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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Thermal Analysis of Iodine Satellite (iSAT)

This paper presents the progress of the thermal analysis and design of the Iodine Satellite (iSAT). The purpose of the iSAT spacecraft (SC) is to demonstrate the ability of the iodine Hall Thruster propulsion system throughout a one year mission in an effort to mature the system for use on future satellites. The benefit of this propulsion system is that it uses a propellant, iodine, that is easy to store and provides a high thrust-to-mass ratio. The spacecraft will also act as a bus for an earth observation payload, the Long Wave Infrared (LWIR) Camera. Four phases of the mission, determined to either be critical to achieving requirements or phases of thermal concern, are modeled. The phases are the Right Ascension of the Ascending Node (RAAN) Change, Altitude Reduction, De-Orbit, and Science Phases. Each phase was modeled in a worst case hot environment and the coldest phase, the Science Phase, was also modeled in a worst case cold environment. The thermal environments of the spacecraft are especially important to model because iSAT has a very high power density. The satellite is the size of a 12 unit cubesat, and dissipates slightly more than 75 Watts of power as heat at times. The maximum temperatures for several components are above their maximum operational limit for one or more cases. The analysis done for the first Design and Analysis Cycle (DAC1) showed that many components were above or within 5 degrees Centigrade of their maximum operation limit. The battery is a component of concern because although it is not over its operational temperature limit, efficiency greatly decreases if it operates at the currently predicted temperatures. In the second Design and Analysis Cycle (DAC2), many steps were taken to mitigate the overheating of components, including isolating several high temperature components, removal of components, and rearrangement of systems. These changes have greatly increased the thermal margin available.

Mauro, Stephanie↗
NASA NTRS1991

Comparative economics of space resource utilization

Physical economic factors such as mass payback ratio, total payback ratio, and capital payback time are discussed and used to compare the economics of using resources from the Moon, Mars and its moons, and near Earth asteroids to serve certain near term markets such as propellant in low Earth orbit or launched mass reduction for lunar and Martian exploration. Methods for accounting for the time cost of money in simple figures of merit such as MPRs are explored and applied to comparisons such as those between lunar, Martian, and asteroidal resources. Methods for trading off capital and operating costs to compare schemes with substantially different capital to operating cost ratio are presented and discussed. Areas where further research or engineering would be extremely useful in reducing economic uncertainty are identified, as are areas where economic merit is highly sensitive to engineering performance - as well as areas where such sensitivity is surprisingly low.

Cutler, Andrew Hall↗
NASA NTRS2002

Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion

The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (lsp~850-1000 s) capability and its attractive engine thrust-to-weight ratio (~3-10). To stay within the available mass and payload volume limits of a “Magnum” heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (~10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as “bimodal” operation), a robust, multiple burn, “power-rich” stage with propulsive Mars capture and reuse capability is possible. A family of modular “bimodal” NTR (BNTR) vehicles are described which utilize a common “core” stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, zero-boiloff liquid hydrogen (LH 2 ) storage, and high data rate communications. An innovative, spine-like “saddle truss” design connects the core stage and payload element and is open underneath to allow supplemental “in-line” propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A “modified” DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step—use of the BNTR for propulsive capture of all payload elements into Mars orbit—the power available in Mars orbit grows to 150 kWe compared to 30 kWe for the DRM. Propulsive capture also eliminates the complex, higher risk aerobraking and capture maneuver which is replaced by a simpler reentry using a standardized, lower mass “aerodescent” shell. The attractiveness of the “all BNTR” option is further increased by the substitution of the lightweight, inflatable “TransHab” module in place of the heavier, hard-shell hab module. Use of TransHab introduces the potential for propulsive recovery and reuse of the BNTR / Earth return vehicle (ERV). It also allows the crew to travel to and from Mars on the same BNTR transfer vehicle thereby cutting the duration of the ERV mission in half—from ~4.7 to 2.5 years. Finally, for difficult Mars options, such as Phobos rendezvous and sample return missions, volume (not mass) constraints limit the performance of the “all LH 2 ” BNTR stage. The use of “LOX-augmented” NTR (LANTR) engines, operating at a modest oxygen-to-hydrogen mixture ratio (MR) of 0.5, helps to increase “bulk” propellant density and total thrust during the TMI burn. On all subsequent burns, the bimodal LANTR engines operate on LH 2 only (MR=0) to maximize vehicle performance while staying within the mass limits of two Magnum launches.

Nuclear thermal rocket↗
NASA NTRS1993

Optimal cooperative power-limited rendezvous between neighboring circular orbits

A minimum-fuel rendezvous of two power-limited spacecraft is investigated. Both vehicles are active and provide thrust to complete the rendezvous. Total propellant consumption is minimized. Analytical solutions are obtained for cooperative rendezvous in the linearized Hill-Clohessy-Wiltshire gravity field. The optimal solutions depend in a complicated way on the power-to-mass ratios of the spacecraft, the initial orbits, and the specified transfer time. For comparison purposes, analytical solutions for active-passive rendezvous in the linearized gravity field are also determined. Cooperative rendezvous requires smaller total propellant consumption, resulting in greater payload capability.

Coverstone-Carroll, Victoria↗
NASA NTRS1990

Ignition and combustion characteristics of metallized propellants

Shakedown and calibration of the experimental apparatus designed to study secondary atomization and ignition characteristics of aluminum slurry propellants was begun and nears completion. The burner was tested over an array of equivalence ratios and oxygen mass fractions to determine the permissible range of operating conditions. A burner flow control system was designed and fabricated which permits independent control of flame stoichiometry and oxidant O2 concentrations. The Sun slurry formulation was selected as the base for initial testing. Laser beam waists at the focal volume of the single particle sizing system were determined to be 396 microns for the Ar-ion and 81 microns for the He-Ne laser. Size calibrations using pinholes were completed and compared to Mie theory for light scattering from particles. A vibrating office droplet generator was used to test the single particle sizing system dynamically by producing a water droplet laden gas flow. The data acquisition and analysis programming was completed and testing is in progress. From initial results, the particle sizing, atomization and ignition detection systems appear to be performing as expected.

Turns, S. R.↗
NASA NTRS2012

Effect of Inductive Coil Geometry on the Operating Characteristics of a Pulsed Inductive Plasma Accelerator

Operational characteristics of two separate inductive thrusters with coils of different cone angles are explored through thrust stand measurements and time-integrated, un- filtered photography. Trends in impulse bit measurements indicate that, in the present experimental configuration, the thruster with the inductive coil possessing a smaller cone angle produced larger values of thrust, in apparent contradiction to results of a previous thruster acceleration model. Areas of greater light intensity in photographs of thruster operation are assumed to qualitatively represent locations of increased current density. Light intensity is generally greater in images of the thruster with the smaller cone angle when compared to those of the thruster with the larger half cone angle for the same operating conditions. The intensity generally decreases in both thrusters for decreasing mass ow rate and capacitor voltage. The location of brightest light intensity shifts upstream for decreasing mass ow rate of propellant and downstream for decreasing applied voltage. Recognizing that there typically exists an optimum ratio of applied electric field to gas pressure with respect to breakdown efficiency, this result may indicate that the optimum ratio was not achieved uniformly over the coil face, leading to non-uniform and incomplete current sheet formation in violation of the model assumption of immediate formation where all the injected propellant is contained in a magnetically-impermeable current sheet.

Hallock, Ashley K.↗
NASA NTRS2012

Effect of Inductive Coil Geometry on the Operating Characteristics of an Inductive Pulsed Plasma Thruster

Operational characteristics of two separate inductive thrusters with conical theta pinch coils of different cone angles are explored through thrust stand measurements and time- integrated, unfiltered photography. Trends in impulse bit measurements indicate that, in the present experimental configuration, the thruster with the inductive coil possessing a smaller cone angle produced larger values of thrust, in apparent contradiction to results of a previous thruster acceleration model. Areas of greater light intensity in photographs of thruster operation are assumed to qualitatively represent locations of increased current density. Light intensity is generally greater in images of the thruster with the smaller cone angle when compared to those of the thruster with the larger half cone angle for the same operating conditions. The intensity generally decreases in both thrusters for decreasing mass flow rate and capacitor voltage. The location of brightest light intensity shifts upstream for decreasing mass flow rate of propellant and downstream for decreasing applied voltage. Recognizing that there typically exists an optimum ratio of applied electric field to gas pressure with respect to breakdown efficiency, this result may indicate that the optimum ratio was not achieved uniformly over the coil face, leading to non-uniform and incomplete current sheet formation in violation of the model assumption of immediate formation where all the injected propellant is contained in a magnetically-impermeable current sheet.

Hallock, Ashley K.↗
NASA NTRSDate not supplied

Analysis of Integrated Spacecraft Performance Using A Rotating Detonation Rocket Engine

For spacecraft with especially stringent mass fraction constraints, the use of a rotating detonation rocket engine (RDRE) may yield potentially significant system mass savings. The pressure gain combustion inherent to RDREs provides higher thrust than traditional rocket engines for a given supply pressure and throat area. This opens multiple avenues for reducing system mass via higher specific impulse, reduced gravity losses, and reduced feed pressures. To highlight the potentially weight-saving characteristics of RDRE systems, the integrated performance of representative spacecraft designs using either a conventional rocket engine or an RDRE was assessed. This paper details the assumptions, methods, and results of this trade study. Two representative spacecraft were assessed. The first is a Liquid Mars Ascent Vehicle (MAV). This spacecraft was designed by NASA in 2011 for the Mars Sample Return Campaign but was not selected for flight due in part to it exceeding the mass requirements by 119 lbm. The second is MIURA-1, a kerolox sounding rocket developed by Payload Aerospace S.L.. The Liquid MAV assessment showed that an RDRE powered design met the mass requirements with a total vehicle mass 46 lbm below the limit. The MIURA-1 assessment showed that wall heat transfer and regenerative cooling are significant challenges for kerolox RDREs, but these challenges may be mitigated by reducing mixture ratio and using both propellants to cool the chamber. The assessment showed that an RDRE powered sounding rocket could achieve 12% more microgravity time than a conventional sounding rocket with the same payload and total vehicle mass. Alternately, if the payload mass and target apogee are held constant, then an RDRE powered sounding rocket can be 8% shorter and 11% lighter than a conventional sounding rocket. These results suggest that the advent of RDREs for in-space propulsion may unlock new missions for which conventional propulsion is not feasible.

Rotating Detonation Rocket Engine↗
NASA NTRSDate not supplied

Analysis of Integrated Spacecraft Performance Using Rotating Detonation Rocket Engine

For spacecraft with especially stringent mass fraction constraints, the use of a rotating detonation rocket engine (RDRE) may yield potentially significant system mass savings. The pressure gain combustion inherent to RDREs provides higher thrust than traditional rocket engines for a given supply pressure and throat area. This opens multiple avenues for reducing system mass via higher specific impulse, reduced gravity losses, and reduced feed pressures. To highlight the potentially weight-saving characteristics of RDRE systems, the integrated performance of representative spacecraft designs using either a conventional rocket engine or an RDRE was assessed. This paper details the assumptions, methods, and results of this trade study. Two representative spacecraft were assessed. The first is a Liquid Mars Ascent Vehicle (MAV). This spacecraft was designed by NASA in 2011 for the Mars Sample Return Campaign but was not selected for flight due in part to it exceeding the mass requirements by 119 lbm. The second is MIURA-1, a kerolox sounding rocket developed by Payload Aerospace S.L.. The Liquid MAV assessment showed that an RDRE powered design met the mass requirements with a total vehicle mass 46 lbm below the limit. The MIURA-1 assessment showed that wall heat transfer and regenerative cooling are significant challenges for kerolox RDREs, but these challenges may be mitigated by reducing mixture ratio and using both propellants to cool the chamber. The assessment showed that an RDRE powered sounding rocket could achieve 12% more microgravity time than a conventional sounding rocket with the same payload and total vehicle mass. Alternately, if the payload mass and target apogee are held constant, then an RDRE powered sounding rocket can be 8% short and 11% lighter than a conventional sounding rocket. These results suggest that the advent of RDREs for in-space propulsion may unlock new missions for which conventional propulsion is not feasible.

Rotating Detonation Rocket Engine↗
NASA NTRS1994

Development of CFD model for augmented core tripropellant rocket engine

The Space Shuttle era has made major advances in technology and vehicle design to the point that the concept of a single-stage-to-orbit (SSTO) vehicle appears more feasible. NASA presently is conducting studies into the feasibility of certain advanced concept rocket engines that could be utilized in a SSTO vehicle. One such concept is a tripropellant system which burns kerosene and hydrogen initially and at altitude switches to hydrogen. This system will attain a larger mass fraction because LOX-kerosene engines have a greater average propellant density and greater thrust-to-weight ratio. This report describes the investigation to model the tripropellant augmented core engine. The physical aspects of the engine, the CFD code employed, and results of the numerical model for a single modular thruster are discussed.

Jones, Kenneth M.↗
NASA NTRS1992

Effect of inert propellant injection on Mars ascent vehicle performance

A Mars ascent vehicle is limited in performance by the propellant which can be brought from Earth. In some cases the vehicle performance can be improved by injecting inert gas into the engine, if the inert gas is available as an in-situ resource and does not have to be brought from Earth. Carbon dioxide, nitrogen, and argon are constituents of the Martian atmosphere which could be separated by compressing the atmosphere, without any chemical processing step. The effect of inert gas injection on rocket engine performance was analyzed with a numerical combustion code that calculated chemical equilibrium for engines of varying combustion chamber pressure, expansion ratio, oxidizer/fuel ratio, and inert injection fraction. Results of this analysis were applied to several candidate missions to determine how the required mass of return propellant needed in low Earth orbit could be decreased using inert propellant injection.

Colvin, James E.↗
NASA NTRS1992

Effect of inert propellant injection on Mars ascent vehicle performance

A Mars ascent vehicle is limited in performance by the amount of propellant which can be brought from earth. In some cases the vehicle performance can be improved by injecting inert gas into the engine, if the inert gas is available as an in situ resource. CO2, N2 and Ar are constituents of the Martian atmosphere which are available at every point on the Martian surface and could be produced by a very simple processing technique, consisting essentially of compressing the atmosphere. The effect of inert gas injection on rocket engine performance was analyzed with a numerical code calculating chemical equilibrium in the engine, for engines of varying combustion chamber pressure, expansion ratio, oxidizer/fuel ratio, and inert injection fraction. Results of this analysis were applied to several candidate missions to determine how the required mass of return propellant needed in LEO could be decreased using inert propellant injection.

Colvin, James E.↗
NASA NTRS1996

Augmentation of Rocket Propulsion: Physical Limits

Rocket propulsion is not ideal when the propellant is not ejected at a unique velocity in an inertial frame. An ideal velocity distribution requires that the exhaust velocity vary linearly with the velocity of the vehicle in an inertial frame. It also requires that the velocity distribution variance as a thermodynamic quantity be minimized. A rocket vehicle with an inert propellant is not optimal, because it does not take advantage of the propellant mass for energy storage. Nor is it logical to provide another energy storage device in order to realize variable exhaust velocity, because it would have to be partly unfilled at the beginning of the mission. Performance is enhanced by pushing on the surrounding because it increases the reaction mass and decreases the reaction jet velocity. This decreases the fraction of the energy taken away by the propellant and increases the share taken by the payload. For an optimal model with the propellant used as fuel, the augmentation realized by pushing on air is greatest for vehicles with a low initial/final mass ratio. For a typical vehicle in the Earth's atmosphere, the augmentation is seen mainly at altitudes below about 80 km. When drag is taken into account, there is a well-defined optimum size for the air intake. Pushing on air has the potential to increase the performance of rockets which pass through the atmosphere. This is apart from benefits derived from "air breathing", or using the oxygen in the atmosphere to reduce the mass of an on-board oxidizer. Because of the potential of these measures, it is vital to model these effects more carefully and explore technology that may realize their advantages.

Taylor, Charles R.↗
NASA NTRSDate not supplied

Framework for Analyzing the Complex Interactions Between Spacecraft Motion and Slosh Dynamics in Low-G Environments

The fuel-to-dry-mass ratio of spacecraft continues to grow as new human spaceflight missions target destinations from Earth. Large amounts of liquid propellant can lead to significant coupling between the rigid-body dynamics of the spacecraft and the motion of the fuel within its storage tank. The present work gives an overview of the dynamic features and a flowchart for a method of simulating the motion of a spacecraft with fuel slosh inside a cylindrical, domed tank in a low-g environment. The method involves modeling the liquid propellant as a particle that transfers momentum to the spacecraft through perfectly inelastic collisions with the tank wall. The foundation of the modeling methodology is the approach taken during the Apollo program used to predict the effect of fuel slosh on the complex motion exhibited by the Service Module following separation from the Crew Module. This paper discusses the motivation, methodology, and conclusions from the Apollo-era method, and subsequently builds upon it by incorporating corrections to the derivation of the dynamics and filling in the gaps due to unavailability of the detailed contractor report and simulation code. The results presented in this paper provide an example that demonstrates the effect that fuel slosh can have on the trajectory of a spacecraft in a low-g environment.

William J Elke III↗
NASA NTRS1979

High performance N2O4/amine elements blowapart

Hypergolic earth-storable N2O4/Amine propellants are used for a wide range of liquid propulsion system applications. These propellants are highly reactive and can experience reactive stream separation (RSS, i.e., blowapart) which can inhibit intraelement mixing, hence reducing the overall spray mixture ratio and mass distribution, which can result in altered combustion efficiency, gas-side heat transfer coefficient, and stability. An experimental and analytical program was conducted to develop an understanding of the mechanisms controlling RSS. The program and product was development of design criteria for coping with RSS to allow the design of high performing, stable injectors. RSS mechanisms were identified using high speed color photography to observe reactive stream mixing of single element injectors tested with N2O4/MMH, N2O4/A-50, and N2O4/N2H4 propellants. The parametric characterization included modeling of the Space Shuttle Orbital Maneuvering System and Reaction Control System engine injectors.

Source record↗
NASA NTRS1974

Mars surface-to-orbit vehicles for sample return missions

Review of the results of a Mars surface-to-orbit launch vehicle study using an upgraded atmospheric density profile and an ascent trajectory optimization program. Orbit insertion was achieved with velocity losses less than those found in an earlier study. Values of thrust-to-mass ratio which minimize velocity losses for dual-burn vehicles were found to be significantly greater than the values of thrust-to-launch-mass ratio for the minimum-velocity-loss continuous-burn vehicle. The use of a two-stage solid-propellant vehicle employing a spin-stabilized second stage offers an attractive design option for a Mars surface-to-orbit launch vehicle.

Weaver, W. L.↗
NASA NTRS2012

Prediction of Liquid Slosh Damping Using a High Resolution CFD Tool

Propellant slosh is a potential source of disturbance critical to the stability of space vehicles. The slosh dynamics are typically represented by a mechanical model of a spring mass damper. This mechanical model is then included in the equation of motion of the entire vehicle for Guidance, Navigation and Control analysis. Our previous effort has demonstrated the soundness of a CFD approach in modeling the detailed fluid dynamics of tank slosh and the excellent accuracy in extracting mechanical properties (slosh natural frequency, slosh mass, and slosh mass center coordinates). For a practical partially-filled smooth wall propellant tank with a diameter of 1 meter, the damping ratio is as low as 0.0005 (or 0.05%). To accurately predict this very low damping value is a challenge for any CFD tool, as one must resolve a thin boundary layer near the wall and must minimize numerical damping. This work extends our previous effort to extract this challenging parameter from first principles: slosh damping for smooth wall and for ring baffle. First the experimental data correlated into the industry standard for smooth wall were used as the baseline validation. It is demonstrated that with proper grid resolution, CFD can indeed accurately predict low damping values from smooth walls for different tank sizes. The damping due to ring baffles at different depths from the free surface and for different sizes of tank was then simulated, and fairly good agreement with experimental correlation was observed. The study demonstrates that CFD technology can be applied to the design of future propellant tanks with complex configurations and with smooth walls or multiple baffles, where previous experimental data is not available.

Yang, H. Q.↗
NASA NTRSDate not supplied

Framework for Analyzing the Complex Interactions Between Spacecraft Motion and Slosh Dynamics in Low-G Environments

The fuel-to-dry-mass ratio of spacecraft continues to grow as new human spaceflight missions target destinations farther from Earth. Large amounts of liquid propellant can lead to significant coupling between the rigid-body dynamics of the spacecraft and the motion of the fuel within its storage tank. The present work gives an overview of the dynamic features and a flowchart for a method of simulating the motion of a spacecraft with fuel slosh inside a cylindrical, domed tank in a low-g environment. The method involves modeling the liquid propellant as a particle that transfers momentum to the spacecraft through perfectly inelastic collisions with the tank wall. The foundation of the modeling methodology is the approach taken during the Apollo program used to predict the effect of fuel slosh on the complex motion exhibited by the Service Module following separation from the Command Module. This paper discusses the motivation, methodology, and conclusions from the Apollo-era method, then presents corrections to the derivation of the dynamics and fills in the gaps left from the unavailability of the detailed contractor report and simulation code. The results presented in this paper provide an example that demonstrates the effect that fuel slosh can have on the trajectory of a spacecraft in a low-g environment.

William J Elke III↗