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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 793 records · Page 44
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Investigation of a pulsed electrothermal thruster system

The performance of an ablative wall Pulsed Electrothermal (PET) thruster is accurately characterized on a calibrated thrust stand, using polyethylene propellant. The thruster is tested for four configurations of capillary length and pulse length. The exhaust velocity is determined with twin time-of-flight photodiode stagnation probes, and the ablated mass is measured from the loss over ten shots. Based on the measured thrust impulse and the ablated mass, the specific impulse varies from 1000 to 1750 seconds. The thrust to power varies from .05 N/kW (quasi-steady mode) to .10 N/kW (unsteady mode). The thruster efficiency varies from .56 at 1000 seconds to .42 at 1750 seconds. A conceptual design is presented for a 40 kW PET propulsion system. The point design system performance is .62 system efficiency at 1000 seconds specific impulse. The system's reliability is enhanced by incorporating 20, 20 kW thruster modules which are fired in pairs. The thruster design is non-ablative, and uses water propellant, from a central storage tank, injected through the cathode.

Burton, R. L.↗
NASA NTRS1954

Free-Flight Zero-Lift Drag Results from a 1/5-Scale Model and Several Small-Scale Equivalent Bodies of Revolution of the Convair F-102 Configuration at Mach Numbers up to 1.34

A 1/5-scale, rocket-propelled model of the Convair F-102 configuration was tested in free flight to determine zero-lift drag at Mach numbers up to 1.34 and at Reynolds numbers comparable to those of the full-scale airplane. This large-scale model corresponded to the prototype airplane and had air flow through the duct. Additional zero-lift drag tests involved a series of small equivalent bodies of revolution which were launched by means of a helium gun. The several small-scale models tested corresponded to: the basic configuration, the 1/5-scale rocket-propelled model configuration, a 2-foot (full-scale) fuselage-extension configuration, and a 7-foot (full-scale) fuselage-extension configuration. Models designed to correspond to the area distribution at a Mach number of 1.0 were flown for each of these 'shapes and, in addition, models designed to correspond to the area distribution at a Mach number of 1.2 were flown for the 1/5-scale rocket-propelled model and the 7-foot-fuselage-extension configuration. The value of external pressure drag coefficient (including base drag) obtained from the large-scale rocket model was 0.0190 at a Mach number of 1..05 and the corresponding values from the equivalent-body tests varied from 0.0183 for the rocket-propelled model shape to 0.0137 for the 7-foot-fuselage-extension configuration. From the results of tests of equivalent bodies designed to correspond to the area distribution at a Mach number of 1.0, it is evident that the small changes in shape incorporated in the basic and 2-foot-fuselage-extension configurations from that of the rocket-propelled model configuration will provide no significant change in pressure drag. On the other hand, the data from the 7-foot-fuselage-extension model indicate a substantial reduction in pressure drag at transonic speeds.

Wallskog, Harvey A.↗
NASA NTRS2010

Igniter Testing and Development for Liquid Oxygen/Liquid Methane (LOX/LCH4) and Liquid Oxygen/Gaseous Hydrogen (LOX/GH2) Injectors

In support of NASA's efforts to advance methane technology for future exploration missions, two different igniter concepts were developed and tested at NASA's Marshall Space Flight Center for ignition of oxygen/methane propellants. The first type of igniter, an impinging injection, spark-initiated torch igniter (impinging igniter), used traditional spark plug technology to produce a small, but very intense, pulsed plasma as the ignition source. The second type of igniter, a microwave-generated plasma igniter (microwave igniter), used a microwave magnetron to produce a continuous, larger-volume plasma as the ignition source. Two new designs for both types of igniters were leveraged off past oxygen/methane igniter designs that were tested at MSFC and documented in a 2008 JANNAF paper. In this most recent work, the two new igniters were to be tested at the component level as well as inside a liquid oxygen/liquid methane (LOX/LCH4) 40-element swirl coaxial injector. Although the microwave igniter achieved ignition in every test, it encountered technical difficulties that prevented it from being tested with the injector. The impinging igniter required several design iterations as well as testing at different mixture ratios and mass flow rates before reliable operation with minimal ignition delay was achieved. The final design of the impinging igniter was successfully tested in the LOX/LCH4 injector. The same igniter was later re-configured and component-tested to demonstrate reliable ignition of oxygen/hydrogen propellants. Future tests of the impinging igniter will include ignition of a LOX/hydrogen two-stage swirl injector, which is one of several injector concepts being tested in the Lunar Lander Descent Engine Test Bed at MSFC. All igniter and injector tests have been funded by the Propulsion & Cryogenics Advanced Development project managed by Glenn Research Center and sponsored by NASA s Exploration Technology Development Program.

Osborne, Robin J.↗
NASA NTRS1992

Large scale cryogenic fluid systems testing

NASA Lewis Research Center's Cryogenic Fluid Systems Branch (CFSB) within the Space Propulsion Technology Division (SPTD) has the ultimate goal of enabling the long term storage and in-space fueling/resupply operations for spacecraft and reusable vehicles in support of space exploration. Using analytical modeling, ground based testing, and on-orbit experimentation, the CFSB is studying three primary categories of fluid technology: storage, supply, and transfer. The CFSB is also investigating fluid handling, advanced instrumentation, and tank structures and materials. Ground based testing of large-scale systems is done using liquid hydrogen as a test fluid at the Cryogenic Propellant Tank Facility (K-site) at Lewis' Plum Brook Station in Sandusky, Ohio. A general overview of tests involving liquid transfer, thermal control, pressure control, and pressurization is given.

Source record↗
NASA NTRS2016

Green Propellant Landing Demonstration at U.S. Range

The Green Propellant Loading Demonstration (GPLD) was conducted December 2015 at Wallops Flight Facility (WFF), leveraging work performed over recent years to bring lower toxicity hydrazine replacement green propellants to flight missions. The objective of this collaboration between NASA Goddard Space Flight Center (GSFC), WFF, the Swedish National Space Board (SNSB), and Ecological Advanced Propulsion Systems (ECAPS) was to successfully accept LMP-103S propellant at a U.S. Range, store the propellant, and perform a simulated flight vehicle propellant loading. NASA GSFC Propulsion (Code 597) managed all aspects of the operation, handling logistics, preparing the procedures, and implementing the demonstration. In addition to the partnership described above, Moog Inc. developed an LMP-103S propellant-compatible titanium rolling diaphragm flight development tank and loaned it to GSFC to act as the GPLD flight vessel. The flight development tank offered the GPLD an additional level of flight-like propellant handling process and procedures. Moog Inc. also provided a compatible latching isolation valve for remote propellant expulsion. The GPLD operation, in concert with Moog Inc. executed a flight development tank expulsion efficiency performance test using LMP-103S propellant. As part of the demonstration work, GSFC and WFF documented Range safety analyses and practices including all elements of shipping, storage, handling, operations, decontamination, and disposal. LMP-103S has not been previously handled at a U.S. Launch Range. Requisite for this activity was an LMP-103S Risk Analysis Report and Ground Safety Plan. GSFC and WFF safety offices jointly developed safety documentation for application into the GPLD operation. The GPLD along with the GSFC Propulsion historical hydrazine loading experiences offer direct comparison between handling green propellant versus safety intensive, highly toxic hydrazine propellant. These described motives initiated the GPLD operation in order to investigate the handling and process safety variances in project resources between LMP-103S and typical in-space propellants. The GPLD risk reduction operation proved successful for many reasons including handling the green propellant at a U.S. Range, loading and pressurizing a flight-like tank, expelling the propellant, measuring the tank expulsion efficiency, and most significantly, GSFC propulsion personnel's new insight into the LMP-103S propellant handling details.

Propellant↗
NASA NTRS2016

Green Propellant Loading Demonstration at U.S. Range

The Green Propellant Loading Demonstration (GPLD) was conducted December 2015 at Wallops Flight Facility (WFF), leveraging work performed over recent years to bring lower toxicity hydrazine replacement green propellants to flight missions. The objective of this collaboration between NASA Goddard Space Flight Center (GSFC), WFF, the Swedish National Space Board (SNSB), and Ecological Advanced Propulsion Systems (ECAPS) was to successfully accept LMP-103S propellant at a U.S. Range, store the propellant, and perform a simulated flight vehicle propellant loading. NASA GSFC Propulsion (Code 597) managed all aspects of the operation, handling logistics, preparing the procedures, and implementing the demonstration. In addition to the partnership described above, Moog Inc. developed an LMP-103S propellant-compatible titanium rolling diaphragm flight development tank and loaned it to GSFC to act as the GPLD flight vessel. The flight development tank offered the GPLD an additional level of flight-like propellant handling process and procedures. Moog Inc. also provided a compatible latching isolation valve for remote propellant expulsion. The GPLD operation, in concert with Moog Inc. executed a flight development tank expulsion efficiency performance test using LMP-103S propellant. As part of the demonstration work, GSFC and WFF documented Range safety analyses and practices including all elements of shipping, storage, handling, operations, decontamination, and disposal. LMP-103S has not been previously handled at a U.S. Launch Range. Requisite for this activity was an LMP-103S Risk Analysis Report and Ground Safety Plan. GSFC and WFF safety offices jointly developed safety documentation for application into the GPLD operation. The GPLD along with the GSFC Propulsion historical hydrazine loading experiences offer direct comparison between handling green propellant versus safety intensive, highly toxic hydrazine propellant. These described motives initiated the GPLD operation in order to investigate the handling and process safety variances in project resources between LMP-103S and typical in-space propellants. The GPLD risk reduction operation proved successful for many reasons including handling the green propellant at a U.S. Range, loading and pressurizing a flight-like tank, expelling the propellant, measuring the tank expulsion efficiency, and most significantly, GSFC propulsion personnel's new insight into the LMP-103S propellant handling details.

Green Propellant↗
NASA NTRS1971

Performance of a 10 mlb biowaste resistojet.

Description of the thrusting performance of a 10-mlb thrust, concentric-tubes biowaste resistojet, which was life tested on hydrogen and ammonia propellants for 8000 hours. Thrusting performance is presented for H2, H2O, CH4, and CO2 as well as typical biopropellant mixtures CO2-CH4, H2O-CH4 and CO2-H2O. Using Pt-Ir alloy heater tubes, 235 and 575 seconds were obtained for water and hydrogen propellants. Overall total power efficiencies were 66%, which includes losses for a preevaporator close-coupled for the water tests. Vibration to 1.0 sq q/Hz with an overall G rms of 36, shock at 30 g's and acceleration to 8 g's simulating a launch were passed by the thrustor.

Phillips, D. G.↗
NASA NTRS1989

Aircraft propeller induced structure-borne noise

A laboratory-based test apparatus employing components typical of aircraft construction was developed that would allow the study of structure-borne noise transmission due to propeller induced wake/vortex excitation of in-wake structural appendages. The test apparatus was employed to evaluate several aircraft installation effects (power plant placement, engine/nacelle mass loading, and wing/fuselage attachment methods) and several structural response modifications for structure-borne noise control (the use of wing blocking mass/fuel, wing damping treaments, and tuned mechanical dampers). Most important was the development of in-flight structure-borne noise transmission detection techniques using a combination of ground-based frequency response function testing and in-flight structural response measurement. Propeller wake/vortex excitation simulation techniques for improved ground-based testing were also developed to support the in-flight structure-borne noise transmission detection development.

Unruh, James F.↗
NASA NTRS2017

Design and Testing of a Hall Effect Thruster with 3D Printed Channel and Propellant Distributor

The UAH-78AM is a low-power Hall effect thruster developed at the University of Alabama in Huntsville with channel walls and a propellant distributor manufactured using 3D printing. The goal of this project is to assess the feasibility of using unconventional materials to produce a low-cost functioning Hall effect thruster and consider how additive manufacturing can expand the design space and provide other benefits. A version of the thruster was tested at NASA Glenn Research Center to obtain performance metrics and to validate the ability of the thruster to produce thrust and sustain a discharge. An overview of the thruster design and transient performance measurements are presented here. Measured thrust ranged from 17.2 millinewtons to 30.4 millinewtons over a discharge power of 280 watts to 520 watts with an anode I (sub SP)(Specific Impulse) range of 870 seconds to 1450 seconds. Temperature limitations of materials used for the channel walls and propellant distributor limit the ability to run the thruster at thermal steady-state.

Hall thrusters↗
DOE OSTI2019

Boundary Layer Transition Models for Naval Applications: Capabilities and Limitations

We describe the implementation of several recently developed boundary layer transition models into the overset computational fluid dynamics code, REX, developed at the University of Iowa, together with an evaluation of its capabilities and limitations for naval hydrodynamics applications. Models based on correlations and on amplification factor transport were implemented in one- and two-equation Reynolds-averaged Navier-Stokes turbulence models, including modifications to operate in crossflow. Extensive validation of the transition models implemented in REX is performed for several 2- and 3-dimensional geometries of naval relevance. Standard tests with extensive available experimental data include flat plates in zero pressure gradient, an airfoil, and sickle wing. More complex test cases include the propeller, P4119, with some experimental data available, and the generic submersible, Joubert BB2, with no relevant experimental data available, to validate the transition models. Finally, simulations for these last two cases show that extensive regions of laminar flow can be present on the bodies at laboratory scale and field scale for small vessels, and the potential effects on resistance and propulsion can be significant.

42 ENGINEERING↗
NASA NTRS1972

Exploratory study of the effects of injection configurations and inlet flow conditions on the characteristics of flow in spherical chambers

Flow visualization studies were conducted to evaluate techniques for injecting simulated-fuel and simulated-propellant gases into a spherical cavity for application to open-cycle gaseous-core nuclear rockets. Preliminary studies were conducted with six simulated-fuel injectors and eight simulated-propellant injection configurations. Additional tests were conducted with the best configuration to determine the effect of weight-flow ratio, gas density ratio, injector location, and flow distribution on the simulated-fuel containment characteristics.

Johnson, B. V.↗
NASA NTRS1981

Analysis of Mach number 0.8 turboprop slipstream wing/nacelle interactions

Data from wind tunnel tests of a powered propeller and nacelle mounted on a supercritical wing are analyzed. Installation of the nacelle significantly affected the wing flow and the flow on the upper surface of the wing is separated near the leading edge under powered conditions. Comparisons of various theories with the data indicated that the Neumann surface panel solution and the Jameson transonic solution gave results adequate for design purposes. A modified wing design was developed (Mod 3) which reduces the wing upper surface pressure coefficients and section lift coefficients at powered conditions to levels below those of the original wing without nacelle or power. A contoured over the wing nacelle that can be installed on the original wing without any appreciable interference to the wing upper surface pressure is described.

Welge, H. R.↗