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NASA NTRSDate not supplied

Risk-Based Methodology for the Determination of Hall Thruster Performance Specifications

This work adapts manufacturing and metrology industry standards to create a risk-based approach for the determination of electric propulsion performance specifications. The developed process is applied to 208 total thrust measurements acquired using three different NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Unit (TDU) 12.5-kW Hall thrusters in three different test facilities in order to generate thrust specifications at the 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW operating conditions. These thrust specifications are used to perform a set of sample acceptance tests to illustrate the importance of adding additional conservatism using a parameter known of the Test Accuracy Ratio in order to control the risk to the thruster producer and consumer from false rejection and acceptance of thruster units.

HERMeS↗

NASA NTRSDate not supplied

Risk-based Methodology for the Determination of Hall Thruster Performance Specifications

HERMeS↗

NASA NTRSDate not supplied

Expanded Performance Characterization of the NASA HERMeS Hall Thruster

This work presents a summary of a detailed performance assessment of the Hall Effect Rocket with Magnetic Shielding (HERMeS) Technology Demonstration Unit 3 (TDU-3) thruster at the throttle conditions most relevant for the Gateway Power and Propulsion Element (PPE). First, an assessment was performed of TDU-3 performance and stability at the previously-tested throttle points of 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW that confirmed TDU-3 was operating nominally and consistent with historical baselines. TDU-3 performance and stability were then assessed throughout the PPE throttle range of 2.6-12 kW and shown to mirror the characteristics observed at the previously tested conditions. Finally, a statistical estimate for the thrust specification at the expanded set of throttle conditions was computed and found to yield similar tolerance ranges to those calculated for the heritage conditions. Taken together, these results provide strong evidence that the demonstrated performance characteristics of the HERMeS design are preserved at the expanded set of throttle conditions relevant to application on PPE.

HERMeS↗

NASA NTRSDate not supplied

Extended Performance Characterization of the NASA HERMeS Hall Thruster

HERMeS↗

NASA NTRSDate not supplied

Design and Comparison of Lunar Sample Return Options

Returning lunar samples to earth is a high priority for NASA in order to better understand Lunar Geology, Lunar formation, and potential resource utilization. Core samples of frozen volatiles from permanently shadowed regions are sought after, but require intensive cryogenic cooling to preserve. Returning samples to Earth using the Orion capsule is restrictive for moderate amounts of mass/volume and thus alternate means of transport are investigated in this paper. The trade-space of potential options includes varying: sample mass amount, sample containment conditions, transfer path from Lunar surface to Earth surface, and various vehicle system trades. An integrated solver is applied with parametric estimators and known system heuristics to effectively estimate size, mass, and power of various subsystems. Modeled subsystems include sample containment, EDL technologies(heat shield, parachute, recovery), power, thermal, avionics, structures, and propulsive systems. Transfer path options include various trajectory options, transfer methods, staging points, and vehicle staging options. A Design of Experiments was conducted to explore the tradespace, to identify sensitivities, and to identify promising designs. Results are detailed for various stakeholder objectives including total mass, sample mass, science return, architecture impacts, and operational flexibility.

Lunar outpost↗

NASA NTRS2016

Heatshield for Extreme Entry Environment Technology (HEEET) Development Status

The Heat shield for Extreme Entry Environment Technology (HEEET) Project is a NASA STMD and SMD co-funded effort. The goal is to develop and mission infuse a new ablative Thermal Protection System that can withstand extreme entry. It is targeted to support NASAs high priority missions, as defined in the latest decadal survey, to destinations such as Venus and Saturn in-situ robotic science missions. Entry into these planetary atmospheres results in extreme heating. The entry peak heat-flux and associated pressure are estimated to be between one and two orders of magnitude higher than those experienced by Mars Science Laboratory or Lunar return missions. In the recent New Frontiers community announcement NASA has indicated that it is considering providing an increase to the PI managed mission cost (PIMMC) for investigations utilizing the Heat Shield for Extreme Entry Environment Technology (HEEET) and in addition, NASA is considering limiting the risk assessment to only their accommodation on the spacecraft and the mission environment. The HEEET ablative TPS utilizes 3D weaving technology to manufacture a dual layer material architecture. The 3-D weaving allows for flat panels to be woven. The dual layer consists of a top layer designed to withstand the extreme external environment while the inner or insulating layer by design, is designed to achieve low thermal conductivity, and it keeps the heat from conducting towards the structure underneath. Both arc jet testing combined with material properties have been used to develop thermal response models that allows for comparison of performance with heritage carbon phenolic. A 50 mass efficiency is achieved by the dual layer construct compared to carbon phenolic for a broad range of missions both to Saturn and Venus. The 3-D woven flat preforms are molded to achieve the shape as they are compliant and then resin infusion with curing forms a rigid panels. These panels are then bonded on to the aeroshell structure. Gaps exist between the panels and these gaps have to be filled with seams. The seam material then has to be bonded on to adjacent panels and also to the structure. The heat-shield assembly is shown in Figure 1. One of the significant challenges we have overcome recently is the design, development and testing of the seam. HEEET material development and the seam concept development have utilized some of the unique test capabilities available in the US. The various test facilities utilized in thermal testing along with the entry environment for Saturn and Venus missions are shown in Figure 2. The HEEET project is currently in its 3rd year of a four-year development. Figure 3 illustrates the key accomplishments to date and the challenges yet to be overcome before the technology is ready for mission infusion. This proposed presentation will cover both progress that has been made in the HEEET project and also the challenges to be overcome that is highlighted in Figure 3. Objective of the HEEET project is to mature the system in time to support the next New Frontiers opportunity and we believe we are well along the way to mission infuse HEEET.

Thermal Protection System↗

NASA NTRSDate not supplied

Shields-1 Dosimetry Measurements in Polar Low Earth Orbit

Shields-1, NORAD ID 43850, has been operating in polar low earth orbit since December 2018 at 85 degrees inclination and an initial altitude of 500 kilometers, beginning in solar minimum activity. It launched with the NASA CubeSat Launch Initiative (CSLI) ELaNaXIX mission. Shields-1 has 8 Teledyne dosimeters on board behind different shielding thicknesses. It is a technology demonstration of an atomic number (Z)-grade shielded electronics enclosure, 3 Z-grade shielding samples baselined with 4 aluminum samples having dosimeters behind each sample and inside the electronics enclosure. In addition to the technology demonstration, Shields-1 has been providing scientific dosimetry of energetic particles in the polar low earth orbit space radiation environment.

Shields-1↗

NASA NTRS2014

Preliminary Development of a Multifunctional Hot Structure Heat Shield

Development of a Multifunctional Hot Structure Heat Shield concept has initiated with the goal to provide advanced technology with significant benefits compared to the current state of the art heat shield technology. The concept is unique in integrating the function of the thermal protection system with the primary load carrying structural component. An advanced carbon-carbon material system has been evaluated for the load carrying structure, which will be utilized on the outer surface of the heat shield, and thus will operate as a hot structure exposed to the severe aerodynamic heating associated with planetary entry. Flexible, highly efficient blanket insulation has been sized for use underneath the hot structure to maintain desired internal temperatures. The approach was to develop a preliminary design to demonstrate feasibility of the concept. The preliminary results indicate that the concept has the potential to save both mass and volume with significantly less recession compared to traditional heat shield designs, and thus provide potential to enable new planetary missions.

Walker, Sandra P.↗

NASA NTRSDate not supplied

Design of an Acoustic Shielding Flap Concept with Prediction and Validation

As part of a NATO task group, NASA has designed a noise reduction technology called a Shielding Flap for application to a hybrid wing body aircraft concept. The design drivers and objectives for this technology are outlined in this paper. Design parameters are defined, and recently-developed NASA software for prediction of acoustic scattering is used to investigate the expected performance of the technology for a range of parameter values. Detailed scattering computational results are presented for several simulated noise sources and for many variations of the shielding flap concept. A selection of computational results is compared with experimental data collected in the NASA Langley Quiet Flow Facility. Overall, the results indicate that the Shielding Flap has strong potential to reduce aft-radiated noise, which was a primary design objective, by shielding sound from the simulated sources while also shielding sound that is scattered from the main airframe.

Propulsion Airframe Aeroacoustics↗

NASA NTRS2001

Shielded Heavy-Ion Environment Linear Detector (SHIELD): an experiment for the Radiation and Technology Demonstration (RTD) Mission

Radiological assessment of the many cosmic ion species of widely distributed energies requires the use of theoretical transport models to accurately describe diverse physical processes related to nuclear reactions in spacecraft structures, planetary atmospheres and surfaces, and tissues. Heavy-ion transport models that were designed to characterize shielded radiation fields have been validated through comparison with data from thick-target irradiation experiments at particle accelerators. With the RTD Mission comes a unique opportunity to validate existing radiation transport models and guide the development of tools for shield design. For the first time, transport properties will be measured in free-space to characterize the shielding effectiveness of materials that are likely to be aboard interplanetary space missions. Target materials composed of aluminum, advanced composite spacecraft structure and other shielding materials, helium (a propellant) and tissue equivalent matrices will be evaluated. Large solid state detectors will provide kinetic energy and charge identification for incident heavy-ions and for secondary ions created in the target material. Transport calculations using the HZETRN model suggest that 8 g cm -2 thick targets would be adequate to evaluate the shielding effectiveness during solar minimum activity conditions for a period of 30 days or more.

NASA Center LaRC↗

NASA NTRS1991

A high-temperature furnace for applications in microgravity

Technology in the area of material processing and crystal growth has been greatly furthered by research in microgravity environments. The role of efficient, lightweight furnaces with reliable performance is crucial in these experiments. A need exists for the development of a readily duplicated, high-temperature furnace satisfying stringent weight, volume, and power constraints. A furnace was designed and is referred to as the UAH SHIELD. Stringent physical and operating characteristics for the system were specified, including a maximum weight of 20 kg, a maximum power requirement of 60 W, and a volume of the furnace assembly, excluding the batteries, limited to half a Get-Away-Special canister. The UAH SHIELD furnace uses radiation shield and vacuum technology applied in the form of a series of concentric cylinders enclosed on either end with disks. Thermal testing of a furnace prototype was performed in addition to some thermal and structural analysis. Results indicate the need for spacing of the shields to accommodate the thermal expansion during furnace operation. In addition, a power dissipation of approximately 100 W and system weight of approximately 30 kg was found for the current design.

Source record↗

NASA NTRS2015

Performance and Facility Background Pressure Characterization Tests of NASAs 12.5-kW Hall Effect Rocket with Magnetic Shielding Thruster

NASA's Space Technology Mission Directorate (STMD) Solar Electric Propulsion Technology Demonstration Mission (SEP/TDM) project is funding the development of a 12.5-kW Hall thruster system to support future NASA missions. The thruster designated Hall Effect Rocket with Magnetic Shielding (HERMeS) is a 12.5-kW Hall thruster with magnetic shielding incorporating a centrally mounted cathode. HERMeS was designed and modeled by a NASA GRC and JPL team and was fabricated and tested in vacuum facility 5 (VF5) at NASA GRC. Tests at NASA GRC were performed with the Technology Development Unit 1 (TDU1) thruster. TDU1's magnetic shielding topology was confirmed by measurement of anode potential and low electron temperature along the discharge chamber walls. Thermal characterization tests indicated that during full power thruster operation at peak magnetic field strength, the various thruster component temperatures were below prescribed maximum allowable limits. Performance characterization tests demonstrated the thruster's wide throttling range and found that the thruster can achieve a peak thruster efficiency of 63% at 12.5 kW 500 V and can attain a specific impulse of 3,000 s at 12.5 kW and a discharge voltage of 800 V. Facility background pressure variation tests revealed that the performance, operational characteristics, and magnetic shielding effectiveness of the TDU1 design were mostly insensitive to increases in background pressure.

Performance↗

NASA NTRS2010

Impact of Spacecraft Shielding on Direct Ionization Soft Error Rates for sub-130 nm Technologies

We use ray tracing software to model various levels of spacecraft shielding complexity and energy deposition pulse height analysis to study how it affects the direct ionization soft error rate of microelectronic components in space. The analysis incorporates the galactic cosmic ray background, trapped proton, and solar heavy ion environments as well as the October 1989 and July 2000 solar particle events.

Pellish, Jonathan A.↗

NASA NTRS2010

Impact of Spacecraft Shielding on Direct Ionization Soft Error Rates for Sub-130 nm Technologies

Pellish, Jonathan A.↗

NASA NTRS2019

Challenges in Qualification of Thermal Protection Systems in Extreme Entry Environments

Planetary entry vehicles employ ablative TPS materials to shield the aeroshell from entry aeroheating environments. To ensure mission success, it must be demonstrated that the heat shield system, including local features such as seams, does not fail at conditions that are suitably margined beyond those expected in flight. Furthermore, its thermal response must be predictable, with acceptable fidelity, by computational tools used in heat shield design. Mission assurance is accomplished through a combination of ground testing and material response modelling. A material's robustness to failure is verified through arcjet testing while its thermal response is predicted by analytical tools that are verified against experimental data. Due to limitations in flight-like ground testing capability and lack of validated high-fidelity computational models, qualification of heat shield materials is often achieved by piecing together evidence from multiple ground tests and analytical simulations, none of which fully bound the flight conditions and vehicle configuration. Extreme heating environments (>2000 W/sq. cm heat flux and >2 atm pressure), experienced during entries at Venus, Saturn and Ice Giants, further stretch the current testing and modelling capabilities for applicable TPS materials. Fully-dense Carbon Phenolic was the material of choice for these applications; however, since heritage raw materials are no longer available, future uses of re-created Carbon Phenolic will require re-qualification. To address this sustainability challenge, NASA is developing a new dual-layer material based on 3D weaving technology called Heat shield for Extreme Entry Environments (HEEET). Regardless of TPS material, extreme environments pose additional certification challenges beyond what has been typical in recent NASA missions. Scope of this presentation: This presentation will give an overview of challenges faced in verifying TPS performance at extreme heating conditions. Examples include: (1) Bounding aeroheating parameters (heat flux, pressure, shear and enthalpy) in ground facilities. How to certify TPS if environments can't be bounded or aeroheating parameters can't be simultaneously achieved. (2) Higher uncertainties in ground test environments (facility calibration and analytical predictions) at extreme conditions. (3) Testing in flows similar to planetary atmosphere composition (H2/He for Gas and Ice Giants). (4) Test sample size limitations for qualifying seam designs. (5) Lack of computational tools capable of simulating all significant aspects of TPS performance (including initiation and propagation of failures). This presentation will provide recommendations on how the EDL community can address these challenges and mitigate some of the risks involved in flying TPS materials at extreme conditions. Examples include: (1) Dedicated activity to understanding TPS failure modes. Develop computational tools capable of modelling fluid interaction with material's thermostructural response. Validate these tools through failure testing. A better understanding of failure mechanisms may eliminate the need to fully bound all aeroheating parameters in ground testing. (2) Enhancements to current testing facilities to simulate flight-like ablation mechanism (ex. testing in Nitrogen at Ames Interaction Heating Facility to limit oxidation in favor of more sublimation). (3) Improved characterization of test conditions with new diagnostic methods and determination of environment uncertainty through rigorous statistical analysis of available data. (4) Design margin policies that are directly tied to uncertainties in ground test environments and modelling fidelity

Mahzari, Milad↗

NASA NTRS1998

Anti-Matter Propulsion for Space. (Latest citations from the Aerospace Database)

The bibliography contains citations concerning techniques for the efficient production, long-term storage and effective utilization of antimatter for space propulsion. Inertial confinement fusion (ICF), magnetic mirror fusion, and liquid-propellant thermal antimatter fusion propulsion are some of the technologies discussed. Radiation shields, cryogenic confinement of plasma and single-stage-to-orbit vehicles are also cited. (Contains 50-250 citations and includes a subject term index and title list.)

ANTIPROTONS↗

NASA NTRS1984

Free jet feasibility study of a thermal acoustic shield concept for AST/VCE application: Single stream nozzles

A technology base for the thermal acoustic shield concept as a noise suppression device for single stream exhaust nozzles was developed. Acoustic data for 314 test points for 9 scale model nozzle configurations were obtained. Five of these configurations employed an unsuppressed annular plug core jet and the remaining four nozzles employed a 32 chute suppressor core nozzle. Influence of simulated flight and selected geometric and aerodynamic flow variables on the acoustic behavior of the thermal acoustic shield was determined. Laser velocimeter and aerodynamic measurements were employed to yield valuable diagnostic information regarding the flow field characteristics of these nozzles. An existing theoretical aeroacoustic prediction method was modified to predict the acoustic characteristics of partial thermal acoustic shields.

Majjigi, R. K.↗

NASA NTRS2018

Optimization of the Magnetic Field Topology in the Hall Effect Rocket with Magnetic Shielding

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kilowatt Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The TDU-1 thruster implements a magnetically shielded field topology and has demonstrated the elimination of the discharge channel erosion. Extensive wear testing of the TDU Hall thrusters has identified the thruster front pole covers as the next life limiting component. This effort aims to explore and investigate alternate magnetic field topologies to assess whether reductions in the front pole cover erosion can be attained while still maintaining very low erosion rates on the discharge channel walls. NASA GRC (Glenn Research Center) and JPL (Jet Propulsion Laboratory) have begun a magnetic field topology characterization and optimization study by designing four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the pole covers. Three of the four candidate magnetic field topologies have been manufactured subjected to an extensive test campaign that includes performance, plume, and stability characterization. In the Phase I test segment, the thruster's oscillation magnitude and laser-induced fluorescence (LIF) measurements of the plasma plume were performed for the three candidate topologies. In Phase I, the thruster's oscillation magnitude and LIF measurements were performed for the three candidate topologies. Phase I test results found that the B1 configuration attained lower oscillation levels than the baseline topology (B0). Additionally, laser induced fluorescence measurements along the discharge chamber centerline found that upstream retraction of the thruster’s peak magnetic field does result in an upstream shift of the acceleration zone but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment will include performing performance, stability, plume, and erosion measurements for the various candidate magnetic field topologies..

Kamhawi, Hani↗