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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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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Determining Mars parking orbits which ensure in-plane arrival and departure burns

A numerical method to find suitable Mars parking orbits is developed which takes into account geometries associated with the asymptotes, along with the nodal precession caused by the oblateness of Mars. A selected orbital plane which contains the arrival asymptote precesses through the stay time to the plane also containing the departure asymptote. The parking orbit is co-planar with both the arrival and departure asymptotes and only in-plane burns are required at both Mars arrival and departure. The need for a plane change at Mars departure to achieve the proper velocity vector for earth return is eliminated. The method requires very little computation time to determine a set of all possible inclinations and right ascensions of the ascending nodes.

Desai, Prasun N.↗
NASA NTRS2000

Smoke-Point Properties of Nonbuoyant Round Laminar Jet Diffusion Flames

The laminar smoke-point properties of nonbuoyant round laminar jet diffusion flames were studied emphasizing results from long duration (100-230 s) experiments at microgravity carried -out on- orbit in the Space Shuttle Columbia. Experimental conditions included ethylene-and propane-fueled flames burning in still air at an ambient temperature of 300 K, initial jet exit diameters of 1.6 and 2.7 mm, jet exit velocities of 170-1630 mm/s, jet exit Reynolds numbers of 46-172, characteristic flame residence times of 40-302 ms, and luminous flame lengths of 15-63 mm. The onset of laminar smoke-point conditions involved two flame configurations: closed-tip flames with first soot emissions along the flame axis and open-tip flames with first soot emissions from an annular ring about the flame axis. Open-tip flames were observed at large characteristic flame residence times with the onset of soot emissions associated with radiative quenching near the flame tip; nevertheless, unified correlations of laminar smoke-point properties were obtained that included both flame configurations. Flame lengths at laminar smoke-point conditions were well-correlated in terms of a corrected fuel flow rate suggested by a simplified analysis of flame shape. The present steady and nonbuoyant flames emitted soot more readily than earlier tests of nonbuoyant flames at microgravity using ground-based facilities and of buoyant flames at normal gravity due to reduced effects of unsteadiness, flame disturbances and buoyant motion. For example, laminar smoke-point flame lengths from ground-based microgravity measurements were up to 2.3 times longer and from buoyant flame measurements were up to 6.4 times longer than the present measurements at comparable conditions. Finally, present laminar smoke-point flame lengths were roughly inversely proportional to pressure, which is a somewhat slower variation than observed during earlier tests both at microgravity using ground-based facilities and at normal gravity.

Urban, D. L.↗
NASA NTRS2001

TIGER Burned Brightly in JAMIC

The Transition From Ignition to Flame Growth Under External Radiation in 3D (TIGER- 3D) experiment, which is slated to fly aboard the International Space Station, conducted a series of highly successful tests in collaboration with the University of Hokkaido using Japan's 10-sec JAMIC drop tower. The tests were conducted to test engineering versions of advanced flight diagnostics such as an infrared camera for detailed surface temperature measurements and an infrared spectroscopic array for gas-phase species concentrations and temperatures based on detailed spectral emissions in the near infrared. Shown in the top figure is a visible light image and in the bottom figure is an infrared image at 3.8 mm obtained during the microgravity tests. The images show flames burning across cellulose samples against a slow wind of a few centimeters per second (wind is from right to left). These flow velocities are typical of spacecraft ventilation systems that provide fresh air for the astronauts. The samples are ignited across the center with a hot wire, and the flame is allowed to spread upwind and/or downwind. As these images show, the flames prefer to spread upwind, into the fresh air, which is the exact opposite of flames on Earth, which spread much faster downwind, or with the airflow, as in forest fires.

Olson, Sandra L.↗
NASA NTRS1986

A lateral guidance algorithm to reduce the post-aerobraking burn requirements for a lift-modulated orbital transfer vehicle

A lateral guidance algorithm which controls the location of the line of intersection between the actual and desired orbital planes (the hinge line) is developed for the aerobraking phase of a lift-modulated orbital transfer vehicle. The on-board targeting algorithm associated with this lateral guidance algorithm is simple and concise which is very desirable since computation time and space are limited on an on-board flight computer. A variational equation which describes the movement of the hinge line is derived. Simple relationships between the plane error, the desired hinge line position, the position out-of-plane error, and the velocity out-of-plane error are found. A computer simulation is developed to test the lateral guidance algorithm for a variety of operating conditions. The algorithm does reduce the total burn magnitude needed to achieve the desired orbit by allowing the plane correction and perigee-raising burn to be combined in a single maneuver. The algorithm performs well under vacuum perigee dispersions, pot-hole density disturbance, and thick atmospheres. The results for many different operating conditions are presented.

Herman, G. C.↗
NASA NTRS1996

Flow Coupling Effects in Jet-in-Crossflow Flowfields

The combustor designer is typically required to design liner orifices that effectively mix air jets with crossflow effluent. CFD combustor analysis is typically used in the design process; however the jets are usually assumed to enter the combustor with a uniform velocity and turbulence profile. The jet-mainstream flow coupling is usually neglected because of the computational expense. This CFD study was performed to understand the effect of jet-mainstream flow coupling, and to assess the accuracy of jet boundary conditions that are commonly used in combustor internal calculations. A case representative of a plenum-fed quick-mix section of a Rich Burn/Quick Mix/Lean Burn combustor (i.e. a jet-mainstream mass-flow ratio of about 3 and a jet-mainstream momentum-flux ratio of about 30) was investigated. This case showed that the jet velocity entering the combustor was very non-uniform, with a low normal velocity at the leading edge of the orifice and a high normal velocity at the trailing edge of the orifice. Three different combustor-only cases were analyzed with uniform inlet jet profile. None of the cases matched the plenum-fed calculations. To assess liner thickness effects, a thin-walled case was also analyzed. The CFD analysis showed the thin-walled jets had more penetration than the thick-walled jets.

Bain, D. B.↗
NASA NTRS2008

Microgravity Flame Spread in Exploration Atmospheres: Pressure, Oxygen, and Velocity Effects on Opposed and Concurrent Flame Spread

Microgravity tests of flammability and flame spread were performed in a low-speed flow tunnel to simulate spacecraft ventilation flows. Three thin fuels were tested for flammability (Ultem 1000 (General Electric Company), 10 mil film, Nomex (Dupont) HT90-40, and Mylar G (Dupont) and one fuel for flame spread testing (Kimwipes (Kimberly-Clark Worldwide, Inc.). The 1g Upward Limiting Oxygen Index (ULOI) and 1g Maximum Oxygen Concentration (MOC) are found to be greater than those in 0g, by up to 4% oxygen mole fraction, meaning that the fuels burned in 0g at lower oxygen concentrations than they did using the NASA Standard 6001 Test 1 protocol. Flame spread tests with Kimwipes were used to develop correlations that capture the effects of flow velocity, oxygen concentration, and pressure on flame spread rate. These correlations were used to determine that over virtually the entire range of spacecraft atmospheres and flow conditions, the opposed spread is faster, especially for normoxic atmospheres. The correlations were also compared with 1g MOC for various materials as a function of pressure and oxygen. The lines of constant opposed flow agreed best with the 1g MOC trends, which indicates that Test 1 limits are essentially dictated by the critical heat flux for ignition. Further evaluation of these and other materials is continuing to better understand the 0g flammability of materials and its effect on the oxygen margin of safety.

Olson, Sandra L.↗
NASA NTRS1976

A stop-restart solid propellant study with salt quench

Experiments were conducted to gain insight into the unsatisfactory performance of the salt quench system of solid propellants in earlier studies. Nine open-air salt spray tests were conducted and high-speed cinematographic coverage was obtained of the events. It is shown that the salt spray by the detonator is generally a two-step process yielding two different fractions. The first fraction consists of finely powdered salt and moves practically unidirectionally at a high velocity (thousand of feet per second) while the second fraction consists of coarse particles and moves randomly at a low velocity (a few feet per second). Further investigation is required to verify the speculation that a lower quench charge ratio (weight of salt/propellant burning area) than previously employed may lead to an efficient quench

Kumar, R. N.↗
NASA NTRS1994

A laser-based sizing/velocimetry technique to investigate the secondary atomization of aluminum gel propellants

A laser-based, forward-scatter diagnostic technique, employing a single laser sheet, has been developed to simultaneously measure the size and velocity of individual 10-150 micron droplets in a dilute polydisperse droplet stream (less than 1000 particles/cc) and to detect the presence of burning aluminum in these same droplets. Spectral emission from aluminum vapor in the 390-400 nm wavelength region is used as an indication of burning aluminum. The technique utilizes a 4-mm uniformly illuminated probe volume, eliminating trajectory-dependent particle sizing and size-dependent system detection bias. Particle sizing is based on a correlation of particle size with near-forward scattered light intensity. Calculations show average particle sizing variation to be within 3.5% over the expected range of refractive indices. Calibrations using a range of optical pinholes (10-100 micron) were used to verify the above sizing correlation.

Mueller, D. C.↗
NASA NTRS1994

The delayed-detonation model of Type Ia supernovae. 2: The detonation phase

The investigation, by use of two-dimensional numerical hydrodynamics simulations, of the 'delayed detonation' mechanism of Khokhlov for the explosion of Type Ia supernovae is continued. Previously we found that the deflagration is insufficient to unbind the star. Expansion shuts off the flame; much of this small production of iron group nuclei occurs at lower densities, which reduces the electron-capture problem. Because the degenerate star has an adiabatic exponent only slightly above 4/3, the energy released by deflagration drives a pulsation of large amplitude. During the first expansion phase, adiabatic cooling shuts off the burning, and a Rayleigh-Taylor instability then gives mixing of high-entropy ashes with low-entropy fuel. During the first contraction phase, compressional heating reignites the material. The burning was allowed to develop into a detonation in these nonspherical models. The detonation grows toward spherical symmetry at late times. At these densities (rho approx. 10(exp 7) to 10(exp 8) g cm(exp -3)), either Ni-56 or nuclei of the Si-Ca group are the dominant products of the burning. The bulk yields are sensitive to the density of the star when the transition to detonation occurs. The relevance of the abundances, velocities, mixing, and total energy release to the theory and interpretation of Type Ia supernovae is discussed.

Arnett, David↗
NASA NTRS2004

Trajectory Analysis

The Systems Analysis Branch performs trajectory and systems analysis for next-generation launch vehicle and space transportation technologies. Currently the branch is supporting the Project Prometheus with analysis of nuclear electric and nuclear thermal propulsion missions to a variety of destinations. Within Project Prometheus a proposed mission to Jupiter and three of the planet sized icy moons that orbit it is developing. The Jupiter Icy Moons Orbiter (JIMO) as the project is being called will enable detailed scientific investigation of Jupiter's moons Callisto, Ganymede, and Europa. These moons were choose to orbit for intensive study in particular because they are each believed to have water, energy, and organic material. The JIMO mission will utilize nuclear fission power and electric propulsion in order to allow the spacecraft to orbit the moons at close range for long durations of time. My assignment this summer was to assist in developing a trajectory analysis for the spacecraft system the Jupiter Icy Moons Orbiter by rewriting an inefficient Excel file into a more efficient FORTRAN program. This program has been created for use planning the trajectory of the Jupiter Icy Moons Orbiter Mission. The program uses a database of thousands of data points representing flight time, burn time, thrust, mass of the propellant used, final mass of the spacecraft, ratio of finals mass to initial mass, and change in velocity that a spacecraft experiences during each phase in the Jupiter Icy Moons Orbiter Mission. The trajectory program is capable of taking a specific user entered specific impulse (isp), final mass fraction, and thrust (P(sub J)/m0) and through the use of cubic splines to fit specific data curves, the program can locate the exact flight time linked to the user specified values of specific impulse and final over initial mass fraction. Currently, the database used by the program to calculate flight times for a given thrust is only for isps in the range of 2000 to 9000 seconds. Although the program is specific to the Jupiter Icy Moons Orbiter mission, it can be easily modified to fit any mission with a propulsion system expected isp range between 2000 and 9000 seconds. This program performs the same tasks as the old Excel file but in a more timely fashion. Finally, this FORTRAN program will replace the Excel file as the tool used to calculate trajectory properties in missions and will provide valuable data to mission analysts for years to come.

Burke, Laura M.↗
NASA NTRS2019

Operations of the Optical Communications Demonstration for the Orion EM-2 Mission

The GSFC implementation of an Optical Communication System to demonstrate an operational optical communication link for Orion EM-2. It will serve as a base for providing an operational optical communications capability for future Orion missions. GSFC plans to maintain a development path for the optical communication flight terminal to allow commercialization and implementation on future Orion missions. The Orion optical module part of the optical communications flight terminal and its control electronics have a common architecture with the ILLUMA-T optical terminal provided by GSFC for use on the ISS. The plan is to flow data from Orion through the Optical Communication Fight Terminal to the Optical Communication Ground Terminal and reverse. NASA's Orion spacecraft is an exploratory vehicle designed for longer-duration flights beyond the Moon. Following Orion's Exploration Mission-1 (EM-1), during which the spacecraft will travel beyond the Moon, enter a distant retrograde orbit around the Moon and return to Earth unmanned, Exploration Mission-2 (EM-2) will see a crewed spacecraft complete a slightly different flight path. First crewed test flight of the Orion spacecraft, currently targeting a June 30, 2022 launch. The mission involves: One revolution in Low Earth Orbit (LEO) parking orbit to verify basic Orion systems functionality and deploy solar arrays. A single 42-hour Highly Elliptical Orbit (HEO) intermediate checkout orbit allows characterization of the Orion vehicle system performance prior to committing to a cis-lunar flight. Trans Lunar Injection (TLI) burn using Orion Service Module (SM) main engine, which sends Orion on a lunar flyby and free return. Skip reentry at lunar return velocities to splashdown off the coast of San Diego. Total mission duration is approximately 10 days. EM-2 is the first crewed mission of the Orion Spacecraft that crew brings more video up/downloads, file transfers, and real-time chats with family back home and high-rate communications enables live HD streaming for Crew conferences, Public Affairs Office (PAO) events, and significant mission events. Also still collecting vehicle data on numerous Orion subsystems via Development Flight Instrumentation (DFI) allows large data volume returns sooner for DFI and future science payloads, as opposed to waiting for end-of-mission. O2O is a demonstration of operational utility system tested as a Developmental Test Objective (DTO) and not required to meet mission requirements/success/ Flight Test Objectives (FTO). EM-2 architecture is as close to future mission operational architecture as possible.Orion subsystems (video, DFI, etc.) expected to generate ~250 GB of data in the first 24 hours of flight. Total data generated over the mission estimated to be more than 400 GBUsing S-Band alone, Orion limited to ~ 6GB of data downlink per day. Because of this limitation, Orion is planning to limit live video downlinks on EM-1 in order to downlink high priority fileswith 1 hour/day of Optical Communication, Orion could downlink ~6x more data per day (~ 36GB/day).The optical communication system is capable of multiple data rates up to at least 80 Mbps downlink for the transfer of Orion data to Earth while Orion is operating in the lunar vicinity.

Optical Communication↗
NASA NTRS1999

Numerical Simulations of Acoustically Driven, Burning Droplets

This computational study focuses on understanding and quantifying the effects of external acoustical perturbations on droplet combustion. A one-dimensional, axisymmetric representation of the essential diffusion and reaction processes occurring in the vicinity of the droplet stagnation point is used here in order to isolate the effects of the imposed acoustic disturbance. The simulation is performed using a third order accurate, essentially non-oscillatory (ENO) numerical scheme with a full methanol-air reaction mechanism. Consistent with recent microgravity and normal gravity combustion experiments, focus is placed on conditions where the droplet is situated at a velocity antinode in order for the droplet to experience the greatest effects of fluid mechanical straining of flame structures. The effects of imposed sound pressure level and frequency are explored here, and conditions leading to maximum burning rates are identified.

Kim, H.-C.↗
NASA NTRS1979

Investigation of critical burning of fuel droplets

The general problem of spray combustion was investigated. The combustion of bipropellent droplets; combustion of hydrozine fuels; and combustion of sprays were studied. A model was developed to predict mean velocities and temperatures in a combusting gas jet.

Faeth, G. M.↗
NASA NTRS1973

Effect of premixing on nitric oxide formation

Emissions from a simple 10-cm (4-in.) diameter tube combustor burning a premixed, gaseous propane/air mixture were measured. Inlet conditions included a temperature of 590 K (600 F), pressure of 5.5 atm, and reference velocity of 23 m/s (75 ft/s) for a range of equivalence ratios from the lean limit to slightly richer than stoichiometric. A nitric oxide emission index of 1 g NO2/kg fuel was measured for an equivalence ratio of 0.57.

Anderson, D. N.↗
NASA NTRS1974

Film cooling in a combustor operating at fuel-rich exit conditions

Data were taken with a film-cooled test plate placed in the exhaust stream of a rectangular combustor. Results showed that in a fuel-rich zone, fuel entrained into the film-cooling air would burn at the conditions tested. Test conditions were cooling-gas flow rates, 9.5 and 5 percent of total gas flow; cooling-gas velocities, 23 and 12 m/sec; ambient-temperature cooling gas; hot-gas velocity, nominally 215 m/sec; fuel-air ratio to stoichiometric fuel-air ratio values of 0.1, 0.8, 1.0, 1.2, and 1.35, resulting in hot-gas temperatures from 590 to 2100 K; and atmospheric pressure. Analytical prediction of wall temperatures agreed reasonably well with experimental results.

Tacina, R. R.↗
NASA NTRS1988

Modelling of time-dependent ion outflows at high geomagnetic latitudes

The output from the NCAR Thermospheric General Circulation Model was used to simulate the time and altitude-dependent F-region ion frictional heating rate. These ion-heating profiles, computed as a functionm of UT along the locus of specific flux tubes, were used as boundary conditions in the Gombosi et al. (1985) polar wind model to simulate the plasma heating history of a flux tube traversing the cusp region. Using derived time-varying frictional heating rates such as those experienced by these flux tubes, it is shown that transverse ion heating below 500 km can provide sufficient energy to perturb the velocity distribution of the major ion species.

Cannata, R. W.↗
DOE OSTI2020

An Eulerian multimaterial framework for simulating high-explosive aquarium tests

Aquarium tests of cylindrical high-explosive charges provide optical data of the detonation front velocity and shape, propagation of the shock wave in the surrounding water, and expansion rates of the detonation products behind the front. Data from aquarium experiments is often used for calibration of reactive burn models based on phenomenological equations of state (EOS) and reaction rate laws. This paper presents a multimaterial numerical modeling framework to solve the 2D axisymmetric reactive Euler equations for high-explosive aquarium tests, in particular for ammonium nitrate - fuel oil (ANFO) explosives. An extension of the Ghost Fluid Method (GFM) is used to handle the dynamic material interfaces for the ANFO explosion products, the charge-confining material (polymethyl methacrylate PMMA), and the surrounding water. This study analyzes the sensitivity of calculations (both computational efficiency and numerical accuracy) to different algorithms for the material interface models including the original GFM versus Riemann solver-based strategies. A novel method for defining the left and right states in the interfacial Riemann problem eliminates the need for sorting or nodal interpolation during the projection along the material interface. Numerical tests indicate that populating the interface node values using the Riemann solution mitigate the overheating error observed in steady-state calculations. Solution convergence and computational efficiency are explored as a function of the spatial and temporal order of the schemes. Results from the computational model with analytical equations of state and fitted reaction rate parameters show very good quantitative agreement with experimentally observed detonation front velocity, reaction products expansion, and shock wave propagation in the surrounding water for a cylindrical ANFO charge encased in PMMA. Finally, the proposed modeling framework, in conjunction with experimental tests, provides a reliable tool to assess equations of state and reaction rate expressions for reactive burn models of confined high explosives.

42 ENGINEERING↗
DOE OSTI2021

SURFplus model calibration for PBX 9012

PBX 9012 is a plastic bonded explosive composed of 90.2 wt % HMX (cyclo-tetramethylene-tetranitramine) and 9.8 % Viton A (vinylidene-hexafluoropropylene copolymer). It has nearly the same composition as LX-07 (90% HMX and 10% Viton A) but with a slightly lower density. Hence it has a higher porosity and is slightly more shock sensitive. Here we describe the SURFplus model calibration for PBX 9012 using data from five shock-to-detonation transition (SDT) gas gun experiments by Burns and Chiquete, detonation front shape data from rate stick experiments and cylinder test experiment reported in Anderson et al., Jackson et al. The SDT experiments provide Pop plot data points (run distance-to-detonation) which characterize shock initiation, Lagrangian velocity time histories from embedded magnetic velocity gauges which give information on the reaction behind the lead shock, and Hugoniot data points for the reactants EOS. The rate stick experiments provide data on the curvature effect which characterize propagating detonation waves. The cylinder test experiments provide data on the detonation release isentrope used to calibrate the products EOS. The HE model uses a Davis reactants EOS with parameters that Burns and Chiquete [2020] fit to Hugoniot data, and Davis products EOS with parameters that Anderson et al. [2020] fit to cylinder test data.

45 MILITARY TECHNOLOGY, WEAPONRY, AND NATIONAL DEF↗