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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 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
DOE OSTI2022

Initial Atmospheric Transport (IAT)

SAND2022-12904 O Initial Atmospheric Transport (IAT) predicts the initial puff/plume rise resulting from either a solid rocket propellant or liquid rocket fuel fire. The code generates initial conditions for subsequent atmospheric transport calculations. Sandia National Laboratories is a multimission laboratory managed and operated by National Technology & Engineering Solutions of Sandia, LLC, a wholly owned subsidiary of Honeywell International Inc., for the U.S. Department of Energy’s National Nuclear Security Administration under contract DE-NA0003525.

Villa, Daniel↗
NASA NTRS2015

Unsteady Three-Dimensional Simulation of a Shear Coaxial GO2/GH2 Rocket Injector with RANS and Hybrid-RAN-LES/DES Using Flamelet Models

Historically, the analysis and design of liquid rocket engines (LREs) has relied on full‐scale testing and one‐dimensional empirical tools. The testing is extremely expensive and the one‐dimensional tools are not designed to capture the highly complex, and multi‐dimensional features that are inherent to LREs. Recent advances in computational fluid dynamics (CFD) tools have made it possible to predict liquid rocket engine performance, stability, to assess the effect of complex flow features, and to evaluate injector‐driven thermal environments, to mitigate the cost of testing. Extensive efforts to verify and validate these CFD tools have been conducted, to provide confidence for using them during the design cycle. Previous validation efforts have documented comparisons of predicted heat flux thermal environments with test data for a single element gaseous oxygen (GO2) and gaseous hydrogen (GH2) injector. The most notable validation effort was a comprehensive validation effort conducted by Tucker et al. [1], in which a number of different groups modeled a GO2/GH2 single element configuration by Pal et al [2]. The tools used for this validation comparison employed a range of algorithms, from both steady and unsteady Reynolds Averaged Navier‐Stokes (U/RANS) calculations, large‐eddy simulations (LES), detached eddy simulations (DES), and various combinations. A more recent effort by Thakur et al. [3] focused on using a state‐of‐the‐art CFD simulation tool, Loci/STREAM, on a two‐dimensional grid. Loci/STREAM was chosen because it has a unique, very efficient flamelet parameterization of combustion reactions that are too computationally expensive to simulate with conventional finite‐rate chemistry calculations. The current effort focuses on further advancement of validation efforts, again using the Loci/STREAM tool with the flamelet parameterization, but this time with a three‐dimensional grid. Comparisons to the Pal et al. heat flux data will be made for both RANS and Hybrid RANSLES/ Detached Eddy simulations (DES). Computation costs will be reported, along with comparison of accuracy and cost to much less expensive two‐dimensional RANS simulations of the same geometry.

Westra, Doug G.↗
NASA NTRS2023

Effects of Strain Rate and Surface Finish on Tensile Properties in High Pressure Hydrogen — A Case Study for LP-DED NASA HR-1

NASA has advanced NASA HR-1 as a solution for liquid rocket engine components that operate in hydrogen rich environments using various AM techniques. The NASA HR-1 for AM applications was specifically formulated for high ductility in high pressure gaseous hydrogen environment in addition to high strength and low cycle fatigue (LCF) resistance. NASA HR-1 meets materials requirements for liquid rocket engine components, including good hydrogen resistance, high conductivity, good LCF performance, and high elongation and strength for components in medium heat flux environments. NASA has completed fabrication of several subscale and full-scale channel wall nozzles in LP-DED NASA HR-1 and completed hot-fire testing. This includes refinement of the process to produce thin-walls and various channel geometries to meet the requirements for channel wall nozzle applications. The main objective of this investigation is to better understand the effects of tensile test strain rate and specimen surface finish on the amount of tensile ductility degradation in high pressure gaseous hydrogen environments. Performing smooth tensile tests in high pressure hydrogen and nitrogen environments can serve as a rapid and economical screening method to determine the relative susceptibility of materials to HEE. This paper presents the tensile test results for LP-DED NASA HR-1 in hydrogen and details the material evaluation on the effects of strain rate and surface finish condition on surface cracking behavior and fracture characteristics. The resistance to HEE for LP-DED NASA HR-1 and the wrought alloy is compared under the same testing conditions. Loss of tensile ductility is quantified in terms of changes in elongation at failure in hydrogen to compare HEE susceptibility. The most suitable strain rate and specimen surface finish condition for future HEE screening testing at MSFC are recommended.

Additive manufacturing↗
NASA NTRS1992

Computational analysis of liquid hypergolic propellant rocket engines

The combustion process in liquid rocket engines depends on a number of complex phenomena such as atomization, vaporization, spray dynamics, mixing, and reaction mechanisms. A computational tool to study their mutual interactions is developed to help analyze these processes with a view of improving existing designs and optimizing future designs of the thrust chamber. The focus of the article is on the analysis of the Variable Thrust Engine for the Orbit Maneuvering Vehicle. This engine uses a hypergolic liquid bipropellant combination of monomethyl hydrazine as fuel and nitrogen tetroxide as oxidizer.

Krishnan, A.↗
NASA NTRS2023

Reactive Additive Manufacturing for Fourth Industrial Revolution Exploration Systems (Ramfire) Aluminum 6061-Ram2 Nozzle Testing

The NASA Marshall Space Flight Center (MSFC) has applied various forms of metallic additive manufacturing (AM) in liquid rocket engine component design, development, and testing since 2010. These AM techniques reduce hardware cost, shorten fabrication schedules, increase reliability by reducing the number of joints, and improve hardware performance by allowing unconventional design. The RAMFIRE project, funded under Space Technology Mission Directorate (STMD) Game Changing Development (GCD) Program, has furthered the use of novel AM liquid rocket nozzles in collaboration with Elementum 3D. The project advanced novel large-scale AM aluminum material technology for significant weight savings in rocket engines and launch vehicles. Previously, aluminum alloys were difficult to weld and print using additive manufacturing. Elementum 3D’s patented Aluminum 6061-RAM2 alloy allows for aluminum alloys to be printed using various AM techniques and at various scales. The alloy can be leveraged for welding wire, showing drastic improvements in aluminum weldability. The RAMFIRE project focuses on five key areas: 1) Laser Powder Directed Energy Deposition (LP-DED) AL6061-RAM2 feedstock specification and verification, 2) LP-DED process development and validation, 3) LP-DED printed AL6061-RAM2 microstructural and mechanical property characterization, 4) Hot-fire test a 7k-lbf thrust class regeneratively cooled nozzle, 5) Print large scale regeneratively cooled nozzle. Hot-fire testing demonstrates potential of advanced space technologies to NASA and potential users by providing the relevant environments to advance TRL levels to the 5/6 range.

Al6061-RAM2↗
NASA NTRS1959

Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48

The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

Fortini, Anthony↗
NASA NTRS1963

MODELING TECHNIQUES FOR LIQUID PROPELLANT ROCKET COMBUSTION PROCESSES

A qualitative physical processes description of stable and unstable combustion in rockets is presented to establish the basis for choosing model hardware design criteria. It is concluded that it is not possible to scale rocket combustion processes in the usual sense of the term. The studies conducted at Rocketdyne have shown that model chambers which satisfactorily model the steady state behavior of large engines must be designed to maintain (1) the propellant injection density, and (2) the chamber to throat Contraction ratio (hence chamber pressure). For studies of destructive acoustic modes of combustion instability a third parameter, the frequency, must also be maintained.

COMBUSTION↗
NASA NTRSDate not supplied

A Review Towards the Design Optimization of High Performance Additively Manufactured Rotating Detonation Rocket Engine Injectors

Rotating Detonation Rocket Engines (RDRE) have been marketed primarily for their higher specific impulse potential over constant pressure (CP) liquid rocket engines. However, several other performance advantages exist with RDREs over CP engines such as heat transfer advantages for gas expander cycle, increased completeness of combustion at low chamber L*, compact engine design, reduced coolant channel pressure drop potential, and improved injector C* performance. NASA has paved the way for liquid engine system performance enhancement since the Apollo program and continues to do so with metal additive manufacturing (AM), super-alloy materials, and advanced propulsion concepts. A team of propulsion development engineers at NASA are in the process of developing high-performance 7K lbf class RDRE hardware for their potential use in lander, upper stage, and even launch vehicle applications. Clear advantages have been demonstrated with AM including program cost and schedule reductions of up to 50%. It is well known that injector performance is integrally linked to the global performance of a combustion device. This is especially the case for RDREs since detonation stability is heavily dependent on the mixedness of propellants. A major program goal is to rapidly produce ultra-high-performance AM injectors. This paper reviews the available literature on liquid rocket injector design optimization as well as the experimental work conducted to date on injectors tested in RDREs. Major lessons learned are document and suggestions given towards the design of high-performance liquid RDRE injectors. In addition, the integration of metal AM into the design of liquid RDRE injector schemes is discussed. Finally, several candidate AM RDRE injector elements were produced to obtain their diodicity and cold flow characteristics.

rotating detonation engine↗
NASA NTRS2015

Method for Determining Optimum Injector Inlet Geometry

A method for determining the optimum inlet geometry of a liquid rocket engine swirl injector includes obtaining a throttleable level phase value, volume flow rate, chamber pressure, liquid propellant density, inlet injector pressure, desired target spray angle and desired target optimum delta pressure value between an inlet and a chamber for a plurality of engine stages. The method calculates the tangential inlet area for each throttleable stage. The method also uses correlation between the tangential inlet areas and delta pressure values to calculate the spring displacement and variable inlet geometry of a liquid rocket engine swirl injector.

Trinh, Huu P.↗
NASA NTRS2016

Additive Manufacturing a Liquid Hydrogen Rocket Engine

Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.

Jones, Carl P.↗
NASA NTRS1998

Injector Design Tool Improvements: User's manual for FDNS V.4.5

The major emphasis of the current effort is in the development and validation of an efficient parallel machine computational model, based on the FDNS code, to analyze the fluid dynamics of a wide variety of liquid jet configurations for general liquid rocket engine injection system applications. This model includes physical models for droplet atomization, breakup/coalescence, evaporation, turbulence mixing and gas-phase combustion. Benchmark validation cases for liquid rocket engine chamber combustion conditions will be performed for model validation purpose. Test cases may include shear coaxial, swirl coaxial and impinging injection systems with combinations LOXIH2 or LOXISP-1 propellant injector elements used in rocket engine designs. As a final goal of this project, a well tested parallel CFD performance methodology together with a user's operation description in a final technical report will be reported at the end of the proposed research effort.

Chen, Yen-Sen↗
NASA NTRS1993

Heat transfer in rocket combustion chambers

Complexities of liquid rocket engine heat transfer which involve the injector faceplate and film cooled walls are being investigated by computational analysis. A conjugate heat transfer analysis was used to describe localized heating phenomena associated with particular injector configurations and film coolant flows. These components were analyzed, and the analyses verified when appropriate test data were available. The component analyses are being synthesized into an overall flowfield/heat transfer model. A Navier-Stokes flow solver, the FDNS code, was used to make the analyses. Particular attention was given to the representation of the thermodynamic properties of the fluid streams. Unit flow models of specific coaxial injector elements have been developed and are being used to describe the flame structure near the injector faceplate.

Anderson, P.↗
NASA NTRS1975

Thrust vector control by liquid injection for solid propellant rockets

In liquid injection thrust vector control, a rocket jet is deflected for steering purposed by injecting a liquid into the nozzle exit cone. The liquid is preferably both dense and reactive so that it adds mass and energy and generates shocks in the supersonic exhaust. This behavior increases thrust in the affected part of the jet producing not only a side force for steering but an addition to axial thrust. This paper presents a summary of current liquid injection thrust vector control technology, including procedures for design, development, analysis, testing and evaluation, together with supporting data and references.

Zeamer, R. J.↗
NASA NTRS1998

Low Cost Propulsion Technology at the Marshall Space Flight Center: Fastrac Engine and the Propulsion Test Article

The need for low cost access to space has initiated the development of low cost liquid rocket engine and propulsion system hardware at the Marshall Space Flight Center (MSFC). The engine, the 60,000 lbf, RP-1 and LOX Fastrac Engine has been designed as a robust, low cost liquid rocket engine with applications for X-34 as well as future low cost booster systems. The engine is a turbopump fed, gas generator cycle, rocket motor with an ablative nozzle. The Propulsion Test Article (PTA) is a test bed for low cost propulsion system hardware including a composite RP-1 tank, flight feedlines and pressurization system, stacked in a booster configuration. A general description of the PTA and the Fastrac engine is given, with emphasis on the technical specification of the hardware including flow rates, pressures and other operating conditions. The process which has been used for the design and integration of this hardware is described.

Fisher, Mark F.↗
NASA NTRS2001

Time-Dependent Simulations of Turbopump Flows

The objective of the current effort is to provide a computational framework for design and analysis of the entire fuel supply system of a liquid rocket engine, including high-fidelity unsteady turbopump flow analysis. This capability is needed to support the design of pump sub-systems for advanced space transportation vehicles that are likely to involve liquid propulsion systems. To date, computational tools for design/analysis of turbopump flows are based on relatively lower fidelity methods. An unsteady, three-dimensional viscous flow analysis tool involving stationary and rotational components for the entire turbopump assembly has not been available for real-world engineering applications. The present effort will provide developers with information such as transient flow phenomena at start up, impact of non-uniform inflows, system vibration and impact on the structure. In the proposed paper, the progress toward the capability of complete simulation of the turbo-pump for a liquid rocket engine is reported. The Space Shuttle Main Engine (SSME) turbo-pump is used as a test case for evaluation of the hybrid MPI/Open-MP and MLP versions of the INS3D code. The relative motion of the grid systems for the rotor-stator interaction was obtained using overset grid techniques. Time-accuracy of the scheme has been evaluated with simple test cases. Unsteady computations for the SSME turbo-pump, which contains 114 zones with 34.5 million grid points, are carried out on Origin 2000 systems at NASA Ames Research Center. Results from these time-accurate simulations with moving boundary capability will be presented along with the performance of parallel versions of the code.

Kris, Cetin C.↗