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Near Earth Asteroid Scout: NASA's Solar Sail Mission to a NEA

NASA is developing a solar sail propulsion system for use on the Near Earth Asteroid (NEA) Scout reconnaissance mission and laying the groundwork for their use in future deep space science and exploration missions. Solar sails use sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like sail made of a lightweight, highly reflective material. This continuous photon pressure provides propellantless thrust, allowing for very high Delta V maneuvers on long-duration, deep space exploration. Since reflected light produces thrust, solar sails require no onboard propellant. The Near Earth Asteroid (NEA) Scout mission, funded by NASA's Advanced Exploration Systems Program and managed by NASA MSFC, will use the sail as primary propulsion allowing it to survey and image Asteroid 1991VG and, potentially, other NEA's of interest for possible future human exploration. NEA Scout uses a 6U cubesat (to be provided by NASA's Jet Propulsion Laboratory), an 86 m(exp. 2) solar sail and will weigh less than 12 kilograms. NEA Scout will be launched on the first flight of the Space Launch System in 2018. The solar sail for NEA Scout will be based on the technology developed and flown by the NASA NanoSail-D and The Planetary Society's Lightsail-A. Four approximately 7 m stainless steel booms wrapped on two spools (two overlapping booms per spool) will be motor deployed and pull the sail from its stowed volume. The sail material is an aluminized polyimide approximately 2.5 microns thick. As the technology matures, solar sails will increasingly be used to enable science and exploration missions that are currently impossible or prohibitively expensive using traditional chemical and electric propulsion systems. This paper will summarize the status of the NEA Scout mission and solar sail technology in general.

Johnson, Les↗

Cyclic fatigue analysis of rocket thrust chambers. Volume 2: Attitude control thruster high cycle fatigue

A finite element stress analysis was performed for the film cooled throat section of an attitude control thruster. The anlaysis employed the RETSCP finite element computer program. The analysis included thermal and pressure loads, and the effects of temperature dependent material properties, to determine the strain range corresponding to the thruster operating cycle. The configuration and operating conditions considered, correspond to a flightweight integrated thruster assembly which was thrust pulse tested. The computed strain range was used in conjuction with Haynes 188 Universal Slopes minimum life data to predict throat section fatigue life. The computed number of cycles to failure was greater than the number of pulses to which the thruster was experimentally subjected without failure.

Miller, R. W.↗

NASA Tech Briefs, April 2010

Topics covered include: Active and Passive Hybrid Sensor; Quick-Response Thermal Actuator for Use as a Heat Switch; System for Hydrogen Sensing; Method for Detecting Perlite Compaction in Large Cryogenic Tanks; Using Thin-Film Thermometers as Heaters in Thermal Control Applications; Directional Spherical Cherenkov Detector; AlGaN Ultraviolet Detectors for Dual-Band UV Detection; K-Band Traveling-Wave Tube Amplifier; Simplified Load-Following Control for a Fuel Cell System; Modified Phase-meter for a Heterodyne Laser Interferometer; Loosely Coupled GPS-Aided Inertial Navigation System for Range Safety; Sideband-Separating, Millimeter-Wave Heterodyne Receiver; Coaxial Propellant Injectors With Faceplate Annulus Control; Adaptable Diffraction Gratings With Wavefront Transformation; Optimizing a Laser Process for Making Carbon Nanotubes; Thermogravimetric Analysis of Single-Wall Carbon Nanotubes; Robotic Arm Comprising Two Bending Segments; Magnetostrictive Brake; Low-Friction, Low-Profile, High-Moment Two-Axis Joint; Foil Gas Thrust Bearings for High-Speed Turbomachinery; Miniature Multi-Axis Mechanism for Hand Controllers; Digitally Enhanced Heterodyne Interferometry; Focusing Light Beams To Improve Atomic-Vapor Optical Buffers; Landmark Detection in Orbital Images Using Salience Histograms; Efficient Bit-to-Symbol Likelihood Mappings; Capacity Maximizing Constellations; Natural-Language Parser for PBEM; Policy Process Editor for P(sup 3)BM Software; A Quality System Database; Trajectory Optimization: OTIS 4; and Computer Software Configuration Item-Specific Flight Software Image Transfer Script Generator.

Source record↗

The High Density Vertiplex Advanced Onboard Automation Overview

While many studies have been performed examining Urban Air Mobility (UAM) operations from UAM Maturity Level (UML) UML-1 to UML-4, [1, 2] some uncertainty exists regarding the integration and role of onboard autonomous systems, airspace management systems, ground control and fleet management systems, and how they integrate with vertiport automation systems to ensure safe high-density future operations. One thrust of the Advanced Air Mobility (AAM) High Density Vertiplex (HDV) sub-project is to perform rapid prototyping and assessment of an Urban Air Mobility (UAM) Ecosystem within the terminal operational area to help inform future research investments and technology development. Another thrust within HDV is to perform integration, testing, and safety risk assessments required to acquire operational credit for several NASA small Unmanned Aerial Systems (sUAS) beyond visual line of sight (BVLOS) enabling technologies to expand test capabilities and to expedite technology transfer and ultimate effective usage. Both thrusts leverage sUAS to serve as surrogates for the highly-technologically-similar envisioned UAM aircraft as well as to provide significant contributions to sUAS Part-135 operators. This report provides an overview of the activities accomplished within the Advanced Onboard Automation (AOA) schedule work package of HDV.

Human Factors, Simulation↗

Longitudinal stability and control characteristics of the Quiet Short-Haul Research Aircraft (QSRA)

Flight experiments were conducted to evaluate various aerodynamic characteristics of the Quiet Short-Haul Research Aircraft (QSRA), an experimental aircraft that makes use of the upper-surface blown (USB) powered-lift concept. Time-history records from maneuvers performed with the aircraft in landing-approach and take-off configurations (with its stability augmentation system disengaged) were analyzed to obtain longitudinal stability and control derivatives and performance characteristics. The experiments included measuring the aircraft responses to variations in the deflection of direct-lift control spoilers and to thrust variations as well as to elevator inputs. The majority of the results are given for the aircraft in a landing configuration with the USB flaps at 50 degrees. For this configuration, if the static longitudinal stability is defined as the variation of the pitching-moment coefficient with the lift coefficient at a constant thrust coefficient, this stability decreases significantly with increasing angle of attack above 9 degrees. For this configuration, at small and negative angles of attack and high levels of thrust, the elevators and the horizontal stabilizer lost effectiveness owing to incipent stalling, but this occurred only during unsteady maneuvers and for brief time intervals.

Stephenson, Jack D.↗

Hydrogen film cooling of a small hydrogen-oxygen thrust chamber and its effect on erosion rates of various ablative materials

An experimental investigation was conducted to determine what arrangement of film-coolant-injection orifices should be used to decrease the erosion rates of small, high temperature, high pressure ablative thrust chambers without incurring a large penalty in combustion performance. All of the film cooling was supplied through holes in a ring between the outer row of injector elements and the chamber wall. The best arrangement, which had twice the number of holes as there were outer row injection elements, was also the simplest. The performance penalties, presented as a reduction in characteristic exhaust velocity efficiency, were 0.8 and 2.8 percentage points for the 10 and 20 percent cooling flows, respectively, The best film-coolant injector was then used to obtain erosion rates for 19 ablative materials. The throat erosion rate was reduced by a factor of 2.5 with a 10 percent coolant flow. Only the more expensive silica phenolic materials had low enough erosion rates to be considered for use in the nozzle throat. However, some of the cheaper materials might qualify for use in other areas of small nozzles with large throat diameters where the higher erosion rates are more acceptable.

Hannum, N.↗

Validation of High Aspect Ratio Cooling in a 89 kN (20,000 lb(sub f)) Thrust Combustion Chamber

In order to validate the benefits of high aspect ratio cooling channels in a large scale rocket combustion chamber, a high pressure, 89 kN (20,000 lbf) thrust, contoured combustion chamber was tested in the NASA Lewis Research Center Rocket Engine Test Facility. The combustion chamber was tested at chamber pressures from 5.5 to 11.0 MPa (800-1600 psia). The propellants were gaseous hydrogen and liquid oxygen at a nominal mixture ratio of six, and liquid hydrogen was used as the coolant. The combustion chamber was extensively instrumented with 30 backside skin thermocouples, 9 coolant channel rib thermocouples, and 10 coolant channel pressure taps. A total of 29 thermal cycles, each with one second of steady state combustion, were completed on the chamber. For 25 thermal cycles, the coolant mass flow rate was equal to the fuel mass flow rate. During the remaining four thermal cycles, the coolant mass flow rate was progressively reduced by 5, 6, 11, and 20 percent. Computer analysis agreed with coolant channel rib thermocouples within an average of 9 percent and with coolant channel pressure drops within an average of 20 percent. Hot-gas-side wall temperatures of the chamber showed up to 25 percent reduction, in the throat region, over that of a conventionally cooled combustion chamber. Reducing coolant mass flow yielded a reduction of up to 27 percent of the coolant pressure drop from that of a full flow case, while still maintaining up to a 13 percent reduction in a hot-gas-side wall temperature from that of a conventionally cooled combustion chamber.

Wadel, Mary F.↗

Graphical method for predicting life of a rocket thrust chamber with half-hard zirconium-copper liner and electroformed nickel closeout

A method for estimating the life of a regeneratively cooled rocket thrust chamber was developed and is based on the hot-gas wall temperature and the temperature difference between the hot-gas wall and the outside surface of the closeout. This method permits a quick estimate of the life of a thrust chamber when design changes or test-cycle variations are considered. Strain range and life are presented graphically as functions of these temperature parameters for a typical high-performance rocket thrust chamber with a half-hard zirconium-copper liner and an electroformed nickel closeout.

Kasper, H. J.↗

An experimental investigation of high-aspect-ratio cooling passages

An experimental investigation was conducted to evaluate the effectiveness of using high-aspect-ratio cooling passenges to improve the life and reduce the colant pressure drop in high-pressure rocket thrust chambers. A plug-nozzle rocket-engine test apparatus was used to test two cylindrical chambers with low-aspect-ratio cooling passages and one with high-aspect-ratio cooling passages. The chambers were cyclically tested and data were taken over a wide range of coolant mass flows. The results showed that for the same coolant pressure drop, the hot-gas-side wall temperature of the high-aspect-ratio chamber was 30 percent lower than the baseline low-aspect-ratio chamber, resulting in no fatigue damage to the wall. The coolant pressure drop for the high-aspect-ratio chamber was reduced in increments to one-half that of the baseline chamber, by reducing the coolant mass flow, and still resulted in a reduction in the hot-gas-side wall temperature when compared to the low-aspect-ratio chambers.

Carlile, Julie A.↗

An experimental investigation of high-aspect-ratio cooling passages

An experimental investigation was conducted to evaluate the effectiveness of using high-aspect-ratio cooling passages to improve the life and reduce the coolant pressure drop in high-pressure rocket thrust chambers. A plug-nozzle rocket-engine test apparatus was used to test two cylindrical chambers with low-aspect-ratio cooling passages and one with high-aspect-ratio cooling passages. The chambers were cyclically tested and data were taken over a wide range of coolant mass flows. The results showed that for the same coolant pressure drop, the hot-gas-side wall temperature of the high-aspect-ratio chamber was 30 percent lower than the baseline low-aspect-ratio chamber, resulting in no fatigue damage to the wall. The coolant pressure drop for the high-aspect-ratio chamber was reduced in increments to one-half that of the baseline chamber, by reducing the coolant mass flow, and still resulted in a reduction in the hot-gas-side wall temperature when compared to the low-aspect-ratio chambers.

Carlile, Julie A.↗

Plasma Wall Interaction with 3-D Plasma Boundaries

The interaction of the edge plasma and the material surfaces is one of the most critical challenges on the path to harness fusion power as new, fundamental energy source. This challenge typically combines the thrust to reach high density, low temperature (detached) plasmas in front of the divertor target plates as well as understanding the plasma material interaction (PMI) in particular in this regime. The combination of both research thrusts represents an extraordinarily challenging subject encompassing spatial scales spanning nanometers to meters in all states of matter and across a broad energy range. Modeling capabilities, which help to interpret data from nowadays experiments and enable extrapolation to future devices are urgently required. This is in particular true for toroidal magnetic confinement devices with three-dimensional (3D) plasma boundaries. Such plasma boundary geometries occur in tokamaks, when small amplitude magnetic perturbations are used to stabilize the unruly edge plasma or in stellarators, that are inherently 3D plasma confinement devices. In this project, the impact of 3D plasma boundaries on the plasma material interaction (PMI) was assessed. This work focused on plasma boundary conditions, in which high-density conditions at the material surfaces yield mitigation of the otherwise immense heat and particle loads that these materials would see. These so-called high recycling and eventually detached plasma regimes are of great interest for future reactor operation. In the project, key features that are unique to 3D boundaries were explored in comparison to canonically assumed axisymmetric plasma edge situations in tokamaks. In particular, the relevance of the 3D boundary situation in the extrapolation to the plasma boundary solution at ITER, the next step fusion energy experiment under construction as a multi-national, world-wide large-science experiment in southern France, has been explored. The EMC3-EIRENE plasma edge fluid and kinetic neutral transport code has been advanced to cope with the challenging and unprecedented conditions in the ITER boundary plasma including 3D fields that are planned to be used to suppress harmful edge instabilities, the so-called edge localized modes. This is a vital integration challenge for ITER and the results from this grant have provide a leading capability for this assessment. It was shown that the detachment process in a 3D edge solution for ITER follows the recycling regimes that are known from axisymmetric solutions, but that multiple plasma exhaust channels connected to the material surfaces are established which feature individual recycling characteristics. Because these channels touch the material surfaces in the divertor in a 3D geometry, the compatibility with the plasma material interaction (PMI), including erosion and impurity generation has been found to be an important part of the integration challenge. To address this, the fully 3D plasma material interaction code ERO2 has been adapted to these ITER specific geometries and a homogeneous mixing model was implemented, that allows to consider the mixing of Be and Was used at ITER in the PMI modeling. This model enhancement has been used to study non-local migration of Be in the JET ITER like wall configuration and it has been shown that with this model such complex migration processes in ITER relevant plasma shapes and with ITER relevant plasma boundary conditions can be addressed. The combined modeling approach using EMC3-EIRENE as a plasma boundary transport code and the ERO2 specialized PMI model will be an asset for the continued preparations of ITER operation as well as for Fusion Pilot Plant efforts that have emerged in the U.S. during the evolution of this grant. The predictive capability of this numerical tool has been validated at the DIII-D US national fusion facility. Here, dedicated plasma edge diagnostics were implemented to measure the impurity household around a 3D edge plasma during ELM suppression by 3D fields. Dedicated experiments with local material probes using these diagnostics and the state-of-the-art suite of boundary measurements at DIII-D have shown that the 3D perturbation of the plasma edge that is excreted by such 3D control fields yield a perturbation of the plasma boundary flux structure and hence also of the resulting PMI. The 3D boundary plasma is composed out of helical magnetic flux channels that intersect the divertor targets at an angle relative to the main guiding field, i.e., the toroidal magnetic field component of the tokamak. A similar effect has been measured as well on limiter surfaces during the startup campaign at the new stellarator experiment Wendelstein 7-X. These experiments ad initial analysis with the ERO plasma material interaction model, suggested that the place of erosion for a given particle from the surface and its re-deposition can be different in such 3D field geometries yielding potentially a significant level of net-erosion. This is not the case for axisymmetric solutions, where it was shown in the past that the eroded particles are effectively re-deposited into gaps produced by erosion at the same position and hence the net-erosion levels are small. For ITER, the quest to suppress the ELMs and at the same time maintain the integrity of the divertor is an issue, which these fundamental findings will help to resolve. The coupling of this work to the extrapolation in the ITER program has been addressed by both the PI and the lead numerical scientist being ITER Science Fellows in the duration of the contract and forward. A second focus in the exploration of 3D boundary effects on tokamaks and stellarators has been set on the measurement of helium exhaust features with such 3D fields. This is important because He represents the ash of the fusion process and needs to be exhausted. It was shown that 3D field application compatible with suppression of ELMs yields an increase of the helium exhaust performance. The ratio of the effective helium confinement time over the energy confinement time was reduced by almost 50% which demonstrated that the impact of helium accumulation in the plasma core with respect to the confinement of energy to sustain the fusion reaction is significantly improved with such 3D control fields. It was shown that this is the case for tokamaks as well as stellarators. At the Large helical Device in Japan, a similar enhancement of the helium exhaust features when small amplitude additional 3D fields were applied was measured. This is an important additional function of 3D field application and its impact on ITER is presently being studied in combination with investigations of helium exhaust in 3D field geometries of stellarator devices.

3D plasma edge transport↗

Chemical propulsion technology

An overview of NASA's low thrust liquid chemical propulsion program is presented with particular emphasis on thrust system technology in the ten to one thousand pound thrust range. Key technology issues include high performance of cooled low thrust engines; small cryogenic pumps; multiple starts-shutdowns (10) with slow ramps (approximately 10 seconds); thrust variation - 4/1 in flight and 20/1 between flights; long life (100 hours); improved system weight and size; and propellant selection.

Priem, R. J.↗

Foundational Methane Propulsion Related Technology Efforts, and Challenges for Applications to Human Exploration Beyond Earth Orbit

Current interest in human exploration beyond earth orbit is driving requirements for high performance, long duration space transportation capabilities. Continued advancement in photovoltaic power systems and investments in high performance electric propulsion promise to enable solar electric options for cargo delivery and pre-deployment of operational architecture elements. However, higher thrust options are required for human in-space transportation as well as planetary descent and ascent functions. While high thrust requirements for interplanetary transportation may be provided by chemical or nuclear thermal propulsion systems, planetary descent and ascent systems are limited to chemical solutions due to their higher thrust to weight and potential planetary protection concerns. Liquid hydrogen fueled systems provide high specific impulse, but pose challenges due to low propellant density and the thermal issues of long term propellant storage. Liquid methane fueled propulsion is a promising compromise with lower specific impulse, higher bulk propellant density and compatibility with proposed in-situ propellant production concepts. Additionally, some architecture studies have identified the potential for commonality between interplanetary and descent/ascent propulsion solutions using liquid methane (LCH4) and liquid oxygen (LOX) propellants. These commonalities may lead to reduced overall development costs and more affordable exploration architectures. With this increased interest, it is critical to understand the current state of LOX/LCH4 propulsion technology and the remaining challenges to its application to beyond earth orbit human exploration. This paper provides a survey of NASA's past and current methane propulsion related technology efforts, assesses the accomplishments to date, and examines the remaining risks associated with full scale development.

Brown, Thomas↗

Preliminary Design of Low-Thrust Interplanetary Missions

For interplanetary missions, highly efficient electric propulsion systems can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems. This technology is being demonstrated on the Deep Space 1 mission - part of NASA's New Millennium Program validating technologies which can lower the cost and risk and enhance the performance of future missions. With the successful demonstration on Deep Space 1, future missions can consider electric propulsion as a viable propulsion option. Electric propulsion systems, while highly efficient, produce only a small amount of thrust. As a result, the engines operate during a significant fraction of the trajectory. This characteristic makes it much more difficult to find optimal trajectories. The methods for optimizing low-thrust trajectories are typically categorized as either indirect, or direct. Indirect methods are based on calculus of variations, resulting in a two-point boundary value problem that is solved by satisfying terminal constraints and targeting conditions. These methods are subject to extreme sensitivity to the initial guess of the variables - some of which are not physically intuitive. Adding a gravity assist to the trajectory compounds the sensitivity. Direct methods parameterize the problem and use nonlinear programming techniques to optimize an objective function by adjusting a set of variables. A variety of methods of this type have been examined with varying results. These methods are subject to the limitations of the nonlinear programming techniques. In this paper we present a direct method intended to be used primarily for preliminary design of low-thrust interplanetary trajectories, including those with multiple gravity assists. Preliminary design implies a willingness to accept limited accuracy to achieve an efficient algorithm that executes quickly.

Sims, Jon A.↗

Two dimensional wedge/translating shroud nozzle

A jet propulsion exhaust nozzle is reported for multi-engine installations which produces high internal/external, thrust-minus-drag, performance for transonic cruise or transonic acceleration as well as improved performance at subsonic and supersonic speeds. A two dimensional wedge/translating shroud provides the variable nozzle exit geometry needed to achieve high engine performance over a wide range of throttle power settings.

Maiden, D. L.↗

The FRC Acceleration Space Thruster (FAST) Experiment

The objective of the FRC (Field Reversed Configuration) Acceleration Space Thruster (FAST) Experiment is to investigate the use of a repetitive FRC source as a thruster, specifically for an NEP (nuclear electric propulsion) system. The Field Reversed Configuration is a plasmoid with a closed poloidal field line structure, and has been extensively studied as a fusion reactor core. An FRC thruster works by repetitively producing FRCs and accelerating them to high velocity. An FRC thruster should be capable of I(sub sp)'s in the range of 5,000 - 25,000 seconds and efficiencies in the range of 60 - 80 %. In addition, they can have thrust densities as high as 10(exp 6) N/m2, and as they are inductively formed, they do not suffer from electrode erosion. The jet-power should be scalable from the low to the high power regime. The FAST experiment consists of a theta-pinch formation chamber, followed by an acceleration stage. Initially, we will produce and accelerate single FRCs. The initial focus of the experiment will be on the ionization, formation and acceleration of a single plasmoid, so as to determine the likely efficiency and I(sub sp). Subsequently, we will modify the device for repetitive burst-mode operation (5-10 shots). A variety of diagnostics are or will be available for this work, including a HeNe interferometer, high-speed cameras, and a Thomson-scattering system. The status of the experiment will be described.

Martin, Adam↗

An investigation at static conditions of nonaxisymmetric nozzle thrust reverser port geometry including effects of rounding and rotating the port corner

An investigation has been conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the effects of rounding and rotating the port corner of a nonaxisymmetric thrust reversing port. High pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 6.0. The investigation indicates that using a rounded corner in the reverser port passage increases the discharge coefficient significantly compared to that of a sharp cornered thrust reversing port. Incorporating a rotating cylinder into the rounded corner reverser port passage provides no additional increase at cylinder revolutions per minute up to 10,000. (Revolutions per minute of 390,000 were necessary to achieve a surface velocity equal to the duct velocity.) Highest levels of discharge coefficient are obtained by providing a large, round, cylindrical surface in the port upstream wall. Locating the reverser port in a constant area duct results in higher discharge coefficients than locating the port in a convergent duct.

Arbiter, D. G.↗

Near Earth Asteroid Scout: NASA's Solar Sail Mission to a NEA

NASA is developing a solar sail propulsion system for use on the Near Earth Asteroid (NEA) Scout reconnaissance mission and laying the groundwork for their use in future deep space science and exploration missions. Solar sails use sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like sail made of a lightweight, highly reflective material. This continuous photon pressure provides propellant-less thrust, allowing for very high delta V maneuvers on long-duration, deep space exploration. Since reflected light produces thrust, solar sails require no onboard propellant. The Near Earth Asteroid (NEA) Scout mission, funded by NASA’s Advanced Exploration Systems Program and managed by NASA MSFC, will use the sail as primary propulsion allowing it to survey and image Asteroid 1991VG and, potentially, other NEA’s of interest for possible future human exploration. The NEA Scout spacecraft is housed in a 6U CubeSat-form factor and utilizes an 86 square meter solar sail for a total mass less than 14 kilograms. The mission is in partnership with the Jet Propulsion Laboratory with support from Langley Research Center and science participants from various institutions. NEA Scout will be launched on the maiden flight of the Space Launch System in 2019. The solar sail for NEA Scout will be based on the technology developed and flown by the NASA NanoSail-D and flown on The Planetary Society’s Lightsail-A. Four approximately-7-meter stainless steel booms wrapped on two spools (two overlapping booms per spool) will be motor driven and pull the sail from its stowed volume. The sail material is an aluminized polyimide approximately 2.5 microns thick. As the technology matures, solar sails will increasingly be used to enable science and exploration missions that are currently impossible or prohibitively expensive using traditional chemical and electric propulsion systems. This paper will summarize the status of the NEA Scout mission and solar sail technology in general.

Near Earth Asteroid↗