TechDemoSat-1 (Technology Demonstration Satellite-1) / TDS-1
In October 2010, SSTL (Surrey Satellite Technology Ltd.) was awarded a grant to commence the design phase of a UK national technology demonstration satellite called TechDemoSat-1. The UK’s Technology Strategy Board (TSB) and the South East England Development Agency (SEEDA) are funding the design of the core elements of the mission. SSTL, UK industry and UK academia are funding the payload technologies. 1) 2) 3) 4)
TechDemoSat-1 (or simply TDS-1) will be the second collaborative UK satellite launched since the establishment of the UK Space Agency (UKSA, April 1, 2010) and will demonstrate the advanced capabilities of state-of-the-art small satellite technology for scientific and commercial applications. It will also be among the first missions to make use of the ground station facilities that are currently under construction at the UK’s new ISIC (International Space Innovation Centre) at Harwell, Oxfordshire.
From the outset, the mission concept has been industry-led — the culmination of several leading British space companies’ and universities’ efforts to demonstrate the UK’s technical lead in space innovation. As the project leader, SSTL will cooperate with the research and commercial partners providing the payloads in building, testing and operating the satellite.
Figure 1: Illustration of the TechDemoSat-1 small satellite (image credit: SSTL)
TechDemoSat-1 has been developed at SSTL under sponsorship of the recently formed UKSA (UK Space Agency) and with contributions from the payload suppliers. It is intended to be the first of a series of UK technology satellites aiming to provide a rapid affordable means of testing and proving the next generation of space hardware in orbit. 5)
TDS-1 is a satellite platform, derived from heritage technology, which will function as an ‘in-orbit test facility’ for innovative UK payloads and software. The minisatellite is based on the SSTL-150 platform with a size of 77 cm x 50 cm x 90 cm. With its 8 payloads, the satellite has a launch mass of ~ 160 kg. It is capable of accommodating around 52 W of on-orbit average power. TDS-1 features 4 wheel slew agility, and a new generation of star trackers, gyros, magnetometers and torque rods.
Figure 2: Block diagram of the TechDemoSat-1 spacecraft (image credit: SSTL)
RF communications: TDS-1 has S-band and X-band downlinks capable of operating with experimental downlink speeds up to 400 Mbit/s. Up to 128 GB of payload data can be stored onboard. -
• The TDS-1 spacecraft passed the PDR (Preliminary Design Review) on April 18, 2011.
Figure 3: Photo of TechDemoSat-1 flight ready in SSTL's cleanroom March 2013 (image credit: SSTL)
• The TechDemoSat-1 spacecraft passed its FRR (Flight Readiness Review) in early July 2013.
Launch: A launch of TechDemoSat-1 is now scheduled for Q1 2014 on a Soyuz-2-1b launch vehicle with Fregat-M upper stage from the Baikonur Cosmodrome, Kazakhstan. SSTL has signed an agreement with Glavkosmos / NPO Lavotchkin for the launch of the UK technology demonstration mission. 6)
The primary payload on this flight is the Russian meteorological satellite Meteor-M2 (or Meteor-M N2). The launch of Meteor-M2 was already delayed several times.
The planned mid-December 2013 launch has been delayed again following the latest series of issues with the primary satellite payload. Russia’s Meteor-M2 polar-orbiting meteorological satellite has faced delays in the past that have kept the secondary payloads — Norway’s AISSat-2, Canada’s M3MSat, Britain’s TechDemoSat-1 and UKube-1 among them — on the ground. 7)
The Soyuz delay follows a two-year grounding of the Russian-Ukrainian Dnepr silo-launched rocket, which small-satellite owners hope may now be returning to the market pending an agreement between Russia’s military space forces and the Russian space agency, Roscosmos.
Orbit of the primary payload: Sun-synchronous circular orbit, altitude of ~ 820 km, inclination = 98.8º.
The secondary payloads on this flight are:
• Baumanets-2, a technology microsatellite (~100 kg) of BMSTU (Bauman Moscow State Technical University)
• Monika-Relek (or MKA-PN2), a Russian microsatellite (solar and magnetosphere research)
• Venta-1 / V1-QSPnP1 (V1-QuadSat-PnP-1) the first nanosatellite (7.5 kg) project of Latvia built by LatSpace SIA of Ventspils.
• TechDemoSat-1 of SSTL, UK with a mass of ~160 kg
• UKube-1, a nanosatellite of UKSA.
• M3MSat, a microsatellite (85 kg) of Canada.
• SkySat-2 of Skybox Imaging Inc. of Mountain View, CA, USA, a commercial remote sensing microsatellite of ~100 kg.
• AISSat-2, a nanosatellite with a mass of 6.5 kg of Norway.
Orbit of the secondary payloads: Sun-synchronous orbit, altitude of ~ 680 km, inclination of 98º.
Sensor/experiment complement: (Maritime Suite, Space Environment Suite, Air and Land Monitoring Suite, Platform Technology Suite)
Table 1: Overview of TDS-1 payloads
Maritime Suite: SSP (Sea State Payload)
The SSP (Sea State Payload), will test technologies that can measure ocean conditions by GPS reflectometry and by radar altimetry.
SGR-ReSI (Space GNSS Receiver-Remote Sensing Instrument)
The SGR-RESI uses an enhanced GPS receiver with the objective to monitor reflected signals to determine ocean roughness. The instrument is a highly versatile and uniquely capable GNSS receiver that has been designed with the low-cost microsatellite market in mind. Based on COTS parts, the SGR-ReSI builds on the expertise SSTL and SSC (Surrey Space Centre) have developed in the field of GNSS signals, navigation and remote sensing from Low Earth Orbit. 8) 9) 10) 11) 12) 13)
Heritage: The UK-DMC small satellite, launched in Sept. 2003, carried already a pioneering experiment, the SGR-10 GPS receiver with 3 RF sections, able to collect samples containing reflected GNSS signals from the ocean surface. This instrument generated data showing the feasibility of GNSS reflectometry as a remote sensing technique, with special application to ocean roughness sensing. The primary limitation of this experiment was the small amount of data that could be collected – only 20 seconds each time. The data was sufficient to demonstrate after post-processing a relationship between reflections and sea state.
The objective of SGR-ReSI is to provide a highly capable yet relatively compact and affordable way of studying the Earth from orbit. The core technology can also be reused for a new family of navigation-grade receivers that will ultimately replace SSTL’s heritage SGR receivers which include, amongst others, the SGR-20, SGR-10, SGR-07 & SGR-GEO.
Table 2: SGR-ReSI advances over the UK-DMC GPS experiment
To allow for a wide range of possible applications, the prototype SGR-ReSI supports up to 8 separate RF inputs, enabling up to 4 dual frequency antennas. Two varieties of COTS-based RF front-ends have been incorporated, one of which is L1 optimized (MAX2769) while the other is re-configurable to any of the GNSS bands (MAX2112). The RF front-ends are housed on daughter boards with a reasonably generic interface scheme, allowing the adaptation of new front-ends without requirement of a motherboard redesign, and other front-ends are also being considered for the future.
A schematic of the SGR-ReSI is shown in Figure 4. The co-processor was specifically included for the real-time processing of the raw reflected GNSS data into DDMs (Delay-Doppler Maps). However, it has flexibility to be programmed in orbit as required for different purposes, for example to track new GNSS signals, or to apply spectral analysis to received signals. 14)
Figure 4: Configuration of the SGR-ReSI GNSS reflectometry instrument (image credit: SSTL)
At its heart, the SGR-ReSI is a highly versatile GNSS navigation receiver. The receiver core is a 24-channel L1 receiver with support for up to 4 antennas. The SSTL developed GNSS correlators are implemented alongside a LEON3 softcore processor in an Actel ProASIC3E FPGA (Field Programmable Gate Array). The Actel ProASIC3E is a non-volatile flash-based FPGA that consumes little power, providing a low-power. It contains the processor and 24 channels of SSTL-developed correlators, plus other peripherals, such as UART, CAN-bus, and SPI (Serial Peripheral Interface). The RTEMS (Real-Time Executive for Multiprocessor System) operating system has been selected to allow control over multitasking application software, portability between different processors, and compatibility with other developments at SSTL.
Figure 5: SGR-ReSI prototype with 8 front-ends (image credit: SSTL)
While the first FPGA contains the core functions of a GNSS receiver, a second FPGA is available as a co-processor. This is a Xilinx Virtex 4 FPGA that is SRAM based, allowing the upload of new co-processing algorithms even while in orbit. It enables special processing algorithms for reflected or occulted signals, allowing the equivalent of hundreds of correlators to map the distorted signals. To allow the storage of both sampled and processed data, a bank of DDR2 SDRAM (Double Data Rate Synchronous Dynamic Random Access Memory) devices with a capacity of 1 GByte is used.
The instrument supports multiple interfaces (CAN, RS422, SpaceWire, USB) so it can be accommodated by a variety of different satellite missions. The unit is around 1 kg in mass, consumes less than 10 W, and fits within half of an SSTL standard satellite micro-tray (approx 300 x 160 x 30 mm).
For the co-processor to generate Delay Doppler Maps of the sampled reflected data, it needs to be primed with the PRN, the estimated delay and the estimated Doppler of the reflection as seen from the satellite. These are calculated by the processor in conjunction with the main navigation solution - the data flow for this is shown in Figure 6. Direct signals (from the zenith antenna) are used to acquire, track GNSS signals. From the broadcast Ephemerides, the GNSS satellite positions are known. Then from the geometry of the position of the user and the satellites, the reflectometry geometry can be calculated, and hence an estimate of the delay and Doppler of the reflection.
Figure 6: GNSS reflectometry dataflow (image credit: SSTL)
The processing of the Delay Doppler Map is performed on the coprocessor using data directly sampled from the nadir antenna (Figure 7). In common with a standard GNSS receiver, the local PRN is generated onboard the co-processor. As an alternative to synchronizing and decoding the reflected signal in a standalone manner, the direct signals can be used to feed the navigation data sense, and assist the synchronization. The sampled data is multiplied by a replica carrier and fed into a matrix that performs an FFT on a row by row basis of the Delay Doppler Map, to achieve in effect a 7000 channel correlator, integrating over 1 millisecond. Each point is then accumulated incoherently over hundreds of milliseconds to bring the weak signals out of the noise.
Figure 7: Delay Doppler map processing (image credit: SSTL)
SGR-ReSI antennas: Although the SGR-ReSI can in principle support up to 4 dual frequency antennas, a reduced subset is being flown on TechDemoSat-1 to support its planned applications. A left hand circularly polarized dual frequency L1/L2 fixed phased array antenna (gain 13 dBiC) sits on the earth facing facet for GNSS reflectometry (Figure 1). It is the opposite polarization to conventional GNSS antennas and provides the higher gain required to receive the weak signals from GNSS reflections. A dual-frequency L1/L2 antenna and two additional L1 antennas will occupy the space facing facet with more typical RHCP and hemispherical patterns. These antennas are intended to provide navigation function for the satellite and also support monitoring of radio occultation events with both the L1 and L2 signals. The provision of three antennas on the space facing facet with suitable baselines between them also enables the SGR-ReSI to support GNSS based attitude determination as previously demonstrated by SSTL on the UoSAT-12 and TopSat satellites.
A new low noise amplifier has been designed that supports both L1 and L2 frequencies, is equipped with a temperature sensor, and a switched load to provide a known noise level when enabled for calibration purposes. One each will be used on the nadir and zenith dual frequency antennas respectively. The two L1-only zenith antennas will use the heritage integrated L1 LNAs (Low Noise Amplifiers) used with previous single frequency GNSS receivers.
Figure 8: Photo of the SGR-ReSI Flight Model (image credit: SSTL)
SSP (Sea State Payload) Altimeter:
Active measurements will be made by an experimental radar altimeter based on one phase center of the NovaSAR-S payload frontend. The altimeter is being built by SSTL/Astrium to: 15)
• Provide flight heritage for SAR hardware
• Demonstrate altimeter functionality
• Provide a stepping stone to a low cost altimeter instrument
• Measure background noise and spurious signals at S-band.
The SSP altimeter is an experimental instrument that will demonstrate altimeter functionality and measurements rather than full operational performance. It will also provide an early flight opportunity for the RF equipments of the NovaSAR-S payload. By leveraging the SAR development the altimeter design is simplified and a working system, although far from optimal, can be realized within the challenging TDS-1 schedule. 16)
By utilizing components from Astrium’s SAR (Synthetic Aperture Radar) to operate as a coarse altimeter, the SSP pulses radio waves onto the ocean. The echo waveforms that return give an independent measurement of the sea state and the information gathered can then be applied to meteorology, oceanography, climate science and ice monitoring. Astrium Portsmouth will also contribute an antenna design using the same technology as the SAR antenna but on a smaller scale. 17)
The PFM altimeter will be implemented using the PCU (Power Conditioning Unit), transmit and receive units built to near flight standard for the SAR payload ground demonstrator. An antenna will be built based on the SAR radiator unit but with a square rather than rectangular array of patches. The transit unit will be one of the first GaN amplifiers in orbit.
The altimeter design has been kept deliberately simple to minimize NRE. It transmits chirp modulated pulses and records the deramped echoes without any tracking or onboard processing. Collected data will be processed offline on the ground. The priority will be demonstrating measurement of SWH (Significant Wave Height). However, relative SSH (Sea Surface Height) and wind speed will also be possible. Absolute SSH measurements of limited accuracy will be possible if good orbit knowledge is available from the on board AOCS system, but this is not a priority. It will also be possible to configure the altimeter as a receiver to take measurements of background noise and characterize the in orbit spectrum at the S-band allocated frequencies.
Figure 9: Block diagram of the TDS-1 SSP altimeter (image credit: SSTL)
Space Environment Suite (MuREM, ChaPS, HMRM, LUCID)
MuREM (Micro Radiation Environment Monitor)
MuREM (or µREM) is designed to provide in-fight data on TID (Total Ionizing Dose), dose rate and the charge deposition spectra of SEE (Single Event Effect) initiating energetic protons and heavy ions, as well as providing radiation effects data on samples of advanced COTS (Commercial-off-the-Shelf) electronic devices. Its low mass (< 1 kg) and low power make it suitable in whole or part as a generic radiation environment monitor for routine fight on micro- or nanosatellites. 18)
The monitor consists of three main boards sized to the PC104 standard to make the board layout consistent with a CubeSat layout, for possible future re-flight. The unit as configured for TDS-1 uses a 10 cm x 10 cm x 4.5 cm housing with mounting feet and has a mass of < 1 kg.
Figure 10: Exploded view of MuREM with 3 board PC104 sized stack and mounting points (image credit: SSTL)
MuREM makes use of RADFET (Radiation-sensitive Field Effect Transistor) solid-state dosimeters, a dose rate sensitive photodiode and has two large area PIN diodes – one to measure proton flux and one to measure heavy ion flux and LET (Linear Energy Transfer). The payload additionally carries a radiation effects board used to gather collateral data on devices exposed to the true environment found in space.
Each RADFET dosimeter consists of a pair of MOSFETs with a large gate oxide layer, increasing their response to TID, which is detected bay measuring the change to the gate threshold voltage. In measurement mode, one of the pair is held biased, whilst the other is held unbiased. In read mode, a fixed small current or approximately 10 µA is passed through the MOSFET channel via the source and drain electrodes, and the required gate voltage is measured. The change in gate voltage varies linearly with accumulated dose and the rate of change of voltage with dose is bias dependent (more change for the biased FET). Unfortunately, the gate voltage is also strongly affected by temperature; however, by measuring the change in the biased and unbiased FET, the common temperature effect can be removed, leaving the dose-dependent change. A nearby temperature sensor provides supplementary temperature data, which can be used in case of failure of this detection methodology and the requirement to calibrate for dose.
One of the RADFET dosimeters on MuREM will have a thin Cu foil of 50 µm, and the other a Al shielding equivalent 165 µm to investigate the effects of using high-Z materials in spacecraft shielding. The output from the RADFET circuit feeds into the 10 bit C515C ADC (Analog Digital Converter). In order to provide high resolution data, an adjustable offset is applied to the circuit via a potentiometer voltage divider to maintain a voltage range compatible with the ADC.
The dark current of the photodiode used in the MuREM is due both to thermal and ionizing radiation effects. The photodiode has a linear thermal response, and so calibrating this out of the dark current is straightforward. The remaining dark current signal is then solely due to ionizing particles depositing energy into the detector and so monitoring this current provides a measure of the dose rate environment. This method was used on the CEDEX payload, flown on GIOVE-A with great success, and provided real time information on the ionizing dose rate.
The particle detectors consist of large area, 300 µm thick PIN diodes whose outputs are run through CR-RC pulse shapers. A peak hold circuit allows the peak of the resulting pulse to be read into the 10 bit C515C ADC accurately, providing a high confidence result. A 3 cm x 3 cm diode is used for detecting heavy ions and a 1 cm x 1 cm diode is used for protons, owing to their relative fluxes. Particles travelling through the diodes will deposit energy dependent on their species and kinetic energy. Each particle event is logged in 16 logarithmically spaced energy deposition channels covering a LET range of approximately 0.1 to 10 MeV cm2 mg-1 for the heavy ions. The proton detector has a different energy deposition threshold and range and is designed to detect protons with energies > 30 MeV, dependent on spacecraft shielding. Each detector chain is capable of measuring up to 50,000 events per second.
MuREM operates on a +5V regulated power bus, with internal +1.5V, +3.3V,-5V, ±10V and +20V lines generated internally. Outputs from all sensors are input into the 10 bit ADC of two internal C515C CAN controller devices. Two CAN bus addresses are used, with data from the particle detectors and radiation effects board delivered via CAN file-transfer protocol (FTP), with RADFET and dose rate diode data delivered as telemetry. The unit interfaces with the spacecraft via a single DA-26M connector containing all data and power links.
ChaPS (Charged Particle Spectrometer):
The objective of ChaPS is to demonstrate a novel payload design that combines the capabilities of multiple analyzers by using four miniaturized sensors to perform simultaneous electron-ion detection. Each of the sensors is optimized to carry out electrostatic analysis of the different space plasma populations expected in LEO (Low Earth Orbit). 19) 20)
The prototype ChaPS instrument is provided by UCL/MSSL (University College London/Mullard Space Science Laboratory). It is the first prototype of a new class of compact instruments to detect electrons and ions, building on 40 years of experience at UCL/MSSL.
Table 3: ChaPS detector specifications 21)
Figure 11: Illustration of a ChaPS detector (image credit: UCL/MSSL)
ChaPS has a form factor of a 1U CubeSat. The instrument will demonstrate the principles on-orbit and open the way to use the techniques on other missions where mass and power are at a premium, for example space weather constellations. ChaPS will operate in three modes, to measure electrons in the auroral regions, electrons and ions in other regions and also to measure the spacecraft potential.
ChaPS is extremely attractive because it saves mass, power and volume - and ultimately mission cost-while providing an enabling technology for future space missions such as ESA's proposed JUICE (Jupiter ICy moon Explorer) mission to Jupiter. Its low cost also opens up new applications for such instrumentation that were simply not feasible in the past.
HMRM (Highly Miniaturized Radiation Monitor):
The HMRM instrument is provided by RAL (Rutherford Appleton Laboratory) and ICL (Imperial College London). It is a lightweight, ultra compact radiation monitor designed to measure total radiation dose, particle flux rate and identify particle species (electrons, protons and ions). The instrument is designed to provide housekeeping data on the radiation environment to spacecraft operators to correlate the performance of spacecraft subsystems, raise alerts during periods of enhanced radiation flux and to assist in diagnosing spacecraft system malfunctions. 22) 23) 24)
The goal of the HMRM program is to develop a “chip sized” prototype radiation monitor suitable for application in:
- Coarse radiation housekeeping
- Alert and saving function
- Support to platform and payload systems
- The mass of HMRM device shall be less than 20 g
- The radiation detector single chip shall have a radiation TID tolerance of at least 100 kRad
- The radiation detector single chip shall be latch up free
- The sensor shall be capable of meeting the performance requirements under full solar illumination at an orbital radius of 1 AU.
Table 4: HMRM design summary
Figure 12: Block diagram of HMRM (image credit: RAL Space, ICL)
Figure 13: Photo of the HMRM structure (image credit: RAL Space, ICL)
A CMOS sensor for the HMRM was selected with the following main specifications: 25)
• 0.18 µm CMOS image sensor technology
• 20 µm 4T-pixels in a 50 x 51 array
• Snapshot and correlated double sampling (CDS)
• Frame rate up to 10,000 fps
• Column-parallel 3-bit single-ramp ADC, with in-column trimming
• Digital readout, plus analog readout for debugging
• Integrated DAC for voltage/current generation
• Band gap
• Temperature sensor.
LUCID (Langton Ultimate Cosmic ray Intensity Detector)
The LUCID instrument is provided by the Langston Star Centre (i.e., the Simon Langton Grammar School for Boys, in Canterbury, England). LUCID is the winning entry of a U.K. space competition. The LUCID device allows the characterization of energy, type, intensity and directionality of high energy particles. 26)
The device makes use of COTS sensor technology developed at CERN (The European Organization for Nuclear Research) using Timepix chips from the Medipix Collaboration. Part of a family of photon counting pixel detectors, Timepix allows for recording time information regarding when events occur relative to when the shutter opens.
The data obtained from LUCID is of interest to NASA in terms of radiation monitoring but also provides inspiration to the next generation of physicists and engineers by giving school students the opportunity to work alongside research scientists and take part in authentic research. 27)
Figure 14: Inside the LUCID payload (image credit: SSTL) 28)
The LUCID experiment is part of a wider project called CERN@school, a program that aims to bring the excitement of CERN into the classroom, and encourage the future generation of scientists. CERN@school covers the effort to get individual Timepix chips into schools for educational and research purposes. We’ve started a number of different activities, the most recent of which is Radiation Around You (RAY), a project to build up a radiation map of South East England. We’re into the final of the Rolls-Royce Science Prize with the RAY pilot program.
Air and Land Monitoring Suite:
CMS (Compact Modular Sounder)
The CMS instrument is being provided by Oxford University’s Planetary Group and RAL (Rutherford Appleton Laboratory). The CMS is a modular infrared remote sensing radiometer unit, designed to easily mix and match subsystems and fly multiple versions on multiple platforms at low cost by tailoring it to specific customer requirements once flight heritage has been proven.
- The CDHU (Command and Data Handling Unit), supplied by RAL, consists of an ARM processor and Actel FPGA, which performs real time spatial and temporal averaging of data.
- In standard operation mode the instrument views sequentially the nadir mapping view, space and a calibration target.
- During payload tests limb viewing and different integration times can be substituted.
- Base dimensions 380 mm x 315 mm x 186 mm
- 6 x M5 fixings to the top of the payload panel
- Mounted such that the unit has a FoV for Earth, space, and limb
- Mass of 4 kg
- 26 W male high density D-type (data)
- 25 W male standard D-type (power & CAN)
- Connected to 2 +28 V low power battery switches (electronics and payload heater)
- 4 Mbit/s synchronous serial data link
- Interfaced to HSDR and FMMU via separate LVDS chips for redundancy
• Command & control:
- Primary and redundant CAN architecture.
Platform Technology Suite:
This experiment is the realization of the stellar gyroscope concept on a CubeSat ADCS (Attitude Determination and Control Subsystem), designed and developed by SSBV SGS (Space & Ground Systems), Portsmouth, Hampshire, UK. The stellar gyroscope can be used to measure attitude changes from a known initial condition without drift while sufficient stars are common across frames, because absolute attitude changes are measured and not angular rates. 29) 30)
To maintain a high quality attitude estimate in eclipse, two current alternatives are employed.
• A star tracker/mapper can be used to identify star constellations and retrieve absolute attitude. However, star trackers add cost and complexity requiring a star database, high update rates, and consequently high quality optics, sensors and a baffle.
• The second method is to use an Earth horizon sensor. For the sensor to work in eclipse, the infrared spectrum corresponding to the H2O absorption bands is used. This typically requires a specialized IR sensor, a detector cooling system, or a chopping or rotation mechanism to generate differential readings of the Earth and space temperatures. Such a system requires significant power and the mechanical systems have reliability concerns.
The alternative, used in this design, is to use a stellar gyroscope, as a relative attitude sensor, to reset the drift from the rate gyroscope propagation in eclipse. The stellar gyroscope can be realized using low cost sensor and optics, where the algorithms can tolerate a large amount of noise, and does not require a star database.
The stellar gyroscope can be used to propagate a spacecraft’s attitude from a known initial condition without drift. Normally, in the absence of an absolute attitude measurement, attitude is propagated by integrating gyroscope angular rate data (typically MEMS based for small satellites). This results in a drift in the attitude estimate, which is essentially a loss of attitude knowledge after a sufficient amount of time. The image based approach can propagate attitude without drift while sufficient stars are common across frames. As the camera pans the sky, after sufficient time all the stars may leave the frame. In that case, some error accumulates as rotation estimates are stacked. However, this happens over a significantly longer period of time compared to a MEMS rate integrator. The image-based rotation estimates can complement a set of MEMS rate gyroscopes to maintain a high accuracy attitude estimate at low angular rates (where MEMS gyroscope drift is most severe).
For the stellar gyroscope, the star correspondence problem across frames is challenging due to spurious false-star detections (false-positives) and missed stars (false-negatives). Correspondence of stars across frames can be done with limited success by proximity for small angular changes, where for short time intervals, the stars are assumed to not have moved much. However, for large attitude changes, the star association algorithm between frames must overcome false stars, missed stars, stars leaving the field of view and new stars entering the field of view. The problem is essentially to fit a mathematical model over data with a large number of outliers, for which the RANSAC (Random Sample Consensus) approach is effective. RANSAC is a popular algorithm in machine vision and stereo vision, and has been proposed for satellite based image registration for geographic applications. 31)
CubeSat ADCS: The attitude determination and control subsystem is designed for CubeSats on a standard PC104 board. In its basic configuration, it integrates a high sensitivity magnetometer, up to 6 sun sensors, 3 axis MEMS gyroscopes, and 3 magnetic torque rods as a 3-axis magnetic attitude control system. In its full configuration for improved attitude knowledge and pointing accuracy, a GPS receiver, a stellar gyroscope and an ADCS control computer are added on a daughter board, still within the PC104 height constraints. A momentum wheel or three reaction wheels can be added from a third party supplier.
The stellar gyroscope complements the MEMS rate gyroscopes in eclipse to maintain an accurate estimate of attitude. However, in order to benefit from accurate propagation in eclipse, accurate knowledge in sunlight is necessary. The system utilizes sun sensors accurate within 0.5º developed by SSBV, as well as a high-accuracy magnetometer that produces magnetic field vector measurements to around 1º of accuracy in combination with an IGRF magnetic model and good knowledge of the position in orbit, which is provided by the GPS receiver. This combination results in a high quality estimate of attitude in sunlight.
Figure 15: Block diagram of the CubeSat ADCS (image credit: SSBV)
The stellar gyroscope on this ADCS system consists of a low-cost camera assembly and processing hardware and is designed to require little mass and volume. The camera is based on the OmniVision OV7725 VGA CMOS sensor and a miniature S-mount lens with a focal length of 6 mm. This configuration results in a 27.6º x 36.7º FOV. Table 5 summarizes the camera specifications. As it is only to be used in eclipse in this application it is not required to operate in sun light and does not need a baffle.
The camera is designed to register stars of magnitude 4 and brighter. With the selected optics, FOV, and an exposure time of 800 ms, at least 4 stars are visible in 97% of the sky, and at least 3 stars are visible in 99% of the sky.
Table 5: Summary of stellar gyroscope CubeSat hardware
Figure 16 shows the camera system and ADCS system as designed for the technology demonstration experiment on TechDemoSat-1. The experiment will take sample images and log other sensor data to tune and validate the attitude determination algorithms.
Figure 16: Camera assembly (left) and SSBV CubeSat ADCS experiment (right) on TechDemoSat-1 (image credit: SSBV)
In future developments, in order to tolerate higher slew rates and improve image quality, SSBN will replace the optics to be able to significantly reduce the exposure time by increasing the aperture. Preliminary analysis shows that slew rates up to 3º/s are feasible with the improved configuration. At higher slew rates, the system uses the MEMS rate gyroscopes as they become more reliable.
DOS (De-Orbit Sail)
DOS, developed at Cranfield University, is intended to demonstrate a novel means for de-orbiting a satellite at the end of its mission lifetime through deploying a sail to increase the aerodynamic drag.
The de-orbit sail is the product of several years of Cranfield University’s work on sustainable approaches to space exploration. SSTL’s TechDemoSat-1 gave the Cranfield team the unique opportunity to take-on the challenge of evolving their ideas from designs on paper, to flight-ready hardware. Maintaining a low mass is always a challenge with space projects; the TechDemoSat-1 de-orbit sail is made from a material called Kapton, which is just 25 µm thick. 32)
The sail will be deployed when TechDemoSat-1 issues a command at the end of its mission. This command will trigger cable cutters to be fired, which will release a restraining belt, and the sail will then be deployed by stored spring energy. Cranfield’s payload will then take up to 25 years to safely guide the TechDemoSat-1 spacecraft into the Earth’s atmosphere to disintegrate.
Exactly how long Cranfield’s de-orbit sail will take to complete the satellite’s course back into our atmosphere will be a subject of great importance to those interested in the management of space debris and the continued exploration and utilization of space. This process will be affected by a range of things including the amount of solar activity that it is being exposed to, and what altitude the satellite is at when the sail is deployed.
Figure 17: De-orbit sail from Cranfield University, deployed on TechDemoSat-1 (image credit: SSTL)
HCT (Hollow Cathode Thruster) demonstration module
HCT research and development in SSC (Surrey Space Centre) is focused on propulsion for small, limited on-board resources spacecraft in LEO (Low Earth Orbit). In that scope, several laboratory prototypes have been built and tested, to optimize their performance as propulsive devices. The optimization aims to increase specific impulse at low power levels, reduce heating requirements and increase the robustness of the device for use in flight opportunities. 33) 34) 35)
The HCT operation is based on field enhanced thermionic emission. A low workfunction insert is heated by an external heater to ~1100ºC. Xenon propellant, supplied to the HCT, is ionized by electrons emitted from the cathode and a plasma is formed internally in the HCT between the insert (ground) and an anode electrode (keeper). The anode electrode is kept at positive voltage and current is drawn from the insert via the plasma formed. The discharge to the keeper electrode is self-sustained once initiated, i.e. no heater power is needed. Direct thrust measurements of the HCT using the SSC thrust and facilities have shown a specific impulse of 80-85 s at a flow rate of 20 cm3/s xenon. The HCT can provide increased specific impulse capability for small satellites with the trade offs being a higher power requirement and a lower thrust than a resistojet. The HCT requires 50-60 W during heating and 40-60 W during discharge at nominal conditions.
Figure 18: HCT FM and QM during integration for TDS-1 (image credit: SSC, SSTL)
A complete HCT module has been developed by the SSC and SSTL flight on TechDemoSat-1 (TDS-1) as an experimental enhancement to the spacecraft platform. The TDS-1 is UK Space Agency funded, and the SSTL built spacecraft is based on the SSTL-150 platform. The spacecraft has a propulsion module with a total mass of 1.5 kg of xenon to be shared between the a standard SSTL resistojet and the HCT.
The HCT module comprises of the HCT (Hollow Cathode Thruster), FCU (Flow Control Unit), and the PPU (Power Processing Unit), developed to require minimal changes to the platform. The HCT flight and qualification models are shown in Figure18, the PPU is shown in Figure 19. The HCT module has been integrated and tested at spacecraft level.
Figure 19: The PPU during testing for TDS-1 (image credit: SSC, SSTL)
The flight operations of the HCT will be invaluable for the future joint EP (Electric Propulsion) developments in SSC as it is the first plasma thruster to operate on an SSTL platform as a 'bolt-on' module to the flight proven xenon feed systems that SSTL typically uses for the resistojet.
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.