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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Design of a Nozzle for the Spyder 2nd Stage Solid Rocket Motor

During the 2019 summer term, the author worked with a group of four interns to complete the preliminary design of a 2nd stage solid rocket motor for Up Aerospace’s Spyder Launch Vehicle. The Spyder vehicle is a four stage, solid fuel rocket designed as part of collaboration between NASA and Up Aerospace to develop a vehicle capable of delivering a 10 kg, 6U CubeSat into a 350 km, circular, low Earth orbit. As part of the agreement, NASA is tasked with designing high performance 2nd, 3rd, and 4th stages while Up Aerospace will provide the 1st stage, derived from the first stage of the company’s sub-orbital Spaceloft XL vehicle. Previous intern teams have designed the 3rd and 4th stages, which left the preliminary design of the 2nd stage motor to be completed this summer. The purpose of this report is to highlight a trade study which the author conducted to determine the nozzle geometry which would most benefit the performance of the 2nd stage motor. In this study, various nozzle parameters such as throat radius (RSI), expansion ratio, mass and their effects on the Isp and Delta V of the 2nd Stage were investigated. From this study, a nozzle geometry providing the necessary performance was chosen and implemented as part of the preliminary design of the 2nd stage motor. To mature the 2nd stage motor design, a trade space was needed to determine the nozzle configuration which would most benefit the performance of the 2nd stage. The trade space established did not only evaluate different expansion ratios for the same throat radius, but also investigated the possible performance gained from decreasing the throat radius to increase the expansion ratio and Isp capable of being delivered by the nozzle. Decreasing the throat radius would cause the chamber pressure to increase, consequently increasing the case and insulation mass required to safely operate a motor at higher pressures. To account for this factor, accurate estimates of inert mass first needed to be established. After doing so, the effects of varying nozzle expansion ratios, exit half angles, and subsequently length and mass were evaluated against motor and nozzle performance factors such as delta V and Isp. For this study, four throat radii ranging from 1.75” to 2.375” and consequently four different chamber pressures ranging 550 psia to 1200 psia were investigated. BACKGROUND To launch into Low Earth Orbit, a payload needs to be accelerated to the orbital velocity necessary to keep it from falling back to Earth. The change in velocity required between launch and orbital insertion is known as Delta V. The Delta V which a rocket or stage can deliver can be calculated using the Ideal Rocket equation, 𝛥𝑉=−𝑔0∗𝐼𝑠𝑝∗ln(𝑀𝑓𝑀𝑖) (3) Where 𝑔0 is the acceleration due to gravity at the earth’s surface, 𝐼𝑠𝑝 is the specific impulse of the rocket, 𝑀𝑖 is the initial mass of the rocket, and 𝑀𝑓 is the final mass of the rocket after burnout. From preliminary calculations beyond the scope of this paper, it was determined that 30500 ft/s of delta V would be required for a payload to be inserted into a 350 km circular orbit around the Earth. Using the known masses and Isp values of the 1st, 3rd, and 4th stages and equation 3, the delta V of each stage was calculated. The delta V required by the 2nd stage could then be found by taking the difference between the total delta V required and the delta V of the 1st, 3rd, and 4th stages. From this, the required delta V of the 2nd Stage was calculated to be 7340 ft/s. Specific impulse is an efficiency factor of the nozzle which defines the impulse delivered by the motor per unit of propellant weight. The main variables of a nozzle’s specific impulse investigated in this trade were exit cone half angle, throat radius, and expansion ratio which is affected by the throat radius. The expansion ratio, ε, of a nozzle is defined as the ratio between the nozzle exit area and throat area, and can be calculated using the equation, ε=𝑅𝑒𝑥𝑖𝑡2𝑅𝑠𝑖2 (2) Where 𝑅𝑒𝑥𝑖𝑡 is the radius of the nozzle’s exit and 𝑅𝑠𝑖 is the radius of the nozzle’s throat. A larger expansion ratio and smaller exit half angle will increase the Isp of a nozzle by allowing the gas to expand more and by allowing more of the exhaust gas to produce thrust in the direction of the motor’s central axis. A cross section view of the 2nd Stage motor with the major components annotated is provided in figure 1.

Bennett, Daniel↗
DOE OSTI2026

Plasma-Assisted Pre-Chamber Ignition System for Highly Dilute Stoichiometric Heavy-Duty Natural Gas Engines (Final Technical Report)

This project explored advanced ignition technologies to significantly enhance efficiency and reduce operating costs for heavy-duty natural gas engines operating at stoichiometric conditions, while meeting ultra-low NOx emission standards. The main goal was to develop and validate a plasma-assisted pre-chamber ignition system that could deliver at least a 2% increase in brake thermal efficiency (BTE) and a 4% decrease in total cost of ownership (TCO) compared to a typical multi-cylinder engine with three-way catalyst aftertreatment, ensuring compatibility with the expected 2027 EPA/CARB regulations. In the first half of the project, the research team concentrated on developing and testing plasma-assisted pre-chamber ignition using nanosecond pulsed discharges. Extensive experiments were conducted in an optically accessible rapid-compression and expansion machine, a constant-volume chamber, and an optical single-cylinder engine. Experiments were coupled with CFD simulations. The work produced unique insights into pre-chamber flame formation, jet ignition, dilution effects, and flame quenching at pressures, temperatures, and dilution levels relevant to engines. Although plasma-assisted ignition showed promise in controlled lab settings, the research also identified fundamental and practical challenges when applying this technology to real engine conditions. Midway through the project, a crucial pivot was made, guided by three key findings. First, the power electronics required for nanosecond plasma discharges were found to be too costly for commercial use, undermining the project’s cost-of-ownership goals. Second, nanosecond plasma ignition was highly sensitive to turbulent flow in the pre-chamber, resulting in lower ignition reliability than traditional spark under engine-like conditions. Third, achieving a truly diffuse low-temperature plasma at high pressures near top dead center was not possible, reducing the anticipated chemical enhancement benefits. These results collectively suggested that continuing with plasma-assisted ignition was unlikely to meet both efficiency and cost objectives. In response, the project shifted focus to a more realistic approach: enhancing traditional spark-based pre-chamber ignition with significantly less spark energy. Using insights gained earlier in the project, the team redesigned the pre-chamber to maintain high dilution tolerance and quick combustion, even with lower ignition energy. Testing confirmed that with optimized pre-chamber design and combustion timing, a lower-energy spark could reliably ignite highly diluted stoichiometric mixtures, reduce burn time, and boost thermal efficiency. Final engine testing and techno-economic analysis verified that this revised approach successfully achieved the project goals. The optimized pre-chamber ignition system provided over a 2% increase in calculated brake thermal efficiency compared to the baseline engine. Notably, the lower ignition energy and simplified hardware reduced component stress, extended maintenance intervals, and lowered the total cost of ownership. When used with stoichiometric operation and traditional three-way aftertreatment, the system remained compatible with near-zero NOx emissions targets without increasing cost or complexity in the emissions control system. In summary, although the project deviated from its initial plasma-assisted ignition idea, the work produced a more practical and commercially viable solution. The results show that precisely optimized, low-energy pre-chamber spark ignition can significantly improve efficiency and reduce overall ownership costs for heavy-duty natural gas engines. This directly aligns with DOE goals for cleaner, more efficient, and cost-effective transportation technologies.

03 NATURAL GAS↗
NASA NTRS1982

Ultimate intrinsic coercivity samarian-cobalt magnet. An Earth based feasibility study for Space Shuttle missions

Techniques for the electromagnetic containerless reaction of samarium with cobalt for the formation of samarium-cobalt alloys are summarized. The effort expended to reduce and instrument the oxygen partial pressure in the reaction chamber and coolant gas system are described as well as the experiments in which these improvements were shown to be partially effective. A stainless steel glove box capable of being evacuated to low 10 to the -6th torr pressure and refilled with ultra-pure argon was built and installed. Necessary accessories to perform SmCo5 powder preparation, compaction and subsequent encapsulation of the powder inside a hot isostatic pressing cannister were designed, built, and incorporated into the chamber. All accessories were tested for proper functioning inside the chamber. Using the facility, the first batch of densified SmCo5 powder was fabricated to near total density. Analysis of the densified compacts shows that oxygen contamination during fabrication was near zero.

Das, D.↗
NASA NTRS1981

Acoustic radiation force on a rigid sphere in a resonance chamber

The acoustic radiation force on a rigid sphere has been measured in a resonance chamber for a range of pressures, positions, sizes, and for various gases. In the low to medium intensity region of less than 150 dB, the measured force is consistent with King's theory (1934) when analyzed in terms of the fundamental pressure. However, in the high intensity region of greater than 150 dB, the measured force starts to deviate systematically from King's calculations.

Leung, E.↗
NASA NTRS1988

Debris prevention analysis for DFI/OFI/OEI (STS-26 configuration only)

The first 3 shuttle flights to use the Redesigned Solid Rocket Motors (RSRMs) will utilize Development Flight Instrumentation (DFI), as well as Operational Flight Instrumentation (OFI), and Operational Environment Instrumentation (OEI). The OFI consists of high pressure transducers used on both RSRMs to monitor the igniter and motor chamber pressure. DFI consists of assorted strain gages, temperature sensors, accelerometers, girth gages, and low level pressure transducers. The latter are installed on the left hand booster to measure post-separation aerodynamic loading. OEI consistes of temperature sensors. After Flight 3, all DFI gages are to be deleted, and only OFI and OEI will be used for subsequent flights. This report deals specifically with debris prevention and hazards concernig the STS-26 flight DFI configuration only. Continued analysis is being done that will adequately address the debris hazards associated with the STS-27 and subsequent flight DFI configurations.

Richi, Glen A.↗
NASA NTRS1989

SSME structural dynamic model development

The high pressure fuel turbopump (HPFTP) is a major component of the Space Shuttle Main Engine (SSME) powerhead. The device is a three stage centrifugal pump that is directly driven by a two stage hot gas turbine. The purpose of the pump is to deliver fuel (liquid hydrogen) from the low pressure fuel turbopump (LPFTP) through the main fuel valve (MFV) to the thrust chamber coolant circuits. In doing so, the pump pressurizes the fuel from an inlet pressure of approximately 178 psi to a discharge pressure of over 6000 psi. At full power level (FPL), the pump rotates at a speed of over 37,000 rpm while generating approximately 77,000 horsepower. Obviously, a pump failure at these speeds and power levels could jeopardize the mission. Results are summarized for work in which the solutions obtained from analytical models of the fuel turbopump impellers are compared with the results obtained from dynamic tests.

Foley, Michael J.↗
NASA NTRS1951

Combustion Instability in an Acid-Heptane Rocket with a Pressurized-Gas Propellant Pumping System

Results of experimental measurements of low-frequency combustion instability of a 300-pound thrust acid-heptane rocket engine were compared to the trends predicted by an analysis of combustion instability in a rocket engine with a pressurized-gas propellant pumping system. The simplified analysis, which assumes a monopropellant model, was based on the concept of a combustion the delay occurring from the moment of propellant injection to the moment of propellant combustion. This combustion time delay was experimentally measured; the experimental values were of approximately half the magnitude predicted by the analysis. The pressure-fluctuation frequency for a rocket engine with a characteristic length of 100 inches and operated at a combustion-chamber pressure of 280 pounds per square inch absolute was 38 cycles per second; the analysis indicated. a frequency of 37 cycles per second. Increasing combustion-chamber characteristic length decreased the pressure-fluctuation frequency, in conformity to the analysis. Increasing the chamber operating pressure or increasing the injector pressure drop increased the frequency. These latter two effects are contrary to the analysis; the discrepancies are attributed to the conflict between the assumptions made to simplify the analysis and the experimental conditions. Oxidant-fuel ratio had no apparent effect on the experimentally measured pressure-fluctuation frequency for acid-heptane ratios from 3.0 to 7.0. The frequencies decreased with increased amplitude of the combustion-chamber pressure variations. The analysis indicated that if the combustion time delay were sufficiently short, low-frequency combustion instability would be eliminated.

Tischler, Adelbert O.↗
NASA NTRS2005

Effects of Multiple Nozzles on Asymmetric Ejector Performance

This paper presents a comparison of a single nozzle and a dual nozzle strut based ejector. The results are focused on the fluid properties in the ejector duct. The research focused on choking mechanisms, mass flow entrainment, and mixing duct pressure distributions. The two ejectors were tests at equivalent primary mass flow rates. This corresponds to chamber pressures ranging from 100 psi to 900 psi in the single nozzle strut and 50 psi to 450 psi in the dual nozzle strut. Secondary flow was drawn from the lab at atmospheric pressure, and was not controlled. The secondary flow was found to choke at a value of 2.3 lb/s for a primary mass flow rate at approximately 2.1 lb/s for both ejectors. This choke was believed to be a mass addition choke rather than a traditional aerodynamic choke. The mixing duct pressure distribution exhibited two distinct trends at "low pressure" trend and at "high pressure" trend. For the low pressure trend, the mixing length for the ejectors remained fixed around 20 inches, regardless of the chamber pressure. For the higher pressure trend, the mixing length was considerably longer and increased with increasing chamber pressure. At high chamber pressures (high mass flow rates), a supersonic core flow was present at the exit of the duct. For these cases, the two streams did not have time to mix by the end of the duct.

Lineberry, D.↗
DOE OSTI2025

Skipper CCD Long-Term Vacuum Chamber

The Skipper CCD is a high-sensitivity Charge Coupled Device (CCD) sensor designed at Fermilab that has extremely low readout noise. The Skipper can only be operated at pressures below 10-4Torr (about a millionth of an atmosphere). The purpose of this project was to investigate whether a long-term vacuum chamber could be used for the Skipper CCD, which would require only being re-pumped once a year. This would allow the Skipper to be fully integrated here at Fermilab and then be transported anywhere for immediate use.

Struven, Wolfgang↗
NASA NTRSDate not supplied

Two-Phase Azimuthal Instability Generated By A Supersonic Jet Impinging on A Granular Bed

A persistent azimuthal pattern featuring alternate high and low concentration of ejecta emanating from the area where a supersonic jet (Mach 5.3) impinges on a bed of particles in a near-lunar vacuum condition is studied experimentally. Although this peculiar phenomenon has been documented in early studies motivated by extraterrestrial landing, the mechanism of this pattern is not clearly understood. Recently, a series of experiments were conducted at NASA Marshall Space Flight Center inside a 4.5 m vacuum chamber over a range of reduced ambient pressure. Experimental results show vibrant azimuthal patterns, which are clearest at low ambient pressure. The pattern is shown to be driven by the flow instability at low Reynolds number, despite the supersonic jet employed. The flow instability leads to azimuthal vortex lines, which expel ejecta into streaks between two neighboring vortex lines. A simple model is proposed to explain the observed phenomenon.

Plume Surface Interaction↗
NASA NTRS1977

Reactions between NO/+/ and metal atoms using magnetically confined afterglows

A new method of studying thermal energy ion-neutral collision processes involving nongaseous neutral atoms is described. A long magnetic field produced by a solenoid in a vacuum chamber confines a thermal-energy plasma generated by photoionization of gas at very low pressure. As the plasma moves toward the end of the field, it is crossed by a metal atom beam. Ionic products of ion-atom reactions are trapped by the field and both the reactant and product ions move to the end of the magnetic field where they are detected by a quadrupole mass filter. The cross sections for charge transfer between NO(+) and Na, Mg, Ca, and Sr and that for rearrangement between NO(+) and Ca have been obtained. The charge-transfer reaction is found strongly dominant over the rearrangement reaction that forms metallic oxide ions.

Lo, H. H.↗
NASA NTRS1979

LSA Large Area Silicon Sheet Task. Continuous Liquid Feed Czochralski Growth

The design and development of equipment and processes to demonstrate continuous growth of crystals by the Czochralski method suitable for producing single silicon crystals for use in solar cells is presented. The growth of at least 150 kg of mono silicon crystal, 150 mm in diameter is continuous from one growth container. A furnace with continuous liquid replenishment of the growth crucible, accomplished by a meltdown system with a continuous solid silicon feed mechanism and a liquid transfer system, with associated automatic feedback controls is discussed. Due to the silicon monoxide build up in the furnace and its retarding effect on crystal growth the furnace conversion for operation in the low pressure range is described. Development of systems for continuous solid recharging of the meltdown chamber for various forms of poly silicon is described.

Fiegl, G.↗
NASA NTRS2009

Sensor Requirements and Roadmaps

The chapter starts with a brief overview of the state-of-the-art of the sensing technologies. Principles of detection and the stage of integration in conjunction with the need are discussed. Based on the projected sensor requirements to accomplish propulsion systems with distributed active control, the existent gaps and selected types of sensor technologies are addressed, which are foreseen to deliver the required sensitivity, the resolution, the range, and the bandwidth in classes of sensors and which will allow operation under harsh environment conditions. By changing the packaging and/or design of the current sensors, operation environments of 750°C may be possible, which would meet requirements for sensors located towards the engine intake, compressor and in some cases low-pressure turbine. However, the majority of sensors for locations close to the engine combustion chamber or afterburner (with operation temperatures up to 1700°C) do not exist. There is also a need for smart sensors, which would enable future distributed control architecture. In addition a number of sensors, which are of interest for more intelligent gas turbine engines, are being explored or do not yet exist, for example turbine emission species sensors, burning patter factor sensors, fuel property sensors, and exhaust gas composition sensors. Meanwhile, rapid emerging technologies were made available during the three years duration of the Task Group. These include tip clearance measurement technologies such as eddy current, and microwave methods. To meet future sensor requirements, new fabrication and material technologies (MEMS, other), advanced sensing principles (spectroscopy, other), and their potential applications to new sensor types need to be explored. The status of current R&D for new sensors and future expectations are summarized in a comprehensive table. Sensors and associated electronics that have to operate in high temperature are foreseen to be largely available by 2015 (SiC) and 2020 (SiCN). Both of these technologies come with the potential of embedding such high temperature sensors within the structure. The information presented in the chapter represents the state-of-the-art of the sensing technology of the years 2005-2007 to the best knowledge of the authors. The authors are also fully aware of the fact that there may be many ongoing research programs as well as individual research that are not publicly supported and public information is not available. Therefore, the report might miss that section of knowledge. Also, progress in research and development is made every day such that often, the recent information may be fast outdated. The reader should bear in mind these facts.

Ion Stiharu↗
NASA NTRS1982

Energy efficient engine: Low-pressure turbine subsonic cascade component development and integration program

A subsonic cascade test program was conducted to provide technical data for optimizing the blade and vane airfoil designs for the Energy Efficient Engine Low-Pressure Turbine component. The program consisted of three parts. The first involved an evaluation of the low-chamber inlet guide vane. The second, was an evaluation of two candidate aerodynamic loading philosophies for the fourth blade root section. The third part consisted of an evaluation of three candidate airfoil geometries for the fourth blade mean section. The performance of each candidate airfoil was evaluated in a linear cascade configuration. The overall results of this study indicate that the aft-loaded airfoil designs resulted in lower losses which substantiated Pratt & Whitney Aircraft's design philosophy for the Energy Efficient Engine low-pressure turbine component.

Sharma, O. P.↗
DOE OSTI2020

Low-diffusion Xe-He gas mixtures for rare-event detection: electroluminescence yield

High pressure xenon Time Projection Chambers (TPC) based on secondary scintillation (electroluminescence) signal amplification are being proposed for rare event detection such as directional dark matter, double electron capture and double beta decay detection. The discrimination of the rare event through the topological signature of primary ionisation trails is a major asset for this type of TPC when compared to single liquid or double-phase TPCs, limited mainly by the high electron diffusion in pure xenon. Helium admixtures with xenon can be an attractive solution to reduce the electron diffusion significantly, improving the discrimination efficiency of these optical TPCs. We have measured the electroluminescence (EL) yield of Xe-He mixtures, in the range of 0 to 30% He and demonstrated the small impact on the EL yield of the addition of helium to pure xenon. For a typical reduced electric field of 2.5 kV/cm/bar in the scintillation region, the EL yield is lowered by ~ 2%, 3%, 6% and 10% for 10%, 15%, 20% and 30% of helium concentration, respectively. This decrease is less than what has been obtained from the most recent simulation framework in the literature. The impact of the addition of helium on EL statistical fluctuations is negligible, within the experimental uncertainties. The present results are an important benchmark for the simulation tools to be applied to future optical TPCs based on Xe-He mixtures.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗
NASA NTRS1971

Low-pressure oxidation of Cb-1Zr alloy.

Resistively heated strip specimens of Cb-1Zr alloy were exposed at 927 C in a vacuum chamber at various levels of total pressure in the 1-microtorr range and at various oxygen partial pressures in the .1-microtorr range. Oxygen reaction rates (sticking probabilities) were found to depend on whether or not the specimens were annealed immediately before the test exposure. It is shown that a normally undetectable oxide film exists on the Cb-1Zr surface as a result of oxidation by ambient air, and this film reduces the sticking probability as compared with a clean metal surface. The alloy is considerably strengthened by addition of oxygen to a level of about 6000 ppm, while still maintaining reasonably good room temperature ductility.

Lyon, T. F.↗
NASA NTRS1987

Pressure scaling effects in a scramjet combustion chamber

The test results obtained for a model scramjet over a range of pressure levels corresponding to different flight altitudes involve enthalpies that vary from the ignition limit, at the low temperature end, to temperatures where the dissociation of combustion products severely limits heat release. The minimum temperature is noted to be highly pressure-sensitive; above the ignition limit, the amount of heat release increased markedly with pressure and with combustion chamber length. A FEM computer code has been used to model the mixing and combustion processes.

Morgan, R. G.↗
NASA NTRS1991

Passive control of supersonic cavity flowfields

A computational investigation has been conducted to study the effect and mechanisms of the passive control of a supersonic flow over a rectangular two-dimensional cavity. The passive control was included through the use of a porous surface over a vent chamber in the floor of the cavity. The passive control effectively suppressed the low-frequency pressure oscillations for the open type cavity, (length-to-depth ratio = 6.0). The mechanism for the suppression was observed to be the stabilization of the motion of the free shear layer. For the closed type cavity flow, (length-to-depth ratio = 17.5), the passive control modified the flowfield to nearly that of an open type cavity flow; further the cavity drag was reduced by a factor of four. The computational results of both cases showed good agreement with the available experimental data and the predictions of a semiempirical formula. This study demonstrates that the passive control concept can be used to improve the aerodynamic characteristics of open and closed cavity flowfields.

Chokani, N.↗