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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 NTRS1987

Conventionally cast and forged copper alloy for high-heat-flux thrust chambers

The combustion chamber liner of the space shuttle main engine is made of NARloy-Z, a copper-silver-zirconium alloy. This alloy was produced by vacuum melting and vacuum centrifugal casting; a production method that is currently now available. Using conventional melting, casting, and forging methods, NASA has produced an alloy of the same composition called NASA-Z. This report compares the composition, microstructure, tensile properties, low-cycle fatigue life, and hot-firing life of these two materials. The results show that the materials have similar characteristics.

Kazaroff, John M.↗
NASA NTRS1987

Space station auxiliary thrust chamber technology

A program to design, fabricate, and test a 50 lb sub f (222 N) thruster was undertaken to demonstrate the applicability of the reverse flow concept as an item of auxillary propulsion for the Space Station. The thruster was to operate at a mixture ratio (O/F) of 4, be capable of operating for 2 million lb sub f-seconds (8.896 million N-seconds) impulse with a chamber pressure of 75 psia (52N/sq cm) and a nozzle area ratio of 40. A successful demonstration of an (0/F) of 4 thruster, was followed by the design objective of operating at (O/F) of 8. The demonstration of this thruster resulted in the order of and additional (O/F) of 8 thruster chamber under the present NAS 3-24883 contract. The effort to fabricate and test the second (0/F) of 8 thruster is documented.

Senneff, J. M.↗
NASA NTRS1989

Liquid Oxygen Cooling of Hydrocarbon Fueled Rocket Thrust Chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the collant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1989

Radiative heat transfer in rocket thrust chambers and nozzles

Numerical models based on the discrete exchange factor (DEF) and the zonal methods for radiative analysis of rocket engines containing a radiatively participating medium have been developed. These models implement a new technique for calculating the direct exchange factors to account for possible blockage by the nozzle throat. Given the gas and surface temperature distributions, engine geometry, and radiative properties, the models compute the wall radiative heat fluxes at different axial positions. The results of sample calculations for a typical rocket engine (engine 700 at NASA), which uses RP-1 (a kerosene-type propellant), are presented for a wide range of surface and gas properties. It is found that the heat transfer by radiation can reach up to 50 percent of that due to convection. The maximum radiative heat flux is at the inner side of the engine, where the gas temperature is the highest. While the results of both models are in excellent agreement, the computation time of the DEF method is found to be much smaller.

Hammad, K. J.↗
NASA NTRS1989

Liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the coolant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1990

A three-dimensional turbulent heat transfer analysis for advanced tubular rocket thrust chambers

Heat transfer was analyzed in the throat region of a plug and spool rocket engine for both smooth and corrugated walls. A three-dimensional, Navier-Stokes code was used for the analysis. The turbulence model in the code was modified to handle turbulence suppression in the crevice region of the corrugated wall. The overall heat transfer at the throat for the corrugated wall was 34 percent higher than it was for the smooth wall for comparable rocket flow conditions.

Kacynski, Kenneth J.↗
NASA NTRS1990

A Three-Dimensional Turbulent Heat Transfer Analysis for Advanced Tubular Rocket Thrust Chambers

Heat transfer was analyzed in the throat region of a plug and spool rocket engine for both smooth and corrugated walls. A three-dimensional, Navier-Strokes code was used for the analysis. The turbulence model in the code was modified to handle turbulence suppression in the crevice region of the corrugated wall. Circumferential variations in the wall heat transfer was predicted for the corrugated wall. The overall heat transfer at the throat of the corrugated wall was 34 percent higher than it was for the smooth wall for comparable rocket flow conditions.

Kacynski, Kenneth J.↗
NASA NTRS1990

Space Station auxiliary thrust chamber technology

The objective was to establish a technical data base to support future development of GO2/GH2 flight thrusters for a Space Station Auxiliary Propulsion System. Specific issues of concern were thruster performance and cycle life. To address these issues, NASA funded Aerojet to design, fabricate, and altitude test two 25-lbf GO2/GH2 thrusters. The first thruster was designed to operate at a nominal mixture ratio (O/F) of 4.0 and expansion area ratio (epsilon) of 100:1. It was tested over a range of O/F from 2.0 to 8.0, achieving a range of specific impulse (Isp) from 440 to 310 lbf-sec/Ibm. The second thruster was optimized for a nominal O/F of 8.0 at a lower nozzle expansion area ratio, epsilon, of 30:1. This second thruster was tested over an O/F range of 3.0 to 9.5, achieving an Isp range of 416 to 3323 lbf-sec/Ibm, respectively. At O/F = 8.0, the Isp was 360 lbf-sec/Ibm, as predicted.

Robinson, Philip J.↗
NASA NTRS1991

Investigation of throat film coolant for advanced LOX/RP-1 thrust chambers

The use of throat film coolant as a method to reduce peak combustion heat flux was investigated. Hot-fire tests were performed using a subscale LOX/RP-1 combustor at a nominal chamber pressure of 2000 psia. A special injection manifold was fabricated to introduce the film coolant into the chamber. Four tests were conducted with RP-1 film coolant injected 4.4 inches upstream of the throat. In order to determine the effect of chamber length on the film coolant, four additional tests were performed with the injection point moved 16.4 inches upstream of the throat. Results from the test program are discussed. Methods for incorporating throat film cooling into a flightweight combustion chamber are also discussed.

Volkmann, Jon C.↗
NASA NTRS1992

Pressure fed thrust chamber technology program

This is the final report for the Pressure Fed Technology Program. It details the design, fabrication, and testing of subscale hardware which successfully characterized Liquid Oxygen Rocket Propulsion (LOX/RP) combustion for low cost pressure fed design. The innovative modular injector design is described in detail as well as hot-fire test results which showed excellent performance. The program summary identifies critical LOX/RP design issues that have been resolved in this testing, and details the low risk development requirements for low cost engines for future Expandable Launch Vehicles (ELV).

Dunn, Glen M.↗
NASA NTRS1992

A numerical model for atomization-spray coupling in liquid rocket thrust chambers

The physical process of atomization is an important consideration in the stable operation of liquid rocket engines. Many spray combustion computational fluid dynamics (CFD) codes do not include an atomization sub-model but assume arbitrary drop size distributions, drop initial locations, and velocities. A method of coupling an atomization model with the spray model in a REFLEQS CFD code is presented. This method is based on a jet-embedding technique in which the equations governing the liquid jet core are solved separately using the surrounding gas phase conditions. The droplet initial conditions are calculated using a stability analysis appropriate for the atomization regime of liquid jet break-up. This novel coupling model is used to analyze the SSME fuel preburner single injector flow. Results of the diffusion flame characteristics in a single injection element will be presented. The effect of relative velocity, mixture ratio and droplet initial conditions will be shown. The predictions of present atomization model is compared with that of the widely used CICM correlation. The results are also compared with the predictions of volume-of-fluid method.

M. G. Giridharan↗
NASA NTRS1992

Pressure fed thrust chamber technology program

This is the final report for the Pressure Fed Technology Program. It details the design, fabrication and testing of subscale hardware which successfully characterized LOX/RP combustion for a low cost pressure fed design. The innovative modular injector design is described in detail as well as hot-fire test results which showed excellent performance. The program summary identifies critical LOX/RP design issues that have been resolved by this testing, and details the low risk development requirements for a low cost engine for future Expendable Launch Vehicles (ELVi).

Dunn, Glenn M.↗
NASA NTRS2005

Fabrication of GRCop-84 Rocket Thrust Chambers

GRCop-84, a copper alloy, Cu-8 at% Cr-4 at% Nb developed at NASA Glenn Research Center for regeneratively cooled rocket engine liners has excellent combinations of elevated temperature strength, creep resistance, thermal conductivity and low cycle fatigue. GRCop-84 is produced from prealloyed atomized powder and has been fabricated into plate, sheet and tube forms as well as near net shapes. Fabrication processes to produce demonstration rocket combustion chambers will be presented and includes powder production, extruding, rolling, forming, friction stir welding, and metal spinning. GRCop-84 has excellent workability and can be readily fabricated into complex components using conventional powder and wrought metallurgy processes. Rolling was examined in detail for process sensitivity at various levels of total reduction, rolling speed and rolling temperature representing extremes of commercial processing conditions. Results indicate that process conditions can range over reasonable levels without any negative impact to properties.

Loewenthal, William S.↗
NASA NTRS2006

Fabrication of GRCop-84 Rocket Thrust Chambers

GRCop-84, a copper alloy, Cu-8 at% Cr-4 at% Nb developed at NASA Glenn Research Center for regenerative1y cooled rocket engine liners has excellent combinations of elevated temperature strength, creep resistance, thermal conductivity and low cycle fatigue. GRCop-84 is produced from pre-alloyed atomized powder and has been fabricated into plate, sheet and tube forms as well as near net shapes. Fabrication processes to produce demonstration rocket combustion chambers will be presented and includes powder production, extruding, rolling, forming, friction stir welding, and metal spinning. GRCop-84 has excellent workability and can be readily fabricated into complex components using conventional powder and wrought metallurgy processes. Rolling was examined in detail for process sensitivity at various levels of total reduction, rolling speed and rolling temperature representing extremes of commercial processing conditions. Results indicate that process conditions can range over reasonable levels without any negative impact to properties.

Loewenthal, William↗
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 NTRS2001

Lightweight Chambers for Thrust Assemblies

The Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) has successfully applied new materials and fabrication techniques to create actively cooled thrust chambers that operate 200-400 degrees hotter and weigh 50% lighter than conventional designs. In some vehicles, thrust assemblies account for as much as 20% of the engine weight. So, reducing the weight of these components and increasing their operating range will benefit many engines and vehicle designs, including Reusable Launch Vehicle (RLV) concepts. Obviously, copper and steel alloys have been used successfully for many years in the chamber components of thrust assemblies. Yet, by replacing the steel alloys with Polymer Matrix Composite (PMC) and/or Metal Matrix Composite (MMC) materials, design weights can be drastically reduced. In addition, replacing the traditional copper alloys with a Ceramic Matrix Composite (CMC) or an advanced copper alloy (Cu-8Cr-4Nb, also known as GRCop-84) significantly increases allowable operating temperatures. Several small MMC and PMC demonstration chambers have recently been fabricated with promising results. Each of these designs included GRCop-84 for the cooled chamber liner. These units successfully verified that designs over 50% lighter are feasible. New fabrication processes, including advanced casting technology and a low cost vacuum plasma spray (VPS) process, were also demonstrated with these units. Hot-fire testing at MSFC is currently being conducted on the chambers to verify increased operating temperatures available with the GRCop-84 liner. Unique CMC chamber liners were also successfully fabricated and prepared for hot-fire testing. Yet, early results indicate these CMC liners need significantly more development in order to use them in required chamber designs. Based on the successful efforts with the MMC and PMC concepts, two full size "lightweight" chambers are currently being designed and fabricated for hot-fire testing at MSFC in 2001. These "full size" chambers will be similar in size to those used on the X33 engine (RS2200). One will be fabricated with a MMC structural jacket, while the other uses a PMC jacket. Each will be designed for thrust levels of 15,000 pounds in an oxygen/hydrogen environment with liquid hydrogen coolant. Both chambers will use GRCop-84 for its channel wall liner. Each unit is expected to be at least 60% lighter than a conventional design with traditional materials. Hot-fire testing on the full size units in late 2001 will directly compare performance results between a conventional chamber design and these "lightweight" alternatives. The technology developed and demonstrated by this effort will not only benefit next generation RLV programs, but it can be applied to other existing and future engine programs, as well. Efforts were sponsored by the Advanced Space Transportation Program for RLV Focused Technologies. The task team was led by MSFC with additional members from NASA-Glenn Research Center and the Rocketdyne Division of The Boeing Company. Specific materials development and fabrication processes were provided by Aerojet, Lockheed Martin Astronautics, Composite Optics, Inc., Hyper-Therm, Ceramic Composites, Inc., MSE Technology Applications, and Plasma Processes, Inc.

Elam, Sandra K.↗