Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “control volume methods”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Methods for determining the internal thrust of scramjet engine modules from experimental data

Methods for calculating zero-fuel internal drag of scramjet engine modules from experimental measurements are presented. These methods include two control-volume approaches, and a pressure and skin-friction integration. The three calculation techniques are applied to experimental data taken during tests of a version of the NASA parametric scramjet. The methods agree to within seven percent of the mean value of zero-fuel internal drag even though several simplifying assumptions are made in the analysis. The mean zero-fuel internal drag coefficient for this particular engine is calculated to be 0.150. The zero-fuel internal drag coefficient when combined with the change in engine axial force with and without fuel defines the internal thrust of an engine.

Voland, Randall T.↗
DOE OSTI2020

Computational model for predicting the performance of fluid-powered longiudinal piston modules and motors.

A Matlab-based computational model line code has been developed for predicting the performance of fluid-powered longitudinal piston modules and motors. A system of equations has been developed based upon applying the conservation of mass to a control volume defined by a reciprocating piston. The method results in a differential equation that predicts the pressure gradient in the control volume as a function of rotor rotation angle. The model uses an integration scheme to predict the pressure state in the control volume. The orifice equation is used to predict the flowrate in and out of the control volume. From the resulting solution the torque and speed characteristics of the motor configuration can be predicted, optimized, or tailored for a specific requirement.SAND2019-13983 M Sandia National Laboratories is a multimission laboratory managed and operated by National Technology & Engineering Solutions of Sandia, LLC, a wholly owned subsidiary of Honeywell International Inc., for the U.S. Department of Energy’s National Nuclear Security Administration under contract DE-NA0003525.

Taylor, Michael↗
DOE OSTI2014

Limitation of parallel flow in double diffusive convection: Two- and three-dimensional transitions in a horizontal porous domain

Two-dimensional (2D) and three-dimensional (3D) numerical simulations of double diffusion natural convection in an elongated enclosure filled with a binary fluid saturating a porous medium are carried out in the present work. The Boussinesq approximation is made in the formulation of the problem, and Neumann boundary conditions for temperature and concentration are adopted, respectively, on vertical and horizontal walls of the cavity. The used numerical method is based on the control volume approach, with the third order quadratic upstream interpolation scheme in approximating the advection terms. A semi implicit method algorithm is used to handle the velocity-pressure coupling. To avoid the excessively high computer time inherent to the solution of 3D natural convection problems, full approximation storage with full multigrid method is used to solve the problem. A wide range of the controlling parameters (Rayleigh-Darcy number Ra, lateral aspect ratio Ay, Lewis number Le, and the buoyancy ration N) is investigated. We clearly show that increasing the depth of the cavity (i.e., the lateral aspect ratio) has an important effect on the flow patterns. The 2D perfect parallel flows obtained for small lateral aspect ratio are drastically destabilized by increasing the cavity lateral dimension. This yields a 3D fluid motion with a much more complex flow pattern and the usually considered 2D parallel flow model cannot be applied.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
DOE OSTI2022

An immersed interface method for the 2D vorticity-velocity Navier-Stokes equations with multiple bodies

We present an immersed interface method for the vorticity-velocity form of the 2D Navier Stokes equations that directly addresses challenges posed by nonconvex immersed bodies, multiply connected domains, and the calculation of force distributions on immersed surfaces. The immersed interface method is re-interpreted as a polynomial extrapolation of flow quantities and boundary conditions into the immersed solid bodies, reducing computational cost and enabling simulations with nonconvex bodies that could not be discretized with previous immersed interface methods. In the flow, the vorticity transport equation is discretized using a conservative finite difference scheme and explicit Runge-Kutta time integration. The velocity reconstruction problem is transformed to a scalar Poisson equation that is discretized with conservative finite differences, and solved using an FFT-accelerated iterative algorithm. The use of conservative differencing throughout leads to exact enforcement of a discrete Kelvin's theorem, allowing for simulations with multiply connected domains and outflow boundaries that have challenged other immersed interface vortex methods. We also explore novel methods for recovering time-dependent pressure distributions on immersed bodies within a vorticity-based method and present a novel control volume formulation for recovering aerodynamic moments from only the vorticity and velocity fields. The method achieves second order spatial accuracy and third order temporal accuracy, and is validated on a variety of 2D flows in internal and free-space domains.

97 MATHEMATICS AND COMPUTING↗
NASA NTRS1989

Control-volume based Navier-Stokes equation solver valid at all flow velocities

A control-volume based finite difference method to solve the Reynolds averaged Navier-Stokes equations is presented. A pressure correction equation valid at all flow velocities and a pressure staggered grid layout are used in the method. Example problems presented herein include: a developing laminar channel flow, developing laminar pipe flow, a lid-driven square cavity flow, a laminar flow through a 90-degree bent channel, a laminar polar cavity flow, and a turbulent supersonic flow over a compression ramp. A k-epsilon turbulence model supplemented with a near-wall turbulence model was used to solve the turbulent flow. It is shown that the method yields accurate computational results even when highly skewed, unequally spaced, curved grids are used. It is also shown that the method is strongly convergent for high Reynolds number flows.

Kim, S.-W.↗
NASA NTRS1989

Calculation of reattaching shear layers in divergent channel with a multiple-time-scale turbulence model

Numerical calculations of turbulent reattaching shear layers in a divergent channel are presented. The turbulence is described by a multiple-time-scale turbulence model. The turbulent flow equations are solved by a control-volume based finite difference method. The computational results are compared with those obtained using k-epsilon turbulence models and algebraic Reynolds stress turbulence models. It is shown that the multiple-time-scale turbulence model yields significantly improved computational results than the other turbulence models in the region where the turbulence is in a strongly inequilibrium state.

Kim, S.-W.↗
NASA NTRS1989

Numerical investigation of an internal layer in turbulent flow over a curved hill

The development of an internal layer in a turbulent boundary layer flow over a curved hill is investigated numerically. The turbulence field of the boundary layer flow over the curved hill is compared with that of a turbulent flow over a symmetric airfoil (which has the same geometry as the curved hill except that the leading and trailing edge plates were removed) to study the influence of the strongly curved surface on the turbulence field. The turbulent flow equations are solved by a control-volume based finite difference method. The turbulence is described by a multiple-time-scale turbulence model supplemented with a near-wall turbulence model. Computational results for the mean flow field (pressure distributions on the walls, wall shearing stresses and mean velocity profiles), the turbulence structure (Reynolds stress and turbulent kinetic energy profiles), and the integral parameters (displacement and momentum thicknesses) compared favorably with the measured data. Computational results show that the internal layer is a strong turbulence field which is developed beneath the external boundary layer and is located very close to the wall. Development of the internal layer was more obviously observed in the Reynolds stress profiles and in the turbulent kinetic energy profiles than in the mean velocity profiles. In this regard, the internal layers is significantly different from wall-bounded simple shear layers in which the mean velocity profile characterizes the boundary layer most distinguishably. Development of such an internal layer, characterized by an intense turbulence field, is attributed to the enormous mean flow strain rate caused by the streamline curvature and the strong pressure gradient. In the turbulent flow over the curved hill, the internal layer begin to form near the forward corner of the hill, merges with the external boundary layer, and develops into a new fully turbulent boundary layer as the fluid flows in the downstream direction. For the flow over the symmetric airfoil, the boundary layer began to form from almost the same location as that of the curved hill, grew in its strength, and formed a fully turbulent boundary layer from mid-part of the airfoil and in the downstream region. Computational results also show that the detailed turbulence structure in the region very close to the wall of the curved hill is almost the same as that of the airfoil in most of the curved regions except near the leading edge. Thus the internal layer of the curved hill and the boundary layer of the airfoil were also almost the same. Development of the wall shearing stress and separation of the boundary layer at the rear end of the curved hill mostly depends on the internal layer and is only slightly influenced by the external boundary layer flow.

Kim, S-W.↗
NASA NTRS1990

Calculation of reattaching shear layers in divergent channel with a multiple-time-scale turbulence model

Numerical calculations of turbulent reattaching shear layers in a divergent channel are presented. The turbulence is described by a multiple-time-scale turbulence model. The turbulent flow equations are solved by a control-volume based finite difference method. The computational results are compared with those obtained using k-epsilon turbulence models and algebraic Reynolds stress turbulence models. It is shown that the multiple-time-scale turbulence model yields significantly improved computational results than the other turbulence models in the region where the turbulence is in a strongly inequilibrium state.

Kim, S.-W.↗
NASA NTRS1995

Resolution of the buoyancy in the 8-foot high temperature tunnel combustor

Currently, the 8-Foot High Temperature Tunnel (8-Ft. HTT) combustor produces a good profile at only one point (2000 psia and 3650 R with oxygen enrichment). Air is enriched with oxygen (liquid) so that the combustor product gas will contain the volumetric amount of oxygen normally found in air. The oxygen enriched air has a large fraction that is not reacted and flows through the outer periphery of the fuel injector. This ring of cold air in addition to the relatively cold walls of the combustor set up buoyancy forces that produce a segregation of relatively cool gases at the bottom of the combustor exit. The basic problem is to produce a test gas that has uniform properties at all combustor conditions. The combustor temperature may be as high as 3700 R or as low as 2000 R. Combustor pressures can be as high as 3500 psia (no oxygen enrichment) and as low as 600 psia. The segregation is most severe with oxygen enriched air, since its temperature is lower and its density is high. The combustor is lined with nickel 201 and can be operated at about 1600 R maximum. A global mixing process is desired that produces an acceptable profile of temperature, species, and velocity at the exit of the combustor. The ultimate goal is a temperature profile with about 100 R variance and about 2 percent variance in oxygen. The exit total temperature must not be lowered significantly by the mixing apparatus or mechanisms employed. If immersed bodies are used, they must also be kept very hot. All combustor wall modifications must be able to survive the heat and structural conditions of the varied operating conditions. Our approach to resolving this issue is being conducted in three stages: (1) Consider mixing exclusively, (2) Resolve the heat transfer concerns resulting from the chosen mixing strategy, and (3) Solve the material and structural problems resulting from stages (1) and (2). Since the 8-Ft. HTT is unavailable for experimentation, the study is conducted exclusively with computational fluid dynamic (CPD) codes (Fluent/Uns and Rampant 3.1) using unstructured grid through body fitted coordinates. Both CFD codes are general purpose Navier-Stokes solution packages that can solve integral conservation equations for conservation of mass, momentum and energy. The governing equations are discretized using a control-volume finite-element method on unstructured triangular 2-D grids. In the interest of time, a 3-D tetrahedral grid was used to check the 2-D results on one mixing strategy and the 2-D results were confirmed. Preliminary results indicate that excellent mixing can be achieved with a body placed in the center of the flowing hot fluid with a minor modification to the combustor wall similar to a model positioned in the test section of a wind tunnel. The concept here, is to create longitudinal vortices strong enough to bring both fluids into intimate contact with each other near the wall where the cold fluid resides. However, there may be a trade-off in the choice of mixing strategy, heat transfer and structural requirements. Currently, we are examining the more promising geometries for heat transfer concerns and developing strategies for the material of construction of the center body.

Loney, Norman W.↗
NASA NTRS2002

Incorporation of Condensation Heat Transfer in a Flow Network Code

Pure water is distilled from waste water in the International Space Station. The distillation assembly consists of an evaporator, a compressor and a condenser. Vapor is periodically purged from the condenser to avoid vapor accumulation. Purged vapor is condensed in a tube by coolant water prior to entering the purge pump. The paper presents a condensation model of purged vapor in a tube. This model is based on the Finite Volume Method. In the Finite Volume Method, the flow domain is discretized into multiple control volumes and a simultaneous analysis is performed.

Anthony, Miranda↗
NASA NTRS1979

Practical optimal flight control system design for helicopter aircraft. Volume 2: Software user's guide

A method by which modern and classical control theory techniques may be integrated in a synergistic fashion and used in the design of practical flight control systems is presented. A general procedure is developed, and several illustrative examples are included. Emphasis is placed not only on the synthesis of the design, but on the assessment of the results as well. The first step is to establish the differences, distinguishing characteristics and connections between the modern and classical control theory approaches. Ultimately, this uncovers a relationship between bandwidth goals familiar in classical control and cost function weights in the equivalent optimal system. In order to obtain a practical optimal solution, it is also necessary to formulate the problem very carefully, and each choice of state, measurement and output variable must be judiciously considered. Once design goals are established and problem formulation completed, the control system is synthesized in a straightforward manner. Three steps are involved: filter-observer solution, regulator solution, and the combination of those two into the controller. Assessment of the controller permits and examination and expansion of the synthesis results.

Riedel, S. A.↗
NASA NTRS2001

A Finite-Volume "Shaving" Method for Interfacing NASA/DAO''s Physical Space Statistical Analysis System to the Finite-Volume GCM with a Lagrangian Control-Volume Vertical Coordinate

Toward the development of a finite-volume Data Assimilation System (fvDAS), a consistent finite-volume methodology is developed for interfacing the NASA/DAO's Physical Space Statistical Analysis System (PSAS) to the joint NASA/NCAR finite volume CCM3 (fvCCM3). To take advantage of the Lagrangian control-volume vertical coordinate of the fvCCM3, a novel "shaving" method is applied to the lowest few model layers to reflect the surface pressure changes as implied by the final analysis. Analysis increments (from PSAS) to the upper air variables are then consistently put onto the Lagrangian layers as adjustments to the volume-mean quantities during the analysis cycle. This approach is demonstrated to be superior to the conventional method of using independently computed "tendency terms" for surface pressure and upper air prognostic variables.

Lin, Shian-Jiann↗
NASA NTRS1986

An explicit finite-volume time-marching procedure for turbulent flow calculations

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite-volume, time-marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier-Stokes equations. The entire calculation is performed in the physical domain. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is not required. Different time steps are used in the different governing equations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes (length/height - 1000). The method is compared with two test cases. Essentially incompressible turbulent boundary layer flow in an adverse pressure gradient is calculated and the computed distributions of mean velocity and shear are in good agreement with the measurements. Transonic viscous flow in a converging diverging nozzle is calculated; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good.

Nicholson, Stephen↗
NASA NTRS1987

An explicit finite-volume time-marching procedure for turbulent flow calculations

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite-volume, time-marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier-Stokes equations. The entire calculation is performed in the physical domain. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is not required. Different time steps are used in the different governing equations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes (length/height - 1000). The method is compared with two test cases. Essentially incompressible turbulent boundary layer flow in an adverse pressure gradient is calculated and the computed distributions of mean velocity and shear are in good agreement with the measurements. Transonic viscous flow in a converging diverging nozzle is calculated; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good.

Nicholson, Stephen↗
NASA NTRS2007

Synthesis from Design Requirements of a Hybrid System for Transport Aircraft Longitudinal Control

Volume I of this report presents a new method for synthesizing hybrid systems directly from design requirements, and applies the method to design of a hybrid system for longitudinal control of transport aircraft. The resulting system satisfies general requirement for safety and effectiveness specified a priori, enabling formal validation to be achieved. Volume II contains seven appendices intended to make the report accessible to readers with backgrounds in human factors, fli ght dynamics and control. and formal logic. Major design goals are (1) system desi g n integrity based on proof of correctness at the design level, (2), significant simplification and cost reduction in system development and certification, and (3) improved operational efficiency, with significant alleviation of human-factors problems encountered by pilots in current transport aircraft. This report provides for the first time a firm technical basis for criteria governing design and certification of avionic systems for transport aircraft. It should be of primary interest to designers of next-generation avionic systems.

Hynes, Charles S.↗
NASA NTRS2007

Synthesis from Design Requirements of a Hybrid System for Transport Aircraft Longitudinal Control

Volume I of this report presents a new method for synthesizing hybrid systems directly from desi gn requirements, and applies the method to design of a hybrid system for longitudinal control of transport aircraft. The resulting system satisfies general requirement for safety and effectiveness specified a priori, enabling formal validation to be achieved. Volume II contains seven appendices intended to make the report accessible to readers with backgrounds in human factors, flight dynamics and control, and formal logic. Major design goals are (1) system design integrity based on proof of correctness at the design level, (2) significant simplification and cost reduction in system development and certification, and (3) improved operational efficiency, with significant alleviation of human-factors problems encountered by pilots in current transport aircraft. This report provides for the first time a firm technical basis for criteria governing design and certification of avionic systems for transport aircraft. It should be of primary interest to designers of next-generation avionic systems.

Hynes, Charles S.↗
NASA NTRS2018

Optimal Control Prediction Method for Control Allocation

This paper proposes a novel prediction method for online optimal control allocation that extends the volume of moments achievable with the Moore-Penrose generalized inverse to the entire Attainable Moment Set. This method formulates the control allocation problem using selected basis vectors and associated gains which reduces the optimization problem dimensions and provides physical insight into the resulting optimal solutions. The proposed algorithm finds the entire family of unique optimal control solutions along the desired moment vector from the origin to the boundary of the Attainable Moment Set. Numerical results for the Moore-Penrose prediction method show that the unique minimal controls obtained yield the desired moment with near machine precision accuracy while maintaining control effectors within specified position limits. This method has been fully validated against the unique solution obtained on the boundary of the Attainable Moment Set using the Durham Direct Allocation method. Minimal control solutions obtained for moments in the interior of the Attainable Moment Set, similarly yield the desired moment to near machine precision while providing control solutions that are smaller (i.e. 2-norm) than solutions found with traditional control allocation algorithms (e.g. interior point methods) applied to the minimal control problem. Numerical simulations using a Matlab® autocoded executable (MEX) for the representative real world problem of 3-moments with 20 individual control effectors and prescribed control position limits show a mean computation speed of approximately 125 Hz which is sufficient to enable real-time flight allocation.

Acheson, Michael J.↗