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NASA NTRS2019

Crazy Antennas

Everyone here is familiar with traditional antennas, time-honored favorites like dipoles and solid parabolic reflectors. But occasionally, circumstances call for something peculiar. This paper will describe a number of unusual antennas for particular communications scenarios that have been developed at the NASA Glenn Research over the past decade or so. The list includes: a K-band scanning ferroelectric reflectarray; a UHF "Vivaldi" for cellular connectivity to unmanned aerial vehicles; a Ku-band array that develops a top-hat pattern to feed a zone plate antenna; an active antenna that toggles between Iridium and GPS bands; a VHF hybrid spiral/dipole for orientation determination on Venus; and a Ku-band deployable reflector that strongly resembles a giant beach ball. Design strategy and performance results will be included, and trends towards cognitive antennas will be discussed.

Romanofsky, Robert↗
NASA NTRS2018

TES-8: Advanced Exo-Brake, VR and COM Experiments

The TES-8 was jettisoned from the International Space Station on January 31, 2019. As an orbital laboratory and 8th in on-going series, the design makes use of a standard set of interfaces and safety features that permit rapid re-flight. On this flight, an advanced Exo-Brake is flown with de-orbit targeting capability that will engender sample return capability from LEO platforms. A Virtual Reality data recording system uses stereo imaging and efficient data-compression with an NVIDIA GPU (Graphics Processing Unit) to permit compression and transmission of very large data files. An SDR (Software Defined Radio) will download data to the NEN (Near Earth Network) for the first time - demonstrating potential use in cis-lunar space using S-band. For the first time, a comparison will be made regarding the functionality of the Iridium and Globalstar short burst data modems - as essential communication tools for future nano-sat projects. Lastly, the 7 micro-processors and 4 cameras provide an excellent learning platform for university students and NASA young professionals.

Exo-Brake↗
NASA NTRS2019

The NASA Orbital Debris Engineering Model 3.1: Development, Verification, and Validation

The NASA Orbital Debris Program Office has developed the Orbital Debris Engineering Model (ORDEM) primarily as a tool for spacecraft designers and other users to understand the long-term risk of collisions with orbital debris. The newest version, ORDEM 3.1, incorporates the latest and highest fidelity datasets available to build and validate representative orbital debris populations encompassing low Earth orbit (LEO) to geosynchronous orbit (GEO) altitudes for the years 2016-2050. ORDEM 3.1 models fluxes for object sizes > 10 μm within or transiting LEO and > 10 cm in GEO. The deterministic portion of the populations in ORDEM 3.1 is based on the U.S. Space Surveillance Network (SSN) catalog, which provides coverage down to approximately 10 cm in LEO and 1 m in GEO. Observational datasets from radar, in situ, and optical sources provide a foundation from which the model populations are statistically extrapolated to smaller sizes and orbit regions that are not well-covered by the SSN catalog, yet may pose the greatest threat to operational spacecraft. Objects in LEO ranging from approximately 5 mm to 10 cm are modeled using observational data from ground-based radar, namely the Haystack Ultrawideband Satellite Imaging Radar (HUSIR – formerly known as Haystack). The LEO population smaller than approximately 3 mm in size is characterized based on a reanalysis of in situ data from impacts to the windows and radiators of the U.S. Space Transportation System orbiter vehicle, i.e., the Space Shuttle. Data from impacts on the Hubble Space Telescope are also used to validate the sub-millimeter model populations in LEO. Debris in GEO with sizes ranging from 10 cm to 1 m is modeled using optical measurement data from the Michigan Orbital DEbris Survey Telescope (MODEST). Specific, major debris-producing events, including the Fengyun-1C, Iridium 33, and Cosmos 2251 debris clouds, and unique populations, such as sodium-potassium droplets, have been re-examined and are modeled and added to the ORDEM environment separately. The debris environment greater than 1 mm is forecast using NASA’s LEO-to- GEO ENvironment Debris model (LEGEND). Future explosions of intact objects and collisions involving objects greater than 10 cm are assessed statistically, and the NASA Standard Satellite Breakup Model is used to generate fragments from these events. Fragments smaller than 10 cm are further differentiated based on material density categories, i.e., high-, medium-, and low-density, to better characterize the potential debris risk posed to spacecraft. The future projection of the sub-millimeter environment is computed using a special small-particle degradation model where small particles are created from intact spacecraft and rocket bodies. This work discusses the development, features, and capabilities of the ORDEM 3.1 model; the ne new data analyses used to build the model populations; and sample verification and validation results.

Matney, M.↗
NASA NTRSDate not supplied

Autonomous Measurement of Electrochemical Redox Couples for In-Situ Characterization of Enceladus Plume Samples

A 3D-printed microfluidic electrochemical cell with integrated electrodes was developed to demonstrate autonomous in-situ electrochemical characterization of terrestrial seawater samples. The system is designed to extend the capabilities of the 2008 Phoenix Lander’s Wet Chemistry Laboratory (WCL) that analyzed the soluble chemistry of the Martian soil, by miniaturizing its fluidic architecture for analysis of µL-volume samples as expected from an Enceladus fly-by mission. The microfluidic device has a total fluid volume of 48 µL and includes dual channels each populated with 7 electrodes. Measurement of pH was measured using an iridium oxide electrode. Cyclic voltammetry with Au, Pt, and glassy carbon working electrodes was used to measure redox couples in synthetic seawater that contained inorganic and bio-organic molecules. Glassy carbon was found to have minimal risk of interference from constituents of synthetic seawater when measuring molecular organic redox species; however, care must be used to prevent interference due to potential oxidation or dissolution in the case of Au working electrodes. A completely autonomous end-to-end run including sample delivery was performed, to demonstrate the feasibility of microfluidic-based cyclic voltammetry measurements for in-situ characterization of redox couples in Ocean World samples.

Planetary Instruments↗
NASA NTRS2021

Goldstone Orbital Debris Radar: A Historical Review from a Decade of Observations (2007 – 2017)

The NASA Orbital Debris Program Office (ODPO) has utilized the Goldstone bistatic radar since the early1990’s to statistically characterize the low Earth orbit(LEO) sub-centimeter population. The radar provides a unique capability to detect orbital debris population sizes down to an approximate size of 3 mm for altitudes less than 1000 km, and even smaller sizes at lower altitudes. Due to the limited number of hours that are available from this radar each year, it is often used to validate measurements obtained by other radar data sources in the range of orbital debris sizes where the radar systems have an overlap. Additionally, the Goldstone radar provides significantly improved sensitivity than is available from other ground-based radar data sources, which enables it to observe orbital debris populations that, although smaller, still represent a significant risk to both robotic and human missions in LEO. Over the decade of observations from 2007 to2017, a number of significant on-orbit events have occurred including the anti-satellite test againstFengyun-1C, and the accidental collision between Cosmos 2251 and Iridium 33. The orbital debris flux as measured by the Goldstone radar over this important time period is reviewed in this paper, as well as the evolution of the environment over the ensuing years since these events. In addition, the measurements provided by this radar in the years prior to the deployment of large constellations into the environment, as well as the significant increase in missions and mission participants in recent years, is of interest since it provides a baseline for monitoring the effects of these and other changes in the coming years. In 2018, the Goldstone bistatic orbital debris radar pointing geometry changed due to the decommissioning of the nearby receiver antenna, Deep Space Station (DSS) 15, located approximately 500 m from the transmitter antenna. The historical combination of the DSS-14 transmitter and DSS-15receiver enabled all of LEO to be observed using a single pointing due to the close proximity of the two antenna stations and the resulting beam overlap of the bistatic radar geometry. The updated bistatic Goldstone geometry, following the decommissioning of DSS-15,utilizes one of the pair of DSS-25 or DSS-26 antennas that are located approximately 10 km from the transmitter station. The resulting beam overlap with this pointing is such that in LEO, only a small fraction of LEO altitude coverage is observable at one time. Given the operational change of the radar, it is of interest to review the data collected by the original Goldstone bistatic radar over the final decade preceding this change.

Rossina Miller↗
NASA NTRSDate not supplied

Additive Manufacture of Refractory Metal Green Propulsion Thrusters

High temperature refractory metals are required for a number of high temperature propulsion applications. Refractory metals are expensive, difficult to manufacture with high buy-to-fly ratios, and few vendors. Additive manufacture (AM) is used to produce C103, Molybdenum (Mo), and Tungsten (W) reaction chamber and thrust stand-off as well as Iridium ultra-fine lattice catalysts for integration into 1 N green propulsion thrusters. After printing components underwent post-processing to include heat treatment, surface finish modification, inspection, integration, and testing. Properly utilized AM overcomes existing manufacture limitations by greatly increasing design flexibility, material options, decrease price, decrease lead-time, and increase the commercial supply base.

Omar Mireles↗
NASA NTRS2004

NASA/JPL Tumbleweed Polar Rover

The Tumbleweed Rover, currently under development at the Jet Propulsion Laboratory (JPL) in Pasadena, California, is a large, wind-blown, inflated ball, which carries an instrument payload in its interior. Such rovers offer an effective and simple means of gathering data over large spatial extents of Earth, Mars, and other solar system bodies. Tumbleweeds could prove to be a safe and economical way of deploying instruments such as a ground penetrating radar or a magnetometer in numerous hostile environments. The latest version of the rover was recently deployed in Greenland, where it completed a more than 130 km autonomous traverse across an ice sheet. Communicating via the Iridium satellite network, the rover in question successfully and reliably relayed live GPS, temperature, and pressure data to a ground station at JPL for nearly ten days. The follow-on rover is currently being readied for a traverse from the South Pole to the coast of Antarctica some 2000 km away. The Antarctic test is set to take place in February of 2004 and will serve to verify Tumbleweed as an effective means of harvesting data in extreme and remote settings.

Jones, Jack↗
NASA NTRSDate not supplied

TECHEDSAT-7 and 10: The Little Spacecraft That Could

The NOW (Nanosatellite Orbital Workshop) of NASA Ames Research Center (ARC) has two cubesats in orbit at this time: 6 U TechEdSat-10 (T-10) and the 3U TechEdSat-7 (T-7). T10 was jettisoned from the ISS via the NANORACKS system 7/13/2020, and T-7 was launched via Virgin Orbit 1/17/2021. Both were built by the Nano-satellite Orbital Workshop (NOW) at NASA ARC, and designed and fabricated by interns and students in collaboration with educational institutions. Prototyping novel technologies for non-powered re-entry and communications from orbit are primary research interests, however all subsystems including power generation and distribution, subsystem control, navigation, positioning, heat management etc. extend current technologies. Use of distributed processors using open software platforms and standards other based technologies and software is integral to all segments of spacecraft design. Here, we will present an overview of the spacecraft, experiments, and accomplishments – as well as the next three flight experiments. Some of these experiments include: The exo-brake re-entry system is being developed to enable sample return and end of life disposal; Internal communications for sensors, inter-subsystem and experiments uses both a Zigbee based PAN and internal Wi-Fi for high-speed inter-device communications; The Iridium small message LEO system (Short Burst Data) is used to both command the spacecraft and send data to the ground; Experimental use of the Global-Star system for L-band system comparison and back-up; Collaborative NOAA an experiment to communicate from LEO to the GOES geostationary satellite using the DCS (Data Collection System) with on-board Doppler correction; Mars and Lunar experimental communication systems for future cis-lunar and interplanetary nano-satellites; First demonstration of the NASA Near Earth Network systems with nano-satellites at NASA/Wallops Island; Solar array design and implementation for unique future flexible structures; Power distribution using Tardigrade rad-hard processor omni-board (designed by the team); Distributed processors with internal Wi-Fi connectivity; and Initial experiments with AI/Machine Learning.

M Murbach↗
NASA NTRSDate not supplied

Additive Manufacture of Refractory Metals for Aerospace Applications

High temperature refractory metals are required for a number of high temperature propulsion applications. Refractory metals are expensive, difficult to manufacture with high buy-to-fly ratios, and few vendors. Additive manufacture (AM) is used to produce C103, Molybdenum (Mo), and Tungsten (W) reaction chamber and thrust stand-off as well as Iridium ultra-fine lattice catalysts for integration into 1 N green propulsion thrusters. Refractory AM is in development and like traditional AM alloys requires substantial post-processing to include powder heat treatment, surface finish enhancement, inspection, and machining before placed in service. The combination of limited feedstock sources, high temperature processing, oxygen sensitivity, fracture prone nature, and need for elevated temperature mechanical testing limit the number of qualified facilities capable of post-processing AM refractory materials, which add to cost and schedule constraints. However, properly implemented refractory metal AM can overcome existing manufacture limitations by greatly increasing design flexibility, new material options, reduced price, decreased lead-time, and leverage the ever growing AM commercial supply base.

Refractory metal additive manufacture↗
NASA NTRSDate not supplied

Additive Manufacture of Refractory Metals for Aerospace Applications

High temperature refractory metals are required for a number of high temperature propulsion applications. Refractory metals are expensive, difficult to manufacture with high buy-to-fly ratios, and few vendors. Additive manufacture (AM) is used to produce C103, Molybdenum (Mo), and Tungsten (W) reaction chamber and thrust stand-off as well as Iridium ultra-fine lattice catalysts for integration into 1 N green propulsion thrusters. Refractory AM is in development and like traditional AM alloys requires substantial post-processing to include powder heat treatment, surface finish enhancement, inspection, and machining before placed in service. The combination of limited feedstock sources, high temperature processing, oxygen sensitivity, fracture prone nature, and need for elevated temperature mechanical testing limit the number of qualified facilities capable of post-processing AM refractory materials, which add to cost and schedule constraints. However, properly implemented refractory metal AM can overcome existing manufacture limitations by greatly increasing design flexibility, new material options, reduced price, decreased lead-time, and leverage the ever growing AM commercial supply base.

Refractory metal additive manufacture↗
NASA NTRS2021

GPX2

GPX2is a technology demonstration mission that will provide a novel test-bed for Commercial-Off-the-Shelf (COTS) differential global positioning systems (dGPS) to enable future on-orbit assembly, docking, and formation-flying small satellite missions. While on-orbit,GPX2will assess the capability of multi-frequency COTS dGPS receivers.GPX2 is passively stabilized using a 2-meter gravity gradient boom, providing a local orbit horizon view to the GPX2 antennae and Iridium communication. By demonstrating dGPS on-orbit, GPX2 paves the way for on-orbit assembly and autonomous robotic operations using high-accuracy dGPS to measure relative proximity and orientation.

CubeSat↗
NASA NTRS2023

Development of Two High-Energy Bus ‘Cores’ for Rapid Support of Low-TRL and Educational Payloads: A Software-Configured EPS Combined with Flexible C&DH

For several years the TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has relied upon an in-house developed unit to serve both EPS (Electrical Power System) and C&DH (Command and Data Handling) roles along with low data-rate telemetry functions, i.e., serving as the ‘core’ of the spacecraft bus. This ‘core’ has a considerable task given the rapid cadence of the TES program and the typically low-TRL of payloads; configurability and compatibility are key to prevent mission-specific hardware. However, at only 15 watts the current core has become insufficient to support the program’s growing missions and increasingly demanding payloads. To this end, the NOW program is developing new cores to support two TES mission classes: a single-PCB ‘MiniCore’ designed to support 80-watt missions 6U or smaller in LEO, and a three-PCB, radiation-tolerant ‘StackCore’ designed to support 6U and larger missions over 500 watts in LEO and beyond. The ‘MiniCore’ design consists of three main segments: a processor-agnostic C&DH, a software-configured EPS, and a backup low data-rate radio. The design philosophy was to enable rapid-manufacture in a turbulent supply chain, hence the design consists of COTS parts with a focus on those able to be drop-in replaced with radiation-tolerant versions when demanded by the mission. As a single PC-104 sized circuit board, power density and ease of integration also dominated design, demanding the use of modern features such as single-point USB-C for easy charging and monitoring of the spacecraft on the ground. The ‘MiniCore’ can support 80 watts of load, 140 watt-hours of storage, and over 20 watts of optimized solar generation with extensive power monitoring throughout. The ‘MiniCore’ supports one battery pack, six solar-panels, six loads, five actuators, Iridium SBD, and an internal 802.15.4 network. Additionally, the processor-agnostic design can accept any PJRC Teensy 3.x or Adafruit Feather microcontroller unit to enable processor scaling with mission requirements or environment. It is expected a development unit of this design will be completed before conference. The ‘StackCore’ design consists of three stacked PC-104 sized circuit boards: one dedicated to power generation and storage, one dedicated to power distribution, and one dedicated to C&DH tasks. This delineation is necessary to support the transition from highly integrated ICs to discrete analog circuitry, enabling a primarily analog control power system able to operate without software in a radiation environment with finer monitoring compared to the ‘MiniCore’ design. The planned base architecture supports over 500 watts of load, 250 watt-hours of storage, and over 80 watts of optimized solar generation. The power distribution board allows for the use of daughter cards hosting custom converters or interfaces for payloads, in addition to the software-configured supplies used on the ‘MiniCore’. This core stack will be managed by a Vorago ARM M4 microcontroller and support the same wireless communications as the ‘MiniCore’, with optional integration of a NOW S-band radio and attitude determination sensors for ‘black box’ functionality. It is expected the prototype will still be in development during conference.

Spacecraft↗
NASA NTRSDate not supplied

Development of Two High-Energy Bus ‘Cores’ for Rapid Support of Low-TRL and Educational Payloads: A Software-Configured EPS Combined with Flexible C&DH

For several years the TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has relied upon an in-house developed unit to serve both EPS (Electrical Power System) and C&DH (Command and Data Handling) roles along with low data-rate telemetry functions, i.e., serving as the ‘core’ of the spacecraft bus. This ‘core’ has a considerable task given the rapid cadence of the TES program and the typically low-TRL of payloads; configurability and compatibility are key to prevent mission-specific hardware. However, at only 15 watts the current core has become insufficient to support the program’s growing missions and increasingly demanding payloads. To this end, the NOW program is developing new cores to support two TES mission classes: a single-PCB ‘MiniCore’ designed to support 80-watt missions 6U or smaller in LEO, and a three-PCB, radiation-tolerant ‘StackCore’ designed to support 6U and larger missions over 500 watts in LEO and beyond. The ‘MiniCore’ design consists of three main segments: a processor-agnostic C&DH, a software-configured EPS, and a backup low data-rate radio. The design philosophy was to enable rapid-manufacture in a turbulent supply chain, hence the design consists of COTS parts with a focus on those able to be drop-in replaced with radiation-tolerant versions when demanded by the mission. As a single PC-104 sized circuit board, power density and ease of integration also dominated design, demanding the use of modern features such as single-point USB-C for easy charging and monitoring of the spacecraft on the ground. The ‘MiniCore’ can support 80 watts of load, 140 watt-hours of storage, and over 20 watts of optimized solar generation with extensive power monitoring throughout. The ‘MiniCore’ supports one battery pack, six solar-panels, six loads, five actuators, Iridium SBD, and an internal 802.15.4 network. Additionally, the processor-agnostic design can accept any PJRC Teensy 3.x or Adafruit Feather microcontroller unit to enable processor scaling with mission requirements or environment. It is expected a development unit of this design will be completed before conference. The ‘StackCore’ design consists of three stacked PC-104 sized circuit boards: one dedicated to power generation and storage, one dedicated to power distribution, and one dedicated to C&DH tasks. This delineation is necessary to support the transition from highly integrated ICs to discrete analog circuitry, enabling a primarily analog control power system able to operate without software in a radiation environment with finer monitoring compared to the ‘MiniCore’ design. The planned base architecture supports over 500 watts of load, 250 watt-hours of storage, and over 80 watts of optimized solar generation. The power distribution board allows for the use of daughter cards hosting custom converters or interfaces for payloads, in addition to the software-configured supplies used on the ‘MiniCore’. This core stack will be managed by a Vorago ARM M4 microcontroller and support the same wireless communications as the ‘MiniCore’, with optional integration of a NOW S-band radio and attitude determination sensors for ‘black box’ functionality. It is expected the prototype will still be in development during conference.

Spacecraft↗
NASA NTRSDate not supplied

Development of Two High-Energy Bus ‘Cores’

For several years the TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has relied upon an in-house developed unit to serve both EPS (Electrical Power System) and C&DH (Command and Data Handling) roles along with low data-rate telemetry functions, i.e., serving as the ‘core’ of the spacecraft bus. This ‘core’ has a considerable task given the rapid cadence of the TES program and the typically low-TRL of payloads; configurability and compatibility are key to prevent mission-specific hardware. However, at only 15 watts the current core has become insufficient to support the program’s growing missions and increasingly demanding payloads. To this end, the NOW program is developing new cores to support two TES mission classes: a single-PCB ‘MiniCore’ designed to support 80-watt missions 6U or smaller in LEO, and a three-PCB, radiation-tolerant ‘StackCore’ designed to support 6U and larger missions over 500 watts in LEO and beyond. The ‘MiniCore’ design consists of three main segments: a processor-agnostic C&DH, a software-configured EPS, and a backup low data-rate radio. The design philosophy was to enable rapid-manufacture in a turbulent supply chain, hence the design consists of COTS parts with a focus on those able to be drop-in replaced with radiation-tolerant versions when demanded by the mission. As a single PC-104 sized circuit board, power density and ease of integration also dominated design, demanding the use of modern features such as single-point USB-C for easy charging and monitoring of the spacecraft on the ground. The ‘MiniCore’ can support 80 watts of load, 140 watt-hours of storage, and over 20 watts of optimized solar generation with extensive power monitoring throughout. The ‘MiniCore’ supports one battery pack, six solar-panels, six loads, five actuators, Iridium SBD, and an internal 802.15.4 network. Additionally, the processor-agnostic design can accept any PJRC Teensy 3.x or Adafruit Feather microcontroller unit to enable processor scaling with mission requirements or environment. It is expected a development unit of this design will be completed before conference. The ‘StackCore’ design consists of three stacked PC-104 sized circuit boards: one dedicated to power generation and storage, one dedicated to power distribution, and one dedicated to C&DH tasks. This delineation is necessary to support the transition from highly integrated ICs to discrete analog circuitry, enabling a primarily analog control power system able to operate without software in a radiation environment with finer monitoring compared to the ‘MiniCore’ design. The planned base architecture supports over 500 watts of load, 250 watt-hours of storage, and over 80 watts of optimized solar generation. The power distribution board allows for the use of daughter cards hosting custom converters or interfaces for payloads, in addition to the software-configured supplies used on the ‘MiniCore’. This core stack will be managed by a Vorago ARM M4 microcontroller and support the same wireless communications as the ‘MiniCore’, with optional integration of a NOW S-band radio and attitude determination sensors for ‘black box’ functionality. It is expected the prototype will still be in development during conference.

Avery Brock↗
NASA NTRSDate not supplied

Overview of Additively Manufactured TPS Proposed Flight Test and Earth Re-Entry Capsule Design

A flight mechanics overview is presented of an Earth flight test designed to investigate a novel, 3D printed thermal protection system (TPS) that is currently in development at NASA as part of the Additive Manufacturing of Thermal Protection Systems project. The project is pioneering a method to print a thermal protection system onto an entry vehicle forebody one layer at a time. This method reduces labor and complexity as compared to traditional manufacturing methods while increasing mission-dependent customization of through-depth materials properties. The flight test has three objectives. First, subject the forebody stagnation point of a capsule equipped with additively manufactured TPS (AMTPS) material to peak heat fluxes in excess of 100 $W/cm^2$. Second, capture in-flight data to enable flight reconstruction and AMTPS material thermal response model improvement. Third, recover the capsule with data storage and forebody AMTPS intact to enable post-flight inspection and analysis of AMTPS performance. The flight test trajectory is designed to achieve a peak stagnation point, cold-wall, entry heat flux of 135~$W/cm^2$. Flight mechanics simulations are performed using the Program to Optimize Simulated Trajectories II (POST2) and Monte-Carlo analysis yields statistical percentiles on vehicle performance at key points along the trajectory. Based on the flight mechanics analysis presented in this paper, a prototype capsule was designed, partially fabricated, and underwent preliminary component stress testing in preparation for fabrication of the flight unit capsule. The capsule outer mold line is a modified version of the heritage Mars Microprobe geometry. The capsule has a 0.356~m diameter, a 30~kg mass, and a hypersonic ballistic coefficient of 300~$kg/m^2$. Sensor selection is guided by flight dynamics simulations with the goal of resolving the re-entry heating pulse. On-board instrumentation include forebody and aftbody pressure sensors and thermocouples, a 9-axis IMU, a GPS receiver, and an Iridium satellite modem, all of which collect and store data throughout flight via on-board avionics systems. A two-stage parachute system is designed to decelerate the capsule to touchdown velocities that will not result in significant fracture or deformation of the charred AMTPS material at ground impact.

Flight Mechanics↗
NASA NTRSDate not supplied

Tracking and Recovery of the Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) Reentry Vehicle (RV)

The LOFTID mission launched from Vandenberg on Nov 10, 2022, and successfully demonstrated the reentry of a 6m diameter inflatable aeroshell from low Earth orbit. This paper will cover the design features implemented to enable recovery of the flight vehicle, and will discuss the splashdown calculations, in-flight tracking, recovery from the ocean, and post-flight inspection of the flight vehicle. To support recovery of the RV and ejected data recorder after splashdown, a recovery ship was pre-positioned near the predicted splashdown ellipse in the Pacific Ocean. The splashdown ellipse was repeatedly updated as launch approached. In-flight tracking included transmission from the RV of GPS data through both the Iridium satellite network and the LoRa direct RF link, along with IR video cameras on the recovery ship and airborne imagery from the SCIFLI Team. Using both the GPS data and the IR imagery, the recovery ship tracked down the RV, and deployed an inflatable boat to approach the RV and attach it to the ship’s crane, after which the RV was hoisted on board and secured in its GSE recovery stand. The ship then tracked down the ejected data recorder, which was also broadcasting its GPS data, and pulled it from the water. Once the ship returned to port, the RV was hoisted ashore for additional inspection, removal of the data recorders, and repackaging for shipment back to NASA Langley.

Robert A. Dillman↗
NASA NTRSDate not supplied

Tracking and Recovery of the LOFTID RV (Low-Earth Orbit Flight Test of an Inflatable Decelerator Reentry Vehicle)

The LOFTID mission launched from Vandenberg on Nov 10, 2022, and successfully demonstrated the reentry of a 6m diameter inflatable aeroshell from low Earth orbit. This paper will cover the design features implemented to enable recovery of the flight vehicle, and will discuss the splashdown calculations, in-flight tracking, recovery from the ocean, and post-flight inspection of the flight vehicle. To support recovery of the RV and ejected data recorder after splashdown, a recovery ship was pre-positioned near the predicted splashdown ellipse in the Pacific Ocean. The splashdown ellipse was repeatedly updated as launch approached. In-flight tracking included transmission from the RV of GPS data through both the Iridium satellite network and the LoRa direct RF link, along with IR video cameras on the recovery ship and airborne imagery from the SCIFLI Team. Using both the GPS data and the IR imagery, the recovery ship tracked down the RV, and deployed an inflatable boat to approach the RV and attach it to the ship’s crane, after which the RV was hoisted on board and secured in its GSE recovery stand. The ship then tracked down the ejected data recorder, which was also broadcasting its GPS data, and pulled it from the water. Once the ship returned to port, the RV was hoisted ashore for additional inspection, removal of the data recorders, and repackaging for shipment back to NASA Langley.

Robert A Dillman↗
NASA NTRSDate not supplied

Advances in Autonomous Communications and Operations: Tes-N Series

Advances in nanosat subsystems over the past decade have taken the CubeSat standard from a communication-limited educational tool to a powerful platform enabling space research. The potential for autonomous operations may greatly increase the capability of downlinking even larger data sets. This is enabled through miniaturization of software-defined radio/cognitive communication solutions and a rapidly growing number of ground stations and satellite network crosslinks. The TechEdSat-n orbital flight series is currently demonstrating experiments using cognitive communication concepts including User Initiated Service (UIS) and High-rate Delay Tolerant Networks (HDTN), which show a significant step toward improved capability. At the core of this is the use of the Iridium L-band Short Burst Data (SBD) modems, pioneered by TES-n for space applications. SBD enables unique rapid command, control, and scheduling to initiate the UIS and HDTN protocols. This occurs by performing GPS-assisted on-orbit ephemeris determination, enabling negotiation with high bandwidth ground assets to repeatedly downlink over a specific commercial or government-owned ground station. The technique is RF band-agnostic and may be extended to higher bandwidth stations and SDRs. This may also include free-space optical communication, through both laser and omnidirectional LEDs, which can provide an attractive protocol for downlinking very large datasets over far fewer ground stations. In addition, this may be extended to lunar applications for such future concepts as LunaNet, whereby scheduling and cognitive technologies can assist in greatly improving Earth downlink capabilities with ground stations which will see greater competition for usage. Lastly, the NASA Communication Service Program (CSP), intended to eventually replace the NASA Tracking and Data Relay Satellite System (TDRSS) will also demonstrate the feasibility of commercially-provided satellite communication capabilities. All of these combined advances, including large advances in on-board computation on small platforms, will result in more remarkable data processing capability – yielding even more as of yet unknown discoveries.

Autonomy↗