed laser is shifted during pulses. The amplitude at the shifted wavelength, around 1550.13 nm, is proportional to the duty cycle and is divided by 2 for each increase in PPM order. The light sent to the EDFA is the amount reflected on the FBG, so the convolution of the insertion loss and the seed laser power.

Figure 9:

CLICK-A Transmitter Seed Laser Performance

In addition, the transmitter peak output power was recorded at various duty cycles during environmental testing, as illustrated in Fig. 10. As a key factor to ensuring the link performance, this measurement was particularly important. To measure the peak output power, a linear photodiode was used; the output power was back-calculated from the output current of the photodiode using the circuit and component specifications. The EDFA behaves in an average-power-limited manner, resulting in high instantaneous power. As shown in Fig. 10, the lower duty cycle waveforms (corresponding to higher orders of PPM) produce higher peak output powers. The measured output power tracks closely with the optimal performance, though some efficiency is lost at the highest order of PPM (PPM-128). These measurements, taken under vacuum and in various thermal conditions, serve to validate the functioning of the transmitter on orbit, and confirms the specifications used in previous link budget analyses. In addition to the recorded peak power, the modulated output of the transmitter was examined, as exemplified in Fig. 11. The output of the photodiode shown in this Figure contains a PPM-4-modulated message, prepended by a 16-symbol preamble. The steady train of pulses before and after the message provide a consistent duty cycle to maintain the stability of the EDFA.

Figure 10:

CLICK-A Transmitter Output Peak Power

Figure 11:

CLICK-A Modulated Output

The CLICK-A PAT system was also thoroughly tested. This involved checking the calibration laser, camera, and FSM actuation, confirming that the beacon detection and standalone calibration procedure could be successfully performed, and measuring pointing performance in vacuum over the operational temperature range of 0-40 °C. The PAT performance was measured using the IR camera in the GSE, which measures the transmit spot and beacon spot locations via a Gaussian fit centroiding algorithm that first fits the larger transmit spot before subtracting it from the image to fit the smaller beacon spot (See Figure 12a). The distance between the two centroids in pixels is measured and converted to angular error using the properties of the IR camera optics: 100 μrad/pixel. The Gaussian statistics for angular pointing error in each axis are computed, and the worst mean and standard deviation values from each axis are input to the pointing budget for link budget analysis. The details of the pointing budget and link analysis can be found in Ref. 7. In brief, a simulated typical LEO-ISS overpass of the optical ground station location at MIT Wallace Astrophysical Observatory was used to generate range data (the minimum range is 517.6 km). The pointing and link budgets were run over the pass to generate the time-varying performance data. Figure 12b shows the link margin performance for the ambient temperature vacuum test. The mission requirement of ≥ 10 Mbps was shown to be satisfied for PPM orders less than or equal to 32. The performance details are listed in Table 2.

Figure 12:

CLiCk-A PAT Performance During TVAC Testing

Table 2:

CLICK-A PAT Performance Summary

It should be noted that the measured pointing error is only one component of the pointing budget, the full details of which are found in Ref. 7. In summary, the PAT performance requirement was met for vacuum tests across the operational temperature range.

In addition to the testing of the CLICK-A payload itself, the interface between the payload and the BCT bus system was tested. This was enabled by the use of an engineering demonstration unit (EDU) of the bus, which not only provided the bus data interface to the payload but was also capable of routing command and telemetry and interfacing with the MOC computer through the COSMOS mission control software application. This allowed for the parallel development of mission operations scripts and the end-to-end testing of the payload using the COSMOS software, simulating the on-orbit operations scenario. In this manner, all commands, telemetry, and the payload reprogramming capability were tested and confirmed.

3.1

CLICK-B/C Status

The bulk of recent developments has focused on the final steps of the CLICK-A payload. However, concurrent work on CLICK-B/C has allowed this effort to progress as well, primarily on the payload electronics. Currently, a complete set of boards has been fabricated and is under test. Software development has seen the FPGA controlling the payload peripherals to be working. In addition, a new tunable seed laser, capable of emitting any wavelength within 1535nm and 1565nm has been demonstrated. Using a semiconductor optical amplifier as a modulator, 10 ns pulses have been achieved. Regarding other elements of the payload, the pointing detector, a quadcell, has been previously tested.¹⁰ The detection, sub-modulation filtering and FSM control are functional. The APD circuit used in the payload is identical to the PorTeL and OGS APD. The optical bench of the payload has been assembled, and optical characterization of the prototype is ongoing.