Peterson, Victor L.; Ballhaus, William F., Jr.
The Numerical Aerodynamic Simulation (NAS) program has reached a milestone with the completion of the initial operating configuration of the NAS Processing System Network. This achievement is the first major milestone in the continuing effort to provide a state-of-the-art supercomputer facility for the national aerospace community and to serve as a pathfinder for the development and use of future supercomputer systems. The underlying factors that motivated the initiation of the program are first identified and then discussed. These include the emergence and evolution of computational aerodynamics as a powerful new capability in aerodynamics research and development, the computer power required for advances in the discipline, the complementary nature of computation and wind tunnel testing, and the need for the government to play a pathfinding role in the development and use of large-scale scientific computing systems. Finally, the history of the NAS program is traced from its inception in 1975 to the present time.
In this article two integral computational fluid dynamics methods for steady-state and transient vehicle aerodynamic simulations are described using a Chevrolet Corvette ZR-1 surface panel model. In the last decade, road-vehicle aerodynamics have become an important design consideration. Originally, the design of low-drag shapes was given high priority due to worldwide fuel shortages that occurred in the mid-seventies. More recently, there has been increased interest in the role aerodynamics play in vehicle stability and passenger safety. Consequently, transient aerodynamics and the aerodynamics of vehicle in yaw have become important issues at the design stage. While there has been tremendous progress in Navier-Stokes methodology in the last few years, the physics of bluff-body aerodynamics are still very difficult to model correctly. Moreover, the computational effort to perform Navier-Stokes simulations from the geometric stage to complete flow solutions requires much computer time and impacts the design cycle time. In the short run, therefore, simpler methods must be used for such complicated problems. Here, two methods are described for the simulation of steady-state and transient vehicle aerodynamics.
Williams, Louis J.; Hessenius, Kristin A.; Corsiglia, Victor R.; Hicks, Gary; Richardson, Pamela F.; Unger, George; Neumann, Benjamin; Moss, Jim
The annual accomplishments is reviewed for the Aerodynamics Division during FY 1991. The program includes both fundamental and applied research directed at the full spectrum of aerospace vehicles, from rotorcraft to planetary entry probes. A comprehensive review is presented of the following aerodynamics elements: computational methods and applications; CFD validation; transition and turbulence physics; numerical aerodynamic simulation; test techniques and instrumentation; configuration aerodynamics; aeroacoustics; aerothermodynamics; hypersonics; subsonics; fighter/attack aircraft and rotorcraft.
Cmaylo, Bohden K.; Foo, Lee
This document is a report of the Numerical Aerodynamic Simulation (NAS) Long Haul Communications Prototype (LHCP). It describes the accomplishments of the LHCP group, presents the results from all LHCP experiments and testing activities, makes recommendations for present and future LHCP activities, and evaluates the remote workstation accesses from Langley Research Center, Lewis Research Center, and Colorado State University to Ames Research Center. The report is the final effort of the Long Haul (Wideband) Communications Prototype Plan (PT-1133-02-N00), 3 October 1985, which defined the requirements for the development, test, and operation of the LHCP network and was the plan used to evaluate the remote user bandwidth requirements for the Numerical Aerodynamic Simulation Processing System Network.
Holmes, Bruce J.; Schairer, Edward; Hicks, Gary; Wander, Stephen; Blankson, Isiaiah; Rose, Raymond; Olson, Lawrence; Unger, George
Presented here is a comprehensive review of the following aerodynamics elements: computational methods and applications, computational fluid dynamics (CFD) validation, transition and turbulence physics, numerical aerodynamic simulation, drag reduction, test techniques and instrumentation, configuration aerodynamics, aeroacoustics, aerothermodynamics, hypersonics, subsonic transport/commuter aviation, fighter/attack aircraft and rotorcraft.
An overview of historical and current numerical aerodynamic simulation (NAS) is given. The capabilities and goals of the Numerical Aerodynamic Simulation Facility are outlined. Emphasis is given to numerical flow visualization and its applications to structural analysis of aircraft and spacecraft bodies. The uses of NAS in computational chemistry, engine design, and galactic evolution are mentioned.
Peterson, V. L.; Ballhaus, W. F., Jr.; Bailey, F. R.
The history of the Numerical Aerodynamic Simulation Program, which is designed to provide a leading-edge capability to computational aerodynamicists, is traced back to its origin in 1975. Factors motivating its development and examples of solutions to successively refined forms of the governing equations are presented. The NAS Processing System Network and each of its eight subsystems are described in terms of function and initial performance goals. A proposed usage allocation policy is discussed and some initial problems being readied for solution on the NAS system are identified.
Given here are selected scientific results from the Numerical Aerodynamic Simulation (NAS) Program's third year of operation. During this year, the scientific community was given access to a Cray-2 and a Cray Y-MP supercomputer. Topics covered include flow field analysis of fighter wing configurations, large-scale ocean modeling, the Space Shuttle flow field, advanced computational fluid dynamics (CFD) codes for rotary-wing airloads and performance prediction, turbulence modeling of separated flows, airloads and acoustics of rotorcraft, vortex-induced nonlinearities on submarines, and standing oblique detonation waves.
NASA has planned a supercomputer for computational fluid dynamics research since the mid-1970's. With the approval of the Numerical Aerodynamic Simulation Program as a FY 1984 new start, Congress requested an assessment of the program's objectives, projected short- and long-term uses, program design, computer architecture, user needs, and handling of proprietary and classified information. Specifically requested was an examination of the merits of proceeding with multiple high speed processor (HSP) systems contrasted with a single high speed processor system. The panel found NASA's objectives and projected uses sound and the projected distribution of users as realistic as possible at this stage. The multiple-HSP, whereby new, more powerful state-of-the-art HSP's would be integrated into a flexible network, was judged to present major advantages over any single HSP system.
NASA created the Numerical Aerodynamic Simulation (NAS) Program in 1987 to focus resources on solving critical problems in aeroscience and related disciplines by utilizing the power of the most advanced supercomputers available. The NAS Program provides scientists with the necessary computing power to solve today's most demanding computational fluid dynamics problems and serves as a pathfinder in integrating leading-edge supercomputing technologies, thus benefitting other supercomputer centers in government and industry. The 1992-93 operational year concluded with 399 high-speed processor projects and 91 parallel projects representing NASA, the Department of Defense, other government agencies, private industry, and universities. This document provides a glimpse at some of the significant scientific results for the year.
NASA created the Numerical Aerodynamic Simulation (NAS) Program in 1987 to focus resources on solving critical problems in aeroscience and related disciplines by utilizing the power of the most advanced supercomputers available. The NAS Program provides scientists with the necessary computing power to solve today's most demanding computational fluid dynamics problems and serves as a pathfinder in integrating leading-edge supercomputing technologies, thus benefiting other supercomputer centers in Government and industry. This report contains selected scientific results from the 1991-92 NAS Operational Year, March 4, 1991 to March 3, 1992, which is the fifth year of operation. During this year, the scientific community was given access to a Cray-2 and a Cray Y-MP. The Cray-2, the first generation supercomputer, has four processors, 256 megawords of central memory, and a total sustained speed of 250 million floating point operations per second. The Cray Y-MP, the second generation supercomputer, has eight processors and a total sustained speed of one billion floating point operations per second. Additional memory was installed this year, doubling capacity from 128 to 256 megawords of solid-state storage-device memory. Because of its higher performance, the Cray Y-MP delivered approximately 77 percent of the total number of supercomputer hours used during this year.
Stromberg, W. J.
An engine was specially prepared with extensive instrumentation to monitor performance, case temperatures, and clearance changes. A special loading device was used to apply known loads on the engine by the use of cables placed around the flight inlet. These loads simulated the estimated aerodynamic pressure distributions that occur on the inlet in various segments of a typical airplane flight. Test results indicate that the engine lost 1.3 percent in take-off thrust specific fuel consumption (TSFC) during the course of the test effort. Permanent clearance changes due to the loads accounted for 1.1 percent; increase in low pressure compressor airfoil roughness and thermal distortion in the high pressure turbine accounted for 0.2 percent. Pretest predicted performance loss due to clearance changes was 0.9 percent in TSFC. Therefore, the agreement between measurement and prediction is considered to be excellent.
Bragg, Michael B.; Broeren, Andy P.; Addy, Harold E.; Potapczuk, Mark G.; Guffond, Didier; Montreuil, E.
NASA Glenn Research Center, ONERA, and the University of Illinois are conducting a major research program whose goal is to improve our understanding of the aerodynamic scaling of ice accretions on airfoils. The program when it is completed will result in validated scaled simulation methods that produce the essential aerodynamic features of the full-scale iced-airfoil. This research will provide some of the first, high-fidelity, full-scale, iced-airfoil aerodynamic data. An initial study classified ice accretions based on their aerodynamics into four types: roughness, streamwise ice, horn ice, and spanwise-ridge ice. Subscale testing using a NACA 23012 airfoil was performed in the NASA IRT and University of Illinois wind tunnel to better understand the aerodynamics of these ice types and to test various levels of ice simulation fidelity. These studies are briefly reviewed here and have been presented in more detail in other papers. Based on these results, full-scale testing at the ONERA F1 tunnel using cast ice shapes obtained from molds taken in the IRT will provide full-scale iced airfoil data from full-scale ice accretions. Using these data as a baseline, the final step is to validate the simulation methods in scale in the Illinois wind tunnel. Computational ice accretion methods including LEWICE and ONICE have been used to guide the experiments and are briefly described and results shown. When full-scale and simulation aerodynamic results are available, these data will be used to further develop computational tools. Thus the purpose of the paper is to present an overview of the program and key results to date.
The role of Ames Research Center in conducting basic aerodynamics research through computer simulations is described. The computer facilities, including supercomputers and peripheral equipment that represent the state of the art, are described. The methodology of computational fluid dynamics is explained briefly. Fundamental studies of turbulence and transition are being pursued to understand these phenomena and to develop models that can be used in the solution of the Reynolds-averaged Navier-Stokes equations. Four applications of computer simulations for aerodynamics problems are described: subsonic flow around a fuselage at high angle of attack, subsonic flow through a turbine stator-rotor stage, transonic flow around a flexible swept wing, and transonic flow around a wing-body configuration that includes an inlet and a tail.
Baize, Daniel G. (Editor)
The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executive summaries for all the Aerodynamic Performance technology areas.
Potapczuk, M. G.; Bragg, M. B.; Kwon, O. J.; Sankar, L. N.
The sectional and total aerodynamic load characteristics of moderate aspect ratio wings with and without simulated glaze leading edge ice were studied both computationally, using a three dimensional, compressible Navier-Stokes solver, and experimentally. The wing has an untwisted, untapered planform shape with NACA 0012 airfoil section. The wing has an unswept and swept configuration with aspect ratios of 4.06 and 5.0. Comparisons of computed surface pressures and sectional loads with experimental data for identical configurations are given. The abrupt decrease in stall angle of attack for the wing, as a result of the leading edge ice formation, was demonstrated numerically and experimentally.
Hahne, David E. (Editor)
NASA's High-Speed Research Program sponsored the 1999 Aerodynamic Performance Technical Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in the areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High Lift, and Flight Controls. The review objectives were to: (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working on HSCT aerodynamics. In particular, single and midpoint optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented, along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program. This Volume 1/Part 1 publication covers configuration aerodynamics.
Baize, Daniel G. (Editor)
The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in area of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodyamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executive summaries for all the Aerodynamic Performance technology areas.
McMillin, S. Naomi (Editor)
NASA's High-Speed Research Program sponsored the 1998 Aerodynamic Performance Technical Review on February 9-13, in Los Angeles, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, and Flight Controls. The review objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working HSCT aerodynamics. In particular, single and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program.
Hahne, David E. (Editor)
NASA's High-Speed Research Program sponsored the 1999 Aerodynamic Performance Technical Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in the areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High Lift, and Flight Controls. The review objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working on HSCT aerodynamics. In particular, single and midpoint optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented, along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program. This Volume 1/Part 2 publication covers the design optimization and testing sessions.
McMillin, S. Naomi (Editor)
NASA's High-Speed Research Program sponsored the 1998 Aerodynamic Performance Technical Review on February 9-13, in Los Angeles, California. The review was designed to bring together NASA and industry HighSpeed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of. Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, and Flight Controls. The review objectives were to: (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working HSCT aerodynamics. In particular, single and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program.
Hair, L. M.
Technology for simulating plumes in wind tunnel tests was not adequate to provide the required confidence in test data where plume induced aerodynamic effects might be significant. A broad research program was undertaken to correct the deficiency. Four tasks within the program are reported. Three of these tasks involve conducting experiments, related to three different aspects of the plume simulation problem: (1) base pressures; (2) lateral jet pressures; and (3) plume parameters. The fourth task involves collecting all of the base pressure test data generated during the program. Base pressures were measured on a classic cone ogive cylinder body as affected by the coaxial, high temperature exhaust plumes of a variety of solid propellant rockets. Valid data were obtained at supersonic freestream conditions but not at transonic. Pressure data related to lateral (separation) jets at M infinity = 4.5, for multiple clustered nozzles canted to the freestream and operating at high dynamic pressure ratios. All program goals were met although the model hardware was found to be large relative to the wind tunnel size so that operation was limited for some nozzle configurations.
There were three major issues examined in the feasibility study. First, the ability of the proposed system architecture to support the anticipated workload was evaluated. Second, the throughput of the computational engine (the flow model processor) was studied using real application programs. Third, the availability reliability, and maintainability of the system were modeled. The evaluations were based on the baseline systems. The results show that the implementation of the Numerical Aerodynamic Simulation Facility, in the form considered, would indeed be a feasible project with an acceptable level of risk. The technology required (both hardware and software) either already exists or, in the case of a few parts, is expected to be announced this year. Facets of the work described include the hardware configuration, software, user language, and fault tolerance.
Broeren, Andy P.; Addy, Harold E., Jr.; Bragg, Michael B.; Busch, Greg T.; Montreuil, Emmanuel
This report describes recent improvements in aerodynamic scaling and simulation of ice accretion on airfoils. Ice accretions were classified into four types on the basis of aerodynamic effects: roughness, horn, streamwise, and spanwise ridge. The NASA Icing Research Tunnel (IRT) was used to generate ice accretions within these four types using both subscale and full-scale models. Large-scale, pressurized windtunnel testing was performed using a 72-in.- (1.83-m-) chord, NACA 23012 airfoil model with high-fidelity, three-dimensional castings of the IRT ice accretions. Performance data were recorded over Reynolds numbers from 4.5 x 10(exp 6) to 15.9 x 10(exp 6) and Mach numbers from 0.10 to 0.28. Lower fidelity ice-accretion simulation methods were developed and tested on an 18-in.- (0.46-m-) chord NACA 23012 airfoil model in a small-scale wind tunnel at a lower Reynolds number. The aerodynamic accuracy of the lower fidelity, subscale ice simulations was validated against the full-scale results for a factor of 4 reduction in model scale and a factor of 8 reduction in Reynolds number. This research has defined the level of geometric fidelity required for artificial ice shapes to yield aerodynamic performance results to within a known level of uncertainty and has culminated in a proposed methodology for subscale iced-airfoil aerodynamic simulation.
Graves, R.A. Jr.
A projection is made of likely improvements in the economics of commercial aircraft operation due to developments in aerodynamics in the next half-century. Notable among these improvements are active laminar flow control techniques' application to third-generation SSTs, in order to achieve an L/D value of about 20; this is comparable to current subsonic transports, and has the further consequence of reducing cabin noise. Wave-cancellation systems may also be used to eliminate sonic boom overpressures, and rapid-combustion systems may be able to eliminate all pollutants from jet exhausts other than CO/sub 2/.
Cavin, R. K., III; Colunga, D.
A manual is presented for correctly submitting program runs in aerodynamics on the UNIVAC 1108 computer system. All major program modules are included. Control cards are documented for the user's convenience, and card parameters are included in order to provide some idea as to reasonable time estimates for the program modules.
De Leon, Nelson; De Leon, Matthew N.
We develop a two-dimensional flight simulator for lightweight (less than 10 g) indoor planes. The simulator consists of four coupled time differential equations describing the plane CG, plane pitch and motor. The equations are integrated numerically with appropriate parameters and initial conditions for two planes: (1) Science Olympiad and (2)…
Broeren, Andy P.; Whalen, Edward A.; Busch, Greg T.; Bragg, Michael B.
This report presents the results of recent investigations into the aerodynamics of simulated runback ice accretion on airfoils. Aerodynamic tests were performed on a full-scale model using a high-fidelity, ice-casting simulation at near-flight Reynolds (Re) number. The ice-casting simulation was attached to the leading edge of a 72-in. (1828.8-mm ) chord NACA 23012 airfoil model. Aerodynamic performance tests were conducted at the ONERA F1 pressurized wind tunnel over a Reynolds number range of 4.7?10(exp 6) to 16.0?10(exp 6) and a Mach (M) number ran ge of 0.10 to 0.28. For Re = 16.0?10(exp 6) and M = 0.20, the simulated runback ice accretion on the airfoil decreased the maximum lift coe fficient from 1.82 to 1.51 and decreased the stalling angle of attack from 18.1deg to 15.0deg. The pitching-moment slope was also increased and the drag coefficient was increased by more than a factor of two. In general, the performance effects were insensitive to Reynolds numb er and Mach number changes over the range tested. Follow-on, subscale aerodynamic tests were conducted on a quarter-scale NACA 23012 model (18-in. (457.2-mm) chord) at Re = 1.8?10(exp 6) and M = 0.18, using low-fidelity, geometrically scaled simulations of the full-scale castin g. It was found that simple, two-dimensional simulations of the upper- and lower-surface runback ridges provided the best representation of the full-scale, high Reynolds number iced-airfoil aerodynamics, whereas higher-fidelity simulations resulted in larger performance degrada tions. The experimental results were used to define a new subclassification of spanwise ridge ice that distinguishes between short and tall ridges. This subclassification is based upon the flow field and resulting aerodynamic characteristics, regardless of the physical size of the ridge and the ice-accretion mechanism.
Robinson, Philip E.; Thompson, James
The Application Programming Interface (API) for the Crew Exploration Vehicle (CEV) Aerodynamic Database has been developed to provide the developers of software an easily implemented, fully self-contained method of accessing the CEV Aerodynamic Database for use in their analysis and simulation tools. The API is programmed in C and provides a series of functions to interact with the database, such as initialization, selecting various options, and calculating the aerodynamic data. No special functions (file read/write, table lookup) are required on the host system other than those included with a standard ANSI C installation. It reads one or more files of aero data tables. Previous releases of aerodynamic databases for space vehicles have only included data tables and a document of the algorithm and equations to combine them for the total aerodynamic forces and moments. This process required each software tool to have a unique implementation of the database code. Errors or omissions in the documentation, or errors in the implementation, led to a lengthy and burdensome process of having to debug each instance of the code. Additionally, input file formats differ for each space vehicle simulation tool, requiring the aero database tables to be reformatted to meet the tool s input file structure requirements. Finally, the capabilities for built-in table lookup routines vary for each simulation tool. Implementation of a new database may require an update to and verification of the table lookup routines. This may be required if the number of dimensions of a data table exceeds the capability of the simulation tools built-in lookup routines. A single software solution was created to provide an aerodynamics software model that could be integrated into other simulation and analysis tools. The highly complex Orion aerodynamics model can then be quickly included in a wide variety of tools. The API code is written in ANSI C for ease of portability to a wide variety of systems. The
Cronvich, L. L.; Barnes, G. A.
The Bumblebee program, was designed to provide a supersonic guided missile. The aerodynamics program included a fundamental research effort in supersonic aerodynamics as well as a design task in developing both test vehicles and prototypes of tactical missiles. An index of aerodynamic missile data developed in this program is presented.
Herraez, I.; Medjroubi, W.; Stoevesandt, B.; Peinke, J.
CFD (Computational Fluid Dynamics) simulations are a very promising method for predicting the aerodynamic behavior of wind turbines in an inexpensive and accurate way. One of the major drawbacks of this method is the lack of validated models. As a consequence, the reliability of numerical results is often difficult to assess. The MEXICO project aimed at solving this problem by providing the project partners with high quality measurements of a 4.5 meters rotor diameter wind turbine operating under controlled conditions. The large measurement data-set allows the validation of all kind of aerodynamic models. This work summarizes our efforts for validating a CFD model based on the open source software OpenFoam. Both steady- state and time-accurate simulations have been performed with the Spalart-Allmaras turbulence model for several operating conditions. In this paper we will concentrate on axisymmetric inflow for 3 different wind speeds. The numerical results are compared with pressure distributions from several blade sections and PIV-flow data from the near wake region. In general, a reasonable agreement between measurements the and our simulations exists. Some discrepancies, which require further research, are also discussed.
Loomis, Mark P.; Venkatapathy, Ethiraj; Papadopoulos, Periklis; Davies, Carol B.; Berry, Scott; Horvath, Tom; Campbell, Charles
This paper presents the results of a computational fluid dynamics validation exercise using the General Aerodynamic Simulation Program (GASP) code for vehicles under consideration for the X-38 program. The scope of the exercise involved simulation and comparison with wind tunnel aeroheating and aerodynamic data in the hypersonic regime. In addition, simulations were performed and results are presented for the full-scale flight vehicle along an expected re- entry trajectory.
Messina, Michael D.
The method described in this report is intended to present an overview of a process developed to extract the forebody aerodynamic increments from flight tests. The process to determine the aerodynamic increments (rolling pitching, and yawing moments, Cl, Cm, Cn, respectively) for the forebody strake controllers added to the F/A - 18 High Alpha Research Vehicle (HARV) aircraft was developed to validate the forebody strake aerodynamic model used in simulation.
Baize, Daniel G. (Editor)
The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag, prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executives summaries for all the Aerodynamic Performance technology areas.
Blackwell, K. L.; Hair, L. M.
A series of parametric wind tunnel tests was conducted to provide a base for developing a simulation of afterbody/base aerodynamics for multibody/multibase rocket-powered vehicles (such as Space Shuttle) which use unheated air as the simulant gas in development wind tunnel tests. The tests described were parameterized on external configuration, nozzle internal configuration, base geometry, propulsion gas type, and freestream Mach number (0.5 to 3.5). The tests were conducted over a 4-year period. Presented in this report are the data and pertinent reference information necessary to perform an analysis which would lead to a simulation procedure. The type of data obtained during the tests described herein include model base afterbody, and nozzle internal surface static pressure distributions, model chamber pressure and temperature, and freestream conditions. Also included is a brief description of simulation procedures that were used by the Space Shuttle program.
Lingard, J. S.; Underwood, J. C.; Darley, M. G.; Marraffa, L.; Ferracina, L.
The ESA Supersonic Parachute program expands the knowledge of parachute inflation and flying characteristics in supersonic flows using wind tunnel testing and fluid structure interaction to develop new inflation algorithms and aerodynamic databases.
Hahne, David E. (Editor)
NASA's High-Speed Research Program sponsored the 1999 Aerodynamic Performance Technical Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in the areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High Lift, and Flight Controls. The review objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among die scientists and engineers working on HSCT aerodynamics. In particular, single and midpoint optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented, along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program. This Volume 2/Part 2 publication covers the tools and methods development session.
McMillin, S. Naomi (Editor)
NASA's High-Speed Research Program sponsored the 1998 Aerodynamic Performance Technical Review on February 9-13, in Los Angeles, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, and Flight Controls. The review objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program.
There were three major issues examined in the feasibility study. First, the ability of the proposed system architecture to support the anticipated workload was evaluated. Second, the throughput of the computational engine (the flow model processor) was studied using real application programs. Third, the availability, reliability, and maintainability of the system were modeled. The evaluations were based on the baseline systems. The results show that the implementation of the Numerical Aerodynamic Simulation Facility, in the form considered, would indeed be a feasible project with an acceptable level of risk. The technology required (both hardware and software) either already exists or, in the case of a few parts, is expected to be announced this year.
Lincoln, N. R.; Bergman, R. O.; Bonstrom, D. B.; Brinkman, T. W.; Chiu, S. H. J.; Green, S. S.; Hansen, S. D.; Klein, D. L.; Krohn, H. E.; Prow, R. P.
A Numerical Aerodynamic Simulation Facility (NASF) was designed for the simulation of fluid flow around three-dimensional bodies, both in wind tunnel environments and in free space. The application of numerical simulation to this field of endeavor promised to yield economies in aerodynamic and aircraft body designs. A model for a NASF/FMP (Flow Model Processor) ensemble using a possible approach to meeting NASF goals is presented. The computer hardware and software are presented, along with the entire design and performance analysis and evaluation.
The production of an optimized design of key elements of the candidate facility was the primary objective of this report. This was accomplished by effort in the following tasks: (1) to further develop, optimize and describe the function description of the custom hardware; (2) to delineate trade off areas between performance, reliability, availability, serviceability, and programmability; (3) to develop metrics and models for validation of the candidate systems performance; (4) to conduct a functional simulation of the system design; (5) to perform a reliability analysis of the system design; and (6) to develop the software specifications to include a user level high level programming language, a correspondence between the programming language and instruction set and outline the operation system requirements.
Kutler, P.; Mehta, U. B.; Andrews, A.
The concept of artificial intelligence as it applies to computational fluid dynamics simulation is investigated. How expert systems can be adapted to speed the numerical aerodynamic simulation process is also examined. A proposed expert grid generation system is briefly described which, given flow parameters, configuration geometry, and simulation constraints, uses knowledge about the discretization process to determine grid point coordinates, computational surface information, and zonal interface parameters.
Bailey, F. R.; Ballhaus, W. F.
The current and projected use of advanced computers for large-scale aerodynamic flow simulation applied to engineering design and research is discussed. The design use of mature codes run on conventional, serial computers is compared with the fluid research use of new codes run on parallel and vector computers. The role of flow simulations in design is illustrated by the application of a three dimensional, inviscid, transonic code to the Sabreliner 60 wing redesign. Research computations that include a more complete description of the fluid physics by use of Reynolds averaged Navier-Stokes and large-eddy simulation formulations are also presented. Results of studies for a numerical aerodynamic simulation facility are used to project the feasibility of design applications employing these more advanced three dimensional viscous flow simulations.
A computing system was designed with the capability of providing an effective throughput of one billion floating point operations per second for three dimensional Navier-Stokes codes. The methodology used in defining the baseline design, and the major elements of the numerical aerodynamic simulation facility are described.
Benson, Thomas J.
A workstation-based interactive flow simulator has been developed to aid in the teaching of undergraduate compressible aerodynamics. By solving the equations found in NACA 1135, the simulator models three basic fluids problems encountered in supersonic flow: flow past a compression corner, flow past two wedges in series, and flow past two opposed wedges. The study can vary the geometry or flow conditions through a graphical user interface and the new conditions are calculated immediately. Various graphical formats present the results of the flow calculations to the student. The simulator includes interactive questions and answers to aid in both the use of the tool and to develop an understanding of some of the complexities of compressible aerodynamics. A series of help screens make the simulator easy to learn and use.
Tweedt, Daniel L.
Quieter working environments for astronauts are needed if future long-duration space exploration missions are to be safe and productive. Ventilation and payload cooling fans are known to be dominant sources of noise, with the International Space Station being a good case in point. To address this issue cost effectively, early attention to fan design, selection, and installation has been recommended, leading to an effort by NASA to examine the potential for small-fan noise reduction by improving fan aerodynamic design. As a preliminary part of that effort, the aerodynamics of a cabin ventilation fan designed by Hamilton Sundstrand has been simulated using computational fluid dynamics codes, and the computed solutions analyzed to quantify various aspects of the fan aerodynamics and performance. Four simulations were performed at the design rotational speed: two at the design flow rate and two at off-design flow rates. Following a brief discussion of the computational codes, various aerodynamic- and performance-related quantities derived from the computed flow fields are presented along with relevant flow field details. The results show that the computed fan performance is in generally good agreement with stated design goals.
Moss, James N.; Glass, Christopher E.; Greene, Francis A.
Direct simulation Monte Carlo DSMC simulations are performed for the Apollo capsule in the hypersonic low density transitional flow regime. The focus is on ow conditions similar to that experienced by the Apollo Command Module during the high altitude portion of its reentry Results for aerodynamic forces and moments are presented that demonstrate their sensitivity to rarefaction that is for free molecular to continuum conditions. Also aerodynamic data are presented that shows their sensitivity to a range of reentry velocity encompasing conditions that include reentry from low Earth orbit lunar return and Mars return velocities to km/s. The rarefied results are anchored in the continuum regime with data from Navier Stokes simulations
Medan, R. T.
A computer program that provides the geometry and boundary conditions appropriate for an analysis of a lifting, thin wing with control surfaces in linearized, subsonic, steady flow is presented. The kernel function method lifting surface theory is applied. The data which is generated by the program is stored on disk files or tapes for later use by programs which calculate an influence matrix, plot the wing planform, and evaluate the loads on the wing. In addition to processing data for subsequent use in a lifting surface analysis, the program is useful for computing area and mean geometric chords of the wing and control surfaces.
Chapman, D. R.; Mark, H.; Pirtle, M. W.
It is pointed out that in other fields of computational physics, such as ballistics, celestial mechanics, and neutronics, computations have already displaced experiments as the principal means of obtaining dynamic simulations. In the case of aerodynamic investigations, the complexity of the computational work involved in solving the Navier-Stokes equations is the reason that such investigations rely currently mainly on wind-tunnel testing. However, because of inherent limitations of the wind-tunnel approach and economic considerations, it appears that at some time in the future aerodynamic studies will chiefly rely on computational flow data provided by the computer. Taking into account projected development trends, it is estimated that computers with the required capabilities for a solution of the complete viscous, time-dependent Navier-Stokes equations will be available in the mid-1980s.
Axelson, J. A.
The AEROX program estimates lift, induced-drag and pitching moments to high angles (typ. 60 deg) for wings and for wingbody combinations with or without an aft horizontal tail. Minimum drag coefficients are not estimated, but may be input for inclusion in the total aerodynamic parameters which are output in listed and plotted formats. The theory, users' guide, test cases, and program listing are presented.
Dillenius, M. F. E.; Hemsch, M. J.; Sawyer, W. C.; Allen, J. M.; Blair, A. B., Jr.
Two different classes of missile aeroprediction programs have been recently developed. The first class of programs provides rapid engineering predictions and includes MISSILE1 and MISSILE2 applicable to missile configurations with axisymmetric bodies. The second class of programs consists of the DEMON series, including a simplified version NSWCDM, designed to calculate detailed loadings acting on supersonic missiles which may have non-circular body cross sections. Both classes account for high angles of attack and track vortices from canard or wing section to the tail section. Extensive comparisons with experimental data are presented including nonlinear effects of canard control.
Zorumski, W. E. (Editor); Weir, D. S. (Editor)
The prediction sequence used in the aircraft noise prediction program (ANOPP) is described. The elements of the sequence are called program modules. The first group of modules analyzes the propeller geometry, the aerodynamics, including both potential and boundary-layer flow, the propeller performance, and the surface loading distribution. This group of modules is based entirely on aerodynamic strip theory. The next group of modules deals with the first group. Predictions of periodic thickness and loading noise are determined with time-domain methods. Broadband noise is predicted by a semiempirical method. Near-field predictions of fuselage surface pressrues include the effects of boundary layer refraction and scattering. Far-field predictions include atmospheric and ground effects.
Papadopoulos, Periklis E.; Prabhu, Dinesh; Wright, Michael; Davies, Carol; McDaniel, Ryan; Venkatapathy, E.; Wercinski, Paul; Gomez, R. J.
Modern Computational Fluid Dynamics (CFD) techniques were used to compute aerodynamic forces and moments of the Space Shuttle Orbiter in specific portions of contingency abort trajectory space. The trajectory space covers a Mach number range of 3.5-15, an angle-of-attack range of 20deg-60deg, an altitude range of 100-190 kft, and several different settings of the control surfaces (elevons, body flap, and speed brake). Presented here are details of the methodology and comparisons of computed aerodynamic coefficients against the values in the current Orbiter Operational Aerodynamic Data Book (OADB). While approximately 40 cases have been computed, only a sampling of the results is provided here. The computed results, in general, are in good agreement with the OADB data (i.e., within the uncertainty bands) for almost all the cases. However, in a limited number of high angle-of-attack cases (at Mach 15), there are significant differences between the computed results, especially the vehicle pitching moment, and the OADB data. A preliminary analysis of the data from the CFD simulations at Mach 15 shows that these differences can be attributed to real-gas/Mach number effects. The aerodynamic coefficients and detailed surface pressure distributions of the present simulations are being used by the Shuttle Program in the evaluation of the capabilities of the Orbiter in contingency abort scenarios.
Boyer, Henry; Sigurdson, Lorenz; Lange, Carlos
A wide range of trailers with very poor aerodynamics are hauled long distances across a vast North American highway system. Our goal was to use preliminary smoke-wire flow visualizations to learn: the characteristic flow patterns over models representing modern Vehicle-Trailer Systems (VTS); what improvements need to be made in the experimental set-up; and if there is an opportunity for reduction in aerodynamic drag. Visualization tests were done in an open circuit wind tunnel, with a cross-sectional area of 0.3 m2. Detailed models of light duty trucks and trailers were used at a Reynolds number of 13,700. Images of the streaklines indicated two characteristic features. One was the presence of a stagnation point on the leading face of the trailer followed by a separation bubble on its top. The other feature was an unexpected separation bubble on the hood of the towing vehicle. We determined that it did not have a significant effect on the downstream flow pattern. By adding a small wedge deflector on the cab of the vehicle it was concluded that there is an opportunity for significant improvement of the VTS aerodynamics. Computational simulation of the flow is underway. Support from the Natural Sciences and Engineering Research Council of Canada Discovery Grant Number 41747 is gratefully acknowledged.
Aftosmis, Michael J.; Nemec, Marian
This work examines the level of discretization error in simulation-based aerodynamic databases and introduces strategies for error control. Simulations are performed using a parallel, multi-level Euler solver on embedded-boundary Cartesian meshes. Discretization errors in user-selected outputs are estimated using the method of adjoint-weighted residuals and we use adaptive mesh refinement to reduce these errors to specified tolerances. Using this framework, we examine the behavior of discretization error throughout a token database computed for a NACA 0012 airfoil consisting of 120 cases. We compare the cost and accuracy of two approaches for aerodynamic database generation. In the first approach, mesh adaptation is used to compute all cases in the database to a prescribed level of accuracy. The second approach conducts all simulations using the same computational mesh without adaptation. We quantitatively assess the error landscape and computational costs in both databases. This investigation highlights sensitivities of the database under a variety of conditions. The presence of transonic shocks or the stiffness in the governing equations near the incompressible limit are shown to dramatically increase discretization error requiring additional mesh resolution to control. Results show that such pathologies lead to error levels that vary by over factor of 40 when using a fixed mesh throughout the database. Alternatively, controlling this sensitivity through mesh adaptation leads to mesh sizes which span two orders of magnitude. We propose strategies to minimize simulation cost in sensitive regions and discuss the role of error-estimation in database quality.
A technology forecast was established for the 1980-1985 time frame and the appropriateness of various logic and memory technologies for the design of the numerical aerodynamic simulation facility was assessed. Flow models and their characteristics were analyzed and matched against candidate processor architecture. Metrics were established for the total facility, and housing and support requirements of the facility were identified. An overview of the system is presented, with emphasis on the hardware of the Navier-Stokes solver, which is the key element of the system. Software elements of the system are also discussed.
Giffin, R. G.; Parker, D. E.; Dunbar, L. W.
This report presents the aerodynamic component test results of Fan C, a high-bypass-ratio, low-aerodynamic-loading, 1550 feet per second (472.4 m/sec), single-stage fan, which was designed and tested as part of the NASA Experimental Quiet Engine Program. The fan was designed to deliver a bypass pressure ratio of 1.60 with an adiabatic efficiency of 84.2 percent at a total fan flow of 915 lb/sec (415.0 kg/sec). It was tested with and without inlet distortion. A bypass total-pressure ratio of 1.61 and an adiabatic efficiency of 83.9 percent at a total fan flow of 921 lb/sec (417.8 kg/sec) were actually achieved. An operating margin in excess of 14.6 percent was demonstrated at design speed.
Kennelly, Robert A., Jr.; Bencze, Daniel P. (Technical Monitor)
Two constrained shape optimization problems relevant to aerodynamics are solved by genetic programming, in which a population of computer programs evolves automatically under pressure of fitness-driven reproduction and genetic crossover. Known optimal solutions are recovered using a small, naive set of elementary operations. Effectiveness is improved through use of automatically defined functions, especially when one of them is capable of a variable number of iterations, even though the test problems lack obvious exploitable regularities. An attempt at evolving new elementary operations was only partially successful.
Carlson, H. W.
The Supersonic Wing Nonlinear Aerodynamics computer program, LTSTAR, was developed to provide for the estimation of the nonlinear aerodynamic characteristics of a wing at supersonic speeds. This corrected linearized-theory method accounts for nonlinearities in the variation of basic pressure loadings with local surface slopes, predicts the degree of attainment of theoretical leading-edge thrust forces, and provides an estimate of detached leading-edge vortex loadings that result when the theoretical thrust forces are not fully realized. Comparisons of LTSTAR computations with experimental results show significant improvements in detailed wing pressure distributions, particularly for large angles of attack and for regions of the wing where the flow is highly three-dimensional. The program provides generally improved predictions of the wing overall force and moment coefficients. LTSTAR could be useful in design studies aimed at aerodynamic performance optimization and for providing more realistic trade-off information for selection of wing planform geometry and airfoil section parameters. Input to the LTSTAR program includes wing planform data, freestream conditions, wing camber, wing thickness, scaling options, and output options. Output includes pressure coefficients along each chord, section normal and axial force coefficients, and the spanwise distribution of section force coefficients. With the chordwise distributions and section coefficients at each angle of attack, three sets of polars are output. The first set is for linearized theory with and without full leading-edge thrust, the second set includes nonlinear corrections, and the third includes estimates of attainable leading-edge thrust and vortex increments along with the nonlinear corrections. The LTSTAR program is written in FORTRAN IV for batch execution and has been implemented on a CDC 6000 series computer with a central memory requirement of approximately 150K (octal) of 60 bit words. The LTSTAR
Divan, P.; Dunn, K.; Kojima, J.
A comprehensive aerodynamic analysis program based on linearized potential theory is described. The solution treats thickness and attitude problems at subsonic and supersonic speeds. Three dimensional configurations with or without jet flaps having multiple nonplanar surfaces of arbitrary planform and open or closed slender bodies or noncircular contour are analyzed. Longitudinal and lateral-directional static and rotary derivative solutions are generated. The analysis is implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis problem. Nominal case computation time of 45 CPU seconds on the CDC 175 for a 200 panel simulation indicates the program provides an efficient analysis for systematically performing various aerodynamic configuration tradeoff and evaluation studies.
Bezos, Gaudy M.; Cambell, Bryan A.; Melson, W. Edward
A research technique to obtain large-scale aerodynamic data in a simulated natural rain environment has been developed. A 10-ft chord NACA 64-210 wing section wing section equipped with leading-edge and trailing-edge high-lift devices was tested as part of a program to determine the effect of highly-concentrated, short-duration rainfall on airplane performance. Preliminary dry aerodynamic data are presented for the high-lift configuration at a velocity of 100 knots and an angle of attack of 18 deg. Also, data are presented on rainfield uniformity and rainfall concentration intensity levels obtained during the calibration of the rain simulation system.
Papadopoulos, Periklis; Prabhu, Dinesh; Wright, Michael; Davies, Carol; McDaniel, Ryan; Venkatapathy, Ethiraj; Wersinski, Paul; Gomez, Reynaldo; Arnold, Jim (Technical Monitor)
Modern Computational Fluid Dynamics (CFD) techniques were used to compute aerodynamic forces and moments of the Space Shuttle Orbiter in specific portions of contingency abort trajectory space. The trajectory space covers a Mach number range of 3.5-15, an angle-of-attack range of 20-60 degrees, an altitude range of 100-190 kft, and several different settings of the control surfaces (elevons, body flap, and speed brake). While approximately 40 cases have been computed, only a sampling of the results is presented here. The computed results, in general, are in good agreement with the Orbiter Operational Aerodynamic Data Book (OADB) data (i.e., within the uncertainty bands) for almost all the cases. However, in a limited number of high angle-of-attack cases (at Mach 15), there are significant differences between the computed results, especially the vehicle pitching moment, and the OADB data. A preliminary analysis of the data from the CFD simulations at Mach 15 shows that these differences can be attributed to real-gas/Mach number effects.
Salas, Manuel D.
The research program of the aerodynamics, aerothermodynamics and plasmadynamics discipline of NASA's Hypersonic Project is reviewed. Details are provided for each of its three components: 1) development of physics-based models of non-equilibrium chemistry, surface catalytic effects, turbulence, transition and radiation; 2) development of advanced simulation tools to enable increased spatial and time accuracy, increased geometrical complexity, grid adaptation, increased physical-processes complexity, uncertainty quantification and error control; and 3) establishment of experimental databases from ground and flight experiments to develop better understanding of high-speed flows and to provide data to validate and guide the development of simulation tools.
Pamadi, Bandu N.; Brauckmann, Gregory J.; Ruth, Michael J.; Fuhrmann, Henri D.
An overview of the aerodynamic characteristics, development of the preflight aerodynamic database and flight simulation of the NASA/Orbital X-34 vehicle is presented in this paper. To develop the aerodynamic database, wind tunnel tests from subsonic to hypersonic Mach numbers including ground effect tests at low subsonic speeds were conducted in various facilities at the NASA Langley Research Center. Where wind tunnel test data was not available, engineering level analysis is used to fill the gaps in the database. Using this aerodynamic data, simulations have been performed for typical design reference missions of the X-34 vehicle.
Lawrence, C.; Reddy, T. S. R.; Spyropoulos, E.
The multiphysics code Spectrum(TM) is applied to calculate the unsteady aerodynamic pressures of oscillating cascade of airfoils representing a blade row of a turbomachinery component. Multiphysics simulation is based on a single computational framework for the modeling of multiple interacting physical phenomena, in the present case being between fluids and structures. Interaction constraints are enforced in a fully coupled manner using the augmented-Lagrangian method. The arbitrary Lagrangian-Eulerian method is utilized to account for deformable fluid domains resulting from blade motions. Unsteady pressures are calculated for a cascade designated as the tenth standard, and undergoing plunging and pitching oscillations. The predicted unsteady pressures are compared with those obtained from an unsteady Euler co-de refer-red in the literature. The Spectrum(TM) code predictions showed good correlation for the cases considered.
Kitchar, A. F.; Steuffen, R. L.
An aerodynamic thermal simulation system (ATSS) is presented. The construction of the system, the maintenance, set up, and operations are reported. System description of the radiant array is given along with the array subsystems modular heating unit, adjustable stanchion frame, cooling water system, and the gaseous nitrogen cooling system. The array configuration procedure outlines the set up and the start up. Maintenance procedures involve both lamp maintenance and reflector maintenance along with weather protection. Drawing codes are included. Descriptions of the 36 zone ATSS controls are also given. Each zone is an independent closed loop temperature control circuit. Procedures are presented for starting and stopping the system. The DATA-TRAK programmer and its operation, instruction manuals for the temperature controller and the power regulator and discriptions of peripheral equipment are discussed.
Dayman, B., Jr.; Fiore, A. W.
The present work discusses in general terms the various kinds of ground facilities, in particular, wind tunnels, which support aerodynamic testing. Since not all flight parameters can be simulated simultaneously, an important problem consists in matching parameters. It is pointed out that there is a lack of wind tunnels for a complete Reynolds-number simulation. Using a computer to simulate flow fields can result in considerable reduction of wind-tunnel hours required to develop a given flight vehicle.
Neuman, F.; Warner, N.
The set of programs described has been used for rapidly introducing, checking out and very efficiently using aerodynamic tables in complex aircraft simulations on the IBM 360. The preprocessor program reads in tables with different names and dimensions and stores then on disc storage according to the specified dimensions. The tables are read in from IBM cards in a format which is convenient to reduce the data from the original graphs. During table processing, new auxiliary tables are generated which are required for table cataloging and for efficient interpolation. In addition, DIMENSION statements for the tables as well as READ statements are punched so that they may be used in other programs for readout of the data from disc without chance of programming errors. A quick data checking graphical output for all tables is provided in a separate program.
A robust technique for performing numerical simulations of nonlinear unsteady aeroelastic behavior is developed. The technique is applied to long-span bridges and the wing of a modern business jet. The heart of the procedure is combining the aerodynamic and structural models. The aerodynamic model is a general unsteady vortex-lattice method. The structural model for the bridges is a rigid roadbed supported by linear and torsional springs. For the aircraft wing, the structural model is a cantilever beam with rigid masses attached at various positions along the span; it was generated with the NASTRAN program. The structure, flowing air, and control devices are considered to be the elements of a single dynamic system. All the governing equations are integrated simultaneously and interactively in the time domain; a predictor-corrector method was adapted to perform this integration. For long-span bridges, the simulation predicts the onset of flutter accurately, and the numerical results strongly suggest that an actively controlled wing attached below the roadbed can easily suppress the wind-excited oscillations. The governing equations for a proposed passive system were developed. The wing structure is modelled with finite elements. The deflections are expressed as an expansion in terms of the free-vibration modes. The time-dependent coefficients are the generalized coordinates of the entire dynamic system. The concept of virtual work was extended to develop a method to transfer the aerodynamic loads to the structural nodes. Depending on the speed of the aircraft, the numerical results show damped responses to initial disturbances (although there are no viscous terms in either the aerodynamic or structural model), merging of modal frequencies, the development of limit-cycle oscillations, and the occurrence of a supercritical Hopf bifurcation leading to motion on a torus.
A framework is developed for the computational fluid dynamics (CFD) analyses of a series of helicopter rotor flowfields in hover and in forward flight. The methodology is based on the unsteady solutions of the three-dimensional, compressible Navier-Stokes equations recast in a rotating frame of reference. The simulations are carried out by solving the developed mathematical model on hybrid meshes that aim to optimally exploit the benefits of both the structured and the unstructured grids around complex configurations. The computer code is prepared for parallel processing with distributed memory utilization in order to significantly reduce the computational time and the memory requirements. The developed model and the simulation methodology are validated for single-rotor-in-hover flowfields by comparing the present results with the published experimental data. The predictive merit of different turbulence models for complex helicopter aerodynamics are tested extensively. All but the kappa-o and LES results demonstrate acceptable agreement with the experimental data. It was deemed best to use the one-equation Spalart-Allmaras turbulence model for the subsequent rotor flowfield computations. First, the flowfield around a single rotor in forward flight is simulated. These time---accurate computations help to analyze an adverse effect of increasing the forward flight speed. A dissymmetry of the lift on the advancing and the retreating blades is observed for six different advance ratios. Since the coaxial rotor is proposed to mitigate the dissymmetry, it is selected as the next logical step of the present investigation. The time---accurate simulations are successfully obtained for the flowfields generated by first a hovering then a forward-flying coaxial rotor. The results for the coaxial rotor in forward flight verify the aerodynamic balance proposed by the previously published advancing blade concept. The final set of analyses aims to investigate if the gap between the
Andrisani, D., II; Daughaday, H.; Dittenhauser, J.; Rynaski, E.
The aerodynamics, control system, instrumentation complement and recording system of the USAF Total In/Flight Simulator (TIFS) airplane are described. A control system that would allow the ailerons to be operated collectively, as well as, differentially to entrance the ability of the vehicle to perform the dual function of maneuver load control and gust alleviation is emphasized. Mathematical prediction of the rigid body and the flexible equations of longitudinal motion using the level 2.01 FLEXSTAB program are included along with a definition of the vehicle geometry, the mass and stiffness distribution, the calculated mode frequencies and mode shapes, and the resulting aerodynamic equations of motion of the flexible vehicle. A complete description of the control and instrumentation system of the aircraft is presented, including analysis, ground test and flight data comparisons of the performance and bandwidth of the aerodynamic surface servos. Proposed modification for improved performance of the servos are also presented.
LSPC is the Loading Simulation Program in C++, a watershed modeling system that includes streamlined Hydrologic Simulation Program Fortran (HSPF) algorithms for simulating hydrology, sediment, and general water quality on land as well as a simplified stream transport model. LSPC ...
Kenner, B. G.; Lincoln, N. R.
The manual is intended to show the revisions and additions to the current STAR FORTRAN. The changes are made to incorporate an FMP (Flow Model Processor) for use in the Numerical Aerodynamic Simulation Facility (NASF) for the purpose of simulating fluid flow over three-dimensional bodies in wind tunnel environments and in free space. The FORTRAN programming language for the STAR-100 computer contains both CDC and unique STAR extensions to the standard FORTRAN. Several of the STAR FORTRAN extensions to standard FOR-TRAN allow the FORTRAN user to exploit the vector processing capabilities of the STAR computer. In STAR FORTRAN, vectors can be expressed with an explicit notation, functions are provided that return vector results, and special call statements enable access to any machine instruction.
Mohrfeld-Halterman, J. A.; Uddin, M.
Presented in this paper is a procedure to develop a high fidelity quasi steady-state aerodynamic model for use in race car vehicle dynamic simulations. Developed to fit quasi steady-state wind tunnel data, the aerodynamic model is regressed against three independent variables: front ground clearance, rear ride height, and yaw angle. An initial dual range model is presented and then further refined to reduce the model complexity while maintaining a high level of predictive accuracy. The model complexity reduction decreases the required amount of wind tunnel data thereby reducing wind tunnel testing time and cost. The quasi steady-state aerodynamic model for the pitch moment degree of freedom is systematically developed in this paper. This same procedure can be extended to the other five aerodynamic degrees of freedom to develop a complete six degree of freedom quasi steady-state aerodynamic model for any vehicle.
Stakolich, E. G.; Stromberg, W. J.
This paper presents the results of testing to identify the effects of simulated aerodynamic flight loads on JT9D engine performance. The test results were also used to refine previous analytical studies on the impact of aerodynamic flight loads on performance losses. To accomplish these objectives, a JT9D-7AH engine was assembled with average production clearances and new seals as well as extensive instrumentation to monitor engine performance, case temperatures, and blade tip clearance changes. A special loading device was designed and constructed to permit application of known moments and shear forces to the engine by the use of cables placed around the flight inlet. The test was conducted in the Pratt and Whitney Aircraft X-Ray Test Facility to permit the use of X-ray techniques in conjunction with laser blade tip proximity probes to monitor important engine clearance changes. Upon completion of the test program, the test engine was disassembled, and the condition of gas path parts and final clearances were documented. The test results indicate that the engine lost 1.1 percent in thrust specific fuel consumption (TSFC), as measured under sea level static conditions, due to increased operating clearances caused by simulated flight loads. This compares with 0.9 percent predicted by the analytical model and previous study efforts.
Stakolich, E. G.; Stromberg, W. J.
The results of testing to identify the effects of simulated aerodynamic flight loads on JT9D engine performance are presented. The test results were also used to refine previous analytical studies on the impact of aerodynamic flight loads on performance losses. To accomplish these objectives, a JT9D-7AH engine was assembled with average production clearances and new seals as well as extensive instrumentation to monitor engine performance, case temperatures, and blade tip clearance changes. A special loading device was designed and constructed to permit application of known moments and shear forces to the engine by the use of cables placed around the flight inlet. The test was conducted in the Pratt & Whitney Aircraft X-Ray Test Facility to permit the use of X-ray techniques in conjunction with laser blade tip proximity probes to monitor important engine clearance changes. Upon completion of the test program, the test engine was disassembled, and the condition of gas path parts and final clearances were documented. The test results indicate that the engine lost 1.1 percent in thrust specific fuel consumption (TSFC), as measured under sea level static conditions, due to increased operating clearances caused by simulated flight loads. This compares with 0.9 percent predicted by the analytical model and previous study efforts.
Batterson, J. G.
The successful parametric modeling of the aerodynamics for an airplane operating at high angles of attack or sideslip is performed in two phases. First the aerodynamic model structure must be determined and second the associated aerodynamic parameters (stability and control derivatives) must be estimated for that model. The purpose of this paper is to document two versions of a stepwise regression computer program which were developed for the determination of airplane aerodynamic model structure and to provide two examples of their use on computer generated data. References are provided for the application of the programs to real flight data. The two computer programs that are the subject of this report, STEP and STEPSPL, are written in FORTRAN IV (ANSI l966) compatible with a CDC FTN4 compiler. Both programs are adaptations of a standard forward stepwise regression algorithm. The purpose of the adaptation is to facilitate the selection of a adequate mathematical model of the aerodynamic force and moment coefficients of an airplane from flight test data. The major difference between STEP and STEPSPL is in the basis for the model. The basis for the model in STEP is the standard polynomial Taylor's series expansion of the aerodynamic function about some steady-state trim condition. Program STEPSPL utilizes a set of spline basis functions.
Richard, M.; Harrison, B. A.
The program input presented consists of configuration geometry, aerodynamic parameters, and modal data; output includes element geometry, pressure difference distributions, integrated aerodynamic coefficients, stability derivatives, generalized aerodynamic forces, and aerodynamic influence coefficient matrices. Optionally, modal data may be input on magnetic file (tape or disk), and certain geometric and aerodynamic output may be saved for subsequent use.
Tejero E, F.; Doerffer, P.; Szulc, O.; Cross, J. L.
The application of an efficient flow control system on helicopter rotor blades may lead to improved aerodynamic performance. Recently, our invention of Rod Vortex Generators (RVGs) has been analyzed for helicopter rotor blades in hover with success. As a step forward, the study has been extended to forward flight conditions. For this reason, a validation of the numerical modelling for a reference helicopter rotor (without flow control) is needed. The article presents a study of the flow-field of the AH-1G helicopter rotor in low-, medium- and high-speed forward flight. The CFD code FLOWer from DLR has proven to be a suitable tool for the aerodynamic analysis of the two-bladed rotor without any artificial wake modelling. It solves the URANS equations with LEA (Linear Explicit Algebraic stress) k-ω model using the chimera overlapping grids technique. Validation of the numerical model uses comparison with the detailed flight test data gathered by Cross J. L. and Watts M. E. during the Tip Aerodynamics and Acoustics Test (TAAT) conducted at NASA in 1981. Satisfactory agreements for all speed regimes and a presence of significant flow separation in high-speed forward flight suggest a possible benefit from the future implementation of RVGs. The numerical results based on the URANS approach are presented not only for a popular, low-speed case commonly used in rotorcraft community for CFD codes validation but preferably for medium- and high-speed test conditions that have not been published to date.
Manhardt, P. D.; Orzechowski, J. A.; Baker, A. J.
The COMOC finite element fluid mechanics computer program system is applicable to diverse problem classes. The two dimensional aerodynamics sequence was established for solution of the potential and/or viscous and turbulent flowfields associated with subsonic flight of elementary two dimensional isolated airfoils. The sequence is constituted of three specific flowfield options in COMOC for two dimensional flows. These include the potential flow option, the boundary layer option, and the parabolic Navier-Stokes option. By sequencing through these options, it is possible to computationally construct a weak-interaction model of the aerodynamic flowfield. This report is the user's guide to operation of COMOC for the aerodynamics sequence.
Lorber, Peter F.
The Rotor-Fuselage Analysis is a method of calculating the aerodynamic reaction between a helicopter rotor and fuselage. This manual describes the structure and operation of the computer programs that make up the Rotor-Fuselage Analysis, programs which prepare the input and programs which display the output.
Matas, R.; Straka, F.; Hoznedl, M.
The contribution subscribes a numerical simulation of a steam flow through a balanced control valve. The influence of some parameters in simulations were tested, analyzed and discussed. As a result of the simulations a graph of aerodynamics forces for a specific turbine characteristic was obtained. The results from numerical simulations were compared with results from experiments. The experiment was performed with an air flow, but the final data were converted with a criterion to steam flow.
Hahne, David E. (Editor)
The High-Speed Research Program sponsored the NASA High-Speed Research Program Aerodynamic Performance Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of: Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization) and High-Lift. The review objectives were to: (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. The HSR AP Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas within the airframe element of the HSR Program. This Volume 2/Part 1 publication presents the High-Lift Configuration Development session.
Lincoln, N. R.
The Ames Research Center of NASA is engaged in the development and investigation of numerical methods and computer technologies to be employed in conjunction with physical experiments, particularly utilizing wind tunnels in the furtherance of the field of aircraft and aerodynamic body design. Several studies, aimed primarily at the areas of development and production of extremely high-speed computing facilities, were conducted. The studies focused on evaluating the aspects of feasibility, reliability, costs, and practicability of designing, constructing, and bringing into effect production of a special-purpose system. An executive summary of the activities for this project is presented in this volume.
Green, F. M.; Resnick, D. R.
An FMP (Flow Model Processor) was designed for use in the Numerical Aerodynamic Simulation Facility (NASF). The NASF was developed to simulate fluid flow over three-dimensional bodies in wind tunnel environments and in free space. The facility is applicable to studying aerodynamic and aircraft body designs. The following general topics are discussed in this volume: (1) FMP functional computer specifications; (2) FMP instruction specification; (3) standard product system components; (4) loosely coupled network (LCN) specifications/description; and (5) three appendices: performance of trunk allocation contention elimination (trace) method, LCN channel protocol and proposed LCN unified second level protocol.
Isogai, Koji; Kawabe, Hiroyasu
The paper describes how to simulate the flight of a flapping-wing micro-aerial vehicle (MAV). It uses an aerodynamic database generated using three-dimensional Navier-Stokes code. The database is composed of the time mean aerodynamic forces and moments generated at various flapping wing motions in various flight modes. Flight is simulated utilizing the database by interpolation. The procedure is applied to transition flight of a dragonfly-type MAV with two-pairs of resonance-type flapping wings. The present MAV attains the mission of hovering, transition and cruising flights successfully with stable attitude.
Chen, Xin; Wang, Dengfeng; Ma, Zhengdong
How to simulate interior aerodynamic noise accurately is an important question of a car interior noise reduction. The unsteady aerodynamic pressure on body surfaces is proved to be the key effect factor of car interior aerodynamic noise control in high frequency on high speed. In this paper, a detail statistical energy analysis (SEA) model is built. And the vibra-acoustic power inputs are loaded on the model for the valid result of car interior noise analysis. The model is the solid foundation for further optimization on car interior noise control. After the most sensitive subsystems for the power contribution to car interior noise are pointed by SEA comprehensive analysis, the sound pressure level of car interior aerodynamic noise can be reduced by improving their sound and damping characteristics. The further vehicle testing results show that it is available to improve the interior acoustic performance by using detailed SEA model, which comprised by more than 80 subsystems, with the unsteady aerodynamic pressure calculation on body surfaces and the materials improvement of sound/damping properties. It is able to acquire more than 2 dB reduction on the central frequency in the spectrum over 800 Hz. The proposed optimization method can be looked as a reference of car interior aerodynamic noise control by the detail SEA model integrated unsteady computational fluid dynamics (CFD) and sensitivity analysis of acoustic contribution.
Walker, Eric L.; Hemsch, Michael J.; Pinier, Jeremy T.; Bibb, Karen L.; Chan, David T.; Hanke, Jeremy L.
The NASA Constellation Program has worked for the past five years to develop a re- placement for the current Space Transportation System. Of the elements that form the Constellation Program, only two require databases that define aerodynamic environments and their respective uncertainty: the Ares launch vehicles and the Orion crew and launch abort vehicles. Teams were established within the Ares and Orion projects to provide repre- sentative aerodynamic models including both baseline values and quantified uncertainties. A technical team was also formed within the Constellation Program to facilitate integra- tion among the project elements. This paper is a summary of the collective experience of the three teams working with the quantification and use of uncertainty in aerodynamic environments: the Ares and Orion project teams as well as the Constellation integration team. Not all of the lessons learned discussed in this paper could be applied during the course of the program, but they are included in the hope of benefiting future projects.
Aerodynamics calculations are treated in routines which concern moments as they vary with flight conditions and attitude. The subroutines discussed: (1) compute component equivalent flat plate and wetted areas and profile drag; (2) print and plot low and high speed drag polars; (3) determine life coefficient or angle of attack; (4) determine drag coefficient; (5) determine maximum lift coefficient and drag increment for various flap types and flap settings; and (6) determine required lift coefficient and drag coefficient in cruise flight.
Winter, O. A.
The B01 version of the United Subsonic Supersonic Aerodynamic Analysis program is the result of numerous modifications and additions made to the B00 version. These modifications and additions affect the program input, its computational options, the code readability, and the overlay structure. The following are described: (1) the revised input; (2) the plotting overlay programs which were also modified, and their associated subroutines, (3) the auxillary files used by the program, the revised output data; and (4) the program overlay structure.
Naziar, J.; Couch, R.; Davis, M.
Traditionally, aeropropulsion structural performance and aerodynamic performance have been designed separately and later mated together via flight testing. In today`s atmosphere of declining resources, it is imperative that more productive ways of designing and verifying aeropropulsion performance and structural interaction be made available to the aerospace industry. One method of obtaining a more productive design and evaluation capability is through the use of numerical simulations. Currently, Lawrence Livermore National Laboratory has developed a generalized fluid/structural interaction code known as ALE3D. This code is capable of characterizing fluid and structural interaction for components such as the combustor, fan/stators, inlet and/or nozzles. This code solves the 3D Euler equations and has been applied to several aeropropulsion applications such as a supersonic inlet and a combustor rupture simulation. To characterize aerodynamic-structural interaction for rotating components such as the compressor, appropriate turbomachinery simulations would need to be implemented within the ALE3D structure. The Arnold Engineering Development Center is currently developing a three-dimensional compression system code known as TEACC (Turbine Engine Analysis Compressor Code). TEACC also solves the 3D Euler equations and is intended to simulate dynamic behavior such as inlet distortion, surge or rotating stall. The technology being developed within the TEACC effort provides the necessary turbomachinery simulation for implementation into ALE3D. This paper describes a methodology to combine three-dimensional aerodynamic turbomachinery technology into the existing aerodynamic-structural interaction simulation, ALE3D to obtain the desired aerodynamic and structural integrated simulation for an aeropropulsion system.
Kerr, P. A.; Petley, D. H.
A user's guide for the computer program that calculates the steady-state aerodynamic loads on the Shuttle thermal-protection tiles is presented. The main element in the program is the MITAS-II, Martin Marietta Interactive Thermal Analysis System. The MITAS-II is used to calculate the mass flow in a nine-tile model designed to simulate conditions duing a Shuttle flight. The procedures used to execute the program using the MITAS-II software are described. A list of the necessry software and data files along with a brief description of their functions is given. The format of the data file containing the surface pressure data is specified. The interpolation techniques used to calculate the pressure profile over the tile matrix are briefly described. In addition, the output from a sample run is explained. The actual output and the procedure file used to execute the program at NASA Langley Research Center on a CDC CYBER-175 are provided in the appendices.
Lincoln, N. R.
Functional requirements and preliminary design data were identified for use in the design of all system components and in the construction of a facility to perform aerodynamic simulation for airframe design. A skeleton structure of specifications for the flow model processor and monitor, the operating system, and the language and its compiler is presented.
Lee, Jungil; Lee, Hansol; Kim, Woojin; Choi, Haecheon
In a ski jumping competition, it is essential to analyze the effect of various posture parameters of a ski jumper to achieve a longer flight distance. For this purpose, we conduct a large eddy simulation (LES) of turbulent flow past a model ski jumper which is obtained by 3D scanning a ski jumper's body (Mr. Chil-Ku Kang, member of the Korean national team). The angle of attack of the jump ski is 30° and the Reynolds number based on the length of the jump ski is 540,000. The flow statistics including the drag and lift coefficients in flight are in good agreements with our own experimental data. We investigate the flow characteristics such as the flow separation and three-dimensional vortical structures and their effects on the drag and lift. In addition to LES, we construct a simple geometric model of a ski jumper where each part of the ski jumper is modeled as a canonical bluff body such as the sphere, cylinder and flat plate, to find its optimal posture. The results from this approach will be compared with those by LES and discussed. Supported by NRF program (2014M3C1B1033848, 2014R1A1A1002671).
Broeren, Andy P.; Addy, Harold E., Jr.; Lee, Sam; Monastero, Marianne C.
Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice-accretion documentation methods have yielded a laser-scanning capability that can measure highly three-dimensional (3-D) features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing Research Tunnel using an 18-in. chord, two-dimensional (2-D) straight wing with NACA 23012 airfoil section. For six ice-accretion cases, a 3-D laser scan was performed to document the ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 × 10(exp 6) and a Mach number of 0.18 with an 18-in. chord NACA 23012 airfoil model that was designed to accommodate the artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For five of the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil maximum lift coefficient were less than 3 percent with corresponding differences in stall angle of approximately 1 deg or less. The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice accretion for a NACA 23012 airfoil. For several
Broeren, Andy P.; Addy, Harold E., Jr.; Lee, Sam; Monastero, Marianne C.
Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice accretion documentation methods have yielded a laser-scanning capability that can measure highly three-dimensional features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing Research Tunnel using an 18-inch chord, 2-D straight wing with NACA 23012 airfoil section. For six ice accretion cases, a 3-D laser scan was performed to document the ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 x 10(exp 6) and a Mach number of 0.18 with an 18-inch chord NACA 23012 airfoil model that was designed to accommodate the artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For four of the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil maximum lift coefficient were less than 3% with corresponding differences in stall angle of approximately one degree or less. The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice accretion for a NACA 23012 airfoil. For several of the ice
Churchfield, M. J.; Lee, S.; Moriarty, P. J.; Martinez, L. A.; Leonardi, S.; Vijayakumar, G.; Brasseur, J. G.
In this work, we present results of a large-eddy simulation of the 48 multi-megawatt turbines composing the Lillgrund wind plant. Turbulent inflow wind is created by performing an atmospheric boundary layer precursor simulation and turbines are modeled using a rotating, variable-speed actuator line representation. The motivation for this work is that few others have done wind plant large-eddy simulations with a substantial number of turbines, and the methods for carrying out the simulations are varied. We wish to draw upon the strengths of the existing simulations and our growing atmospheric large-eddy simulation capability to create a sound methodology for performing this type of simulation. We have used the OpenFOAM CFD toolbox to create our solver.
Churchfield, M. J.; Lee, S.; Moriarty, P. J.; Martinez, L. A.; Leonardi, S.; Vijayakumar, G.; Brasseur, J. G.
In this work, we present results of a large-eddy simulation of the 48 multi-megawatt turbines composing the Lillgrund wind plant. Turbulent inflow wind is created by performing an atmospheric boundary layer precursor simulation, and turbines are modeled using a rotating, variable-speed actuator line representation. The motivation for this work is that few others have done large-eddy simulations of wind plants with a substantial number of turbines, and the methods for carrying out the simulations are varied. We wish to draw upon the strengths of the existing simulations and our growing atmospheric large-eddy simulation capability to create a sound methodology for performing this type of simulation. We used the OpenFOAM CFD toolbox to create our solver. The simulated time-averaged power production of the turbines in the plant agrees well with field observations, except with the sixth turbine and beyond in each wind-aligned. The power produced by each of those turbines is overpredicted by 25-40%. A direct comparison between simulated and field data is difficult because we simulate one wind direction with a speed and turbulence intensity characteristic of Lillgrund, but the field observations were taken over a year of varying conditions. The simulation shows the significant 60-70% decrease in the performance of the turbines behind the front row in this plant that has a spacing of 4.3 rotor diameters in this direction. The overall plant efficiency is well predicted. This work shows the importance of using local grid refinement to simultaneously capture the meter-scale details of the turbine wake and the kilometer-scale turbulent atmospheric structures. Although this work illustrates the power of large-eddy simulation in producing a time-accurate solution, it required about one million processor-hours, showing the significant cost of large-eddy simulation.
Zhu, Xiaowei; Anderson, William
The inherent spatial heterogeneity exhibited by real urban environments complicates a priori estimation of the roughness height needed to parameterize the inertial layer mean streamwise velocity. A large-eddy simulation study of turbulent flow over 3-D random urban-like topographies is conducted to explore the effects of surface geometry on bulk aerodynamic characterization. In a mean sense, we find that statistical attributes including surface height root mean square and skewness can adequately capture the spatial heterogeneities and randomness of real urban geometries. We find, however, that higher-order statistical moments have a negligible affect on aerodynamic drag (i.e. kurtosis may be omitted). The results enable exploration of applicability of some recently-proposed roughness parameterizations that are relevant to complex, urban-like roughness (including the model proposed by Flack and Schultz, 2010: J. Fluids Eng. 132, 041203-1). We evaluate empirical parameters needed in these models for the present urban-like cases. We find that two empirical parameters (relevant to height rms and skewness) can characterize the bulk aerodynamic roughness of topographies with statistical attributes comparable to dense urban environments. This work was supported by the Army Research Office, Atmospheric Sciences Program (PM: Dr. S. Collier) under Grant # W911NF-13-1-0474. Computational resources were provided by the Texas Advanced Computing Center at the University of Texas.
Busch, Greg T.
Determining the aerodynamic effects of ice accretion on aircraft surfaces is an important step in aircraft design and certification. The goal of this work was to develop a complete sub-scale wind tunnel simulation methodology based on knowledge of the detailed iced-airfoil flowfield that allows the accurate measurement of aerodynamic penalties associated with the accretion of ice on an airfoil and to validate this methodology using full-scale iced-airfoil performance data obtained at near-flight Reynolds numbers. In earlier work, several classifications of ice shape were developed based on key aerodynamic features in the iced-airfoil flowfield: ice roughness, streamwise ice, horn ice, and tall and short spanwise-ridge ice. Castings of each of these classifications were acquired on a full-scale NACA 23012 airfoil model and the aero-dynamic performance of each was measured at a Reynolds number of 12.0 x 106 and a Mach number = 0.20. In the current study, sub-scale simple-geometry and 2-D smooth simulations of each of these castings were constructed based on knowledge of iced-airfoil flowfields. The effects of each simulation on the aerodynamic performance of an 18-inch chord NACA 23012 airfoil model was measured in the University of Illinois 3 x 4 ft. wind tunnel at a Reynolds number of 1.8 x 106 and a Mach number of 0.18 and compared with that measured for the corresponding full-scale casting at high Reynolds number. Geometrically-scaled simulations of the horn-ice and tall spanwise-ridge ice castings modeled C l,maxto within 2% and Cd,min to within 15%. Good qualitative agreement in the Cp distributions suggests that important geometric features such as horn and ridge height, surface location, and angle with respect to the airfoil chordline were appropriately modeled. Geometrically-scaled simulations of the ice roughness, streamwise ice, and short-ridge ice tended to have conservative C l,max and Cd. The aerodynamic performance of simulations of these types of
Jackson, E. Bruce; Cruz, Christopher I.
A nonlinear, six-degree-of-freedom aerodynamic model for an early version of the HL-20 lifting body is described and compared with wind tunnel data upon which it is based. Polynomial functions describing most of the aerodynamic parameters are given and tables of these functions are presented. Techniques used to arrive at these functions are described. Basic aerodynamic coefficients were modeled as functions of angles of attack and sideslip. Vehicle lateral symmetry was assumed. Compressibility (Mach) effects were ignored. Control-surface effectiveness was assumed to vary linearly with angle of deflection and was assumed to be invariant with the angle of sideslip. Dynamic derivatives were obtained from predictive aerodynamic codes. Landing-gear and ground effects were scaled from Space Shuttle data. The model described is provided to support pilot-in-the-loop simulation studies of the HL-20. By providing the data in tabular format, the model is suitable for the data interpolation architecture of many existing engineering simulation facilities. Because of the preliminary nature of the data, however, this model is not recommended for study of the absolute performance of the HL-20.
Draxl, C.; Churchfield, M.; Mirocha, J.; Lee, S.; Lundquist, J.; Michalakes, J.; Moriarty, P.; Purkayastha, A.; Sprague, M.; Vanderwende, B.
Wind plant aerodynamics are influenced by a combination of microscale and mesoscale phenomena. Incorporating mesoscale atmospheric forcing (e.g., diurnal cycles and frontal passages) into wind plant simulations can lead to a more accurate representation of microscale flows, aerodynamics, and wind turbine/plant performance. Our goal is to couple a numerical weather prediction model that can represent mesoscale flow [specifically the Weather Research and Forecasting model] with a microscale LES model (OpenFOAM) that can predict microscale turbulence and wake losses.
Studwell, R. E.
User instructions for the Tektronix Graphics Package of the Automated Paneling Technique (APT) and the Wing and Body Aerodynamic Technique (WABAT) Programs are provided. Responses to plot package messages which the user must make to activate plot package operations and options are described. Modifications to the APT and WABAT input run streams, to affect the graphic interface, are also covered.
The computer program for a comprehensive analytical model of rotorcraft aerodynamics and dynamics is described. This analysis is designed to calculate rotor performance, loads, and noise; the helicopter vibration and gust response; the flight dynamics and handling qualities; and the system aeroelastic stability. The analysis is a combination of structural, inertial, and aerodynamic models that is applicable to a wide range of problems and a wide class of vehicles. The analysis is intended for use in the design, testing, and evaluation of rotors and rotorcraft and to be a basis for further development of rotary wing theories.
Mohrfeld-Halterman, J. A.; Uddin, M.
We described in this paper the development of a high fidelity vehicle aerodynamic model to fit wind tunnel test data over a wide range of vehicle orientations. We also present a comparison between the effects of this proposed model and a conventional quasi steady-state aerodynamic model on race vehicle simulation results. This is done by implementing both of these models independently in multi-body quasi steady-state simulations to determine the effects of the high fidelity aerodynamic model on race vehicle performance metrics. The quasi steady state vehicle simulation is developed with a multi-body NASCAR Truck vehicle model, and simulations are conducted for three different types of NASCAR race tracks, a short track, a one and a half mile intermediate track, and a higher speed, two mile intermediate race track. For each track simulation, the effects of the aerodynamic model on handling, maximum corner speed, and drive force metrics are analysed. The accuracy of the high-fidelity model is shown to reduce the aerodynamic model error relative to the conventional aerodynamic model, and the increased accuracy of the high fidelity aerodynamic model is found to have realisable effects on the performance metric predictions on the intermediate tracks resulting from the quasi steady-state simulation.
Staszak, T.; Brede, M.; Strelnikov, B.
The only way to do in-situ measurements, which are very important experimental studies for atmospheric science, in the mesoshere/lower thermosphere (MLT) is to use sounding rockets. The drawback of using rockets is the shock wave appearing because of the very high speed of the rocket motion (typically about 1000 mIs). This shock wave disturbs the density, the temperature and the velocity fields in the vicinity of the rocket, compared to undisturbed values of the atmosphere. This effect, however, can be quantified and the measured data has to be corrected not just to make it more precise but simply usable. The commonly accepted and widely used tool for this calculations is the Direct Simulation Monte Carlo (DSMC) technique developed by GA. Bird which is available as stand-alone program limited to use a single processor. Apart from complications with simulations of flows around bodies related to different flow regimes in the altitude range of MLT, that rise due to exponential density change by several orders of magnitude, a particular hardware configuration introduces significant difficulty for aerodynamical calculations due to choice of the grid sizes mainly depending on the demands on adequate DSMCs and good resolution of geometries with scale differences of factor of iO~. This makes either the calculation time unreasonably long or even prevents the calculation algorithm from converging. In this paper we apply the free open source software OpenFOAM (licensed under GNU GPL) for a three-dimensional CFD-Simulation of a flow around a sounding rocket instrumentation. An advantage of this software package, among other things, is that it can run on high performance clusters, which are easily scalable. We present the first results and discuss the potential of the new tool in applications for sounding rockets.
Bower, W. W.
The critical issues in viscous flow simulations, such as boundary-layer separation, entrainment, turbulence modeling, and compressibility, are discussed with regard to the ground effects problem for vertical-takeoff-and-landing (VTOL) aircraft. A simulation of the two-dimensional incompressible lift jet in ground proximity is based on solution of the Reynolds-averaged Navier-Stokes equations and a turbulence-model equation which are written in stream function-vorticity form and are solved using Hoffman's augmented-central-difference algorithm. The resulting equations and their shortcomings are discussed when the technique is extended to two-dimensional compressible and three-dimensional incompressible flows.
Glaab, Louis J.; Riley, Donald R.; Brandon, Jay M.; Person, Lee H., Jr.; Glaab, Patricia C.
As part of an effort between NASA and private industry to reduce airport-community noise for high-speed civil transport (HSCT) concepts, a piloted simulation study was initiated for the purpose of predicting the noise reduction benefits that could result from improved low-speed high-lift aerodynamic performance for a typical HSCT configuration during takeoff and initial climb. Flight profile and engine information from the piloted simulation were coupled with the NASA Langley Aircraft Noise Prediction Program (ANOPP) to estimate jet engine noise and to propagate the resulting source noise to ground observer stations. A baseline aircraft configuration, which also incorporated different levels of projected improvements in low-speed high-lift aerodynamic performance, was simulated to investigate effects of increased lift and lift-to-drag ratio on takeoff noise levels. Simulated takeoff flights were performed with the pilots following a specified procedure in which either a single thrust cutback was performed at selected altitudes ranging from 400 to 2000 ft, or a multiple-cutback procedure was performed where thrust was reduced by a two-step process. Results show that improved low-speed high-lift aerodynamic performance provides at least a 4 to 6 dB reduction in effective perceived noise level at the FAA downrange flyover measurement station for either cutback procedure. However, improved low-speed high-lift aerodynamic performance reduced maximum sideline noise levels only when using the multiple-cutback procedures.
An experimental study was conducted to investigate the effect of simulated glaze ice shapes on the aerodynamic performance characteristics of two-dimensional airfoils and a swept finite tail. The two dimensional tests involved two NACA 0011 airfoils with chords of 24 and 12 inches. Glaze ice shapes computed with the LEWICE code that were representative of 22.5-min and 45-min ice accretions were simulated with spoilers, which were sized to approximate the horn heights of the LEWICE ice shapes. Lift, drag, pitching moment, and surface pressure coefficients were obtained for a range of test conditions. Test variables included Reynolds number, geometric scaling, control deflection and the key glaze ice features, which were horn height, horn angle, and horn location. For the three-dimensional tests, a 25%-scale business jet empennage (BJE) with a T-tail configuration was used to study the effect of ice shapes on the aerodynamic performance of a swept horizontal tail. Simulated glaze ice shapes included the LEWICE and spoiler ice shapes to represent 9-min and 22.5-min ice accretions. Additional test variables included Reynolds number and elevator deflection. Lift, drag, hinge moment coefficients as well as boundary layer velocity profiles were obtained. The experimental results showed substantial degradation in aerodynamic performance of the airfoils and the swept horizontal tail due to the simulated ice shapes. For the two-dimensional airfoils, the largest aerodynamic penalties were obtained when the 3-in spoiler-ice, which was representative of 45-min glaze ice accretions, was set normal to the chord. Scale and Reynolds effects were not significant for lift and drag. However, pitching moments and pressure distributions showed great sensitivity to Reynolds number and geometric scaling. For the threedimensional study with the swept finite tail, the 22.5-min ice shapes resulted in greater aerodynamic performance degradation than the 9-min ice shapes. The addition of 24
Peele, E. L.; Adams, W. M., Jr.
A computer program, ISAC, is described which calculates the stability and response of a flexible airplane equipped with active controls. The equations of motion relative to a fixed inertial coordinate system are formulated in terms of the airplane's rigid body motion and its unrestrained normal vibration modes. Unsteady aerodynamic forces are derived from a doublet lattice lifting surface theory. The theoretical basis for the program is briefly explained together with a description of input data and output results.
Meakin, Robert L.
A prototype method for time-accurate simulation of multiple aerodynamic bodies in relative motion is presented. The method is general and features unsteady chimera domain decomposition techniques and an implicit approximately factored finite-difference procedure to solve the time-dependent thin-layer Navier-Stokes equations. The method is applied to a set of two- and three- dimensional test problems to establish spatial and temporal accuracy, quantify computational efficiency, and begin to test overall code robustness.
Handschuh, Robert; Hill, Matthew; Kunz, Robert; Long, Lyle; Morris, Philip; Noack, Ralph
A computational fluid dynamics (CFD) method is adapted, validated and applied to spinning gear systems with emphasis on predicting windage losses. Several spur gears and a disc are studied. The CFD simulations return good agreement with measured windage power loss. Turbulence modeling choices, the relative importance of viscous and pressure torques with gear speed and the physics of the complex 3-D unsteady flow field in the vicinity of the gear teeth are studied.
Newsome, Richard W.; Kandil, Osama A.
Progress in the numerical simulation of vortical flow due to three-dimensional flow separation about flight vehicles at high angles of attack and quasi-steady flight conditions is surveyed. Primary emphasis is placed on Euler and Reynolds-averaged Navier-Stokes methods where the vortices are 'captured' as a solution to the governing equations. A discussion of the relevant flow physics provides a perspective from which to assess numerical solutions. Current numerical prediction capabilities and their evolutionary development are surveyed. Future trends and challenges are identified and discussed.
Batina, John T.
A new computer program has been developed called ASP3D (Advanced Small Perturbation - 3D), which solves the small perturbation potential flow equation in an advanced form including mass-consistent surface and trailing wake boundary conditions, and entropy, vorticity, and viscous effects. The purpose of the program is for unsteady aerodynamic and aeroelastic analyses, especially in the nonlinear transonic flight regime. The program exploits the simplicity of stationary Cartesian meshes with the movement or deformation of the configuration under consideration incorporated into the solution algorithm through a planar surface boundary condition. The paper presents unsteady aerodynamic and aeroelastic applications of ASP3D to assess the time dependent capability and demonstrate various features of the code.
Wilder, M. C.; Bogdanoff, D. W.
The Hypervelocity Free Flight Aerodynamic Facility at NASA Ames Research Center provides a potential platform for the experimental simulation of meteor breakup at conditions that closely match full-scale entry condition for select parameters. The poster describes the entry environment simulation capabilities of the Hypervelocity Free Flight Aerodynamic Facility (HFFAF) at NASA Ames Research Center and provides example images of the fragmentation of a hypersonic projectile for which break-up was initiated by mechanical forces (impact with a thin polymer diaphragm).
Status report on the Aeronautical Research Institute of Sweden version of the missile aerodynamics program LARV, for calculation of static aerodynamic properties and longitudinal aerodynamic damping derivatives. Part 1: Theory
Improvements to a missile aerodynamics program which enable it to (a) calculate aerodynamic coefficients as input for a flight mechanics model, (b) check manufacturers' data or estimate performance from photographs, (c) reduce wind tunnel testing, and (d) aid optimization studies, are discussed. Slender body theory is used for longitudinal damping derivatives prediction. Program predictions were compared to known values. Greater accuracy is required in the estimation of drag due to excrescences on actual missile configurations, the influence of a burning motor, and nonlinear effects in the stall region. Prediction of pressure centers on wings and on bodies in presence of wings must be improved.
Described is a computer program written in FORTRAN IV which offers considerable flexibility in generating simulated data pertinent to education and educational psychology. The user is allowed to specify the number of samples, data sets, and variables, together with the population means, standard deviations and intercorrelations. In addition the…
Harrison, B. A.; Richard, M.
The information necessary for execution of the digital computer program L216 on the CDC 6600 is described. L216 characteristics are based on the doublet lattice method. Arbitrary aerodynamic configurations may be represented with combinations of nonplanar lifting surfaces composed of finite constant pressure panel elements, and axially summetric slender bodies composed of constant pressure line elements. Program input consists of configuration geometry, aerodynamic parameters, and modal data; output includes element geometry, pressure difference distributions, integrated aerodynamic coefficients, stability derivatives, generalized aerodynamic forces, and aerodynamic influence coefficient matrices. Optionally, modal data may be input on magnetic field (tape or disk), and certain geometric and aerodynamic output may be saved for subsequent use.
Conley, Joseph L.
The computer program AeroFcn is discussed. AeroFcn is a utility program that computes the following aerodynamic parameters: geopotential altitude, Mach number, true velocity, dynamic pressure, calibrated airspeed, equivalent airspeed, impact pressure, total pressure, total temperature, Reynolds number, speed of sound, static density, static pressure, static temperature, coefficient of dynamic viscosity, kinematic viscosity, geometric altitude, and specific energy for a standard- or a modified standard-day atmosphere using compressible flow and normal shock relations. Any two parameters that define a unique flight condition are selected, and their values are entered interactively. The remaining parameters are computed, and the solutions are stored in an output file. Multiple cases can be run, and the multiple case solutions can be stored in another output file for plotting. Parameter units, the output format, and primary constants in the atmospheric and aerodynamic equations can also be changed.
Agarwal, R. K.
The source code for the SPICE 2 program was deblocked in order to isolate and compile the subroutine in an effort to provide a software simulation of discrete and combinatorial electronic components. Incompatibilities between the UNIVAC 1180 FORTRAN and the Sigma V CP-V FORTRAN 4 were resolved. The SPICE 2 model is to be used to determine gate and fan-out delays, logic state conditions, and signal race conditions for transistor array elements and circuit logic to be patterned in the (SPI) 7101 CMOS silicon gate semicustom array. The simulator is to be operable from the CP-V time sharing terminals.
Melvin, Arron Hector
A simulation-based aerodynamic design tool is developed for multi-element high-lift airfoils operating in ground effect. A control theory approach is adopted, using the compressible Navier-Stokes equations as the basis for viscous design of airfoil element shapes and relative positioning. Particular considerations of aerodynamic design, high-lift systems, and the ground effect are described, and the suitability of aerodynamic shape optimization of such systems is discussed. The model of fluid flow and its discretization for solution on digital computers is investigated. A cell-centered finite-volume explicit multigrid method is used to solve both the flow and adjoint systems utilizing structured multiblock meshes. The adjoint equations for shape optimization are developed using a continuous adjoint formulation, and implemented with a moving ground boundary condition for the first time. A suite of test cases verified and validated the numerical algorithms and implementation. Realistic case studies were performed, demonstrating significant performance improvements over the baseline configurations. These included two free-air multi-element airfoil drag minimizations, and in addition two inverted two-element airfoil drag minimizations in ground effect.
Lai, Y. G. J.; Przekwas, A. J.; So, R. M. C.
In the past, most aerodynamic flow calculations were carried out with density-based numerical methods and zero-equation turbulence models. However, pressure-based methods and more advanced turbulence models have been routinely used in industry for many internal flow simulations and for incompressible flows. Unfortunately, their usefulness in calculating aerodynamic flows is still not well demonstrated and accepted. In this study, an advanced pressure-based numerical method and a recently proposed near-wall compressible two-equation turbulence model are used to calculate external aerodynamic flows. Several TVD-type schemes are extended to pressure-based method to better capture discontinuities such as shocks. Some improvements are proposed to accelerate the convergence of the numerical method. A compressible near-wall two-equation turbulence model is then implemented to calculate transonic turbulent flows over NACA 0012 and RAE 2822 airfoils with and without shocks. The calculated results are compared with wind tunnel data as well as with results obtained from the Baldwin-Lomax model. The performance of the two-equation turbulence model is evaluated and its merits or lack thereof are discussed.
In helicopters, vortices (generated at the tip of the rotor blades) interact with the next advancing blades during certain flight and manoeuvring conditions, generating undesirable levels of acoustic noise and vibration. These Blade-Vortex Interactions (BVIs), which may cause the most disturbing acoustic noise, normally occur in descent or high-speed forward flight. Acoustic noise characterization (and potential reduction) is one the areas generating intensive research interest to the rotorcraft industry. Since experimental investigations of BVI are extremely costly, some insights into the BVI or AVI (2-D Airfoil-Vortex Interaction) can be gained using Computational Fluid Dynamics (CFD) numerical simulations. Numerical simulation of BVI or AVI has been of interest to CFD for many years. There are still difficulties concerning an accurate numerical prediction of BVI. One of the main issues is the inherent dissipation of CFD turbulence models, which severely affects the preservation of the vortex characteristics. Moreover this is not an issue only for aerodynamic and aeroacoustic analysis but also for aeroelastic investigations as well, especially when the strong (two-way) aeroelastic coupling is of interest. The present investigation concentrates mainly on AVI simulations. The simulations are performed for Mach number, Ma = 0.3, resulting in a Reynolds number, Re = 1.3 x 106, which is based on the chord, c, of the airfoil (NACA0012). Extensive literature search has indicated that the present work represents the first comprehensive investigation of AVI using the LES numerical approach, in the rotorcraft research community. The major factor affecting the aerodynamic coefficients and aeroacoustic field as a result of airfoil-vortex interaction is observed to be the unsteady pressure generated at the location of the interaction. The present numerical results show that the aerodynamic coefficients (lift, moment, and drag) and aeroacoustic field are strongly dependent on
Trussell, Donald H.; Weidman, Deene J.
A recently developed radiant-heating test technique for simulation of aerodynamic heating in wind tunnel is describes. The heating device, which utilized quartz-tube lamps, was operated successfully while exposed directly to a supersonic airstream. Tests were made on a calibration panel, and experimental temperature and pressure data are presented. Results indicate that initial heating rates of about 26 Btu/(sq ft) (sec) are obtainable at a distance of 12 inches from the heater. Further applications of the basic design are discussed briefly.
Crill, W.; Dale, B.
The input data required to execute the computer program ISCON are described. The program generates a numerical procedure for the determination of unsteady aerodynamic forces on arbitrarily interacting wings and tails in supersonic flow. A velocity potential gradient method is used. Constant Mach number is assumed throughout the flow field. Lifting surfaces are represented by trapezoidal elements which can be generated automatically by the program. The wake field is represented by rectangular strip elements. The formulation is reviewed as well as input overview and input format. Instruction on how to use ISCON, a sample problem, and the restart feature are discussed. Program size limitations, computer program flow, and error messages are also included along with a description of the SS31 program used to compute the coefficients of surface spline.
Cao, Chunqin; Xiang, Jingbo; Zhang, Xiaoyang; Wang, Weiqiang
The infrared (IR) imaging missile's dome will be heated when fly at high speed in the atmosphere because of the friction of the air flow blocking. The detector's performance will be decline if the dome surface is heated to a certain temperature. In this paper, we find a right way to evaluate the aerothermal effects in the imaging and information processing algorithm. Which have three steps including the aerothermal radiation calculation, quantization and image reconstruction. Firstly, the aerothermal radiation is calculated by using a combination of both methods of theoretical analysis and experiment data. Secondly, the relationship between aerothermal radiation and IR images background mean gray and noise can be calculated through the analysis of the experiment data. At last, we can rebuild an aerodynamic heating effect of infrared images fusion with target and decoy, which can be used for virtual prototyping platform missile trajectory simulation. It can be found that the above constructed images have good agreements with the actual image according to comparison between the simulation data and experiment data. It is an economic method that can solve the lab aerodynamic heating simulation and modeling problems.
Many others in the fusion energy and advanced scientific computing communities participated in the development of this plan. The core planning team is grateful for their important contributions. This summary is meant as a quick overview the Fusion Simulation Program's (FSP's) purpose and intentions. There are several additional documents referenced within this one and all are supplemental or flow down from this Program Plan. The overall science goal of the DOE Office of Fusion Energy Sciences (FES) Fusion Simulation Program (FSP) is to develop predictive simulation capability for magnetically confined fusion plasmas at an unprecedented level of integration and fidelity. This will directly support and enable effective U.S. participation in International Thermonuclear Experimental Reactor (ITER) research and the overall mission of delivering practical fusion energy. The FSP will address a rich set of scientific issues together with experimental programs, producing validated integrated physics results. This is very well aligned with the mission of the ITER Organization to coordinate with its members the integrated modeling and control of fusion plasmas, including benchmarking and validation activities. . Initial FSP research will focus on two critical Integrated Science Application (ISA) areas: ISA1, the plasma edge; and ISA2, whole device modeling (WDM) including disruption avoidance. The first of these problems involves the narrow plasma boundary layer and its complex interactions with the plasma core and the surrounding material wall. The second requires development of a computationally tractable, but comprehensive model that describes all equilibrium and dynamic processes at a sufficient level of detail to provide useful prediction of the temporal evolution of fusion plasma experiments. The initial driver for the whole device model will be prediction and avoidance of discharge-terminating disruptions, especially at high performance, which are a critical
Orlov, Dmitri; Font, Gabriel
This work presents different approaches to modeling of the plasma actuator, an electrical flow control device, which is now widely used in aerodynamics for separation control, lift enhancement, drag reduction and flight control without moving surfaces. Study of the physics of the discharge in air at atmospheric pressure was performed using particle (PIC-DSMC) and fluid plasma simulations. Based on the experimentally obtained data electro-static and lumped-element circuit models were developed for engineering purposes. Numerical flow simulations were performed to study the effect of the plasma body force on the neutral fluid. The results agreed well with the experiments. An application of the plasma actuators to the leading-edge separation control on the NACA 0021 airfoil was studied numerically. The results were obtained for a range of angles of attack. Improvement in the airfoil characteristics was observed in numerical simulations at high angles of attack in cases with plasma actuation.
Matsushima, Kisa; Takanashi, Susume; Fujii, Kozo; Obayashi, Shigeru
The slightly supersonic viscous flow about the space-plane under development at the National Aerospace Laboratory (NAL) in Japan was simulated numerically using the LU-ADI algorithm. The wind-tunnel testing for the same plane also was conducted with the computations in parallel. The main purpose of the simulation is to capture the phenomena which have a great deal of influence to the aerodynamic force and efficiency but is difficult to capture by experiments. It includes more accurate representation of vortical flows with high angles of attack of an aircraft. The space-plane shape geometry simulated is the simplified model of the real space-plane, which is a combination of a flat and slender body and a double-delta type wing. The comparison between experimental results and numerical ones will be done in the near future. It could be said that numerical results show the qualitatively reliable phenomena.
Blade fatigue life is an important element in determining the economic viability of the Vertical-Axis Wind Turbine (VAWT). A principal source of blade fatigue is thought to be the stochastic (i.e., random) aerodynamic loads created by atmospheric turbulence. This report describes the theoretical background of the VAWT Stochastic Aerodynamic Loads (VAWT-SAL) computer code, whose purpose is to numerically simulate these random loads, given the rotor geometry, operating conditions, and assumed turbulence properties. A Double-Multiple-Stream Tube (DMST) analysis is employed to model the rotor's aerodynamic response. The analysis includes the effects of Reynolds number variations, different airfoil sections and chord lengths along the blade span, and an empirical model for dynamic stall effects. The mean ambient wind is assumed to have a shear profile which is described by either a power law or a logarithmic variation with height above ground. Superimposed on this is a full 3-D field of turbulence: i.e., in addition to random fluctuations in time, the turbulence is allowed to vary randomly in planes perpendicular to the mean wind. The influence of flow retardation on the convection of turbulence through the turbine is also modeled. Calculations are presented for the VAWT 34-m Test Bed currently in operation at Bushland, Texas. Predicted time histories of the loads, as well as their Fourier spectra, are presented and discussed. Particular emphasis is placed on the differences between so-called steady-state'' (mean wind only) predictions, and those produced with turbulence present. Somewhat surprisingly, turbulence is found to be capable of either increasing or decreasing the average output power, depending on the turbine's tip-speed ratio. A heuristic explanation for such behavior is postulated, and a simple formula is derived for predicting the magnitude of this effect without the need for a full stochastic simulation. 41 refs., 32 figs., 1 tab.
The project tasks and deliverables are as follows: Computations and Experiments--(1) Simulation and analysis of a range of generic shapes, simplified to more complex, representative of tractor and integrated tractor-trailer flow characteristics using computational tools, (2) The establishment of an experimental data base for tractor-trailer models for code/computational method development and validation. The first shapes to be considered will be directed towards the investigation of tractor-trailer gaps and mismatch of tractor-trailer heights. (3) The evaluation and documentation of effective computational approaches for application to heavy vehicle aerodynamics based on the benchmark results with existing and advanced computational tools compared to experimental data, and (4) Computational tools and experimental methods for use by industry, National Laboratories, and universities for the aerodynamic modeling of heavy truck vehicles. Evaluation of current and new technologies--(1) The evaluation and documentation of current and new technologies for drag reduction based on published literature and continued communication with the heavy vehicle industry (e.g., identification and prioritization of tractor-trailer drag-sources, blowing and/or suction devices, body shaping, new experimental methods or facilities), and the identification and analysis of tractor and integrated tractor-trailer aerodynamic problem areas and possible solution strategies. (2) Continued industrial site visits. It should be noted that ''CFD tools'' are not only the actual computer codes, but descriptions of appropriate numerical solution methods. Part of the project effort will be to determine the restrictions or avenues for technology transfer.
Under this project, General Atomics (GA) was tasked to develop the experimental validation plans for two high priority ISAs, Boundary and Pedestal and Whole Device Modeling in collaboration with the theory, simulation and experimental communities. The following sections have been incorporated into the final FSP Program Plan (www.pppl.gov/fsp), which was delivered to the US Department of Energy (DOE). Additional deliverables by GA include guidance for validation, development of metrics to evaluate success and procedures for collaboration with experiments. These are also part of the final report.
Campbell, Bryan A.; Bezos, Gaudy M.
The steady-state and transient effects of simulated heavy rain on the subsonic aerodynamic characteristics of a wing model were determined in the Langley 14- by 22-Foot Subsonic Tunnel. The 1.29 foot chord wing was comprised of a NACA 23015 airfoil and had an aspect ratio of 6.10. Data were obtained while test variables of liquid water content, angle of attack, and trailing edge flap angle were parametrically varied at dynamic pressures of 10, 30, and 50 psf (i.e., Reynolds numbers of .76x10(6), 1.31x10(6), and 1.69x10(6)). The experimental results showed reductions in lift and increases in drag when in the simulated rain environment. Accompanying this was a reduction of the stall angle of attack by approximately 4 deg. The transient aerodynamic performance during transition from dry to wet steady-state conditions varied between a linear and a nonlinear transition.
Bezos, Gaudy M.; Dunham, R. Earl, Jr.; Gentry, Garl L., Jr.; Melson, W. Edward, Jr.
The effects of simulated heavy rain on the aerodynamic characteristics of an NACA 64-210 airfoil section equipped with leading-and trailing-edge high-lift devices were investigated in the Langley 14- by 22-Foot Subsonic Tunnel. The model had a chord of 2.5 ft, a span of 8 ft, and was mounted on the tunnel centerline between two large endplates. Aerodynamic measurements in and out of the simulated rain environment were obtained for dynamic pressures of 30 and 50 psf and an angle-of-attack range of 0 to 20 degrees for the cruise configuration. The rain intensity was varied to produce liquid water contents ranging from 16 to 46 gm/cu m. The results obtained for various rain intensity levels and tunnel speeds showed significant losses in maximum lift capability and increases in drag for a given lift as the liquid water content was increased. The results obtained on the landing configuration also indicate a progressive decrease in the angle of attack at which maximum lift occurred and an increase in the slope of the pitching-moment curve as the liquid water content was increased. The sensitivity of test results to the effects of the water surface tension was also investigated. A chemical was introduced into the rain environment that reduced the surface tension of water by a factor of 2. The reduction in the surface tension of water did not significantly alter the level of performance losses for the landing configuration.
Birckelbaw, Lourdes G.; Mcneil, Walter E.; Wardwell, Douglas A.
This report describes the aerodynamics model used in a simulation model of an advanced short takeoff and vertical landing (ASTOVL) lift-fan fighter aircraft. The simulation model was developed for use in piloted evaluations of transition and hover flight regimes, so that only low speed (M approximately 0.2) aerodynamics are included in the mathematical model. The aerodynamic model includes the power-off aerodynamic forces and moments and the propulsion system induced aerodynamic effects, including ground effects. The power-off aerodynamics data were generated using the U.S. Air Force Stability and Control Digital DATCOM program and a NASA Ames in-house graphics program called VORVIEW which allows the user to easily analyze arbitrary conceptual aircraft configurations using the VORLAX program. The jet-induced data were generated using the prediction methods of R. E. Kuhn et al., as referenced in this report.
Ladeinde, Foluso; Alabi, Ken; Li, Wenhai
The problem of generating design data for the operation of a farm of wind turbines for clean energy production is quite complicated, if properly done. Potential flow theories provide some models, but these are not suitable for the massive aerodynamic separation and turbulence that characterize many realistic wind turbine applications. Procedures, such as computational fluid dynamics (CFD), which can potentially resolve some of the accuracy problems with the purely theoretical approach, are quite expensive to use, and often prohibit real-time design and control. In our work, we seek affordable and acceptably-accurate models derived from the foregoing approaches. The simulation used in our study is based on high-fidelity CFD, meaning that we use high-order (compact-scheme based), mostly large-eddy simulation methods, with due regards for the proper treatment of the stochastic inflow turbulence data. Progress on the project described herein will be presented.
Bezos, Gaudy M.; Dunham, R. Earl, Jr.; Campbell, Bryan A.; Melson, W. Edward, Jr.
The NASA Langley Research Center has developed a large-scale ground testing capability for evaluating the effect of heavy rain on airfoil lift. The paper presents the results obtained at the Langley Aircraft Landing Dynamics Facility on a 10-foot cord NACA 64-210 wing section equipped with a leading-edge slat and double-slotted trailing-edge flap deflected to simulate landing conditions. Aerodynamic lift data were obtained with and without the rain simulation system turned on for an angle-of-attack range of 7.5 to 19.5 deg and for two rainfall conditions: 9 in/hr and 40 in/hr. The results are compared to and correlated with previous small-scale wind tunnel results for the same airfoil section. It appears that to first order, scale effects are not large and the wind tunnel research technique can be used to predict rain effects on airplane performance.
Broeren, Andy P.; Lee, Sam; Shah, Gautam H.; Murphy, Patrick C.
An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5 percent scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from alpha = -5deg to 85deg and beta = -45 deg to 45 deg at a Reynolds number of 0.24 x10(exp 6) and Mach number of 0.06. The 3.5 percent scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5 percent scale GTM. The addition of the large, glaze-horn type ice shapes did result in an increase in airplane drag coefficient but had little effect on the lift and pitching moment. The lateral-directional characteristics showed mixed results with a small effect of the ice shapes observed in some cases. The flow visualization images revealed the presence and evolution of a spanwise-running vortex on the wing that was the dominant feature of the flowfield for both clean and iced configurations. The lack of ice-induced performance and flowfield effects observed in this effort was likely due to Reynolds number effects for the clean configuration. Estimates of full-scale baseline performance were included in this analysis to illustrate the potential icing effects.
Piazolla, Sabino; Hemmati, Hamid; Tratt, David
Laser Ranging Simulation Program (LRSP) is a computer program that predicts selected aspects of the performances of a laser altimeter or other laser ranging or remote-sensing systems and is especially applicable to a laser-based system used to map terrain from a distance of several kilometers. Designed to run in a more recent version (5 or higher) of the MATLAB programming language, LRSP exploits the numerical and graphical capabilities of MATLAB. LRSP generates a graphical user interface that includes a pop-up menu that prompts the user for the input of data that determine the performance of a laser ranging system. Examples of input data include duration and energy of the laser pulse, the laser wavelength, the width of the laser beam, and several parameters that characterize the transmitting and receiving optics, the receiving electronic circuitry, and the optical properties of the atmosphere and the terrain. When the input data have been entered, LRSP computes the signal-to-noise ratio as a function of range, signal and noise currents, and ranging and pointing errors.
Hughes, Christopher E.; Podboy, Gary, G.; Woodward, Richard P.; Jeracki, Robert, J.
The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise signature. Therefore, a comprehensive aeroacoustic wind tunnel test program was conducted called the Fan Broadband Source Diagnostic Test as part of the NASA Quiet Aircraft Technology program. The test was performed in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 1/5 scale model turbofan simulator which represented a current generation, medium pressure ratio, high bypass turbofan aircraft engine. The investigation focused on simulating in model scale only the bypass section of the turbofan engine. The test objectives were to: identify the noise sources within the model and determine their noise level; investigate several component design technologies by determining their impact on the aerodynamic and acoustic performance of the fan stage; and conduct detailed flow diagnostics within the fan flow field to characterize the physics of the noise generation mechanisms in a turbofan model. This report discusses results obtained for one aspect of the Source Diagnostic Test that investigated the effect of the bypass or fan nozzle exit area on the bypass stage aerodynamic performance, specifically the fan and outlet guide vanes or stators, as well as the farfield acoustic noise level. The aerodynamic performance, farfield acoustics, and Laser Doppler Velocimeter flow diagnostic results are presented for the fan and four different fixed-area bypass nozzle configurations. The nozzles simulated fixed engine operating lines and encompassed the fan stage operating envelope from near stall to cruise. One nozzle was selected as a baseline reference, representing the nozzle area which would achieve the design point operating conditions and fan stage performance. The total area change from
Bensassi, Khalil; Laguna, Alejandro A.; Lani, Andrea; Mansour, Nagi N.
Numerical simulations of an arc heated flow inside NASA's 20 [MW] Aerodynamics heating facility (AHF) are performed in order to investigate the three-dimensional swirling flow and the current distribution inside the wind tunnel. The plasma is considered in Local Thermodynamics Equilibrium(LTE) and is composed of Air-Argon gas mixture. The governing equations are the Navier-Stokes equations that include source terms corresponding to Joule heating and radiative cooling. The former is obtained by solving an electric potential equation, while the latter is calculated using an innovative massively parallel ray-tracing algorithm. The fully coupled system is closed by the thermodynamics relations and transport properties which are obtained from Chapman-Enskog method. A novel strategy was developed in order to enable the flow solver and the radiation calculation to be preformed independently and simultaneously using a different number of processors. Drastic reduction in the computational cost was achieved using this strategy. Details on the numerical methods used for space discretization, time integration and ray-tracing algorithm will be presented. The effect of the radiative cooling on the dynamics of the flow will be investigated. The complete set of equations were implemented within the COOLFluiD Framework. Fig. 1 shows the geometry of the Anode and part of the constrictor of the Aerodynamics heating facility (AHF). Fig. 2 shows the velocity field distribution along (x-y) plane and the streamline in (z-y) plane.
Sawyer, W. C.; Allen, J. M.; Hernandez, G.; Dillenius, M. F. E.; Hemsch, M. J.
This paper presents a survey of engineering computational methods and experimental programs used for estimating the aerodynamic characteristics of missile configurations. Emphasis is placed on those methods which are suitable for preliminary design of conventional and advanced concepts. An analysis of the technical approaches of the various methods is made in order to assess their suitability to estimate longitudinal and/or lateral-directional characteristics for different classes of missile configurations. Some comparisons between the predicted characteristics and experimental data are presented. These comparisons are made for a large variation in flow conditions and model attitude parameters. The paper also presents known experimental research programs developed for the specific purpose of validating analytical methods and extending the capability of data-base programs.
Chang, Choong Seock
New York University, Courant Institute of Mathematical Sciences, participated in the ÃÂÃÂ¢ÃÂÃÂÃÂÃÂFusion Simulation Program (FSP) Planning ActivitiesÃÂÃÂ¢ÃÂÃÂÃÂÃÂ [http://www.pppl.gov/fsp], with C.S. Chang as the institutional PI. FSPÃÂÃÂ¢ÃÂÃÂÃÂÃÂs mission was to enable scientific discovery of important new plasma phenomena with associated understanding that emerges only upon integration. This requires developing a predictive integrated simulation capability for magnetically-confined fusion plasmas that are properly validated against experiments in regimes relevant for producing practical fusion energy. Specific institutional goal of the New York University was to participate in the planning of the edge integrated simulation, with emphasis on the usage of large scale HPCs, in connection with the SciDAC CPES project which the PI was leading. New York University successfully completed its mission by participating in the various planning activities, including the edge physics integration, the edge science drivers, and the mathematical verification. The activity resulted in the combined report that can be found in http://www.pppl.gov/fsp/Overview.html. Participation and presentations as part of this project are listed in a separation file.
Johnson, W. G., Jr.
The low-speed aerodynamic characteristics are investigated of a general research model - a swept-wing, jet-powered STOL transport with externally blown flaps. The model was tested with four-engine simulators mounted on pylons under the 9.3-percent-thick supercritical airfoil wing. Two sets of air ejectors were used to provide data with large and small engines. Tests were conducted in the Langley V/STOL tunnel over an angle-of-attack range of -4 deg to 22 deg and a thrust-coefficient range from 0 to approximately 4. The effects are described of power, wing leading-edge slat configuration, T-tail and low horizontal-tail positions, and double-slotted flap deflection. Additional untrimmed and trimmed engine-out data and tail-body data are included.
Feng, Fang; Li, Shengmao; Li, Yan; Tian, Wenqiang
To invest the effects of blade surface icing on the aerodynamics performance of the straight-bladed vertical-axis wind turbine (SB-VAWT), wind tunnel tests were carried out on a static straight blade using a simple icing wind tunnel. Firstly, the icing situations on blade surface at some kinds of typical attack angle were observed and recorded under different cold water flow fluxes. Then the iced blade airfoils were combined into a SB-VAWT model with two blades. Numerical simulations were carried out on this model, and the static and dynamic torque coefficients of the model with and without icing were computed. Both the static and dynamic torque coefficients were decreased for the icing effects.
Taghavi, R.; Dupont, A.; Dupont, J.F. )
This paper reports on a computational fluid dynamics code used as a guide during the development stage of a passenger car spark ignition engine. The focus is on the flow proiperties of the inlet port as well as the heat transfer characteristics of the proposed cylinder head design. In the first part of this study, the aerodynamic characteristics of two slightly different inlet ports are considered and their effect on the development of in-cylinder flow is examined. The collected information is used to estimate geometric sensitivity and assess the effects of drifts between design and actual production specifications of inlet ports. In the second part, the same computational code is used to simulate in-cylinder combustion and determine the resulting temperature and heat flux distribution on the cylinder head walls. A comparison is thn carried out between numerical results and experimental measurements and good agreement is obtained.
Batina, John T.
A new computer program has been developed called ASP3D (Advanced Small Perturbation 3D), which solves the small perturbation potential flow equation in an advanced form including mass-consistent surface and trailing wake boundary conditions, and entropy, vorticity, and viscous effects. The purpose of the program is for unsteady aerodynamic and aeroelastic analyses, especially in the nonlinear transonic flight regime. The program exploits the simplicity of stationary Cartesian meshes with the movement or deformation of the configuration under consideration incorporated into the solution algorithm through a planar surface boundary condition. The new ASP3D code is the result of a decade of developmental work on improvements to the small perturbation formulation, performed while the author was employed as a Senior Research Scientist in the Configuration Aerodynamics Branch at the NASA Langley Research Center. The ASP3D code is a significant improvement to the state-of-the-art for transonic aeroelastic analyses over the CAP-TSD code (Computational Aeroelasticity Program Transonic Small Disturbance), which was developed principally by the author in the mid-1980s. The author is in a unique position as the developer of both computer programs to compare, contrast, and ultimately make conclusions regarding the underlying formulations and utility of each code. The paper describes the salient features of the ASP3D code including the rationale for improvements in comparison with CAP-TSD. Numerous results are presented to demonstrate the ASP3D capability. The general conclusion is that the new ASP3D capability is superior to the older CAP-TSD code because of the myriad improvements developed and incorporated.
The tilt-duct vertical takeoff and landing (VTOL) concept has been around since the early 1960s; however, to date the design has never passed the research phase and development phase. Nearly 50 years later, American Dynamics Flight Systems (ADFS) is developing the AD-150, a 2,250lb weight class unmanned aerial vehicle (UAV) configured with rotating ducts on each wingtip. Unlike its predecessor, the Doak VZ-4, the AD-150 features a V tail and wing sweep -- both of which affect the aerodynamic behavior of the aircraft. Because no aircraft of this type has been built and tested, vital aerodynamic research was conducted on the bare airframe behavior (without wingtip ducts). Two weeks of static and dynamic testing were performed on a 3/10th scale model at the University of Maryland's 7' x 10' low speed wind tunnel to facilitate the construction of a nonlinear flight simulator. A total of 70 dynamic tests were performed to obtain damping parameter estimates using the ordinary least squares methodology. Validation, based on agreement between static and dynamic estimates of the pitch and yaw stiffness terms, showed an average percent error of 14.0% and 39.6%, respectively. These inconsistencies were attributed to: large dynamic displacements not encountered during static testing, regressor collinearity, and, while not conclusively proven, differences in static and dynamic boundary layer development. Overall, the damping estimates were consistent and repeatable, with low scatter over a 95% confidence interval. Finally, a basic open loop simulation was executed to demonstrate the instability of the aircraft. As a result, it is recommended that future work be performed to determine trim points and linear models for controls development.
Dippold, Vance F., III
Computational Fluid Dynamics (CFD) simulations were completed for a series of convergent nozzles in participation of the American Institute of Aeronautics and Astronautics (AIAA) 1st Propulsion Aerodynamics Workshop. The simulations were performed using the Wind-US flow solver. Discharge and thrust coefficients were computed for four axisymmetric nozzles with nozzle pressure ratios (NPR) ranging from 1.4 to 7.0. The computed discharge coefficients showed excellent agreement with available experimental data; the computed thrust coefficients captured trends observed in the experimental data, but over-predicted the thrust coefficient by 0.25 to 1.0 percent. Sonic lines were computed for cases with NPR >= 2.0 and agreed well with experimental data for NPR >= 2.5. Simulations were also performed for a 25 deg. conic nozzle bifurcated by a flat plate at NPR = 4.0. The jet plume shock structure was compared with and without the splitter plate to the experimental data. The Wind-US simulations predicted the shock structure well, though lack of grid resolution in the plume reduced the sharpness of the shock waves. Unsteady Reynolds-Averaged Navier-Stokes (URANS) simulations and Detached Eddy Simulations (DES) were performed at NPR = 1.6 for the 25 deg conic nozzle with splitter plate. The simulations predicted vortex shedding from the trailing edge of the splitter plate. However, the vortices of URANS and DES solutions appeared to dissipate earlier than observed experimentally. It is believed that a lack of grid resolution in the region of the vortex shedding may have caused the vortices to break down too soon
Holland, Scott D.; Woods, William C.; Engelund, Walter C.
This paper provides an overview of the experimental aerodynamics test program to ensure mission success for the autonomous flight of the Hyper-X Research Vehicle (HXRV). The HXRV is a 12-ft long, 2700 lb lifting body technology demonstrator designed to flight demonstrate for the first time a fully airframe integrated scramjet propulsion system. Three flights are currently planned, two at Mach 7 and one at Mach 10, beginning in the fall of 2000. The research vehicles will be boosted to the prescribed scramjet engine test point where they will separate from the booster, stabilize, and initiate engine test. Following 5+ seconds of powered flight and 15 seconds of cow-open tares, the cowl will close and the vehicle will fly a controlled deceleration trajectory which includes numerous control doublets for in-flight aerodynamic parameter identification. This paper reviews the preflight testing activities, wind tunnel models, test rationale, risk reduction activities, and sample results from wind tunnel tests supporting the flight trajectory of the HXRV from hypersonic engine test point through subsonic flight termination.
Holland, Scott D.; Woods, William C.; Engelund, Walter C.
This paper provides an overview of the experimental aerodynamics test program to ensure mission success for the autonomous flight of the Hyper-X Research Vehicle (HXRV). The HXRV is a 12-ft long, 2700 lb lifting body technology demonstrator designed to flight demonstrate for the first time a fully airframe integrated scramjet propulsion system. Three flights are currently planned, two at Mach 7 and one at Mach 10, beginning in the fall of 2000. The research vehicles will be boosted to the prescribed scramjet engine test point where they will separate from the booster, stabilize. and initiate engine test. Following 5+ seconds of powered flight and 15 seconds of cowl-open tares, the cowl will close and the vehicle will fly a controlled deceleration trajectory which includes numerous control doublets for in-flight aerodynamic parameter identification. This paper reviews the preflight testing activities, wind tunnel models, test rationale. risk reduction activities, and sample results from wind tunnel tests supporting the flight trajectory of the HXRV from hypersonic engine test point through subsonic flight termination.
The D8 ``double bubble'' is a proposed design for quieter and more efficient domestic passenger aircraft of the Boeing 737 class. It features boundary layer-ingesting engines located under a non-load-bearing π-tail and a lightweight low-sweep wing for flight around Mach 0.7. The D8's wide lifting body is expected to supply 15% of its total lift, while a Boeing 737's fuselage contributes only 8%. The tapering rear of the fuselage is also predicted to experience a negative moment resulting in positive pitch, produce a thicker boundary layer for ingestion by distortion-tolerant engines, and act as a noise shield. To investigate these predictions, unstructured grids generated over a fine surface triangulation using Star-CCM+ are used to model the unpowered D8 with flow conditions mimicking those in the MIT Wright brothers wind tunnel at angles of attack from - 2 to 14 degrees. LAVA, the recently developed Launch Ascent and Vehicle Aerodynamics solver, is used to carry out simulations on an unstructured grid. The results are compared to wind tunnel data, and to data from structured grid simulations using the LAVA, Overflow, and Cart3D solvers. Applied Modeling and Simulation Branch, NASA Advanced Supercomputing Division, funded by New York Space Grant.
Deck, Sébastien; Gand, Fabien; Brunet, Vincent; Ben Khelil, Saloua
This paper provides an up-to-date survey of the use of zonal detached eddy simulations (ZDES) for unsteady civil aircraft applications as a reflection on the stakes and perspectives of the use of hybrid methods in the framework of industrial aerodynamics. The issue of zonal or non-zonal treatment of turbulent flows for engineering applications is discussed. The ZDES method used in this article and based on a fluid problem-dependent zonalization is briefly presented. Some recent landmark achievements for conditions all over the flight envelope are presented, including low-speed (aeroacoustics of high-lift devices and landing gear), cruising (engine–airframe interactions), propulsive jets and off-design (transonic buffet and dive manoeuvres) applications. The implications of such results and remaining challenges in a more global framework are further discussed. PMID:25024411
Deck, Sébastien; Gand, Fabien; Brunet, Vincent; Ben Khelil, Saloua
This paper provides an up-to-date survey of the use of zonal detached eddy simulations (ZDES) for unsteady civil aircraft applications as a reflection on the stakes and perspectives of the use of hybrid methods in the framework of industrial aerodynamics. The issue of zonal or non-zonal treatment of turbulent flows for engineering applications is discussed. The ZDES method used in this article and based on a fluid problem-dependent zonalization is briefly presented. Some recent landmark achievements for conditions all over the flight envelope are presented, including low-speed (aeroacoustics of high-lift devices and landing gear), cruising (engine-airframe interactions), propulsive jets and off-design (transonic buffet and dive manoeuvres) applications. The implications of such results and remaining challenges in a more global framework are further discussed. PMID:25024411
Hartill, W. R.
A hypersonic wind tunnel test method for obtaining credible aerodynamic data on a complete hypersonic vehicle (generic X-24c) with scramjet exhaust flow simulation is described. The general problems of simulating the scramjet exhaust as well as accounting for scramjet inlet flow and vehicle forces are analyzed, and candidate test methods are described and compared. The method selected as most useful makes use of a thrust-minus-drag flow-through balance with a completely metric model. Inlet flow is diverted by a fairing. The incremental effect of the fairing is determined in the testing of two reference models. The net thrust of the scramjet module is an input to be determined in large-scale module tests with scramjet combustion. Force accounting is described, and examples of force component levels are predicted. Compatibility of the test method with candidate wind tunnel facilities is described, and a preliminary model mechanical arrangement drawing is presented. The balance design and performance requirements are described in a detailed specification. Calibration procedures, model instrumentation, and a test plan for the model are outlined.
Enomoto, F.; Keller, P.
The Computer Aided Design (CAD) system's common geometry database was used to generate input for theoretical programs and numerically controlled (NC) tool paths for wind tunnel part fabrication. This eliminates the duplication of work in generating separate geometry databases for each type of analysis. Another advantage is that it reduces the uncertainty due to geometric differences when comparing theoretical aerodynamic data with wind tunnel data. The system was adapted to aerodynamic research by developing programs written in Design Analysis Language (DAL). These programs reduced the amount of time required to construct complex geometries and to generate input for theoretical programs. Certain shortcomings of the Design, Drafting, and Manufacturing (DDM) software limited the effectiveness of these programs and some of the Calma NC software. The complexity of aircraft configurations suggests that more types of surface and curve geometry should be added to the system. Some of these shortcomings may be eliminated as improved versions of DDM are made available.
Carlson, H. W.; Walkley, K. B.
The expanded capabilities for analysis and design of low speed flap systems afforded by recent modifications of an existing computer program is described. The program provides for the simultaneous analysis of up to 25 pairs of leading-edge and trailing-edge flap deflection schedules. Among other new features of the program are a revised attainable thrust estimation method to provide more accurate predictions for low Mach numbers, and a choice of three options for estimation of leading-edge separation vortex flow effects. Comparison of program results with low speed experimental data for an arrow wing supersonic cruise configuration with leading-edge and trailing-edge flaps showed good agreement over most of the range of flap deflections. Other force data comparisons and an independent study of airfoil and wing pressure distributions indicated that wind-tunnel measurements of the aerodynamic performance of twisted and cambered wings and wings with leading-edge flaps can be very sensitive to Reynolds number effects.
VSAERO is a computer program used to predict the nonlinear aerodynamic characteristics of arbitrary three-dimensional configurations in subsonic flow. Nonlinear effects of vortex separation and vortex surface interaction are treated in an iterative wake-shape calculation procedure, while the effects of viscosity are treated in an iterative loop coupling potential-flow and integral boundary-layer calculations. The program employs a surface singularity panel method using quadrilateral panels on which doublet and source singularities are distributed in a piecewise constant form. This user's manual provides a brief overview of the mathematical model, instructions for configuration modeling and a description of the input and output data. A listing of a sample case is included.
Schroer, Bernard J.; Tseng, Fan T.; Zhang, Shou X.; Dwan, Wen S.
The objective of automatic programming is to improve the overall environment for describing the program. This improved environment is realized by a reduction in the amount of detail that the programmer needs to know and is exposed to. Furthermore, this improved environment is achieved by a specification language that is more natural to the user's problem domain and to the user's way of thinking and looking at the problem. The goal of this research is to apply the concepts of automatic programming (AP) to modeling discrete event simulation system. Specific emphasis is on the design and development of simulation tools to assist the modeler define or construct a model of the system and to then automatically write the corresponding simulation code in the target simulation language, GPSS/PC. A related goal is to evaluate the feasibility of various languages for constructing automatic programming simulation tools.
Bibb, Karen L.; Alter, Stephen J.; Mcdaniel, Ryan D.
The aerodynamic effects of the recession of the ablative thermal protection system for the Orion Command Module of the Crew Exploration Vehicle are important for the vehicle guidance. At the present time, the aerodynamic effects of recession being handled within the Orion aerodynamic database indirectly with an additional safety factor placed on the uncertainty bounds. This study is an initial attempt to quantify the effects for a particular set of recessed geometry shapes, in order to provide more rigorous analysis for managing recession effects within the aerodynamic database. The aerodynamic forces and moments for the baseline and recessed geometries were computed at several trajectory points using multiple CFD codes, both viscous and inviscid. The resulting aerodynamics for the baseline and recessed geometries were compared. The forces (lift, drag) show negligible differences between baseline and recessed geometries. Generally, the moments show a difference between baseline and recessed geometries that correlates with the maximum amount of recession of the geometry. The difference between the pitching moments for the baseline and recessed geometries increases as Mach number decreases (and the recession is greater), and reach a value of -0.0026 for the lowest Mach number. The change in trim angle of attack increases from approx. 0.5deg at M = 28.7 to approx. 1.3deg at M = 6, and is consistent with a previous analysis with a lower fidelity engineering tool. This correlation of the present results with the engineering tool results supports the continued use of the engineering tool for future work. The present analysis suggests there does not need to be an uncertainty due to recession in the Orion aerodynamic database for the force quantities. The magnitude of the change in pitching moment due to recession is large enough to warrant inclusion in the aerodynamic database. An increment in the uncertainty for pitching moment could be calculated from these results and
Sullivan, Robert; Greeley, Ronald
This study evaluates how well values of aerodynamic surface roughness, z sub 0, measured over scale models in wind tunnels correlate with values of z sub 0 measured at full scale in the field. A field experiment was conducted in which values of z sub 0 and u* (wind friction speed) were measured over three arrays of non-erodible roughness elements on a dry lake bed. Wind profiles were measured by ten anemometers on a 15 m mast under thermally neutral atmospheric conditions. Values of z sub 0 increased from .00014 m (dry lake bed only) to .026 m with increasing roughness element density. The three roughness element arrays were simulated at 1/10 and 1/20 scale in an open-circuit atmospheric boundary-layer wind tunnel. Velocities were measured with a boundary-layer pitot-tube rake from the same relative position within the scale model arrays as the anemometers were relative to the field arrays. Each array at each scale was sampled three times at five freestream velocities. Average values of z sub 0 for each model array at each scale were compared with full-scale values of z sub 0 obtained in the field. The field vs. wind tunnel correspondence of z sub 0 is found to be z sub 0 field = 0.2661 x (z sub(0 model) x scale(exp -1))exp .8159.
Bailey, F. R.; Hathaway, A. W.
Critical to the advancement of computational aerodynamics capability is the ability to simulate flows about three-dimensional configurations that contain both compressible and viscous effects, including turbulence and flow separation at high Reynolds numbers. Analyses were conducted of two solution techniques for solving the Reynolds averaged Navier-Stokes equations describing the mean motion of a turbulent flow with certain terms involving the transport of turbulent momentum and energy modeled by auxiliary equations. The first solution technique is an implicit approximate factorization finite-difference scheme applied to three-dimensional flows that avoids the restrictive stability conditions when small grid spacing is used. The approximate factorization reduces the solution process to a sequence of three one-dimensional problems with easily inverted matrices. The second technique is a hybrid explicit/implicit finite-difference scheme which is also factored and applied to three-dimensional flows. Both methods are applicable to problems with highly distorted grids and a variety of boundary conditions and turbulence models.
Donohue, Paul F.
The results of an aerodynamic performance evaluation of the National Aeronautics and Space Administration (NASA)/Ames Research Center Advanced Concepts Flight Simulator (ACFS), conducted in association with the Navy-NASA Joint Institute of Aeronautics, are presented. The ACFS is a full-mission flight simulator which provides an excellent platform for the critical evaluation of emerging flight systems and aircrew performance. The propulsion and flight dynamics models were evaluated using classical flight test techniques. The aerodynamic performance model of the ACFS was found to realistically represent that of current day, medium range transport aircraft. Recommendations are provided to enhance the capabilities of the ACFS to a level forecast for 1995 transport aircraft. The graphical and tabular results of this study will establish a performance section of the ACFS Operation's Manual.
Medan, R. T.; Ray, K. S.
A description of and users manual for a USA FORTRAN IV computer program which plots the planform and control points of a wing are presented. The program also plots some of the configuration data such as the aspect ratio. The planform data is stored on a disc file which is created by a geometry program. This program, the geometry program, and several other programs are used together in the analysis of lifting, thin wings in steady, subsonic flow according to a kernel function lifting surface theory.
Matha, D.; Schlipf, M.; Cordle, A.; Pereira, R.; Jonkman, J.
This paper presents the current major modeling challenges for floating offshore wind turbine design tools and describes aerodynamic and hydrodynamic effects due to rotor and platform motions and usage of non-slender support structures.
The VSAERO low order panel method formulation is described for the calculation of subsonic aerodynamic characteristics of general configurations. The method is based on piecewise constant doublet and source singularities. Two forms of the internal Dirichlet boundary condition are discussed and the source distribution is determined by the external Neumann boundary condition. A number of basic test cases are examined. Calculations are compared with higher order solutions for a number of cases. It is demonstrated that for comparable density of control points where the boundary conditions are satisfied, the low order method gives comparable accuracy to the higher order solutions. It is also shown that problems associated with some earlier low order panel methods, e.g., leakage in internal flows and junctions and also poor trailing edge solutions, do not appear for the present method. Further, the application of the Kutta conditions is extremely simple; no extra equation or trailing edge velocity point is required. The method has very low computing costs and this has made it practical for application to nonlinear problems requiring iterative solutions for wake shape and surface boundary layer effects.
Kaylor, J. T.; Powell, L. F.; Powell, R. W.
The automatic reentry flight dynamics simulator, a nonlinear, six-degree-of-freedom simulation, digital computer program, has been developed. The program includes a rotating, oblate earth model for accurate navigation calculations and contains adjustable gains on the aerodynamic stability and control parameters. This program uses a real-time simulation system and is designed to examine entries of vehicles which have constant mass properties whose attitudes are controlled by both aerodynamic surfaces and reaction control thrusters, and which have automatic guidance and control systems. The program has been used to study the space shuttle orbiter entry. This report includes descriptions of the equations of motion used, the control and guidance schemes that were implemented, the program flow and operation, and the hardware involved.
Describes a sequence of five computer programs (listings for Apple II available from author) on tactic responses (oriented movement of a cell, cell group, or whole organism in reponse to stimuli). The simulation programs are useful in helping students examine mechanisms at work in real organisms. (JN)
Land, Kenneth P.; Best, Robert E.; Steel, Douglas J.; Gadd, William C.
Dynamic Environment Communications Alalysis Testbed (DECAT) computer program is modular simulation program. Computes effects of motion, antenna radiation patterns, noise, interference, and other phenomena. Flexibility enables users to analyze many communications scenarios quickly and easily, eliminating need for users to create specific computer programs. Users create simulations involving any number of vehicles, receivers, transmitters, and antennas via graphical user interface (GUI). DECAT GUI implemented by use of software tool called "Transportable Applications Environment Plus" (TAE Plus). Written in C language. Graphical user interface requires TAE Plus, Version 5.1 package available from COSMIC (GSC-13463).
Deadrick, F.J.; Hawke, R.S.; Scudder, J.K.
A computer simulation code at the Lawrence Livermore National Laboratory (LLNL) to predict the performance of a railgun electromagnetic accelerator has been developed and validated. The code, called MAGRAC (MAGnetic Railgun ACcelerator), models the performance of a railgun driven by a magnetic flux compression current generator (MFCG). The MAGRAC code employs a time-step solution of the nonlinear time-varying element railgun circuit to determine rail currents. From the rail currents, the projectile acceleration, velocity, and position are found. The MAGRAC code was validated through a series of eight railgun tests conducted jointly with the Los Alamos National Laboratory. The formulation of the MAGRAC railgun model is described and the predicted current waveforms compared with those obtained from full-scale experiments.
Klose, G. J.
The purpose of the VEEP simulation program was to: (1) predict vehicle fuel economy and relative emissions over any specified driving cycle; (2) calculate various measures of vehicle performance (acceleration, passing manuevers, gradeability, top speed), and (3) give information on the various categories of energy dissipation (rolling friction, aerodynamics, accessories, inertial effects, component inefficiences, etc.). The vehicle is described based on detailed subsystem information and numerical parameters characterizing the components of a wide variety of self-propelled vehicles. Conventionally arranged heat engine powered automobiles were emphasized, but with consideration in the design toward the requirement of other types of vehicles.
Gustafson, Karl; Leben, Robert
Dragonflies are seen to hover and dart, seemingly at will and in remarkably nimble fashion, with great bursts of speed and effectively discontinuous changes of direction. In their short lives, their gossamer flight provides us with glimpses of an aerodynamics of almost extraterrestrial quality. Here we present the first computer simulations of such aerodynamics.
Dillenius, Marnix F. E.
Program LRCDM2 was developed for supersonic missiles with axisymmetric bodies and up to two finned sections. Predicted are pressure distributions and loads acting on a complete configuration including effects of body separated flow vorticity and fin-edge vortices. The computer program is based on supersonic panelling and line singularity methods coupled with vortex tracking theory. Effects of afterbody shed vorticity on the afterbody and tail-fin pressure distributions can be optionally treated by companion program BDYSHD. Preliminary versions of combined shock expansion/linear theory and Newtonian/linear theory have been implemented as optional pressure calculation methods to extend the Mach number and angle-of-attack ranges of applicability into the nonlinear supersonic flow regime. Comparisons between program results and experimental data are given for a triform tail-finned configuration and for a canard controlled configuration with a long afterbody for Mach numbers up to 2.5. Initial tests of the nonlinear/linear theory approaches show good agreement for pressures acting on a rectangular wing and a delta wing with attached shocks for Mach numbers up to 4.6 and angles of attack up to 20 degrees.
Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. Comprehensive data report. Volume 2: Tabulated aerodynamic data book 2
Nelson, D. P.
Tabulated aerodynamic data from coannular nozzle performance tests are given for test runs 26 through 37. The data include nozzle thrust coefficient parameters, nozzle discharge coefficients, and static pressure tap measurements.
Quinn, Robert D.; Gong, Leslie
A real-time heating algorithm was derived and installed on the Ames Research Center Dryden Flight Research Facility real-time flight simulator. This program can calculate two- and three-dimensional stagnation point surface heating rates and surface temperatures. The two-dimensional calculations can be made with or without leading-edge sweep. In addition, upper and lower surface heating rates and surface temperatures for flat plates, wedges, and cones can be calculated. Laminar or turbulent heating can be calculated, with boundary-layer transition made a function of free-stream Reynolds number and free-stream Mach number. Real-time heating rates and surface temperatures calculated for a generic hypersonic vehicle are presented and compared with more exact values computed by a batch aeroheating program. As these comparisons show, the heating algorithm used on the flight simulator calculates surface heating rates and temperatures well within the accuracy required to evaluate flight profiles for acceptable heating trajectories.
The Water Quality Analysis Simulation Program (WASP6), an enhancement of the original WASP (Di Toro et al., 1983; Connolly and Winfield,1984; Ambrose, R.B. et al.,1988). This model helps users interpret and predict water quality responses to natural phenomena and man-made polluti...
Persing, T. Ray; Bellish, Christine A.; Brandon, Jay; Kenney, P. Sean; Carzoo, Susan; Buttrill, Catherine; Guenther, Arlene
Several aircraft airframe modeling approaches are currently being used in the DoD community for acquisition, threat evaluation, training, and other purposes. To date there has been no clear empirical study of the impact of airframe simulation fidelity on piloted real-time aircraft simulation study results, or when use of a particular level of fidelity is indicated. This paper documents a series of piloted simulation studies using three different levels of airframe model fidelity. This study was conducted using the NASA Langley Differential Maneuvering Simulator. Evaluations were conducted with three pilots for scenarios requiring extensive maneuvering of the airplanes during air combat. In many cases, a low-fidelity modified point-mass model may be sufficient to evaluate the combat effectiveness of the aircraft. However, in cases where high angle-of-attack flying qualities and aerodynamic performance are a factor or when precision tracking ability of the aircraft must be represented, use of high-fidelity models is indicated.
Schroer, Bernard J.; Tseng, Fan T.; Zhang, Shou X.; Dwan, Wen S.
The concepts of software engineering were used to improve the simulation modeling environment. Emphasis was placed on the application of an element of rapid prototyping, or automatic programming, to assist the modeler define the problem specification. Then, once the problem specification has been defined, an automatic code generator is used to write the simulation code. The following two domains were selected for evaluating the concepts of software engineering for discrete event simulation: manufacturing domain and a spacecraft countdown network sequence. The specific tasks were to: (1) define the software requirements for a graphical user interface to the Automatic Manufacturing Programming System (AMPS) system; (2) develop a graphical user interface for AMPS; and (3) compare the AMPS graphical interface with the AMPS interactive user interface.
Petrarca, J. R.; Harrison, B. A.; Redman, M. C.; Rowe, W. S.
A digital computer program was developed to calculate unsteady loadings caused by motions of lifting surfaces with leading edge and trailing edge controls based on the subsonic kernel function approach. The pressure singularities at hinge line and side edges were extracted analytically as a preliminary step to solving the integral equation of collocation. The program calculates generalized aerodynamic forces for user supplied deflection modes. Optional intermediate output includes pressure at an array of points, and sectional generalized forces. From one to six controls on the half span can be accomodated.
Campbell, Bryan A.; Bezos, Gaudy M.; Dunham, R. Earl, Jr.; Melson, W. Edward, Jr.
One of the necessary areas of consideration for outdoor heavy rain testing is the effect of wind on both the simulated rain field and the quality and repeatability of the aerodynamic data. This paper discusses the data acquisition and subsequent reduction to nondimensional coefficients of lift and drag, with the appropriate correction for wind and rain field. Sample force data showing these effects are presented, along with estimates for accuracy and repeatability. The capability to produce high-quality data for rain drop size distribution using photographic and computerized image processing techniques was developed. Sample photographs depicting rain drop size are shown.
Crouse, J. E.; Gorrell, W. T.
A code for computing the aerodynamic design of a multistage axial-flow compressor and, if desired, the associated blading geometry input for internal flow analysis codes is presented. Compressible flow, which is assumed to be steady and axisymmetric, is the basis for a two-dimensional solution in the meridional plane with viscous effects modeled by pressure loss coefficients and boundary layer blockage. The radial equation of motion and the continuity equation are solved with the streamline curvature method on calculation stations outside the blade rows. The annulus profile, mass flow, pressure ratio, and rotative speed are input. A number of other input parameters specify and control the blade row aerodynamics and geometry. In particular, blade element centerlines and thicknesses can be specified with fourth degree polynomials for two segments. The output includes a detailed aerodynamic solution and, if desired, blading coordinates that can be used for internal flow analysis codes.
Bland, M. P.; Fajfar, B.; Konsewicz, R. K.
Simulation data are reported for the purpose of programming the flight simulator for advanced aircraft for tests of the lift/cruise fan V/STOL Research Technology Aircraft. These simulation tests are to provide insight into problem areas which are encountered in operational use of the aircraft. A mathematical model is defined in sufficient detail to represent all the necessary pertinent aircraft and system characteristics. The model includes the capability to simulate two basic versions of an aircraft propulsion system: (1) the gas coupled configuration which uses insulated air ducts to transmit power between gas generators and fans in the form of high energy engine exhaust and (2) the mechanically coupled power system which uses shafts, clutches, and gearboxes for power transmittal. Both configurations are modeled such that the simulation can include vertical as well as rolling takeoff and landing, hover, powered lift flight, aerodynamic flight, and the transition between powered lift and aerodynamic flight.
Bragg, M. B.
An experimental study was conducted in the Ohio State University subsonic wind tunnel to measure the detailed aerodynamic characteristics of an airfoil with a simulated glaze ice accretion. A NACA 0012 model with interchangeable leading edges and pressure taps every one percent chord was used. Surface pressure and wake data were taken on the airfoil clean, with forced transition and with a simulated glaze ice shape. Lift and drag penalties due to the ice shape were found and the surface pressure clearly showed that large separation bubbles were present. Both total pressure and split-film probes were used to measure velocity profiles, both for the clean model and for the model with a simulated ice accretion. A large region of flow separation was seen in the velocity profiles and was correlated to the pressure measurements. Clean airfoil data were found to compare well to existing airfoil analysis methods.
Srivastava, R.; Reddy, T. S. R.
The program DuctE3D is used for steady or unsteady aerodynamic and aeroelastic analysis of ducted fans. This guide describes the input data required and the output files generated, in using DuctE3D. The analysis solves three dimensional unsteady, compressible Euler equations to obtain the aerodynamic forces. A normal mode structural analysis is used to obtain the aeroelastic equations, which are solved using either the time domain or the frequency domain solution method. Sample input and output files are included in this guide for steady aerodynamic analysis and aeroelastic analysis of an isolated fan row.
Raj, P.; Long, L. N.
The current capabilities and the future plans for a three dimensional Euler Aerodynamic Method are described. The basic solution algorithm is based on the finite volume, Runge-Kutta pseudo-time-stepping scheme of FLO-57. Several modifications to improve accuracy and computational efficiency were incorporated and others are being investigated. The computer code is used to analyze a cropped delta wing at 0.6 Mach number and an arrow wing at 0.85 Mach number. Computed aerodynamic parameters are compared with experimental data. In all cases, the configuration is impulsively started and no Kutta condition is applied at sharp edges. The results indicate that with additional development and validation, the present method will be a useful tool for engineering analysis of high speed aircraft.
Keith, T. G., Jr.; Afjeh, A. A.; Jeng, D. R.; White, J. A.
A description of a computer program entitled VORTEX that may be used to determine the aerodynamic performance of horizontal axis wind turbines is given. The computer code implements a vortex method from finite span wind theory and determines the induced velocity at the rotor disk by integrating the Biot-Savart law. It is assumed that the trailing helical vortex filaments form a wake of constant diameter (the rigid wake assumption) and travel downstream at the free stream velocity. The program can handle rotors having any number of blades which may be arbitrarily shaped and twisted. Many numerical details associated with the program are presented. A complete listing of the program is provided and all program variables are defined. An example problem illustrating input and output characteristics is solved.
Spearman, M. Leroy
The achievement of flight with manned, powered, heavier-than-air aircraft in 1903 marked the beginning of a new era in the means of transportation. A special advantage for aircraft was in speed. However, when an aircraft penetrates the air at very high speeds, the disturbed air is compressed and there are changes in the density, pressure and temperature of the air. These compressibility effects change the aerodynamic characteristics of an aircraft and introduce problems in drag, stability and control. Many aircraft designed in the post-World War II era were plagued with the effects of compressibility. Accordingly, the study of the aerodynamic behavior of aircraft, spacecraft and missiles at high-speed became a major part of the research activity of the NACA/NASA. The intent of the research was to determine the causes and provide some solutions for the aerodynamic problems resulting from the effects of compressibility. The purpose of this paper is to review some of the high-speed aerodynamic research work conducted at the Langley Research Center from the viewpoint of the author who has been active in much of the effort.
Vadyak, Joseph; Shrewsbury, George E.; Narramore, Jim C.; Montry, Gary; Holst, Terry; Kwak, Dochan (Technical Monitor)
The paper will describe the Development of a general three-dimensional multiple grid zone Navier-Stokes flowfield simulation program (ENS3D-MPP) designed for efficient execution on the Intel Paragon Massively Parallel Processor (MPP) supercomputer, and the subsequent application of this method to the prediction of the viscous flowfield about the V-22 Osprey tiltrotor vehicle. The flowfield simulation code solves the thin Layer or full Navier-Stoke's equation - for viscous flow modeling, or the Euler equations for inviscid flow modeling on a structured multi-zone mesh. In the present paper only viscous simulations will be shown. The governing difference equations are solved using a time marching implicit approximate factorization method with either TVD upwind or central differencing used for the convective terms and central differencing used for the viscous diffusion terms. Steady state or Lime accurate solutions can be calculated. The present paper will focus on steady state applications, although time accurate solution analysis is the ultimate goal of this effort. Laminar viscosity is calculated using Sutherland's law and the Baldwin-Lomax two layer algebraic turbulence model is used to compute the eddy viscosity. The Simulation method uses an arbitrary block, curvilinear grid topology. An automatic grid adaption scheme is incorporated which concentrates grid points in high density gradient regions. A variety of user-specified boundary conditions are available. This paper will present the application of the scalable and superscalable versions to the steady state viscous flow analysis of the V-22 Osprey using a multiple zone global mesh. The mesh consists of a series of sheared cartesian grid blocks with polar grids embedded within to better simulate the wing tip mounted nacelle. MPP solutions will be shown in comparison to equivalent Cray C-90 results and also in comparison to experimental data. Discussions on meshing considerations, wall clock execution time
Bragg, Michael B.
This is the second volume of a report documenting the effect of simulated ice accretion on the aerodynamic performance of a NACA 0012 airfoil. Both an experimentally measured and a computer generated ice shape are studied. The purpose of this report is to present the results of the measurements, not an analysis of the data. Surface pressure, integrated lift and pitching moment data are presented as well as drag from a wake survey. A split hot film probe was used to document the flow-field about the airfoil with simulated ice. Data in the separation bubbles, reattached boundary layer and wake are presented. Both tabulated and graphical data are presented in the paper. The data are also available on computer disk for easy access.
Brooks, B.R.; Brooks, C.L.; MacKerell, A.D.; Nilsson, L.; Petrella, R.J.; Roux, B.; Won, Y.; Archontis, G.; Bartels, C.; Boresch, S.; Caflisch, A.; Caves, L.; Cui, Q.; Dinner, A.R.; Feig, M.; Fischer, S.; Gao, J.; Hodoscek, M.; Im, W.; Kuczera, K.; Lazaridis, T.; Ma, J.; Ovchinnikov, V.; Paci, E.; Pastor, R.W.; Post, C.B.; Pu, J.Z.; Schaefer, M.; Tidor, B.; Venable, R. M.; Woodcock, H. L.; Wu, X.; Yang, W.; York, D.M.; Karplus, M.
CHARMM (Chemistry at HARvard Molecular Mechanics) is a highly versatile and widely used molecular simulation program. It has been developed over the last three decades with a primary focus on molecules of biological interest, including proteins, peptides, lipids, nucleic acids, carbohydrates and small molecule ligands, as they occur in solution, crystals, and membrane environments. For the study of such systems, the program provides a large suite of computational tools that include numerous conformational and path sampling methods, free energy estimators, molecular minimization, dynamics, and analysis techniques, and model-building capabilities. In addition, the CHARMM program is applicable to problems involving a much broader class of many-particle systems. Calculations with CHARMM can be performed using a number of different energy functions and models, from mixed quantum mechanical-molecular mechanical force fields, to all-atom classical potential energy functions with explicit solvent and various boundary conditions, to implicit solvent and membrane models. The program has been ported to numerous platforms in both serial and parallel architectures. This paper provides an overview of the program as it exists today with an emphasis on developments since the publication of the original CHARMM paper in 1983. PMID:19444816
Xiao, Haosu; Zuo, Baojun; Tian, Yi; Zhang, Wang; Hao, Chenglong; Liu, Chaofeng; Li, Qi; Li, Fan; Zhang, Li; Fan, Zhigang
We investigated the joint influences exerted by the nonuniform aerodynamic flow field surrounding the optical dome and the aerodynamic heating of the dome on imaging quality degradation of an airborne optical system. The Spalart-Allmaras model provided by FLUENT was used for flow computations. The fourth-order Runge-Kutta algorithm based ray tracing program was used to simulate optical transmission through the aerodynamic flow field and the dome. Four kinds of imaging quality evaluation parameters were presented: wave aberration of the exit pupil, point spread function, encircled energy, and modulation transfer function. The results show that the aero-optical disturbance of the aerodynamic flow field and the aerodynamic heating of the dome significantly affect the imaging quality of an airborne optical system. PMID:23262604
Castner, Raymond; Wyzykowski, John; Chiapetta, Santo; Adamczyk, John
A comprehensive test program was performed in the Propulsion Systems Laboratory at the NASA Glenn Research Center, Cleveland Ohio using a highly instrumented Pratt and Whitney Canada PW 545 turbofan engine. A key objective of this program was the development of a high-altitude database on small, high-bypass ratio engine performance and operability. In particular, the program documents the impact of altitude (Reynolds Number) on the aero-performance of the low-pressure turbine (fan turbine). A second objective was to assess the ability of a state-of-the-art CFD code to predict the effect of Reynolds number on the efficiency of the low-pressure turbine. CFD simulation performed prior and after the engine tests will be presented and discussed. Key findings are the ability of a state-of-the art CFD code to accurately predict the impact of Reynolds Number on the efficiency and flow capacity of the low-pressure turbine. In addition the CFD simulations showed the turbulent intensity exiting the low-pressure turbine to be high (9%). The level is consistent with measurements taken within an engine.
Dutheil, Sylvain; Pibarot, Julien; Tran, Dac; Vallee, Jean-Jacques; Tribot, Jean-Pierre
With the aim of placing Europe among the world's space players in the strategic area of atmospheric re-entry, several studies on experimental vehicle concepts and improvements of critical re-entry technologies have paved the way for the flight of an experimental space craft. The successful flight of the Intermediate eXperimental Vehicle (IXV), under ESA's Future Launchers Preparatory Programme (FLPP), is definitively a significant step forward from the Atmospheric Reentry Demonstrator flight (1998), establishing Europe as a key player in this field. The IXV project objectives were the design, development, manufacture and ground and flight verification of an autonomous European lifting and aerodynamically controlled reentry system, which is highly flexible and maneuverable. The paper presents, the role of aerodynamics aerothermodynamics as part of the key technologies for designing an atmospheric re-entry spacecraft and securing a successful flight.
Levy, D. W.; Finn, P.; Roskam, J.
The use of aerodynamic pressure differentials to position a control surface is evaluated. The system is a differential pressure command loop, analogous to a position command loop, where the surface is commanded to move until a desired differential pressure across the surface is achieved. This type of control is more direct and accurate because it is the differential pressure which causes the control forces and moments. A frequency response test was performed in a low speed wind tunnel to measure the performance of the system. Both pressure and position feedback were tested. The pressure feedback performed as well as position feedback implying that the actuator, with a break frequency on the order of 10 Rad/sec, was the limiting component. Theoretical considerations indicate that aerodynamic lags will not appear below frequencies of 50 Rad/sec, or higher.
Nasir, Rizal E. M.; Ali, Zurriati; Kuntjoro, Wahyu; Wisnoe, Wirachman
Previous wind tunnel test has proven the improved aerodynamic charasteristics of Baseline-II E-2 Blended Wing-Body (BWB) aircraft studied in Universiti Teknologi Mara. The E-2 is a version of Baseline-II BWB with modified outer wing and larger canard, solely-designed to gain favourable longitudinal static stability during flight. This paper highlights some results from current investigation on the said aircraft via computational fluid dynamics simulation as a mean to validate the wind tunnel test results. The simulation is conducted based on standard one-equation turbulence, Spalart-Allmaras model with polyhedral mesh. The ambience of the flight simulation is made based on similar ambience of wind tunnel test. The simulation shows lift, drag and moment results to be near the values found in wind tunnel test but only within angles of attack where the lift change is linear. Beyond the linear region, clear differences between computational simulation and wind tunnel test results are observed. It is recommended that different type of mathematical model be used to simulate flight conditions beyond linear lift region.
Chamberlain, Robert G.; Zendejas, Silvino; Malhotra, Shan
Costs and profits estimated for models based on user inputs. Standard Assembly-line Manufacturing Industry Simulation (SAMIS) program generalized so useful for production-line manufacturing companies. Provides accurate and reliable means of comparing alternative manufacturing processes. Used to assess impact of changes in financial parameters as cost of resources and services, inflation rates, interest rates, tax policies, and required rate of return of equity. Most important capability is ability to estimate prices manufacturer would have to receive for its products to recover all of costs of production and make specified profit. Written in TURBO PASCAL.
Bragg, Michael B.
Two semispan wings, one with a rectangular planform and one with 30 degrees of leading edge sweep were tested. Both had a NACA 0012 airfoil section, and both were tested clean and with simulated glaze ice shapes on their leading edges. Several surface roughness were tested. Each model geometry is documented and each surface roughness is explained. Aerodynamic performance of the wing in the form of sectional lift and integrated three-dimensional lift is documented through pressure measurements obtained from rows of surface pressure taps placed at five span locations on the wing. For the rectangular wing, sectional drag near the midspan is obtained from wake total pressure profiles. The data is presented in tabular and graphical form and is also available on computer disk.
Nelson, D. P.
Wind tunnel model tests were conducted to demonstrate the aerodynamic performance improvements of a refined actuated inlet ejector nozzle. Models of approximately one-tenth scale were configured to simulate nozzle operation at takeoff, subsonic cruise, transonic cruise and supersonic cruise. Variations of model components provided a performance evaluation of ejector inlet and exit area, forebody boattail angle and ejector inlet operation in the open and closed mode. Approximately 700 data points were acquired at Mach numbers of 0, 0.36, 0.9, 1.2, and 2.0 for a wide range of nozzle flow conditions. Results show that relative to two ejector nozzles previously tested performance was improved significantly at takeoff and subsonic cruise performance, a C sub f of 0.982, was attained equal to the high performance of the previous tests. The established advanced supersonic transport propulsion study performance goals were met or closely approached at takeoff and supersonic cruise.
Nielsen, Jack N.
The fundamental aerodynamics of slender bodies is examined in the reprint edition of an introductory textbook originally published in 1960. Chapters are devoted to the formulas commonly used in missile aerodynamics; slender-body theory at supersonic and subsonic speeds; vortices in viscid and inviscid flow; wing-body interference; downwash, sidewash, and the wake; wing-tail interference; aerodynamic controls; pressure foredrag, base drag, and skin friction; and stability derivatives. Diagrams, graphs, tables of terms and formulas are provided.
Mehrotra, S. C.; Lan, C. E.
The necessary information for using a computer program to predict distributed and total aerodynamic characteristics for low aspect ratio wings with partial leading-edge separation is presented. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the Quasi-Vortex-Lattice method. The leading edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at midpoints to satisfy the force free condition. The wake behind the trailing edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading and trailing edges. The program is restricted to delta wings with zero thickness and no camber. It is written in FORTRAN language and runs on CDC 6600 computer.
Choo, Yung K.; Vickerman, Mary B.
SmaggIce 2D version 1.8 is a software toolkit developed at the NASA Glenn Research Center that consists of tools for modeling the geometry of and generating the grids for clean and iced airfoils. Plans call for the completed SmaggIce 2D version 2.0 to streamline the entire aerodynamic simulation process--the characterization and modeling of ice shapes, grid generation, and flow simulation--and to be closely coupled with the public-domain application flow solver, WIND. Grid generated using version 1.8, however, can be used by other flow solvers. SmaggIce 2D will help researchers and engineers study the effects of ice accretion on airfoil performance, which is difficult to do with existing software tools because of complex ice shapes. Using SmaggIce 2D, when fully developed, to simulate flow over an iced airfoil will help to reduce the cost of performing flight and wind-tunnel tests for certifying aircraft in natural and simulated icing conditions.
Majumdar, Alok Kumar (Inventor); Bailey, John W. (Inventor); Schallhorn, Paul Alan (Inventor); Steadman, Todd E. (Inventor)
A general purpose program implemented on a computer analyzes steady state and transient flow in a complex fluid network, modeling phase changes, compressibility, mixture thermodynamics and external body forces such as gravity and centrifugal force. A preprocessor provides for the inter- active development of a fluid network simulation having nodes and branches. Mass, energy, and specie conservation equations are solved at the nodes, and momentum conservation equations are solved in the branches. Contained herein are subroutines for computing "real fluid" thermodynamic and thermophysical properties for 12 fluids, and a number of different source options are provided for model- ing momentum sources or sinks in the branches. The system of equations describing the fluid network is solved by a hybrid numerical method that is a combination of the Newton-Raphson and successive substitution methods. Application and verification of this invention are provided through an example problem, which demonstrates that the predictions of the present invention compare most reasonably with test data.
GRASP (General Reactive Atomistic Simulation Program) is primarily intended as a molecular dynamics package for complex force fields, The code is designed to provide good performance for large systems, either in parallel or serial execution mode, The primary purpose of the code is to realistically represent the structural and dynamic properties of large number of atoms on timescales ranging from picoseconds up to a microsecond. Typically the atoms form a representative sample of some material,more » such as an interface between polycrystalline silicon and amorphous silica. GRASP differs from other parallel molecular dynamics codes primarily due to its ability to handle relatively complicated interaction potentials and its ability to use more than one interaction potential in a single simulation. Most of the computational effort goes into the calculation of interatomic forces, which depend in a complicated way on the positions of all the atoms. The forces are used to integrate the equations of motion forward in time using the so-called velocity Verlet integration scheme. Alternatively, the forces can be used to find a minimum energy configuration, in which case a modified steepest descent algorithm is used.« less
Tiffany, Sherwood H.; Adams, William M., Jr.
The approximation of unsteady generalized aerodynamic forces in the equations of motion of a flexible aircraft are discussed. Two methods of formulating these approximations are extended to include the same flexibility in constraining the approximations and the same methodology in optimizing nonlinear parameters as another currently used extended least-squares method. Optimal selection of nonlinear parameters is made in each of the three methods by use of the same nonlinear, nongradient optimizer. The objective of the nonlinear optimization is to obtain rational approximations to the unsteady aerodynamics whose state-space realization is lower order than that required when no optimization of the nonlinear terms is performed. The free linear parameters are determined using the least-squares matrix techniques of a Lagrange multiplier formulation of an objective function which incorporates selected linear equality constraints. State-space mathematical models resulting from different approaches are described and results are presented that show comparative evaluations from application of each of the extended methods to a numerical example.
Liu, Yi; Sankar, Lakshmi N.; Englar, Robert J.; Ahuja, Krishan K.
The aerodynamic characteristics of a Circulation Control Wing (CCW) airfoil have been numerically investigated, and comparisons with experimental data have been made. The configuration chosen was a supercritical airfoil with a 30 degree dual-radius CCW flap. Steady and pulsed jet calculations were performed. It was found that the use of steady jets, even at very small mass flow rates, yielded a lift coefficient that is comparable or superior to conventional high-lift systems. The attached flow over the flap also gave rise to lower drag coefficients, and high L/D ratios. Pulsed jets with a 50% duty cycle were also studied. It was found that they were effective in generating lift at lower reduced mass flow rates compared to a steady jet, provided the pulse frequency was sufficiently high. This benefit was attributable to the fact that the momentum coefficient of the pulsed jet, during the portions of the cycle when the jet was on, was typically twice as much as that of a steady jet.
Hasan, Mohammad Khatim; Sulaiman, Jumat; Karim, Samsul Arifin Abdul
This paper discussed the used of an open source sofware called Scilab to develop a heat simulator. In this paper, heat equation was used to simulate heat behavior in an object. The simulator was developed using finite difference method. Numerical experiment output show that Scilab can produce a good heat behavior simulation with marvellous visual output with only developing simple computer code.
Edmunson, J.; Betts, W.; Rickman, D.; McLemore, C.; Fikes, J.; Stoeser, D.; Wilson, S.; Schrader, C.
Lunar regolith simulant production is absolutely critical to returning man to the Moon. Regolith simulant is used to test hardware exposed to the lunar surface environment, simulate health risks to astronauts, practice in situ resource utilization (ISRU) techniques, and evaluate dust mitigation strategies. Lunar regolith simulant design, production process, and management is a cooperative venture between members of the NASA Marshall Space Flight Center (MSFC) and the U.S. Geological Survey (USGS). The MSFC simulant team is a satellite of the Dust group based at Glenn Research Center. The goals of the cooperative group are to (1) reproduce characteristics of lunar regolith using simulants, (2) produce simulants as cheaply as possible, (3) produce simulants in the amount needed, and (4) produce simulants to meet users? schedules.
Prokhorov, V. B.; Grigorev, I. V.; Fomenko, M. V.; Kaverin, A. A.
Power generating unit no. 4 of Troitskaya State District Power Plant (SDPP) is incapable of operating with a nominal load of 278 MW because of high aerodynamic drag of the gas path. At present, the maximum load of the two-boiler single-turbine unit is 210 MW practically without a possibility of adjustment. The results of numerical simulation of the gas flow for the existing gas path from the electrostatic precipitator (EP) to the smoke exhausts (SEs) and two flue designs proposed for renovation of this section are presented. The results of simulation show that the existing flue section has high aerodynamic drag, which is explained by poor, as regards aerodynamics, design. The local loss coefficient, in terms of the dynamic pressure in the sucker pocket of the smoke exhaust is equal to 4.57. The local aerodynamic loss coefficient after renovation at the considered section according to the first version would make 1.48, and according to the second version 1.325, which would reduce losses at this section by more than a factor of three, and ensure the power unit operation with the rated load.
Gomez, A. V.
The program was designed to provide solutions of engineering accuracy for determining the aerodynamic loads on single- or multiple-lifting-surface configurations that represent vehicles in subsonic flight, e.g., wings, wing-tail, wing-canard, lifting bodies, etc. The preparation is described of the input data, associated input arrangement, and the output format for the program data, including specification of the various operational details of the program such as array sizes, tape numbers utilized, and program dumps. A full description of the underlying theory used in the program development and a review of the program qualification tests are included.
Researchers at the National Renewable Energy Laboratory (NREL) develop a high-fidelity large-eddy simulation model designed to predict the performance of large wind plants with a higher degree of accuracy than current models.
The Aerodynamics Wing Curriculum is a high school program that combines basic physics, aerodynamics, pre-engineering, 3D visualization, computer-assisted drafting, computer-assisted manufacturing, production, reengineering, and success in a 15-hour, 3-week classroom module. (JOW)
Middleton, W. D.; Lundry, J. L.; Coleman, R. G.
The computer program documentation for the design and analysis of supersonic configurations is presented. Schematics and block diagrams of the major program structure, together with subroutine descriptions for each module are included.
Choo, Yung K.; Slater, John W.; Vickerman, Mary B.; VanZante, Judith F.; Wadel, Mary F. (Technical Monitor)
Issues associated with analysis of 'icing effects' on airfoil and wing performances are discussed, along with accomplishments and efforts to overcome difficulties with ice. Because of infinite variations of ice shapes and their high degree of complexity, computational 'icing effects' studies using available software tools must address many difficulties in geometry acquisition and modeling, grid generation, and flow simulation. The value of each technology component needs to be weighed from the perspective of the entire analysis process, from geometry to flow simulation. Even though CFD codes are yet to be validated for flows over iced airfoils and wings, numerical simulation, when considered together with wind tunnel tests, can provide valuable insights into 'icing effects' and advance our understanding of the relationship between ice characteristics and their effects on performance degradation.
Churchfield, M. J.; Michalakes, J.; Vanderwende, B.; Lee, S.; Sprague, M. A.; Lundquist, J. K.; Moriarty, P. J.
Wind plant aerodynamics are directly affected by the microscale weather, which is directly influenced by the mesoscale weather. Microscale weather refers to processes that occur within the atmospheric boundary layer with the largest scales being a few hundred meters to a few kilometers depending on the atmospheric stability of the boundary layer. Mesoscale weather refers to large weather patterns, such as weather fronts, with the largest scales being hundreds of kilometers wide. Sometimes microscale simulations that capture mesoscale-driven variations (changes in wind speed and direction over time or across the spatial extent of a wind plant) are important in wind plant analysis. In this paper, we present our preliminary work in coupling a mesoscale weather model with a microscale atmospheric large-eddy simulation model. The coupling is one-way beginning with the weather model and ending with a computational fluid dynamics solver using the weather model in coarse large-eddy simulation mode as an intermediary. We simulate one hour of daytime moderately convective microscale development driven by the mesoscale data, which are applied as initial and boundary conditions to the microscale domain, at a site in Iowa. We analyze the time and distance necessary for the smallest resolvable microscales to develop.
Lin, Qiao; Jiang, Fukang; Wang, Xuan-Qi; Xu, Yong; Han, Zhigang; Tai, Yu-Chong; Lew, James; Ho, Chih-Ming
MEMS thermal shear-stress sensors exploit heat-transfer effects to measure the shear stress exerted by an air flow on its solid boundary, and have promising applications in aerodynamic control. Classical theory for conventional, macroscale thermal shear-stress sensors states that the rate of heat removed by the flow from the sensor is proportional to the 1/3-power of the shear stress. However, we have observed that this theory is inconsistent with experimental data from MEMS sensors. This paper seeks to develop an understanding of MEMS thermal shear-stress sensors through a study including both experimental and theoretical investigations. We first obtain experimental data that confirm the inadequacy of the classical theory by wind-tunnel testing of prototype MEMS shear-stress sensors with different dimensions and materials. A theoretical analysis is performed to identify that this inadequacy is due to the lack of a thin thermal boundary layer in the fluid flow at the sensor surface, and then a two-dimensional MEMS shear-stress sensor theory is presented. This theory incorporates important heat-transfer effects that are ignored by the classical theory, and consistently explains the experimental data obtained from prototype MEMS sensors. Moreover, the prototype MEMS sensors are studied with three-dimensional simulations, yielding results that quantitatively agree with experimental data. This work demonstrates that classical assumptions made for conventional thermal devices should be carefully examined for miniature MEMS devices.
VanZante, Dale; Envia, Edmane
As turbofan engine bypass ratios continue to increase, the contribution of the turbine to the engine noise signature is receiving more attention. Understanding the relative importance of the various turbine noise generation mechanisms and the characteristics of the turbine acoustic transmission loss are essential ingredients in developing robust reduced-order models for predicting the turbine noise signature. A computationally based investigation has been undertaken to help guide the development of a turbine noise prediction capability that does not rely on empiricism. As proof-of-concept for this approach, two highly detailed numerical simulations of the unsteady flow field inside the first stage of a modern high-pressure turbine were carried out. The simulations were computed using TURBO, which is an unsteady Reynolds-Averaged Navier-Stokes code capable of multi-stage simulations. Spectral and modal analysis of the unsteady pressure data from the numerical simulation of the turbine stage show a circumferential modal distribution that is consistent with the Tyler-Sofrin rule. Within the high-pressure turbine, the interaction of velocity, pressure and temperature fluctuations with the downstream blade rows are all possible tone noise source mechanisms. We have taken the initial step in determining the source strength hierarchy by artificially reducing the level of temperature fluctuations in the turbine flowfield. This was accomplished by changing the vane cooling flow temperature in order to mitigate the vane thermal wake in the second of the two simulations. The results indicated that, despite a dramatic change in the vane cooling flow, the computed modal levels changed very little indicating that the contribution of temperature fluctuations to the overall pressure field is rather small compared with the viscous and potential field interaction mechanisms.
CASE, DAVID A.; CHEATHAM, THOMAS E.; DARDEN, TOM; GOHLKE, HOLGER; LUO, RAY; MERZ, KENNETH M.; ONUFRIEV, ALEXEY; SIMMERLING, CARLOS; WANG, BING; WOODS, ROBERT J.
We describe the development, current features, and some directions for future development of the Amber package of computer programs. This package evolved from a program that was constructed in the late 1970s to do Assisted Model Building with Energy Refinement, and now contains a group of programs embodying a number of powerful tools of modern computational chemistry, focused on molecular dynamics and free energy calculations of proteins, nucleic acids, and carbohydrates. PMID:16200636
Povinelli, L. A.
The drag of a simulated scramjet combustion module was measured at Mach 2, 2.5, and 3. The combustor was rectangular in cross section and incorporated six swept fuel injector struts. The effect of strut leading edge radius, position of maximum thickness, thickness ratio, sweep angle, and strut length on the drag was determined. Reduction in thickness ratio had the largest effect on drag reduction. Sweeping the struts upstream yielded the same drag as sweeping the struts downstream and potentially offers the advantages of increased mixing time for the fuel. Helium injection was used to simulate hydrogen fuel. The interstrut spacing required to achieve good distribution of fuel was was found to be about 10 jet diameters. The contribution of helium injection to drag reduction was small.
Tweedt, Daniel L.
Computational Aerodynamic simulations of an 840 ft/sec tip speed, Advanced Ducted Propulsor fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15- foot Low Speed Wind Tunnel at the NASA Glenn Research Center, resulting in quality, detailed aerodynamic and acoustic measurement data. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating conditions simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, excluding a long core duct section downstream of the core inlet guide vane. As a result, only fan rotational speed and system bypass ratio, set by specifying static pressure downstream of the core inlet guide vane row, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. The computed blade row flow fields for all five fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive boundary layer separations or related secondary-flow problems. A few spanwise comparisons between
Carlson, H. W.; Walkley, K. B.
This paper describes methodology and an associated computer program for the design of wing lifting surfaces with attainable thrust taken into consideration. The approach is based on the determination of an optimum combination of a series of candidate surfaces rather than the more commonly used candidate loadings. Special leading-edge surfaces are selected to provide distributed leading-edge thrust forces which compensate for any failure to achieve the full theoretical leading-edge thrust, and a second series of general candidate surfaces is selected to minimize drag subject to constraints on the lift coefficient and, if desired, on the pitching moment coefficient. A primary purpose of the design approach is the introduction of attainable leading-edge thrust considerations so that relatively mild camber surfaces may be employed in the achievement of aerodynamic efficiencies comparable to those attainable if full theoretical leading-edge thrust could be achieved. The program provides an analysis as well as a design capability and is applicable to both subsonic and supersonic flow.
Ngo, Khiem Viet; Tumer, Irem Y.
The unsteady compressible inviscid flow is characterized by the conservations of mass, momentum, and energy; or simply the Euler equations. In this paper, a study of the subsonic one-dimensional Euler equations with local preconditioning is presented with a modal analysis approach. Specifically, this study investigates the behavior of airflow in a gas turbine engine using the specified conditions at the inflow and outflow boundaries of the compressor, combustion chamber, and turbine, under the impact of variations in pressure, velocity, temperature, and density at low Mach numbers. Two main questions that motivate this research are: 1) Is there any aerodynamic problem with the existing gas turbine engines that could impact aircraft performance? 2) If yes, what aspect of a gas turbine engine could be improved via design to alleviate that impact and to optimize aircraft performance. This paper presents an initial attempt to the flow behavior in terms (perturbation) using simulation outputs from a customer-deck model obtained from Pratt&Whitney, (i.e., pressure, temperature, velocity, density) about their mean states at the inflow and outflow boundaries of the compressor, combustion chamber, and turbine. Flow behavior is analyzed for the high pressure compressor and combustion chamber employing the conditions on their left and right boundaries. In the same fashion, similar analyses are carried out for the high and low-pressure turbines. In each case, the eigenfrequencies that are obtained for different boundary conditions are examined closely based on their probabilistic distributions, a result of a Monte Carlo 10,000-sample simulation. Furthermore, the characteristic waves and eave response are analyzed and contrasted among different cases, with and without preconditioners. The results reveal the existence of flow instabilities due to the combined effect of variations and excessive pressures; which are clearly the case in the combustion chamber and high
Middleton, W. D.; Lundry, J. L.; Coleman, R. G.
The computer program for the design and analysis of supersonic aircraft configurations is presented. The schematics of the program structure are provided. The individual overlays and subroutines are described. The system is useful in determining surface pressures and supersonic area rule concepts.
Ruo, S. Y.
A computer program was developed to account approximately for the effects of finite wing thickness in transonic potential flow over an oscillation wing of finite span. The program is based on the original sonic box computer program for planar wing which was extended to account for the effect of wing thickness. Computational efficiency and accuracy were improved and swept trailing edges were accounted for. Account for the nonuniform flow caused by finite thickness was made by application of the local linearization concept with appropriate coordinate transformation. A brief description of each computer routine and the applications of cubic spline and spline surface data fitting techniques used in the program are given, and the method of input was shown in detail. Sample calculations as well as a complete listing of the computer program listing are presented.
Chu, Wei; Lin, Peifeng; Zhang, Li; Jin, Yingzi; Wang, Yanping; Kim, Heuy Dong; Setoguchi, Toshiaki
Interaction between rotor and struts has great effect on the performance of small axial fan systems. The small axial fan systems are selected as the studied objects in this paper, and four square struts are downstream of the rotor. The cross section of the struts is changed to the cylindrical shapes for the investigation: one is in the same hydraulic diameter as the square struts and another one is in the same cross section as the square struts. Influence of the shape of the struts on the static pressure characteristics, the internal flow and the sound emission of the small axial fans are studied. Standard K-ɛ turbulence model and SIMPLE algorithm are applied in the calculation of the steady fluid field, and the curves of the pressure rising against the flow rate are obtained, which demonstrates that the simulation results are in nice consistence with the experimental data. The steady calculation results are set as the initial field in the unsteady calculation. Large eddy simulation and PISO algorithm are used in the transient calculation, and the Ffowcs Williams-Hawkings model is introduced to predict the sound level at the eight monitoring points. The research results show that: the static pressure coefficients of the fan with cylindrical struts increase by about 25% compared to the fan with square struts, and the efficiencies increase by about 28.6%. The research provides a theoretical guide for shape optimization and noise reduction of small axial fan with struts.
Stowers, S. T.; Bass, J. M.; Oden, J. T.
A phase 2 research and development effort was conducted in area transonic, compressible, inviscid flows with an ultimate goal of numerically modeling complex flows inherent in advanced helicopter blade designs. The algorithms and methodologies therefore are classified as adaptive methods, which are error estimation techniques for approximating the local numerical error, and automatically refine or unrefine the mesh so as to deliver a given level of accuracy. The result is a scheme which attempts to produce the best possible results with the least number of grid points, degrees of freedom, and operations. These types of schemes automatically locate and resolve shocks, shear layers, and other flow details to an accuracy level specified by the user of the code. The phase 1 work involved a feasibility study of h-adaptive methods for steady viscous flows, with emphasis on accurate simulation of vortex initiation, migration, and interaction. Phase 2 effort focused on extending these algorithms and methodologies to a three-dimensional topology.
Medan, R. T.; Ray, K. S.
A description of and users manual are presented for a U.S.A. FORTRAN 4 computer program which evaluates spanwise and chordwise loading distributions, lift coefficient, pitching moment coefficient, and other stability derivatives for thin wings in linearized, steady, subsonic flow. The program is based on a kernel function method lifting surface theory and is applicable to a large class of planforms including asymmetrical ones and ones with mixed straight and curved edges.
Dill, C. C.; Young, J. C.; Roberts, B. B.; Craig, M. K.; Hamilton, J. T.; Boyle, W. W.
The phase B Space Shuttle systems definition studies resulted in a generic configuration consisting of a delta wing orbiter, and two solid rocket boosters (SRB) attached to an external fuel tank (ET). The initial challenge facing the aerodynamic community was aerodynamically optimizing, within limits, this configuration. As the Shuttle program developed and the sensitivities of the vehicle to aerodynamics were better understood the requirements of the aerodynamic data base grew. Adequately characterizing the vehicle to support the various design studies exploded the size of the data base to proportions that created a data modeling/management challenge for the aerodynamicist. The ascent aerodynamic data base originated primarily from wind tunnel test results. The complexity of the configuration rendered conventional analytic methods of little use. Initial wind tunnel tests provided results which included undesirable effects from model support tructure, inadequate element proximity, and inadequate plume simulation. The challenge to improve the quality of test results by determining the extent of these undesirable effects and subsequently develop testing techniques to eliminate them was imposed on the aerodynamic community. The challenges to the ascent aerodynamics community documented are unique due to the aerodynamic complexity of the Shuttle launch. Never before was such a complex vehicle aerodynamically characterized. The challenges were met with innovative engineering analyses/methodology development and wind tunnel testing techniques.
Britt, Charles L.
The NASA Radar Simulation Program is a comprehensive calculation of the expected output of an airborne coherent pulse Doppler radar system viewing a low level microburst along or near the approach path. Inputs to the program include the radar system parameters and data files that contain the characteristics of the microbursts to be simulated, the ground clutter map, and the discrete target data base which provides a simulation of the moving ground clutter. For each range bin, the simulation calculates the received signal amplitude level by integrating the product of the antenna gain pattern and the scattering source amplitude and phase of a spherical shell volume segment defined by the pulse width, radar range, and ground plane intersection. A series of in-phase and quadrature pulses are generated and stored for further processing if desired. In addition, various signal processing techniques are used to derive the simulated velocity and hazard measurements, and store them for use in plotting and display programs.
Reddy, T. S. R.; Srivastava, R.
This guide describes the input data required for using MSAP2D (Multi Stage Aeroelastic analysis Program - Two Dimensional) computer code. MSAP2D can be used for steady, unsteady aerodynamic, and aeroelastic (flutter and forced response) analysis of bladed disks arranged in multiple blade rows such as those found in compressors, turbines, counter rotating propellers or propfans. The code can also be run for single blade row. MSAP2D code is an extension of the original NPHASE code for multiblade row aerodynamic and aeroelastic analysis. Euler equations are used to obtain aerodynamic forces. The structural dynamic equations are written for a rigid typical section undergoing pitching (torsion) and plunging (bending) motion. The aeroelastic equations are solved in time domain. For single blade row analysis, frequency domain analysis is also provided to obtain unsteady aerodynamic coefficients required in an eigen analysis for flutter. In this manual, sample input and output are provided for a single blade row example, two blade row example with equal and unequal number of blades in the blade rows.
Eckert, W. T.; Mort, K. W.; Jope, J.
General guidelines are given for the design of diffusers, contractions, corners, and the inlets and exits of non-return tunnels. A system of equations, reflecting the current technology, has been compiled and assembled into a computer program (a user's manual for this program is included) for determining the total pressure losses. The formulation presented is applicable to compressible flow through most closed- or open-throat, single-, double-, or non-return wind tunnels. A comparison of estimated performance with that actually achieved by several existing facilities produced generally good agreement.
Weinstein, I.; Avery, D. E.; Chapman, A. J.
An experimental investigation was made on a simulated reusable-surface-insulation tile array in a turbulent boundary layer to determine aerodynamic-heating distributions representative of those expected on the surface of the shuttle orbiter during earth entry due to the presence of longitudinal and transverse surface gaps. The tests were conducted in an 8-foot high-temperature structures tunnel in a test medium of methane-air combustion products at a nominal Mach number of 6.6 and over a free-stream Reynolds number range from 2,000,000 to 4,900,000 per meter (600,000 to 1,500,000 per foot). The results were used to assess the aerodynamic heating effects produced by parameters that include gap width, boundary-layer displacement thickness, in-line and staggered tile arrangement, and tile protrusion.
Perumalla, Kalyan S; Park, Alfred J
In simulating large parallel systems, bottom-up approaches exercise detailed hardware models with effects from simplified software models or traces, whereas top-down approaches evaluate the timing and functionality of detailed software models over coarse hardware models. Here, we focus on the top-down approach and significantly advance the scale of the simulated parallel programs. Via the direct execution technique combined with parallel discrete event simulation, we stretch the limits of the top-down approach by simulating message passing interface (MPI) programs with millions of tasks. Using a timing-validated benchmark application, a proof-of-concept scaling level is achieved to over 0.22 billion virtual MPI processes on 216,000 cores of a Cray XT5 supercomputer, representing one of the largest direct execution simulations to date, combined with a multiplexing ratio of 1024 simulated tasks per real task.
Yuen, Allan H. K.
Extensive studies have been conducted on programming knowledge, instructional environments and effective pedagogies in order to improve the teaching and learning of computer programming. However, there have been few studies on the usefulness and effectiveness of using simulation. The aim of this paper is two-fold: to describe the design and…
Blanchard, Robert C.; Larman, Kevin T.; Moats, Christina D.
A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low-frequency (less than 1Hz), low-level (approximately 1 x 10(exp -6) g) acceleration measurements are made during atmospheric re-entry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP re-entry mission data sets spanning a period of about 10 years have been processed. The HiRAP-derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as function of angle of attack, body-flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle of attack are also presented, along with flight-derived rarefied-flow transition bridging formulae. Comparisons are made between the aerodynamics model, data from the latest Orbiter Operational Aerodynamic Design Data Book, applicable computer simulations, and wind-tunnel data.
Maksymiuk, C. M.; Watson, S. A.
A computer program written in a table ""look-up'' format, is presented which provides a comprehensive data base on NACA 16-series airfoils. The geometry covered is limited to cambers for a design-lift coefficient from 0.0 to 0.7 and thickness ratios from 4 to 21%. The data include Mach numbers from 0.3 to 1.6, angles of attack from -4 to 8 degrees, and lift coefficients from 0.0 to 0.8. Extrapolation is used to obtain data from Mach numbers, angles of attack, and lift coefficients beyond those for which data are available. A routine to adjust the lift and drag coefficients beyond stall is included. The uses and limitations of the program are also discussed.
Spearman, M. L.
The World War 2 time period ushered in a new era in aeronautical research and development. The air conflict during the war highlighted the need of aircraft with agility, high speed, long range, large payload capability, and in addition, introduced a new concept in air warfare through the use of guided missiles. Following the war, the influx of foreign technology, primarily German, led to rapid advances in jet propulsion and speed, and a host of new problem areas associated with high-speed flight designs were revealed. The resolution of these problems led to a rash of new design concepts and many of the lessons learned, in principle, are still effective today. In addition to the technical lessons learned related to aircraft development programs, it might also be noted that some lessons involving the political and philosophical nature of aircraft development programs are worth attention.
Lan, C. E.; Mehrotra, S. C.; Fox, C. H., Jr.
The necessary information for using a computer program to calculate the aerodynamic characteristics under symmetrical flight conditions and the lateral-directional stability derivatives of wing-body combinations with upper-surface-blowing (USB) or over-wing-blowing (OWB) jets are described. The following new features were added to the program: (1) a fuselage of arbitrary body of revolution has been included. The effect of wing-body interference can now be investigated, and (2) all nine lateral-directional stability derivatives can be calculated. The program is written in FORTRAN language and runs on CDC Cyber 175 and Honeywell 66/60 computers.
Melton, John E.
EGADS is a comprehensive preliminary design tool for estimating the performance of light, single-engine general aviation aircraft. The software runs on the Apple Macintosh series of personal computers and assists amateur designers and aeronautical engineering students in performing the many repetitive calculations required in the aircraft design process. The program makes full use of the mouse and standard Macintosh interface techniques to simplify the input of various design parameters. Extensive graphics, plotting, and text output capabilities are also included.
Driggers, H. H.
A conceptual design study was performed of a vertical attitude takeoff and landing (VATOL) fighter/attack aircraft. The configuration has a close-coupled canard-delta wing, side two-dimensional ramp inlets, and two augmented turbofan engines with thrust vectoring capability. Performance and sensitivities to objective requirements were calculated. Aerodynamic characteristics were estimated based on contractor and NASA wind tunnel data. Computer simulations of VATOL transitions were performed. Successful transitions can be made, even with series post-stall instabilities, if reaction controls are properly phased. Principal aerodynamic uncertainties identified were post-stall aerodynamics, transonic aerodynamics with thrust vectoring and inlet performance in VATOL transition. A wind tunnel research program was recommended to resolve the aerodynamic uncertainties.
Villarreal, James; Mcintire, Gary
Neural Network Environment on Transputer System (NNETS) computer program provides users high degree of flexibility in creating and manipulating wide variety of neural-network topologies at processing speeds not found in conventional computing environments. Supports back-propagation and back-propagation-related algorithms. Back-propagation algorithm used is implementation of Rumelhart's generalized delta rule. NNETS developed on INMOS Transputer(R). Predefines back-propagation network, Jordan network, and reinforcement network to assist users in learning and defining own networks. Also enables users to configure other neural-network paradigms from NNETS basic architecture. Small portion of software written in OCCAM(R) language.
Hrabak, R. R.; Levy, D. W.; Finn, P.; Roskam, J.
The use of pressure differentials in a flight control system was evaluated. The pressure profile around the test surface was determined using two techniques: (1) windtunnel data (actual); and (2) NASA/Langley Single Element Airfoil Computer Program (theoretical). The system designed to evaluate the concept of using pressure differentials is composed of a sensor drive and power amplifiers, actuator, position potentiometer, and a control surface. The characteristics (both desired and actual) of the system and each individual component were analyzed. The desired characteristics of the system as a whole are given. The flight control system developed, the testing procedures and data reduction methods used, and theoretical frequency response analysis are described.
Deadrick, F.J.; Hawke, R.S.; Scudder, J.K.
A computer simulation code has been developed and validated at the Lawrence Livermore National Laboratory (LLNL) to predict the performance of a railgun electromagnetic accelerator. The code, called MAGRAC (MAGnetic Railgun ACcelerator), models the performance of a railgun driven by a magnetic flux compression current generator (MFCG). The MAGRAC code employs a time-step solution of the nonlinear time-varying element railgun circuit to determine rail currents. From the rail currents, the projectile acceleration, velocity, and position is found. The MAGRAC code was validated through a series of eight railgun tests conducted jointly with the Los Alamos Scientific National Laboratory. This paper describes the formulation of the MAGRAC railgun model and compares the predicted current waveforms with those obtained from full-scale experiments.
Fasching, W. A.
The improved single shank high pressure turbine design was evaluated in component tests consisting of performance, heat transfer and mechanical tests, and in core engine tests. The instrumented core engine test verified the thermal, mechanical, and aeromechanical characteristics of the improved turbine design. An endurance test subjected the improved single shank turbine to 1000 simulated flight cycles, the equivalent of approximately 3000 hours of typical airline service. Initial back-to-back engine tests demonstrated an improvement in cruise sfc of 1.3% and a reduction in exhaust gas temperature of 10 C. An additional improvement of 0.3% in cruise sfc and 6 C in EGT is projected for long service engines.
Tweedt, Daniel L.
Computational Aerodynamic simulations of a 1215 ft/sec tip speed transonic fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, low-noise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating points simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, which for this model did not include a split flow path with core and bypass ducts. As a result, it was only necessary to adjust fan rotational speed in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. Computed blade row flow fields at all fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the flow fields at all operating conditions reveals no excessive boundary layer separations or related secondary-flow problems.
Hedin, J.; Gumbel, J.; Rapp, M.
There is a great interest in nanometer-sized particles in the mesosphere at the moment with the recent launches of the MAGIC and ECOMA payloads. However, rocket-borne measurements of these particles are far from trivial. Since rocket payloads move through the measurement volume with supersonic speeds they can introduce aerodynamic perturbations that complicate the collection of e.g. smoke particle measurements in this region. Nanometer-sized particles tend to follow the gas flow around the payload and do not reach the detector if the aerodynamic design of the instrument has not been considered carefully. The analysis is further complicated by the fact that the payload moves from continuum flow conditions to free molecular flow conditions via the transition regime. Therefore, aerodynamics simulations are of critical importance for the success of these projects. To simulate the gas flow around the rocket payload a Direct Simulation Monte Carlo program is used. A simple model has been developed to introduce smoke particles in the gas flow and determine their trajectories. The result from this is a specific lower limit to the size of smoke particles detectable by various detector designs.
Wang, Qing; He, Kai-Feng; Qian, Wei-Qi; Zhang, Tian-Jiao; Cheng, Yan-Qing; Wu, Kai-Yuan
In view of engineering application, it is practicable to decompose the aerodynamics into three components: the static aerodynamics, the aerodynamic increment due to steady rotations, and the aerodynamic increment due to unsteady separated and vortical flow. The first and the second components can be presented in conventional forms, while the third is described using a one-order differential equation and a radial-basis-function (RBF) network. For an aircraft configuration, the mathematical models of 6-component aerodynamic coefficients are set up from the wind tunnel test data of pitch, yaw, roll, and coupled yawroll large-amplitude oscillations. The flight dynamics of an aircraft is studied by the bifurcation analysis technique in the case of quasi-steady aerodynamics and unsteady aerodynamics, respectively. The results show that: (1) unsteady aerodynamics has no effect upon the existence of trim points, but affects their stability; (2) unsteady aerodynamics has great effects upon the existence, stability, and amplitudes of periodic solutions; and (3) unsteady aerodynamics changes the stable regions of trim points obviously. Furthermore, the dynamic responses of the aircraft to elevator deflections are inspected. It is shown that the unsteady aerodynamics is beneficial to dynamic stability for the present aircraft. Finally, the effects of unsteady aerodynamics on the post-stall maneuverability are analyzed by numerical simulation.
An aerodynamic analysis system based on potential theory at subsonic/supersonic speeds and impact type finite element solutions at hypersonic conditions is described. Three dimensional configurations having multiple nonplanar surfaces of arbitrary planform and bodies of noncircular contour may be analyzed. Static, rotary, and control longitudinal and lateral directional chracteristics may be generated. The analysis has been implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis problem. Typical simulation indicates that program provides an efficient analysis for systematically performing various aerodynamic configuration tradeoff and evaluation studies.
Redman, M. C.; Rowe, W. S.
A digital computer program has been developed to calculate unsteady loadings caused by motions of lifting surfaces with leading edge or trailing edge controls based on the subsonic kernel function approach. The pressure singularities at hinge line and side edges have been extracted analytically as a preliminary step to solving the integral equation by collocation. The program calculates generalized aerodynamic forces for user supplied deflection modes. Optional intermediate output includes pressure at an array of points, and sectional generalized forces. From one to six controls on the half span can be accommodated.
Campbell, R. L.
Tests were conducted in the Langley High-Speed 7- by 10-Foot Tunnel using a 1/10-scale model of an executive jet to examine the effects of the nacelles on the wing pressures and model longitudinal aerodynamic characteristics. For the present investigation, each wing panel was modified with a simulated, partial-chord, laminar-flow-control glove. Horizontal-tail effects were also briefly examined. The tests covered a range of Mach numbers from 0.40 to 0.82 and lift coefficients from 0.20 to 0.55. Oil-flow photographs of the wing at selected conditions are included.
Mazhul, I. I.; Volkov, V. F.; Zvegintsev, V. I.; Ivanov, I. V.
A possible influence of the deflection of control surfaces on the aerodynamics of an axisymmetric slender configuration at supersonic flow speeds is considered. A classical configuration consisting from the fuselage in the form of a body of revolution and having cross frontal fins and six-blade trailing stabilizers is considered as the investigation object. The physical flow pattern at the deflection of horizontal fin consoles is investigated and the estimates are obtained for the influence of this deflection on both the characteristics of elements (the body and stabilizers) as well as on the integral aerodynamic characteristics of the entire configuration. Numerical computations of the flow have been done at the freestream Mach number M = 3 in the range of attack angles α = 0-10° and the angles of the control surfaces deflection δ cs = ±5° on the basis of the averaged Navier-Stokes equations and the SST k-ω turbulence model.