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At least 1,279 records · Page 71

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NASA NTRS2008

Mars Rocket Propulsion System

A report discusses the methane and carbon monoxide/LOX (McLOx) rocket for ascent from Mars as well as other critical space propulsion tasks. The system offers a specific impulse over 370 s roughly 50 s higher than existing space-storable bio-propellants. Current Mars in-situ propellant production (ISPP) technologies produce impure methane and carbon monoxide in various combinations. While separation and purification of methane fuel is possible, it adds complexity to the propellant production process and discards an otherwise useful fuel product. The McLOx makes such complex and wasteful processes unnecessary by burning the methane/CO mixtures produced by the Mars ISPP systems without the need for further refinement. Despite the decrease in rocket-specific impulse caused by the CO admixture, the improvement offered by concomitant increased propellant density can provide a net improvement in stage performance. One advantage is the increase of the total amount of propellant produced, but with a decrease in mass and complexity of the required ISPP plant. Methane/CO fuel mixtures also may be produced by reprocessing the organic wastes of a Moon base or a space station, making McLOx engines key for a human Lunar initiative or the International Space Station (ISS) program. Because McLOx propellant components store at a common temperature, very lightweight and compact common bulkhead tanks can be employed, improving overall stage performance further.

Zubrin, Robert↗

NASA NTRS2007

Modeling and Parameter Estimation of Spacecraft Fuel Slosh with Diaphragms Using Pendulum Analogs

Prediction and control of liquid slosh in moving containers is an important consideration in the design of spacecraft and launch vehicle control systems. Even with modern computing systems, CFD type simulations are not fast enough to allow for large scale Monte Carlo analyses of spacecraft and launch vehicle dynamic behavior with slosh included. It is still desirable to use some type of simplified mechanical analog for the slosh to shorten computation time. Analytic determination of the slosh analog parameters has met with mixed success and is made even more difficult by the introduction of propellant management devices such as elastomeric diaphragms. By subjecting full-sized fuel tanks with actual flight fuel loads to motion similar to that experienced in flight and measuring the forces experienced by the tanks, these parameters can be determined experimentally. Currently, the identification of the model parameters is a laborious trial-and-error process in which the hand-derived equations of motion for the mechanical analog are evaluated and their results compared with the experimental results. This paper will describe efforts by the university component of a team comprised of NASA's Launch Services Program, Embry Riddle Aeronautical University, Southwest Research Institute and Hubert Astronautics to improve the accuracy and efficiency of modeling techniques used to predict these types of motions. Of particular interest is the effect of diaphragms and bladders on the slosh dynamics and how best to model these devices. The previous research was an effort to automate the process of slosh model parameter identification using a MATLAB/SimMechanics-based computer simulation. These results are the first step in applying the same computer estimation to a full-size tank and vehicle propulsion system. The introduction of diaphragms to this experimental set-up will aid in a better and more complete prediction of fuel slosh characteristics and behavior. Automating the parameter identification process will save time and thus allow earlier identification of potential vehicle performance problems.

Chatman, Yadira↗

NASA NTRSDate not supplied

Development of A Maximum Anti-Slosh Baffle Pressure Load Model

The sloshing of propellants can affect the stability of a spacecraft and the integrity of the tank structure. Undesirable sloshing can be controlled by the addition of anti-slosh baffles, and the spacing and configuration of baffles are driven by damping requirements. The structural design of the baffle is determined after consideration of many factors, such as the strength and rigidity needed to support the baffle for its lifetime. Therefore, knowledge of distributed pressure loading is important for detailed structural design. In addition, the resultant force and moment produced by the distributed pressure are of direct importance to the design of a vehicle’s control system. Previous experimental investigations have been conducted to determine the liquid pressure loads and slosh damping associated with a rigid ring baffle. The results suggested that when the nondimensional velocity parameter is larger than 1.0, the theories agree with the test. However, when the velocity parameter is less than 1.0, all theories are nonconservative and under-predict the pressure loads. The present study has derived a maximum pressure load on the slosh baffle based on the energy conservation principle. It is verified from CFD that pressure in the slosh flow field can be decomposed into static and transient components. The CFD results confirm that there is a phase shift in pressure across the baffle, which depends on the fluid damping. Higher damping leads to a higher phase shift. The CFD investigation further verifies the proposed theory: the maximum pressure load occurs when the phase shift is 90 degrees. A comparison of the present computational results to the previous comprehensive experimental data validates the maximum pressure theory. When the baffle is submerged, the maximum pressure theory envelopes all the experimental data points.

CFD↗

NASA NTRS2018

Demonstration of a 600 Watt Hybrid Iodine-Xenon Electric Propulsion System

This paper reviews recent iodine compatible electric propulsion technologies evaluated at the NASA Glenn Research Center. The work culminated in a 1,174-hour hybrid iodine-xenon propulsion system durability demonstration (iodine fed thruster with xenon fed cathode). The test demonstrated that (i) a Hall-effect thruster operates with nearly identical performance whether employing iodine or xenon propellant, (ii) careful selection of propulsion system materials and coatings can result in durable iodine-compatible hardware, and (iii) implementation of appropriate facility improvements and procedures can limit negative impacts of iodine on test hardware and ground support equipment. The work was motivated by advancements in capability of very dense small-spacecraft (<500 kg), but spacecraft still lacking very dense in-space propulsion to provide the large delta-v capabilities (3 to 10 km/s) required for many high-value NASA exploration and science missions. Volume limitations of small-spacecraft not only require dense propulsion systems, but more so dense propellants. NASA identifies xenon and iodine as having both favorable storage densities and propulsive properties to enable many NASA small-spacecraft mission scenarios. While xenon is inert and well-proven in spaceflight applications, iodine has triple the storage density of xenon and stores at low pressures, permitting use of conformal tank designs. Unlike xenon, iodine does raise concerns related to its reactivity with most materials, potential spacecraft-propellant interactions, impact on ground test facilities, and challenges to acceptance test iodine propulsion systems prior to flight. This work begins to address these challenges. This work was conducted under the Advanced In-Space Propulsion (AISP) project funded through the Game Changing Development (GCD) program within NASA's Science Technology Mission Directorate (STMD).

Benavides, G. F.↗

NASA NTRS1974

A feasibility study of developing toroidal tanks for a spinning spacecraft

The use of toroidal tanks in an advanced Pioneer, spin-stabilized Jupiter orbiter was investigated. Fluid behavior in spinning toroids was evaluated and the need for a propellant acquisition system was established. A common surface tension acquisition system design was selected from four candidates for gas-free liquid feed to a single outlet in both MMH and N2O4 tanks. Fabrication of toroidal tanks using either titanium or aluminum alloys is feasible. Integration of two toroidal tanks with surface tension systems in place of four spherical tanks provides spacecraft weight, packaging, structural mounting, thermal control and stability advantages.

Anderson, J. E.↗

NASA NTRS1966

Stability of nonlinear vehicle systems

Nonlinear effects of liquid motion in fixed and moving containers on stability and control of liquid rocket powered missiles and spacecraft

LIQUID PROPELLANT↗

NASA NTRS1974

Developments in low-gravity cryogenic propellant acquisition and thermal control systems

Development programs are described for two low-gravity cryogenic propellant acquisition systems: a full-screen liner system, and a refillable start-tank. First, based on experimental LH2 screen flow data, the fluid-dynamic characteristics of an integrated screen/thermodynamic vent system were analyzed for a design compared to 17.5 cu fg supercritical life-support systems, and for a Tug orbital LH2/LQ2 resupply system. Second, the analysis, design, fabrication, and test of a 10 cu ft, annular-screen start-tank with a wall-mounted thermodynamic vent was performed. The test program demonstrated that the system expelled LH2 against 1-g and performed properly over a range of conditions.

Cady, E. C.↗

NASA NTRS1986

Modular Firewalls for Storage Areas

Giant honeycomb structures assembled in modular units. Flammable materials stored in cells. Walls insulated with firebrick to prevent spread of fire among cells. Portable, modular barrier withstands heat of combustion for limited time and confines combustion products horizontally to prevent fire from spreading. Barrier absorbs heat energy by ablation and not meant to be reused. Designed to keep fires from spreading among segments of solid rocket propellant in storage, barrier erected between storage units of other flammable or explosive materials; tanks of petroleum or liquid natural gas. Barrier adequate for most industrial purposes.

Fedor, O. H.↗

NASA NTRS1985

Development and life-testing of 10 kW class thermal arcjet engines

A facility to develop and test thermal arcjet engines over extended periods of time has been constructed and is described in this paper. It consists of a large vacuum tank, high capacity vacuum pumps, a 100 kW power supply and a large ammonia propellant storage and delivery system. The facility is instrumented to measure electrical power dissipated in the engine, propellant mass flow rate and developed thrust. Pressures and temperatures up to 2400 K can also be measured. The entire facility is computer-controlled and can be operated unattended for many weeks. Two 30 kW thermal arcjet engines that have been designed, built and are being tested in this facility are also described.

Pivirotto, T. J.↗

NASA NTRS1993

Advanced transportation system studies technical area 3: Alternate propulsion system concepts. SSME upper stage use

The main objective was to determine viable methods for starting the Space Shuttle Main Engine (SSME) in an altitude environment and restarting it in an orbit environment with minimum changes in utilization of the engine system or hardware. The study concluded that the use of the SSME in an upper stage is feasible with minimal changes to the engine systems. The altitude start case requires only a change in the valve sequencing during start and reorificing of the ASI lines. Inlet pressures can be moderately low at 40 psia for the LOX and 32 psia for the H2. The orbital restart case adds the need to recirculate propellant and thermal control paint (to keep the turbomachinery inlets cold to minimize the tank pressures needed), and the need to heat two small components (to maintain acceptable mixture ratios during the early part of the start). These actions allow start anytime after approximately 120 minutes. Earlier starts (approximately one hour) are also possible but would require additional component heating for mixture ratio control during the early portion of the start sequence.

Strangeland, Eric↗

NASA NTRS1995

A Summary of the Slush Hydrogen Technology Program for the National Aero-Space Plane

Slush hydrogen, a mixture of solid and liquid hydrogen, offers advantages of higher density (16 percent) and higher heat capacity (18 percent) than normal boiling point hydrogen. The combination of increased density and heat capacity of slush hydrogen provided a potential to decrease the gross takeoff weight of the National Aero-Space Plane (NASP) and therefore slush hydrogen was selected as the propellant. However, no large-scale data was available on the production, transfer and tank pressure control characteristics required to use slush hydrogen as a fuel. Extensive testing has been performed at the NASA Lewis Research Center K-Site and Small Scale Hydrogen Test Facility between 1990 and the present to provide a database for the use of slush hydrogen. This paper summarizes the results of this testing.

Mcnelis, Nancy B.↗

NASA NTRS2000

Helium Evolution from the Transfer of Helium Saturated Propellant in Space

Helium evolution from the transfer of helium saturated propellant in space is quantified to determine its impact from creating a two-phase mixture in the transfer line. The transfer line is approximately 1/2 inch in diameter and 2400 inches in length comprised of the Fluid Interconnect System (FICS), the Orbiter Propellant Transfer System (OPTS) and the International Space Station (ISS) Propulsion Module (ISSPM). The propellant transfer rate is approximately two to three gallons per minute, and the supply tank pressure is maintained at approximately 250 psig.

Nguyen, Bich N.↗

NASA NTRS2003

Numerical Modeling, Thermomechanical Testing, and NDE Procedures for Prediction of Microcracking Induced Permeability of Cryogenic Composites

Reusable Space Vehicles will include light cryogenic composite fuel tanks that must not leak excessively even after multiple launches. Damage in cryogenic composite fuel tanks induced during manufacturing and advanced by thermomechanical cycling can accelerate leakage of the propellant. Whether the leakage exceeds tolerable levels depends on many factors, including pressure gradients, microcrack density, other damage such as delamination, connectivity of the cracks, residual stresses from manufacture, service-induced stresses from thermal and mechanical loads, and composite lay-up. Although it is critical to experimentally characterize permeability during various thermal and mechanical load histories, optimal design depends on having analytical models that can predict the effect of various parameters on performance. Our broad goal is to develop such models that are experimentally validated by destructive and non-destructive evaluation means.

Noh, Jae↗

NASA NTRS2006

Development and Implementation of Non-Newtonian Rheology Into the Generalized Fluid System Simulation Program (GFSSP)

One of the goals of this program was to develop the experimental and analytical/computational tools required to predict the flow of non-Newtonian fluids through the various system components of a propulsion system: pipes, valves, pumps etc. To achieve this goal we selected to augment the capabilities of NASA's Generalized Fluid System Simulation Program (GFSSP) software. GFSSP is a general-purpose computer program designed to calculate steady state and transient pressure and flow distributions in a complex fluid network. While the current version of the GFSSP code is able to handle various systems components the implicit assumption in the code is that the fluids in the system are Newtonian. To extend the capability of the code to non-Newtonian fluids, such as silica gelled fuels and oxidizers, modifications to the momentum equations of the code have been performed. We have successfully implemented in GFSSP flow equations for fluids with power law behavior. The implementation of the power law fluid behavior into the GFSSP code depends on knowledge of the two fluid coefficients, n and K. The determination of these parameters for the silica gels used in this program was performed experimentally. The n and K parameters for silica water gels were determined experimentally at CFDRC's Special Projects Laboratory, with a constant shear rate capillary viscometer. Batches of 8:1 (by weight) water-silica gel were mixed using CFDRC s 10-gallon gelled propellant mixer. Prior to testing the gel was allowed to rest in the rheometer tank for at least twelve hours to ensure that the delicate structure of the gel had sufficient time to reform. During the tests silica gel was pressure fed and discharged through stainless steel pipes ranging from 1", to 36", in length and three diameters; 0.0237", 0.032", and 0.047". The data collected in these tests included pressure at tube entrance and volumetric flowrate. From these data the uncorrected shear rate, shear stress, residence time, and viscosity were evaluated using formulae for non-Newtonian, power law fluids. The maximum shear rates (corrected for entrance effects) obtained in the rheometer with the current setup were in the 150,000 to 170,000sec- range. GFSSP simulations were performed with a flow circuit simulating the capillary rheometer and using Power Law gel viscosity coefficients from the experimental data. The agreement between the experimental data and the simulated flow curves was within +/-4% given quality entrance effect data.

DiSalvo, Roberto↗

NASA NTRS2017

Additive Manufacturing: An Enabling Technology for the MoonBEAM 6U CubeSat Missions

The Advanced Concepts Office at the NASA Marshall Space Flight Center completed a mission concept study for the Moon Burst Energetics All-sky Monitor (MoonBEAM). The goal of the concept study was to show the enabling aspects that additive manufacturing can provide to CubeSats. In addition to using the additively manufactured tanks as part of the spacecraft structure, the main propulsion system uses a green propellant, which is denser than hydrazine. Momentum unloading is achieved with electric microthrusters, eliminating much of the propellant plumbing. The science mission, requirements, and spacecraft design are described.

Hopkins, R. C.↗

NASA NTRS1981

Propellant management for low thrust chemical propulsion systems

Low-thrust chemical propulsion systems (LTPS) will be required for orbital transfer of large space systems (LSS). The work reported in this paper was conducted to determine the propellant requirements, preferred propellant management technique, and propulsion system sizes for the LTPS. Propellants were liquid oxygen (LO2) combined with liquid hydrogen (LH2), liquid methane or kerosene. Thrust levels of 100, 500, and 1000 lbf were combined with 1, 4, and 8 perigee burns for transfer from low earth orbit to geosynchronous earth orbit. This matrix of systems was evaluated with a multilayer insulation (MLI) or a spray-on-foam insulation. Vehicle sizing results indicate that a toroidal tank configuration is needed for the LO2/LH2 system. Multiple perigee burns and MLI allow far superior LSS payload capability. Propellant settling, combined with a single screen device, was found to be the lightest and least complex propellant management technique.

Hamlyn, K. M.↗

NASA NTRS1979

Hydrazine monopropellant reciprocating engine development

A hydrazine fueled piston engine for providing 11.2 kW was developed to satisfy the need for an efficient power supply in the range from 3.7 to 74.6 kW where existing nonair-breathing power supplies such as fuel cells or turbines are inappropriate. The engine was developed for an aircraft to fly to 21.3 km and above and cruise for extended periods. A remotely piloted aircraft and the associated flight control techniques for this application were designed. The engine is geared down internally (2:1) to accommodate a 1.8 m diameter propeller. An alternator is included to provide electrical power. The pusher-type engine is mounted onto the aft closure of the fuel tank, which also provides mounting for all other propulsion equipment. About 20 hrs of run time demonstrated good efficiency and adequate life. One flight test to 6.1 km was made using the engine with a small fixed-pitch four-bladed propeller. The test was successful in demonstrating operational characteristics and future potential.

Akkerman, J. W.↗

NASA NTRS1966

Flexible baffles for slosh damping.

Flexible baffle effect on nature and magnitude of damping and sloshing motions of liquid propellants over wide range of conditions

FUEL TANK↗