Search NASA⌕ Search

SEARCH · Search NASA

Results for “Propellant testing”

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 1,459 records · Page 81
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Oxygen Compatibility Testing of Composite Materials

Composite materials offer significant weight-saving potential for aerospace applications in propellant and oxidizer tanks. This application for oxygen tanks presents the challenge of being oxygen compatible in addition to complying with the other required material characteristics. This effort reports on the testing procedures and data obtained in examining and selecting potential composite materials for oxygen tank usage. Impact testing of composites has shown that most of these materials initiate a combustion event when impacted at 72 ft-lbf in the presence of liquid oxygen, though testing has also shown substantial variability in reaction sensitivities to impact. Data for screening of 14 potential composites using the Bruceton method is given herein and shows that the 50-percent reaction frequencies range from 17 to 67 ft-lbf. The pressure and temperature rises for several composite materials were recorded to compare the energy releases as functions of the combustion reactions with their respective reaction probabilities. The test data presented are primarily for a test pressure of 300 psia in liquid oxygen. The impact screening process is compared with oxygen index and autogenous ignition test data for both the composite and the basic resin. The usefulness of these supplemental tests in helping select the most oxygen compatible materials is explored. The propensity for mechanical impact ignition of the composite compared with the resin alone is also examined. Since an ignition-free composite material at the peak impact energy of 72 ft-lbf has not been identified, composite reactivity must be characterized over the impact energy level and operating pressure ranges to provide data for hazard analyses in selecting the best potential material for liquid tank usage.

Engel, Carl D.↗
NASA NTRS1986

Advanced turboprop noise prediction: Development of a code at NASA Langley based on recent theoretical results

The development of a high speed propeller noise prediction code at Langley Research Center is described. The code utilizes two recent acoustic formulations in the time domain for subsonic and supersonic sources. The structure and capabilities of the code are discussed. Grid size study for accuracy and speed of execution on a computer is also presented. The code is tested against an earlier Langley code. Considerable increase in accuracy and speed of execution are observed. Some examples of noise prediction of a high speed propeller for which acoustic test data are available are given. A brisk derivation of formulations used is given in an appendix.

Farassat, F.↗
NASA NTRS1987

Advanced turboprop noise prediction based on recent theoretical results

The development of a high speed propeller noise prediction code at Langley Research Center is described. The code utilizes two recent acoustic formulations in the time domain for subsonic and supersonic sources. The structure and capabilities of the code are discussed. Grid size study for accuracy and speed of execution on a computer is also presented. The code is tested against an earlier Langley code. Considerable increase in accuracy and speed of execution are observed. Some examples of noise prediction of a high speed propeller for which acoustic test data are available are given. A brisk derivation of formulations used is given in an appendix.

Farassat, F.↗
NASA NTRS1973

Liquid hydrogen turbopump rapid start program

This program was to analyze, test, and evaluate methods of achieving rapid-start of a liquid hydrogen feed system (inlet duct and turbopump) using a minimum of thermal preconditioning time and propellant. The program was divided into four tasks. Task 1 includes analytical studies of the testing conducted in the other three tasks. Task 2 describes the results from laboratory testing of coating samples and the successful adherence of a KX-635 coating to the internal surfaces of the feed system tested in Task 4. Task 3 presents results of testing an uncoated feed system. Tank pressure was varied to determine the effect of flowrate on preconditioning. The discharge volume and the discharge pressure which initiates opening of the discharge valve were varied to determine the effect on deadhead (no through-flow) start transients. Task 4 describes results of testing a similar, internally coated feed system and illustrates the savings in preconditioning time and propellant resulting from the coatings.

Wong, G. S.↗
NASA NTRS2004

Fuel Flexible Gas Turbine Combustor Flametube Facility Upgraded

In fiscal year 2003, test cell 23 of the Research Combustion Laboratory (RCL 23) at the NASA Glenn Research Center was upgraded with the addition of gaseous hydrogen as a working propellant and the addition of a 450-psig air-supply system. Test flexibility was further enhanced by upgrades to the facility control systems. RCL 23 can now test with gaseous hydrogen flow rates up to 0.05 lbm/sec and jet fuel flow rates up to 0.62 lbm/sec. Research airflow rates up to 3 lbm/sec are possible with the 450-psig supply system over a range of inlet temperatures. Nonvitiated, heated air is supplied from a shell and tube heat exchanger. The maximum nonvitiated facility air temperature is 1100 F at 1.5 lbm/sec. Research-section exhaust temperatures are limited to 3200 F because of material and cooling capacity limits. A variety of support systems are available depending on the research hardware configuration. Test section ignition can be provided via either a hydrogen air torch system or an electronic spark system. Emissions measurements are obtained with either pneumatically or electromechanically actuated gas sample probes, and the electromechanical system allows for radial measurements at a user-specified axial location for measurement of emissions profiles. Gas analysis data can be obtained for a variety of species, including carbon monoxide (CO), carbon dioxide (CO2), nitrogen oxides (NO and NOx), oxygen (O2), unburnt hydrocarbons, and unburnt hydrogen. Facility control is accomplished with a programmable logic control system. Facility operations have been upgraded to a system based on graphical user interface control screens. A data system is available for real-time acquisition and monitoring of both measurements in engineering units and performance calculations. The upgrades have made RCL 23 a highly flexible facility for research into low emissions gas turbine combustor concepts, and the flame tube configuration inherently allows for a variety of fuel nozzle configurations to be tested in a cost-effective manner. RCL 23 is poised to be a leading facility for developing modern low-emission fuel nozzles for use with jet fuel and alternative fuels.

Little, James E.↗
NASA NTRS1954

Summary of Free-Flight Zero-Lift Drag Results from Tests of 1/5-Scale Models of the Convair YF-102 and F-102A Airplanes and Several Related Small Equivalent Bodies at Mach Numbers from 0.70 to 1.46

One-fifth-scale rocket-propelled models of the Convair YF-102 and F-102A airplanes were tested to determine free-flight zero-lift drag coefficients through the transonic speed range at Reynolds numbers near those to be encountered by the full-scale airplane. Trim and duct characteristics were obtained along with measurements of total-, internal-, and base-drag coefficients. Additional zero-lift drag tests involved a series of small equivalent-body-of-revolution models which were launched to low supersonic speeds by means of a helium gun. The several small models tested corresponded to the following full-scale airplanes: basic, YF-102, 2-foot (full-scale) fuselage extension, F-102A, F-102A (relocated inlets), F-102A (faired nose), and F-102A (parabolic nose) . Equivalent-body models corresponding to the normal area distribution (derived for Mach number 1.0) of each of these airplane shapes were flown and, in addition, equivalent-body models designed to represent the YF-102 and F-102A airplanes at Mach number 1.2 were tested. External-drag coefficients obtained from the 115-scale tests ranged from 0.0094 to 0.0273 for the YF-102 model and from 0.0100 to 0.0255 for the F-102A model. Forebody external-pressure-drag coefficients (drag rise) at Mach number 1.05 of 0.0183 and 0.0134 were obtained from the 115-scale models of the YF-102 and F-102A, respectively, a 16-percent reduction for the F-102A model. Values of drag rise at Mach number 1.05 from the small equivalent-body tests were nearly the same for the basic, YF-102, and 2-foot-fuselage-extension airplane shapes. Equivalent-body tests of the YF-102 and F-102A shapes showed the latter to have about 25 percent less drag rise as compared with a 16-percent reduction illustrated by the 1/5-scale tests. Additional equivalent-body tests illustrating effects of modifications to the F-102A airplane shape shared that relocating the inlets on the fuselage or altering the nose shape to provide a smoother cross-sectional area progression reduced the drag rise by approximately 16 percent. Replacing a major portion of the nose of the F-102A equivalent-body model with one of parabolic shape resulted in about a 21-percent reduction in drag rise. The drag-rise data from the equivalent-body tests include base drag.

Wallskog, Harvey A.↗
NASA NTRS2006

Demonstration of a Non-Toxic Reaction Control Engine

Three non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration s (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration engine utilized Liquid Oxygen (LOX) and Ethanol as propellants to produce 870 lbf thrust. The Aerojet RCE s were successfully acceptance tested over a broad range of operating conditions. Steady state tests evaluated engine response to varying chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was also tested over a range of percent duty cycles (DC), and bit impulse and pulsing specific impulse were determined for each of these conditions. White Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test article that will be utilized in an altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty cycle map of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance test results and a status of the engine performance as part of the system level testing. Subsequent to acceptance testing at Aerojet, these three engines were delivered to the NASA

Robinson, Philip J.↗
NASA NTRS2007

Demonstration of a Non-Toxic Reaction Control Engine

T:hree non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration engine utilized Liquid Oxygen (LOX) and Ethanol as propellants to produce 870 lbf thrust. The Aerojet RCE's were successfully acceptance tested over a broad range of operating conditions. Steady state tests evaluated engine response to varying chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was also tested over a range of percent duty cycles (DC), and bit impulse and pulsing specific impulse were determined for each of these conditions. Subsequent to acceptance testing at Aerojet, these three engines were delivered to the NASA White Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test article that will be utilized in an altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty cycle map of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance test results and a status of the engine performance as part of the system level testing.

Robinson, Philip J.↗
NASA NTRS2016

Development of Detonation Modeling Capabilities for Rocket Test Facilities: Hydrogen-Oxygen-Nitrogen Mixtures

The objective of the presented work was to develop validated computational fluid dynamics (CFD) based methodologies for predicting propellant detonations and their associated blast environments. Applications of interest were scenarios relevant to rocket propulsion test and launch facilities. All model development was conducted within the framework of the Loci/CHEM CFD tool due to its reliability and robustness in predicting high-speed combusting flow-fields associated with rocket engines and plumes. During the course of the project, verification and validation studies were completed for hydrogen-fueled detonation phenomena such as shock-induced combustion, confined detonation waves, vapor cloud explosions, and deflagration-to-detonation transition (DDT) processes. The DDT validation cases included predicting flame acceleration mechanisms associated with turbulent flame-jets and flow-obstacles. Excellent comparison between test data and model predictions were observed. The proposed CFD methodology was then successfully applied to model a detonation event that occurred during liquid oxygen/gaseous hydrogen rocket diffuser testing at NASA Stennis Space Center.

Allgood, Daniel C.↗
NASA NTRS1996

Reusable Launch Vehicle Technology Program

Industry/NASA Reusable Launch Vehicle (RLV) Technology Program efforts are underway to design, test, and develop technologies and concepts for viable commercial launch systems that also satisfy national needs at acceptable recurring costs. Significant progress has been made in understanding the technical challenges of fully reusable launch systems and the accompanying management and operational approaches for achieving a low-cost program. This paper reviews the current status of the Reusable Launch Vehicle Technology Program including the DC-XA, X-33 and X-34 flight systems and associated technology programs. It addresses the specific technologies being tested that address the technical and operability challenges of reusable launch systems including reusable cryogenic propellant tanks, composite structures, thermal protection systems, improved propulsion, and subsystem operability enhancements. The recently concluded DC-XA test program demonstrated some of these technologies in ground and flight tests. Contracts were awarded recently for both the X-33 and X-34 flight demonstrator systems. The Orbital Sciences Corporation X-34 flight test vehicle will demonstrate an air-launched reusable vehicle capable of flight to speeds of Mach 8. The Lockheed-Martin X-33 flight test vehicle will expand the test envelope for critical technologies to flight speeds of Mach 15. A propulsion program to test the X-33 linear aerospike rocket engine using a NASA SR-71 high speed aircraft as a test bed is also discussed. The paper also describes the management and operational approaches that address the challenge of new cost-effective, reusable launch vehicle systems.

Freeman, Delma C., Jr.↗
NASA NTRS1997

Reusable Launch Vehicle Technology Program

Industry/NASA reusable launch vehicle (RLV) technology program efforts are underway to design, test, and develop technologies and concepts for viable commercial launch systems that also satisfy national needs at acceptable recurring costs. Significant progress has been made in understanding the technical challenges of fully reusable launch systems and the accompanying management and operational approaches for achieving a low cost program. This paper reviews the current status of the RLV technology program including the DC-XA, X-33 and X-34 flight systems and associated technology programs. It addresses the specific technologies being tested that address the technical and operability challenges of reusable launch systems including reusable cryogenic propellant tanks, composite structures, thermal protection systems, improved propulsion and subsystem operability enhancements. The recently concluded DC-XA test program demonstrated some of these technologies in ground and flight test. Contracts were awarded recently for both the X-33 and X-34 flight demonstrator systems. The Orbital Sciences Corporation X-34 flight test vehicle will demonstrate an air-launched reusable vehicle capable of flight to speeds of Mach 8. The Lockheed-Martin X-33 flight test vehicle will expand the test envelope for critical technologies to flight speeds of Mach 15. A propulsion program to test the X-33 linear aerospike rocket engine using a NASA SR-71 high speed aircraft as a test bed is also discussed. The paper also describes the management and operational approaches that address the challenge of new cost effective, reusable launch vehicle systems.

Freeman, Delma C., Jr.↗
NASA NTRS1947

Force Tests of the Boeing XB-47 Full-Scale Empennage in the Ames 40- by 80-Foot Wind Tunnel

A wind-tunnel investigation of the Boeing XB-47 full-scale empennage was conducted to provide, prior to flight tests, data required on the effectiveness of the elevator and rudder. The XB-47 airplane is a jet-propelled medium bomber having wing and tail surfaces swept back 35 degrees. The investigation included tests of the effectiveness of the elevator with normal straight sides, with a buldged trailing edge, and with a modified hinge-line gap and tests of the effectiveness of the rudder with a normal straight-sided tab and with a bulged tab.

Hunton, Lynn W.↗
NASA NTRS1968

Solid propellant engineering

Surface temperature relations for ignition material deflagration onset, and static tests of Applications Technology Satellite motor

Source record↗
NASA NTRS2006

Metallized Gelled Propellants Combustion Experiments in a Pulse Detonation Engine

A series of combustion tests were performed with metallized gelled JP 8/aluminum fuels in a Pulse Detonation Engine (PDE). Nanoparticles of aluminum were used in the 60 to 100 nanometer diameter. Gellants were also of a nanoparticulate type composed of hydrocarbon alkoxide materials. Using simulated air (a nitrogen-oxygen mixture), the ignition potential of metallized gelled fuels with nanoparticle aluminum was investigated. Ignition of the JP 8/aluminum was possible with less than or equal to a 23-wt% oxygen loading in the simulated air. JP 8 fuel alone was unable to ignite with less than 30 percent oxygen loaded simulated air. The tests were single shot tests of the metallized gelled fuel to demonstrate the capability of the fuel to improve fuel detonability. The tests were conducted at ambient temperatures and with maximal detonation pressures of 1340 psia.

Palaszewski, Bryan↗
NASA NTRS2012

Fast Paced, Low Cost Projects at MSFC

What does an orbiting microsatellite, a robotic lander and a ruggedized camera and telescope have in common? They are all fast paced, low cost projects managed by Marshall Space Flight Center (MSFC) teamed with successful industry partners. MSFC has long been synonymous with human space flight large propulsion programs, engineering acumen and risk intolerance. However, there is a growing portfolio/product line within MSFC that focuses on these smaller, fast paced projects. While launching anything into space is expensive, using a managed risk posture, holding to schedule and keeping costs low by stopping at egood enough f were key elements to their success. Risk is defined as the possibility of loss or failure per Merriam Webster. The National Aeronautics and Space Administration (NASA) defines risk using procedural requirement 8705.4 and establishes eclasses f to discern the acceptable risk per a project. It states a Class D risk has a medium to significant risk of not achieving mission success. MSFC, along with industry partners, has created a niche in Class D efforts. How did the big, cautious MSFC succeed on these projects that embodied the antithesis of its heritage in human space flight? A key factor toward these successful projects was innovative industry partners such as Dynetics Corporation, University of Alabama in Huntsville (UAHuntsville), Johns Hopkins Applied Physics Laboratory (JHU APL), Teledyne Brown Engineering (TBE), Von Braun Center for Science and Innovation (VCSI), SAIC, and Jacobs. Fast Affordable Satellite Technology (FastSat HSV01) is a low earth orbit microsatellite that houses six instruments with the primary scientific objective of earth observation and technology demonstration. The team was comprised of Dynetics, UAHuntsvile, SAIC, Goddard Space Flight Center (GSFC) and VCSI with the United States Air Force Space Test Program as the customer. The team completed design, development, manufacturing, environmental test and integration in one year. FastSat HSV01 also deployed a Poly Picosatellite Orbital Deployer (PPOD) for a separate nano ]satellite class spacecraft (Cubesat: Nano Sail Demonstration) in partnership with Ames Research Center. The Robotic lunar lander is a MSFC JHU APL partnership that led to the development of a flexible architecture for landers to support robotic missions to a wide range of lunar and asteroid destinations. The team started with the goal of meeting NASA agency directives that led to the creation of a test bed focusing on GN&C and software to demonstrate the descent and landing on any airless body for the final 30 to 60 meters. The team created a complex technology demonstration as well as Guidance Control and Navigation (GN&C) algorithms providing autonomous control of the lander. The team uses a green propellant of 90% hydrogen peroxide and has completed 18 successful test flights. The International Space Station (ISS) SERVIR Environmental Research and Visualization System (ISERV) is a technology demonstration payload to assist the SERVIR project with environmental monitoring for disaster relief and humanitarian efforts. The ISERV project was a partnership with TBE. The ISERV payload consists of a commercial off the shelf camera, telescope, and MSFC developed power distribution box and interfaces on ISS with the Window Observational Research Facility in the US Lab. MSFC has identified three key areas that enabled the low cost mission success to include culture, partnering, and cost/schedule control. This paper will briefly discuss these three Class D efforts, FastSat HSV-01, the Robotic Lunar Lander and the ISERV camera system, the lessons learned, their successes and challenges.

Watson-Morgan, Lisa↗