SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Study examines scaling parameters of three fluid amplifier elements - a bistable device, a boundary layer control device, and a vortex device. Variations in performance due to size, fluid, and other conditions are studied. Even with restricted examples the large number of variables impedes the establishment of these scaling laws.
Urban Air Mobility (UAM) is an up and coming topic in the world of aerosciences. A variety of companies, most notably Uber, has begun working towards making air taxis a regular part of day to day life. However, the implementation of this system into society will require a significant amount of work: making helicopters that can handle these flights and operate safely in urban environments, having landing zones, and working out new technologies to lower costs. NASA has also taken interest in this new idea and has formed a focus group of interns to tackle some of these problems. They have started by designing a few concept models for what these air taxis could look like, focusing on low noise, multiple passenger configurations, and the potential for electric or hybrid helicopters. One particular model is known as NR2, or the "Side by Side," which features two rotors spinning in opposite directions. Though these models are purely conceptual at this point, understanding the real-world aerodynamics and performance of this model is crucial to its future development. Using a program known as Rotorcraft Computational Fluid Dynamics (RotCFD), the performance characteristics of the rotors and aircraft, such as the lift, drag, thrust and power can be determined. These simulations can be run in a variety of flight configurations, with hover, forward flight, and climb being focused on in this study. The rotor collective angle can also be changed and used to get results about the lift and drag for the blades and the model body. These simulations can provide information for future model construction and wind tunnel testing.
Nuclear Electric Propulsion (NEP) is an in-space propulsion technology capable of enabling opposition and conjunction class crewed Mars missions. NEP subsystems include the reactor for heat generation, a power conversion system (PCS), power management and distribution (PMAD), electric propulsion subsystem (EPS), and a primary heat rejection system. Specific mass, or αe (kg / kWe), is a key performance parameter (KPP) of the propulsion system which is directly scalable with the performance and mass estimates for individual components. To inform technology maturation planning, full system and component level parametric modeling is ongoing to explore the design trade space and illustrate the effect of subsystem design choices on the system KPPs. In this study, scaling of high-assay, low-enriched uranium (HALEU) reactor designs is assessed through coupled reactor physics and thermal hydraulics analyses. Scaling analyses evaluate the impact of system performance parameters (power level, interface temperatures) on mass for direct gas cooled, pumped liquid metal, and passively cooled heat pipe reactor concepts. Each concept requires specific geometries and working fluids to reach the performance goals of PCS interface conditions (temperature, pressure, flow rate) and system mass. The reactor assembly includes the active core (fuel, moderator, cladding, working fluid), axial and radial neutron reflectors, control drums, structural support / pressure vessel, and external radiation shielding. Each of these components are parametrically sized based on performance parameters for a megawatt-class power cycle. Results of this scaling analysis increase NEP propulsion system modeling fidelity and ultimately aim to support technology down-selection along with related technology development planning. The reactor and shield αe are a function of several PCS design choices, and reactor scaling with these parameters must be considered to enable an informed decision on reactor geometry and working fluid combination.
Experiments performed on Space Shuttle flights have emphasized study of the earliest effects of the cephalad fluid shift resulting from microgravity. Analysis of one subject's urine collected during flight showed that a sharp increase in antidiuretic hormone occurred within 2 h of launch, followed by an increase in cortisol excretion. Although this subject had symptoms of the space adaptation syndrome (SAS), inflight data from Spacelab missions suggested that these transient changes were not caused by SAS. Unpaired t-tests and Mann-Whitney tests showed that before and after flight, plasma thyroxine and urine osmolality were significantly higher in Shuttle crewmembers who exhibited more severe symptoms of SAS than in asymptomatic crewmembers.
An analysis of the performance of a low-temperature Brayton space radiator is presented. The cylindrical radiator with circumferential tubes was subjected to an equivalent sink temperature environment that varied along the fluid flow path. The radiator performance, as measured by fluid outlet temperature, was found to be dependent on the orbital position and the location of the radiator inlet relative to the sink temperature environment. An approximate dynamic analysis of sun-shade variations in performance is included.
The intermediate temperature regime of heat pipes, 450-700 K, has not been well characterized due in part to the corrosive nature of many of the working fluids, as well as the poor performance in many conventional metal envelope systems. Spacecraft nuclear electric propulsion, as well as other emerging technologies, can benefit from heat pipes with optimal performance in this intermediate regime. In this work, we have explored halide working fluids coupled with aluminum nitride shells to operate at 500-600 K and interface with a nuclear electric propulsion system. The shells have been additively manufactured and tested to ensure compatibility with the halides. Specifically, AlBr 3 , AlCl 3 , FeCl 3 , and SbBr 3 were selected owing to their higher performance and comparatively safe handling characteristics compared to other intermediate temperature working fluids. The additive manufacturing process enables highly customizable designs that integrate the heat pipe and radiator fin together, optimized for each fluid. We have evaluated the performance limits of such a heat pipe design in order to establish the feasibility of AlBr 3 in particular, and present design that achieves an areal density of 5.2 kg/m 2 and mass per unit heat performance of 0.71 kg/kW. Analysis is ongoing for all working fluids.
The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQAIR code was applied to compute the convective and radiative heating rates over high-mass Mars entry vehicles. Two vehicle geometries were considered - a 15 m diameter 70-degree sphere cone configuration and a slender, mid-L/D vehicle with a diameter of 5 m called an Ellipsled. The entry masses ranged from 100 to 165 metric tons. Solutions were generated for entry velocities ranging from 6.5 to 9.1 km/s. The coupled fluids-radiation solutions were performed at the peak heating location along trajectories generated by the Traj trajectory analysis code. The impact of fluids-radiation coupling is a function of the level of radiative heating and the freestream density and velocity. For the high-mass Mars vehicles examined in this study, coupling effects were greatest for entry velocities above 8.5 km/s where the surface radiative heating was reduced by up 17%. Generally speaking, the Ellipsled geometry experiences a lower peak radiative heating rate but a higher peak turbulent convective heating rate than the MSL-based vehicle.
The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQAIR code was applied to compute the convective and radiative heating rates over high-mass Mars entry vehicles. Two vehicle geometries were considered - a 15 m diameter 70-degree sphere cone configuration and a slender, mid-L/D vehicle with a diameter of 5 m called an Ellipsled. The entry masses ranged from 100 to 165 metric tons. Solutions were generated for entry velocities ranging from 6.5 to 9.1 km/s. The coupled fluids-radiation solutions were performed at the peak heating location along trajectories generated by the Traj trajectory analysis code. The impact of fluids-radiation coupling is a function of the level of radiative heating and the freestream density and velocity. For the high-mass Mars vehicles examined in this study, coupling effects were greatest for entry velocities above 8.5 km/s where the surface radiative heating was reduced by up 17%. Generally speaking, the Ellipsled geometry experiences a lower peak radiative heating rate but a higher peak turbulent convective heating rate than the MSL-based vehicle.
A Neptune Aerocapture Systems Analysis is completed to determine the feasibility, benefit and risk of an aeroshell aerocapture system for Neptune and to identify technology gaps and technology performance goals. The high fidelity systems analysis is completed by a five center NASA team and includes the following disciplines and analyses: science; mission design; aeroshell configuration screening and definition; interplanetary navigation analyses; atmosphere modeling; computational fluid dynamics for aerodynamic performance and database definition; initial stability analyses; guidance development; atmospheric flight simulation; computational fluid dynamics and radiation analyses for aeroheating environment definition; thermal protection system design, concepts and sizing; mass properties; structures; spacecraft design and packaging; and mass sensitivities. Results show that aerocapture can deliver 1.4 times more mass to Neptune orbit than an all-propulsive system for the same launch vehicle. In addition aerocapture results in a 3-4 year reduction in trip time compared to all-propulsive systems. Aerocapture is feasible and performance is adequate for the Neptune aerocapture mission. Monte Carlo simulation results show 100% successful capture for all cases including conservative assumptions on atmosphere and navigation. Enabling technologies for this mission include TPS manufacturing; and aerothermodynamic methods and validation for determining coupled 3-D convection, radiation and ablation aeroheating rates and loads, and the effects on surface recession.
A program was carried out to develop and test advanced electrochemical cells/modules and critical electromechanical components for a static feed (alkaline electrolyte) water electrolysis oxygen generation subsystem. The accomplishments were refurbishment of a previously developed subsystem and successful demonstration for a total of 2980 hours of normal operation; achievement of sustained one-person level oxygen generation performance with state-of-the-art cell voltages averaging 1.61 V at 191 ASF for an operating temperature of 128F (equivalent to 1.51V when normalized to 180F); endurance testing and demonstration of reliable performance of the three-fluid pressure controller for 8650 hours; design and development of a fluid control assembly for this subsystem and demonstration of its performance; development and demonstration at the single cell and module levels of a unitized core composite cell that provides expanded differential pressure tolerance capability; fabrication and evaluation of a feed water electrolyte elimination five-cell module; and successful demonstration of an electrolysis module pressurization technique that can be used in place of nitrogen gas during the standby mode of operation to maintain system pressure and differential pressures.
A Systems Analysis was completed to determine the feasibility, benefit and risk of an aeroshell aerocapture system for Neptune and to identify technology gaps and technology performance goals. The systems analysis includes the following disciplines: science; mission design; aeroshell configuration; interplanetary navigation analyses; atmosphere modeling; computational fluid dynamics for aerodynamic performance and aeroheating environment; stability analyses; guidance development; atmospheric flight simulation; thermal protection system design; mass properties; structures; spacecraft design and packaging; and mass sensitivities. Results show that aerocapture is feasible and performance is adequate for the Neptune mission. Aerocapture can deliver 1.4 times more mass to Neptune orbit than an all-propulsive system for the same launch vehicle and results in a 3-4 year reduction in trip time compared to all-propulsive systems. Enabling technologies for this mission include TPS manufacturing; and aerothermodynamic methods for determining coupled 3-D convection, radiation and ablation aeroheating rates and loads.
NASA possesses a wealth of historical cryogenic experiments that provide valuable insights into the design, troubleshooting, and understanding involved in the complex fluid and thermodynamics of managing cryogenic propellants. One approach to leveraging these historical datasets is by simulating these experiments to validate the accuracy of simulation environments and constituent models. This study focused on simulating a selection of the K site test series for both static and sloshing pressurized liquid methane draining experiments conducted at NASA in the 1970’s, utilizing computational fluid dynamics. The simulations were performed in the ANSYS FLUENT environment using the Volume of Fluid (VOF) numerical approach. A k-omega turbulence model was used with interfacial turbulence damping, and accurately predicted the amount of pressurant needed to maintain the required tank pressure throughout the static drain. The static simulation predicted the temperature stratification in the ullage observed at the end of the drain. During the methane expulsion with sloshing test, many features were successfully captured using the k-omega turbulence model with interfacial turbulence damping included. The rate of phase change and liquid temperature was overpredicted compared to the experimental measurements. The overprediction may be attributed to uncertainties in the vessel geometry, methane pressurant temperature and composition, and methodological differences in how the sloshing frequency was adjusted during the expulsion.
A preliminary design study was performed to seek a fluid-film thrust bearing design intended to be part of a high-speed, hybrid (rolling element/fluid film) bearing configuration. The base line used is a design previously tested. To improve the accuracy of theoretical predictions of load capacity, flow rate, and friction power loss, an analytical procedure was developed to include curvature effects inherent in thrust bearings and to allow for the temperature rise in the fluid due to viscous heating. Also, a narrow-groove approximation in the treatment of the temperature field was formulated to apply the procedure to the Whipple thrust bearing. A comparative trade-off study was carried out assuming isothermal films; its results showed the shrouded-step design to be superior to the Whipple design for the intended application. An extensive parametric study was performed, employing isoviscous calculations, to determine the optimized design, which was subsequently recalculated allowing for temperature effects.
Supersonic axisymmetric, external-compression inlets designed for freestream Mach numbers of Mach 1.4, 1.7, and 2.0 were evaluated to characterize their off-design performance. The inlets were characterized as isolated from the airframe with freestream conditions, engine-face geometry, and flow rates established from a reference NASA commercial aircraft concept. The off-design performance was characterized using computational fluid dynamics (CFD) simulations. The study characterized the off-design performance at the cruise condition for variations in the inlet flow rate, angle-of-attack, and freestream Mach number. Also characterized was the inlet performance at take-off and approach-to-landing conditions for which the use of auxiliary intakes was included. The inlet performance was characterized by the flow rates, total pressure recovery and distortion, and drag. An understanding of the off-design performance of the axisymmetric inlets provides information for the selection of inlets for commercial supersonic aircraft.
Supersonic axisymmetric, external-compression inlets designed for freestream Mach numbers of Mach 1.4, 1.7, and 2.0 were evaluated to characterize their off-design performance. The inlets were characterized as isolated from the airframe with freestream conditions, engine-face geometry, and flow rates established from a reference NASA commercial aircraft concept. The off-design performance was characterized using computational fluid dynamics (CFD) simulations. The study characterized the off-design performance at the cruise condition for variations in the inlet flow rate, angle-of-attack, and freestream Mach number. Also characterized was the inlet performance at take-off and approach-to-landing conditions for which the use of auxiliary intakes was included. The inlet performance was characterized by the flow rates, total pressure recovery and distortion, and drag. An understanding of the off-design performance of the axisymmetric inlets provides information for the selection of inlets for commercial supersonic aircraft.
A ground-based experimental facility to perform flow boiling and condensation experiments is built in support of the development of the long duration Flow Boiling and Condensation Experiment (FBCE) destined for operation on board of the International Space Station (ISS) Fluid Integrated Rack (FIR). We performed tests with the condensation test module oriented horizontally and vertically. Using FC-72 as the test fluid and water as the cooling fluid, we evaluated the operational characteristics of the condensation module and generated ground based data encompassing the range of parameters of interest to the condensation experiment to be performed on the ISS. During this testing, we also evaluated the pressure drop profile across different components of the fluid subsystem, heater performance, on-orbit degassing subsystem, and the heat loss from different components. In this presentation, we discuss representative results of performance testing of the FBCE flow loop. These results will be used in the refinement of the flight system design and build-up of the FBCE which is scheduled for flight in 2019.
This report presents a study of numerical simulations of mixing layers developing between opposing flows of paired fluids under supercritical conditions, the purpose of the study being to elucidate chemical-species-specific aspects of turbulence. The simulations were performed for two different fluid pairs O2/H2 and C7H16/N2 at similar reduced initial pressures (reduced pressure is defined as pressure divided by critical pressure). Thermodynamically, O2/H2 behaves more nearly like an ideal mixture and has greater solubility, relative to C7H16/N2, which departs strongly from ideality. Because of a specified smaller initial density stratification, the C7H16/N2 layers exhibited greater levels of growth, global molecular mixing, and turbulence. However, smaller density gradients at the transitional state for the O2/H2 system were interpreted as indicating that locally, this system exhibits enhanced mixing as a consequence of its greater solubility and closer approach to ideality. These thermodynamic features were shown to affect entropy dissipation, which was found to be larger for O2/H2 and concentrated in high-density-gradient-magnitude regions that are distortions of the initial density-stratification boundary. In C7H16/N2, the regions of largest dissipation were found to lie in high-density-gradient-magnitude regions that result from mixing of the two fluids.
EXECUTIVE SUMMARY Mechanically connected joints used in cryogenic fluid lines as part of space flight elements need to survive launch vibrations and remain leak-free to minimize the loss of on-board commodity and hazardous gas accumulation. In 2020, a cryogenic test apparatus was developed which can evaluate the leak performance of pressurized threaded fluid fittings. The fittings were mounted in the TVAC and cooled to cryogenic test temperature and pressurized with helium while the leak rate was measured using a calibrated GHe leak detector. The test articles for the initial proof of concept testing were ¼ and 1 inch Swagelok VCR fittings with three different types of seal rings copper, nickel, and Ni. Each fitting configuration (size/seal material) was subjected to two consecutive cryogenic thermal cycles, followed by exposure to a launch vibration profile at ambient temperature, after which two additional TVAC cycle tests were performed. The testing reported here is a continuation of the 2020 tests with a statistically significant large number of samples and test runs. Three Swagelok VCR fitting sizes were tested ¼, ½ and 1 inch, and five (5) samples of each fitting size, each sample was tested with SST and Ni seal rings (Total of 30 unique test articles). Each test article was subjected to four (4) thermal cycles. Half of these cycles were performed before vibration testing and half were performed after vibration testing. The vibration testing was performed to evaluate the ability of the fittings to survive launch-type vibration profiles and remain leak-free. Leak checking of each fitting was completed at temperatures between 20K – 30K. The test procedure in Section 8.0 was designed to facilitate a qualification test program by allowing a higher test throughput rate coupled with repeatable test profiles. Results were very positive and show that out of the 30 samples they all passed with leak rates a factor of 2-3 lower than the established 10-6sccs GHe leak threshold. The result showed the Ni seals had lower leak rate, but the SST was more rugged. There were two deviations where damage to the Ni seal ring during assembly resulted in a leaky fitting, this is discussed in Section 10.7 Test Deviations. These fittings show great promise for space flight use and further testing is recommended to fully qualify the fittings per the ASTM F1387-19 and/or other relevant NASA specifications. The test equipment hardware and software capability developed for this testing is generic and not restricted to VCR fittings. It can be employed to evaluate/qualify the leak performance of other types of fittings and a wide range of other cryogenic fluid components such as valves, gages, connectors, etc.