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NASA NTRS2010

Establishing Consensus Turbulence Statistics for Hot Subsonic Jets

Many tasks in fluids engineering require knowledge of the turbulence in jets. There is a strong, although fragmented, literature base for low order statistics, such as jet spread and other meanvelocity field characteristics. Some sources, particularly for low speed cold jets, also provide turbulence intensities that are required for validating Reynolds-averaged Navier-Stokes (RANS) Computational Fluid Dynamics (CFD) codes. There are far fewer sources for jet spectra and for space-time correlations of turbulent velocity required for aeroacoustics applications, although there have been many singular publications with various unique statistics, such as Proper Orthogonal Decomposition, designed to uncover an underlying low-order dynamical description of turbulent jet flow. As the complexity of the statistic increases, the number of flows for which the data has been categorized and assembled decreases, making it difficult to systematically validate prediction codes that require high-level statistics over a broad range of jet flow conditions. For several years, researchers at NASA have worked on developing and validating jet noise prediction codes. One such class of codes, loosely called CFD-based or statistical methods, uses RANS CFD to predict jet mean and turbulent intensities in velocity and temperature. These flow quantities serve as the input to the acoustic source models and flow-sound interaction calculations that yield predictions of far-field jet noise. To develop this capability, a catalog of turbulent jet flows has been created with statistics ranging from mean velocity to space-time correlations of Reynolds stresses. The present document aims to document this catalog and to assess the accuracies of the data, e.g. establish uncertainties for the data. This paper covers the following five tasks: Document acquisition and processing procedures used to create the particle image velocimetry (PIV) datasets. Compare PIV data with hotwire and laser Doppler velocimetry (LDV) data published in the open literature. Compare different datasets acquired at roughly the same flow conditions to establish uncertainties. Create a consensus dataset for a range of hot jet flows, including uncertainty bands. Analyze this consensus dataset for self-consistency and compare jet characteristics to those of the open literature. One final objective fulfilled by this work was the demonstration of a universal scaling for the jet flow fields, at least within the region of interest to aeroacoustics. The potential core length and the spread rate of the half-velocity radius were used to collapse of the mean and turbulent velocity fields over the first 20 jet diameters in a highly satisfying manner.

Bridges, James↗
NASA NTRS2003

Influence of Shock Wave on the Flutter Behavior of Fan Blades Investigated

Modern fan designs have blades with forward sweep; a lean, thin cross section; and a wide chord to improve performance and reduce noise. These geometric features coupled with the presence of a shock wave can lead to flutter instability. Flutter is a self-excited dynamic instability arising because of fluid-structure interaction, which causes the energy from the surrounding fluid to be extracted by the vibrating structure. An in-flight occurrence of flutter could be catastrophic and is a significant design issue for rotor blades in gas turbines. Understanding the flutter behavior and the influence of flow features on flutter will lead to a better and safer design. An aeroelastic analysis code, TURBO, has been developed and validated for flutter calculations at the NASA Glenn Research Center. The code has been used to understand the occurrence of flutter in a forward-swept fan design. The forward-swept fan, which consists of 22 inserted blades, encountered flutter during wind tunnel tests at part speed conditions.

Srivastava, Rakesh↗
NASA NTRS2005

Airborne Shaped Sonic Boom Demonstration Pressure Measurements with Computational Fluid Dynamics Comparisons

The Shaped Sonic Boom Demonstration project showed for the first time that by careful design of aircraft contour the resultant sonic boom can maintain a tailored shape, propagating through a real atmosphere down to ground level. In order to assess the propagation characteristics of the shaped sonic boom and to validate computational fluid dynamics codes, airborne measurements were taken of the pressure signatures in the near field by probing with an instrumented F-15B aircraft, and in the far field by overflying an instrumented L-23 sailplane. This paper describes each aircraft and their instrumentation systems, the airdata calibration, analysis of the near- and far-field airborne data, and shows the good to excellent agreement between computational fluid dynamics solutions and flight data. The flights of the Shaped Sonic Boom Demonstration aircraft occurred in two phases. Instrumentation problems were encountered during the first phase, and corrections and improvements were made to the instrumentation system for the second phase, which are documented in the paper. Piloting technique and observations are also given. These airborne measurements of the Shaped Sonic Boom Demonstration aircraft are a unique and important database that will be used to validate design tools for a new generation of quiet supersonic aircraft.

Haering, Edward A., Jr.↗
NASA NTRS2004

Computational Study of Primary Electrons in the Cusp Region of an Ion Engine's Discharge Chamber

In this work a computer code called PRIMA is used to study the motion of primary electrons in the magnetic cusp region of the discharge chamber of an ion engine. Even though the amount of wall area covered by the cusps is very small, the cusp regions are important because prior computational analyses have indicated that most primary electrons leave the discharge chamber through the cusps. The analysis presented here focuses on the cusp region only. The affects of the shape and size of the cusp region on primary electron travel are studied as well as the angle and location at which the electron enters the cusp region. These affects are quantified using the confinement length and the number density distributions of the primary electrons. In addition to these results comparisons of the results from PRIMA are made to experimental results for a cylindrical discharge chamber with two magnetic rings. These comparisons indicate the validity of the computer code called PRIMA.

Stueber, Thomas J.↗
NASA NTRS1995

Simulation of patch and slot antennas using FEM with prismatic elements and investigations of artificial absorber mesh termination schemes

Year 1 progress can be characterized with four major achievements which are crucial toward the development of robust, easy to use antenna analysis code on doubly conformal platforms. (1) A new FEM code was developed using prismatic meshes. This code is based on a new edge based distorted prism and is particularly attractive for growing meshes associated with printed slot and patch antennas on doubly conformal platforms. It is anticipated that this technology will lead to interactive, simple to use codes for a large class of antenna geometries. Moreover, the codes can be expanded to include modeling of the circuit characteristics. An attached report describes the theory and validation of the new prismatic code using reference calculations and measured data collected at the NASA Langley facilities. The agreement between the measured and calculated data is impressive even for the coated patch configuration. (2) A scheme was developed for improved feed modeling in the context of FEM. A new approach based on the voltage continuity condition was devised and successfully tested in modeling coax cables and aperture fed antennas. An important aspect of this new feed modeling approach is the ability to completely separate the feed and antenna mesh regions. In this manner, different elements can be used in each of the regions leading to substantially improved accuracy and meshing simplicity. (3) A most important development this year has been the introduction of the perfectly matched interface (PMI) layer for truncating finite element meshes. So far the robust boundary integral method has been used for truncating the finite element meshes. However, this approach is not suitable for antennas on nonplanar platforms. The PMI layer is a lossy anisotropic absorber with zero reflection at its interface. (4) We were able to interface our antenna code FEMA_CYL (for antennas on cylindrical platforms) with a standard high frequency code. This interface was achieved by first generating equivalent magnetic currents across the antenna aperture using the FEM code. These currents were employed as the sources in the high frequency code.

Gong, J.↗
NASA NTRS2001

Overview of the NCC

A multi-disciplinary design/analysis tool for combustion systems is critical for optimizing the low-emission, high-performance combustor design process. Based on discussions between then NASA Lewis Research Center and the jet engine companies, an industry-government team was formed in early 1995 to develop the National Combustion Code (NCC), which is an integrated system of computer codes for the design and analysis of combustion systems. NCC has advanced features that address the need to meet designer's requirements such as "assured accuracy", "fast turnaround", and "acceptable cost". The NCC development team is comprised of Allison Engine Company (Allison), CFD Research Corporation (CFDRC), GE Aircraft Engines (GEAE), NASA Glenn Research Center (LeRC), and Pratt & Whitney (P&W). The "unstructured mesh" capability and "parallel computing" are fundamental features of NCC from its inception. The NCC system is composed of a set of "elements" which includes grid generator, main flow solver, turbulence module, turbulence and chemistry interaction module, chemistry module, spray module, radiation heat transfer module, data visualization module, and a post-processor for evaluating engine performance parameters. Each element may have contributions from several team members. Such a multi-source multi-element system needs to be integrated in a way that facilitates inter-module data communication, flexibility in module selection, and ease of integration. The development of the NCC beta version was essentially completed in June 1998. Technical details of the NCC elements are given in the Reference List. Elements such as the baseline flow solver, turbulence module, and the chemistry module, have been extensively validated; and their parallel performance on large-scale parallel systems has been evaluated and optimized. However the scalar PDF module and the Spray module, as well as their coupling with the baseline flow solver, were developed in a small-scale distributed computing environment. As a result, the validation of the NCC beta version as a whole was quite limited. Current effort has been focused on the validation of the integrated code and the evaluation/optimization of its overall performance on large-scale parallel systems.

Liu, Nan-Suey↗
NASA NTRS1991

Use of residual redundancy in the design of joint source/channel coders

A technique for providing error protection without the additional overhead required for channel coding is developed, starting from the premise that, during source coder design, for the sake of simplicity or due to imperfect knowledge, assumptions have to be made about the source which are often incorrect. This results in residual redundancy at the output of the source coder. The residual redundancy can then be used to provide error protection in much the same way as the insertion of redundancy in convolutional coding provides error protection. To show the validity of this approach, it is applied to image coding using differential pulse code modulation (DPCM). Substantial performance gains are obtained, in terms of both objective and subjective measures.

Sayood, Khalid↗
NASA NTRS1980

Validation of the solar heating and cooling high speed performance (HISPER) computer code

Developed to give a quick and accurate predictions HISPER, a simplification of the TRNSYS program, achieves its computational speed by not simulating detailed system operations or performing detailed load computations. In order to validate the HISPER computer for air systems the simulation was compared to the actual performance of an operational test site. Solar insolation, ambient temperature, water usage rate, and water main temperatures from the data tapes for an office building in Huntsville, Alabama were used as input. The HISPER program was found to predict the heating loads and solar fraction of the loads with errors of less than ten percent. Good correlation was found on both a seasonal basis and a monthly basis. Several parameters (such as infiltration rate and the outside ambient temperature above which heating is not required) were found to require careful selection for accurate simulation.

Wallace, D. B.↗
NASA NTRS1981

Basic research in wake vortex alleviation using a variable twist wing

The variable twist wing concept was used to investigate the relative effects of lift and turbulence distribution on the rolled up vortex wake. Several methods of reducing the vortex strength behind an aircraft were identified. These involve the redistribution of lift spanwise on the wing and drag distribution along the wing. Initial attempts to use the variable twist wing velocity data to validate the WAKE computer code have shown a strong correlation, although the vorticity levels were not exactly matched.

Morris, D. J.↗
NASA NTRS1981

Transonic swept wings studied by the lifting-line theory

Transonic swept wings are analyzed as a lifting-line problem under a small-disturbance approximation. Basic concepts and principal results of the asymptotic theory are discussed. The study focuses on straight oblique wings and V-shaped swept wings, of which the local centerline curvature can be equated to zero. The three-dimensional (3-D) perturbation of the nonlinear component flow admits a similarity flow structure but requires that all wing sections are generated from a single airfoil profile; the reduced 2-D problems in this case are solved only once for all span stations. Examples of solutions involving high subcritical and slightly supercritical component flows are demonstrated and compared with surface pressure data from 3-D computer codes based on the full-potential equation (FLO 22). Except in the neighborhood of leading edges, where the small-disturbance assumption breaks down, and in the vicinities of wing tips and the symmetry plane, where neither the theory nor the 3-D codes may claim full validity, reasonable agreement is consistently found. The explicit results from the upwash analysis, along with the similarity flow structure, provides a rational approach to the control of 3-D effects in transonic aerodynamic design studies.

Cheng, H. K.↗
NASA NTRS1986

Addressing the hypersonic simulation problem

The hypersonic simulation problem is solved by the meshing of bench-marked experimental data with results from validated computational fluid dynamic codes. The example presented is for the reentry of the Space Shuttle Orbiter. Mach number effects were assessed by parametrically varying free-stream Mach number and angle of attack in a series of inviscid, perfect-gas computations carried out on a modified Orbiter geometry. Real-gas effects were determined by making calculations at specific points of the reentry trajectory using equilibrium air thermodynamics and comparing with corresponding perfect-gas computations. Viscous computations were also made for both the basic Orbiter body and control surfaces. A methodology is developed that permits the extrapolation of wind tunnel data to flight situations, providing a solution to the hypersonic simulation problem.

Griffith, B. J.↗
NASA NTRS1987

Simplified curve fits for the transport properties of equilibrium air

New, improved curve fits for the transport properties of equilibruim air have been developed. The curve fits are for viscosity and Prandtl number as functions of temperature and density, and viscosity and thermal conductivity as functions of internal energy and density. The curve fits were constructed using grabau-type transition functions to model the tranport properties of Peng and Pindroh. The resulting curve fits are sufficiently accurate and self-contained so that they can be readily incorporated into new or existing computational fluid dynamics codes. The range of validity of the new curve fits are temperatures up to 15,000 K densities from 10 to the -5 to 10 amagats (rho/rho sub o).

Srinivasan, S.↗
NASA NTRS1988

Direct simulations of chemically reacting turbulent mixing layers, part 2

The results of direct numerical simulations of chemically reacting turbulent mixing layers are presented. This is an extension of earlier work to a more detailed study of previous three dimensional simulations of cold reacting flows plus the development, validation, and use of codes to simulate chemically reacting shear layers with heat release. Additional analysis of earlier simulations showed good agreement with self similarity theory and laboratory data. Simulations with a two dimensional code including the effects of heat release showed that the rate of chemical product formation, the thickness of the mixing layer, and the amount of mass entrained into the layer all decrease with increasing rates of heat release. Subsequent three dimensional simulations showed similar behavior, in agreement with laboratory observations. Baroclinic torques and thermal expansion in the mixing layer were found to produce changes in the flame vortex structure that act to diffuse the pairing vortices, resulting in a net reduction in vorticity. Previously unexplained anomalies observed in the mean velocity profiles of reacting jets and mixing layers were shown to result from vorticity generation by baroclinic torques.

Metcalfe, Ralph W.↗
NASA NTRS1988

Navier-Stokes simulation of 3-D hypersonic equilibrium air flow

A computationally efficient three-dimensional conservative supracharacteristic Navier-Stokes method has been extended to simulate complex external chemically reacting flows of hypersonic reentry vehicles at angle-of-attack. Numerical simulation results of the flow around a sphere-cone-cone-flare reentry vehicle at 10 deg angle-of-attack are presented, in addition to the results of a well-validated two-dimensional code with which the 0-deg axisymmetric flow has been computed. A method for obtaining compositions of species in equilibrium ionized air is proposed.

Nagaraj, N.↗
NASA NTRS1988

A comparative study and validation of upwind and central-difference Navier-Stokes codes for high-speed flows

A comparative study was made using 4 different computer codes for solving the compressible Navier-Stokes equations. Three different test problems were used, each of which has features typical of high speed internal flow problems of practical importance in the design and analysis of propulsion systems for advanced hypersonic vehicles. These problems are the supersonic flow between two walls, one of which contains a 10 deg compression ramp, the flow through a hypersonic inlet, and the flow in a 3-D corner formed by the intersection of two symmetric wedges. Three of the computer codes use similar recently developed implicit upwind differencing technology, while the fourth uses a well established explicit method. The computed results were compared with experimental data where available.

Rudy, David H.↗
NASA NTRS1990

Comparison of 3-D viscous flow computations of Mach 5 inlet with experimental data

A time marching 3-D full Navier-Stokes code, called PARC3D, is validated for an experimental Mach 5 inlet configuration using the data obtained in the 10 x 10 ft supersonic wind tunnel at the NASA Lewis Research Center. For the first time, a solution is obtained for this configuration with the actual geometry, the tunnel conditions, and all the bleed zones modeled in the computation. Pitot pressure profiles and static pressures at various locations in the inlet are compared with the corresponding experimental data. The effect of bleed zones, located in different places on the inlet walls, in eliminating the low energy vortical flow generated from the 3-D shock-boundary layer interaction is simulated very well even though some approximations are used in applying the bleed boundary conditions and in the turbulence model. A further detailed study of the effect of individual bleed ports is needed to understand fully the actual mechanism of efficiently eliminating the vortical flow from the inlet. A better turbulence model would help to improve the accuracy even further in predicting the corner flow boundary layer profiles.

Reddy, D. R.↗
NASA NTRS1990

Experimental and computational study of the viscoplastic response of high temperature structures

The first nine months of experimental and computational studies of the thermal-structural behavior of thin panels subjected to localized heating are described. Initial experimental studies have focused on developing an experimental set-up with well-defined thermal-structure boundary conditions. Preliminary tests with a Heldenfels panel have demonstrated out of plane bending (thermal buckling) due to panel initial imperfections. Initial computational studies have focused on: (1) validation of a thermoviscoplastic code to predict thermal stresses in the unbuckled panel, and (2) investigating in-plane stresses for test panels under transient thermal loading. Plans for future research are described.

Thornton, Earl A.↗
NASA NTRS1991

Experimental water droplet impingement data on modern aircraft surfaces

An experimental method has been developed to determine the water droplet impingement characteristics on two- and three-dimensional aircraft surfaces. The experimental water droplet impingement data are used to validate particle trajectory analysis codes that are used in aircraft icing analyses and engine inlet particle separator analyses. The aircraft surface is covered with thin strips of blotter paper in areas of interest. The surface is then exposed to an airstream that contains a dyed-water spray cloud. The water droplet impingement data are extracted from the dyed blotter paper strips by measuring the optical reflectance of each strip with an automated reflectometer. Preliminary experimental and analytical impingement efficiency data are presented for a NLF(1)-0414F airfoil, s swept MS(1)-0317 airfoil, a swept NACA 0012 wingtip and for a Boeing 737-300 engine inlet model.

Papadakis, Michael↗