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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 1,027 records · Page 57
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Computational study of inlet injection for a Pre-Mixed, Shock-Induced Combustion (PM/SIC) engine

A computational simulation of reacting 2-D and 3-D flowfields in a model inlet section of a Pre-Mixed, Shock-Induced Combustion (PM/SIC) engine concept was performed. LARCK, a multi-dimensional Navier-Stokes code with finite-rate kinetics chemistry developed at NASA LaRC by J.A. White, was adapted for this simulation. The flow conditions in the simulation match those envisioned for the PM/SIC engine experiments currently planned at LaRC. The reacting flowfields were Mach 6.3 freestream air and Mach 2 hydrogen at various pressure and temperature conditions injected through a slot injector at the base of the inlet section. In the PM/SIC engine, fuel is injected at the inlet section upstream of the combustor, and reaction is initiated by the shock wave at the inlet which increases the gas temperature and pressure beyond the kinetic limits for reaction. Many challenges exist prior to establishing shock-controlled combustion as a practical engine concept. These challenges include fuel injection schemes that can provide proper fuel-air mixing without creating large losses in the inlet section, and control of the combustion process so that early ignition or combustion propagation through the inlet boundary layer does not occur. For this project, a parametrics study was carried out to model the fuel injection of hydrogen at different flow conditions. It was found that, as the fuel temperature and pressure were increased, the potential for pre-ignition was high at a short distance downstream of the slot injector. The next stage of this work will investigate injection techniques for enhancing mixing of fuel and air in a manner that prevents or reduces the potential for premature ignition observed numerically.

Gonzalez, D. E.↗
NASA NTRS1979

Evaluation of two inflow control devices for flight simulation of fan noise using a JT15D engine

The program was developed to accurately simulate flight fan noise on ground static test stands. The results generally indicated that both the induct and external ICD's were effective in reducing the inflow turbulence and the fan blade passing frequency tone generated by the turbulence. The external ICD was essentially transparent to the propagating fan tone but the induct ICD caused attenuation under most conditions.

Jones, W. L.↗
NASA NTRS2001

Computational Analysis of a Pylon-Chevron Core Nozzle Interaction

In typical engine installations, the pylon of an engine creates a flow disturbance that interacts with the engine exhaust flow. This interaction of the pylon with the exhaust flow from a dual stream nozzle was studied computationally. The dual stream nozzle simulates an engine with a bypass ratio of five. A total of five configurations were simulated all at the take-off operating point. All computations were performed using the structured PAB3D code which solves the steady, compressible, Reynolds-averaged Navier-Stokes equations. These configurations included a core nozzle with eight chevron noise reduction devices built into the nozzle trailing edge. Baseline cases had no chevron devices and were run with a pylon and without a pylon. Cases with the chevron were also studied with and without the pylon. Another case was run with the chevron rotated relative to the pylon. The fan nozzle did not have chevron devices attached. Solutions showed that the effect of the pylon is to distort the round Jet plume and to destroy the symmetrical lobed pattern created by the core chevrons. Several overall flow field quantities were calculated that might be used in extensions of this work to find flow field parameters that correlate with changes in noise.

Thomas, Russell H.↗
NASA NTRS1987

Impeller shroud to casing leakage flow simulations in the Space Shuttle Main Engine high pressure fuel pump

Quasi-three-dimensional Navier-Stokes calculations were carried out for the Space Shuttle Main Engine high-pressure fuel pump to simulate the impeller shroud to casing leakage flow. This flow geometry was modeled as an axisymmetric cavity flow with a stationary surface representing the casing, and a rotating surface denoting the impeller. A 63 x 81-node mesh provided sufficient resolution in the regions of greatest flow variations and reduced the effects of numerical diffusion. The turbulence field was closed with the high Reynolds number form of the k-epsilon model supplemented with wall functions in the vicinity of the walls. Finally, a parametric study quantified the effects of through mass flow changes on this leakage flow.

Sindir, Munir M.↗
NASA NTRS2021

NASA Engineering and Safety Center Technical Bulletin No. 21-02: Genesis Flight Mechanics Simulation

Genesis Flight Mechanics Simulation The NASA Engineering and Safety Center (NESC) consolidated and modernized a suite of legacy flight mechanics simulations, including the Flight Analysis and Simulation Tool (FAST), resulting in Genesis, a generic, multi-vehicle, variable-degree-of-freedom flight mechanics simulation for ascent, aerocapture, entry, descent, and landing (A2EDL) trajectory design. Genesis is more flexible, capable, and performant than FAST. It enables trajectory optimization and interactive trajectory generation. Its interoperability with Copernicus, an exo-atmospheric and interplanetary trajectory design tool, facilitates end-to-end trajectory optimization across all mission phases. Genesis is implemented in Julia, a new language for technical computing that combines the ease of use of scripting languages with the run-time performance of compiled languages.

EDL↗
NASA NTRS1997

GSFC Space Simulation Laboratory Contamination Philosophy: Efficient Space Simulation Chamber Cleaning Techniques

This paper will provide a general overview of the molecular contamination philosophy of the Space Simulation Test Engineering Section and how the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) space simulation laboratory controls and maintains the cleanliness of all its facilities, thereby, minimizing down time between tests. It will also briefly cover the proper selection and safety precautions needed when using some chemical solvents for wiping, washing, or spraying thermal shrouds when molecular contaminants increase to unacceptable background levels.

Roman, Juan A.↗
NASA NTRS1959

A Combined Water-Bromotrifluoromethane Crash-Fire Protection System for a T-56 Turbopropeller Engine

A crash-fire protection system is described which will suppress the ignition of crash-spilled fuel that may be ingested by a T-56 turbo-propeller engine. This system includes means for rapidly extinguishing the combustor flame, means for cooling and inerting with water the hot engine parts likely to ignite engine ingested fuel, and means for blanketing with bromotrifluoromethane massive metal parts that may reheat after the engine stops rotating. Combustion-chamber flames were rapidly extinguished at the engine fuel nozzles by a fuel shutoff and drain valve. Hot engine parts were inerted and cooled by 42 pounds of water discharged at seven engine stations. Massive metal parts that could reheat were inerted with 10 pounds of bromotrifluoromethane discharged at two engine stations. Performance trials of the crash-fire protection system were conducted by bringing the engine up to takeoff temperature, actuating the crash-fire protection system, and then spraying fuel into the engine to simulate crash-ingested fuel. No fires occurred during these trials, although fuel was sprayed into the engine from 0.3 second to 15 minutes after actuating the crash-fire protection system.

Campbell, John A.↗
NASA NTRS1959

Crash-Fire Protection System for T-56 Turbopropeller Engine Using Water as Cooling and Inerting Agent

A crash-fire protection system to suppress the ignition of crash-spilled fuel that may be ingested by a T-56 turbopropeller engine is described. This system includes means for rapidly extinguishing the combustor flame and means for cooling and inerting with water the hot engine parts likely to ignite engine-ingested fuel. Combustion-chamber flames were extinguished in 0.07 second at the engine fuel manifold. Hot engine parts were inerted and cooled by 52 pounds of water discharged at ten engine stations. Performance trials of the crash-fire prevention system were conducted by bringing the engine up to takeoff temperature, stopping the normal fuel flow to the engine, starting the water discharge, and then spraying fuel into the engine to simulate crash-ingested fuel. No fires occurred during these trials, although fuel was sprayed into the engine from 0.3 second to 15 minutes after actuating the crash-fire protection system.

Busch, Arthur M.↗
NASA NTRS1987

Performance limitations in parallel processor simulations

A jet-engine model is partitioned and simulated on a parallel processor system consisting of five 8086/8087 floating-point computers. The simulation uses Heun's integration method. A near-optimal parallel simulation (in the sense of minimum execution time) achieves speedup of only 2.13 and efficiency of 42.6 percent, in effect wasting 57.4 percent of the available processing power. A detailed analysis identifies and graphically demonstrates why the system fails to achieve ideal performance (viz., speedup of 5 and efficiency of 100 percent). Inherent characteristics of the problem equations and solution algorithm account for the loss of nearly half of the available processing power. Overheads associated with interprocessor communication and processor synchronization account for only a small fraction of the lost processing power. The effects of these and other factors which limit parallel processor performance are illustrated through real-time timing-analyzer tracers describing the run/idle status of the parallel processors during the simulation.

O'Grady, E. Pearse↗
NASA NTRS1996

A numerical investigation of premixed combustion in wave rotors

Wave rotor cycles which utilize premixed combustion processes within the passages are examined numerically using a one-dimensional CFD-based simulation. Internal-combustion wave rotors are envisioned for use as pressure-gain combustors in gas turbine engines. The simulation methodology is described, including a presentation of the assumed governing equations for the flow and reaction in the channels, the numerical integration method used, and the modeling of external components such as recirculation ducts. A number of cycle simulations are then presented which illustrate both turbulent-deflagration and detonation modes of combustion. Estimates of performance and rotor wall temperatures for the various cycles are made, and the advantages and disadvantages of each are discussed.

Nalim, M. Razi↗
NASA NTRS2006

Flight-Tested Prototype of BEAM Software

Researchers at JPL have completed a software prototype of BEAM (Beacon-based Exception Analysis for Multi-missions) and successfully tested its operation in flight onboard a NASA research aircraft. BEAM (see NASA Tech Briefs, Vol. 26, No. 9; and Vol. 27, No. 3) is an ISHM (Integrated Systems Health Management) technology that automatically analyzes sensor data and classifies system behavior as either nominal or anomalous, and further characterizes anomalies according to strength, duration, and affected signals. BEAM (see figure) can be used to monitor a wide variety of physical systems and sensor types in real time. In this series of tests, BEAM monitored the engines of a Dryden Flight Research Center F-18 aircraft, and performed onboard, unattended analysis of 26 engine sensors from engine startup to shutdown. The BEAM algorithm can detect anomalies based solely on the sensor data, which includes but is not limited to sensor failure, performance degradation, incorrect operation such as unplanned engine shutdown or flameout in this example, and major system faults. BEAM was tested on an F-18 simulator, static engine tests, and 25 individual flights totaling approximately 60 hours of flight time. During these tests, BEAM successfully identified planned anomalies (in-flight shutdowns of one engine) as well as minor unplanned anomalies (e.g., transient oil- and fuel-pressure drops), with no false alarms or suspected false-negative results for the period tested. BEAM also detected previously unknown behavior in the F- 18 compressor section during several flights. This result, confirmed by direct analysis of the raw data, serves as a significant test of BEAM's capability.

Mackey, Ryan↗
NASA NTRS1997

Flutter Boundary Identification From Simulation Time Histories

While there has been much recent progress in simulating nonlinear aeroelastic systems, and in predicting many of the aeroelastic phenomena of concern in transport aircraft design (i.e. transonic flutter buckets), the utility of a simulation in generating an understanding of the flutter behavior is limited. This is due in part to the high cost of generating these simulations; and the implied limitation on the number of conditions that can be analyzed, but there are also some difficulties introduced by the very nature of a simulation. Flutter engineers have traditionally worked in the frequency domain, and are accustomed to describing the flutter behavior of an airplane in terms of its V-G and V-F (or Q-G and Q-F) plots and flutter mode shapes. While the V-G and V-F plots give information about how the dynamic response of an airplane changes as the airspeed is increased, the simulation only gives information about one isolated condition (Mach, airspeed, altitude, etc.). Therefore, where a traditional flutter analysis can let the engineer determine an airspeed at which an airplane becomes unstable, while a simulation only serves as a binary check: either the airplane is fluttering at this condition, or it is not. In this document, a new technique is described in which system identification is used to easily extract modal frequencies and damping ratios from simulation time histories, and shows how the identified parameters can be used to determine the variation in frequency and dampin,o ratio as the airspeed is changed. This technique not only provides the flutter engineer with added insight into the aeroelastic behavior of the airplane, but it allows calculation of flutter mode shapes, and allows estimation of flutter boundaries while minimizing the number of simulations required.

Baker, Myles↗
NASA NTRS1989

Real-time simulation of an F110/STOVL turbofan engine

A traditional F110-type turbofan engine model was extended to include a ventral nozzle and two thrust-augmenting ejectors for Short Take-Off Vertical Landing (STOVL) aircraft applications. Development of the real-time F110/STOVL simulation required special attention to the modeling approach to component performance maps, the low pressure turbine exit mixing region, and the tailpipe dynamic approximation. Simulation validation derives by comparing output from the ADSIM simulation with the output for a validated F110/STOVL General Electric Aircraft Engines FORTRAN deck. General Electric substantiated basic engine component characteristics through factory testing and full scale ejector data.

Drummond, Colin K.↗
NASA NTRS1957

Preliminary investigation of the combustion of a 50 percent pentaborane - 50 percent JP-4 fuel blend in a turbojet combustor at simulated altitude conditions

A preliminary investigation was conducted to determine the combustion characteristics of a fuel composed of 50 percent pentaborane and 50 percent JP-4 (MIL-F-5624A) by weight in a turbojet combustor. A combustor designed to fit the housing of a J33-A-23 turbojet engine was selected for convenience. The fuel was evaluated at two engine conditions simulating altitudes of 40,000 and 57,000 feet, an engine speed of 85 percent of rated rpm, and a flight Mach number of 0.6. The pentaborane blend was initially evaluated in combustors developed for pure pentaborane and diborane reported in NACA RM E53B18 and RM E52L15. The performance of the blend was unsatisfactory in these combustors. A new combustor was then developed which provided combustor efficiencies measured from 91 to 101 percent as compared with efficiencies of 92 to 94 percent previously obtained for pentaborane at comparable conditions. Additional refinements of design details are needed to obtain lower oxide deposits and a more uniform outlet temperature profile; however, the combustor is believed to incorporate some of the design principles required to obtain satisfactory over-all performance with the fuel blend investigated.

FUELS - JP-4-*PENTABORANE↗
NASA NTRS1932

The Effect of Valve Timing Upon the Performance of a Supercharged Engine at Altitude and an Unsupercharged Engine at Sea Level

This investigation was conducted to determine the comparative effects of valve timing on the performance of an unsupercharged engine at sea level and a supercharged engine at altitude. The tests were conducted on the NACA universal test engine. The timing of the four valve events was varied over a wide range; the engine speeds were varied between 1,050 and 1,500 r.p.m.; the compression ratios were varied between 4.35:1 and 7.35:1. The conditions of exhaust pressure and carburetor pressure of a supercharged engine were simulated for altitudes between 0 and 18,000 feet. The results show that optimum valve timing for a supercharged engine at an altitude of 18,000 feet differs slightly from that for an unsupercharged engine at sea level. A small increase in power is obtained by using the optimum timing for 18,000 feet for altitudes above 5,000 feet. The timing of the intake opening and exhaust closing becomes more critical as the compression ratio is increased.

Schey, Oscar W↗
NASA NTRSDate not supplied

Advanced Materials Development under NASA’s Hybrid Thermally Efficient Core (HyTEC) Project

The Hybrid Thermally Efficient Core (HyTEC) project aims to develop small core turbofan engine technologies that will enable fuel burn reductions, additional use of electric airplane systems through power extraction, and to advance engine operability and compatibility with sustainable aviation fuels. As such, a portfolio of technologies that contribute to raising the pressure, temperature and efficiency of turbofan engine cores needs developed. Included in HyTEC’s technology portfolio to enable these advancements are enhanced combustor liner materials and higher temperature capable turbine blades and vanes manufactured using ceramic matrix composites (CMCs) and environmental barrier coatings (EBCs). These technologies are being developed and tested in laboratory-scale relevant environments to advance the technology readiness level (TRL) to 4 or 5, before moving into an engine core demonstrator to raise the TRL to 6. A new natural gas/oxygen burner rig facility will be used to simulate turbine engine relevant environments at the laboratory/coupon scale. In addition, improvements to enable sub-element, and more complex component testing in NASA Glenn’s combustor facility (CE-5) will be presented.

Ceramic Matrix Composites (CMCs)↗
NASA NTRS2011

Flow Matching Results of an MHD Energy Bypass System on a Supersonic Turbojet Engine Using the Numerical Propulsion System Simulation (NPSS) Environment

Flow matching has been successfully achieved for an MHD energy bypass system on a supersonic turbojet engine. The Numerical Propulsion System Simulation (NPSS) environment helped perform a thermodynamic cycle analysis to properly match the flows from an inlet employing a MHD energy bypass system (consisting of an MHD generator and MHD accelerator) on a supersonic turbojet engine. Working with various operating conditions (such as the applied magnetic field, MHD generator length and flow conductivity), interfacing studies were conducted between the MHD generator, the turbojet engine, and the MHD accelerator. This paper briefly describes the NPSS environment used in this analysis. This paper further describes the analysis of a supersonic turbojet engine with an MHD generator/accelerator energy bypass system. Results from this study have shown that using MHD energy bypass in the flow path of a supersonic turbojet engine increases the useful Mach number operating range from 0 to 3.0 Mach (not using MHD) to a range of 0 to 7.0 Mach with specific net thrust range of 740 N-s/kg (at ambient Mach = 3.25) to 70 N-s/kg (at ambient Mach = 7). These results were achieved with an applied magnetic field of 2.5 Tesla and conductivity levels in a range from 2 mhos/m (ambient Mach = 7) to 5.5 mhos/m (ambient Mach = 3.5) for an MHD generator length of 3 m.

Benyo, Theresa L.↗
NASA NTRS2011

The Challenge of Configuring Model-Based Space Mission Planners

Mission planning is central to space mission operations, and has benefited from advances in model-based planning software. Constraints arise from many sources, including simulators and engineering specification documents, and ensuring that constraints are correctly represented in the planner is a challenge. As mission constraints evolve, planning domain modelers need help with modeling constraints efficiently using the available source data, catching errors quickly, and correcting the model. This paper describes the current state of the practice in designing model-based mission planning tools, the challenges facing model developers, and a proposed Interactive Model Development Environment (IMDE) to configure mission planning systems. We describe current and future technology developments that can be integrated into an IMDE.

model-based systems↗