Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “impulse”

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.

At least 865 records · Page 48
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

A Plasma Rocket Demonstration on the International Space Station

The Advanced Space Propulsion Laboratory at the NASA Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One feature of this concept is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. For this reason the system is called the Variable Specific Impulse Magneto-plasma Rocket or VASIMR. This ability to vary specific impulse and thrust will allow for optimum low thrust interplanetary trajectories and results in shorter trip times than is possible with fixed specific impulse systems while preserving adequate payload margins. In the development of the VASIMR technology, a series of ground-based experiments and space demonstrations are envisioned. A ground-based experiment of a low-power VASIMR prototype rocket is currently underway at the Advanced Space Propulsion Laboratory. The next step is a proposal to build and fly a 25-kilowatt VASIMR rocket as an external payload on the International Space Station. This experiment will provide an opportunity to demonstrate the performance of the rocket in space and measure the induced environment. The experiment will also utilize the space station for its intended purpose as a laboratory with vacuum conditions that cannot be matched by any laboratory on Earth. The VASIMR experiment will also blaze the trail for the wider application of advanced electric propulsion on the space station. An electric propulsion system like VASIMR, if provided with sufficient electrical power, could provide continuous drag force compensation for the space station. Drag compensation would eliminate the need for reboosting the station, an operation that will consume about 60 metric tons of propellant in a ten-year period. In contrast, an electric propulsion system would require very little propellant. In fact, a system like VASIMR can use waste hydrogen from the station's life support system as its propellant. This waste hydrogen is otherwise dumped overboard. Continuous drag compensation would also improve the microgravity conditions on the station. So electric propulsion can reduce propellant delivery requirements and thereby increase available payload capacity and at the same time improve the conditions for scientific research.

Petro, Andrew J.↗
NASA NTRS2016

Propulsion Estimates for High Energy Lunar Missions Using Future Propellants

High energy propellants for human lunar missions are analyzed, focusing on very advanced ozone and atomic hydrogen. One of the most advanced launch vehicle propulsion systems, such as the Space Shuttle Main Engine (SSME), used hydrogen and oxygen and had a delivered specific impulse of 453 seconds. In the early days of the space program, other propellants (or so called metapropellants) were suggested, including atomic hydrogen and liquid ozone. Theoretical and experimental studies of atomic hydrogen and ozone were conducted beginning in the late 1940s. This propellant research may have provided screenwriters with the idea of an atomic hydrogen-ozone rocket engine in the 1950 movie, Rocketship X-M. This paper presents analyses showing that an atomic hydrogen-ozone rocket engine could produce a specific impulse over a wide range of specific impulse values reaching as high as 1,600 s. A series of single stage and multistage rocket vehicle analyses were conducted to find the minimum specific impulse needed to conduct high energy round trip lunar missions.

rocket propulsion↗
NASA NTRS2016

Atmospheric Mining in the Outer Solar System: Outer Planet Orbital Transfer and Lander Analyses

High energy propellants for human lunar missions are analyzed, focusing on very advanced ozone and atomic hydrogen. One of the most advanced launch vehicle propulsion systems, such as the Space Shuttle Main Engine (SSME), used hydrogen and oxygen and had a delivered specific impulse of 453 seconds. In the early days of the space program, other propellants (or so called metapropellants) were suggested, including atomic hydrogen and liquid ozone. Theoretical and experimental studies of atomic hydrogen and ozone were conducted beginning in the late 1940s. This propellant research may have provided screenwriters with the idea of an atomic hydrogen-ozone rocket engine in the 1950 movie, Rocketship X-M. This paper presents analyses showing that an atomic hydrogen-ozone rocket engine could produce a specific impulse over a wide range of specific impulse values reaching as high as 1,600 seconds. A series of single stage and multistage rocket vehicle analyses were conducted to find the minimum specific impulse needed to conduct high energy round trip lunar missions.

lunar mission design↗
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↗
NASA NTRSDate not supplied

Hydrogen Peroxide-Water-Ethanol Monopropellant Blend for CubeSat Propulsion

Monopropellant propulsion that provides high performance and low hazards compared to state-of-art hydrazine are being investigated under NASA, DOD and commercially sponsored green propellant research and development programs. This NASA effort looks at phosphate stabilized 50% hydrogen peroxide/water –ethanol blends to evaluate their stability and combustion behavior in submillimeter diameter microtube combustion apparatus for cubesat propulsion applications. Exploratory 7-day stability tests find no pressure rise in a sealed test tube environment at room temperature indicating at least short term stability of the mixture. Temperature rise during combustion testing in resistively heated nickel microtube apparatus shows that micro-combustion is achieved in the tube. Propellant density is approximately 1.2 g/cu.cm compared to 1.0 g/cu.cm for hydrazine. Theoretical specific impulse of the stoichiometric blend is 219 s at a chamber pressure of 20 psia and an area ratio of 30:1. Density-specific impulse of this blend is 263 s while hydrazine is only about 215-220 s in small reaction control system thrusters. This provides about 20-22% more density-specific impulse for the cube sat compared to hydrazine. The the oretical combustion temperature is 2168°F, which is thermally compatible with stainless steel materials of construction and the nickel catalyst used in the tests. Combustion test results are given for stoichiometric, as well as fuel-lean and fuel-rich propellant formulations. These tests focus on thermally and catalytically supported combustion in the microtube, which requires a microtube preheat temperature of about 900°F and sufficient residence time to achieve combustion that is stable and anchored in the tube without blowing through the tube unreacted. Thrust class is expected to be in the 10 m N range based on flow rate and estimated specific impulse. Power requirements are generally below 25 W to heat the tube. Steady state combustion testing is typically conducted for 3 minutes and shows little degradation of the hardware over multiple test cycles indicating that there is little catalyst degradation over time and little catalyst deactivation due to the phosphate stabilizer in the hydrogen peroxide.

Monopropellant↗
DOE OSTI2021

Radiative classical gravitational observables at $ \mathcal{O} $(G 3 ) from scattering amplitudes

We compute classical gravitational observables for the scattering of two spinless black holes in general relativity and N =8 supergravity in the formalism of Kosower, Maybee, and O’Connell (KMOC). We focus on the gravitational impulse with radiation reaction and the radiated momentum in black hole scattering at $ \mathcal{O} $(G 3 ) to all orders in the velocity. These classical observables require the construction and evaluation of certain loop-level quantities which are greatly simplified by harnessing recent advances from scattering amplitudes and collider physics. In particular, we make use of generalized unitarity to construct the relevant loop integrands, employ reverse unitarity, the method of regions, integration-by-parts (IBP), and (canonical) differential equations to simplify and evaluate all loop and phase-space integrals to obtain the classical gravitational observables of interest to two-loop order. The KMOC formalism naturally incorporates radiation effects which enables us to explore these classical quantities beyond the conservative two-body dynamics. From the impulse and the radiated momentum, we extract the scattering angle and the radiated energy. Finally, we discuss universality of the impulse in the high-energy limit and the relation to the eikonal phase.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
DOE OSTI2024

Molecular beam scattering of ammonia from a dodecane flat liquid jet

The evaporation and scattering of ND 3 from a dodecane flat liquid jet are investigated and the results are compared with previous studies on molecular beam scattering from liquid surfaces. Evaporation is well-described by a Maxwell–Boltzmann flux distribution with a cos θ angular distribution at the liquid temperature. Scattering experiments at E i = 28.8 kJ mol -1 over a range of deflection angles show evidence for impulsive scattering and thermal desorption. Further, at a deflection angle of 90°, the thermal desorption fraction is 0.49, which is higher than that of other molecules previously scattered from dodecane and consistent with work performed on NH 3 scattering from a squalane-wetted wheel. ND 3 scattering from dodecane results in super-specular scattering, as seen in previous experiments on dodecane. The impulsive scattering channel is fitted to a “soft-sphere” model, yielding an effective surface mass of 55 amu and an internal excitation of 5.08 kJ mol -1 . Overall, impulsively scattered ND 3 behaves similarly to other small molecules scattered from dodecane.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
DOE OSTI2026

Titanium alloy response sensitivity to variations in spectral reconstructions of National Ignition Facility xenon line-emission x-ray sources

Thermomechanical shock experiments on the National Ignition Facility (NIF) aim to study high strain rate dynamic material response. In such experiments, the NIF laser is used to generate high fluence x-ray emission sources, which irradiate material samples of interest. Under sufficiently high x-ray energy deposition, thermomechanical impulses are generated in the materials. While it is known that the characteristics of x-ray generated impulses vary as a function of incident x-ray spectra, it remains unclear how spectral assumptions and uncertainties in NIF spectral reconstructions affect our interpretation of impulsive loading. Here, in this paper, we simulate the response of a standard titanium alloy baseline sample to synthetic analytically derived and measured NIF xenon line-emission x-ray sources with a radiation hydrodynamics code. We vary the source spectral characteristics based on different source reconstruction techniques to understand the resulting variation in baseline sample response and compare the simulated response with experimental results. We find that the response is highly sensitive to assumptions made about the spectral contents and that knowledge of spectral uncertainties bounds our understanding of the resulting material response. The results of this effort help to extend our ability to use baseline material samples to extract quantitative properties from x-ray experiments on the NIF.

Alloys↗
DOE OSTI2025

Combining quantum noise reduction resources: A practical approach

Optomechanical sensors are capable of transducing external perturbations to resolvable optical signals. A particular regime of interest is that of high-bandwidth force detection, where an impulse is delivered to the system over a short period of time. Exceedingly sensitive impulse detection has been proposed to observe very weak signals like those due to long-range interactions with dark matter that require much higher sensitivities than current sensors can provide. Quantum resources to go beyond the traditional standard quantum limit of these sensors include squeezing of the light used to transduce the signal, backaction evasion by measuring the optimal quadrature, and quantum nondemolition (QND) measurements that reduce backaction directly. These methods have been developed in the context of gravitational wave detection for target frequencies in the audio band range. Here, we provide the theoretical limits to quantum noise reduction for higher and broader frequency targets, such as those from dark matter signals, while combining quantum-enhanced readout techniques based on squeezed light and QND measurements with optomechanical sensors. Here, we demonstrate that backaction evasion through QND techniques dramatically reduces the technical challenges presented when using squeezed light for broadband force detection, paving the way for combining multiple quantum noise reduction techniques for enhanced sensitivity in the context of impulse metrology.

Optomechanics↗
DOE OSTI2024

GraphCH: A Deep Framework for Assessing Cyber-Human Aspects in Insider Threat Detection

Insider threat is one of the most damaging cyber attacks that could cause the loss of intellectual property and enterprise data security breaches. Action sequence data such as host logs are used to investigate such threats and develop anomaly-based AI detectors. However, insider threat actions are similar to legitimate user activities, causing AI detectors to fail and suffer from high false alarm rates. Therefore, user cyber activity logs are inadequate to fully unfold insider threats. In this study, we adopt human psychological principles of risk-taking and impulsiveness along with host data to assess the influence and usefulness of human behavioral aspects in insider threat detection. Here, we hypothesize that individuals' impulsive and risk-taking behavior correlates with cyberspace activities. To validate our hypothesis, we conducted an IRB-approved study recruiting 35 participants who work in a large U.S. university and collected their cyber and psychological data for 90 days. Host and human-behavioral data analysis and mapping indicate that impulsive and risk-taking users trigger more system errors causing (un)intentional insider threats and are susceptible to attackers' social engineering and cognitive hacking. Utilizing cyber-human aspects, we introduce a Cyber-Human Graph Neural Network (GNN) based framework GraphCH to identify abnormal user behaviors and detect insider threats.

97 MATHEMATICS AND COMPUTING↗
DOE OSTI2026

Topographic effects on reflected acoustic waves from the OSIRIS-REx reentry observed from stratospheric balloons

During long-distance sound propagation in planetary atmospheres, acoustic waves may reflect off the air/surface interface one or more times. For low sound frequencies and flat interfaces, the incident and reflected wave tend to be nearly identical. However, this may not be the case when the downgoing acoustic wave encounters topography. Here, we describe a set of direct and reflected acoustic signals recorded on free-flying balloons during the hypersonic entry of the OSIRIS-REx sample return capsule (SRC). In two of the three cases presented here, an impulsive reflected arrival similar in form to the direct sonic boom of the SRC is observed, followed by a diffuse coda. In contrast, one of the floating stations lacked an impulsive reflection entirely, with only the coda present. We use full waveform modeling to show how reflection in the presence of complex topography can explain coda seen in all three examples as well as the lack of impulsive arrival on the third. Our results indicate that the complex signals often observed in long range acoustic propagation could be due, in part, to interactions with topography during transmission.

Lees, Jonathan M. [University of North Carolina, C↗
DOE OSTI2021

Spatial Resolution of a microCT Scanner

We have measured the spatial resolution of the 450 kV test-bed microCT scanner. We have also modeled the resolution analytically. The model and measurement show correspondence within 5%. Our measurement method was to use a thin tungsten wire to act as an impulse function. Therefore, the image is the impulse response, also called the point spread function (PSF) of the system. The Fourier transform (FT) of the impulse response is the modulation transfer function (MTF) of the system. We report the MTF here. By moving the wire at different positions in the scan field of view (FOV), we can measure MTF at different positions in the FOV. Since our imaging system is radially symmetric, we report MTF in two directions, radial and tangential as shown in Figure 1. The radial direction is along the line from isocenter to the wire, and the tangential is perpendicular to it.

42 ENGINEERING↗
DOE OSTI2026

Analysis of Slow Spill Data for Mu2e

The Mu2e experiment requires a constant, relatively low intensity muon beam to produce data with high clarity, which can be achieved using slow extraction. Slow spills/extractions in the Delivery Ring involve contracting and expanding the stable region, which is bordered by the separatrix, of the beam pipe. While this does lower the beam intensity, it is very inconsistent. To help mitigate future inconsistencies, data from many trial spills (some including various magnet impulses to influence the beam intensity) was examined. This involved cutting low quality spills that have abnormal peak and integrated intensities, as well as spills with unusually low magnet ramping. Then, the remaining spills in the datasets were analyzed for trends within spills and across many spills. The findings from this analysis were then given to the FAN-C team to help them develop their simulations, as well as provide training data for their machine learning models that will use beam and impulse data to apply corrective impulses during future slow extractions.

Osborn, Thomas [Purdue U., West Lafayette]↗
NASA NTRS1972

Propellant sidefeed-short pulse discharge thruster studies

The technique of feeding a solid propellant into the discharge from the sides of the discharge was evaluated. The thrust/power ratio could be significantly effected by the included angle of V-shaped propellants and by the electrode length. This result implies that when results are compared at the same specific impulse it is possible to obtain higher thrust efficiencies. In particular, it was found that for a given discharge energy the thrust/power ratio correlated with propellant mass. Increasing the integral simultaneously increases both the gas dynamic and electromagnetic thrust. An analytic expression was formulated for ablated mass which comprehensively describes experimental data in terms of geometry and electrical parameters. The correlation of the product impulse x specific impulse with discharge energy was also described. It is suggested that the reliability of dry energy storage capacitors does not equal the reliability of liquid impregnated units when the comparison is made at the same joules/Kg rating.

Palumbo, D. J.↗
NASA NTRS1972

Optimal trajectories between Earth and Mars in their true planetary orbits.

The optimal transfers from Earth to Mars and from Mars to Earth, considering the actual planetary orbits, are presented as functions of the corresponding idealized Hohmann transfers. The numerically exact two-impulse optimal trajectories are given in graphical form for all possible Hohmann windows. The two-impulse transfers which are absolute optimals and those for which a third impulse provides the absolute optimal are delineated. These data are designed to provide all the information necessary for quick orbit calculations for preliminary Martian mission analysis. In this form, they are as easy to use as the standard Hohmann transfer approximations and provide much greater accuracy.

Gravier, J. P.↗
NASA NTRS1972

Electrothermal hydrazine thruster development.

Results of testing of several electrically-heated, thermal-decomposition hydrazine thrustors in the 5- to 70-mlb thrust range. Propellant supply pressures are compatible with those typical of larger catalytic thrustors. With a 5-W power input, pulsed specific impulse varies from about 165 to 215 sec, depending on duty cycle. The thrustors yield highly reproducible 1-mlb-sec impulse bits with 50-msec command pulse widths. The steady-state specific impulse exceeds 230 sec for thrust levels above 10 mlb.

Murch, C. K.↗
NASA NTRS1972

Location of the electron acceleration region in solar flares.

Observations of impulsive solar flare X rays (energy greater than 10 keV) by the OGO-5 satellite and the measurements of energetic solar electrons made with the Explorer-35 and Explorer-41 (IMP-5) satellites during the period March 1968-September 1969 have been analyzed in order to determine the ion density in the X-ray source region as well as the location of the electron acceleration region in the solar atmosphere. The ion density in the X-ray source region varies from event to event and lies between 1 and 100 billion ions per cu cm for those events in which the impulsive X-ray emission could be detected; for those events in which no impulsive emission was detected above threshold, the ion density in the X-ray source was less than one billion ions per cu cm. At least in some small solar flares, the region where the electrons are accelerated during the flash phase is located in the lower corona.

Kane, S. R.↗
NASA NTRS1973

Simplified procedures for correlation of experimentally measured and predicted thrust chamber performance

Thrust chamber performance is evaluated in terms of an analytical model incorporating all the loss processes that occur in a real rocket motor. The important loss processes in the real thrust chamber were identified, and a methodology and recommended procedure for predicting real thrust chamber vacuum specific impulse were developed. Simplified equations for the calculation of vacuum specific impulse are developed to relate the delivered performance (both vacuum specific impulse and characteristic velocity) to the ideal performance as degraded by the losses corresponding to a specified list of loss processes. These simplified equations enable the various performance loss components, and the corresponding efficiencies, to be quantified separately (except that interaction effects are arbitrarily assigned in the process). The loss and efficiency expressions presented can be used to evaluate experimentally measured thrust chamber performance, to direct development effort into the areas most likely to yield improvements in performance, and as a basis to predict performance of related thrust chamber configurations.

Powell, W. B.↗