Advances in Small Satellite Technologies

BeliefSat-1 is the first student satellite of K. J. Somaiya Institute of Engineering and Information Technology, Sion, Mumbai being developed as a sub-part of the institute’s proposed payload under ISRO’s PS4OP programme. It is being developed as per PocketQube standard with 2P (50 × 50 × 114 mm³, ≤ 500 g) unit specifications for satisfying mechanical, power and tracking requirements. The major goals of the satellite include the demonstration of innovative fabrication approach based on interlocked PCBs serving as a structural material, providing digipeater service to amateur radio community, flight proving communication and power system based on COTS electronic for usage in future missions. The design has been evolved around the Arduino ecosystem, to make it easily adaptable by future missions. The paper first highlights certain calculations for getting a range of operational parameters and then describes different subsystems implemented within the satellite design. The mission assumes 550 km sun-synchronous orbit. Overall, it uses Atmega1284p as on-board computer, IXYS monocrystalline solar cells for power generation, SPV1040 for maximum power point tracking, LORA1268F30 as transceiver, Arducam 2MP camera as imaging payload, Samsung 18,650 batteries and COTS sensors (HMC5883L magnetometer, MPU6050 gyro-sensor). The paper also describes minimal setups for enabling reception by students all over India.

This paper reports the design, operating principle, testing and qualification of a compact demonstrator payload designed for evaluating the performance RF micro-electromechanical systems (MEMS) switches in low earth orbit conditions on an experimental small satellite platform. Key aspects calibration of RF level detectors and the approach for fault detection from telemetry data are discussed. The payload has a weight of 800 g, foot print of 100 mm × 110 mm, height 100 mm and power consumption 4.2 W. Telemetry requirements are suitable for typical small satellite missions. The electronics components used are commercial grade which have undergone payload level qualification tests for the intended mission using the proto flight model test philosophy.

Student-built space missions offer prospects to commence small space missions with small financial budgets. The significant outcome along with technical intrinsic worth is the cross-disciplinary training for students that endow a future career in the space industry. The experience in learning and working with hands-on, project-based education proves effective for talent management among the student community. Programs such as the CubeSat standard have dramatically changed spacecraft engineering education, fostering aerospace education resulting in designing and building small satellites. The paper explores the proliferation of the status of the University/College supported student-built space missions over the last years and the various levels of applications of small satellite platforms in India. Moreover, how the student satellite projects translated into useful small satellite commercial missions is also discussed.

Satellite consists of many electronic packages, and during the launch of a satellite, it will be subjected to high random vibration loads which are generated by the launch vehicle. These electronic packages also known as modules, systems and/or subsystems will experience these vibration loads when mounted in the satellite and must be designed and qualified to meet these launch loads. These modules house the Printed Circuit Board (PCB) on which critical components such as HMCs, capacitors, CCGAs, flat packs and other leaded components, which are prone to vibration failures, need to be taken care during the design phase. Above all these components are high reliable and high cost items, this puts the constraint of first time right, and no failure is allowed. Hence, structural design and analysis become most important stage before manufacturing and testing the packages. This paper discusses one of the electronic packages designed for Indian communication satellite to meet such vibration loads and methods used to analyze the package virtually (through analysis software) before testing which saves the time and cost by reducing the number of prototypes. Three analyses (Eigen value analysis, quasi static analysis and random vibration analysis) have been carried out to qualify the module virtually by finite element method. Modal analysis has been carried out to determine the natural frequency and modal effective mass of the module. Quasi static analysis has been carried out to determine the stresses. Random vibration analysis has been carried out to see the acceleration levels and overall g_rms along with the stress levels experienced by the components. Based on the analyses, the qualification model (QM) was built and validated through testing after which the proto-flight model (PFM) and flight models were built.

The Transmit–Receive module is one of the prime sub-systems required for designing an airborne or satellite imaging radar. This paper presents an overall detail of a highly optimized Ku-band Transmit–Receive module consisting of exciter and receiver sub-systems. The exciter receives the IF signal with necessary modulation like Linear Frequency Modulation (LFM) from signal processor module. The IF signal is upconverted to the transmit frequency using coherent LO signal and suitably amplified to the level required by the HPA. Suitable Ku-band calibration signal is also generated in this module for carrying out end-to-end calibration of the entire Transmit–Receive system. A common LO signal for both up-converter and down-converter is generated using a frequency synthesizer which is locked to high-stability Oven Controlled X-tal Oscillator (OCXO) for achieving high stability as well as extreme low phase noise. Various clock signals required for ADC/DAC/FPGA, etc. are also generated in this unit and are locked to the OCXO frequency. The received signal from antenna is amplified using a Low Noise Amplifier (LNA) after necessary protection with a limiter and downconverted to IF frequency using same LO used for up-conversion. The baseband section of the receiver provides necessary amplification to ensure optimum operation over the entire dynamic range. The signal is fed to the baseband data handling processor unit where it is digitized and processed. A compact Electronic Power Conditioner (EPC) is also housed along with the Transmit–Receive module for providing conditioned power supply. The paper briefly describes various options considered before freezing the final configuration. All the circuits are realized using Coplanar Waveguide (CPW) transmission lines on RT 6002, 10Mil substrate. To ensure stable and spurious free performance, the entire Transmit–Receive module is realized using several compartments to reduce coupling among various stages. Special attention has been given to reduce the DC power requirement. Both thermal as well as structural issues are addressed while finalizing the mechanical housing. As the system is designed for satellite application, the paper also addresses various guidelines followed for component selection, fabrication processes, testing, qualification, etc. Various techniques adopted for the realization of low-cost/high-rel hardware are also addressed in the paper.

Automatic interpretation of remote sensing images is a fundamental but challenging problem in the field of aerial and satellite image analysis. It plays a vital role in a wide range of applications and is receiving significant attention in recent years. Even though many great progress has been made in this field, the detection of multi-scale objects, especially small objects in high-resolution satellite (HRS) and drone images, has not been adequately explored. As a result, detection performance both in terms of detection speed and accuracy turns out to be poor. To address this problem, we propose a convolutional neural network (CNN)-based single-stage object detector for the real-time and accurate recognition of remote sensing images. Our model predicts bounding boxes and corresponding class probabilities directly from images in a single assessment. This will result in a real-time object detection of images without compromising accuracy.

PS4-Orbital Platform (PS4-OP) refers to a novel idea by ISRO (Indian Space Research Organisation) to use the spent PS4 stage (fourth stage of PSLV) as a three-axis-stabilized platform for small scientific payloads to carry out in-orbit scientific experiments for an extended duration of 4–6 months. The PS4 stage has standard interfaces and packages for power generation, telemetry, tele-command, stabilization, orbit keeping and orbit maneuvering. The scientific community/research organizations can design the scientific payload and utilize the OP interfaces offered by spent PS4 stage for powering, data management and specific experimental requirements. In short, ISRO is extending its expertise in space technology to the scientific community as a platform to design, develop and validate their experiments in an effective manner.

An Antenna Deployment System has become an essential component of any pico- or nano-satellite design due to space constraints during launch. The Sanket mission is a technology demonstration designed to be flown on the Indian Space Research Organization’s PSLV Stage 4 Orbital Platform (PS4-OP) (Announcement of opportunity for orbital platform. Indian Space Research Organisation, Bangalore [1]) and aims to qualify the team’s Antenna Deployment System in Ultra High Frequency band to a TRL-7 (Technology Readiness Level) in Low Earth Orbit (LEO). Sanket, i.e. the complete system, comprises an ADS and an auxiliary system. The purpose of the auxiliary system is to test the ADS on PS4-OP simulating a 1U CubeSat mission life cycle and conditions. Sanket will be mounted on PS4-OP which remains in LEO for around 6 months. Our Antenna Deployment System is developed as an independent module that is compatible with standard CubeSat sizes 1U, 2U and 3U.

Following the Indian Space Research Organisation’s (ISRO) novel idea of using the spent PS4 stage of the PSLV as a platform for scientific experiments and technology demonstrations, a unique opportunity has been created to develop and test small-satellite subsystems and instruments as PS4-OP (PSLV 4th Stage Orbital Platform) payloads, prior to the launch of the small satellite itself. The PiLOT (PS4 in-orbital OBC and TTC) payload has been developed with an engineering objective of flight qualifying the in-house developed small-satellite subsystems, which are the OBC (On-Board Computer) and TTC (Telemetry and Telecommand) subsystems. The PiLOT payload also contains a RADFET (Radiation Field Effect Transistor) sensor for monitoring radiation dosage. Mapping the temporal and spatial distribution of the radiation dosage will help improve the understanding of the space weather in the low earth orbit. This paper describes in detail the development of the PiLOT PS4-OP payload including a description of the different subsystems of the payload, which are the RADFET sensor board, the OBC, the TTC board, and the PS4-OP interface board. The paper concludes with a description of the integration, testing methodology, and the current developmental status of the payload.

In this work, design, development and testing of a sensor module for small satellite application using commercial-off-the-shelf (COTS) components was performed in connection with the student satellite programme of College of Engineering Trivandrum (CETSAT). Various initiatives to democratize space missions have been implemented all over the world. Indeed, COTS components are finding their way into small satellites lately as they lead to significant cost reductions and are easily available. It involves using industrial and automotive-grade electronics and other elements from nontraditional space markets. Such components are particularly useful in missions on a small-time frame or budget. Testing the space-worthiness and long-time durable data collection while in orbit of these COTS is an important preliminary step to be carried out prior to the launch of a full-fledged satellite. The payload constructed by College Of Engineering Trivandrum is a sensor module composed of a 3-axis magnetometer with associated breakout board as well as four temperature sensors to obtain the attitude and temperature information, respectively. The deployment of the magnetometer also helps in confirming the effectiveness or lack thereof in the magnetometer-only approach. A ground station was designed and set up to receive data from the satellite module and calibrate its performance with the data obtained by the ground station of Indian Space Research Organization. The chosen sensors were passive sensors and it is hypothesized that such sensors have a lower chance of failure in the Thermovac and Vibration tests as well as in the Low Earth Orbit (LEO). The details of the design, fabrication and testing are presented in this paper and inferences of the performance of COTS sensors in simulated environments of LEO were made.

Future human-space exploration efforts aim to achieve maximal synergy between human and robotic missions, where in human-scale robots shall supplement astronaut exploration activities and also undertake robotic precursor missions. Towards this, design and development of a flying Smart Space Robot (SSR) is initiated. The Smart Space Robot (SSR) is a space flying robot which will be tethered to the unmanned orbital platform on the fourth stage (PS4-UOP) of Polar Satellite Launch Vehicle (PSLV) for micro g experiments. SSR is of nano-satellite class with dimensions of ~350 × 350 × 350 mm, weight of ~12 kg and power of ~30 W. The nanosat has to undergo different phases of operations which include deployment, station keeping, retrieval and docking. These manoeuvres calls for a robust system capable of operating for an extended duration with multiple restart and pulsing capability. This paper details out the development and design process of a cold gas system within the constraints of space, volume, mass, voltage and power as designated by the nanosat specification which meets the mission requirements.

Ring Laser Gyroscopes (RLGs) offer highly accurate angular measurement capability for inertial class applications. They offer high stability in scale factor and bias drift. The gyro being a voluminous module with at least ten specific components or interfaces which are sensitive to temperature, thermal compensation becomes a complicated problem. Further, the direct thermal measurements are limited and sensitivity factors have to be derived from available monitoring parameters, which increases the complexity of the problem manifold. Published literature is available on many compensation algorithms; however, they limit their research mostly to  using temperature reading of the optical block in order to predict bias. In this paper, we study the drift variations in the gyro as a function of temperature monitored from the optical block as well as the control signals used to sustain the optimal mode of the laser including path length control signal, high frequency oscillator signal and dithering frequency variations. We applied these data on linear and polynomial regression models and obtained up to 76% and 88% reduction in bias drift, respectively. Further, the denoising filter is studied by varying its type and the window length. Training and testing sensors in INS cluster configuration gave bias drift reduction of up to 74%. All these improvements are achieved in the navigation loop cycle timing itself enabling a smooth compensation for INS applications.

The reaction wheels play a major role in attitude control of the spacecraft. As they work on the principle of conservation of angular momentum, any ripple in reaction torque will affect the performance of critical on-board instruments such as atomic clock, high-resolution cameras, etc. in the spacecraft. The three-axis stabilization system using reaction wheels aims to provide high-pointing accuracy, low jitter and long-term attitude stability. The stringent attitude and micro-vibration requirement of future missions cannot be met practically with the existing reaction wheels with conventional trapezoidal brushless direct current motor (BLDC) drive. A suitable upgrade in reaction wheels with permanent magnet synchronous motor (PMSM) instead of BLDC motor and field-oriented control instead of six-step commutation will be enabled to meet these requirements. Improvements in torque ripple will give additional benefit in losses and hence lesser load on thermal system. This study investigates the possible usage of PMSM and field-oriented control for future generation reaction wheels and its implementation in FPGA. Simulation results showed improved steady-state performance in terms of torque ripple without any degradation in the transient response of the wheels. The study can be well extended to gimbal control of CMG and other high-precision servo applications in the spacecraft.

Miniaturisation of electronics, newer materials especially composites and advanced manufacturing techniques including 3D printing has encouraged the development of small satellites. This enabled many commercial companies to venture in to this area since the development can be made possible with small budget. The market size of small satellite is expected to flourish in the next few years. Small satellites are used mainly in the areas of Earth observation services, monitoring of agricultural fields, detection of climatic changes, disaster mitigation, meteorology, etc. In the current day scenario of launch vehicle industry, every mission of a launch vehicle is unique since they are used for putting satellites of different mass, size, shape, purpose and to various orbits. Satellites are exposed to dynamic/fluctuating excitations due to different dynamic environment from lift off to the completion of its intended mission. The dynamic environment of the satellite will be different for different launch vehicle configuration and its mounting scheme. This paper addresses quasi-static and dynamic analyses carried out for a typical small satellite using indigenously developed finite element software FEAST^SMT by Vikram Sarabhai Space Centre (VSSC). This study helps in assessing the structural adequacy of small satellite as well as the acceleration level experienced by different components and packages. This will help the satellite team to modify the structure, change the mounting location or by adopting new mounting schemes of components/packages in order to optimise the design well before realising the hardware.

Tape spring mechanisms have been widely used to design deployable structures for satellites. In this paper, the dynamics of the deployment of VHF tape spring antennas from a CubeSat have been discussed. A set of nonlinear equations of motion representing the coupled dynamics of the deployment and the overall motion of the CubeSat were obtained using Kane’s method. These were integrated numerically to obtain parameters of interest like the time taken for deployment, the trajectory of the antenna, and variation of the CubeSat angular velocities during the interval of deployment. These models have been used to analyze the design of Advitiy (1U CubeSat conceptualized by IIT Bombay’s Student Satellite Team). Based on simulations, comments have been made on this design of the antenna deployment system, and few performance specifications have been laid down for the CubeSat.

Continuous development in electronics industries results in the introduction of many complex electronics packages. Nowadays, digital and frequency modulation technology occupies a major part in commercial, telecommunication, space, and defence electronics applications. Space industries use high-reliable electronics packages like the Ceramic Column Grid Array package (CCGA) which is programmable as per application in satellite modules. These packages have more number of columns for electronics as well as for mechanical strength. As the number of columns increases, mechanical strength increases, but same time screening of columns that are very closely constructed is very difficult. When these packages undergo screening tests like thermal cycling, there may be chances of solder cracks and bending of columns due to temperature variation. Thereafter, the high-level vibration test makes the solder cracks grow faster and bent columns to fail. These grid array packages shall be integrated onto PCB through reflow technology; reflow creates the solder fillets at the bottom portion of columns which mounts upon PCB. During these reflow processes, in any chance the solder fillet is not properly formed, and vibration test followed by thermal cycling may lead to damage of columns in-turn whole package. As the cost of one package itself is very high, damage to multiple packages may result in huge losses to industries. In this paper, two numbers of 1752 columns CCGA (VERTEX 5 CN1752 CCGA) packages are considered which are mounted on 2.6-mm thick, 12-layer polymide PCB. The board along with CCGA packages is placed inside an aluminium housing, and this assembly is considered to perform dynamic analysis like eigen value, sine response and random response analysis and is subjected to vibration tests to study the solder joint behaviour of CCGA columns; observations are discussed in detail.

Emerging small satellite market faces many challenges mainly to protect their payload during the launch. Satellite team has to share their finite element model to the launch vehicle team for the various analysis like launch vehicle dynamic characteristics, Coupled Load Analysis (CLA), dynamic envelop studies, etc. FE model contains sensitive details about the satellite, and dynamic condensation of the satellite is the only way to protect satellite details. This also helps in reducing the number of degrees of freedom of the satellite finite element model. Craig-Bampton technique is the most popular method to reduce the satellite into a mass and stiffness matrix with the required number of modes. Satellite team needs a physical node at critical locations to study the responses, and these nodes need to be retained while reduction. This paper examines the Craig-Bampton reduction method and discusses how to compute the response at various locations of the satellite. A frequency response analysis with reduced model and full model is also compared in this paper.

The demand for a small satellite launch vehicle for ISRO has been ever increasing due to growing launch requirements from the rising number of industrial and academic players who are venturing into design and fabrication of smaller satellites for their specific applications. Therefore, a new lightweight navigation system needs to be designed for catering to the requirements of a low weight small satellite launch vehicle. The major mechanical component of a navigation package is the cluster (chassis) which is used for mounting gyroscopes and accelerometers required for providing information about the position and attitude of the launch vehicle. The cluster adds a good percentage to the total mass of the package. As a first step towards proving the technical feasibility of AM route, the CNC machinable cluster was 3D printed along with test coupons, and their structural and thermal characteristics were studied. Then, improvements were made on the cluster to increase one of the normal mode frequencies using design for additive manufacturing philosophy. This modification also helped to reduce the mass, which was made possible with the flexibilities offered by additive manufacturing in terms of manufacturing constraints. The modal analysis carried out on the modified cluster proved the improvement in performance of the cluster. The modified cluster was successfully 3D printed, followed by testing to meet all geometric and performance requirements.

Spacecraft elements like reaction wheels are subjected to random vibration tests. These tests are done to simulate launch loads during ascend phase of the launch, to establish the design margins, and to bring out workmanship-related issues. Random vibration tests typically excite all natural frequencies of the structure within the excitation bandwidth. The excitation bandwidth typically ranges between 20 and 2000 Hz. The excited range of frequencies is significantly wide, so that it is practically impossible to design a structure without any natural frequencies within the range of excited frequencies. Hence, they must be designed in such a way that they must perform satisfactorily during these tests even when they are resonated at their natural frequencies. Flywheels of reaction wheels are subjected to very high vibration loads; these loads are transferred to bearing unit through the assembly joints. The integrity of these joints plays a critical role in change in unbalance of reaction wheels during vibration. The reaction wheels used for spacecraft applications are balanced to very fine levels such that the on-orbit vibrations generated by a running reaction wheel does not affect the performance of payloads onboard a spacecraft (especially optical payloads). Random vibration tests may slightly increase the unbalance of a reaction wheel, but the change in unbalance of a wheel after vibration test must be limited within specified values. As part of qualification of a new generation reaction wheel it was subjected to random vibration test. After the test, it was seen that the dynamic unbalance value after vibration test has gone out of specification. This study indents to find the reason for this deviation and propose design improvements for limiting the change in dynamic unbalance values within the acceptable range.

Understanding the radiation environment of the low earth orbit (LEO) is of considerable scientific interest, due to its unpredictable nature and variability. Temporal and spatial distribution of the ionizing as well as non-ionizing radiation levels in the LEO are especially important considering the negative impact these radiation have on the spacecraft avionics by causing Single Event Upsets (SEUs) and single-event latchups (SELs). With in-situ data about the radiation environment, we can gain an insight on the reliability of various Commercial off-the-shelf (COTS) in the space environment, as ground-based testing with simulated radiation levels for a long duration is often not practical. Further, the data collected about the ionizing radiation levels can also prove useful for the upcoming human space flight mission planning, wherein shielding against ionizing radiation is an important aspect to consider regarding the safety of the astronaut. With this primary objective of in-situ radiation dosimetry, the Ahan 3U CubeSat has been developed with a mass of around 2 kg and a nominal power consumption of less than 3 Watts. The satellite contains two different sensors, a Geiger-Muller Counter and a Radiation Field Effect Transistor (RADFET) to monitor radiation dosage in the LEO. This paper describes the design and development of the above-mentioned CubeSat mission.

Digital technology is trending in modern electronic applications. Due to its high reliability, space technology uses electronic packages like CCGA which are molded inside ceramic case hence named ceramic packages. Some of these packages have thermal path from junction (die) to top surface of the package. Thus, external thermal coupling is required to drain out excess of heat to maintain the temperature within allowable limits for safe operation of satellite. Since, these packages are high in cost and one-time programmable devices, cannot be used for R&D purpose; hence, equivalent thermal component which dissipates the way CCGA dissipates is employed to duplicate the model. Aluminum heat sink is used as a thermal coupler to drain out excess of heat. In this paper, effectiveness of heat sink is studied. Thermal analysis has been performed using CFD software by considering different heatsink configurations. Based on analysis results, the best heatsink design is selected and fabricated. Experiments are performed under different boundary conditions as per the analysis. Assembled unit is subjected to thermal testing, and the results are compared with that of thermal simulation, and they are well within the acceptable limit. Using these results, the same heat sink is implemented with actual CCGA to maintain junction temperature within allowable limit specified by the manufacturer.

In recent years, the space sector is going through an incredible transformation. The services from large satellites are actively considered for replacement with smaller satellites which have several advantages compared to the traditional satellites. These are designed and manufactured at a lower cost with flexibility in reducing the lead times. In addition, several low earth orbit satellite Iinternet constellations using small satellites are already being developed for low-cost and low-latency Internet access from space. The services are comparable with addition of the newer functions. The experimentation possibilities using new technologies are more feasible resulting in an accelerated pace of innovation. Starting from the selection of components including those which are rated non-space Commercial Off-The-Shelf (COTS), following the revised derating guidelines, up screening the modules and subsystems to the required environmental stress levels, and complying with the modified product realization and packaging guidelines are some of the path-breaking practices that are becoming common in faster realization of small satellite. Modular designs and application of best practices of a manufacturing industry set up dominate the cost expectations from the small satellite programs. The miniaturization trends in multiple technologies such as in sensors, computing and electronic subsystems have benefited different types of small satellites. Communication, data, earth observation and navigation areas are benefitted in small satellites with the technological advancements in the commercial and industrial domains and their chipsets. While the small satellites are generally in low earth orbits (LEOs), they can offer services comparable to larger satellites in medium earth orbits (MEOs) or geosationary earth orbit (GEOs). This paper describes the small satellite classification, drivers for their growth, technology trends, and applications. It also explains the eco-system available and the players involved in the growth of the small satellite markets.

Coherent population trapping (CPT) is a promising technique to realize miniaturized atomic frequency standard (clocks) and magnetic field sensor (magnetometer) for precision measurements in space. Atomic clocks in particular are the reference source for precision time and frequency signals in navigational satellites. Frequency stability of atomic clocks and sensitivity of magnetometer depend primarily on the quality factor of CPT resonance which is the function of amplitude (contrast) and line width (\(\Delta \nu_{{\text{FWHM}}}\)). In this paper, the CPT resonance characteristics are studied with respect to various critical operating parameters such as laser excitation intensity (atomic excitation), cell temperature, and radio frequency (RF) power. Our study shows that FWHM increases linearly with laser intensity, whereas resonance signal contrast increases with cell temperature up to an optimum temperature and then after decreases with increase in temperature. The numerical values for these operating parameters are then derived by analyzing the experimental results and then arrived at an optimized quality factor for the atomic system.

Global navigation satellite system has become a de-facto standard for navigation, with enormous applications worldwide. Currently, four global navigation satellite systems are operational and a number of regional navigation satellite systems are also in place including Indian Regional Navigation Satellite System (IRNSS) which is also known as Navigation with Indian Constellation (NavIC). NavIC is a regional navigation satellite system owned and operated by ISRO. The application of GNSS ranges from commercial applications to intercontinental ballistic missiles or heavy lifters of ISRO/NASA. High dynamic NavIC receivers are used in ISRO launch vehicles for preliminary orbit determination/closed loop guidance along with inertial navigation system in an integrated manner. When such critical applications of GNSS are considered, the integrity of navigation data in real time is a prime concern. Different methods are developed by different agencies over time in this area. Usage of inbuilt health information from satellite is one of the possible methods of integrity check. Receiver autonomous integrity monitoring (RAIM) provides integrity checking of GNSS without an external aid. NavIC/GPS receiver architecture is presented in this paper, the satellite signal analysis is done, and possible integrity threats are identified. The different approaches are considered in RAIM algorithm, and a simplified RAIM approach is developed which has good real-time performance. The simulation studies and flight software development are completed.

Protective shields are indispensable for protecting the satellites components from debris impact. Proper evaluation of the response of these shields is important. High velocity impact simulations were conducted to ascertain the response of single bumper Whipple shields used as satellite shields. The study was conducted with the Abaqus explicit solver with the smooth particle hydrodynamics (SPH) framework. A simple damage-based failure model was able to capture the fragmentation as well as the penetration of the shield under the impact. The energy as well as the residual velocity were seen to match with the similar experiments.