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At least 739 records · Page 41

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NASA NTRS1989

Spectroscopic study of combustion diagnostics on hydroxyl radicals

Experimental observations of propane/air flames have been carried out. Measurements of hydroxyl (OH) radical concentration have been made using resonance line absorption techniques. A microwave-pumped low pressure discharge in argon and water vapor is employed to produce strong OH radical band radiation in the 308 nm region. This radiation is transmitted through the plume and absorption data is taken at various radial positions using an optical multichannel analyzer. This absorption data is used to compute OH number density using a model for the absorption band characteristics as a function of temperature based on an atlas of line strengths. A numerical computation of flow fields, temperature profile, OH number density and water vapor mass fraction profiles was carried out by using a technique of computational fluid dynamics (CFD). The results of CFD computation are compared with experimental observation with a good agreement.

Eskridge, Richard↗

NASA NTRS2008

An Idealized, Single Radial Swirler, Lean-Direct-Injection (LDI) Concept Meshing Script

To easily study combustor design parameters using computational fluid dynamics codes (CFD), a Gridgen Glyph-based macro (based on the Tcl scripting language) dubbed BladeMaker has been developed for the meshing of an idealized, single radial swirler, lean-direct-injection (LDI) combustor. BladeMaker is capable of taking in a number of parameters, such as blade width, blade tilt with respect to the perpendicular, swirler cup radius, and grid densities, and producing a three-dimensional meshed radial swirler with a can-annular (canned) combustor. This complex script produces a data format suitable for but not specific to the National Combustion Code (NCC), a state-of-the-art CFD code developed for reacting flow processes.

Iannetti, Anthony C.↗

NASA NTRS2017

Structural Design Exploration of an Electric Powered Multi-Propulsor Wing Configuration

Advancements in aircraft electric propulsion may enable an expanded operational envelope for electrically powered vehicles compared to their internal combustion engine counterparts. High aspect ratio wings provide additional lift and drag reduction for a proposed multi-propulsor design, however, the challenge is to reduce the weight of wing structures while maintaining adequate structural and aeroelastic margins. Design exploration using a conventional design-and-build philosophy coupled with a finite element method (FEM)-based design of experiments (DOE) strategy are presented to examine high aspect ratio wing structures that have spanwise distributed electric motors. Multiple leading-edge-mounted engine masses presented a challenge to design a wing within acceptable limits for dynamic and aeroelastic stability. Because the first four primary bending eigenmodes of the proposed wing structure are very sensitive to outboard motor placement, safety-of-flight requirements drove the need for multiple spars, rib attachments, and outboard structural reinforcements in the design. Global aeroelasticity became an increasingly important design constraint during the on-going design process, with outboard motor pod flutter ultimately becoming a primary design constraint. Designers successively generated models to examine stress, dynamics, and aeroelasticity concurrently. This research specifically addressed satisfying multi-disciplinary design criteria to generate fluid-structure interaction solution sets, and produced high aspect ratio primary structure designs for the NASA Scalable Convergent Electric Propulsion Technology and Operations Research (SCEPTOR) project in the Aeronautic Research Mission Directorate at NASA. In this paper, a dynamics-driven, quasi-inverse design methodology is presented to address aerodynamic performance goals and structural challenges encountered for the SCEPTOR demonstrator vehicle. These results are compared with a traditional computer aided design based approach.

Moore, James B.↗

NASA NTRS1992

Transient/structural analysis of a combustor under explosive loads

The 8-Foot High Temperature Tunnel (HTT) at NASA Langley Research Center is a combustion-driven blow-down wind tunnel. A major potential failure mode that was considered during the combustor redesign was the possibility of a deflagration and/or detonation in the combustor. If a main burner flame-out were to occur, then unburned fuel gases could accumulate and, if reignited, an explosion could occur. An analysis has been performed to determine the safe operating limits of the combustor under transient explosive loads. The failure criteria was defined and the failure mechanisms were determined for both peak pressures and differential pressure loadings. An overview of the gas dynamics analysis was given. A finite element model was constructed to evaluate 13 transient load cases. The sensitivity of the structure to the frequency content of the transient loading was assessed. In addition, two closed form dynamic analyses were conducted to verify the finite element analysis. It was determined that the differential pressure load or thrust load was the critical load mechanism and that the nozzle is the weak link in the combustor system.

Gregory, Peyton B.↗

NASA NTRSDate not supplied

A Summary of Recent SBIR and STTR Projects in Support of Cryogenic Fluid Management Modeling

Over the last five years, NASA has invested in development of sub-models to improve the current capabilities of Computational Fluid Dynamics (CFD) codes utilized for cryogenic fluid management (CFM) predictions. The goal is for the submodels to be modular so that they can easily be implemented into NASA’s CFD codes such as ANSYS Fluent, Siemens STAR-CCM+, and Streamline Numerics Loci-Stream. These technologies are being developed through Small Business Innovative Research (SBIR) and Small Business Technology Transfer (STTR) contracts under the CFM subtopic solicitation and funded by the CFM Portfolio Project. All contracts presented include cryogenic experiments to enable model validation. A sub-grid boiling model has been developed by Combustion Research and Flow Technology Inc. and Massachusetts Institute of Technology for cryogenic line chilldown predictions and has completed Phase III. A spray chill-and-fill model for cryogenic tankage is being developed by Combustion Research and Flow Technology Inc. and University of Connecticut and is currently in Phase II. Last, film condensation modeling to support cryogenic liquefaction was progressed through a Phase I contract with Combustion Research and Flow Technology Inc. and University of Connecticut. A summary of accomplishments for each contract is presented.

cryogenics↗

NASA NTRSDate not supplied

A Simple Model for Rotating Detonation Rocket Engine Sizing and Performance Estimates

A Rotating Detonation Rocket Engine (RDRE) model is described which characterizes the device as an essentially infinite number of circumferentially arranged, sequentially firing pulse detonation engine (PDE) tubes. The PDE tubes are in turn treated as lumped-parameter chambers, each of which executes a dynamic (i.e. time dependent) Atkinson cycle with an allowance for a finite Mach number during the refill portion of the cycle. The formulation results in several free parameters which are used to essentially tune the model such that it closely matches higher fidelity computational fluid dynamic RDRE simulations in terms of performance, flow rates, etc. The simplicity of the model allows for straightforward combination with component models (e.g. turbopumps, cooling jackets, etc.) as well as implementation within larger vehicle simulations. Such capability allows for realistic mission analyses and benefits studies which will help to determine the best application for the RDRE concept.

detonation↗

NASA NTRSDate not supplied

A Simple Model for Rotating Detonation Rocket Engine Sizing and Performance Estimates

detonation↗

NASA NTRS1976

Composite propellant combustion modeling with a porous plate burner

A burner is designed to model on a large scale (millimeters) the complex vapor phase processes that take place on awkwardly small scales (about a hundred microns) in the combustion of practical AP/ composite propellants at conventional pressures. Binder vapor evolution is modeled with gaseous fuel (ethane in the experiments reported) flow through a porous plate and oxidizer vapor with the flow of a gaseous oxidizer (air and enriched air in the experiments reported) through discrete holes in the porous plate. Measured flame standoff distance and surface temperature variations are consistent with theoretical predictions at atmospheric pressure. Data obtained at several atmospheres are discussed in the light of the current theories of propellant burning that place varying emphasis on the roles of chemical kinetics and fluid dynamic diffusion/mixing in the vapor phase of a burning composite propellant. The potential and proposed future applications of the burner are indicated.

Kumar, R. N.↗

NASA NTRS1990

Carbon deposition model for oxygen-hydrocarbon combustion. Task 6: Data analysis and formulation of an empirical model

The formation and deposition of carbon (soot) was studied in the Carbon Deposition Model for Oxygen-Hydrocarbon Combustion Program. An empirical, 1-D model for predicting soot formation and deposition in LO2/hydrocarbon gas generators/preburners was derived. The experimental data required to anchor the model were identified and a test program to obtain the data was defined. In support of the model development, cold flow mixing experiments using a high injection density injector were performed. The purpose of this investigation was to advance the state-of-the-art in LO2/hydrocarbon gas generator design by developing a reliable engineering model of gas generator operation. The model was formulated to account for the influences of fluid dynamics, chemical kinetics, and gas generator hardware design on soot formation and deposition.

Makel, Darby B.↗

NASA NTRS1991

The direct simulation of high-speed mixing-layers without and with chemical heat release

A direct numerical simulation of high speed reacting and non-reacting flows for H2-air systems is presented. The calculations are made for a convective Mach number of 0.38 with hyperbolic tangent initial profile and finite rate chemical reactions. A higher-order numerical method is used in time accurate mode to time advance the solution to a statistical steady state. About 600 time slices of all the variables are then stored for statistical analysis. It is shown that most of the problems of high-speed combustion with air are characterized by relatively weak heat release. The present study shows that: (1) the convective speed is reduced by heat release by about 10 percent at this convective Mach number M(sub c) = 0.38; (2) the variation of the mean and rms fluctuation of temperature can be explained on the basis of temperature fluctuation between the flame temperature and the ambient; (3) the growth rate with heat release is reduced by 7 percent; and (4) the entrainment is reduced by 25 percent with heat release. These differences are small in comparison with incompressible flow dynamics, and are argued to be due to the reduced importance of heat release in comparison with the large enthalpy gradients resulting from the large-scale vortex dynamics. It is finally suggested that the problems of reduced mixing in high-speed flows are not severely complicated by heat release.

Sekar, B.↗

NASA NTRS1987

Computational fluid dynamics research at the United Technologies Research Center requiring supercomputers

An overview of research activities at the United Technologies Research Center (UTRC) in the area of Computational Fluid Dynamics (CFD) is presented. The requirement and use of various levels of computers, including supercomputers, for the CFD activities is described. Examples of CFD directed toward applications to helicopters, turbomachinery, heat exchangers, and the National Aerospace Plane are included. Helicopter rotor codes for the prediction of rotor and fuselage flow fields and airloads were developed with emphasis on rotor wake modeling. Airflow and airload predictions and comparisons with experimental data are presented. Examples are presented of recent parabolized Navier-Stokes and full Navier-Stokes solutions for hypersonic shock-wave/boundary layer interaction, and hydrogen/air supersonic combustion. In addition, other examples of CFD efforts in turbomachinery Navier-Stokes methodology and separated flow modeling are presented. A brief discussion of the 3-tier scientific computing environment is also presented, in which the researcher has access to workstations, mid-size computers, and supercomputers.

Landgrebe, Anton J.↗

NASA NTRS2004

Light Microscopy Module Imaging Tested and Demonstrated

The Fluids Integrated Rack (FIR), a facility-class payload, and the Light Microscopy Module (LMM), a subrack payload, are integrated research facilities that will fly in the U.S. Laboratory module, Destiny, aboard the International Space Station. Both facilities are being engineered, designed, and developed at the NASA Glenn Research Center by Northrop Grumman Information Technology. The FIR is a modular, multiuser scientific research facility that is one of two racks that make up the Fluids and Combustion Facility (the other being the Combustion Integrated Rack). The FIR has a large volume dedicated for experimental hardware; easily reconfigurable diagnostics, power, and data systems that allow for unique experiment configurations; and customizable software. The FIR will also provide imagers, light sources, power management and control, command and data handling for facility and experiment hardware, and data processing and storage. The first payload in the FIR will be the LMM. The LMM integrated with the FIR is a remotely controllable, automated, on-orbit microscope subrack facility, with key diagnostic capabilities for meeting science requirements--including video microscopy to observe microscopic phenonema and dynamic interactions, interferometry to make thin-film measurements with nanometer resolution, laser tweezers to manipulate micrometer-sized particles, confocal microscopy to provide enhanced three-dimensional visualization of structures, and spectrophotometry to measure the photonic properties of materials. Vibration disturbances were identified early in the LMM development phase as a high risk for contaminating the science microgravity environment. An integrated FIR-LMM test was conducted in Glenn's Acoustics Test Laboratory to assess mechanical sources of vibration and their impact to microscopic imaging. The primary purpose of the test was to characterize the LMM response at the sample location, the x-y stage within the microscope, to vibration emissions from the FIR and LMM support structures.

Gati, Frank↗

NASA NTRS2008

Forebody and Inlet Design for the HIFiRE 2 Flight Test

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

Ferlemann, Paul G.↗

NASA NTRS2015

Effects of Burning Alternative Fuel in a 5-Cup Combustor Sector

A goal of NASA's Environmentally Responsible Aviation (ERA) program is to develop a combustor that will reduce the NOx emissions and that can burn both standard and alternative fuels. To meet this goal, NASA partnered with General Electric Aviation to develop a 5-cup combustor sector; this sector was tested in NASA Glenn's Advanced Subsonic Combustion Rig (ASCR). To verify that the combustor sector was fuel-flexible, it was tested with a 50-50 blend of JP-8 and a biofuel made from the camelina sativa plant. Results from this test were compared to results from tests where the fuel was neat JP-8. Testing was done at three combustor inlet conditions: cruise, 30% power, and 7% power. When compared to burning JP-8, burning the 50-50 blend did not significantly affect emissions of NOx, CO, or total hydrocarbons. Furthermore, it did not significantly affect the magnitude and frequency of the dynamic pressure fluctuations.

alternative fuel↗

NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E^3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P(41) = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E^3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P(41) = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗

NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P41 = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P41 = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗

NASA NTRS1994

Multi-dimensional combustor flowfield analyses in gas-gas rocket engine

The objectives of the present research are to improve design capabilities for low thrust rocket engines through understanding of the detailed mixing and combustions processes. Of particular interest is a small gaseous hydrogen-oxygen thruster which is considered as a coordinated part of an on-going experimental program at NASA LeRC. Detailed computational modeling requires the application of the full three-dimensional Navier Stokes equations, coupled with species diffusion equations. The numerical procedure is performed on both time-marching and time-accurate algorithms and using an LU approximate factorization in time, flux split upwinding differencing in space. The emphasis in this paper is focused on using numerical analysis to understand detailed combustor flowfields, including the shear layer dynamics created between fuel film cooling and the core gas in the vicinity on the nearby combustor wall; the integrity and effectiveness of the coolant film; three-dimensional fuel jets injection/mixing/combustion characteristics; and their impacts on global engine performance.

Tsuei, Hsin-Hua↗

NASA NTRS2013

OH PLIF Visualization of the UVa Supersonic Combustion Experiment: Configuration C

Non-intrusive hydroxyl radical (OH) planar laser-induced fluorescence (PLIF) measurements were obtained in configuration C of the University of Virginia supersonic combustion experiment. The combustion of hydrogen fuel injected through an unswept compression ramp into a supersonic cross-flow was imaged over a range of streamwise positions. Images were corrected for optical distortion, variations in the laser sheet profile, and different camera views. Results indicate an effect of fuel equivalence ratio on combustion zone shape and local turbulence length scale. The streamwise location of the reaction zone relative to the fuel injector was also found to be sensitive to the fuel equivalence ratio. The flow boundary conditions in the combustor section, which are sensitive to the fuel flow rate, are believed to have caused this effect. A combination of laser absorption and radiative trapping effects are proposed to have caused asymmetry observed in the images. The results complement previously published OH PLIF data obtained for configuration A along with other non-intrusive measurements to form a database for computational fluid dynamics (CFD) model validation.

McRae, Colin D.↗