Results for “ENGINES - FUEL INJECTION” — Page 17

NASA NTRS1986

Improved silicon nitride for advanced heat engines

Silicon nitride is a high temperature material currently under consideration for heat engine and other applications. The objective is to improve the net shape fabrication technology of Si3N4 by injection molding. This is to be accomplished by optimizing the process through a series of statistically designed matrix experiments. To provide input to the matrix experiments, a wide range of alternate materials and processing parameters was investigated throughout the whole program. The improvement in the processing is to be demonstrated by a 20 percent increase in strength and a 100 percent increase in the Weibull modulus over that of the baseline material. A full characterization of the baseline process was completed. Material properties were found to be highly dependent on each step of the process. Several important parameters identified thus far are the starting raw materials, sinter/hot isostatic pressing cycle, powder bed, mixing methods, and sintering aid levels.

Yeh, H. C.↗

NASA NTRS2011

Advanced Development of a Compact 5-15 lbf Lox/Methane Thruster for an Integrated Reaction Control and Main Engine Propulsion System

This paper describes the advanced development and testing of a compact 5 to 15 lbf LOX/LCH4 thruster for a pressure-fed integrated main engine and RCS propulsion system to be used on a spacecraft "vertical" test bed (VTB). The ability of the RCS thruster and the main engine to operate off the same propellant supply in zero-g reduces mass and improves mission flexibility. This compact RCS engine incorporates several features to dramatically reduce mass and parts count, to ease manufacturing, and to maintain acceptable performance given that specific impulse (Isp) is not the driver. For example, radial injection holes placed on the chamber body for easier drilling, and high temperature Haynes 230 were selected for the chamber over other more expensive options. The valve inlets are rotatable before welding allowing different orientations for vehicle integration. In addition, the engine design effort selected a coil-on-plug ignition system which integrates a relay and coil with the plug electrode, and moves some exciter electronics to avionics driver board. The engine injector design has small dribble volumes to target minimum pulse widths of 20 msec. and an efficient minimum impulse bit of less than 0.05 lbf-sec. The propellants, oxygen and methane, were chosen because together they are a non-toxic, Mars-forward, high density, space storable, and high performance propellant combination that is capable of pressure-fed and pump-fed configurations and integration with life support and power subsystems. This paper will present the results of the advanced development testing to date of the RCS thruster and the integration with a vehicle propulsion system.

Hurlbert, Eric A.↗

NASA NTRS1978

Effects of film injection on performance of a cooled turbine

Tests were conducted in a 20 inch diameter single stage air cooled turbine designed to evaluate the effects of film cooling air on turbine aerodynamic performance. A comparison was made of the experimental results and an analytical method of evaluating film injection effects on turbine performance. The results are used to determine the effects of film cooling on overall engine performance for selected cycle conditions. The engine performance studies are used to show the cycle benefits of increased gas temperature at various coolant flow rates.

Mcdonel, J. D.↗

NASA NTRS1986

Computing Cooling Flows in Turbines

Algorithm developed for calculating both quantity of compressor bleed flow required to cool turbine and resulting decrease in efficiency due to cooling air injected into gas stream. Program intended for use with axial-flow, air-breathing, jet-propulsion engines with variety of airfoil-cooling configurations. Algorithm results compared extremely well with figures given by major engine manufacturers for given bulk-metal temperatures and cooling configurations. Program written in FORTRAN IV for batch execution.

Gauntner, J.↗

NASA NTRS1989

A Rocket Engine Design Expert System

The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are described. The expert system incorporates a JANNAF standard reference computer code to determine rocket engine performance and a state of the art finite element computer code to calculate the interactions between propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb heuristics were incorporated for the H2-O2 coaxial injector design, including a minimum gap size constraint on the total number of injector elements. One dimensional equilibrium chemistry was used in the energy release analysis of the combustion chamber. A 3-D conduction and/or 1-D advection analysis is used to predict heat transfer and coolant channel wall temperature distributions, in addition to coolant temperature and pressure drop. Inputting values to describe the geometry and state properties of the entire system is done directly from the computer keyboard. Graphical display of all output results from the computer code analyses is facilitated by menu selection of up to five dependent variables per plot.

Davidian, Kenneth J.↗

NASA NTRS2008

Influence of Reduced Mass Flow Rate and Chamber Backpressure on Swirl Injector Fluid Mechanics

Industry interest in variable-thrust liquid rocket engines places a demand on engine injector technology to operate over a wide range of liquid mass flow rates and chamber backpressures. One injection technology of current interest for variable thrust applications is an injector design with swirled fluids. Current swirl injector design methodologies do not take into account how swirl injector design parameters respond to elevated chamber backpressures at less than design mass flow rates. The current work was created to improve state-of-the-art swirl injector design methods in this area. The specific objective was to study the effects of elevated chamber backpressure and off-design mass flow rates on swirl injector fluid mechanics. Using a backpressure chamber with optical access, water was flowed through a swirl injector at various combinations of chamber backpressure and mass flow rates. The film thickness profile down the swirl injector nozzle section was measured through a transparent nozzle section of the injector. High speed video showed measurable increases in the film thickness profile with application of chamber backpressure and mass flow rates less than design. At prescribed combinations of chamber backpressure and injected mass flow rate, a discrete change in the film thickness profile was observed. Measured injector discharge coefficient values showed different trends with increasing chamber backpressure at low mass flow rates as opposed to near-design mass flow rates. Downstream spray angles showed classic changes in morphology as the mass flow rate was decreased below the design value. Increasing chamber backpressure decreased the spray angle at any injection mass flow rate. Experimental measurements and discussion of these results are reported in this paper.

Kenny, R Jeremy↗

NASA NTRS2009

Preliminary Axial Flow Turbine Design and Off-Design Performance Analysis Methods for Rotary Wing Aircraft Engines: Applications - Part 2

In this paper, preliminary studies on two turbine engine applications relevant to the tilt-rotor rotary wing aircraft are performed. The first case-study is the application of variable pitch turbine for the turbine performance improvement when operating at a substantially lower shaft speed. The calculations are made on the 75 percent speed and the 50 percent speed of operations. Our results indicate that with the use of the variable pitch turbines, a nominal (3 percent (probable) to 5 percent (hypothetical)) efficiency improvement at the 75 percent speed, and a notable (6 percent (probable) to 12 percent (hypothetical)) efficiency improvement at the 50 percent speed, without sacrificing the turbine power productions, are achievable if the technical difficulty of turning the turbine vanes and blades can be circumvented. The second casestudy is the contingency turbine power generation for the tilt-rotor aircraft in the One Engine Inoperative (OEI) scenario. For this study, calculations are performed on two promising methods: throttle push and steam injection. By isolating the power turbine and limiting its air mass flow rate to be no more than the air flow intake of the take-off operation, while increasing the turbine inlet total temperature (simulating the throttle push) or increasing the air-steam mixture flow rate (simulating the steam injection condition), our results show that an amount of 30 to 45 percent extra power, to the nominal take-off power, can be generated by either of the two methods. The methods of approach, the results, and discussions of these studies are presented in this paper.

Chen, Shu-cheng, S.↗

NASA NTRS1971

Mixing of hydrogen injected from multiple injectors normal to a supersonic airstream

Mixing of hydrogen downstream from multiple injectors normal to supersonic jet flow, for supersonic combustion on ramjet engines

Rogers, R. C.↗

NASA NTRS2019

NASA's Space Launch System: Payload Opportunities for Lunar Exploration, Science Missions

Nearing its first launch, the Space Launch System (SLS), NASA’s new super heavy-lift launch vehicle, offers highly energetic launches that deliver more mass and provide more volume in 8.4 m-diameter and potentially larger fairings to make a new generation of deep space missions possible. An evolvable launcher available in crew, crew with a co-manifested payload (CPL) and cargo-only configurations, SLS is a crucial capability to enable astronauts to return to the Moon, but it also offers key benefits for science missions. NASA’s 21st-century return to the Moon recently received a formal name: the Artemis program. In addition to the core enabling capabilities of SLS and the Orion crew spacecraft, NASA will also enlist international and commercial partnerships for Artemis. The Agency intends to build a scientific outpost in lunar orbit, the Gateway, from which human and robotic missions to and from the surface can rendezvous. SLS will launch Orion on a series of missions leading to landing the first woman and the next man on the Moon as part of Artemis. SLS uses proven propulsion systems: two solid rocket boosters and four RS-25s engines that have been upgraded to provide more thrust and operate in the SLS environment. SLS Block 1 uses a modified Delta IV Heavy upper stage, called the Interim Cryogenic Propulsion Stage (ICPS) and lifts at least 26 metric tons (t) to trans-lunar injection (TLI). The Block 1 vehicle can also be outfitted with a 5 m-diameter fairing. Block 1B, the next major variant, also uses solid rocket boosters and RS-25 engines to achieve Earth escape velocity, but replaces the single-engine liquid hydrogen (LH2)/liquid oxygen (LOX) ICPS with a four-engine LH2/LOX Exploration Upper Stage (EUS) to increase mass to TLI to 34-37 t, depending on crew or cargo configuration. In the Block 1B crew configuration, a 10 m-tall Universal Stage Adapter (USA) connects the vehicle to Orion and can carry a CPL up to 10 t. The USA provides 286 m3 of unpressurized volume for payloads. For large payloads, 8.4 m- and 10 m-diameter cargo fairings in 19.1 m and 27.4 m lengths are possible. The ultimate SLS vehicle, Block 2, incorporates evolved boosters to reach a lift capacity of more than 45 t to TLI. The capabilities of SLS not only make new missions to the Moon possible, but also game-changing science missions, such as deployment of large-aperture space telescopes, spacecraft to the ice giants or even probes to the interstellar medium. This paper will discuss the capabilities of SLS, the vehicle’s planned evolution, missions that can effectively utilize the vehicle and manufacturing status of the vehicle.

Creech, Stephen D.↗

NASA NTRS2019

NASA's Space Launch System: Payload Opportunities for Lunar Exploration, Science Missions

Nearing its first launch, the Space Launch System (SLS), NASA’s new super heavy-lift launch vehicle, offers highly energetic launches that deliver more mass and provide more volume in 8.4 m-diameter and potentially larger fairings to make a new generation of deep space missions possible. An evolvable launcher available in crew, crew with a co-manifested payload (CPL) and cargo-only configurations, SLS is a crucial capability to enable astronauts to return to the Moon, but it also offers key benefits for science missions. NASA’s 21st-century return to the Moon recently received a formal name: the Artemis program. In addition to the core enabling capabilities of SLS and the Orion crew spacecraft, NASA will also enlist international and commercial partnerships for Artemis. The Agency intends to build a scientific outpost in lunar orbit, the Gateway, from which human and robotic missions to and from the surface can rendezvous. SLS will launch Orion on a series of missions leading to landing the first woman and the next man on the Moon as part of Artemis. SLS uses proven propulsion systems: two solid rocket boosters and four RS-25s engines that have been upgraded to provide more thrust and operate in the SLS environment. SLS Block 1 uses a modified Delta IV Heavy upper stage, called the Interim Cryogenic Propulsion Stage (ICPS) and lifts at least 26 metric tons (t) to trans-lunar injection (TLI). The Block 1 vehicle can also be outfitted with a 5 m-diameter fairing. Block 1B, the next major variant, also uses solid rocket boosters and RS-25 engines to achieve Earth escape velocity, but replaces the single-engine liquid hydrogen (LH2)/liquid oxygen (LOX) ICPS with a four-engine LH2/LOX Exploration Upper Stage (EUS) to increase mass to TLI to 34-37 t, depending on crew or cargo configuration. In the Block 1B crew configuration, a 10 m-tall Universal Stage Adapter (USA) connects the vehicle to Orion and can carry a CPL up to 10 t. The USA provides 286 m3 of unpressurized volume for payloads. For large payloads, 8.4 m- and 10 m-diameter cargo fairings in 19.1 m and 27.4 m lengths are possible. The ultimate SLS vehicle, Block 2, incorporates evolved boosters to reach a lift capacity of more than 45 t to TLI. The capabilities of SLS not only make new missions to the Moon possible, but also game-changing science missions, such as deployment of large-aperture space telescopes, spacecraft to the ice giants or even probes to the interstellar medium. This paper will discuss the capabilities of SLS, the vehicle’s planned evolution, missions that can effectively utilize the vehicle and manufacturing status of the vehicle.

Creech, Steve↗

NASA NTRS1972

Experimental evaluation of a 600 lbf spacecraft rocket engine.

Experimental results are presented for a long-duration-capability (1000-sec), space-storable, bipropellant liquid rocket motor burning fluorine/hydrazine or FLOX/monomethylhydrazine. The interrelationship between injected mixture ratio and the per cent film cooling on vacuum specific impulse performance and chamber heat transfer is given. Experimental sea level measurements are used to predict space vacuum performance based upon simplified JANNAF reference procedures. Dynamic combustion stability is demonstrated over a wide range of operating conditions. Analytical results of char penetration, erosion, and ablative wall temperature distributions are presented for prototype chamber designs.

Hoehn, F. W.↗

NASA NTRS1987

Advanced Gas Turbine (AGT) Technology Development Project

This report is the eleventh in the series of Technical Summary reports for the Advanced Gas Turbine (AGT) Technology Development Project, authorized under NASA Contract DEN3-167, and sponsored by the Department of Energy (DOE). This report was prepared by Garrett Turbine Engine Company, A Division of the Garrett Corporation, and includes information provided by Ford Motor Company, the Standard Oil Company, and AiResearch Casting Company. This report covers plans and progress for the period July 1, 1985 through June 30, 1986. Technical progress during the reported period was highlighted by the 85-hour endurance run of an all-ceramic engine operating in the 2000 to 2250 F temperature regime. Component development continued in the areas of the combustion/fuel injection system, regenerator and seals system, and ceramic turbine rotor attachment design. Component rig testing saw further refinements. Ceramic materials showed continued improvements in required properties for gas turbine applications; however, continued development is needed before performance and reliability goals can be set.

Source record↗

NASA NTRS1989

A rocket engine design expert system

The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are described. The expert system incorporates a JANNAF standard reference computer code to determine rocket engine performance and a state-of-the-art finite element computer code to calculate the interactions between propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb heuristics were incorporated for the hydrogen-oxygen coaxial injector design, including a minimum gap size constraint on the total number of injector elements. One-dimensional equilibrium chemistry was employed in the energy release analysis of the combustion chamber and three-dimensional finite-difference analysis of the regenerative cooling channels was used to calculate the pressure drop along the channels and the coolant temperature as it exits the coolant circuit. Inputting values to describe the geometry and state properties of the entire system is done directly from the computer keyboard. Graphical display of all output results from the computer code analyses is facilitated by menu selection of up to five dependent variables per plot.

Davidian, Kenneth J.↗

NASA NTRS1989

Transient fault behavior in a microprocessor: A case study

An experimental analysis is described which studies the susceptibility of a microprocessor based jet engine controller to upsets caused by current and voltage transients. A design automation environment which allows the run time injection of transients and the tracing from their impact device to the pin level is described. The resulting error data are categorized by the charge levels of the injected transients by location and by their potential to cause logic upsets, latched errors, and pin errors. The results show a 3 picoCouloumb threshold, below which the transients have little impact. An Arithmetic and Logic Unit transient is most likely to result in logic upsets and pin errors (i.e., impact the external environment). The transients in the countdown unit are potentially serious since they can result in latched errors, thus causing latent faults. Suggestions to protect the processor against these errors, by incorporating internal error detection and transient suppression techniques, are also made.

Duba, Patrick↗

NASA NTRS1990

An experimental study of fault propagation in a jet-engine controller

An experimental analysis of the impact of transient faults on a microprocessor-based jet engine controller, used in the Boeing 747 and 757 aircrafts is described. A hierarchical simulation environment which allows the injection of transients during run-time and the tracing of their impact is described. Verification of the accuracy of this approach is also provided. A determination of the probability that a transient results in latch, pin or functional errors is made. Given a transient fault, there is approximately an 80 percent chance that there is no impact on the chip. An empirical model to depict the process of error exploration and degeneration in the target system is derived. The model shows that, if no latch errors occur within eight clock cycles, no significant damage is likely to happen. Thus, the overall impact of a transient is well contained. A state transition model is also derived from the measured data, to describe the error propagation characteristics within the chip, and to quantify the impact of transients on the external environment. The model is used to identify and isolate the critical fault propagation paths, the module most sensitive to fault propagation and the module with the highest potential of causing external pin errors.

Choi, Gwan Seung↗

NASA NTRS1948

Performance of J-33-A-21 and J-33-A-23 Compressors with and without Water Injection

In an investigation of the J-33-A-21 and the J-33-A-23 compressors with and without water injection, it was discovered that the compressors reacted differently to water injection although they were physically similar. An analysis of the effect of water injection on compressor performance and the consequent effect on matching of the compressor and turbine components in the turbojet engine was made. The analysis of component matching is based on a turbine flow function defined as the product of the equivalent weight flow and the reciprocal of the compressor pressure ratio.

Beede, William L.↗

NASA NTRSDate not supplied

Glenn Extreme Environments Rig (GEER) Facility Users Guide

The Glenn Extreme Environment Rig (GEER) is a large capacity pressure vessel capable of simulating high pressure and temperature conditions with a specified, custom mixture of gas. The chamber is certified for operation in physical and chemical conditions representative of the surface conditions of Venus. It can accommodate large scientific or engineering test articles while maintaining the specified conditions indefinitely. The user-configurable gas supply is capable of handling a mixture of up to nine independent streams for injection directly into the vessel. GEER is supported in part by NASA’s Planetary Science Division under the Science Mission Directorate and is available for use by the commercial, academic, and research communities. GEER was operationally commissioned in November of 2014. At that time, it successfully maintained the temperature, pressure, and chemistry of the Venus surface atmosphere. In the spring of 2015, GEER began operations for engineering and scientific research. Since then, over 20 tests have been successfully conducted in the chamber with a variety of atmospheric test conditions.

Blake W. Beckley↗

NASA NTRS1977

The alleged contributions of Pedro E. Paulet to liquid-propellant rocketry

The first practical working liquid propellant rocket motor was claimed by Pedro E. Paulet, a South American engineer from Peru (1895). He operated a conical motor, 10 centimeters in diameter, using nitrogen peroxide and gasoline as propellants and measuring thrust up to 90 kilograms, and apparently used spark ignition and intermittent propellant injection. The test device which he used contained elements of later test stands, such as a spring thrust-measuring device. However, he did not publish his work until twenty-five years later. Evidence is examined concerning this only known claim to liquid propellant rocket engine experiments in the nineteenth century.

Ordway, F. I., III↗