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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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Monograph treats valve design technology problems as they were solved in successful development of flightweight operational valves for liquid rocket systems. General practices for cleaning and contamination prevention are summarized. Balance of information is arranged by topic, since detail design requirements apply to most types of valves.
Turbine airfoils are subjected to increasingly higher heat loads which escalate the cooling requirements in order to satisfy life goals for the component materials. If turbine efficiency is to be maintained, however, cooling requirements should be as low as possible. To keep the quantity of cooling air bounded, a more efficient internal cooling scheme must be developed. One approach is to employ airfoils with multipass cooling passages that contain devices to augment internal heat transfer while limiting pressure drop. Design experience with multipass cooling passage airfoils has shown that a surplus of cooling air must be provided as a margin of safety. This increased cooling air leads to a performance penalty. Reliable methods for predicting the internal thermal and aerodynamic performance of multipass cooling passage airfoils would reduce or eliminate the need for the safety margin of surplus cooling air. The objective of the program is to develop and verify improved analytical methods that will form the basis for design technology which will result in efficient turbine components with improved durability without sacrificing performance. The objective will be met by: (1) establishing a comprehensive experimental data base that can form the basis of an empirical design system; (2) developing computational fluid dynamic techniques; and (3) analyzing the information in the data base with both phenomenological modeling and mathematical modeling to derive a suitable design and analysis procedure.
Achieving a luminosity of $\gtrsim 10^{34} cm^{-2} s^{-1}$ in a $10 \text{ } TeV$ Muon Collider, given the short lifetime of a muon, requires reducing the 6D emittance of the muon beam through a process known as ionization cooling. In the final stage of this cooling process, the transverse emittance must be reduced to $22 \text{ } μm$, typically by allowing longitudinal emittance growth up to downstream acceptance limits. While the current International Muon Collider Collaboration designs involve $40 \text{ } T$ solenoids to reach the transverse emittance target, such high-field solenoids come with several challenges, including mechanical stress management, quench protection, and potential limitations in relying on High Temperature Superconductor technology. Designed as an alternative to using such solenoids while simultaneously reaching target transverse emittance, the previously proposed wedge-based, reverse emittance-exchange cooling scheme requires excellent dispersion suppression. In this study, we design and simulate a dispersion suppressor channel for the wedge-based final cooling design that reduces dispersion in the target direction to a target value of $D_x \sim 0.001 \text{ } m$.
A high voltage bandgap reference circuit has been designed and implemented in an SOI CMOS technology.
An experimental program comprising model nozzle and full-scale engine tests was undertaken to acquire parametric data for acoustically lined ejectors applied to primary jet noise suppression. Ejector lining design technology and acoustical scaling of lined ejector configurations were the major objectives. Ground static tests were run with a J-75 turbojet engine fitted with a 37-tube, area ratio 3.3 suppressor nozzle and two lengths of ejector shroud (L/D = 1 and 2). Seven ejector lining configurations were tested over the engine pressure ratio range of 1.40 to 2.40 with corresponding jet velocities between 305 and 610 M/sec. One-fourth scale model nozzles were tested over a pressure ratio range of 1.40 to 4.0 with jet total temperatures between ambient and 1088 K. Scaling of multielement nozzle ejector configurations was also studied using a single element of the nozzle array with identical ejector lengths and lining materials. Acoustic far field and near field data together with nozzle thrust performance and jet aerodynamic flow profiles are presented.
Several technology investments, beyond those used for robotic missions, are required to develop Mars human scale Entry, Descent, and Landing (EDL) systems. In a resource-constrained environment, studies play the critical role of identifying the most feasible technical paths and high payoff investments. One such NASA multi-directorate, multi-center study, is called the Entry, Descent and Landing Architecture Study (EDLAS). This paper presents a summary of phase two of the study conducted from October 2016 to September 2017. Study ground rules and assumptions are provided by NASA?s Human Exploration and Operations Mission Directorate (HEOMD) and include a Mars architecture and Mars surface lander payload manifests. Four unique entry technologies with the potential to deliver a specified 20 t human scale payload to the Mars surface are considered for analysis. Two of these techniques are evolutionary rigid vehicles, derived from robotic capsule missions and the space shuttle. Likewise, two revolutionary deployable vehicles are considered, the Adaptable Deployable Entry Placement Technology (ADEPT) and Hypersonic Inflatable Aerodynamic Decelerator (HIAD). This paper summarizes updates to the entry technology designs, the analysis motivation and approach, figures of merit by which the configurations are compared, and results of the analyses. Finally, findings are presented with the recommendation that two of the four configurations, one evolutionary and one revolutionary technology, continue to be studied in future human scale EDL studies.
Nephele is a small atmospheric probe mission concept to determine whether the Venus clouds contain organic matter. Nephele complements larger missions that seek to conduct Venus atmospheric analyses, such as DAVINCI, Venera-D, and the Morning Star missions, by directly sampling and analyzing cloud and haze particles. AERACEPT is an aerosol sampling technology designed to minimize the mass, volume, and complexity required for in situ planetary atmosphere characterization. A single aeroshell functions as a combined entry vehicle, descent probe, and instrumented sonde, combining recent advances in thermal protection materials with heritage technology from planetary and airborne science. The velocity of descent drives aerosol through inlets embedded in the aeroshell’s thermal protection material; inertial separation isolates the particles from the gas stream; and the particle capture surface doubles as the analysis substrate. This method can work without heat shield separation, descent control, or active sample handling; these advantages trade against a smaller sample volume, higher risk of sample heating, and shorter time for analysis and data return. AERACEPT further benefits from modeling of the internal and external flow to understand the particle capture efficiency and bias for size and concentration distributions. AERACEPT is well suited for a Venus mission such as Nephele, where a passive descent trajectory is both subsonic over the altitudes of interest and provides sufficient velocity and pressure differential for the particle sizes of interest. A thermal and flow toolchain developed in-house (Traj→ DPLR→ PATO→ openFOAM) was used to model an 80-cm aeroshell with science operations 63–39 km. These models indicate that AERACEPT will provide ~24 µL accumulated sample of particles ≥ 0.2 µm, substantially above the 2 µL threshold requirement. The worst-case particle heating will be ~14 K above ambient, for the smallest particles at the lowest altitude; most particles will experience ≤ 1 K rise, at which almost all analytes of interest will be unaffected. Plasma jet testing of a small-scale aeroshell-material inlet was recently conducted to validate the thermal model parameters; wind tunnel testing to validate the flow model parameters is underway.
The Green Flight Challenge is one of the National Aeronautics and Space Administration s Centennial Challenges designed to push technology and make passenger aircraft more efficient. Airliners currently average around 50 passenger-miles per gallon and this competition will push teams to greater than 200 passenger-miles per gallon. The aircraft must also fly at least 100 miles per hour for 200 miles. The total prize money for this competition is $1.65 Million. The Green Flight Challenge will be run by the Comparative Aircraft Flight Efficiency (CAFE) Foundation September 25 October 1, 2011 at Charles M. Schulz Sonoma County Airport in California. Thirteen custom aircraft were developed with electric, bio-diesel, and other bio-fuel engines. The aircraft are using various technologies to improve aerodynamic, propulsion, and structural efficiency. This paper will explore the feasibility of the rule set, competitor vehicles, design approaches, and technologies used.
This report is to present the work that was performed during the summer in the Advance Computing Application office. The NFFP (NASA Faculty Fellow Program) had ten summer faculty members working on IVHM (Integrated Vehicle Health Management) technologies. The objective of this project was two-fold: 1) to become familiar with IVHM concepts and key demonstrated IVHM technologies; and 2) to integrate the research that has been performed by IVHM faculty members into the MASTLAB (Marshall Avionic Software Test Lab). IVHM is a NASA-wide effort to coordinate, integrate and apply advanced software, sensors and design technologies to increase the level of intelligence, autonomy, and health state of future vehicles. IVHM is an important concept because it is consistent with the current plan for NASA to go to the moon, mars, and beyond. In order for NASA to become more involved in deep exploration, avionic systems will need to be highly adaptable and autonomous.
Propellants mass gauging technologies designed to work in an accelerated environment, where the propellant remains settled at one end of the propellant tank, do not work well in a microgravity environment because the propellant is not necessarily settled. While some microgravity mass gauging technologies exist at various TRLs, most of them have major disadvantages. Improvements in microgravity propellant mass gauging will result in improvements to many areas of propellant management, which influences performance and mission assurance. Electrical Capacitance Tomography (ECT) is a sensing technology that is able to reconstruct the liquid distribution inside of a tank, which can then be integrated to obtain mass. ECT mass gauging recently achieved high accuracy in laboratory testing, <0.1% liquid volume measurement error, even during sloshing. While ECT mass gauging will theoretically work during all phases of flight, it had not yet been tested in microgravity. The NASA KSC Launch Services Program, with support from the Flight Opportunities Program, successfully flew an ECT liquid mass gauging system experiment on a Zero-G parabolic flight aircraft. Basics of ECT measurement theory, details of the experiment setup and flights, and preliminary results will be discussed. The preliminary results suggest that, even in the current prototype generation, ECT sensor systems will be useful as a propellant mass gauging technology in both an accelerated and microgravity environment.
Propellants mass gauging technologies designed to work in an accelerated environment, where the propellant remains settled at one end of the propellant tank, do not work well in a microgravity environment because the propellant is not necessarily settled. While some microgravity mass gauging technologies exist at various TRLs, most of them have major disadvantages. Improvements in microgravity propellant mass gauging will result in improvements to many areas of propellant management, which influences performance and mission assurance. Electrical Capacitance Tomography (ECT) is a sensing technology that has been used in the oil/gas industry for decades to measure multi-phase flow rate in pipes. ECT, when installed in a tank, is able to reconstruct the liquid distribution inside of the tank, which can then be integrated to obtain mass. ECT mass gauging recently achieved high accuracy in laboratory testing on the ground, and while ECT mass gauging will theoretically work during all phases of flight, it had not yet been tested in microgravity. The NASA KSC Launch Services Program, with support from the Flight Opportunities Program, successfully flew an ECT tank liquid mass gauging system experiment on a Zero-G parabolic flight aircraft. The experiment hardware was rented from a company; NASA LSP did no technology development work for this project. This STRIVES entry is for the entire raw data set, which consists of approximately 25GB of csv files of capacitance and motion data from the ground and flight tests of the experiment. The data set will be available on the NASA LSP Electronic Slosh Data Catalog (ESDC), and the public may request the data set from the authors (POC: Jed Storey, jedediah.m.nasa.gov).
Advancements in microgravity propellant mass gauging will result in improvements to many areas of propellant management, which influences space vehicle performance and mission assurance. Propellant mass gauging technologies designed to work in an accelerated environment, where the propellant remains settled at one end of the propellant tank, do not always work well in a microgravity environment because the propellant is not necessarily settled. While some microgravity mass gauging technologies exist at various stages of development, most of them have major disadvantages. Electrical Capacitance Tomography (ECT) is a sensing technology that is able to reconstruct the liquid distribution inside of a tank, which can then be integrated to obtain mass. While ECT mass gauging will theoretically work during all phases of flight, it had not yet been tested in microgravity. The NASA Kennedy Space Center Launch Services Program, with support from the Flight Opportunities Program, successfully tested an ECT liquid mass gauging system experiment on a parabolic flight aircraft in May 2022. Basics of ECT measurement theory, details of the experiment setup, data processing, ground test results, and the flight test result will be discussed. The results suggest that ECT sensor systems will be useful as a propellant mass gauging technology in both accelerated and microgravity environments.
Computer design assistance for evolving large scale integrated circuit technology
A flight investigation has been conducted to study the behavior of three advanced-technology airfoils in the three-dimensional, unsteady-flow environment of the helicopter main rotor. Three sets of instrumented main-rotor blades were flown on an AH-1G helicopter. Each set employed one of three airfoils developed with significantly different design technologies. Data are given on performance, rotor loads, and the measurements of chordwise pressure distribution at 90 percent blade radius. The pressure data are compared with theoretical distributions calculated for two-dimensional, steady flow.
A number of aircraft and spacecraft configurations are being considered for future development. These include high-speed turboprop aircraft, advanced vertical take-off and landing fighter aircraft, and aerospace planes for hypersonic intercontinental cruise or flight to orbit and return. Review of the acoustic environment expected for these vehicles indicates levels high enough that acoustic fatigue must be considered. Unfortunately, the sonic fatique design technology used for current aircraft may not be adequate for these future vehicles. This has resulted in renewed emphasis on acoustic fatigue research at the NASA Langley Research Center. The overall objective of the Langley program is to develop methods and information for design of aerospace vehicles that will resist acoustic fatigue. The program includes definition of the acoustic loads acting on structures due to exhaust jets of boundary layers, and subsequent determination of the stresses within the structure due to these acoustic loads. Material fatigue associated with the high frequency structural stress reversal patterns resulting from acoustic loadings is considered to be an area requiring study, but no activity is currently underway.
Some of the design options under consideration for providing on-board electric power for the Solar Probe Mission are discussed. Five spacecraft configurations were evaluated with slightly different power demands and volumetric constraints on the power system. This resulted in three different baseline power system designs to satisfy the five spacecraft configurations. These three current baseline power system designs use modified general-purpose heat source (GPHS) radioisotope thermoelectric generators (RTGs) similar to those launched on the Galileo and Ulysses spacecraft. The modular RTG currently under development and testing is a potential advanced alternative to the current baseline GPHS-RTG technology design.
This paper will outline the design and test results for the Europa Lander motor controller’s Power Conversion Slice. Europa Lander is a proposed NASA mission concept to send a stationary spacecraft to Jupiter’s Icy Moon, Europa, to search for subsurface biosignatures. The Power Conversion Card is one of the many technologies designed for this mission. The Power Conversion Card’s primary function is to convert the power supplied by the spacecraft battery to usable power for use within the motor controller. In addition, this card provides power system telemetry to the spacecraft computer. This card tolerates Europa’s harsh radiation environment within a protected vault to enable it to survive the external temperatures which can be as low as -184°C during the three-week mission. This design makes use of multiple commercially available radiation hardened electronics, including Cobham’s UT32M0R500 Arm Microcontroller along with VPT’s DC-DC Converters for power conversion. The microcontroller is used to collect local power system telemetry and assess card health autonomously without intervention from the spacecraft’s main computer. This paper will showcase the design behind the prototype and electrical testing results.
A design technology study was performed to identify a high speed, multistage, variable geometry fan configuration capable of achieving wide flow modulation with near optimum efficiency at the important operating condition. A parametric screening study of the front and rear block fans was conducted in which the influence of major fan design features on weight and efficiency was determined. Key design parameters were varied systematically to determine the fan configuration most suited for a double bypass, variable cycle engine. Two and three stage fans were considered for the front block. A single stage, core driven fan was studied for the rear block. Variable geometry concepts were evaluated to provide near optimum off design performance. A detailed aerodynamic design and a preliminary mechanical design were carried out for the selected fan configuration. Performance predictions were made for the front and rear block fans.