Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “zirconium”

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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1965

Solubility studies in alkali metals

Solubility parameters for niobium and zirconium in lithium and potassium, and calcium solubility in eutectic sodium potassium alloy

SOLUBILITY↗
NASA NTRSDate not supplied

Additive Manufacturing of ZrC for Nuclear Thermal Propulsion Applications

Looking into the future, we are seeing that nuclear is becoming a more and more viable fuel source for space applications due to it being one of the safest energy forms, low greenhouse gas emissions, and inexpensive maintenance. National Aeronautics and Space Administration (NASA) and a few others are looking at nuclear fuels in relation to space travel, specifically in the form of Nuclear Thermal Propulsion (NTP) [1]. NTP provides the answer for space travel to Mars where chemical rockets would not complete the job. A concept art of what these spacecrafts could look like is shown (Figure 1). Uranium Carbide fuel has been chosen to heat a liquid propellant for propulsion. The complex fuel geometries required by space reactors can be achieved by novel techniques fairly easily, cheaply, and safely. In this specific study, we are looking at using Zirconium Carbide (ZrC) to use as a surrogate for Uranium Carbide (UC) due to their similar behaviors under high temperatures. An ink made with ZrC will be made and will be extruded into unique, complex geometries. This is a great starting point to test additive manufacturing on these ceramics.

Additive Manufacturing↗
NASA NTRS2023

Sintering Behaviors of ZrC, NbC, and TaC Mono-and Binary Carbides

Nuclear Thermal Propulsion (NTP) has undergone development as an alternate in-space propulsion system to traditional chemical propulsion methods since the 1950s. In an NTP system, the energy released from fission in the core is utilized as the heat source to directly heat a propellant for propulsion, rather than chemical combustion in a traditional rocket. NTP has many desirable capabilities including flexible mission launch dates and reduced transit times due to it’s capability for high specific impulse. One of the main challenges with NTP systems is the structural integrity of the fuel. The fuel form required in the reactor core must withstand temperatures above 2700 Kelvin. Ceramic-metallic matrix fuel, ceramic-ceramic matrix fuel, and solid solution carbide fuels are the three strongest candidates for the extreme environments. Solid solution carbides have the potential to exhibit promising behavior as a fuel form in an NTP system. Of the multiple refractory metal carbides of interest, here we focus on zirconium carbide (ZrC), niobium carbide (NbC), and tantalum carbide (TaC). ZrC, NbC, and TaC powders were consolidated in monocarbide (ZrC, NbC, and TaC) and bi-carbide (ZrC-NbC, ZrC-TaC, NbC-TaC) forms using spark plasma sintering (SPS). In addition to the pure endpoint carbides, the examined compositions of the various bi-carbides ranged from 25-75 mol%. The sintering temperatures, pressures, and hold times were varied to determine the ideal sintering conditions. Grain size analysis, Archimedes’ density, scanning electron microscopy (SEM), X-ray diffraction (XRD), and energy dispersion spectroscopy (EDS) were used to determine and characterize the grain size, ideal density, porosity, phase stability, and chemical composition of each sample. The data from each sample was then used to generate a Master Sintering Curve (MSC) unique to each monocarbide or bi-carbide.

nuclear thermal propulsion↗
NASA NTRS2023

Homogeneity Study of ZrC, NbC, and TaC Binary Carbide Fuels for the Application of Nuclear Thermal Propulsion

In alignment with NASA and DARPA goals, efforts towards Nuclear Thermal Propulsion (NTP) have increased in recent years. As progress is made, a main challenge still exists; the fuel in an NTP system must survive the extreme environment experienced during operating conditions. Three fuel forms have been studied including a ceramic-metallic matrix, a ceramic-ceramic matrix, and solid solution carbides. Solid solution carbides were studied previously as a fuel for NTP and was shown to be a promising technology; however, the programs were canceled before demonstration in an NTP engine test. Of the refractory metal carbides of interest, zirconium carbide (ZrC), niobium carbide (NbC) and tantalum carbide (TaC) were studied. ZrC, NbC, and TaC powders were consolidated in monocarbide (ZrC, NbC, TaC) and bi-carbide (ZrC-NbC, ZrC-TaC, NbC-TaC) forms using spark plasma sintering (SPS), with 50-50 mol% compositions of the bi-carbides of interest. The time and temperature conditions to achieve a completely homogeneous solid solution for the mixed carbides is of interest for NTP fuel applications, since this homogeneity is considered to be essential for this fuel form. Homogeneity is typically not achieved during sintering, as surface diffusion (predominant for sintering to near ideal density) tends to occur at lower temperatures than bulk volumetric diffusion (required for solute interdiffusion to achieve a homogeneous solid solution). High temperature isochronal annealing was performed to identify the conditions to achieve a chemically homogeneous sample. The solute interpenetration vs annealing temperature was quantified using energy dispersion spectroscopy (EDS) in a scanning electron microscope (SEM). X-ray diffraction (XRD) was used to accurately measure the phase stability and lattice parameter as a function of annealing conditions. The diffusion data from each sample was then used to determine self-diffusion coefficients for interdiffusion of the Nb, Zr, and Ta solutes and compared to existing literature values. The minimum annealing conditions to achieve complete chemical homogeneity for each carbide was determined.

nuclear thermal propulsion↗
NASA NTRS1962

PYROLYTIC GRAPHITE ROCKET THRUST CHAMBER INVESTIGATION PROJECT. MONTHLY LETTER

This is the Monthly Letter Progress Report for the eighth month of a twelve month technical study directed toward the following objectives: A. Determining analytically and experimentally the feasibility and attractiveness of using free-standing pyrolytic graphite for a radiation cooled liquid rocket thrust chamber and exit nozzle for space application. B. Establishing through analysis and experimental demonstration, a set of design principles to facilitate the design of a reliable pyrolytic graphite thrust chamber assembly for a specific range of operating conditions. C. Demonstrate by means of test firings of complete thrust chamber assemblies, the validity of the design data and principles developed during the program. Preparations have been completed for test firing the two 100 lb. thrust Boron Pyralloy chambers at Marquardt during the first week in March 1962. These two chambers will first be static pressure checked with water to a pressure of 135 psig. The chambers will then be test fired with N204/N2H4-MMH propellants . Test samples of eleven different pyrolytic carbide and nitride materials have been ordered for high temperature oxidation tests and should be delivered by the end of March for evaluation in the plasma torch facility. The specific materials ordered in free- standing condition in thicknesses from 0.030 inch to 0.065 inch were as follows: (a) Pyrolytic graphite (on hand) (b) Pyrolytic graphite-boron alloy (on hand) (c) Pyrolytic graphite-tungsten alloy (d)Boron nitride (e) Titanium nitride (f) Silicon carbide (g) Titanium carbide (h) Zirconium carbide (i) Hafnium carbide (j) Niobium carbide (k) Tantalum carbide (l) High density (1 1-90) graphite (m) High density graphite (heat treated) These materials were ordered from Raytheon, HTM, and American Meta 1 Products . A summary of the available data on the oxidation rates of pyrolytic materials under slow air and rocket motor conditions is shown in Figure 1. High Temperature Materials, Inc. has been contacted relative to providing Pyrographite thrust chambers with extra-thick walls . They believe they can deliver chambers with substantially thicker walls than were possible a short time ago. Such a configuration, with acceptable residual stresses, would provide a longer firing life since it is ultimately limited by the oxidation rate of the chamber wall material. Stress and structural studies have continued. The analytical stress studies point out some of the problems associated with the experimental determination of the elastic constants of this highly anisotropic material. Further analysis may shed light on improved techniques for making these experimental determinations. The large potential weight saving available with the use of pyrolytic graphite as a thrust chamber wall material in a radiation cooled rocket motor has been calculated and is shown graphically in Figure 2. Figure A presents the program progress schedule and the actual expenditures as of 7 February 1962.

ROCKET ENGINE↗