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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 NTRS2017

The Lightweight Integrated Solar Array and anTenna (LISA-T) Big Power for Small Spacecraft

NASA is developing a space power system using lightweight, flexible photovoltaic devices originally developed for use here on Earth to provide low cost power for spacecraft. The Lightweight Integrated Solar Array and anTenna (LISA-T) is a launch stowed, orbit deployed array on which thin-film photovoltaic and antenna elements are embedded. The LISA-T system is deployable, building upon NASA's expertise in developing thin-film deployable solar sails such the one being developed for the Near Earth Asteroid Scout project which will fly in 2018. One of the biggest challenges for the NEA Scout, and most other spacecraft, is power. There simply isn't enough of it available, thus limiting the range of operation of the spacecraft from the Sun (due to the small surface area available for using solar cells), the range of operation from the Earth (low available power with inherently small antenna sizes tightly constrain the bandwidth for communication), and the science (you can only power so many instruments with limited power). The LISA-T has the potential to mitigate each of these limitations, especially for small spacecraft. Inherently, small satellites are limited in surface area, volume, and mass allocation; driving competition between their need for power and robust communications with the requirements of the science or engineering payload they are developed to fly. LISA-T is addressing this issue, deploying large-area arrays from a reduced volume and mass envelope - greatly enhancing power generation and communications capabilities of small spacecraft and CubeSats. The problem is that these CubeSats can usually only generate between 7W and 50W of power. The power that can be generated by the LISA-T ranges from tens of watts to several hundred watts, at a much higher mass and stowage efficiency. A matrix of options are in development, including planar (pointed) and omnidirectional (non-pointed) arrays. The former is seeking the highest performance possible while the latter is seeking GN&C simplicity. Options for leveraging both high performance, 'typical cost' triple junction thin-film solar cells as well as moderate performance, low cost cells are being developed. Alongside, UHF (ultrahigh frequency), S-band, and X-band antennas are being integrated into the array to move their space claim away from the spacecraft and open the door for more capable multi-element antenna designs such as those needed for spherical coverage and electronically steered phase arrays.

Space Exploration↗
NASA NTRS2010

Applications of Multi-Body Dynamical Environments: The ARTEMIS Transfer Trajectory Design

The application of forces in multi-body dynamical environments to pennit the transfer of spacecraft from Earth orbit to Sun-Earth weak stability regions and then return to the Earth-Moon libration (L1 and L2) orbits has been successfully accomplished for the first time. This demonstrated transfer is a positive step in the realization of a design process that can be used to transfer spacecraft with minimal Delta-V expenditures. Initialized using gravity assists to overcome fuel constraints; the ARTEMIS trajectory design has successfully placed two spacecraft into EarthMoon libration orbits by means of these applications.

Folta, David C.↗
NASA NTRSDate not supplied

Ultraviolet/Visible Reflectance Spectra of Returned Samples of (162173) Ryugu Unexposed to Earth’s Atmosphere

Following the successful return of JAXA’s Hayabusa2 sample capsule, the material from asteroid (162173) Ryugu retrieved from the spacecraft retrieval container has been the subject of multi-pronged analyses [e.g., 1]. We undertook an effort to acquire reflectance spectra covering 200- to- 1100 nm of particles from Ryugu. This region probes the near-ultraviolet (NUV) to acquire reflectance spectra that cover the 250-nm Fresnel peak that can result from increased graphitization (reduced H content) of carbon-bearing materials. Laboratory tests show that increased graphitization could result from the effects of increased temperature, an effect of space weathering of materials closer to the Sun. Some ground-based reflectance spectra and space-based photometry of Ryugu show that reflectance is increasing with decreasing wavelength, suggesting the beginning of a UV upturn [2]. We sought evidence for space weathering (graphitization) of the Ryugu material at NUV wavelengths, by acquiring reflectance spectra of surface samples.

F Vilas↗
NASA NTRS2001

Optimizing Orbit-Instrument Configuration for Global Precipitation Mission (GPM) Satellite Fleet

Following the scientific success of the Tropical Rainfall Measuring Mission (TRMM) spearheaded by a group of NASA and NASDA scientists, their external scientific collaborators, and additional investigators within the European Union's TRMM Research Program (EUROTRMM), there has been substantial progress towards the development of a new internationally organized, global scale, and satellite-based precipitation measuring mission. The highlights of this newly developing mission are a greatly expanded scope of measuring capability and a more diversified set of science objectives. The mission is called the Global Precipitation Mission (GPM). Notionally, GPM will be a constellation-type mission involving a fleet of nine satellites. In this fleet, one member is referred to as the "core" spacecraft flown in an approximately 70 degree inclined non-sun-synchronous orbit, somewhat similar to TRMM in that it carries both a multi-channel polarized passive microwave radiometer (PMW) and a radar system, but in this case it will be a dual frequency Ku-Ka band radar system enabling explicit measurements of microphysical DSD properties. The remainder of fleet members are eight orbit-synchronized, sun-synchronous "constellation" spacecraft each carrying some type of multi-channel PMW radiometer, enabling no worse than 3-hour diurnal sampling over the entire globe. In this configuration the "core" spacecraft serves as a high quality reference platform for training and calibrating the PMW rain retrieval algorithms used with the "constellation" radiometers. Within NASA, GPM has advanced to the pre-formulation phase which has enabled the initiation of a set of science and technology studies which will help lead to the final mission design some time in the 2003 period. This presentation first provides an overview of the notional GPM program and mission design, including its organizational and programmatic concepts, scientific agenda, expected instrument package, and basic flight architecture. Following this introduction, we focus specifically on the last topic, that being an analysis which leads to an optimal flight architecture dictated in part by science requirements but constrained by allowable orbital mechanics, instrument scan patterns, and antenna aperture properties. Because the optimal architecture involves an interplay between orbit mechanics and instrument specifications, it is important to recognize that in attempting to serve various scientific themes, the final optimal architecture will represent a compromise concerning dynamic range, spatial resolution, sampling interval, pointing, beam coincidence, and measurement uncertainty. Moreover, cost becomes a major factor in seeking the optimal architecture through the pathways of antenna and instrument scan designs, as well as propulsion requirements associated with the orbit heights of various "constellation" members. Although the results presented at the IGARSS-2001 meeting will likely not be the fully refined flight architecture specifications, they are expected to be nearly complete.

Smith, Eric A.↗
NASA NTRS2018

Multi-Scale Modeling of Low-Density Carbon-Phenolic Ablators

Protecting a spacecraft during atmospheric entry is one of highest risk factors that needs to be mitigated during design of a space exploration mission. At entry speeds from space, air turns into high-temperature plasma, and spacecraft Thermal Protection Systems (TPS) are needed to protect the vehicle payload. Modern successful material architectures of spacecraft shields use a porous carbon fiber substrate impregnated with phenolic as an ablator material. In the lecture, efforts to build a Predictive Material Modeling framework for porous ablators from micro-scale to macro-scale will be presented. Several numerical methods and techniques will be summarized that use voxelized images to compute geometrical properties of the porous substrate. These computed properties include porosity, specific surface area and tortuosity that are otherwise indirectly measured through experimental techniques. Direct simulation Monte Carlo (DSMC), a particle-based method for approximating the Boltzmann equation, is used to compute the permeability coefficient of the porous substrate based on its digitized representation. The method computes the flow within the microstructure, where the size of the pores may approach the mean-free-path of the flow. Finally, a high-fidelity model implemented in PATO (Porous-material Analysis Toolbox) is discussed, and some examples of ablative material response are presented, including for the first time 3D simulations of the full tiled heat shield for the Mars Science Laboratory (MSL) capsule.

Ablators↗
NASA NTRS2011

Verification and Validation Plan for Flight Performance Requirements on the CEV Parachute Assembly System

The Crew Exploration Vehicle Parachute Assembly System (CPAS) is engaged in a multi-year design and test campaign aimed at qualifying a parachute recovery system for human use on the Orion Spacecraft. Orion has parachute flight performance requirements that will ultimately be verified through the use of Monte Carlo multi-degree of freedom flight simulations. These simulations will be anchored by real world flight test data and iteratively improved to provide a closer approximation to the real physics observed in the inherently chaotic inflation and steady state flight of the CPAS parachutes. This paper will examine the processes necessary to verify the flight performance requirements of the human rated spacecraft. The focus will be on the requirements verification and model validation planned on CPAS.

Morris, Aaron L.↗
NASA NTRS1987

A demonstration of high precision GPS orbit determination for geodetic applications

High precision orbit determination of Global Positioning System (GPS) satellites is a key requirement for GPS-based precise geodetic measurements and precise low-earth orbiter tracking, currently under study at JPL. Different strategies for orbit determination have been explored at JPL with data from a 1985 GPS field experiment. The most successful strategy uses multi-day arcs for orbit determination and includes fine tuning of spacecraft solar pressure coefficients and station zenith tropospheric delays using the GPS data. Average rms orbit repeatability values for 5 of the GPS satellites are 1.0, 1.2, and 1.7 m in altitude, cross-track, and down-track componenets when two independent 5-day fits are compared. Orbit predictions up to 24 hours outside the multi-day arcs agree within 4 m of independent solutions obtained with well tracked satellites in the prediction interval. Baseline repeatability improves with multi-day as compared to single-day arc orbit solutions. When tropospheric delay fluctuations are modeled with process noise, significant additional improvement in baseline repeatability is achieved. For a 246-km baseline, with 6-day arc solutions for GPS orbits, baseline repeatability is 2 parts in 100 million (0.4-0.6 cm) for east, north, and length components and 8 parts in 100 million for the vertical component. For 1314 and 1509 km baselines with the same orbits, baseline repeatability is 2 parts in 100 million for the north components (2-3 cm) and 4 parts in 100 million or better for east, length, and vertical components.

Lichten, S. M.↗
NASA NTRS2018

The Second European Service Module (ESM-2) Evolutions, Production and Challenges

This paper presents an overview of the Second European Service Module (ESM-2), the second in a series of European Service Modules produced as part of the Barter agreement between NASA and ESA for the Orion Program. The European Industrial consortium is led by the ESA prime contractor Airbus Defence and Space in Bremen. ESA and Airbus signed the ESM-2 contract on 16 February 2017, for this key element of the Orion Exploration Mission 2 (EM-2). EM-2 is the first crewed mission for Orion and will take astronauts farther into the solar system than humanity has ever travelled. EM-2 will also be a historic mission for Europe, as the ESM-2 will be the first European spacecraft to be part of a human transportation system carrying humans beyond low Earth orbit. ESM-2 is mainly a recurring production following ESM-1. Nevertheless, there are a number of important changes being implemented, for example, to incorporate upgrades to further enhance safety and reliability. The challenging delivery schedule for ESM-2 has driven the need to commence manufacturing prior to completion of the qualification on ESM-1. In addition, some requirement deviations and non-compliances approved for ESM-1 have resulted in modifications for ESM-2. In order to manage the competing constraints effectively, the ESM-2 Team has put in place a number of novel approaches to manage schedule, risk, and technical changes. Airbus has set up multi-functional teams according to an approach known as "Major Spacecraft Deliveries" consisting of quality assurance, engineering and procurement. The risk of starting manufacturing prior to qualification is managed through a special risk share agreement. This agreement necessitates rigorous risk reviews across the board for all manufacturing, assembly, integration and test milestones. The ESM-2 changes are managed by Configuration Management, but Airbus has also introduced the Technical Baseline Matrix to provide a transparent top-level overview of the changes from ESM-1 to ESM-2. The tool provides the basis for ESM-2 design and development needs, decisions, as well as the input for the Orion EM-2 Critical Design Review (CDR). The main technical evolutions, status of the production and the novel management approaches for ESM-2 are presented and discussed in the paper.

Orion European Service Module Programme↗
NASA NTRS1992

Distributed digital signal processors for multi-body flexible structures

Multi-body flexible structures, such as those currently under investigation in spacecraft design, are large scale (high-order) dimensional systems. Controlling and filtering such structures is a computationally complex problem. This is particularly important when many sensors and actuators are located along the structure and need to be processed in real time. This report summarizes research activity focused on solving the signal processing (that is, information processing) issues of multi-body structures. A distributed architecture is developed in which single loop processors are employed for local filtering and control. By implementing such a philosophy with an embedded controller configuration, a supervising controller may be used to process global data and make global decisions as the local devices are processing local information. A hardware testbed, a position controller system for a servo motor, is employed to illustrate the capabilities of the embedded controller structure. Several filtering and control structures which can be modeled as rational functions can be implemented on the system developed in this research effort. Thus the results of the study provide a support tool for many Control/Structure Interaction (CSI) NASA testbeds such as the Evolutionary model and the nine-bay truss structure.

Lee, Gordon K. F.↗
NASA NTRS2017

Orbit Maintenance and Navigation of Human Spacecraft at Cislunar Near Rectilinear Halo Orbits

Multiple studies have concluded that Earth-Moon libration point orbits are attractive candidates for staging operations. The Near Rectilinear Halo Orbit (NRHO), a member of the Earth-Moon halo orbit family, has been singularly demonstrated to meet multi-mission architectural constraints. In this paper, the challenges associated with operating human spacecraft in the NRHO are evaluated. Navigation accuracies and human vehicle process noise effects are applied to various station keeping strategies in order to obtain a reliable orbit maintenance algorithm. Additionally, the ability to absorb missed burns, construct phasing maneuvers to avoid eclipses and conduct rendezvous and proximity operations are examined.

Davis, Diane↗
NASA NTRS2004

Orbit Optimization For The Geospace Electrodynamics Connections (GEC) Mission

Part of NASA's Solar Terrestrial Probe line of missions, the Geospace Electrodynamics Connections (GEC) mission will deploy a formation of three spacecraft to perform in-situ atmospheric research in the low Ionosphere-Thermosphere region. These spacecraft will fly together in a %tring-of-pearls formation with variable spacings ranging from 10 seconds to one-quarter of an orbit at perigee. Over the course of its two-year mission, the three spacecraft will perform ten, 1-week dipping campaigns whereby they maneuver to lower their perigee to near 134 km. Using available launch vehicle performance data, an optimal parking orbit of 222 x 1525 km was found to maximize the dry mass available while providing enough propellant to perform the ten deep-dipping campaigns over its two-year mission. The results were used to create multi-variable contour plots containing the orbit perigee, the orbit apogee, spacecraft dry mass, propellant mass, and T500 (a science data collection figure of merit that tabulates the cumulative time spent below 500 km). These plots illustrate how the mission can trade off science return relative to the cost in dry mass and propellant. Other optimal solutions such as minimum propellant or maximum T500 were found to either limit the science data collection or to be dry mass limiting, respectively. Sensitivity analyses were performed to find new optimal (maximum dry mass) solutions if the number of campaigns changed, if the coefficient of drag (CD) were different, and if the propellant specific impulse were increased. A surprising result showed that the dry mass and T500 were both increased if the number of campaigns decreased. Changes in CD provided the expected results - raising CD lowered both the dry mass and T500 while lowering CD raised both the dry mass and T500. Increases in the propellant specific impulse had the expected outcome of raising the dry mass and lowering the propellant load but there was no change in the T500 figure of merit. The orbit optimization was performed parametrically using a Matlab(TradeMark) script and validated using FreeFlyer(TradeMark), a commercial orbit analysis tool. ,

Mesarch, Michael A.↗
NASA NTRS1994

Development of an Aerosol Opacity Retrieval Algorithm for Use with Multi-Angle Land Surface Images

In 1998, the Multi-angle Imaging SpectroRadiometer (MISR) will fly aboard the EOS-AM1 spacecraft. MISR will enable unique methods for retrieving the properties of atmospheric aerosols, by providing global imagery of the Earth at nine viewing angles in four visible and near-IR spectral bands. As part of the MISR algorithm development, theoretical methods of analyzing multi-angle, multi-spectral data are being tested using images acquired by the airborne Advanced Solid-State Array Spectroradiometer (ASAS). In this paper we derive a method to be used over land surfaces for retrieving the change in opacity between spectral bands, which can then be used in conjunction with an aerosol model to derive a bound on absolute opacity.

Advanced Solid-State Array Spectroradiometer ASAS ↗
NASA NTRS2015

Deployment Technology of a Heliogyro Solar Sail for Long Duration Propulsion

Interplanetary, multi-mission, station-keeping capabilities will require that a spacecraft employ a highly efficient propulsion-navigation system. The majority of space propulsion systems are fuel-based and require the vehicle to carry and consume fuel as part of the mission. Once the fuel is consumed, the mission is set, thereby limiting the potential capability. Alternatively, a method that derives its acceleration and direction from solar photon pressure using a solar sail would eliminate the requirement of onboard fuel to meet mission objectives. MacNeal theorized that the heliogyro-configured solar sail architecture would be lighter, less complex, cheaper, and less risky to deploy a large sail area versus a masted sail. As sail size increases, the masted sail requires longer booms resulting in increased mass, and chaotic uncontrollable deployment. With a heliogyro, the sail membrane is stowed as a roll of thin film forming a blade when deployed that can extend up to kilometers. Thus, a benefit of using a heliogyro-configured solar sail propulsion technology is the mission scalability as compared to masted versions, which are size constrained. Studies have shown that interplanetary travel is achievable by the heliogyro solar sail concept. Heliogyro solar sail concept also enables multi-mission missions such as sample returns, and supply transportation from Earth to Mars as well as station-keeping missions to provide enhanced warning of solar storm. This paper describes deployment technology being developed at NASA Langley Research Center to deploy and control the center-of-mass/center-of-pressure using a twin bladed heliogyro solar sail 6-unit (6U) CubeSat. The 6U comprises 2x2U blade deployers and 2U for payload. The 2U blade deployers can be mounted to 6U or larger scaled systems to serve as a non-chemical in-space propulsion system. A single solar sail blade length is estimated to be 2.4 km with a total area from two blades of 720 m2; total allowable weight of a 6U CubeSat is approximately 8 kg. This makes the theoretical characteristic acceleration of approximately 0.75 mm/s2 at I AU (astronomical unit), when compared to IKAROS (0.005 mm/s2) and NanoSail-D (0.02 mm/s2).

Peerawan, Wiwattananon↗
NASA NTRS2016

Paving the Path for Human Space Exploration: The Challenges and Opportunities

Lauri Hansen, Director of Engineering at NASA Johnson Space Center will discuss the challenges of human space exploration. The future of human exploration begins with our current earth reliant missions in low earth orbit. These missions utilize the International Space Station to learn how to safely execute deep space missions. In addition to serving as an exploration test bed and enabling world class research, the International Space Station enables NASA to build international and commercial partnerships. NASA's next steps will be to enable the commercialization of low earth orbit while concentrating on developing the spacecraft and infrastructure necessary for deep space exploration and long duration missions. The Orion multi-purpose crew vehicle and the Space Launch System rocket are critical building blocks in this next phase of exploration. There are many challenges in designing spacecraft to perform these missions including safety, complex vehicle design, and mass challenges. Orion development is proceeding well, and includes a significant partnership with the European Space Agency (ESA) to develop and build the Service Module portion of the spacecraft. Together, NASA and ESA will provide the capability to take humans further than we have ever been before - 70,000 km past the moon. This will be the next big step in expanding the frontiers of human exploration, eventually leading to human footprints on Mars.

Hansen, Lauri↗
NASA NTRS2018

Spacecraft Micro-Vibration: A Survey of Problems, Experiences, Potential Solutions, and Some Lessons Learned

Predicting, managing, controlling, and testing spacecraft microvibrations due to on-board internal disturbance sources is a formidable multi-disciplinary systems engineering challenge, especially for those observatories hosting extremely sensitive optical sensor payloads with stringent requirements on allowable line-of-sight jitter. In this paper some specific spacecraft micro-vibration engineering challenges will be introduced and described. Technical background context is provided with the inclusion of several illustrative examples of NASA and European Space Agency missions (both past and present) where dynamic interactions have to be addressed and which have demanding payload instrument line-of-sight jitter requirements. A general modeling, analysis, simulation, and test approach to address and solve the overall problem of spacecraft micro-vibrations is outlined. Recommended rules of thumb are presented to provide guidance for analysts on where to initiate and how to approach a new spacecraft micro-vibration design problem. A set of experience-based spacecraft micro-vibration lessons learned are also presented in the hope they can be leveraged on new system development projects to help overcome unfamiliarity with previously identified microvibration technical pitfalls and challenges.

Dennehy, Cornelius↗
NASA NTRS2015

Coupled Low-thrust Trajectory and System Optimization via Multi-Objective Hybrid Optimal Control

The optimization of low-thrust trajectories is tightly coupled with the spacecraft hardware. Trading trajectory characteristics with system parameters ton identify viable solutions and determine mission sensitivities across discrete hardware configurations is labor intensive. Local independent optimization runs can sample the design space, but a global exploration that resolves the relationships between the system variables across multiple objectives enables a full mapping of the optimal solution space. A multi-objective, hybrid optimal control algorithm is formulated using a multi-objective genetic algorithm as an outer loop systems optimizer around a global trajectory optimizer. The coupled problem is solved simultaneously to generate Pareto-optimal solutions in a single execution. The automated approach is demonstrated on two boulder return missions.

Spacecraft Systems↗
NASA NTRSDate not supplied

Preliminary Experimental Investigation of Multi-Shock Shield Performance Against Meteoritic Other Lithic Projectiles

The protection capabilities of micrometeoroid and orbital debris (MMOD) shields are invariably linked to the physical properties of the impactor. Regardless of the shield type, or bumper material employed, the general principles of purposefully designed multi-walled MMOD protection relies on the ability to disrupt, shock, melt, and/or vaporize the projectile before it interacts with a critical spacecraft component. Here, we begin to study the influence of projectile heterogeneity on the performance of multi-shock shields by conducting hypervelocity impact tests using meteorite and a variety of other rock-like (i.e., lithic) materials. Testing was conducted using a .17 caliber two-stage light-gas gun, launching cylindrical lithic projectiles between 5.66 and 7.06 km/s. Shield layups included three layers of Nextel and one layer of graphite composite as a rear wall, with each layer having a 35 mm standoff. After testing a comparison of rear wall and witness plate damage was made to separate tests conducted using spherical Al projectiles of similar mass. It was observed that the more heterogenous meteorite projectile gave rise to notably more damage to the witness plate than an Al projectile of equivalent mass, while the other, more homogeneous lithic projectiles, did not.

meteorite↗
NASA NTRS2013

The Status of Spacecraft Bus and Platform Technology Development under the NASA ISPT Program

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (electric and chemical), Entry Vehicle Technologies (aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for near-term flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; and aerothermal effect models. Two component technologies being developed with flight infusion in mind are the Advanced Xenon Flow Control System and ultralightweight propellant tank technologies. Future directions for ISPT are technologies that relate to sample return missions and other spacecraft bus technology needs like: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV); and 3) electric propulsion. These technologies are more vehicles and mission-focused, and present a different set of technology development and infusion steps beyond those previously implemented. The Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion and spacecraft bus technologies to a wide variety of mission concepts. These inspace propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

Anderson, David J.↗