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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Application of CFD to Explain Anomalous Stall Behavior of the SSME Flowmeter

The Space Shuttle Main Engine (SSME) Fuel Flowmeter is located in the duct between the low and high pressure fuel turbopumps. In the flowmeter the rotation rate of a 4-blade rotor positioned downstream of two flow straighteners is employed to measure the engine fuel flow rate and thereby control the engine mixture ratio via the engine controller. Hence, inaccurate operation of the flowmeter could have serious consequences for SSME engine operation and performance, forcing, for example, tanking of extra fuel to allow for inaccurate flowmeter measurement. Since the current flight flowmeter configuration was incorporated into the SSME in the early eighties, some anomalies in flowmeter behavior have been observed. The initial flowmeter incorporated an "egg crate" design for the two flow straighteners which turn the duct flow to make it more uniform and parallel after it has come out of the 90 degree bend just upstream of the flowmeter.

Ascoli, E.↗
NASA NTRS2022

An Analysis of the Impact of Extreme Attitude Operation on a Turbofan Engine in a Regional Jet Aircraft

The objective of this work is to characterize the variability in engine performance at extreme attitudes, resulting from the use of different standard control variables, evaluated at different points in the engine’s lifecycle. This paper describes the procedure used to create a dynamic model of an aft-mounted jet engine on a T-tail regional jet aircraft. This model enables simulation of engine operation at extreme attitudes. The model was subsequently evaluated at various altitudes, Mach numbers, and power settings over a range of angles of attack and sideslip, and deterioration levels. Each case was simulated using both fan speed and engine pressure ratio as the engine control variable. The results show that the wing has a very large impact on the engine operation, overwhelming other sources of variation.

loss of control - inflight↗
NASA NTRS1979

Orbital transfer vehicle advanced expander cycle engine point design study

A detailed computer model of the expander cycle engine system steady-state operation was developed. Payload performance data for the expander cycle engine for 100K pounds and 65K pounds Space Shuttle missions were examined to establish peak payload thrusts. It was found that peak payload for the 100K pound gross weight single engine case occurred at nearly 15,000 pounds thrust for both engine mixture ratios of 6:1 and 7:1. Peak payload at engine mixture ratio of 6:1 was 4 percent greater than at 7:1 mixture ratio. The two engine configuration peak paYloads occurred at engine thrust below 7,500 pounds (extrapolated) and were lower than single engine payloads. For the 65K pound gross weight mission, peak payloads for the single engine configuration occurred at nearly 10,000 pounds for both mixture ratios. Mixture ratio of 7:1 payloads were 5 percent lower than 6:1 payloads. Two engine configuration payloads occurred at engine thrusts considerably below 10K pounds.

Martinez, A.↗
NASA NTRS1977

Effect of forward motion on engine noise

Methods used to determine a procedure for correcting static engine data for the effects of forward motion are described. Data were analyzed from airplane flyover and static-engine tests with a JT8D-109 low-bypass-ratio turbofan engine installed on a DC-9-30, with a CF6-6D high-bypass-ratio turbofan engine installed on a DC-10-10, and with a JT9D-59A high-bypass-ratio turbofan engine installed on a DC-10-40. The observed differences between the static and the flyover data bases are discussed in terms of noise generation, convective amplification, atmospheric propagation, and engine installation. The results indicate that each noise source must be adjusted separately for forward-motion and installation effects and then projected to flight conditions as a function of source-path angle, directivity angle, and acoustic range relative to the microphones on the ground.

Blankenship, G. L.↗
NASA NTRS1989

Effect of heavy rain on aviation engines

High bypass ratio gas turbine engines may ingest water and hail during flight in an environment of thunderstorms, and the performance and the operation-handling characteristics of the engine and its control become affected often substantially and critically. It is, therefore, of interest to establish predictive schemes for determining changes in performance of components and the total system. The current status of development of such predictive schemes is discussed along with illustrative examples. The needs for additional research are discussed, that are essential for improving: (1) the predictive schemes and (2) the methods of simulating in-rain flight conditions on ground.

Murthy, S. N. B.↗
NASA NTRS1992

A neural network-based estimator for the mixture ratio of the Space Shuttle Main Engine

In order to properly utilize the available fuel and oxidizer of a liquid propellant rocket engine, the mixture ratio is closed loop controlled during main stage (65 percent - 109 percent power) operation. However, because of the lack of flight-capable instrumentation for measuring mixture ratio, the value of mixture ratio in the control loop is estimated using available sensor measurements such as the combustion chamber pressure and the volumetric flow, and the temperature and pressure at the exit duct on the low pressure fuel pump. This estimation scheme has two limitations. First, the estimation formula is based on an empirical curve fitting which is accurate only within a narrow operating range. Second, the mixture ratio estimate relies on a few sensor measurements and loss of any of these measurements will make the estimate invalid. In this paper, we propose a neural network-based estimator for the mixture ratio of the Space Shuttle Main Engine. The estimator is an extension of a previously developed neural network based sensor failure detection and recovery algorithm (sensor validation). This neural network uses an auto associative structure which utilizes the redundant information of dissimilar sensors to detect inconsistent measurements. Two approaches have been identified for synthesizing mixture ratio from measurement data using a neural network. The first approach uses an auto associative neural network for sensor validation which is modified to include the mixture ratio as an additional output. The second uses a new network for the mixture ratio estimation in addition to the sensor validation network. Although mixture ratio is not directly measured in flight, it is generally available in simulation and in test bed firing data from facility measurements of fuel and oxidizer volumetric flows. The pros and cons of these two approaches will be discussed in terms of robustness to sensor failures and accuracy of the estimate during typical transients using simulation data.

Guo, T. H.↗
NASA NTRS1978

Preliminary study of optimum ductburning turbofan engine cycle design parameters for supersonic cruising

The effect of turbofan engine overall pressure ratio, fan pressure ratio, and ductburner temperature rise on the engine weight and cruise fuel consumption for a mach 2.4 supersonic transport was investigated. Design point engines, optimized purely for the supersonic cruising portion of the flight where the bulk of the fuel is consumed, are considered. Based on constant thrust requirements at cruise, fuel consumption considerations would favor medium by pass ratio engines (1.5 to 1.8) of overall pressure ratio of about 16. Engine weight considerations favor low bypass ratio (0.6 or less) and low wverall pressure ratio (8). Combination of both effects results in bypass ratios of 0.6 to 0.8 and overall pressure ratio of 12 being the overall optimum.

Fishbach, L. H.↗
NASA NTRS1993

The prediction of nozzle performance and heat transfer in hydrogen/oxygen rocket engines with transpiration cooling, film cooling, and high area ratios

An advanced engineering computational model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multi-species, chemically reacting and diffusing Navier-Stokes equations are modelled, finite difference approach that is tailored to be conservative in an axisymmetric coordinate system for both the inviscid and viscous terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and transpiration cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent plug-and-spool rocket engine analysis cases performed. Further, the Soret term was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration cooled rocket engine.

Kacynski, Kenneth J.↗
NASA NTRS1994

The Prediction of Nozzle Performance and Heat Transfer in Hydrogen/Oxygen Rocket Engines with Transpiration Cooling, Film Cooling, and High Area Ratios

An advanced engineering computational model has been developed to aid in the analysis of chemical rocket engines. The complete multispecies, chemically reacting and diffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and Dufour energy transfer terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film-cooled nozzle, and a transpiration-cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film-cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. It is demonstrated that thermal diffusion has a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle and was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration-cooled rocket engine.

Kacynski, Kenneth J.↗
NASA NTRS1992

Application of Taguchi methods to dual mixture ratio propulsion system optimization for SSTO vehicles

The application of advanced technologies to future launch vehicle designs would allow the introduction of a rocket-powered, single-stage-to-orbit (SSTO) launch system early in the next century. For a selected SSTO concept, a dual mixture ratio, staged combustion cycle engine that employs a number of innovative technologies was selected as the baseline propulsion system. A series of parametric trade studies are presented to optimize both a dual mixture ratio engine and a single mixture ratio engine of similar design and technology level. The effect of varying lift-off thrust-to-weight ratio, engine mode transition Mach number, mixture ratios, area ratios, and chamber pressure values on overall vehicle weight is examined. The sensitivity of the advanced SSTO vehicle to variations in each of these parameters is presented, taking into account the interaction of each of the parameters with each other. This parametric optimization and sensitivity study employs a Taguchi design method. The Taguchi method is an efficient approach for determining near-optimum design parameters using orthogonal matrices from design of experiments (DOE) theory. Using orthogonal matrices significantly reduces the number of experimental configurations to be studied. The effectiveness and limitations of the Taguchi method for propulsion/vehicle optimization studies as compared to traditional single-variable parametric trade studies is also discussed.

Stanley, Douglas O.↗
NASA NTRS1959

Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48

The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

Fortini, Anthony↗
NASA NTRS2009

Analysis of Turbofan Design Options for an Advanced Single-Aisle Transport Aircraft

The desire for higher engine efficiency has resulted in the evolution of aircraft gas turbine engines from turbojets, to low bypass ratio, first generation turbofans, to today's high bypass ratio turbofans. It is possible that future designs will continue this trend, leading to very-high or ultra-high bypass ratio (UHB) engines. Although increased bypass ratio has clear benefits in terms of propulsion system metrics such as specific fuel consumption, these benefits may not translate into aircraft system level benefits due to integration penalties. In this study, the design trade space for advanced turbofan engines applied to a single-aisle transport (737/A320 class aircraft) is explored. The benefits of increased bypass ratio and associated enabling technologies such as geared fan drive are found to depend on the primary metrics of interest. For example, bypass ratios at which fuel consumption is minimized may not require geared fan technology. However, geared fan drive does enable higher bypass ratio designs which result in lower noise. Regardless of the engine architecture chosen, the results of this study indicate the potential for the advanced aircraft to realize substantial improvements in fuel efficiency, emissions, and noise compared to the current vehicles in this size class.

Guynn, Mark D.↗
NASA NTRS1992

The future challenge for aeropropulsion

NASA's research in aeropropulsion is focused on improving the efficiency, capability, and environmental compatibility for all classes of future aircraft. The development of innovative concepts, and theoretical, experimental, and computational tools provide the knowledge base for continued propulsion system advances. Key enabling technologies include advances in internal fluid mechanics, structures, light-weight high-strength composite materials, and advanced sensors and controls. Recent emphasis has been on the development of advanced computational tools in internal fluid mechanics, structural mechanics, reacting flows, and computational chemistry. For subsonic transport applications, very high bypass ratio turbofans with increased engine pressure ratio are being investigated to increase fuel efficiency and reduce airport noise levels. In a joint supersonic cruise propulsion program with industry, the critical environmental concerns of emissions and community noise are being addressed. NASA is also providing key technologies for the National Aerospaceplane, and is studying propulsion systems that provide the capability for aircraft to accelerate to and cruise in the Mach 4-6 speed range. The combination of fundamental, component, and focused technology development underway at NASA will make possible dramatic advances in aeropropulsion efficiency and environmental compatibility for future aeronautical vehicles.

Rosen, Robert↗
NASA NTRS1988

A 1050 K Stirling space engine design

As part of the NASA CSTI High Capacity Power Program on Conversion Systems for Nuclear Applications, Sunpower, Inc. completed for NASA Lewis a reference design of a single-cylinder free-piston Stirling engine that is optimized for the lifetimes and temperatures appropriate for space applications. The NASA effort is part of the overall SP-100 program which is a combined DOD/DOE/NASA project to develop nuclear power for space. Stirling engines have been identified as a growth option for SP-100 offering increased power output and lower system mass and radiator area. Superalloy materials are used in the 1050 K hot end of the engine; the engine temperature ratio is 2.0. The engine design features simplified heat exchangers with heat input by sodium heat pipes, hydrodynamic gas bearings, a permanent magnet linear alternator, and a dynamic balance system. The design shows an efficiency (including the alternator) of 29 percent and a specific mass of 5.7 kg/kW. This design also represents a significant step toward the 1300 K refractory Stirling engine which is another growth option of SP-100.

Penswick, L. Barry↗
NASA NTRS1992

SSME Automated Engine Calibration System (AECS)

An algorithm is derived for the real-time calibration of the engine fuel flowmeter and the engine mixture ratio during Space Shuttle Main Engine (SSME) ground testing. Because currently used calibration methods are post-test operations, there exists no fail-safe way of predicting at what mixture ratio a planned test will run. It is proposed that the algorithm developed here be used as part of an Automated Engine Calibration System (AECS) which could ensure that nearly all SSME tests are run at the proper mixture ratio. In this way, AECS has the potential of increasing the efficiency of the SSME ground test program. In addition to the derivation of the algorithm, an overview of this calibration system is presented along with the list of test stand facility instrumentation necessary for AECS implementation.

Greene, William D.↗
NASA NTRS1972

Tabulated pressure measurements on a large subsonic transport model airplane with high bypass ratio, powered, fan jet engines

An experimental wind-tunnel investigation to determine the aerodynamic interference and the jet-wake interference associated with the wing, pylon, and high-bypass-ratio, powered, fan-jet model engines has been conducted on a typical high-wing logistics transport airplane configuration. Pressures were measured on the wing and pylons and on the surfaces of the engine fan cowl, turbine cowl, and plug. Combinations of wing, pylons, engines, and flow-through nacelles were tested, and the pressure coefficients are presented in tabular form. Tests were conducted at Mach numbers from 0.700 to 0.825 and angles of attack from -2 to 4 deg.

Flechner, S. G.↗
NASA NTRS1998

Incompressible Turbulent Wing-Body Junction Flow

The overall objective of this study is to contribute to the optimized design of fan bypass systems in advanced turbofan engines. Increasing the engine bypass ratios have provided a major boost in engine performance improvement over the last fifty years. An engine with high bypass ratio (11-16:1) such as the Advanced Ducted Propulsion (ADP) is being developed and is expected to provide an additional 25% improvement in overall efficiency over the early turbofans. Such significant improvements in overall efficiency would reduce the cost per seat mile, which is a major government and Industry challenge for the 21th century. The research is part of the Advanced Subsonic Technology (AST) program that involves a NASA, U.S. Industry and FAA partnership with the goal of a safe and highly productive global air transportation system. The immediate objective of the study is to perform numerical simulation of duct-strut interactions to elucidate the loss mechanisms associated with this configuration that is typical of advanced turbofan engines such as ADP. However, at present experimental data for a duct-strut configuration are not available. Thus, as a first step a wing-body junction flow would be studied and is the specific objective of the present study. At the outset it is to be recognized that while duct-strut interaction flow is similar to that of wing-body junction flows, there are some differences owing to the presence of a wall at both ends of the strut. Likewise, some differences are due to the sheared inflow (as opposed to a uniform inflow) velocity profile. It is however expected that some features of a wing-body junction flow would persist. Next, some of the salient aspects of the complex flow near a wing-body junction, as revealed by various studies reported in the literature will be reviewed. One of the principle characteristics of the juncture flow, is the presence of the mean flow components in a plane perpendicular to the direction of the oncoming free-stream flow. The lateral curvature of the wing/strat causes the oncoming turbulent layer to skew about am axis (x-axis) parallel to the plane (xz-plane) of the mean shear. This is the principle mechanism for the generation of secondary flow. Such skew-induced secondary flows are slow to be attenuated by Reynolds stresses. Additional contribution to the generation of secondary flow comes from anisotropies in Reynolds stresses. Upstream of the strut, the mean-vorticity is directed span wise (along the y-direction). The presence of secondary flow in the vicinity of the strut causes the vorticity to stretch around the obstacle in a horse-shoe shape, with each leg having a vorticity of the opposite sense. The blockage effect of the strut imposes a severe adverse pressure gradient on the oncoming turbulent shear layer, causing boundary layer separation ahead of the leading edge, resulting in a vortex that rolls up and flows downstream into the juncture region. The separation vortices trailing in the wake of the wing can alter the lift or drag characteristics of the surfaces downstream of the wing-body juncture. Likewise, on submarines, the wake flow behind the appendage can degrade the performance of the propeller located downstream. The complex nature of this flow is caused by the presence of all six components of Reynolds stresses. Devenport and Simpson report that in the vicinity of the horse-shoe vortex there is intense recirculation with turbulent stresses being much larger than those normally observed in turbulent flows. These features contribute to making this flow a challenge to predict numerically. Some of the past studies provide useful insights into this flow that would guide our numerical efforts. In measurements reported by Shabaka and Bradshaw, the eddy viscosity tensor is seen to be non-isotropic and has negative components in certain regions. In an effort to evaluate the closure assumptions of various turbulence models, Devenport and Simpson used their own extensive measurements in juncture flows around the nose of a wing-body junction. Measured values of mean-velocity and/or turbulence kinetic energy was used to predict the magnitude of the shear stress vector. Algebraic stress models performed the best followed by Cebeci-Smith eddy viscosity model. The flow is reported to be dominated by a pressure field produced by the wing and the velocity field generated by the horseshoe vortex that is wrapped around the junction between the wing and wall. Kubendran et al. conclude from an experimental study that the shape of leading edge of the wing as characterized by its slenderness ratio is a major factor in determining the flow fields in the juncture region. The more thinner the leading edge of the juncture, the weaker the horseshoe vortex is. Also, with a slender leading edge, the secondary flow in the juncture would be due mainly to the cross-stream gradients of Reynolds stresses rather than due to a lateral skewing of the shear layer.

Krishnamurthy, R.↗