Advances in Small Satellite Technologies

In the telemetry system, multiple types of sensors are used for monitoring the physical and structural parameters of a system. These parameters are helpful for understanding the behavior of the system and to improve the performance of the system. Telemetry system integrates all the signal conditioning circuits for different types of sensors and provides digital data in the form of serial interface. To provide a compact and energy-efficient solution for telemetry, a 44 channels multi-sensors interface is implemented in single CMOS IC. In telemetry system, different types of sensors have been used which provide an output voltage in range of micro volt to 10 V. Each sensor has different noise, bandwidth and filtering requirement. This sensor interface chip (SIC) provides an integrated solution for different types of sensors used in the telemetry system. SIC used advanced circuit techniques to achieve better sensitivity and accuracy and also used the combination of high-voltage and low-voltage breakdown transistors for high level of integration. SIC has been designed and fabricated using 180 nm CMOS technology. Die area for this chip is 8 mm x 8 mm, and the package area is 20 mm × 20 mm. This chip has been tested for 10-bit accuracy of all sensor channel interface. Total power consumption of this chip is 800 mW.

Vibratory gyroscopes with piezoelectric transducers have found substantial number of applications where low drift rate is easily achievable with low cost, low weight and high reliability. However, piezoelectric transducers are prone to aging effects that cannot be compensated and hence affect gyro drift and scale factor stability. In this paper, two methods of using same piezoelectric transducer element for actuation and sensing are designed and evaluated for improving gyro performance.

The computation of optimal route for various kinds of automatic or manned vehicles on hilly terrain is an important task in route planning applications. In the absence of a road network, the topography factors of the terrain (slope, elevation, etc.) and the climbing angle of the vehicle play an important role in the computation of the optimal route between two points. This kind of problem has been addressed in the artificial intelligence domain, and the graph search algorithms can be applied to find a solution. The A* algorithm and various versions of A* have been reported in the literature to achieve the faster results while maintaining the optimality criteria of the solution. The speed-up in all the versions of A* algorithms is achieved through the use of heuristic knowledge available in the problem domain. There may exist various heuristic functions that fulfil the admissibility criteria but produce different results as far as the speed and the optimality are concerned. This paper presents a performance measure for quantitative comparison of the various heuristics functions that can be used for optimal route selection in hilly terrain.

This paper presents the simulation studies of the dynamics of short tethered satellite system during the deployment phase. The deployment dynamics of the tethered system is dominated by the tether oscillations and the three-dimensional attitude dynamics of the sub-satellite. The dynamic characters of the system are investigated, and probable control strategies that can be adopted for the faithful and hassle-free deployment of the tether are discussed. Tether is proposed to be released with an initial velocity with constant deceleration using tension in tether. Dynamics of a short tether of 10 m length connected to a 10 kg 3U sub-satellite is considered in the study. Deployment with tether in tension and without tension in tether is discussed.

Chandrayaan-2 lander propulsion system employs a bipropellant thruster with a throttleable flow control valve (TCV) to vary the thrust produced by 800 N engines. A latching solenoid valve was proposed downstream of TCV to precisely admit/cut-off propellant supply to the engine as the latter does not inherently possess stringent leak tightness characteristics. The proposed valve employs a flexure guided plunger assembly to facilitate both continuous and pulse mode operation of the engine while preventing the possibility of cold welding/stiction between the sliding parts under space environment. A latching type of solenoid configuration was selected with balanced loading and permanent magnets to complement the solenoid in reducing the mass when compared to a conventional solenoid valve for the same application. In the selected configuration, switching of the valve position is accomplished by energising the solenoid coil and latching in the commanded position with the help of permanent magnets. Thus, the solenoid valve does not require continuous external supply, which results in negligible coil heating and very low power consumption. Two sets of proto models were realised which had undergone performance tests, 10,000 cycles of operation and 100 s duration hot test satisfactorily.

NIUSAT was developed by the students of Noorul Islam Centre for Higher Education under the guidance of various faculties of our college and support from the ISRO experts. One of the most important problems that a satellite faces after being separated from the launcher is the tumbling. When the satellite has relatively high angular velocity, it is said to be tumbling. Tumbling occurs right after the release of the satellite from the launcher and also during the deployment of the nano-satellite from the launcher. The first task of attitude control system is to de-tumble the satellite. We have to bring down the angular velocities before starting to stabilize the satellite because the stabilization controllers may not have the capacity stabilize the satellite with high initial velocities. So taking considerations of these things, we have to design the de-tumbling controller in such a way that it takes reasonable measures to effectively reduce the tumbling of the satellite in a reasonable time. In this paper, we will describe the design details of the de-tumbling controller, digital simulation results and the on-orbit performance on board the NIUSAT nano-satellite.

Satellite design involves the development of subsystems like telecommand, telemetry, attitude control, thermal control, structure, and payload. Testing the satellite on ground and qualification is an essential aspect to be considered while designing a satellite. A test system should be built to test and qualify all subsystems and their integrated behavior on ground. RSAT is a small satellite being developed at PES University with automatic identification system (AIS) as a payload. This paper presents the design aspects of such a test system for testing the attitude control system (ACS), which is flexible in terms of adding or removing any sensor or actuator. This paper is an attempt to present details on how the test system is made configurable for manual testing, automated testing, open-loop bench test, closed-loop attitude control testing and hardware-in-the-loop tests. The test system has been built as a generic tool making the system configurable for any satellite with ease.

Design of flight control systems for an aerospace vehicle is a significant challenge to the control designer. Since the aerospace vehicle is highly nonlinear system, the linear control design may not be effective as well as does lead to large operational domains. Once the number of operating points increases, gain scheduling process in linear control technique becomes cumbersome. To overcome the demerits of linear control design techniques, the nonlinear control strategy techniques can be applied. This paper first discusses mathematical model of the aerospace vehicle. In the present work, the nonlinear control methods are designed to achieve effective performance of an aerospace vehicle without linearizing the plant dynamics. The vehicle dynamics are separated using timescale separation principle, and control synthesis is designed using nonlinear dynamic inversion (NDI) methodology as well as optimal dynamic inversion (ODI). Nonlinear dynamic inversion is the control synthesis technique in which the inherent dynamics of a dynamical system are cancelled out and replaced by desired performance parameters which are selected by the designer. Two-loop control structure is adopted based on vehicle dynamics. In this approach, the outer loop control law generates yaw rate, pitch rate and roll rate commands for the inner loop, based on the side slip angle, angle of attack and bank angle commands. These body rate commands are further tracked by inner loop control law by generating necessary control surface deflections. Since the number of control surfaces is more than the tracking objectives, the inner loop control law is designed by using optimal dynamic inversion (ODI) technique. Nonlinear dynamic inversion and optimal dynamic inversion methods are popular in the application of control design. The detailed generic formulation for the NDI and ODI is discussed. Different perturbation studies on aerodynamic parameters are completed, and robustness of the design is validated. The second-order actuator model for the aerodynamic control surface (fins) is considered in this design. The nominal control law is designed and compared the simulation results for a step input of guidance profiles. Later nominal control design results are compared with respect to off-nominal cases like upper, lower bounds. All the design specification parameters show the satisfactory performance, and design is validated using different guidance profiles. To study the robustness of the control design against parameter uncertainty or model inaccuracies, the design is validated using different guidance profiles.

This document describes the implementation of a state estimator, and an attitude controller for Advitiy, IIT Bombay’s 2nd Student Satellite. The algorithms used by the estimator and the controller are described. A closed-loop simulation (CLS) framework for the testing of these algorithms is explained. A detailed description of each of the blocks in CLS framework has been provided. The problems faced due to actuation with magnetorquer are described. It also contains the description of practices followed to ensure the quality (correctness of results obtained and repeatability of the process).

The continuous communication link to the ground station is the prime requirement of any space mission. Presently, the LEO is primarily used for remote sensing, science, and human in space program. The visibility of the low Earth orbit (LEO) to the ground station is limited for a maximum of ~20 min, with four to five orbital visibilities over a day. This requires a huge network of ground stations. About two-third coverage of the orbit of LEO satellite is possible, with about 40 stations located geographically over different places on the globe. The most of the Earth’s surface is covered by water; a majority of ground stations need to be located on ships resulting in exorbitant costs of the ground network. Data relay satellite system with three satellite geosynchronous orbit (GEO) presents a very elegant solution in this context and provides a stable link between satellite in LEO and ground station. In the global scenario, the evolution of Tracking and Data Relay Satellite System (TDRSS) is developed by NASA to support communications and tracking services for its LEO spacecraft. The modern space technologies with low-cost microelectronics and microelectromechanical systems (MEMS) technology have enabled building small satellites in a short time. They have the advantage of low mass, low power consumption, low cost, and quick accessibility to space with launch-on-demand services. Small satellites have successfully demonstrated ground satellite and inter-satellite laser communication, which transmits data bit rates orders of magnitudes greater than the radio communication. The efforts are being made to design onboard controller as a generic 6-DOF control system capable of autonomous orientation, orbit, and formation control of multiple satellites in orbit. We use the small data relay satellites at GEO to keep in contact with LEO satellite for continuous data transfer coverage and better navigation accuracy.

Describes the mission, control and attitude determination meant for spin stabilization to three-axis stabilization by using momentum bias system. Advantage of the proposed minimal configuration is better for small satellite using off-the-shelf component. There is clear advantage to spin up and spin down of the wheel in this configuration. Attitude determination is now simpler and reliable for the payload pointing to earth or solar panel tracking the sun when the pitch is held to orbit normal.

Attitude maneuvers are required for reorienting the satellite for meeting various mission objectives. A higher satellite slew rate during maneuvers is a desirable characteristic for any mission design. At the same time, a higher slew rate demands a higher actuator torque capacity which is always constrained due to the physical limits of the actuator. This paper discusses the satellite attitude maneuver problem under slew rate and actuator capacity constraints. A quaternion feedback controller with quaternion and angular rate feedback is implemented to achieve a rest to rest reorientation maneuver. The control gains are computed based on the slew rate limits while the control commands incorporate the actuator limits. Three different configurations of four reaction wheels, (a) pyramid configuration, (b) skew configuration with three wheels along the body axes and the fourth wheel along body diagonal and (c) skew configuration with skew angle of 45°, are studied. The paper further discusses single reaction wheel failure conditions for all the three configurations and demonstrates satellite rest to rest maneuver within the slew rate and actuator limits.

In recent years, dual polarimetry has gained much importance among other earth observation missions due to its less complex architecture, larger swath coverage and low battery consumption. In this paper, a novel technique for automatic initial seed point selection for dual-pol data is proposed and its performance for different dual-pol combinations has been assessed. The proposed technique is independent of polarization combinations. Impact on classification when only 50% pixel is considered for selecting initial seed point is demonstrated. RISAT-1 quad-polarimetric data has been used for validating our result.

The purpose of this paper is to discuss the criteria to select the orbit for satellite with SAR payload which is best suitable for the mission. The requirements of the SAR payload are spelled out, and in turn how we convert those requirements to design of orbit is discussed. The constraints of SAR payload imaging over normal electro-optical payload will be detailed. Mission requirements and payload constraints are converted into specific orbital parameters. SAR payload geometry is discussed, and off-nadir angle, repeat cycle and revisit time relation is detailed. The superiority of the selected orbit is simulated and detailed in the paper. Global coverage analysis studies are simulated and presented.

This paper presents a study of a deterministic camouflaged target detection using hyperspectral images. Hyperspectral imagery has hundreds of spectral bands. This plethora of spectral information enables the detection of targets concealed by tree canopy of camouflage net. This problem cannot be addressed using high-resolution visible data alone because ‘shape information’ provided by such imagery cannot be used reliably in case of concealed targets. Hence, this problem is modeled as a target detection problem using ‘spectral information’ present in hyperspectral imagery. Moreover, the presence of concealing material such as canopy or cam net over the target necessitates the use of ‘sub-pixel’ target detectors.

An exciting part of the astrophysical spectrum is the far-ultraviolet (FUV: 90 to 180 nm), with a great density of absorption and emission lines. We are planning to develop an imaging spectrograph incorporated into a 6U (300 × 200 × 100 mm3) satellite to observe di use nebulae. The field of view is 4 × 1′ with a spatial resolution of 13 and a spectral resolution of 0.6 Å (Resolution 2000). In this work, we present the selection of the components and the optical design of the spectrograph using hand calculations and Zemax software. To determine the positions of the components, ray tracing (sequential) is done.

In this paper, design of electronics for VNIR hyperspectral payload for the small satellite is presented. The payload comprises of imaging unit, linear variable filters (LVF), a CMOS sensor and electronics module. The LVF is used as a spectral filter before the sensor. The rectangular sensor being an array detector captures images in different bands in pushbroom fashion utilising satellite motion along the track. VNIR electronics captures sensor data, processes, rearranges and writes in memory to form hyperspectral data cube (H-cube). This high-speed data is transferred in real time to bus data handler (BDH) of satellite bus. This paper covers challenges and important aspects required for designing such payload electronics.

Military applications such as reconnaissance, surveillance and detection of camouflaged targets require imaging payloads with state-of-the-art sensors in spectral wavebands covering visible NIR (VNIR), SWIR, and thermal IR (MWIR and LWIR) region. Hyperspectral imaging sensor is one of them. This type of sensor acquires both the spatial and spectral signatures of the scene simultaneously. These sensors can be designed using variable filter, prism, grating, or interferometric means. This paper presents the design of a hyperspectral imaging camera (HSIC) for the VNIR (450–900 nm) and SWIR (900–2500 nm) wavebands using linear variable filter (LVF). This camera has advanced features such as high spatial resolution (12 m at 500 km) and high spectral resolution (10–20 nm) and can be used for military applications from small satellites operating in LEO.

The objective of this paper is to design and analyse the opto-mechanical design of the SWIR camera for space applications. The opto-mechanical design includes, mounting of various optical elements using compliant flexures so that the optical elements are isolated from mechanical and thermal loads. To mitigate the effects of thermal cycles in space environment, special structural materials are used. Joining of the opto-mechanical is done using space grade adhesives to avoid various issues with the friction-based joining schemes. Opto-mechanical design approach is focused on ease of manufacturing, optical performance and survival in space environment. Finite element model of the entire assembly was established to analyse the design for static and dynamic design requirements.

Presently, small satellites with hyperspectral imaging payloads have been used for target detection, environmental monitoring, and forest mapping. All these applications involve the same process of recording the spectra of target/object of interest over densely sampled contiguous spectrum, usually from 400 to 2500 nm. Once a target spectra is recorded, it is matched with reference spectra library for detection. The accuracy of spectra matching highly depends upon the accuracy of radiometric calibration of the hyperspectral imaging payload. This paper presents a study of three methods of absolute radiometric calibration used by various hyperspectral imaging systems in VNIR region, namely Hihara method, Mitsunaga method, and Xiujuan method. Out of the analyzed methods, Hihara method performs the best.

The paper is aimed at obtaining the optimum mount configuration for a Zerodur primary mirror for a Cassegrain spaceborne electro-optical payload. A bipod flexure design procedure is presented while taking various optomechanical design principles into account. Finite element model of the flexure is established in ANSYS APDL to analyse the design w.r.t optical and mechanical design parameters. In the design process of the mounts, thickness of the flexure notches; angle between the bipods; and height of the mounts are optimized for the best optical performance.

This paper talks about the weight optimization scheme of the primary mirror of a space telescope. Comparative study of three lightweight structural schemes (hexagonal, triangular and square) for the primary mirror is done. For the study, finite element modelling method is used using as tool. While comparing the weight reduction, the comparison between the three schemes is carried out based on static deflection, surface figure of the mirror and fundamental modal frequency.

Stabilized electro-optical system (SEOS) housed with electro-optical (EO) payloads is required for battlefield surveillance, positioning, pointing and tracking of ground, aerial and naval targets. SEOS are required for line-of-sight (LOS) stabilization by rejecting the host platform motion for accurate pointing and tracking. In such complex systems, the design of real-time embedded software becomes very challenging. Hard real-time implementation of stabilization and position and track control algorithms make it more critical. Interfacing and real-time handling of heterogeneous interfaces simultaneously and further make the embedded software complex. In this paper, we present an innovative real-time embedded software design and implementation for stabilized EO system.

Recent solar flare explorations reveal the presence of high-frequency components in sub-THz and THz regime, which are much more intense, compared to the RF counterpart, and understanding of origin is still ambiguous. Solar flare observations in THz region need to be carried out from space since the atmospheric attenuation hinders measurements from ground. Present paper discusses, in detail, the development of a photometer that operates in THz region from 0.1 to 7 THz in different spectral bands. Entire electro-opto-mechanical design was carried to meet the size, mass, and power constraints posed by inherent resource limitations of Nanosatellite platforms. A Cassegrain optical configuration was designed to collect the sun radiation with aperture effective area of 0.0027 m2 with HDPE windows and resonant mesh filter for selective perception of THz radiation, rejecting the huge background of visible and NIR with a spectral selectivity of 107:1. Sensitivity of the photometer was estimated to be 9 K/mV by calibration against the blackbody radiator. The minimum detectable temperature transient of the instrument is 30.6 SFU (1 SFU = 1×10−22 Wm−2 Hz−1).

Defense satellites with hyperspectral imaging (HSI) and synthetic-aperture radar (SAR) payloads capture high-resolution images. The voluminous data captured by payload needs to be processed and stored for transmission to ground station upon visibility for analysis. Data from payloads has to be acquired by data handling system and to be stored in memory after required processing. This paper describes novel and efficient hardware architecture of high speed, high capacity, miniaturized, non-volatile, redundant and radiation tolerant integrated base-band data handling (BDH)—solid-state recorder (SSR) system for polar lower earth observation (LEO) small satellite applications.

This paper presents a novel space-based automatic identification system (AIS) receiver architecture meant for maritime surveillance. The receiver performance has been simulated in the presence of three AIS Tx signal messages colliding in different combinations of frequency, time and power. Doppler shift due to the LEO satellite moving at a velocity of 7.5 km/s at an altitude of 600–900 km, and the delay in signal transmission was also induced in the AIS Tx signals to fully characterize the AIS receiver. The architecture proposed in this paper is suited for on-board processing on any nanosatellite with enhanced message and ship detection capability. The performance parameters of the proposed AIS receiver in the collision environment have been tabulated in the results’ section.

This paper aims at the overview in design and development of reliable and robust space grade power supplies for satellite subsystems. The satellite is using six types of power supplies modules (PSM) which provide total of 20 isolated outputs. All converters are designed by using forward converter topology followed by post-Mag-Amp regulator for each output to meet the required specifications. Voltage mode with feed-forward control technique has been adopted to provide better CS101 performance. The converters deliver the output power from 15 to 110 W for different PSMs for the lower input voltage range of 32–42 V at a switching frequency of 150 kHz with sync. This paper also covers specification, design procedure and practical test results. Flight models have been developed, and practical electrical results have been validated in environmental tests like thermo-vacuum cycling test, operational temperature test, vibration test, etc.

Grounding is very essential part of any electrical and electronic system design. Spacecraft is a complex system consisting of multiple complex subsystems electrically connected to each other. Most operations in satellites take place autonomously based on measurements; improper measurements may lead to unintended operation. Spurious triggering of switches and relays due to EMI is also very dangerous in satellites as it consists pyros and propulsion. Spacecraft charging will cause arcing in satellite which may lead to failure of subsystems. These situations sometimes lead to loss of whole satellite and make it irretrievable. Proper grounding is very essential to avoid these kinds of problems. The associated costs of any failure of on orbit satellite are very huge; hence, minute care shall be taken from the design phase to avoid any such problems. The objective of grounding is to ensure the proper operation of the spacecraft and reduction in electromagnetic interference. A good grounding design critically looks into these aspects and tries to minimize such occurrences and henceforth helps in reducing the noise coupling. This paper briefly presents the spacecraft grounding requirements, design approaches, grounding scheme philosophy at satellite level and discusses the different types of grounding schemes in small satellites. This paper also deals with advantages and disadvantages of different grounding schemes in small satellites.

Evaluation studies were carried out to assess the suitability of high energy and high-power 75 Ah lithium polymer cells for space and satellite applications. The study includes evaluation of cell nominal capacity, Ah efficiency at different discharge rates, loss of capacity on storage and performance of the cell at different discharge rates. It also explains about the fast chargeability, self-discharge characteristic of the cell and effect of temperature on the performance of the cells. The study has been carried out at different temperatures ranging −30 to 60 °C at different C rates (0.5C to 6C). The cells were charged at constant current of 0.5C until the on-charge voltage reaches 4.2 V, then soaked at different temperatures from −30 to 60 °C for 16 h and discharged at different currents varying 0.5C to 6C. The above experiment was repeated by charging at different currents and discharging at 1C at 23 °C. The self-discharge characteristics were analysed by charging at constant current of 0.5C then stored for the period of 1 day, 7 days, 14 days and 28 days and discharged at 0.5C at 23 °C. The suitability of low temperature operation for satellite and space application was discussed and the results were presented.

Spacecraft power systems reliability is critical parameter for mission success. Multiple checks and inspections are carried out for each component for power subsystems. In this paper, a novel automated solar cell micro-crack inspection tool is presented which is based on convolutional neural network (CNNs) to classify space-grade multi-junction solar cells taken under electroluminescence condition. The whole system is named ELSIS, which stands for “Electroluminescence Smart Inspection System”. It is an end-to-end automated system that acquires images under electroluminescence condition as arrays, identifies each cell and classifies them into two classes namely the cells that exhibit micro-cracks and cells and those that are free of micro-crack. ELSIS is developed with the objective to reduce the tedious and time-consuming manual effort required to identify micro-cracks and to increase the reliability of the solar cell modules by minimizing the human errors which may arise during the manual inspection of thousands of solar cells for spacecraft by automating the whole process. The CNNs for ELSIS have been developed in TensorFlow framework in Python based on InceptionV3 architecture. ELSIS is augmented with the latest image processing techniques that are applied to acquire EL images, which are modules containing array of solar cells (Fig. 1). ELSIS thus identifies and produces individual solar cells from the arrays and indexed and stored. The deep learning network was trained on a large number of solar cell >6000 images such that cross-entropy of the network settles within an accepted constant value. The trained network when tested for a large sample size of ~3000 new solar cells yielded a very reliable >98% accuracy. ELSIS helps in capacity building for scale manufacturing of satellites with stringent power budget, making it imperative to have zero defect solar cells.

Microsat is the third satellite in the series of remote sensing micro-satellites by ISRO. It is a low earth orbit (LEO) spacecraft, launched into a 350-km circular orbit on 12 January 2018. The orbit of 350 km poses lot of constraints on the satellite with respect to space environment as well as atmospheric drag. The major concern is from atomic oxygen present in the LEO environment. The interaction between atomic oxygen and spacecraft surfaces produces significant changes in the mass, surface morphology, electrical and optical properties of the spacecraft materials through erosion and oxidation. The major exposed materials of solar panels, viz. silver, Kapton and epoxy, are vulnerable to atomic oxygen effect. For LEO missions (<500 km), erosion rates of solar array materials can be high and can cause degradation in their mechanical, thermal and electrical properties. In recent years, many nanosatellites and student satellites are launched into low earth orbits for technology demonstration. Hence, understanding of the performance degradation of solar array exposed materials is very critical for the low-altitude satellites. To protect the exposed surfaces, mitigation technique needs to be adopted. In this chapter, the effect of atomic oxygen on the solar array exposed materials, details of the protection coatings, its qualification and on-board performance are presented.

Recently, the occupancy of small satellite in space is increasing, and one potential application of this is satellite imaging. The difference in acquisition conditions of satellite images makes it susceptible to much degradation compared to normal images. The common degradations are motion blur, low resolution and noise due to low light conditions, rain, snow, etc. Since small satellites are planned to be launched in big numbers, the number of image data to be analyzed will be higher, and this calls for some automatic techniques for assessing the quality of acquired images. Even though perceptual image quality assessment is a much explored area, there is not much work done in the area of satellite image quality assessment. In this paper, we review different techniques used for satellite image quality assessment and suggest some techniques which can also be used. The first method that we propose extracts GIST feature from satellite images and then uses a shallow neural network to predict the quality of the image from these features. The second method uses a deep convolutional neural network (CNN) with a shallow fully connected neural network to predict satellite image quality.

Many challenging space application requirements can be met using COTS. The time comes (better sooner than later) for the policymakers to reach a consensus on applying the COTS philosophy to space applications. This policy change is critical for the space industry. This paper attempts to suggest different issues to be tackled in order to meet technical, quality assurance, and cost requirements using COTS. A less heavy package results in a smaller payload for the same board dimension; this provides a huge weight advantage for space missions because the payload mass defines the launch vehicle requirements. Potential reduction in cost can be achieved by using various methods like screening, accelerated testing, or partial qualification techniques to complement the existing commercial qualification as well as by the reduced package materials costs. Assessment of the risk associated with potentially lower reliability devices, engineers within the commercial and aerospace industries are using trade-off and risk analysis to aid in reducing satellite system cost while increasing performance and maintaining high reliability.

Different characteristics of two or more images can be combined to achieve the merits of both in the resultant image. In proposed work image fusion is performed on pair of images, LISS-III and AWiFS of Resourcesat-I satellite (ISRO) LISS-III sensor data have high-spatial resolution, whereas AWiFS sensor data has high temporal resolution and larger swath. Different techniques namely regression, support vector machines (SVM), non-subsampled contourlet transform (NSCT) are used for acquiring images with high temporal, finer spatial and larger swath. The results of the resultant images with high resolutions are found to be satisfactory.

In Chandrayaan I Mission and Mars Orbiter Mission, ISRO put the spacecraft from Earth Parking Orbit to the trans-lunar/trans-planetary injection point through a series of orbit-raising manoeuvres around Earth. The start time and duration of each burn, along with the size of the intermediate orbits and the burn sequence, must be designed, such that at the end of manoeuvres, the trans-lunar/trans-planetary injection epoch is achieved. This study aims to design the optimal multiple finite burn orbit-raising manoeuvres for trans-lunar and trans-planetary manoeuvres, using differential evolution as the optimization technique. The reduction in the computational time is achieved through multi-threading.

A multichannel digital receiver is required for estimating several signal parameters such as instantaneous frequency, pulse width, pulse amplitude, PRF as well as direction of arrival of signals. In this paper, a five-channel signal conditioning module is developed for the multichannel digital receiver that is capable of operating in 160–180 MHz and 750–1250 MHz bands with 40 dB dynamic range at the input (−38 to 2 dBm).

Small satellite constellations are becoming space systems capable of operational space application products delivery and services. As the number of satellites are becoming large in each constellation of satellites, many of the operations of the satellite need to be smart with intelligence or expert systems. To achieve this, technology of the satellite sub-systems need to be improved for enabling such operations. One of the crucial sub-systems for data transfer is the communication sub-system and antenna design. With constellations, the inter-satellite links become very essential for exchanging health keeping, command, control, scheduling and mission data so that wherever the satellite is in the globe, for crucial link establishes with ground at any time. Beam switching with phased array concept, though used in many missions, a continuous development to make it small, light weight and more efficient, is a continuous effort.

Missile early warning system enables the defence services to detect the missile launches. IR imaging sensors on satellite platform can detect vehicle plumes from the launcher, and it can send the data to the ground station. Orbit design and constellation study analysis of missile early warning system is carried out to get the global coverage. The constellation consists of three or four GEO stationary satellites, four highly elliptical orbit (HEO) satellite orbits in two planes to cover higher latitudes on earth and low earth orbit satellites to cover the specific area of interest to track the launcher in middle stage. Number of satellites in the constellation estimated and optimized to cover area of interest. GEO satellite, HEO satellite coverage analysis and LEO satellite orbit parameters analysed. The complete scenario of missile defence scheme is presented.

Continuous coverage of a fixed location on Earth with the constellation of LEO satellite is considered. Target’s relative velocity on Earth with respect to the satellite is analyzed for different orbit altitudes and inclinations. Characteristics of footprint with respect to central angle, footprint overlap, and coverage area are analyzed. A closed-form solution for the selection of true anomaly and RAAN separations is obtained. Time period and access time are analyzed with respect to the ground location of the target for different inclinations in order to obtain maximum possible access time of the satellite. Effects of RAAN rate and orbit inclinations are analyzed with respect to the Earth relative velocity for continuously covering certain latitude zones of the Earth. The geometry of footprint overlap is analyzed, and a closed-form expression for the maximum effective coverage angle is obtained.

Space and defence electronics system demands higher reliability for its operation in harsh (Maurer et al. in John Hopkins APL Tech Dig 28, 2008 [1]) and tough application environments. Development of processor (Azambuja et al. in IEEE Trans Nucl Sci 59, 2012 [2]) based electronics systems has advantages over sequencer logics, viz. high operating frequency, handle complex algorithm for autonomous operations, supports multiinput and multioutput systems, low form factor, ultra-low-power consumption and increased reliability. The operation of processor under various environmental conditions and its uninterrupted operations pave way to achieve successful missions. Most of the on-board computers for the space and defence systems are subjected to harsh environments due to cosmic radiations(aerial), high temperature, vibrations, sand and dust. Modular redundancy concepts aid to improve reliability however at the cost of increased form factor and enhanced power requirements. Researchers across the globe evolved with various methods and techniques to increase the ruggedness of the electronics systems. Traditional design hardening techniques viz. modular redundancies, parity algorithms, cyclic redundancy checks, single error correction and double error detection are in use to develop the fault-tolerant on-board computers. Since most of the fault-tolerant processors developed using in-house design, process hardening techniques for specific purpose, the availability of the same for the strategic programs and critical applications is limited and export controlled. To mitigate the scenarios, an indigenous effort has been made to architect, design, develop and validate the fault-tolerant processor using DRDO processor core (Series-A). The core is indigenously developed by DRDO. Using the fault-tolerant design methods, the indigenous processor core is upgraded to a fault-tolerant core. The core is RISC-based 32-bit architecture. The non-fault tolerant (FT) core is architected to incorporate the FT feature using the in-house and traditional techniques, viz. modular redundancy, single error correction and double error detection for FFs and registers, parity logics for memories. To provide the operational compatibility and to estimate the performance of FT-core using the user’s application algorithm, an integrated development environment (IDE) is designed, developed and validated. The FT-core with a frequency of up to 50 MHz in FPGA is validated using the application program for the function-A (aerial) and function-B (ground) strategic systems. The functional performance was on par with the industry architecture. The timing performance is satisfactory and met the user’s functional requirement. To meet the high performance and computational intensive requirements of the futuristic programs, the FT-core to be developed using dual or quad-core approach with appropriate fault-tolerant techniques.

Capability for on-orbit inspection using autonomous robots is important in manned space station/capsule operations. Due to the unpredictable lighting conditions on orbit, machine learning-based package identification and detection is a good candidate for inspection operations. In this work, the AI-based technique has been used to identify packages of interest from a set of packages in the orbital platforms. Images of different packages on the platform are used for the training using a simple neural network for fast inference of packages in the frame.

Pratham, IIT Bombay’s first student satellite was launched on September 26, 2016 onboard PSLV C-35. The intended mission life was of 4 months. The satellite was operational but the mission was not successful as only beacon signal was received and not the downlink signal. During the post-launch analysis, the team realized the importance of maintaining the quality. The IIT Bombay Student Satellite Team then started working on instituting quality assurance and management practices within the team. In the process, we listed and made few guidelines that can be useful for initiating and managing quality assurance in Student Satellite Projects across the world. This paper talks about our initial analysis and then lists down various activities to be performed to assure quality.

This paper chronicles the design of a very high-speed data acquisition system (DAS) which is qualified with a novel architecture for high-altitude applications. The state-of-the-art system is realized with DDR SDRAM modules as the storage media and the controller are designed with AX2000 FPGAs from Microsemi. The design of this high-altitude system matches the reliability requirements using system-level methodologies. The main features of the data acquisition system are the capability to acquire and retrieve up to 12,000 pulses with variable real-time programmable pre-pulse and post-pulse data. The system has modular design approach. Some of the main design challenges are highlighted here.

Indian Space Research Organization (ISRO)’s satellite technology has emerged on the global scale with its own IRNSS (Indian Regional Navigation Satellite System)—NavIC (Navigation with Indian Constellation) in outer space, and this paper aims to exploit one of its applications in traffic management for the complete city with inter-connected squares. This paper projects a novel design towards using NavIC to determine position of the vehicles, and as per their directional motion, number of vehicles moving towards particular road of any designated traffic square is determined. Suitable state machine and flow diagram are designed considering the real-time data. “Hand-over” mechanism is proposed to reduce the design complexity in terms of software implementation. Accordingly, the counter is made adaptive and dynamic. The finite-state machine (FSM) consists of 50 states, and very high-speed integrated circuit hardware description language (VHDL) coding and simulation are done using Libero SoC v11.9, Xilinx ISE 14.7 and ModelSim Microsemi 10.5c.

Internationally, space industry is going under transformational evolution which in turn is giving rise to new business model and opportunities to entrepreneurs, innovators and startups. Space sector in India is ready to take a next leap as its pioneer Indian Space Research Organization (ISRO) has come a long way by securing the podium place on launching 104 satellites into the intended orbit successfully. This exponential growth has propelled and fuelled the dream of many startups and innovators who are eager to contribute and explore the endless boundaries of space. This paper envisages that the opportunities startups can grab in space segment by learning to explore technological innovation and become commercially revenue garnering entity. Startup India initiative by the Indian Government and establishment of six space technology-based incubators by ISRO will build a strong ecosystem for nurturing the innovation and startups.

Power supply module is designed to provide input power to digital system and radio frequency system for space application. High power and package density in electronics are continuously increasing, which create more heat in the system. Thermal management in space application is even more challenging since the cooling of these systems primarily relies on conduction and radiation by ignoring convection for space application. The effective heat removal from a component and PCB to mechanical housing is through conduction and partly with radiation. In fact, thermal design in electronics through thermal modelling and simulation is becoming an integral part of the electronics system design process because it is normally less time consuming and also low expensive compared to the experimental cut-and-try approach. In paper, thermal design, analysis and testing of DC–DC power supplies module have been explained in detailed. Thermal analysis has been carried out using computational fluid dynamics technique, by considering total heat dissipation of 40 W. Thermal model was created using Ansys Icepak, and all active and highly dissipating component are modelled. Conduction and radiation mode of heat transfer are considered during the analysis by maintaining the base plate temperature at 55 °C. Based on the analysis results, the convertor has been manufactured, assembled and electrically tested. Thermal testing has been carried out at ambient air temperature of 25 °C for identifying the hot spot in the convertor using thermal camera before it is tested at high temperature. Thermal cycle testing of space grade DC–DC convertor is performed in thermo vacuum chamber at extreme cold temperature of negative 15 °C and extreme hot temperature of 55 °C under 1 × 10−5 torr pressure condition, during which electrical functionality of the module is monitored and temperature of active component is measured and the results of the thermal analysis and testing are elaborated and compared.

This paper involves a detailed thermal mathematical modeling of 300 mm × 300 mm deployed and body mounted solar panel of 10 kg sun-pointing Low Earth Orbit small satellite. An analytical calculation of solar panel both deployed and body mounted is performed using energy conservation principle. The system-level satellite thermal model along with solar panels is generated and solved numerically. In this paper, two configuration of body mounted solar panel in detailed system-level thermal model is studied. Firstly, the solar panel is fully conduction coupled with body structure panel. Secondly, conduction coupling is restricted by using 3 mm thick GFRP insulator at fastening points and resulting heat transfer by radiation is also studied. Also, the worst hot case temperatures of solar panel both deployed and body mount predicted by analytical approach is compared with numerical approach of the detail system-level thermal model. Hence, the presented method is beneficial for predicting the temperature profile of solar panel under worst hot condition.

This paper presents an approach to design an antenna deployment system (ADS) for nano-satellite applications. The known material properties, simulations on ANSYS for structural loads, mission requirements and simulations on HFSS for antenna characteristics were used to reach a design for the current version of the ADS. Mechanical testing was carried out to test the recovery of the antenna from the coiled state. The test ascertained the thickness and fabrication process to be used. The antenna was decided to be a dipole antenna operating at 145 MHz based on simulations as well as mission requirements.

The Student Satellite Team, IIT Bombay aims to develop India’s lightest actively controlled satellite. Major concern for the team is to keep the mass as low as possible while making the satellite structurally capable to bear random vibration, static and harmonic loads. The work below presents an approach towards designing satellite structure by considering mass minimization while maintaining the structural integrity of the satellite in launch and in-orbit environment.