COMMUTE: Cubesat Swarm Orbital Maneuvers for a Mission to Study Uranus’ aTmospheric Environment

Following recommendations from the 2023-2032 Planetary Science and Astrobiology Decadal Survey, new mission concepts are being developed with the focus of launching Uranus’ exploration missions in the early 2030s. To minimize both fuel consumption and cruise time on our way to Uranus, we propose a Jupiter-Uranus gravity assist trajectory using a Falcon Heavy Expendable Launcher to deliver a 3000 kg spacecraft to Uranus orbit in under seven years. The spacecraft will be composed of a mothership of 2000 kg wet mass and a swarm of CubeSats with a combined wet mass of 1000 kg. Using the ephemerides data of Earth, Jupiter and Uranus, and numerical solutions to the Lambert’s problem for a Jupiter flyby, we found that, with an initial launch window around April 15th, 2032, we reach Jupiter’s sphere of influence and perform a gravitational slingshot maneuver on December 31st, 2034, allowing the spacecraft to reach Uranus on December 31st, 2038. This proposed mission trajectory reaches Uranus with a relatively short cruise period of seven years, compared to the 13-year transfer period of the mission plan detailed in the decadal survey. This shorter transfer time could allow for significant extensions of the scientific mission nominal operations period and, potentially, reduce the cost of the overall mission. The swarm of 16 CubeSats of approximately 62 kg each will be divided into 4 groups of 4 identical spacecraft. Each group will be equipped with specialized instrumentation, exploring Uranus more extensively and performing planned plunges into its atmosphere while using the mothership as a communications relay with the Earth. This research demonstrates that a CubeSat swarm mission to Uranus can be not only viable, but also a fuel and cruise time optimization opportunity, delivering 16 exploration spacecraft to Uranus in under seven years

Control of Small Spacecraft by Optimal Output Regulation: A Reinforcement Learning Approach

The growing number of noncooperative flying objects has prompted interest in sample-return and space debris removal missions. Current solutions are both costly and largely dependent on specific object identification and capture methods. In this paper, a low-cost modular approach for control of a swarm flight of small satellites in rendezvous and capture missions is proposed by solving the optimal output regulation problem. By integrating the theories of tracking control, adaptive optimal control, and output regulation, the optimal control policy is designed as a feedback-feedforward controller to guarantee the asymptotic tracking of a class of reference input generated by the leader. The estimated state vector of the space object of interest and communication within satellites is assumed to be available. The controller rejects the nonvanishing disturbances injected into the follower satellite while maintaining the closed-loop stability of the overall leader-follower system. The simulation results under the Basilisk-ROS2 framework environment for high-fidelity space applications with accurate spacecraft dynamics, are compared with those from a classical linear quadratic regulator controller, and the results reveal the efficiency and practicality of the proposed method

A Methodology for Measuring the Air Infiltration Rates into Refrigerated Compartments

This work focuses on the development of a methodology for measuring the air infiltration rates into refrigerated compartments based on the tracer gas technique. To this end an experimental apparatus was designed and constructed. The apparatus is comprised of a device to supply the tracer gas uniformly into the zone, a device to ensure a uniform concentration both in the vertical and horizontal directions, a device to collect the air samples and a gas analyzer to measure the gas concentration according to the infrared photoacoustic spectroscopy technique. Tests were carried out with three distinct refrigerators, with the compressor on and off, using both the concentration decay and constant concentration measurement techniques. A high level of repeatability in the air infiltration rate measurements has been found, with deviations always lower than 5%. It was also noted that the two measurement techniques provide practically the same results. Furthermore, it was observed that the air infiltration rate is approximately five times higher when the compressor is turned on due to the higher pressure difference created by the higher temperature difference. Adsorption tests were also carried out indicating that the adsorption of the tracer gas by the refrigerator materials and gasket is insignificant when compared to the concentration variations caused by air infiltration. Manufacturing deviations were also simulated and it was found that they greatly increase the air infiltration rate through the gasket. The developed technique is reliable and practical and can become a useful tool for the design and analysis of household refrigerators

Avaliação comparativa da perda de retenção de quatro sistemas de encaixes do tipo ERA e O-Ring empregados sob overdentures em função do tempo de uso

The aim of this study was to evaluate and compare the retentive capacity between two O-ring and O-SO system (Group I), and two ERA system types - white and gray retention caps - (Group II), respectively, in simulated function in database, 6 months, 1, 2, 3, 4, and 5 years later, with insertion and removal cycles. Two Brånemark implants were fixed in two trapezoidal metallic bases for the tests. Removal and insertion tests were done in a sewing machine, adjusted for this purpose using a belt and a pulley system, moving a steel crankshaft. A delineator platinum hold was used for body trial fixation to the metallic base of the sewing machine. Resistance test to axial movement of the caps by tension was done in a Universal test machine in an established period of time, before and after the cycle accomplishments in the adapted sewing machine. Based on the results, this study concludes that all the attachment systems tested showed retention loss during the experiment; the ERA system showed, since the beginning, higher retention compared to the other systems and the gray colored attachment showed the best result in the end of the simulated use test.O objetivo deste trabalho foi avaliar e comparar a capacidade retentiva entre dois sistemas O-ring e O-SO (Grupo I) e dois sistemas ERA - cápsulas de retenção cinza e brancas - (Grupo II), em função simulada, com ciclos de inserção e remoção, do início, 6 meses, 1, 2, 3, 4 e 5 anos depois. Dois implantes do tipo Brånemark foram fixados em duas bases metálicas trapezoidais, sendo os testes de inserção e remoção feitos numa máquina de costura ajustada para este propósito, usando um sistema de correias e polias, movendo um virabrequim. Uma platina de delineador foi usada para fixação dos corpos de prova às bases metálicas na máquina de costura. Os testes de resistência ao movimento axial das cápsulas por tensão foram feitos numa máquina de ensaios universal em períodos de tempo pré-estabelecidos, antes e após o término do ciclo na máquina de costura. Baseado nos resultados, esse estudo pôde concluir que houve perda de retenção de todos os sistemas testados, porém o sistema ERA apresentou, desde o início, maior retenção quando comparado aos outros sistemas e a cápsula cinza mostrou o melhor resultado no final da simulação

The Deformable Mirror Demonstration Mission (DeMi) CubeSat: optomechanical design validation and laboratory calibration

Coronagraphs on future space telescopes will require precise wavefront correction to detect Earth-like exoplanets near their host stars. High-actuator count microelectromechanical system (MEMS) deformable mirrors provide wavefront control with low size, weight, and power. The Deformable Mirror Demonstration Mission (DeMi) payload will demonstrate a 140 actuator MEMS deformable mirror (DM) with \SI{5.5}{\micro\meter} maximum stroke. We present the flight optomechanical design, lab tests of the flight wavefront sensor and wavefront reconstructor, and simulations of closed-loop control of wavefront aberrations. We also present the compact flight DM controller, capable of driving up to 192 actuator channels at 0-250V with 14-bit resolution. Two embedded Raspberry Pi 3 compute modules are used for task management and wavefront reconstruction. The spacecraft is a 6U CubeSat (30 cm x 20 cm x 10 cm) and launch is planned for 2019.Comment: 15 pages, 10 figues. Presented at SPIE Astronomical Telescopes + Instrumentation, Austin, Texas, US

Folded Lightweight Actuator Positioning System (FLAPS)

Precision actuation of mechanical structures on small spacecraft is challenging. Current solutions include single-use actuators, which rely on pyrotechnics and springs, and multiple-use actuators, which typically consume more size, weight, and power than available on CubeSats. The Folded Lightweight Actuated Positioning System (FLAPS) demonstrates the use of a simple rotary shape memory alloy (SMA) actuator in a bending architecture, along with a feedback control loop for repeatable and precise deployment. The FLAPS mechanism consists of a pair of SMA strips mounted to a hinge assembly, with one side attached to the CubeSat bus and the other to the deployable element. A custom actuator shape was manufactured using oven annealing. SMA actuation is achieved using joule heating. Feedback control is provided by a closed-loop PID control scheme, feedback sensor, and controller board. The FLAPS actuator is currently being developed for CubeSat solar panel positioning and drag control. Other potential FLAPS applications include aperture repositioning, deployable radiators, and steerable antennas. The FLAPS team will validate the actuator system in a microgravity environment on a parabolic fight in late 2019

Calibration and Testing of the Deformable Mirror Demonstration Mission (DeMi) CubeSat Payload

The Deformable Mirror Demonstration Mission (DeMi) is a 6U CubeSat that will operate and characterize the on-orbit performance of a Microelectromechanical Systems (MEMS) deformable mirror (DM) with both an image plane and a Shack-Hartmann wavefront sensor (SHWFS). Coronagraphs on future space telescopes will require precise wavefront control to detect and characterize Earth-like exoplanets. High-actuator count MEMS deformable mirrors can provide wavefront control with low size, weight, and power. The DeMi payload will characterize the on-orbit performance of a 140 actuator MEMS DM with 5.5 _m maximum stroke, with a goal of measuring individual actuator wavefront displacement contributions to a precision of 12 nm. The payload will be able to measure low order aberrations to l/10 accuracy and l/50 precision, and will correct static and dynamic wavefront phase errors to less than 100 nm RMS. The DeMi team developed miniaturized DM driver boards to fit within the CubeSat form factor, and two cross-strapped Raspberry Pi 3 boards are used as payload computers. We present an overview of the payload design, the assembly, integration and test progress, and the miniaturized DM driver characterization process. Launch is planned for late 2019

Testing of the CubeSat Laser Infrared CrosslinK (CLICK-A) Payload

The CubeSat Laser Infrared CrosslinK (CLICK-A) is a risk-reduction mission that will demonstrate a miniaturized optical transmitter capable of ≥10 Mbps optical downlinks from a 3U CubeSat to aportable 30 cm optical ground telescope. The payload is jointly developed by MIT and NASA ARC, and is on schedule for a 2020 bus integration and 2021 launch. The mission purpose is to reduce risk to its follow-up in 2022, called CLICK-B/C, that plans to demonstrate ≥20 Mbps intersatellite optical crosslinks and precision ranging between two 3U CubeSats. The 1.4U CLICK-A payload will fly on a Blue Canyon Technologies 3U bus inserted into a 400 km orbit. The payload will demonstrate both the transmitter optoelectronics and the fine-pointing system based on a MEMS fast steering mirror, which enables precision pointing of its 1300 μrad full-width half-maximum (FWHM) downlink beam with anestimated error of 136.9 μrad (3-σ) for a pointing loss of -0.134 dB (3-σ) at the time of link closure. We present recent test results of the CLICK-A payload, including results from thermal-vacuum testing, beam characterization, functional testing of the transmitter, and thermal analyses including measurement of deformation due to the thermal loading of the MEMS FSM

Thermomechanical design and testing of the Deformable Mirror Demonstration Mission (DeMi) CubeSat

The Deformable Mirror Demonstration Mission (DeMi) is a 6U CubeSat that will operate and characterize the on-orbit performance of a Microelectromechanical Systems (MEMS) deformable mirror (DM) with both an image plane and a Shack-Hartmann wavefront sensor (SHWFS). Coronagraphs on future space telescopes will require precise wavefront control to detect and characterize Earth-like exoplanets. High-actuator count MEMS deformable mirrors can provide wavefront control with low size, weight, and power. The DeMi payload will characterize the on-orbit performance of a 140 actuator MEMS DM with 5.5 μm maximum stroke, with a goal of measuring individual actuator wavefront displacement contributions to a precision of 12 nm. The payload is designed to measure low order aberrations to λ/10 accuracy and λ/50 precision, and correct static and dynamic wavefront phase errors to less than 100 nm RMS. The thermal stability of the payload is key to maintaining the errors below that threshold. To decrease mismatches between coefficients of thermal expansion, the payload structure is made out of a single material, aluminum 7075. The gap between the structural components of the payload was filled with a thermal gap filler to increase the temperature homogeneity of the payload. The fixture that holds the payload into the bus is a set of three titanium flexures, which decrease the thermal conductivity between the bus and the payload while providing flexibility for the payload to expand without being deformed. The mounts for the optical components are attached to the main optical bench through kinematic coupling to allow precision assembly and location repeatability. The MEMS DM is controlled by miniaturized high-voltage driver electronics. Two cross-strapped Raspberry Pi 3 payload computers interface with the DM drive electronics. Each Raspberry Pi is paired to read out one of the wavefront sensor cameras. The DeMi payload is ~4.5U in volume, 2.5 kg in mass, and is flying on a 6U spacecraft built by Blue Canyon Technologies. The satellite launch was on February15,2020 onboard a Northrop Grumman Antares rocket, lifting off from the NASA Wallops Flight Facility. We present the mechanical design of the payload, the thermal considerations and decisions taken into the design, the manufacturing process of the flight hardware, and the environmental testing results