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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Application of High Speed Digital Image Correlation in Rocket Engine Hot Fire Testing

Hot fire testing of rocket engine components and rocket engine systems is a critical aspect of the development process to understand performance, reliability and system interactions. Ground testing provides the opportunity for highly instrumented development testing to validate analytical model predictions and determine necessary design changes and process improvements. To properly obtain discrete measurements for model validation, instrumentation must survive in the highly dynamic and extreme temperature application of hot fire testing. Digital Image Correlation has been investigated and being evaluated as a technique to augment traditional instrumentation during component and engine testing providing further data for additional performance improvements and cost savings. The feasibility of digital image correlation techniques were demonstrated in subscale and full scale hotfire testing. This incorporated a pair of high speed cameras to measure three-dimensional, real-time displacements and strains installed and operated under the extreme environments present on the test stand. The development process, setup and calibrations, data collection, hotfire test data collection and post-test analysis and results are presented in this paper.

Gradl, Paul R.↗
NASA NTRS1976

Ames T-3 fire test facility - Aircraft crash fire simulation

There is a need to characterize the thermal response of materials exposed to aircraft fuel fires. Large scale open fire tests are costly and pollute the local environment. This paper describes the construction and operation of a subscale fire test that simulates the heat flux levels and thermochemistry of typical open pool fires. It has been termed the Ames T-3 Test and has been used extensively by NASA since 1969 to observe the behavior of materials exposed to JP-4 fuel fires.

Fish, R. H.↗
NASA NTRS1992

Hot fire test results of subscale tubular combustion chambers

Advanced, subscale, tubular combustion chambers were built and test fired with hydrogen-oxygen propellants to assess the increase in fatigue life that can be obtained with this type of construction. Two chambers were tested: one ran for 637 cycles without failing, compared to a predicted life of 200 cycles for a comparable smooth-wall milled-channel liner configuration. The other chamber failed at 256 cycles, compared to a predicted life of 118 cycles for a comparable smooth-wall milled-channel liner configuration. Posttest metallographic analysis determined that the strain-relieving design (structural compliance) of the tubular configuration was the cause of this increase in life.

Kazaroff, John M.↗
NASA NTRS2018

NASA Plum Brook Station In-Space Propulsion Facility Test Stand Characterization Hot Fire Test

A test facility modification to enable small scale altitude propulsion testing at the NASA Glenn Research Center's In-Space Propulsion (ISP) Facility was verified with a hot fire test campaign. As the facility's primary steam supply system undergoes refurbishment, the alternate facility configuration, known as the "vacuum accumulator" mode, would enable rocket engine testing up to 10,000 lbf thrust. The NASA Johnson Space Center developed the vehicle for the verification test campaign: the Integrated Cryogenic Propulsion Test Article (ICPTA). Constructed primarily from assets of the former Morpheus Project, the ICPTA provided an integrated liquid oxygen (LOX) / liquid methane (LCH4) propulsion system including a 2,800 lbf thrust main engine. The ISP Facility's vacuum accumulator configuration leveraged the large test volume of the facility and a diffuser insert to maintain altitude conditions. During hot fire, the ICPTA main engine "started" the diffuser insert constructed for the test campaign. As a result, the test chamber upstream of the diffuser insert remained at altitude conditions throughout the hot fire. Upon engine shut down, a backflow deflector mitigated blow back into the test chamber by restricting the mass flow and redirecting it away from the test article. The test campaign successfully characterized the performance of the vacuum accumulator configuration. In addition, it provided an opportunity to collect data for an integrated LOX / LCH4 propulsion system in an altitude and thermal vacuum environment.

Rocket Engine Testing↗
NASA NTRS2016

Cold Helium Pressurization for Liquid Oxygen/Liquid Methane Propulsion Systems: Fully-Integrated Hot-Fire Test Results

Hot-fire test demonstrations were successfully conducted using a cold helium pressurization system fully integrated into a liquid oxygen (LOX) / liquid methane (LCH4) propulsion system (Figure 1). Cold helium pressurant storage at near liquid nitrogen (LN2) temperatures (-275 F and colder) and used as a heated tank pressurant provides a substantial density advantage compared to ambient temperature storage. The increased storage density reduces helium pressurant tank size and mass, creating payload increases of ~35% for small lunar-lander sized applications. This degree of mass reduction also enables pressure-fed propulsion systems for human-rated Mars ascent vehicle designs. Hot-fire test results from the highly-instrumented test bed will be used to demonstrate system performance and validate integrated models of the helium and propulsion systems. A pressurization performance metric will also be developed as a means to compare different active pressurization schemes.

Morehead, R. L.↗
NASA NTRS1978

Development of fire test methods for airplane interior materials

Fire tests were conducted in a 737 airplane fuselage at NASA-JSC to characterize jet fuel fires in open steel pans (simulating post-crash fire sources and a ruptured airplane fuselage) and to characterize fires in some common combustibles (simulating in-flight fire sources). Design post-crash and in-flight fire source selections were based on these data. Large panels of airplane interior materials were exposed to closely-controlled large scale heating simulations of the two design fire sources in a Boeing fire test facility utilizing a surplused 707 fuselage section. Small samples of the same airplane materials were tested by several laboratory fire test methods. Large scale and laboratory scale data were examined for correlative factors. Published data for dangerous hazard levels in a fire environment were used as the basis for developing a method to select the most desirable material where trade-offs in heat, smoke and gaseous toxicant evolution must be considered.

Tustin, E. A.↗
NASA NTRS2015

Hot-Fire Testing of a 1N AF-M315E Thruster

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends. NASA completed a hot-fire test of a 1N AF-M315E monopropellant thruster at the Marshall Space Flight Center in the small altitude test stand located in building 4205. The thruster is a ground test article used for basic performance determination and catalyst studies. The purpose of the hot-fire testing was for performance determination of a 1N size thruster and form a baseline from which to study catalyst performance and life with follow-on testing to be conducted at a later date. The thruster performed as expected. The result of the hot-fire testing are presented in this paper and presentation.

Burnside, Christopher G.↗
NASA NTRS1980

Analysis of NASA JP-4 fire tests data and development of a simple fire model

The temperature, velocity and species concentration data obtained during the NASA fire tests (3m, 7.5m and 15m diameter JP-4 fires) were analyzed. Utilizing the data analysis, a sample theoretical model was formulated to predict the temperature and velocity profiles in JP-4 fires. The theoretical model, which does not take into account the detailed chemistry of combustion, is capable of predicting the extent of necking of the fire near its base.

Raj, P.↗
NASA NTRS2015

Hot-Fire Testing of 5N and 22N HPGP Thrusters

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends.NASA completed hot-fire testing of 5N and 22N HPGP thrusters at the Marshall Space Flight Center’s Component Development Area altitude test stand in April 2015. Both thrusters are ground test articles and not flight ready units, but are representative of potential flight hardware with a known path towards flight application. The purpose of the 5N testing was to perform facility check-outs and generate a small set of data for comparison to ECAPS and Orbital ATK data sets. The 5N thruster performed as expected with thrust and propellant flow-rate data generated that are similar to previous testing at Orbital ATK. Immediately following the 5N testing, and using the same facility, the 22N testing was conducted on the same test stand with the purpose of demonstrating the 22N performance. The results of 22N testing indicate it performed as expected.The results of the hot-fire testing are presented in this paper and presentation.

Burnside, Christopher G.↗
NASA NTRS1980

Fire tests for airplane interior materials

Large scale, simulated fire tests of aircraft interior materials were carried out in salvaged airliner fuselage. Two "design" fire sources were selected: Jet A fuel ignited in fuselage midsection and trash bag fire. Comparison with six established laboratory fire tests show that some laboratory tests can rank materials according to heat and smoke production, but existing tests do not characterize toxic gas emissions accurately. Report includes test parameters and test details.

Tustin, E. A.↗
NASA NTRS2016

Hot-Fire Test Results of Liquid Oxygen/RP-2 Multi-Element Oxidizer-Rich Preburners

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. To supply the oxidizer-rich combustion products to the main injector of the integrated test article, existing subscale preburner injectors from a previous NASA-funded oxidizer-rich staged combustion engine development program were utilized. For the integrated test article, existing and newly designed and fabricated inter-connecting hot gas duct hardware were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. However, before one of the preburners was used in the integrated test article, it was first hot-fire tested at length to prove it could provide the hot exhaust gas mean temperature, thermal uniformity and combustion stability necessary to perform in the integrated test article experiment. This paper presents results from hot-fire testing of several preburner injectors in a representative combustion chamber with a sonic throat. Hydraulic, combustion performance, exhaust gas thermal uniformity, and combustion stability data are presented. Results from combustion stability modeling of these test results are described in a companion paper at this JANNAF conference, while hot-fire test results of the preburner injector in the integrated test article are described in another companion paper.

Protz, C. S.↗
NASA NTRSDate not supplied

Hot-Fire Testing Results of Additively Manufactured NASA HR-1 towards Large Scale Implementation of Channel Wall Nozzles

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. One material that has been the focus of this development is NASA HR-1, an Iron-Nickel based superalloy optimized for use in high pressure hydrogen environments. Material properties and process characterization has been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also 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. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

Paul Gradl↗
NASA NTRS2016

Hot-Fire Test Results of an Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged-Combustion Integrated Test Article

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. For the thrust chamber assembly of the test article, several configurations of new main injectors, using relatively conventional gas-centered swirl coaxial injector elements, were designed and fabricated. The design and fabrication of these main injectors are described in a companion paper at this JANNAF meeting. New ablative combustion chambers were fabricated based on hardware previously used at NASA for testing at similar size and pressure. An existing oxygen/RP-1 oxidizer-rich subscale preburner injector from a previous NASA-funded program, along with existing and new inter-connecting hot gas duct hardware, were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Results from independent hot-fire tests of the preburner injector in a combustion chamber with a sonic throat are described in companion papers at this JANNAF conference. The resulting integrated test article - which includes the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was assembled at Test Stand 116 at the East Test Area of the NASA Marshall Space Flight Center. The test article was well instrumented with static and dynamic pressure, temperature, and acceleration sensors to allow the collected data to be used for combustion analysis model development. Hot-fire testing was conducted with main combustion chamber pressures ranging from 1400 to 2100 psia, and main combustion chamber mixture ratios ranging from 2.4 to 2.9. Different levels of fuel film cooling injected from the injector face were examined ranging from none to about 12% of the total fuel flow. This paper presents the hot-fire test results of the integrated test article. Combustion performance, stability, thermal, and compatibility characteristics of both the preburner and the thrust chamber are described. Another companion paper at this JANNAF meeting includes additional and more detailed test data regarding the combustion dynamics and stability characteristics.

Hulka, J. R.↗
NASA NTRSDate not supplied

Process Development and Hot-fire Testing of Additively Manufactured NASA HR-1 for Liquid Rocket Engine Applications

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. Optimized materials in the extreme high pressure and hydrogen environment for liquid rocket engines remains a key challenge. NASA has advanced an enabling material called NASA HR-1 (Hydrogen Resistant -1) as a solution using AM techniques. NASA HR-1 is a high-strength Fe-Ni superalloy designed to resist high pressure, hydrogen environment embrittlement, oxidation, and corrosion. NASA HR-1 meets materials requirements for liquid rocket engine components, including good hydrogen resistance, high conductivity, good low cycle fatigue performance, and high elongation and strength for components in high heat flux environments. Material properties and process characterization have been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also 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. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

DED↗
NASA NTRSDate not supplied

Process Development and Hot-fire Testing of Additively Manufactured NASA HR-1 for Liquid Rocket Engine Applications

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. Optimized materials in the extreme high pressure and hydrogen environment for liquid rocket engines remains a key challenge. NASA has advanced an enabling material called NASA HR-1 (Hydrogen Resistant -1) as a solution using AM techniques. NASA HR-1 is a high-strength Fe-Ni superalloy designed to resist high pressure, hydrogen environment embrittlement, oxidation, and corrosion. NASA HR-1 meets materials requirements for liquid rocket engine components, including good hydrogen resistance, high conductivity, good low cycle fatigue performance, and high elongation and strength for components in high heat flux environments. Material properties and process characterization have been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also 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. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

Additive Manufacturing↗
NASA NTRS1979

Airport flammability, full scale fire tests

The primary objectives of the fire tests are: (1) conduct full scale test with 737 fuselages by the end of 1980; (2) demonstrate evacuation time increase to five minutes minimum; (3) show that exterior fire will not penetrate an intact cabin for five minutes; (4) show that closed cabin will not have excess smoke or temperatures above 400 F; and (5) demonstrate that fire in cabin opening will not propagate throughout cabin. The test program has three phases. A test schedule and status of required materials are presented.

Bricker, R. W.↗
NASA NTRS1977

Systems for animal exposure in full-scale fire tests

Two systems for exposing animals in full-scale fire tests are described. Both systems involve the simultaneous exposure of two animal species, mice and rats, in modular units; determination of mortality, morbidity, and behavioral response; and analysis of the blood for carboxyhemoglobin. The systems described represent two of many possible options for obtaining bioassay data from full-scale fire tests. In situations where the temperatures to which the test animals are exposed can not be controlled, analytical techniques may be more appropriate than bioassay techniques.

Hilado, C. J.↗