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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2020

Snow and Water Imaging Spectrometer: Final Instrument Characterization

The Snow and Water Imaging Spectrometer (SWIS) is a science-grade imaging spectrometer and telescope system suitable for CubeSat applications, spanning a 350-1700 nm spectral range with 5.7 nm sampling, a 10 degree field of view and 0.3 mrad spatial resolution. The system operates at F/1.8, providing high throughput for low-reflectivity water surfaces, while avoiding saturation over bright snow or clouds. The SWIS design utilizes heritage from previously demonstrated instruments on airborne platforms, while advancing the state of the art in compact sensors of this kind in terms of size and spectral coverage. Through frequent repeat observations from space at a moderate spatial resolution, SWIS can address key science questions concerning aquatic and terrestrial ecosystem changes, cryosphere warming and melt behavior, cloud and atmospheric science, and potential impacts of climate change and human activities on the environment. We review the optical design and innovations and key technologies developed for this instrument, as well as its measured optical performance. We discuss the radiometric calibration characterization, including detector linearity, flat field correction, and SNR. Finally, we discuss stray light modeling and the development of a focused ghost removal algorithm, which is tested and supported by laboratory results.

Bellardo, John↗
NASA NTRS2020

Free-Space Optical Communication for CubeSats in Low Lunar Orbit (LLO)

A fine pointing capability has been developed for laser beam pointing to augment body pointing by CubeSats. An application is made to CubeSats in Low Lunar Orbit (LLO), at 100 km. Body pointing was used by Aerospace Corporation for CubeSats in LEO in NASA’s Optical Communications and Sensors Demonstration (OCSD) program. Computer simulations of this fine pointing capability have been applied to the OCSD program. With fine pointing, the spot size on the Earth could be reduced by a factor of eight with a reduction in laser output power by a factor of sixty-four, thereby mitigating the thermal load challenge on the CubeSats. The same reductions in spot size and laser output power can be achieved for CubeSats in LLO. The new method uses laser arrays for fine laser beam pointing and does not use moving parts. It combines a lens system and a VCSEL/Photodetector Array. For these electro-optical systems, reaction times to pointing changes and vibrations are on a nanosecond time scale, much faster than those for mechanical systems. Results from computer simulations will be presented.

Laser beam pointing↗
NASA NTRS2014

A CubeSat Asteroid Mission: Design Study and Trade-Offs

There is considerable interest in expanding the applicability of cubesat spacecraft into lightweight, low cost missions beyond Low Earth Orbit. A conceptual design was done for a 6-U cubesat for a technology demonstration to demonstrate use of electric propulsion systems on a small satellite platform. The candidate objective was a mission to be launched on the SLS test launch EM-1 to visit a Near-Earth asteroid. Both asteroid fly-by and asteroid rendezvous missions were analyzed. Propulsion systems analyzed included cold-gas thruster systems, Hall and ion thrusters, incorporating either Xenon or Iodine propellant, and an electrospray thruster. The mission takes advantage of the ability of the SLS launch to place it into an initial trajectory of C3=0. Targeting asteroids that fly close to earth minimizes the propulsion required for fly-by/rendezvous. Due to mass constraints, high specific impulse is required, and volume constraints mean the propellant density was also of great importance to the ability to achieve the required deltaV. This improves the relative usefulness of the electrospray salt, with higher propellant density. In order to minimize high pressure tanks and volatiles, the salt electrospray and iodine ion propulsion systems were the optimum designs for the fly-by and rendezvous missions respectively combined with a thruster gimbal and wheel system For the candidate fly-by mission, with a mission deltaV of about 400 m/s, the mission objectives could be accomplished with a 800s electrospray propulsion system, incorporating a propellant-less cathode and a bellows salt tank. This propulsion system is planned for demonstration on 2015 LEO and 2016 GEO DARPA flights. For the rendezvous mission, at a ΔV of 2000 m/s, the mission could be accomplished with a 50W miniature ion propulsion system running iodine propellant. This propulsion system is not yet demonstrated in space. The conceptual design shows that an asteroid mission is possible using a cubesat platform with high-efficiency electric propulsion.

cubesat↗
NASA NTRS2015

Space Optical Communications Using Laser Beam Amplification

The Space Optical Communications Using Laser Beam Amplification (SOCLBA) project will provide a capability to amplify a laser beam that is received in a modulating retro-reflector (MRR) located in a satellite in low Earth orbit. It will also improve the pointing procedure between Earth and spacecraft terminals. The technology uses laser arrays to strengthen the reflected laser beam from the spacecraft. The results of first year's work (2014) show amplification factors of 60 times the power of the signal beam. MMRs are mirrors that reflect light beams back to the source. In space optical communications, a high-powered laser interrogator beam is directed from the ground to a satellite. Within the satellite, the beam is redirected back to ground using the MMR. In the MMR, the beam passes through modulators, which encode a data signal onto the returning beam. MMRs can be used in small spacecraft for optical communications. The SOCLBA project is significant to NASA and small spacecraft due to its application to CubeSats for optical data transmission to ground stations, as well as possible application to spacecraft for optical data transmission.

Laser↗
NASA NTRSDate not supplied

Hydrogen Peroxide-Water-Ethanol-Monopropellant Blend for CubeSat Propulsion

Monopropellant propulsion that provides high performance and low hazards compared to state-of-art hydrazine are being investigated under NASA, DOD and commercially sponsored green propellant research and development programs. This NASAeffort looks at phosphate stabilized 50% hydrogen peroxide/water –ethanol blends to evaluate their stability and combustion behavior in submillimeter diameter microtube combustion apparatus for cubesat propulsion applications. Exploratory 7-day stabilitytests find no pressure rise in a sealed test tube environment at room temperature indicating at least short term stability of the mixture. Temperature rise during combustion testing in resistively heated nickel microtube apparatus shows that micro-combustion is achieved in the tube. Propellant density is approximately 1.2 g/cm3compared to 1.0 g/cm3for hydrazine. Theoretical specific impulse of the stoichiometric blend is 219 s at a chamber pressure of 20 psia and an area ratio of 30:1. Density-specific impulseof this blendis263 swhile hydrazine is only about 215-220 sin small reaction control system thrusters. This provides about 20-22% more density-specific impulse for the cubesat compared to hydrazine. The theoreticalcombustion temperature is2168°F, whichis thermally compatible with stainless steel materials of constructionand the nickel catalystused in the tests. Combustion test results are given for stoichiometric, as well as fuel-leanand fuel-richpropellant formulations. Thesetests focus on thermally and catalytically supported combustionin the microtube, whichrequires a microtube preheat temperature of about 900°F and sufficient residence time to achieve combustion that is stable and anchored in the tubewithout blowing through the tube unreacted. Thrust class is expected to be in the 10 mN range based on flow rate and estimated specific impulse.Power requirements are generally below 25 Wto heat the tube. Steady state combustion testing is typically conducted for 3 minutes and shows littledegradation of the hardware over multiple test cyclesindicating thatthere is little catalyst degradation over time and littlecatalyst deactivation due tothe phosphate stabilizer inthe hydrogen peroxide.

Monopropellant↗
NASA NTRS2014

Integration of a MicroCAT Propulsion System and a PhoneSat Bus into a 1.5U CubeSat

NASA Ames Research Center and the George Washington University have developed an electric propulsion subsystem that can be integrated into the PhoneSat bus. Experimental tests have shown a reliable performance by firing three different thrusters at various frequencies in vacuum conditions. The three thrusters were controlled by a SmartPhone that was running the PhoneSat software. The subsystem is fully operational and it requires low average power to function (about 0.1 W). The interface consists of a microcontroller that sends a trigger pulses to the PPU (Plasma Processing Unit), which is responsible for the thruster operation. Frequencies ranging from 1 to 50Hz have been tested, showing a strong flexibility. A SmartPhone acts as the main user interface for the selection of commands that control the entire system. The micro cathode arc thruster MicroCAT provides a high 1(sub sp) of 3000s that allows a 4kg satellite to obtain a (delta)V of 300m/s. The system mass is only 200g with a total of volume of 200(cu cm). The propellant is based on a solid cylinder made of Titanium, which is the cathode at the same time. This simplicity in the design avoids miniaturization and manufacturing problems. The characteristics of this thruster allow an array of MicroCATs to perform attitude control and orbital correcton maneuvers that will open the door for the implementation of an extensive collection of new mission concepts and space applications for CubeSats. NASA Ames is currently working on the integration of the system to fit the thrusters and PPU inside a 1.5U CubeSat together with the PhoneSat bus into a 1.5U CubeSat. This satellite is intended to be deployed from the ISS in 2015 and test the functionality of the thrusters by spinning the satellite around its long axis and measure the rotational speed with the phone byros. This test flight will raise the TRL of the propulsion system from 5 to 7 and will be a first test for further CubeSats with propulsion systems, a key subsystem for long duration or interplanetary CubeSat missions.

PhoneSat↗
NASA NTRSDate not supplied

Hydrogen Peroxide-Water-Ethanol Monopropellant Blend for CubeSat Propulsion

Monopropellant propulsion that provides high performance and low hazards compared to state-of-art hydrazine are being investigated under NASA, DOD and commercially sponsored green propellant research and development programs. This NASA effort looks at phosphate stabilized 50% hydrogen peroxide/water –ethanol blends to evaluate their stability and combustion behavior in submillimeter diameter microtube combustion apparatus for cubesat propulsion applications. Exploratory 7-day stability tests find no pressure rise in a sealed test tube environment at room temperature indicating at least short term stability of the mixture. Temperature rise during combustion testing in resistively heated nickel microtube apparatus shows that micro-combustion is achieved in the tube. Propellant density is approximately 1.2 g/cu.cm compared to 1.0 g/cu.cm for hydrazine. Theoretical specific impulse of the stoichiometric blend is 219 s at a chamber pressure of 20 psia and an area ratio of 30:1. Density-specific impulse of this blend is 263 s while hydrazine is only about 215-220 s in small reaction control system thrusters. This provides about 20-22% more density-specific impulse for the cube sat compared to hydrazine. The the oretical combustion temperature is 2168°F, which is thermally compatible with stainless steel materials of construction and the nickel catalyst used in the tests. Combustion test results are given for stoichiometric, as well as fuel-lean and fuel-rich propellant formulations. These tests focus on thermally and catalytically supported combustion in the microtube, which requires a microtube preheat temperature of about 900°F and sufficient residence time to achieve combustion that is stable and anchored in the tube without blowing through the tube unreacted. Thrust class is expected to be in the 10 m N range based on flow rate and estimated specific impulse. Power requirements are generally below 25 W to heat the tube. Steady state combustion testing is typically conducted for 3 minutes and shows little degradation of the hardware over multiple test cycles indicating that there is little catalyst degradation over time and little catalyst deactivation due to the phosphate stabilizer in the hydrogen peroxide.

Monopropellant↗
NASA NTRS2014

Remote Recession Sensing of Ablative Heat Shield Materials

Material recession and charring are two major processes determining the performance of ablative heat shield materials. Even in ground testing, the characterization of these two mechanisms relies on measurements of material thickness before and after testing, thus providing only information integrated over the test time. For recession measurements, optical methods such as imaging the sample surface during testing are under investigation but require high alignment and instrument effort, therefore being not established as a standard measurement method. For char depth measurements, the most common method so far consists in investigation of sectioned samples after testing or in the case of Stardust where core extractions were performed to determine char information. In flight, no reliable recession measurements are available, except total recession after recovering the heat shield on ground. Developments of mechanical recession sensors have been started but require substantial on board instrumentation adding mass and complexity. In this work, preliminary experiments to evaluate the feasibility of remote sensing of material recession and possibly char depth through optically observing the emission signatures of seeding materials in the post shock plasma is investigated. It is shown that this method can provide time resolved recession measurements without the necessity of accurate alignment procedures of the optical set-up and without any instrumentation on board of a spacecraft. Furthermore, recession data can be obtained without recovering flight hardware which would be a huge benefit for inexpensive heat shield material testing on board of small re-entry probes, e.g. on new micro-satellite re-entry probes as a possible future application of Cubesats or RBR

instrumentation↗
NASA NTRS2014

Micro Cathode Arc Thruster for PhoneSat: Development and Potential Applications

NASA Ames Research Center and the George Washington University are developing an electric propulsion subsystem that will be integrated into the PhoneSat bus. Experimental tests have shown a reliable performance by firing three different thrusters at various frequencies in vacuum conditions. The interface consists of a microcontroller that sends a trigger pulse to the Pulsed Plasma Unit that is responsible for the thruster operation. A Smartphone is utilized as the main user interface for the selection of commands that control the entire system. The propellant, which is the cathode itself, is a solid cylinder made of Titanium. This simplicity in the design avoids miniaturization and manufacturing problems. The characteristics of this thruster allow an array of μCATs to perform attitude control and orbital correction maneuvers that will open the door for the implementation of an extensive collection of new mission concepts and space applications for CubeSats. NASA Ames is currently working on the integration of the system to fit the thrusters and the PPU inside a 1.5U CubeSat together with the PhoneSat bus. This satellite is intended to be deployed from the ISS in 2015 and test the functionality of the thrusters by spinning the satellite around its long axis and measure the rotational speed with the phone gyros. This test flight will raise the TRL of the propulsion system from 5 to 7 and will be a first test for further CubeSats with propulsion systems, a key subsystem for long duration or interplanetary small satellite missions.

CUBESAT↗
NASA NTRS2019

Overview of Electric Propulsion Projects at NASA Glenn Research Center

NASA Glenn Research Center (GRC) is currently leading the development of multiple electric propulsion systems to flight readiness. The Advanced Electric Propulsion System is a 12.5 kW Hall thruster system that is being developed by the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM) project, under the sponsorship of the Space Technology Mission Directorate. NASA's Evolutionary Xenon Thruster-Commercial (NEXT-C) is 7 kW class gridded ion thruster system that being developed under the sponsorship of the Science Mission Directorate. NASA GRC is also providing electric propulsion discipline support to the Power and Propulsion Element and the Double Asteroid Redirection Test (DART) missions, which will be the first applications for these technologies, respectively. Lower technology readiness level (TRL) projects are underway for applications including CubeSats, small spacecraft and Mars exploration vehicles. Under the sponsorship of the Small Spacecraft Technology Program, NASA GRC has performed numerous independent verification and validation tests of CubeSat class electric propulsion systems in support of a growing number of small US businesses that are developing these systems. Lastly, three technology development efforts focused on 100 kW EP strings led by Aerojet Rocketdyne, Ad Astra and MSNW were recently completed.

Smith, Bryan K.↗
NASA NTRS2021

CubeSat Technology Past and Present: Current State-Of-The-Art Survey

A cube satellite (CubeSat) is an evolving and emerging technology that gives a novice or advanced researcher relatively affordable access to space research experiments and applications. The initial CubeSat standard was created in 1999 by California Polytechnic State University, San Luis Obispo and Stanford University’s Space Systems Development Lab to facilitate direct access to space for university students. This initial CubeSat standard has now been adopted by hundreds of organizations worldwide and includes not only universities, educational institutions, but private firms and government organizations. Dozens of CubeSats have been launched since 2003 and have come from more than 29 states in the United States. The CubeSat standard facilitates frequent and affordable access to space with launch opportunities available on most launch vehicles. CubeSats are a class of research spacecraft called nanosatellites and are built to standard CubeSat Units or U dimensions of 10 by 10 by 10 cm and are formally classified as 1U, 2U, 3U, or 6U in size. Most CubeSats are deployed from a Poly-Picosatellite Orbital Deployer called a P-POD. Partnerships among NASA, U.S. industry, and educational institutions are being formed to build upon existing successful CubeSat initiatives with a goal to expand and include launching 50 small satellites from 50 states within the next several years. An extensive and detailed literature review that includes over 830 citations has been conducted to provide a comprehensive resource on both NASA and non-NASA CubeSat experiments and applications that can serve as a guide for background information on CubeSats as well as a valuable resource of lessons learned from CubeSats that have been launched in the past. CubeSats are currently being launched from all over the world on different launch vehicle platforms. Some organizations providing launch opportunities are California Polytechnic State University (http://www.cubesat.org/contactus), ISISPACE Group (https://www.isispace.nl/), Nanoracks (http://nanoracks.com/), Spaceflight Industries, Inc. (http://spaceflight.com/), TriSept Corporation (https://trisept.com/), and Tyvak Nano-Satellite Systems, Inc. (https://www.tyvak.com/).

Cubesat↗
NASA NTRS2015

Performance Comparisons and Down Selection of Small Motors for Two-Blade Heliogyro Solar Sail 6U CubeSat

This report compiles a review of 130 commercial small scale motors (piezoelectric and electric motors) and almost 20 researched-type small scale piezoelectricmotors for potential use in a 2 blades Heliogyro Solar Sail 6U CubeSat. In this application, a motor and gearhead (drive system) will deploy a roll of solar sailthin film (2 um thick)accommodated in a 2U CubeSat (100 x 200 x 100 mm) housing. The application requirements are: space rated, output torque at fulldeployment of 0.8 Nm, reel speed of 3 rpm, drive system weight limited to 150 grams, diameter limited to 50 mm, and the length not to exceed 40 mm. The 50mm diameter limit was imposed as motors with larger diameters would likely weigh too much and use more space on the satellite wall. This would limit theamount of the payload. The motors performance are compared between small scale, volume within 3x102 cm3 (3x105 mm3), commercial electric DC motors,commercial piezoelectric motors, and researched-type (non-commercial) piezoelectric motors extracted from scientific and product literature. The comparisonssuggest that piezoelectric motors without a gearhead exhibit larger output torque with respect to their volume and weight and require less input power toproduce high torque. A commercially available electric motor plus a gearhead was chosen through a proposed selection process to meet the applications designrequirements.

Wiwattananon, Peerawan↗
NASA NTRS2018

Technology and Opportunities of Photon Sieve CubeSat with Deployable Optical Membrane

A photon sieve (PS) is a revolutionary optical instrument that provides high resolution imaging at a fraction of the weight of typical telescopes, with an areal density of 0.3 kg/m2 compared to 25 kg/m2 for the James Webb Space Telescope. The photon sieve is a variation of a Fresnel Zone Plate consisting of many small holes arranged in a series of concentric rings. The sieve works by diffracting light of a certain wavelength to a specified focal point for imaging, so that only a specific wavelength can be imaged. Moreover, the better image contrast and higher signal-to-noise ratios come from suppressing higher diffracted orders by apodizing the number of pinholes in the outer rings. Finally, a photon sieve requires less supports and can withstand more deformation without a reduction in the imaging qualities. Due to these properties, various groups have created PS CubeSats for Earth and Sun imaging at a low cost and weight specifically using deployable technology. A team at the Air Force Research Laboratory created a design and prototype of a mechanism that deploys a 20 cm diameter photon sieve. The United States Air Force Laboratory used a similar design to create a CubeSat-based deployable photon sieve. The team at NASA Langley Research Center has researched photon sieves for conducting an Earth-observing experiment using LIDAR (Light Detection and Ranging) with a higher signal-to-noise ratio benefit from the PS. This paper provides a state of the art overview on existing PS CubeSat technology with deployable structures and applications. Especially, the paper introduces PS for LIDAR applications and discusses the CubeSat-based PS challenge being worked at the NASA Langley Research Center.

Hyun Jung Kim↗
NASA NTRS2014

Laser Photonic Propulsion Force for Station-Keeping Applications

Small satellites, e.g. cubesats, do not tend to incorporate propulsion subsystems that can compensate for perturbation forces, which causes orbital decay. Cubesats are especially susceptible to the phenomenon of orbital decay, which limits their potential performance, since these effects are more noticeable in Low Earth Orbit (LEO). We postulate that a network of ground-based lasers could extend the operational lifetimes of these satellites by applying a photonic force onto their surfaces. This boosting force would help to counteract the degrading force, which is mainly produced by the drag of the atmosphere. This solution may present an advantage for low cost missions, in that it would enable longer mission durations without the need to incorporate a propulsion system, which comprises a large part of the mass budget and the power constraints of a satellite. This poster presents an analysis of the trade space for both the required network of laser ground stations and the satellite orbits. The analysis is based on simulations of the orbital decay of model satellites.

laser↗
NASA NTRS2018

ASTERIA Operations Demonstrates the Value of Combining the Mission Assurance and Fault Protection Roles on CubeSats

On November 20, 2017, ASTERIA (Arcsecond Space Telescope Enabling Research in Astrophysics), a 6U CubeSat performing a technology demonstration of astrophysical measurements, deployed from the ISS. The technology demonstration goals to achieve precision photometry via arcsecond-level line-of-sight pointing error and highly stable focal plane temperature control were met by February 2018. Extended mission operations are ongoing, with the primary focus on observing nearby stars for transiting exoplanets. Throughout development and operations, the roles of mission assurance and fault protection have proven critical to achieving the primary technical goals and to maintaining a healthy spacecraft through multiple extended missions. Given the budget and schedule constraints typical of a CubeSat, innovative tailoring of processes has been critical to success throughout both development and operations of ASTERIA. Mission assurance plays an important role in identifying and evaluating risk and developing cost-effective mitigations. Flexibility in the fault protection design offers a variety of options for implementing risk mitigations as risks have been uncovered both in pre-delivery testing and in mission operations. This paper will discuss the approach taken on ASTERIA to implement mission assurance and fault protection and the resulting benefits to operational efficiency and success. It will briefly address the advantages of this approach during development, in which the combination of the roles provided mission assurance significant insight to system risks, which feeds back into testing methodologies and directly into fault protection design. Operations will be discussed in detail. During this phase, the roles merge to identify in-flight fault protection updates to efficiently respond to anomalies and improve the likelihood of successful technology demonstrations. The paper will also detail the tools that are used to analyse data, identify anomalies, and develop the updates to uplink to the spacecraft. Finally, the general operational approach will be discussed to highlight the usefulness of the ASTERIA processes and their applicability to future CubeSat missions.

Knapp, Mary↗
NASA NTRS2021

A Spectrum Sensor for CubeSat Radios

Cube satellite (CubeSat) launches have increased exponentially over the last 20 years. This class of miniature spacecraft is well-suited for a set of nonconventional satellite architectures collectively known as formation flying. With the exponential pace of launches expected to continue, the prospect of spectrum management for these complex formations arises. In previous work, investigators focus on terrestrial applications of spectrum sensing, which have the luxury to utilize hardware with high size, weight, and power (SWaP) resources. In this work, we develop and test a spectrum sensor for CubeSat radio applications. Given that CubeSat radios are inherently designed for low SWaP, they cannot implement the computationally expensive spectral correlation analyzer (SCA) algorithms for signal detection. To that end, our investigation focuses on the application of the SCA to square-root-raised-cosine (SRRC) pulse-shaped quadrature amplitude modulation (QAM) waveforms using a field-programmable gate array (FPGA). This model requires no prior knowledge of the radio-frequency (RF) channel. We show that this model can consistently and accurately detect the symbol rate and center frequencies of waveforms located in a spectrum.

cubesat↗
NASA NTRS2023

The IAU Recommended Photometric System for Ultraviolet Astronomy

In the current era when access to space is becoming easier and at a lower cost thanks to the standardised cubesat technology, numerous missions are expected to be launched to observe, particularly, at ultraviolet wavelengths. Given the reduced dimensions of the telescope that a cubesat can carry, most of these missions will be focused on photometric surveys of a reduced sample of targets of interest, and therefore each mission will define their own photometric bands according to their scientific objectives and orbital constraints. However, in order to provide a coherent view of the ultraviolet sky, the data should be post-processed under a common framework. In 2017, the IAU working group on ultraviolet astronomy identified the need to define such a common framework for the upcoming ultraviolet missions, and coordinated the definition of a standard set of photometric bands that could serve for homogenizing the current and future data. This paper presents the procedure adopted by the working group for the definition of the standard photometric system, that was approved by the IAU during the General Assembly Business Sessions held in August, 2021. The photometric system consists of seven bands, denoted as UV1-UV7, all included in the range 115 - 400 nm. Some of these bands are based on existing filters, while others have been defined as theoretical bands with constant throughput. This system is to be regarded as a set of synthetic bands for post-processing the data of any mission, and an example of its application to the SPARCS cubesat is also included. The photometric bands are publicly available and can be downloaded from https://www.nuva.eu/uv-photometry/.

Ultraviolet: ISM↗