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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 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Green Propellant Test Capabilities of the Altitude Combustion Stand at the NASA Glenn Research Center

The NASA Glenn Research Center (GRC) is committed to providing simulated altitude rocket test capabilities to NASA programs, other government agencies, private industry partners, and academic partners. A primary facility to support those needs is the Altitude Combustion Stand (ACS). ACS provides the capability to test combustion components at a simulated altitude up to 100,000 ft. (approx.0.2 psia/10 Torr) through a nitrogen-driven ejector system. The facility is equipped with an axial thrust stand, gaseous and cryogenic liquid propellant feed systems, data acquisition system with up to 1000 Hz recording, and automated facility control system. Propellant capabilities include gaseous and liquid hydrogen, gaseous and liquid oxygen, and liquid methane. A water-cooled diffuser, exhaust spray cooling chamber, and multi-stage ejector systems can enable run times up to 180 seconds to 16 minutes. The system can accommodate engines up to 2000-lbf thrust, liquid propellant supply pressures up to 1800 psia, and test at the component level. Engines can also be fired at sea level if needed. The NASA GRC is in the process of modifying ACS capabilities to enable the testing of green propellant (GP) thrusters and components. Green propellants are actively being explored throughout government and industry as a non-toxic replacement to hydrazine monopropellants for applications such as reaction control systems or small spacecraft main propulsion systems. These propellants offer increased performance and cost savings over hydrazine. The modification of ACS is intended to enable testing of a wide range of green propellant engines for research and qualification-like testing applications. Once complete, ACS will have the capability to test green propellant engines up to 880 N in thrust, thermally condition the green propellants, provide test durations up to 60 minutes depending on thrust class, provide high speed control and data acquisition, as well as provide advanced imaging and diagnostics such as infrared (IR) imaging.

Testing↗
NASA NTRS2023

Passive Rocket Diffuser Testing: Reacting Flow Performance of Advanced Configurations

Upper stage rocket engines are optimized for operation at the edge of space. Ground testing often requires simulation of high-altitude environments to prevent damage from off-design operation. Passive supersonic diffusers are an effective means of establishing the partial vacuum needed around the nozzle exit. The low cost of diffusers relative to active vacuum sources (ejectors, pumps, etc.) provides a perpetual incentive to improve their performance and expand the envelope of passively testable engines. Second-throat diffusers have been the de facto standard for rocket testing since the 1950s, and numerous studies have focused on parametric optimization of the topology. Yet, even the best second-throat designs may require more driving pressure than an upper stage engine can provide. To overcome conventional limits, engineers at NASA’s Stennis Space Center (SSC) adapted an obscure cold-flow diffuser topology to the extremes of immersion in supersonic combustion products. Aerodynamic experimentation was conducted at SSC’s E-3 test stand to quantify the performance of prototypical hot-fire spike diffuser hardware and enable direct comparison to prior states of the art. A LO X /GH 2 thruster was contoured to approximate the regeneratively-cooled nozzle of an upper-stage engine and fired into 28 downstream diffuser configurations: 16 second-throat, 2 centerbody, and 10 spike. Temperature and pressure were recorded along the test article walls for a range of chamber pressures spanning 3.5 - 5.0 MPa. The second-throat configurations with the lowest start and unstart pressure ratios were shown to outperform equivalent systems in literature and taken as baselines for evaluation of advanced diffuser performance. Compared to these exemplars, centerbody diffusers decreased start pressure ratio by 11% but increased unstart pressure ratio by 15%. Spike diffusers reduced start and unstart pressure ratios by 24% and 9%, respectively.

Rocket↗
NASA NTRS1967

Air sampler collects and protects minute particles

Air ejector impactor sampler collects and protects samples of particles greater than 0.1 micron in diameter. In operation, it causes impaction of particle-laden air onto several collection surfaces within a collection cylinder. When not operating, the collector cylinder is maintained in a retracted state within a protective envelope.

Wood, R. C.↗
NASA NTRS1968

Two-fluid, impinging-sheet injector

Two-fluid, impinging-sheet propellant injector reduces the severe erosion found to occur when ejector elements are directly exposed during throttling without the benefits of a cooling flow of the propellant liquids. It greatly improves combustion efficiency by venting the secondary stream of combustion gases generated by backspray reaction.

Riebling, R. W.↗
NASA NTRS1972

Peak axial-velocity decay with single- and multi-element nozzles

Jet peak-velocity decay data were obtained for a variety of circular and noncircular single-element and multi-element nozzles for application to externally-blown-flap (EBF) STOL aircraft. These data permit a rational approach, in terms of element type and element spacing, to nozzles designed to promote mixing of the jet exhaust with the surrounding air. Rapid mixing and the resulting lower axial jet velocity decreases the noise caused by the interaction of jet impingement on the flap assembly of EBF STOL aircraft. Empirical relationships are presented that permit the prediction of peak axial-velocity decay curves for a wide spectrum of mixer-type nozzles. The data are useful also in the design of ejector-type noise suppressors and for the suppression of VTOL downwash velocities caused by vertically oriented exhaust nozzles.

Vonglahn, U. H.↗
NASA NTRS1972

Analysis of a supersonic-combustion rocket concept

A preliminary analysis has been made of a supersonic-combustion rocket engine concept using hydrogen and oxygen propellants. The ejector action of a separate small rocket motor is employed to pump the propellants to high stagnation pressures and supersonic velocities. Therefore complicated heavy turbopumps are eliminated and cooling problems of a sonic throat are reduced. The results of the study show that vacuum specific impulse levels as high as a conventional rocket having the same chamber pressure as the drive motor are possible. The supersonic-combustion rocket would be an attractive alternate for a high-altitude low-thrust conventional rocket operating with a pressure feed propellant system. It would also be a convenient technique for obtaining extremely high thrusts without the need for developing corresponding large turbopumps.

Franciscus, L. C.↗
NASA NTRS1972

Peak axial-velocity decay with single- and multi-element nozzles.

Jet peak-velocity decay data were obtained for a variety of circular and noncircular single-element and multi-element nozzles for application to externally-blown-flap STOL aircraft. These data permit a rational approach, in terms of element type and element spacing, to nozzles designed to promote mixing of the jet exhaust with the surrounding air. Rapid mixing and the resulting lower axial jet velocity decreases the noise caused by the interaction of jet impingement on the flap assembly of EBF STOL aircraft. Empirical relationships are presented that permit the prediction of peak axial-velocity decay curves for a wide spectrum of mixer-type nozzles. The data are useful also in the design of ejector-type noise suppressors and for the suppression of VTOL downwash velocities caused by vertically oriented exhaust nozzles.

Von Glahn, U. H.↗
NASA NTRS1972

Jet noise

The main conclusions reached are: (1) At low subsonic jet exhaust velocities, jet noise varies as the velocity to the eighth power; (2) at high subsonic exhaust velocities, jet noise approaches a variation with velocity to the third power; (3) use of the jet density squared overcorrects density effects on jet noise; (4) subsonic and supersonic jet noise levels can be predicted; and (6) use of multielement nozzles and acoustically lined ejectors significantly suppress jet noise at intermediate jet exhaust velocities.

Vonglahn, U. H.↗
NASA NTRS1972

Flyover and static tests to investigate external flow effect on jet noise from nonsuppressor and suppressor exhaust nozzles

The effect of external air flowing across exhaust nozzles on the jet noise characteristics of supersonic transport aircraft at high takeoff speeds was investigated. A series of flyover and static tests were conducted using an F-106B aircraft modified with two underwing nacelles each containing a calibrated J85-GE-13 turbojet engine. Comparison of flyover and static data indicated that external flow reduces the noise of an auxiliary inlet ejector nozzle. An unsuppressed plug nozzle was not affected while the plug suppressor configurations were not as effective in flight.

Burley, R. R.↗
NASA NTRS1973

Apparatus for sampling particulates in gases

A device for sampling particulates in gases is described. The device is used to obtain samples of the upper atmosphere. The equipment used a common source of gas pressure to provide the driving gas of an air ejector pump. The sample collection cylinder has many slit impactors running longitudinally on the outer surface of a cylinder and terminating just short of each end of the cylinder.

Wood, R. C.↗
NASA NTRS1973

Flight investigation of acoustic and thrust characteristics of several exhaust nozzles installed on underwing nacelles on an F106 airplane

To determine flyover noise and thrust and to investigate whether flight velocity significantly affects the noise of exhaust nozzles, a series of flight tests was conducted on three different exhaust nozzles of a type suitable for supersonic transport aircraft. The tests were conducted using an F-106B aircraft modified to carry two underwing nacelles, each containing a calibrated turbojet engine. A flyover altitude of 91 meters (300 ft) and a Mach number of 0.4 provided acoustic data that were repeatable to within + or -1.5 PNdB. Flyover results showed that an auxiliary inlet ejector nozzle was the quietest of the nozzles tested; flight velocity appeared to reduce its noise.

Burley, R. R.↗