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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

3-D CFD Simulation and Validation of Oxygen-Rich Hydrocarbon Combustion in a Gas-Centered Swirl Coaxial Injector using a Flamelet-Based Approach

Injector design is a critical part of the development of a rocket Thrust Chamber Assembly (TCA). Proper detailed injector design can maximize propulsion efficiency while minimizing the potential for failures in the combustion chamber. Traditional design and analysis methods for hydrocarbon-fuel injector elements are based heavily on empirical data and models developed from heritage hardware tests. Using this limited set of data produces challenges when trying to design a new propulsion system where the operating conditions may greatly differ from heritage applications. Time-accurate, Three-Dimensional (3-D) Computational Fluid Dynamics (CFD) modeling of combusting flows inside of injectors has long been a goal of the fluid analysis group at Marshall Space Flight Center (MSFC) and the larger CFD modeling community. CFD simulation can provide insight into the design and function of an injector that cannot be obtained easily through testing or empirical comparisons to existing hardware. However, the traditional finite-rate chemistry modeling approach utilized to simulate combusting flows for complex fuels, such as Rocket Propellant-2 (RP-2), is prohibitively expensive and time consuming even with a large amount of computational resources. MSFC has been working, in partnership with Streamline Numerics, Inc., to develop a computationally efficient, flamelet-based approach for modeling complex combusting flow applications. In this work, a flamelet modeling approach is used to simulate time-accurate, 3-D, combusting flow inside a single Gas Centered Swirl Coaxial (GCSC) injector using the flow solver, Loci-STREAM. CFD simulations were performed for several different injector geometries. Results of the CFD analysis helped guide the design of the injector from an initial concept to a tested prototype. The results of the CFD analysis are compared to data gathered from several hot-fire, single element injector tests performed in the Air Force Research Lab EC-1 test facility located at Edwards Air Force Base.

Richardson, Brian↗
NASA NTRS1966

Secondary electron emission

Secondary electron emission from porous tungsten surfaces to determine mechanisms of beam enhancement in ion rocket thrust augmentor

ION BEAM↗
NASA NTRS1968

Fundamental problems of injector design

The principal control of the combustion within a liquid-propellant rocket thrust chamber is provided by the injector. Only when precise control of the injection process is attained can optimum performance, dynamic stability, heat transfer compatibility and reproducibility be provided. Control must be designed into each step of the injection process with sufficient isolation provided to insure stability. In the design of the injector manifolding, the main desire is to provide mass and mixture ratio regulation across the injector face. The design of the injector orifices must be such that this mass and mixture ratio control is retained through the orifices and imparted to the droplet sprays that are produced. Furthermore, the orifice design must control the size and distribution of the droplets generated so that combustion will be completed within the chamber. Mechanical damping devices are often required to protect the early mixing zone from transverse velocities and associated displacement effects in the chamber. Only when each phase of the injection process is carefully controlled can the design aims be reached.

David T Harrje↗
NASA NTRS1972

Mariner 9 propulsion subsystem performance during interplanetary cruise and Mars orbit insertion

On 14 November 1971 the Mariner 9 1334-N-(300-lbf)-thrust rocket engine was fired for just over 15 min to place the first man-made satellite into orbit about Mars. Propulsion subsystem data gathered during the 5-month interplanetary cruise and orbit insertion are of significance to future missions of this type. Specific results related to performance predictability, zero g heat transfer, and nitrogen permeation, diffusion, and solubility values are presented.

Cork, M. J.↗
NASA NTRS1971

Fabrication methods for TD nickel.

Spin-forming, welding, machining, electric discharge machining, and electroforming operations in th e construction of a TD Nickel rocket thrust chamber are described. TD Nickel was one of the first commercially available dispersion hardened alloys. Dispersion hardened alloys are being given prominent consideration in the construction of the space shuttle; therefore, fabrication experience gained with the TD Nickel material is of current interest.

Leach, A. E.↗
NASA NTRS1972

Boundary cooled rocket engines for space storable propellants

An evaluation of an existing analytical heat transfer model was made to develop the technology of boundary film/conduction cooled rocket thrust chambers to the space storable propellant combination oxygen difluoride/diborane. Critical design parameters were identified and their importance determined. Test reduction methods were developed to enable data obtained from short duration hot firings with a thin walled (calorimeter) chamber to be used quantitatively evaluate the heat absorbing capability of the vapor film. The modification of the existing like-doublet injector was based on the results obtained from the calorimeter firings.

Kesselring, R. C.↗
NASA NTRS1973

Coolant-side heat-transfer rates for a hydrogen-oxygen rocket and a new technique for data correlation

An experimental investigation was conducted to determine the coolant-side, heat transfer coefficients for a liquid cooled, hydrogen-oxygen rocket thrust chamber. Heat transfer rates were determined from measurements of local hot gas wall temperature, local coolant temperature, and local coolant pressure. A correlation incorporating an integration technique for the transport properties needed near the pseudocritical temperature of liquid hydrogen gives a satisfactory prediction of hot gas wall temperatures.

Schacht, R. L.↗
NASA NTRS1973

Mechanical property determination of high conductivity metals and alloys

Pertinent mechanical properties of three high conductivity metals and alloys; namely, vacuum hot pressed grade S-200E beryllium, OFHC copper and beryllium-copper alloy no. 10 were determined. These materials were selected based on their possible use in rocket thrust chamber and nozzle hardware. They were procured in a form and condition similar to that which might be ordered for actual hardware fabrication. The mechanical properties measured include (1) tension and compression stress strain curves at constant strain rate (2) tensile and compressive creep, (3) tensile and compressive stress-relaxation behavior and (4) elastic properties. Tests were conducted over the temperature range of from 75 F to 1600 F. The resulting data is presented in both graphical and tabular form.

Harrod, D. L.↗
NASA NTRS1973

Property investigation and sputter deposition of dispersion-hardened copper for fatigue specimen fabrication

Sputter-deposited alloys of dispersion-hardenable Cu-0.25 vol% SiC and Cu-0.50 vol% SiC and precipitation-hardenable Cu-0.15 wt% Zr and Cu-0.05 wt% Mg-0.15 wt% Zr-0.40 wt% Cr were investigated for selection to evaluate fatigue specimen performance with potential application in fabricating regeneratively cooled rocket thrust chambers. Yield strengths in the 700 to 1000-MN/sq m range were observed with uniform elongation ranging from 0.5 to 1.5% and necking indicative of greater ductility. Electrical conductivity measured as an analog to thermal conductivity gave values 90% IACS for Cu-0.15 wt% Zr and Cu-0.05 wt% Mg-0.15 wt% Zr-0.40 wt% Cr. A 5500-g sputtered deposit of Cu-0.15 wt% Zr alloy, 12.29 mm (0.484 in.) average thickness in the fatigue specimen gage length, was provided to NASA on one of their substrates.

Mcclanahan, E. D.↗
NASA NTRS1974

Castable thermal insulation for use as heat shields

Structural members supporting the afterburners of high thrust rocket engines are subjected to extreme heating, along with severe vibration and high acceleration levels during early lift-off. Chemically-bonded, castable, zircon composite foams were developed and successfully tested to meet specific, laboratory simulated lift-off conditions.

Mountvala, A. J.↗