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Results for “Liquid propellant rocket engine”

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
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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 NTRS1994

Optical measurements in rocket engine liquid sprays

The performance of liquid propellant rocket engines is dependent upon many elements of the entire system. One of the most fundamental and most critical is the performance of the injector elements. Their characterization is an important part of the development of combustion devices. Optical measurements within these environments have proven to be invaluable tools in quantifying the physical environment of two phase flows. The effort reported herein involves the measurement of drop velocity, drop size, and most importantly mass flux using Phase-Doppler Particle Anemometry within a spray generated by a single swirl injector element operating in atmospheric pressure conditions. The mass flux has been determined and validated by mechanical patternation methods and by profile integration of the mass flux.

Feikema, Douglas A.↗
NASA NTRS1993

NASA Propulsion Engineering Research Center, volume 1

Over the past year, the Propulsion Engineering Research Center at The Pennsylvania State University continued its progress toward meeting the goals of NASA's University Space Engineering Research Centers (USERC) program. The USERC program was initiated in 1988 by the Office of Aeronautics and Space Technology to provide an invigorating force to drive technology advancements in the U.S. space industry. The Propulsion Center's role in this effort is to provide a fundamental basis from which the technology advances in propulsion can be derived. To fulfill this role, an integrated program was developed that focuses research efforts on key technical areas, provides students with a broad education in traditional propulsion-related science and engineering disciplines, and provides minority and other under-represented students with opportunities to take their first step toward professional careers in propulsion engineering. The program is made efficient by incorporating government propulsion laboratories and the U.S. propulsion industry into the program through extensive interactions and research involvement. The Center is comprised of faculty, professional staff, and graduate and undergraduate students working on a broad spectrum of research issues related to propulsion. The Center's research focus encompasses both current and advanced propulsion concepts for space transportation, with a research emphasis on liquid propellant rocket engines. The liquid rocket engine research includes programs in combustion and turbomachinery. Other space transportation modes that are being addressed include anti-matter, electric, nuclear, and solid propellant propulsion. Outside funding supports a significant fraction of Center research, with the major portion of the basic USERC grant being used for graduate student support and recruitment. The remainder of the USERC funds are used to support programs to increase minority student enrollment in engineering, to maintain Center infrastructure, and to develop research capability in key new areas. Significant research programs in propulsion systems for air and land transportation complement the space propulsion focus. The primary mission of the Center is student education. The student program emphasizes formal class work and research in classical engineering and science disciplines with applications to propulsion.

Source record↗
NASA NTRS2013

Review of Combustion Stability Characteristics of Swirl Coaxial Element Injectors

Liquid propellant rocket engine injectors using coaxial elements where the center liquid is swirled have become more common in the United States over the past several decades, although primarily for technology or advanced development programs. Currently, only one flight engine operates with this element type in the United States (the RL10 engine), while the element type is very common in Russian (and ex-Soviet) liquid propellant rocket engines. In the United States, the understanding of combustion stability characteristics of swirl coaxial element injectors is still very limited, despite the influx of experimental and theoretical information from Russia. The empirical and theoretical understanding is much less advanced than for the other prevalent liquid propellant rocket injector element types, the shear coaxial and like-on-like paired doublet. This paper compiles, compares and explores the combustion stability characteristics of swirl coaxial element injectors tested in the United States, dating back to J-2 and RL-10 development, and extending to very recent programs at the NASA MSFC using liquid oxygen and liquid methane and kerosene propellants. Included in this study are several other relatively recent design and test programs, including the Space Transportation Main Engine (STME), COBRA, J-2X, and the Common Extensible Cryogenic Engine (CECE). A presentation of the basic data characteristics is included, followed by an evaluation by several analysis techniques, including those included in Rocket Combustor Interactive Design and Analysis Computer Program (ROCCID), and methodologies described by Hewitt and Bazarov.

Hulka, J. R.↗
NASA NTRS1966

Propulsion System

Adaptation of Titan II liquid propellant rocket engine for Gemini program

LIQUID PROPELLANT ROCKET ENGINE↗
NASA NTRS2010

Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using techniques from Wenzel and Szuch. High-frequency combustion instability also occurred at the first tangential (1T) mode, at about 4500 Hz, with several multi-element injectors with liquid/liquid propellants. Analyses of the transverse mode instability were conducted by evaluating injector resonances and empirical methods developed by Hewitt.

Hulka, James R.↗
NASA NTRS2010

Performance and Stability Analyses of Rocket Combustion Devices Using Liquid Oxygen/Liquid Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented programs with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations on these programs. This paper summarizes these analyses. Test and analysis results of impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Several cases with gaseous methane are included for reference. Several different thrust chamber configurations have been modeled, including thrust chambers with multi-element like-on-like and swirl coax element injectors tested at NASA MSFC, and a unielement chamber with shear and swirl coax injectors tested at The Pennsylvania State University. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interaction Design and Analysis (ROCCID), and the Coaxial Injector Combustion Model (CICM). Significant effort was applied to show how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied or developed in the future. This paper describes the test hardware configurations, presents the results of all the analyses, and compares the results from the two analytical methods.

Hulka, James R.↗
NASA NTRS2010

Combustion and Performance Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, the NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations. This paper summarizes the analyses of combustion and performance as a follow-up to a paper published in the 2008 JANNAF/LPS meeting. Combustion stability analyses are presented in a separate paper. The current paper includes test and analysis results of coaxial element injectors using liquid oxygen and liquid methane or gaseous methane propellants. Several thrust chamber configurations have been modeled, including thrust chambers with multi-element swirl coax element injectors tested at the NASA MSFC, and a uni-element chamber with shear and swirl coax injectors tested at The Pennsylvania State University. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interaction Design and Analysis (ROCCID), and the Coaxial Injector Combustion Model (CICM). Significant effort was applied to show how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied or developed in the future. This paper describes the test hardware configurations, presents the results of all the analyses, and compares the results from the two analytical methods

Hulka, J. R.↗