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  • 1
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    Gas analysis system for the Eight Foot High Temperature Tunnel (1992)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Description: This paper describes the development of a gas collection and analysis system that is to be installed in the Eight-Foot High Temperature Tunnel (8' HTT) at NASA's Langley Research Center. This system will be used to analyze the test gas medium that results after burning a methane-air mixture to achieve the proper tunnel test parameters. The system consists of a sampling rake, a gas sample storage array, and a gas chromatographic system. Gas samples will be analyzed after each run to assure that proper combustion takes place in the tunnel resulting in a correctly balanced composition of the test gas medium. The proper ratio of gas species is critically necessary in order for the proper operation and testing of scramjet engines in the tunnel. After a variety of methane-air burn conditions have been analyzed, additional oxygen will be introduced into the combusted gas and the enriched test gas medium analyzed. The pre/post enrichment sets of data will be compared to verify that the gas species of the test gas medium is correctly balanced for testing of air-breathing engines.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: In: International Instrumentation Symposium, 38th, Las Vegas, NV, Apr. 26-30, 1992, Proceedings (A93-37851 15-35); p. 581-596.
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  • 2
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    Fuel Line Based Acoustic Flame-Out Detection System (1997)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An acoustic flame-out detection system that renders a large high pressure combustor safe in the event of a flame-out and possible explosive reignition. A dynamic pressure transducer is placed in the fuel and detects the stabilizing fuel pressure oscillations, caused by the combustion process. An electric circuit converts the signal from the combustion vortices, and transmitted to the fuel flow to a series of pulses. A missing pulse detector counts the pulses and continuously resets itself. If three consecutive pulses are missing, the circuit closes the fuel valve. With fuel denied the combustor is shut down or restarted under controlled conditions.
    Keywords: Instrumentation and Photography
    Format: application/pdf
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  • 3
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    Device for quickly sensing the amount of O2 in a combustion product gas (1990)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A sensing device comprising an O2 sensor, a pump, a compressor, and a heater is provided to quickly sense the amount of O2 in a combustion product gas. A sample of the combustion product gas is compressed to a pressure slightly above one atmosphere by the compressor. Next, the heater heats the sample between 800 C and 900 C. Next, the pump causes the sample to be flushed against the electrode located in O2 sensor 6000 to 10,000 times per second. Reference air at approximately one atmosphere is provided to the electrode of O2 sensor. Accordingly, the O2 sensor produces a voltage which is proportional to the amount of oxygen in the combustion product gas. This voltage may be used to control the amount of O2 entering into the combustion chamber which produces the combustion product gas.
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Format: application/pdf
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  • 4
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    Method and device for determining heats of combustion of gaseous hydrocarbons (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method and device is provided for a quick, accurate and on-line determination of heats of combustion of gaseous hydrocarbons. First, the amount of oxygen in the carrier air stream is sensed by an oxygen sensing system. Second, three individual volumetric flow rates of oxygen, carrier stream air, and hydrocrabon test gas are introduced into a burner. The hydrocarbon test gas is fed into the burner at a volumetric flow rate, n, measured by a flowmeter. Third, the amount of oxygen in the resulting combustion products is sensed by an oxygen sensing system. Fourth, the volumetric flow rate of oxygen is adjusted until the amount of oxygen in the combustion product equals the amount of oxygen previously sensed in the carrier air stream. This equalizing volumetric flow rate is m and is measured by a flowmeter. The heat of combustion of the hydrocrabon test gas is then determined from the ratio m/n.
    Keywords: INORGANIC AND PHYSICAL CHEMISTRY
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  • 5
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    Modification to the Langley 8-Foot High Temperature Tunnel for hypersonic propulsion testing (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: This paper describes the modifications currently underway to the Langley 8-Foot High Temperature Tunnel to produce a new, unique national resource for testing of hypersonic air-breathing propulsion systems. The current tunnel, which has been used for aerothermal loads and structures research since its inception, is being modified with the addition of a LOX system to bring the oxygen content of the test medium up to that of air, the addition of alternate Mach number capability to augment the current M = 7 capability, improvements to the tunnel hardware to reduce maintenance downtime, the addition of a hydrogen system to allow the testing of hydrogen powered engines, and a new data system to increase both the quantity and quality of the data obtained. The paper discusses both the modifications and the development thereof.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: AIAA PAPER 87-1887
    Format: text
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  • 6
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    The NASA Langley Scramjet Test Complex (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The NASA Langley Scramjet Test Complex consists of five propulsion facilities which cover a wide spectrum of supersonic combustion ramjet (scramjet) test capabilities. These facilities permit observation of the effects on scramjet performance of speed and dynamic pressure from Mach 3.5 to near-orbital speeds, engine size from Mach 4 to 7, and test gas composition from Mach 4 to 7. In the Mach 3.5 to 8 speed range, the complex includes a direct-connect combustor test facility, two small-scale complete engine test facilities, and a large-scale complete engine test facility. In the hypervelocity speed range, a shock-expansion tube is used for combustor tests from Mach 12 to Mach 17+. This facility has recently been operated in a tunnel mode, to explore the possibility of semi-free-jet testing of complete engine modules at hypervelocity conditions. This paper presents a description of the current configurations and capabilities of the facilities of the NASA Langley Scramjet Test Complex, reviews the most recent scramjet tests in the facilities, and discusses comparative engine tests designed to gain information about ground facility effects on scramjet performance.
    Keywords: Research and Support Facilities (Air)
    Type: NASA-TM-111658 , NAS 1.15:111658 , AIAA Paper 96-3243 , AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 01, 1996; Lake Buena Vista, FL; United States|AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
    Format: text
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  • 7
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    Research capabilities of the NASA Langley 8-foot high temperature tunnel (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The NASA Langley Research Center's 8-Foot High Temperature Tunnel (8' HTT) has been modified to facilitate the testing of hypersonic airbreathing propulsion systems in addition to aerothermal load definition and structural concept verification at Mach 4, 5, and 7. The 8' HTT simulates flight from 60 to 125 kft with run times of 1 to 2 min. The 8-ft diameter and 12-ft-long free jet length enables the testing of large engines or multiple subscale engines. Methane and air combustion products provide the true temperature environment; an oxygen system enriches the combustion products to the same volume fraction of oxygen as air to enable the testing of airbreathing engines. A hydrogen system provides model fuel for cooling and combustion. High use tunnel components have been upgraded or replaced.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: Structural Testing Technology at High Temperature; Nov 04, 1991 - Nov 06, 1991; Dayton, OH; United States
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  • 8
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    Improved Insert For Variable Mach Number (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Nozzle insert lowers wind-tunnel mach number while maintaining excellent flow quality. Essential components of improved design include sonic first throat, expansion surface, variable boundary-layer bleed, insert itself, and existing, unchanged, contour of nozzle. Modification involves creation of secondary throat and critical addition of boundary-layer bleed path between insert and original tunnel wall. Enables quick change of mach number in existing facility at relatively low cost and requires no special contouring of existing nozzle. Represents simple and cost-effective method of altering mach number in any supersonic wind tunnel.
    Keywords: MECHANICS
    Type: LAR-13548 , NASA Tech Briefs (ISSN 0145-319X); 14; 7; P. 63
    Format: text
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  • 9
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    Compact Analyzer/Controller For Oxygen-Enrichment System (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: System controls hypersonic air-breathing engine tests. Compact analyzer/controller developed, built, and tested in small-scale wind tunnel prototype of the 8' HTT (High-Temperature Tunnel). Monitors level of oxygen and controls addition of liquid oxygen to enrich atmosphere for combustion. Ensures meaningful ground tests of hypersonic engines in range of speeds from mach 4 to mach 7.
    Keywords: ELECTRONIC SYSTEMS
    Type: LAR-14016 , NASA Tech Briefs (ISSN 0145-319X); 14; 6; P. 50
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  • 10
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    Hypersonic Airbreathing Propulsion: An Aerodynamics, Aerothermodynamics, and Acoustics Competency White Paper (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: This White Paper examines the current state of Hypersonic Airbreathing Propulsion at the NASA Langley Research Center and the factors influencing this area of work and its personnel. Using this knowledge, the paper explores beyond the present day and suggests future directions and strategies for the field. Broad views are first taken regarding potential missions and applications of hypersonic propulsion. Then, candidate propulsion systems that may be applicable to these missions are suggested and discussed. Design tools and experimental techniques for developing these propulsion systems are then described, and approaches for applying them in the design process are considered. In each case, current strategies are reviewed and future approaches that may improve the techniques are considered. Finally, the paper concentrates on the needs to be addressed in each of these areas to take advantage of the opportunities that lay ahead for both the NASA Langley Research Center and the Aerodynamic Aerothermodynamic, and Aeroacoustics Competency. Recommendations are then provided so that the goals set forth in the paper may be achieved.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-2002-211951 , L-18110 , NAS 1.15:211951
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