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All sourcesNASA NTRSDOE OSTI
NASA NTRS1998

Studying Turbulence Using Numerical Simulation Databases: Proceedings of the Summer Program - No. 7

The Seventh Summer Program of the Center for Turbulence Research took place in the four-week period, July 5 to July 31, 1998. This was the largest CTR Summer Program to date, involving thirty-six participants from the U. S. and nine other countries. Thirty-one Stanford and NASA-Ames staff members facilitated and contributed to most of the Summer projects. A new feature, and perhaps a preview of the future programs, was that many of the projects were executed on non-NASA computers. These included supercomputers located in Europe as well as those operated by the Departments of Defense and Energy in the United States. In addition, several simulation programs developed by the visiting participants at their home institutions were used. Another new feature was the prevalence of lap-top personal computers which were used by several participants to carry out some of the work that in the past were performed on desk-top workstations. We expect these trends to continue as computing power is enhanced and as more researchers (many of whom CTR alumni) use numerical simulations to study turbulent flows. CTR's main role continues to be in providing a forum for the study of turbulence for engineering analysis and in facilitating intellectual exchange among the leading researchers in the field. Once again the combustion group was the largest. Turbulent combustion has enjoyed remarkable progress in using simulations to address increasingly complex and practically more relevant questions. The combustion group's studies included such challenging topics as fuel evaporation, soot chemistry, and thermonuclear reactions. The latter study was one of three projects related to the Department of Energy's ASCI Program (www.llnl.gov/asci); the other two (rocket propulsion and fire safety) were carried out in the turbulence modeling group. The flow control and acoustics group demonstrated a successful application of the so-called evolution algorithms which actually led to a previously unknown forcing strategy for jets yielding increased spreading rate. A very efficient algorithm for flow in complex geometries with moving boundaries based on the immersed boundary forcing technique was tested with very encouraging results. Also a new strategy for the destruction of aircraft trailing vortices was introduced and tested. The Reynolds Averaged Modeling (RANS) group demonstrated that the elliptic relaxation concept for RANS calculations is also applicable to transonic flows with shocks; however, prediction of laminar/turbulent transition remains an important pacing item. A large fraction of the LES effort was devoted to the development and testing of a new algorithmic procedure (as opposed to phenomenological model) for subgrid scale modeling based on regularized de-filtering of the flow variables. This appears to be a very promising approach, and a significant effort is currently underway to assess its robustness in high Reynolds number flows and in conjunction with numerical methods for complex flows. As part of the Summer Program two review tutorials were given on Turbulent structures in hydrocarbon pool fires (Sheldon Tieszen), and Turbulent combustion modeling: from RANS to LES via DNS (Luc Vervisch); and two seminars entitled Assessment of turbulence models for engineering applications (Paul Durbin) and Subgrid-scale modeling for non-premixed, turbulent reacting flows (James Riley) were presented. A number of colleagues from universities, government agencies, and industry attended the final presentations of the participants on July 31 and participated in the discussions. There are twenty-six papers in this volume grouped in five areas. Each group is preceded with an overview by its coordinator.

Source record↗
NASA NTRS2000

A Design Tool for Liquid Rocket Engine Injectors

A practical design tool which emphasizes the analysis of flowfields near the injector face of liquid rocket engines has been developed and used to simulate preliminary configurations of NASA's Fastrac and vortex engines. This computational design tool is sufficiently detailed to predict the interactive effects of injector element impingement angles and points and the momenta of the individual orifice flows and the combusting flow which results. In order to simulate a significant number of individual orifices, a homogeneous computational fluid dynamics model was developed. To describe sub- and supercritical liquid and vapor flows, the model utilized thermal and caloric equations of state which were valid over a wide range of pressures and temperatures. The model was constructed such that the local quality of the flow was determined directly. Since both the Fastrac and vortex engines utilize RP-1/LOX propellants, a simplified hydrocarbon combustion model was devised in order to accomplish three-dimensional, multiphase flow simulations. Such a model does not identify drops or their distribution, but it does allow the recirculating flow along the injector face and into the acoustic cavity and the film coolant flow to be accurately predicted.

Farmer, R.↗
NASA NTRS2002

Assessment of Various Flow Solvers Used to Predict the Thermal Environment inside Space Shuttle Solid Rocket Motor Joints

It is very important to accurately predict the gas pressure, gas and solid temperature, as well as the amount of O-ring erosion inside the space shuttle Reusable Solid Rocket Motor (RSRM) joints in the event of a leak path. The scenarios considered are typically hot combustion gas rapid pressurization events of small volumes through narrow and restricted flow paths. The ideal method for this prediction is a transient three-dimensional computational fluid dynamics (CFD) simulation with a computational domain including both combustion gas and surrounding solid regions. However, this has not yet been demonstrated to be economical for this application due to the enormous amount of CPU time and memory resulting from the relatively long fill time as well as the large pressure and temperature rising rate. Consequently, all CFD applications in RSRM joints so far are steady-state simulations with solid regions being excluded from the computational domain by assuming either a constant wall temperature or no heat transfer between the hot combustion gas and cool solid walls.

Wang, Qun-Zhen↗
NASA NTRS2001

Experimental and Computational Study of Trapped Vortex Combustor Sector Rig with Tri-pass Diffuser

The Trapped Vortex Combustor (TVC) potentially offers numerous operational advantages over current production gas turbine engine combustors. These include lower weight, lower pollutant emissions, effective flame stabilization, high combustion efficiency, excellent high altitude relight capability, and operation in the lean burn or RQL (Rich burn/Quick mix/Lean burn) modes of combustion. The present work describes the operational principles of the TVC, and provides detailed performance data on a configuration featuring a tri-pass diffusion system. Performance data include EINOx (NO(sub x) emission index) results for various fuel-air ratios and combustor residence times, combustion efficiency as a function of combustor residence time, and combustor lean blow-out (LBO) performance. Computational fluid dynamics (CFD) simulations using liquid spray droplet evaporation and combustion modeling are performed and related to flow structures observed in photographs of the combustor. The CFD results are used to understand the aerodynamics and combustion features under different fueling conditions. Performance data acquired to date are favorable in comparison to conventional gas turbine combustors. Further testing over a wider range of fuel-air ratios, fuel flow splits, and pressure ratios is in progress to explore the TVC performance. In addition, alternate configurations for the upstream pressure feed, including bi-pass diffusion schemes, as well as variations on the fuel injection patterns, are currently in test and evaluation phases.

Hendricks, Robert C.↗
NASA NTRS2004

Experimental and Computational Study of Trapped Vortex Combustor Sector Rig With Tri-Pass Diffuser

The Trapped Vortex Combustor (TVC) potentially offers numerous operational advantages over current production gas turbine engine combustors. These include lower weight, lower pollutant emissions, effective flame stabilization, high combustion efficiency, excellent high altitude relight capability, and operation in the lean burn or RQL modes of combustion. The present work describes the operational principles of the TVC, and extends diffuser velocities toward choked flow and provides system performance data. Performance data include EINOx results for various fuel-air ratios and combustor residence times, combustion efficiency as a function of combustor residence time, and combustor lean blow-out (LBO) performance. Computational fluid dynamics (CFD) simulations using liquid spray droplet evaporation and combustion modeling are performed and related to flow structures observed in photographs of the combustor. The CFD results are used to understand the aerodynamics and combustion features under different fueling conditions. Performance data acquired to date are favorable compared to conventional gas turbine combustors. Further testing over a wider range of fuel-air ratios, fuel flow splits, and pressure ratios is in progress to explore the TVC performance. In addition, alternate configurations for the upstream pressure feed, including bi-pass diffusion schemes, as well as variations on the fuel injection patterns, are currently in test and evaluation phases.

Hendricks, R. C.↗
NASA NTRS1988

Modeling of combustion processes of stick propellants via combined Eulerian-Lagrangian approach

This research is motivated by the improved ballistic performance of large-caliber guns using stick propellant charges. A comprehensive theoretical model for predicting the flame spreading, combustion, and grain deformation phenomena of long, unslotted stick propellants is presented. The formulation is based upon a combined Eulerian-Lagrangian approach to simulate special characteristics of the two phase combustion process in a cartridge loaded with a bundle of sticks. The model considers five separate regions consisting of the internal perforation, the solid phase, the external interstitial gas phase, and two lumped parameter regions at either end of the stick bundle. For the external gas phase region, a set of transient one-dimensional fluid-dynamic equations using the Eulerian approach is obtained; governing equations for the stick propellants are formulated using the Lagrangian approach. The motion of a representative stick is derived by considering the forces acting on the entire propellant stick. The instantaneous temperature and stress fields in the stick propellant are modeled by considering the transient axisymmetric heat conduction equation and dynamic structural analysis.

Kuo, K. K.↗
NASA NTRS1990

Numerical analysis of bipropellant combustion in orbit maneuvering vehicle thrust chamber

The combustion and aerodynamic processes of liquid-liquid bipropellant; monomethylhydrazine and nitrogen tetroxide in the combustion chamber of a variable thrust engine (VTE) have been examined to assess the flow structure, combustion characteristics, and wall heat transfer under the simulated engine operations at two power levels. The Generalized Elliptic Multi-phase flow with Chemical reaction, Heat transfer and Interfacial Processes (GEM-CHIP) code and the Bipropellant Combustion (BICOMB) code, which is the GEMCHIP algorithm reformulated in the frame work of a body fitted coordinate system, have been used for the numerical study. Combustion flow structure with non-premixed type injection that simulates the impingement type injector is predicted and the results are compared with that of the premixed injector.

Chiu, H. H.↗
NASA NTRS1999

JANNAF 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee Joint Meeting

Volume 1, the first of three volumes is a compilation of 16 unclassified/unlimited-technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee held jointly with the 181 Propulsion Systems Hazards Subcommittee. The meeting was held on 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Topics covered include overviews of RBCC and PDE hypersonic technology, Hyper-X propulsion ground testing, development of JP-8 for hypersonic vehicle applications, numerical simulation of dual-mode SJ combustion, V&V of M&S computer codes, MHD SJ and Rocket Based Combined Cycle (RBCC) launch vehicle concepts, and Pulse Detonation Engine (PDE) propulsion technology development including fundamental investigations, modeling, aerodynamics, operation and performance.

Fry, Ronald S.↗
NASA NTRSDate not supplied

Thermal Protection System P50 Cork Char Response Characterization to Ascent Flight Shear and Heating Environments

NASA’s Space Launch System (SLS) experienced deflagration of the acreage of the Core Stage base heat shield during the Green Run hot-fire test campaigns and Artemis I flight. The burning products from the base heat shield during Green Run hot-fire test operation led to concerns in aerothermal environments and thermal protection system (TPS) performance prior to Artemis I flight. Updated design thermal environments accounting for cork combustion were developed using Green Run measurements and assessed prior to Artemis I, resulting in increased TPS thickness for the base heat shield. This work documents a highly controlled ground test within the NASA Marshall Space Flight Center’s Propulsion Test Branch Hot Gas Facility (HGF) and investigates the P50 cork TPS response due to sensitivity changes in radiative heat rates and surface shear stress. Prior to testing at the HGF, most of the P50 cork panels had a thick layer of char due to the cork combustion process replicated in a thermal vacuum chamber. SLS ascent shear stress and heating rate profile were simulated within the HGF by running the combustion-driven wind tunnel at off-nominal conditions and Mach numbers below 1. This test data provided insight into aerothermal measurements from Artemis I flight as well as informing updated base heating design environments for Artemis II. An in-depth understanding of the TPS P50 cork char response and observations is presented for launch vehicle flight ascent environments.

aerothermodynamics↗
NASA NTRS2004

Quantitative Analysis of Spectral Interference of Spontaneous Raman Scattering in High-Pressure Fuel-Rich H2-Air Combustion

We present a theoretical study of the spectral interferences in the spontaneous Raman scattering spectra of major combustion products in 30-atm fuel-rich hydrogen-air flames. An effective methodology is introduced to choose an appropriate line-shape model for simulating Raman spectra in high-pressure combustion environments. The Voigt profile with the additive approximation assumption was found to provide a reasonable model of the spectral line shape for the present analysis. The rotational/vibrational Raman spectra of H2, N2, and H2O were calculated using an anharmonic-oscillator model using the latest collisional broadening coefficients. The calculated spectra were validated with data obtained in a 10-atm fuel-rich H2-air flame and showed excellent agreement. Our quantitative spectral analysis for equivalence ratios ranging from 1.5 to 5.0 revealed substantial amounts of spectral cross-talk between the rotational H2 lines and the N2 O-/Q-branch; and between the vibrational H2O(0,3) line and the vibrational H2O spectrum. We also address the temperature dependence of the spectral cross-talk and extend our analysis to include a cross-talk compensation technique that removes the nterference arising from the H2 Raman spectra onto the N2, or H2O spectra.

Kojima, Jun↗
NASA NTRS1990

Chemical reacting flows

Future aerospace propulsion concepts involve the combustion of liquid or gaseous fuels in a highly turbulent internal airstream. Accurate predictive computer codes which can simulate the fluid mechanics, chemistry, and turbulence-combustion interaction of these chemical reacting flows will be a new tool that is needed in the design of these future propulsion concepts. Experimental and code development research is being performed at LeRC to better understand chemical reacting flows with the long-term goal of establishing these reliable computer codes. Our approach to understand chemical reacting flows is to look at separate, more simple parts of this complex phenomenon as well as to study the full turbulent reacting flow process. As a result, we are engaged in research on the fluid mechanics associated with chemical reacting flows. We are also studying the chemistry of fuel-air combustion. Finally, we are investigating the phenomenon of turbulence-combustion interaction. Research, both experimental and analytical, is highlighted in each of these three major areas.

Mularz, Edward J.↗
NASA NTRSDate not supplied

RDE Nozzle Computational Design Methodology Development and Application

The Open National Combustion Code (OpenNCC) is used to simulate the aerospike exhaust nozzle region of a rotating detonation engine (RDE).The main objectives of the current study are twofold. The first goal is to validate the proposed computational methodology using the experimental data. The second goal is to demonstrate how the validated prediction tool can be used to optimize the nozzle geometry. To achieve a significant speed-up of computational time, the computational domain in divided into two parts: a combustion region and a throat-nozzle section. In the combustion region, a validated quasi-two-dimensional in-house code is utilized to generate an unsteady RDE flow field solution just upstream of the combustor throat. Subsequently, the unsteady flow data is fed into the three-dimensional throat-nozzle section as the inflow boundary condition. This facilitates the design optimization process since the unsteady inflow can be reused, and a relatively coarse mesh (i.e., larger time-step) can be used to analyze the flow fields around the nozzle. Five nozzle designs were studied and the predicted performance (i.e., thrust)compared. The optimized nozzle was found to produce 3.2% more overall thrust than a baseline nozzle design. Without the nozzle, there is a large low-pressure region at the throat exit, which significantly reduces the overall performance. This methodology is shown to be a promising approach to explore a wide variety of nozzle geometries in a relatively short amount of time.

CFD↗
NASA NTRS1994

Development and application of computational aerothermodynamics flowfield computer codes

Research was performed in the area of computational modeling and application of hypersonic, high-enthalpy, thermo-chemical nonequilibrium flow (Aerothermodynamics) problems. A number of computational fluid dynamic (CFD) codes were developed and applied to simulate high altitude rocket-plume, the Aeroassist Flight Experiment (AFE), hypersonic base flow for planetary probes, the single expansion ramp model (SERN) connected with the National Aerospace Plane, hypersonic drag devices, hypersonic ramp flows, ballistic range models, shock tunnel facility nozzles, transient and steady flows in the shock tunnel facility, arc-jet flows, thermochemical nonequilibrium flows around simple and complex bodies, axisymmetric ionized flows of interest to re-entry, unsteady shock induced combustion phenomena, high enthalpy pulsed facility simulations, and unsteady shock boundary layer interactions in shock tunnels. Computational modeling involved developing appropriate numerical schemes for the flows on interest and developing, applying, and validating appropriate thermochemical processes. As part of improving the accuracy of the numerical predictions, adaptive grid algorithms were explored, and a user-friendly, self-adaptive code (SAGE) was developed. Aerothermodynamic flows of interest included energy transfer due to strong radiation, and a significant level of effort was spent in developing computational codes for calculating radiation and radiation modeling. In addition, computational tools were developed and applied to predict the radiative heat flux and spectra that reach the model surface.

Venkatapathy, Ethiraj↗
NASA NTRS2016

Green Propellant Test Capabilities of the Altitude Combustion Stand at the NASA Glenn Research Center

The NASA Glenn Research Center (GRC) is committed to providing simulated altitude rocket test capabilities to NASA programs, other government agencies, private industry partners, and academic partners. A primary facility to support those needs is the Altitude Combustion Stand (ACS). ACS provides the capability to test combustion components at a simulated altitude up to 100,000 ft. (approx.0.2 psia/10 Torr) through a nitrogen-driven ejector system. The facility is equipped with an axial thrust stand, gaseous and cryogenic liquid propellant feed systems, data acquisition system with up to 1000 Hz recording, and automated facility control system. Propellant capabilities include gaseous and liquid hydrogen, gaseous and liquid oxygen, and liquid methane. A water-cooled diffuser, exhaust spray cooling chamber, and multi-stage ejector systems can enable run times up to 180 seconds to 16 minutes. The system can accommodate engines up to 2000-lbf thrust, liquid propellant supply pressures up to 1800 psia, and test at the component level. Engines can also be fired at sea level if needed. The NASA GRC is in the process of modifying ACS capabilities to enable the testing of green propellant (GP) thrusters and components. Green propellants are actively being explored throughout government and industry as a non-toxic replacement to hydrazine monopropellants for applications such as reaction control systems or small spacecraft main propulsion systems. These propellants offer increased performance and cost savings over hydrazine. The modification of ACS is intended to enable testing of a wide range of green propellant engines for research and qualification-like testing applications. Once complete, ACS will have the capability to test green propellant engines up to 880 N in thrust, thermally condition the green propellants, provide test durations up to 60 minutes depending on thrust class, provide high speed control and data acquisition, as well as provide advanced imaging and diagnostics such as infrared (IR) imaging.

Testing↗
NASA NTRS1992

An analysis of combustion studies in shock expansion tunnels and reflected shock tunnels

The effect of initial nonequilibrium dissociated air constituents on the combustion of hydrogen in high-speed flows for a simulated Mach 17 flight condition was investigated by analyzing the results of comparative combustion experiments performed in a reflected shock tunnel test gas and in a shock expansion tunnel test gas. The results were analyzed and interpreted with a one-dimensional quasi-three-stream combustor code that includes finite rate combustion chemistry. The results of this study indicate that the combustion process is kinetically controlled in the experiments in both tunnels and the presence of the nonequilibrium partially dissociated oxygen in the reflected shock tunnel enhances the combustion. Methods of compensating for the effect of dissociated oxygen are discussed.

Jachimowski, Casimir J.↗
NASA NTRS2015

Unsteady Three-Dimensional Simulation of a Shear Coaxial GO2/GH2 Rocket Injector with RANS and Hybrid-RAN-LES/DES Using Flamelet Models

Historically, the analysis and design of liquid rocket engines (LREs) has relied on full‐scale testing and one‐dimensional empirical tools. The testing is extremely expensive and the one‐dimensional tools are not designed to capture the highly complex, and multi‐dimensional features that are inherent to LREs. Recent advances in computational fluid dynamics (CFD) tools have made it possible to predict liquid rocket engine performance, stability, to assess the effect of complex flow features, and to evaluate injector‐driven thermal environments, to mitigate the cost of testing. Extensive efforts to verify and validate these CFD tools have been conducted, to provide confidence for using them during the design cycle. Previous validation efforts have documented comparisons of predicted heat flux thermal environments with test data for a single element gaseous oxygen (GO2) and gaseous hydrogen (GH2) injector. The most notable validation effort was a comprehensive validation effort conducted by Tucker et al. [1], in which a number of different groups modeled a GO2/GH2 single element configuration by Pal et al [2]. The tools used for this validation comparison employed a range of algorithms, from both steady and unsteady Reynolds Averaged Navier‐Stokes (U/RANS) calculations, large‐eddy simulations (LES), detached eddy simulations (DES), and various combinations. A more recent effort by Thakur et al. [3] focused on using a state‐of‐the‐art CFD simulation tool, Loci/STREAM, on a two‐dimensional grid. Loci/STREAM was chosen because it has a unique, very efficient flamelet parameterization of combustion reactions that are too computationally expensive to simulate with conventional finite‐rate chemistry calculations. The current effort focuses on further advancement of validation efforts, again using the Loci/STREAM tool with the flamelet parameterization, but this time with a three‐dimensional grid. Comparisons to the Pal et al. heat flux data will be made for both RANS and Hybrid RANSLES/ Detached Eddy simulations (DES). Computation costs will be reported, along with comparison of accuracy and cost to much less expensive two‐dimensional RANS simulations of the same geometry.

Westra, Doug G.↗
NASA NTRS2005

Control of Combustion-Instabilities Through Various Passive Devices

It is well known that under some operating conditions, rocket engines (using solid or liquid fuels) exhibit unstable modes of operation that can lead to engine malfunction and shutdown. The sources of these instabilities are diverse and are dependent on fuel, chamber geometry and various upstream sources such as pumps, valves and injection mechanism. It is believed that combustion-acoustic instabilities occur when the acoustic energy increase due to the unsteady heat release of the flame is greater than the losses of acoustic energy from the system [1, 2]. Giammar and Putnam [3] performed a comprehensive study of noise generated by gasfired industrial burners and made several key observations; flow noise was sometimes more intense than combustion roar, which tended to have a characteristic frequency spectrum. Turbulence was amplified by the flame. The noise power varied directly with combustion intensity and also with the product of pressure drop and heat release rate. Karchmer [4] correlated the noise emitted from a turbofan jet engine with that in the combustion chamber. This is important, since it quantified how much of the noise from an engine originates in the combustor. A physical interpretation of the interchange of energy between sound waves and unsteady heat release rates was given by Rayleigh [5] for inviscid, linear perturbations. Bloxidge et al [6] extended Rayleigh s criterion to describe the interaction of unsteady combustion with one-dimensional acoustic waves in a duct. Solutions to the mass, momentum and energy conservation equations in the pre- and post-flame zones were matched by making several assumptions about the combustion process. They concluded that changes in boundary conditions affect the energy balance of acoustic waves in the combustor. Abouseif et al [7] also solved the one-dimensional flow equations, but they used a onestep reaction to evaluate the unsteady heat release rate by relating it to temperature and velocity perturbations. Their analysis showed that oscillations arise from coupling between entropy waves produced at the flame and pressure waves originating from the nozzle. Yang and Culick [8] assumed a thin flame sheet, which is distorted by velocity and pressure oscillations. Conservation equations were expressed in integral form and solutions for the acoustic wave equations and complex frequencies were obtained. The imaginary part of the frequency indicated stability regions of the flame. Activation energy asymptotics together with a one-step reaction were used by McIntosh [9] to study the effects of acoustic forcing and feedback on unsteady, one-dimensional flames. He found that the flame stability was altered by the upstream acoustic feedback. Shyy et al [10] used a high-accuracy TVD scheme to simulate unsteady, one-dimensional longitudinal, combustion instabilities. However, numerical diffusion was not completely eliminated. Recently, Prasad [11] investigated numerically the interactions of pressure perturbations with premixed flames. He used complex chemistry to study responses of pressure perturbations in one-dimensional combustors. His results indicated that reflected and transmitted waves differed significantly from incident waves.

Frendi, Kader↗