Schlesinger's Telescope: A Brief History of the Yale 26-inch Refractor
NASA Astrophysics Data System (ADS)
van Altena, W. F.; Hoffleit, E. D.
2003-08-01
Frank Schlesinger began planning for the establishment of a southern observatory when he arrived at Yale in 1920. After discussing the possibility of a location in Auckland, New Zealand and gathering site survey observations for a four-month period, he decided to select a site in Johannesburg, South Africa. A large photographically corrected 26-inch objective was ground, polished and completed by James McDowell in 1923, while the telescope was built largely in the Yale shops in New Haven. Schlesinger left New Haven in 1924 with the lens, and the telescope followed shortly thereafter. Installation of the 26-inch refractor was completed in early June and dedicated by the Prince of Wales on June 22, 1924. The principal observational program for the 26-inch refractor was the determination of parallaxes of the bright stars and it was continued until 1952, when the telescope was moved to Mt. Stromlo due to the deteriorating sky conditions in Johannesburg. The parallax program continued at Mt. Stromlo until 1963 when the telescope was donated to the Commonwealth Observatory and the Yale-Columbia project moved to Argentina, Columbia having joined with Yale in 1943. Approximately 70,000 plates were taken with the refractor for the parallax programs and about 2,000 stellar parallaxes determined. The Schlesinger 26-inch refractor was destroyed by a firestorm on January 18, 2003, along with the other telescopes on Mount Stromlo, the workshop, library and many of the residences. This event was a sad ending to a telescope that played a major role in defining our knowledge of the distances, motions and masses of the brighter stars during the first half of the 20th century.
Astrometric observations of visual binaries using 26-inch refractor during 2007-2014 at Pulkovo
NASA Astrophysics Data System (ADS)
Izmailov, I. S.; Roshchina, E. A.
2016-04-01
We present the results of 15184 astrometric observations of 322 visual binaries carried out in 2007-2014 at Pulkovo observatory. In 2007, the 26-inch refractor ( F = 10413 mm, D = 65 cm) was equipped with the CCD camera FLI ProLine 09000 (FOV 12' × 12', 3056 × 3056 pixels, 0.238 arcsec pixel-1). Telescope automation and weather monitoring system installation allowed us to increase the number of observations significantly. Visual binary and multiple systems with an angular distance in the interval 1."1-78."6 with 7."3 on average were included in the observing program. The results were studied in detail for systematic errors using calibration star pairs. There was no detected dependence of errors on temperature, pressure, and hour angle. The dependence of the 26-inch refractor's scale on temperature was taken into account in calculations. The accuracy of measurement of a single CCD image is in the range of 0."0005 to 0."289, 0."021 on average along both coordinates. Mean errors in annual average values of angular distance and position angle are equal to 0."005 and 0.°04 respectively. The results are available here http://izmccd.puldb.ru/vds.htmand in the Strasbourg Astronomical Data Center (CDS). In the catalog, the separations and position angles per night of observation and annual average as well as errors for all the values and standard deviations of a single observation are presented. We present the results of comparison of 50 pairs of stars with known orbital solutions with ephemerides.
Astrometric observations of satellites of Uranus using a 26-inch refractor in 2007-2011
NASA Astrophysics Data System (ADS)
Roshchina, E. A.; Izmailov, I. S.; Kiseleva, T. P.
2015-05-01
This paper reports CCD observations of Uranus and its main satellites using a 26-inch refractor at the Pulkovo Observatory in 2007-2011. These are 2450 CCD frames with images of Uranus and its four main satellites, i.e., Ariel, Umbriel, Titania, and Oberon. The field of view of the FLI Proline 9000 CCD camera is 12' × 12', which allows us to obtain stars and perform astrometric reduction by Turner's method to determine the satellites' equatorial coordinates. UCAC2 is used as a reference catalogue. The equatorial coordinates are compared with the GUST 06 theory. The average accuracy of normal places is 0.030″-0.040″ in right ascension and declination. The positions of the satellites and their theoretical uranocentric coordinates by GUST 06 are used to calculate the equatorial coordinates of Uranus. The positions of Uranus are compared with the INPOP10 planetary theory. The paper also presents the satellites' differential coordinates relative to one another.
Blowdown Wind Tunnels: Latest Citations from the Aerospace Database
NASA Technical Reports Server (NTRS)
1996-01-01
The bibliography contains citations concerning the design, construction, operation, and performance of blowdown wind tunnels. The use of compressed gas, mechanical piston, or combustion exhaust to provide continuous or short-duration operation from transonic to hypersonic approach velocities is discussed. Also covered are invasive and non-invasive aerothermodynamic instrumentation, data acquisition and reduction techniques, and test reports on aerospace components. Comprehensive coverage of wind tunnel force balancing systems and supersonic wind tunnels are covered in separate bibliographies.
BWR drywell behavior under steam blowdown
NguyenLe, Q.A.; Ishii, Mamoru
1998-12-31
Historically, the focus of thermal-hydraulics analyses on large-break loss-of-coolant accidents (LOCAs) has been on the transients within the reactor or steam generator. Few have studied the effects of steam blowdown on the containment building. The authors present some numerical and experimental results of the blowdown tests performed at the Purdue University multidimensional integrated test assembly (PUMA).
Time-dependent measurement of base pressure in a blowdown tunnel with varying unit Reynolds number
NASA Technical Reports Server (NTRS)
Kangovi, S.; Rao, D. M.
1978-01-01
An operational characteristic of blowdown-type of wind tunnels is the drop in the stagnation temperature with time and the accompanying change in the test-section unit Reynolds number at constant stagnation pressure and Mach number. This apparent disadvantage can be turned to advantage in some cases where a Reynolds number scan is desired in order to study the effect of unit Reynolds number variation on a particular viscous flow phenomenon. This note presents such an instance arising from recent investigations on base pressure at transonic speeds conducted in the NAL 1-ft tunnel.
BWR drywell behavior under steam blowdown.
NguyenLe, Q.
1998-05-08
Historically, thermal hydraulics analyses on Large Break Loss of Coolant Accidents (LOCA) have been focused on the transients within the reactor or steam generator. Few have studied the effects of steam blowdown on the containment building. This paper discusses some theoretical issues as well as presenting numerical and experimental results of the blowdown tests performed at the Purdue University Multi-Dimensional Integrated Test Assembly (PUMA).
Calibration of transonic and supersonic wind tunnels
NASA Technical Reports Server (NTRS)
Reed, T. D.; Pope, T. C.; Cooksey, J. M.
1977-01-01
State-of-the art instrumentation and procedures for calibrating transonic (0.6 less than M less than 1.4) and supersonic (M less than or equal to 3.5) wind tunnels were reviewed and evaluated. Major emphasis was given to transonic tunnels. Continuous, blowdown and intermittent tunnels were considered. The required measurements of pressure, temperature, flow angularity, noise and humidity were discussed, and the effects of measurement uncertainties were summarized. A comprehensive review of instrumentation currently used to calibrate empty tunnel flow conditions was included. The recent results of relevant research are noted and recommendations for achieving improved data accuracy are made where appropriate. It is concluded, for general testing purposes, that satisfactory calibration measurements can be achieved in both transonic and supersonic tunnels. The goal of calibrating transonic tunnels to within 0.001 in centerline Mach number appears to be feasible with existing instrumentation, provided correct calibration procedures are carefully followed. A comparable accuracy can be achieved off-centerline with carefully designed, conventional probes, except near Mach 1. In the range 0.95 less than M less than 1.05, the laser Doppler velocimeter appears to offer the most promise for improved calibration accuracy off-centerline.
Redistribution of Carbon During Forest Blowdowns
NASA Astrophysics Data System (ADS)
Wohl, E. E.
2013-12-01
Numerous blowdowns in subalpine and montane forests of the Southern Rocky Mountains during the winter of 2011-12 present an opportunity to evaluate how this type of disturbance affects the distribution of organic carbon. Patch blowdowns covering 0.1 to 33 ha are an episodic event with an unknown recurrence interval. Blowdowns influence carbon partitioning in a forested ecosystem by transferring live to dead biomass and exposing soil on uprooted trees. Wood recruited to streams via blowdowns can cause channel-spanning jams that enhance overbank flows and channel avulsion in wider valley segments. This can lead to a multithread channel planform and increased floodplain storage of carbon, as well as altered stream metabolism and animal (insect and fish) production. This talk examines a 33-ha blowdown that occurred along Glacier Creek in Rocky Mountain National Park, Colorado during February 2012. Estimated carbon redistribution ranged as high as 308 Mg C/ha in high-severity patches to 106 Mg C/ha in low-severity patches. Volumes of carbon redistributed from living to dead biomass at high-severity sites are close to average total biomass in subalpine forests in the region. Blowdowns are likely to increase under a warming climate as part of an accelerated disturbance regime involving intense storms and wind, wildfire, and insect infestations. The consequences for carbon partitioning across the landscape, and for riverine ecosystems, depend partly on geomorphic setting, which creates path-dependence and hysteresis. In wider valley segments, downed trees (carbon transferred to dead biomass by blowdowns) may enhance retention of carbon in transport within the stream, facilitating both burial in sedimentary reservoirs and uptake by stream organisms.
An airfoil flutter model suspension system to accommodate large static transonic airloads
NASA Technical Reports Server (NTRS)
Reed, W. H., III
1985-01-01
A pitch/plunge flutter model suspension system and associated two-dimensional MBB-A3 airfoil models is described. The system is designed for installation in the Langley 6-by-19-inch and 6-by-18-inch transonic blowdown wind tunnels to enable systematic study of the transonic flutter characteristics and static pressure distributions of supercritical airfoils at transonic Mach numbers. A compound spring suspension concept is introduced which simultaneously meets requirements for low plunge-mode stiffness, lightweight suspended model, and large steady lift due to angle of attack without the need for excessive static deflections of the plunge spring. The system features variable pitch and plunge frequencies, changeable airfoil rotation axes, and a self aligning control system to maintain a constant mean position of the model with changing airload.
Blowdown transients and implication for leak protection shutdown strategy
Magee, P M
1980-04-01
The creep-fatigue damage due to a blowdown transient is independent of the rate of water-side blowdown following isolation. The damage depends primarily upon plant system characteristics which set the rate of IHTS sodium cooldown after reactor scram and main pump trip. The time delay between a small leak alarm and reactor scram should be considered as a tradeoff between (1) providing time to confirm the existence of the leak to a reasonable level of assurance and (2) the potential for secondary tube wastage damage. In general, there appears to be no incentive to delay water-side blowdown following scram because the additional stress damage due to blowdown is insignificant and the potential for wastage damage should be eliminated as quickly as possible. One exception is the case of multiple evaporators feeding a superheater (as for CRBRP) where isolation and blowdown of one evaporator shortly after scram can cause significant additional stress damage to the superheater upper tubesheet. When initiated, the water-side blowdown process should be rapid. About 30 seconds blowdown time appears practical and adequate, based on the CRBRP design.
NASA Technical Reports Server (NTRS)
Bauer, F.; Garabedian, P.; Korn, D.
1980-01-01
Program aids in design of shockless airfoils, assists development of fuel-conserving, supercritical wings. Algorithm calculates approximate airfoil shape given prescribed pressure distribution. This allows design of families of transonic airfoils for use in aircraft wings or turbine and compressor blades. Program is written in FORTRAN IV for batch execution on CDC-6000.
46 CFR 162.018-5 - Blow-down adjustment and popping tolerance.
Code of Federal Regulations, 2011 CFR
2011-10-01
... 46 Shipping 6 2011-10-01 2011-10-01 false Blow-down adjustment and popping tolerance. 162.018-5... Compressed Gas § 162.018-5 Blow-down adjustment and popping tolerance. (a) Safety relief valves shall be so... adjustible blow-down construction shall be adjusted to close after blowing down not more than 5 percent...
46 CFR 162.018-5 - Blow-down adjustment and popping tolerance.
Code of Federal Regulations, 2013 CFR
2013-10-01
... 46 Shipping 6 2013-10-01 2013-10-01 false Blow-down adjustment and popping tolerance. 162.018-5... Compressed Gas § 162.018-5 Blow-down adjustment and popping tolerance. (a) Safety relief valves shall be so... adjustible blow-down construction shall be adjusted to close after blowing down not more than 5 percent...
46 CFR 162.018-5 - Blow-down adjustment and popping tolerance.
Code of Federal Regulations, 2014 CFR
2014-10-01
... 46 Shipping 6 2014-10-01 2014-10-01 false Blow-down adjustment and popping tolerance. 162.018-5... Compressed Gas § 162.018-5 Blow-down adjustment and popping tolerance. (a) Safety relief valves shall be so... adjustable blow-down construction shall be adjusted to close after blowing down not more than 5 percent...
46 CFR 162.018-5 - Blow-down adjustment and popping tolerance.
Code of Federal Regulations, 2012 CFR
2012-10-01
... 46 Shipping 6 2012-10-01 2012-10-01 false Blow-down adjustment and popping tolerance. 162.018-5... Compressed Gas § 162.018-5 Blow-down adjustment and popping tolerance. (a) Safety relief valves shall be so... adjustible blow-down construction shall be adjusted to close after blowing down not more than 5 percent...
Containment steam blowdown analysis : experimental and numerical comparisons.
NguyenLe, Q.; Ishii, M.; Reactor Analysis; Purdue Univ.
1999-01-01
This paper compares the numerical simulation with the experimental data of a steam blowdown event in a light water reactor containment building. A three step approach was used to analyze the steam jet behavior. First, the temperature and pressure data of a steam blowdown event was measured at the Purdue University Multi-Dimensional Integrated Test Assembly (PUMA), a scaled model of the General Electric Simplified Boiling Water Reactor. Second, a 1-dimensonial, system level RELAP5/Mod3.2 model of the steam blowdown event was created and the results used to set the initial conditions for the PUMA blowdown experiments. Finally, 2-dimensional and 3-dimensional CFD models of the discharged steam jets were computed using PHOENICS, a commercially available CFD package. It was found that RELAP5 is reasonably capable in predicting the general temperature and pressure trends in the RPV. However, due to modeling compromises and the code's built-in capabilities, RELAP5 1-dimensional predictions of containment temperature and pressure did not compare well with measured data. On the other hand, with minor modfications to the k-{var_epsilon} turbulence model, the 2-dimensional and 3-dimensional PHOENICS CFD solutions compared extremely well with the measured data.
Transonic swirling nozzle flow
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Pawlas, Gary E.
1991-01-01
A numerical model of viscous transonic swirling flow in axisymmetric nozzles is developed. MacCormack's implicit Gauss-Seidel method is applied to the thin-layer Navier-Stokes equations in transformed coordinates. Numerical results are compared with experimental data to validate the method. The effect of swirl and viscosity on nozzle performance are demonstrated by examining wall pressures, Mach contours, and integral parameters.
Transonic Flow Past Cone Cylinders
NASA Technical Reports Server (NTRS)
Solomon, George E
1955-01-01
Experimental results are presented for transonic flow post cone-cylinder, axially symmetric bodies. The drag coefficient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental results are compared with theoretical predictions. Interferometric results for several typical flow configurations are shown and an example of shock-free supersonic-to-subsonic compression is experimentally demonstrated. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the axially symmetric transonic equations, valid for a semi-infinite cone, is presented.
NASA Technical Reports Server (NTRS)
Agopian, K. G.
1974-01-01
The problem of inviscid, steady transonic conical flow, formulated in terms of the small disturbance theory, is studied. The small disturbance equation and similarity rules are presented, and a boundary value problem is formulated for the case of a supersonic freestream Mach number. The equation for the perturbation potential is solved numerically using an elliptic finite difference system. The difference equations are solved with a point relaxation algorithm that is also capable of capturing the shock wave during the iteration procedure by using the boundary conditions at the shock. Numerical calculations, for shock location, pressure distribution and drag coefficient, are presented for a family of nonlifting conical wings. The theory of slender wings is also presented and analytical results for pressure and drag coefficients are obtained.
Unsteady design-point flow phenomena in transonic compressors
NASA Technical Reports Server (NTRS)
Gertz, J. B.; Epstein, A. H.
1986-01-01
High-frequency response probes which had previously been used exclusively in the MIT Blowndown Facility were successfully employed in two conventional steady state axial flow compressor facilities to investigate the unsteady flowfields of highly loaded transonic compressors at design point operation. Laser anemometry measurements taken simultaneously with the high response data were also analyzed. The time averaged high response data of static and total pressure agreed quite well with the conventional steady state instrumentation except for flow angle which showed a large spread in values at all radii regardless of the type of instrumentation used. In addition, the time resolved measurements confirmed earlier test results obtained in the MIT Blowdown Facility for the same compressor. The results of these tests have further revealed that the flowfields of highly loaded transonic compressors are heavily influenced by unsteady flow phenomena. The high response measurements exhibited large variations in the blade to blade flow and in the blade passage flow. The observed unsteadiness in the blade wakes is explained in terms of the rotor blades' shed vorticity in periodic vortex streets. The wakes were modeled as two-dimensional vortex streets with finite size cores. The model fit the data quite well as it was able to reproduce the average wake shape and bi-modal probability density distributions seen in the laser anemometry data. The presence of vortex streets in the blade wakes also explains the large blade to blade fluctuations seen by the high response probes which is simply due to the intermittent sampling of the vortex street as it is swept past a stationary probe.
Dynamic Loads Due to Blowdown Through a Sparger
Yoon-Yeong Bae; Choon-Kyung Park; Seok Cho; Se-Young Chun; Chul-Hwa Song; Jong-Kyun Park
2002-07-01
An experiment has been performed using a facility, which simulates the safety depressurization system (SDS) and in-containment refueling water storage tank (IRWST) of APR1400, an advanced PWR being developed in Korea, to investigate the dynamic load resulting from the blowdown of steam from a steam generator through a sparger. The influence of the key parameters, such as air mass, steam pressure, submergence, valve opening time, and pool temperature, on frequency and peak loads was investigated. The blowdown phenomenon was analyzed to find out the real cause of the initiation of bubble oscillation and discrepancy in frequencies between the experiment and calculation by conventional equation for bubble oscillation. The cause of significant damping was discussed and is presumed to be the highly tortuous flow path around bubble. The Rayleigh-Plesset equation modified by the introduction of method of image reasonably reproduce the bubble oscillation in a confined tank. Right after the completion of air discharge the steam discharge immediately follows and it condenses abruptly to provide low-pressure pocket. It may contribute to the negative maximum being greater than positive maximum. The subsequently discharging steam does not play as the driving force anymore. The effect of various parameters on the peak pressure and frequency was analyzed. The air mass, valve opening time, and submergence give significant effect on both of frequency and peak pressure while steam mass flux gives questionable effect and pool temperature gives almost none. (authors)
Transonic airframe propulsion integration
NASA Technical Reports Server (NTRS)
Coltrin, Robert E.; Sanders, Bobby W.; Bencze, Daniel P.
1992-01-01
This chart shows the time line for HSR propulsion/airframe integration program. HSR Phase 1 efforts are underway in both propulsion and aerodynamics. The propulsion efforts focus on cycles, inlets combustors and nozzles that will be required to reduce nitrogen oxide (NOX) at cruise and noise at takeoff and landing to acceptable levels. The aerodynamic efforts concentrate on concepts that will reduce sonic booms and increase the lift/drag (L/D) ratio for the aircraft. The Phase 2 critical propulsion component technology program will focus on large scale demonstrators of the inlet, fan, combustor, and nozzle. The hardware developed here will feed into the propulsion system program which will demonstrate overall system technology readiness, particularly in the takeoff and supersonic cruise speed ranges. The Phase 2 aerodynamic performance and vehicle integration program will provide a validated data base for advanced airframe/control/integration concepts over the full HSR speed range. The results of this program will also feed into the propulsion system demonstration program, particularly in the critical transonic arena.
Jump conditions in transonic equilibria
Guazzotto, L.; Betti, R.; Jardin, S. C.
2013-04-15
In the present paper, the numerical calculation of transonic equilibria, first introduced with the FLOW code in Guazzotto et al.[Phys. Plasmas 11, 604 (2004)], is critically reviewed. In particular, the necessity and effect of imposing explicit jump conditions at the transonic discontinuity are investigated. It is found that 'standard' (low-{beta}, large aspect ratio) transonic equilibria satisfy the correct jump condition with very good approximation even if the jump condition is not explicitly imposed. On the other hand, it is also found that high-{beta}, low aspect ratio equilibria require the correct jump condition to be explicitly imposed. Various numerical approaches are described to modify FLOW to include the jump condition. It is proved that the new methods converge to the correct solution even in extreme cases of very large {beta}, while they agree with the results obtained with the old implementation of FLOW in lower-{beta} equilibria.
Design considerations of the national transonic facility
NASA Technical Reports Server (NTRS)
Baals, D. D.
1976-01-01
The inability of existing wind tunnels to provide aerodynamic test data at transonic speeds and flight Reynolds numbers was examined. The proposed transonic facility is a high Reynolds number transonic wind tunnel designed to meet the research and development needs of industry, and the scientific community. The facility employs the cryogenic approach to achieve high transonic Reynolds numbers at acceptable model loads and tunnel power. By using temperature as a test variable, a unique capability to separate scale effects from model aeroelastic effects is provided. The performance envelope of the facility is shown to provide a ten fold increase in transonic Reynolds number capability compared to currently available facilities.
Design optimization of transonic airfoils
NASA Technical Reports Server (NTRS)
Joh, C.-Y.; Grossman, B.; Haftka, R. T.
1991-01-01
Numerical optimization procedures were considered for the design of airfoils in transonic flow based on the transonic small disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented with an accurate approximation of the wave drag based on the Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without remeshing the grid. Two effective design procedures producing converged designs in approximately 10 global iterations were developed: interchanging the role of the objective function and constraint and the direct lift maximization with move limits which were fixed absolute values of the design variables.
Blow-down analysis of helium from a cryogenic dewar
NASA Astrophysics Data System (ADS)
Khan, H. J.; Zhang, Q. Q.; Rhee, M.; Figueroa, O.
NASA is currently developing Space Shuttle-based refilling of helium using superfluid helium on-orbit transfer (SHOOT). All the critical components of SHOOT need to be developed through ground-based tests. The helium dewar is one of these components. The Dewar consists of a vacuum vessel enclosing a superinsulated tank. The space between the vacuum vessel and the liquid tank is considered a common vacuum space. In the event that the vacuum is lost, the heat transfers to the dewar and the pressure inside the dewar increases rapidly, resulting in rupture of the dewar due to excessive pressure. Therefore, an emergency vent line is required for release of helium to prevent the dewar from rupturing. The study describes a numerical model for blow-down analysis in an emergency. This qualifies the design of the emergency vent line to be adequate for the assumed heat loads to the helium dewar.
A Hydraulic Blowdown Servo System For Launch Vehicle
NASA Astrophysics Data System (ADS)
Chen, Anping; Deng, Tao
2016-07-01
This paper introduced a hydraulic blowdown servo system developed for a solid launch vehicle of the family of Chinese Long March Vehicles. It's the thrust vector control (TVC) system for the first stage. This system is a cold gas blowdown hydraulic servo system and consist of gas vessel, hydraulic reservoir, servo actuator, digital control unit (DCU), electric explosion valve, and pressure regulator etc. A brief description of the main assemblies and characteristics follows. a) Gas vessel is a resin/carbon fiber composite over wrapped pressure vessel with a titanium liner, The volume of the vessel is about 30 liters. b) Hydraulic reservoir is a titanium alloy piston type reservoir with a magnetostrictive sensor as the fluid level indicator. The volume of the reservoir is about 30 liters. c) Servo actuator is a equal area linear piston actuator with a 2-stage low null leakage servo valve and a linear variable differential transducer (LVDT) feedback the piston position, Its stall force is about 120kN. d) Digital control unit (DCU) is a compact digital controller based on digital signal processor (DSP), and deployed dual redundant 1553B digital busses to communicate with the on board computer. e) Electric explosion valve is a normally closed valve to confine the high pressure helium gas. f) Pressure regulator is a spring-loaded poppet pressure valve, and regulates the gas pressure from about 60MPa to about 24MPa. g) The whole system is mounted in the aft skirt of the vehicle. h) This system delivers approximately 40kW hydraulic power, by contrast, the total mass is less than 190kg. the power mass ratio is about 0.21. Have finished the development and the system test. Bench and motor static firing tests verified that all of the performances have met the design requirements. This servo system is complaint to use of the solid launch vehicle.
ASSESSMENT OF THREE TECHNOLOGIES FOR THE TREATMENT OF COOLING TOWER BLOWDOWN
The report gives results of analyses of three methods for treating cooling tower blowdown: vapor compression evaporation (VCE), reverse osmosis (RO), and vertical tube foaming evaporation (VTFE). The two evaporative processes produce pure water (approximately 10 ppm dissolved sol...
TAIR: A transonic airfoil analysis computer code
NASA Technical Reports Server (NTRS)
Dougherty, F. C.; Holst, T. L.; Grundy, K. L.; Thomas, S. D.
1981-01-01
The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.
Transonic Flutter Investigation of Models of T-Tail of Blackburn NA-39 Airplane
NASA Technical Reports Server (NTRS)
Jones, George W., Jr.; Farmer, Moses G.
1959-01-01
A transonic flutter investigation has been made of models of the T-tail of the Blackburn NA-39 airplane. The models were dynamically and elastically scaled from measured airplane data in accordance with criteria which include a flutter safety margin. The investigation was made in the Langley transonic blowdown tunnel and covered a Mach number range from 0.73 to 1.09 at simulated altitudes extending to below sea level. The results of the investigation indicated that, if differences between the measured model and scaled airplane properties are disregarded, the airplane with the normal value of stabilizer pitching stiffness should have a stiffness margin of safety of at least 32 percent at all Mach numbers and altitudes within the flight boundary. However, the airplane with the emergency value of stabilizer pitching stiffness would not have the required margin of safety from symmetrical flutter at Mach numbers greater than about 0.85 at low altitudes. First-order corrections for some differences between the measured model and scaled airplane properties indicated that the airplane with the normal value of stabilizer pitching stiffness would still have an adequate margin of safety from flutter and that the flutter safety margin for the airplane with the emergency value of stabilizer pitching stiffness would be changed from inadequate to adequate. However, the validity of the corrections is questionable.
NASA Technical Reports Server (NTRS)
Ruhlin, Charles L.; Tuovila, W. J.
1961-01-01
A transonic and a supersonic flutter investigation of 1/2-size models of the all-movable canard surface of an expendable powered target has been conducted in the Langley transonic blowdown tunnel and in the Langley 9- by 18-inch supersonic aeroelasticity tunnel, respectively. The transonic investigation covered a Mach number range from 0.7 to 1.3, and the supersonic investigation was made at Mach numbers 1.3, 2.O, and 2.55. The effects on the flutter characteristics of the models of different levels of stiffness and of free play in the pitch control linkage were examined. The semispan models, which were tested at an angle of attack of 0 deg, had pitch springs with the scaled design and 1/2 the scaled design pitch stiffness and total free play in pitch ranging from 0 to 1 deg. An additional model configuration which had a pitch spring 1/4 the scaled design pitch stiffness and no free play in pitch was included in the supersonic tests. All model configurations investigated were flutter free up to dynamic pressures 32 percent greater than those required for flight throughout the Mach number range. Several model configurations were tested to considerably higher dynamic pressures without obtaining flutter at both transonic and supersonic speeds.
Semidirect computations for transonic flow
NASA Technical Reports Server (NTRS)
Swisshelm, J. M.; Adamczyk, J. J.
1983-01-01
A semidirect method, driven by a Poisson solver, was developed for inviscid transonic flow computations. It is an extension of a recently introduced algorithm for solving subsonic rotational flows. Shocks are captured by implementing a form of artificial compressibility. Nonisentropic cases are computed using a shock tracking procedure coupled with the Rankine-Hugoniot relationships. Results are presented for both subsonic and transonic flows. For the test geometry, an unstaggered cascade of 20 percent thick circular arc airfoils at zero angle of attack, shocks are crisply resolved in supercritical situations and the algorithm converges rapidly. In addition, the convergence rate appears to be nearly independent of the entropy and vorticity production at the shock.
Transonic CFD applications at Boeing
NASA Technical Reports Server (NTRS)
Tinoco, E. N.
1989-01-01
The use of computational methods for three dimensional transonic flow design and analysis at the Boeing Company is presented. A range of computational tools consisting of production tools for every day use by project engineers, expert user tools for special applications by computational researchers, and an emerging tool which may see considerable use in the near future are described. These methods include full potential and Euler solvers, some coupled to three dimensional boundary layer analysis methods, for transonic flow analysis about nacelle, wing-body, wing-body-strut-nacelle, and complete aircraft configurations. As the examples presented show, such a toolbox of codes is necessary for the variety of applications typical of an industrial environment. Such a toolbox of codes makes possible aerodynamic advances not previously achievable in a timely manner, if at all.
Transient critical heat flux and blowdown heat-transfer studies
Leung, J.C.
1980-05-01
Objective of this study is to give a best-estimate prediction of transient critical heat flux (CHF) during reactor transients and hypothetical accidents. To accomplish this task, a predictional method has been developed. Basically it involves the thermal-hydraulic calculation of the heated core with boundary conditions supplied from experimental measurements. CHF predictions were based on the instantaneous ''local-conditions'' hypothesis, and eight correlations (consisting of round-tube, rod-bundle, and transient correlations) were tested against most recent blowdown heat-transfer test data obtained in major US facilities. The prediction results are summarized in a table in which both CISE and Biasi correlations are found to be capable of predicting the early CHF of approx. 1 s. The Griffith-Zuber correlation is credited for its prediction of the delay CHF that occurs in a more tranquil state with slowly decaying mass velocity. In many instances, the early CHF can be well correlated by the x = 1.0 criterion; this is certainly indicative of an annular-flow dryout-type crisis. The delay CHF occurred at near or above 80% void fraction, and the success of the modified Zuber pool-boiling correlation suggests that this CHF is caused by flooding and pool-boiling type hydrodynamic crisis.
1984-01-01
The flushing and blowdown system of an EDS plant provides the means of removing viscous coal products and slurry streams from plant vessels and lines. In addition, it provides the flushing oil needed during normal operations for purging instruments in slurry service, for flushing slurry pump and slurry agitator seals, and for flushing slurry safety valve inlet lines. It contains a blowdown system for collecting material from washing operations, including the transport of the collected material to slop tankage. The rerun options for depleting the inventory of collected slop are a related aspect of the flushing and blowdown system design although specific equipment for handling slop is not part of the flushing and blowdown system facilities. This report documents the results of a study which evaluates the flushing and blowdown requirements for a commercial-scale EDS plant. The work was conducted as part of the EDS Consolidation Program. The design recommendations represent a consolidation of learnings accrued during previous phases of the EDS Project including results obtained from ECLP operations, from the ECLP Test Program, and from past EDS Study Design preparations. 1 reference, 4 figures, 2 tables.
National Transonic Facility Characterization Status
NASA Technical Reports Server (NTRS)
Bobbitt, C., Jr.; Everhart, J.; Foster, J.; Hill, J.; McHatton, R.; Tomek, W.
2000-01-01
This paper describes the current status of the characterization of the National Transonic Facility. The background and strategy for the tunnel characterization, as well as the current status of the four main areas of the characterization (tunnel calibration, flow quality characterization, data quality assurance, and support of the implementation of wall interference corrections) are presented. The target accuracy requirements for tunnel characterization measurements are given, followed by a comparison of the measured tunnel flow quality to these requirements based on current available information. The paper concludes with a summary of which requirements are being met, what areas need improvement, and what additional information is required in follow-on characterization studies.
Calculation of transonic aileron buzz
NASA Technical Reports Server (NTRS)
Steger, J. L.; Bailey, H. E.
1979-01-01
An implicit finite-difference computer code that uses a two-layer algebraic eddy viscosity model and exact geometric specification of the airfoil has been used to simulate transonic aileron buzz. The calculated results, which were performed on both the Illiac IV parallel computer processor and the Control Data 7600 computer, are in essential agreement with the original expository wind-tunnel data taken in the Ames 16-Foot Wind Tunnel just after World War II. These results and a description of the pertinent numerical techniques are included.
Simplified Analysis of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance
NASA Technical Reports Server (NTRS)
Morris, C. I.; Rodgers, Stephen L. (Technical Monitor)
2002-01-01
Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modellng task. A simplified model for an idealized, straight-tube, single-shot PDRE blowdown process and thrust determination is described and implemented. In order to form an assessment of the accuracy of the model, the flowfield time history is compared to experimental data from Stanford University. Parametric Studies of the effect of mixture stoichiometry, initial fill temperature, and blowdown pressure ratio on the performance of a PDRE are performed using the model. PDRE performance is also compared with a conventional steady-state rocket engine over a range of pressure ratios using similar gasdynamic assumptions.
Euler solvers for transonic applications
NASA Technical Reports Server (NTRS)
Vanleer, Bram
1989-01-01
The 1980s may well be called the Euler era of applied aerodynamics. Computer codes based on discrete approximations of the Euler equations are now routinely used to obtain solutions of transonic flow problems in which the effects of entropy and vorticity production are significant. Such codes can even predict separation from a sharp edge, owing to the inclusion of artificial dissipation, intended to lend numerical stability to the calculation but at the same time enforcing the Kutta condition. One effect not correctly predictable by Euler codes is the separation from a smooth surface, and neither is viscous drag; for these some form of the Navier-Stokes equation is needed. It, therefore, comes as no surprise to observe that the Navier-Stokes has already begun before Euler solutions were fully exploited. Moreover, most numerical developments for the Euler equations are now constrained by the requirement that the techniques introduced, notably artificial dissipation, must not interfere with the new physics added when going from an Euler to a full Navier-Stokes approximation. In order to appreciate the contributions of Euler solvers to the understanding of transonic aerodynamics, it is useful to review the components of these computational tools. Space discretization, time- or pseudo-time marching and boundary procedures, the essential constituents are discussed. The subject of grid generation and grid adaptation to the solution are touched upon only where relevant. A list of unanswered questions and an outlook for the future are covered.
Assessment of TRAC-PF1 and RELAP5/MOD1 codes with GE large-vessel blowdown test
NASA Astrophysics Data System (ADS)
Jo, J. H.
1983-06-01
The large vessel blowdown Test No. 5801-15 was simulated with the TRAC-PF1 (Version 7.0) and RELAP5/MOD1 (Cycle 14) codes. The test facility consisted of a pressure vessel, 49 in. in diameter by 14 ft long, a 2.5 in. diameter converging-diverging nozzle and a blowdown line connected to the center of the upper part of the vessel (elevation from the bottom of the vessel 10.5 ft). The vessel was filled with saturated water up to 5.5 ft at 1060 psia. The test was initiated by rupturing a disc attached at the end of the nozzle. Blowdown phenomena such as critical blowdown flow and the level swell during blowdown from a partially water filled vessel was studied. Understanding of these phenomena is essential for the analysis of Loss-of-Coolant and steam generator steam line break accidents.
Numerical computation of aeroelastically corrected transonic loads
NASA Technical Reports Server (NTRS)
Chipman, R.; Waters, C.; Mackenzie, D.
1979-01-01
A numerical scheme is presented for the computation of transonic aerodynamic loads on flexible wings. The method consists of iteratively applying the loads computed by a 3D transonic aerodynamics code to a structural model to obtain elastic twist, and then recomputing the loads. Because this iteration is performed concurrently with the iterations performed in computing the aerodynamics, flexible loads are obtained in roughly the same amount of computing time as required to obtain rigid loads. Applications of this method to a flexible supercritical transonic transport wing are presented and compared with model test data.
Transonic interactions of unsteady vortical flows
NASA Technical Reports Server (NTRS)
Mccroskey, W. J.; Srinivasan, G. R.
1984-01-01
Unsteady interactions of strong concentrated vortices, distributed gusts, and sharp-edged gusts with stationary airfoils were analyzed in two-dimensional transonic flow. A simple and efficient method for introducing such vortical disturbances was implemented in numerical codes that range from inviscid transonic small disturbance to thin-layer Navier Stokes. The numerical results demonstrate the large distortions in the overall flow field and in the surface air loads that are produced by various vortical interactions. The results of the different codes are in excellent qualitative agreement, but, as might expected, the transonic small-disturbance calculations are deficient in the important region near the leading edge.
Simulation Of Unsteady, Inviscid, Rotational, Transonic Flow
NASA Technical Reports Server (NTRS)
Damodaran, Murali
1992-01-01
Report describes numerical simulation of two-dimensional, unsteady, inviscid rotational, transonic flow about rigid airfoil in such motions as pitching or plunging oscillations. Study demonstrates potential utility of computation in analyses of aeroelasticity of airfoils.
Computed Flows In A Transonic Turbine
NASA Technical Reports Server (NTRS)
Rangwalla, A. A.; Madavan, N. K.; Johnson, P. D.
1993-01-01
Report presents computational study of flow in first stage of three alternative versions of proposed transonic turbine. Study demonstrates application of computational fluid dynamics to predict performance and analyze effects of changes in designs of these advanced machines.
Recent advances in transonic computational aeroelasticity
NASA Technical Reports Server (NTRS)
Batina, John T.; Bennett, Robert M.; Seidel, David A.; Cunningham, Herbert J.; Bland, Samuel R.
1988-01-01
A transonic unsteady aerodynamic and aeroelasticity code called CAP-TSD was developed for application to realistic aircraft configurations. The code permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis in the flutter critical transonic speed range. The CAP-TSD code uses a time accurate approximate factorization algorithm for solution of the unsteady transonic small disturbance potential equation. An overview is given of the CAP-TSD code development effort and results are presented which demonstrate various capabilities of the code. Calculations are presented for several configurations including the General Dynamics 1/9 scale F-16 aircraft model and the ONERA M6 wing. Calculations are also presented from a flutter analysis of a 45 deg sweptback wing which agrees well with the experimental data. Descriptions are presented of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate these recent developments in transonic computational aeroelasticity.
Inviscid transonic flow computations with shock fitting
NASA Technical Reports Server (NTRS)
Yu, N. J.; Seebass, A. R.
1975-01-01
First-and second-order numerical procedures are presented for calculating two-dimensional transonic flows that treat shock waves as discontinuities. Their application to a simple but nontrivial problem for which there are limited theoretical results is discussed.
NASA Astrophysics Data System (ADS)
Main, A. J.; Day, C. R. B.; Lock, G. D.; Oldfield, M. L. G.
1996-08-01
A four-hole pyramid probe has been calibrated for use in a short-duration transonic turbine cascade tunnel. The probe is used to create area traverse maps of total and static pressure, and pitch and yaw angles of the flow downstream of a transonic annular cascade. This data is unusual in that it was acquired in a short-duration (5 s of run time) annular cascade blowdown tunnel. A four-hole pyramid probe was used which has a 2.5 mm section head, and has the side faces inclined at 60° to the flow to improve transonic performance. The probe was calibrated in an ejector driven, perforated wall transonic tunnel over the Mach number range 0.5 1.2, with pitch angles from -20° to + 20° and yaw angles from-23° to +23°. A computer driven automatic traversing mechanism and data collection system was used to acquire a large probe calibration matrix (˜ 10,000 readings) of non dimensional pitch, yaw, Mach number, and total pressure calibration coefficients. A novel method was used to transform the probe calibration matrix of the raw coefficients into a probe application matrix of the physical flow variables (pitch, yaw, Mach number etc.). The probe application matrix is then used as a fast look-up table to process probe results. With negligible loss of accuracy, this method is faster by two orders of magnitude than the alternative of global interpolation on the raw probe calibration matrix. The blowdown tunnel (mean nozzle guide vane blade ring diameter 1.1 m) creates engine representative Reynolds numbers, transonic Mach numbers and high levels (≈ 13%) of inlet turbulence intensity. Contours of experimental measurements at three different engine relevant conditions and two axial positions have been obtained. An analysis of the data is presented which includes a necessary correction for the finite velocity of the probe. Such a correction is non trivial for the case of fast moving probes in compressible flow.
Transonic airfoil design using Cartesian coordinates
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1976-01-01
A numerical technique for designing transonic airfoils having a prescribed pressure distribution (the inverse problem) is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that inverse boundary conditions and Cartesian coordinates are used. The method is a direct-inverse approach that controls trailing-edge closure. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.
A model for transonic plasma flow
Guazzotto, Luca; Hameiri, Eliezer
2014-02-15
A linear, two-dimensional model of a transonic plasma flow in equilibrium is constructed and given an explicit solution in the form of a complex Laplace integral. The solution indicates that the transonic state can be solved as an elliptic boundary value problem, as is done in the numerical code FLOW [Guazzotto et al., Phys. Plasmas 11, 604 (2004)]. Moreover, the presence of a hyperbolic region does not necessarily imply the presence of a discontinuity or any other singularity of the solution.
Transonic rotor noise: Theoretical and experimental comparisons
NASA Technical Reports Server (NTRS)
Schmitz, F. H.; Yu, Y. H.
1980-01-01
Two complementary methods of describing the high speed rotor noise problem are discussed. The first method uses the second order transonic potential equation to define and characterize the nature of the aerodynamic and acoustic fields and to explain the appearance of radiating shock waves. The second employs the Ffowcs Williams and Hawkings equation to successfully calculate the acoustic far field. Good agreement between theoretical and experimental waveforms is shown for transonic hover tip Mach numbers from 0.8 to 0.9.
Calculations Of Transonic Flow About A Wing
NASA Technical Reports Server (NTRS)
Holst, Terry L.; Gundy, Karen L.; Flores, Jolen; Chaderjian, Neal; Kaynak, Univer; Thomas, Scott D.
1988-01-01
Report describes calculations of transonic airflows about wing in wind tunnel. Basic equations of flow used in study are Reynolds-averaged Navier-Stokes equations in strong conservation-law form. Equations of flow incorporated into finite-difference computer code called TNS (Transonic Navier-Stokes). Computational grid generated by solution of partial differential equations yielding smooth meshes conforming to surfaces of wing and wind tunnel.
Quasi 1-D Study of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance
NASA Technical Reports Server (NTRS)
Morris, Christopher I.
2002-01-01
Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE is described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of several different PDRE nozzle configurations is reported.
N-231 High Reynolds Number Channel I is a blowdown Facility that utilizes interchangeable test
NASA Technical Reports Server (NTRS)
1980-01-01
N-231 High Reynolds Number Channel I is a blowdown Facility that utilizes interchangeable test sections and nozzles. The facility provides experimental support for the fluid mechanics research, including experimental verification of aerodynamic computer codes and boundary-layer and airfoil studies that require high Reynolds number simulation. (Tunnel 1)
Application of RELAP5 to a pipe blowdown experiment. [PWR; BWR
Carlson, K.E.; Ransom, V.H.; Wagner, R.J.
1980-01-01
The application of the RELAP5 computer program to a pipe blowdown experiment is described in this paper. The basic hydrodynamic model, constitutive relations, and special process models included in RELAP5 are also briefly discussed. The results of this application confirm the effectiveness of using a choked flow model.
Study of design and analysis methods for transonic flow
NASA Technical Reports Server (NTRS)
Murman, E. M.
1977-01-01
An airfoil design program and a boundary layer analysis were developed. Boundary conditions were derived for ventilated transonic wind tunnels and performing transonic windtunnel wall calculations. A computational procedure for rotational transonic flow in engine inlet throats was formulated. Results and conclusions are summarized.
Experience with transonic unsteady aerodynamic calculations
NASA Technical Reports Server (NTRS)
Edwards, J. W.; Bland, S. R.; Seidel, D. A.
1984-01-01
Comparisons of calculated and experimental transonic unsteady pressures and airloads for four of the AGARD Two Dimensional Aeroelastic Configurations and for a rectangular supercritical wing are presented. The two dimensional computer code, XTRAN2L, implementing the transonic small perturbation equation was used to obtain results for: (1) pitching oscillations of the NACA 64A010A; NLR 7301 and NACA 0012 airfoils; (2) flap oscillations for the NACA 64A006 and NRL 7301 airfoils; and (3) transient ramping motions for the NACA 0012 airfoils. Results from the three dimensional code XTRAN3S are compared with data from a rectangular supercritical wing oscillating in pitch. These cases illustrate the conditions under which the transonic inviscid small perturbation equation provides reasonable predictions.
Turbulence and modeling in transonic flow
NASA Technical Reports Server (NTRS)
Rubesin, Morris W.; Viegas, John R.
1989-01-01
A review is made of the performance of a variety of turbulence models in the evaluation of a particular well documented transonic flow. This is done to supplement a previous attempt to calibrate and verify transonic airfoil codes by including many more turbulence models than used in the earlier work and applying the calculations to an experiment that did not suffer from uncertainties in angle of attack and was free of wind tunnel interference. It is found from this work, as well as in the earlier study, that the Johnson-King turbulence model is superior for transonic flows over simple aerodynamic surfaces, including moderate separation. It is also shown that some field equation models with wall function boundary conditions can be competitive with it.
Transonic wing analysis using advanced computational methods
NASA Technical Reports Server (NTRS)
Henne, P. A.; Hicks, R. M.
1978-01-01
This paper discusses the application of three-dimensional computational transonic flow methods to several different types of transport wing designs. The purpose of these applications is to evaluate the basic accuracy and limitations associated with such numerical methods. The use of such computational methods for practical engineering problems can only be justified after favorable evaluations are completed. The paper summarizes a study of both the small-disturbance and the full potential technique for computing three-dimensional transonic flows. Computed three-dimensional results are compared to both experimental measurements and theoretical results. Comparisons are made not only of pressure distributions but also of lift and drag forces. Transonic drag rise characteristics are compared. Three-dimensional pressure distributions and aerodynamic forces, computed from the full potential solution, compare reasonably well with experimental results for a wide range of configurations and flow conditions.
Critical air/water blow-down in safety valves at low qualities.
Moncalvo, D; Friedel, L
2011-02-28
Critical air/water blow-downs in safety valves for qualities from 0.01 to 0.113 and mass flow rates from 1.5 up to 4.3 kg/s have been observed in our test facility. These critical blow-downs are characterized by a large void fraction and by an intense mixing of the phases both in the valve body and in the outlet pipe. A qualitative estimation of the flow pattern in the outlet pipe using the map of Taitel and Dukler suggests that these air/water flows are intermittent flows--presumably slug flows--evolving to annular flows for qualities above 0.1. Intermittent flows are also predicted for critical air/water and air/glycerine flows taken from the literature for the same safety valve at slightly larger relieving pressures. PMID:21227579
Performance of Blowdown Turbine Driven by Exhaust Gas of Nine-Cylinder Radial Engine
NASA Technical Reports Server (NTRS)
Turner, L Richard; Desmon, Leland G
1944-01-01
An investigation was made of an exhaust-gas turbine having four separate nozzle boxes each covering a 90 degree arc of the nozzle diaphragm and each connected to a pair of adjacent cylinders of a nine-cylinder radial engine. This type of turbine has been called a "blowdown" turbine because it recovers the kinetic energy developed in the exhaust stacks during the blowdown period, that is the first part of the exhaust process when the piston of the reciprocating engine is nearly stationary. The purpose of the investigation was to determine whether the blow turbine could develop appreciable power without imposing any large loss in engine power arising from restriction of the engine exhaust by the turbine.
Silling, S.A.; Gross, M.B.; Santee, G.E. Jr.; Chang, F.H.
1981-01-01
The STEALTH 3D and WHAMSE 3D computer codes have been combined to perform three-dimensional coupled fluid/structure calculations of the blowdown response of a pressure vessel with internal structures typical of a pressurized water reactor. The fluid/structure coupling, which is performed cycle by cycle during a calculation, is described. The coupled fluid/structure code, STEALTH/WHAMSE 3D, has been used to simulate the decompression of test V31.1 from the HDR blowdown test series. Calculations of fluid pressure, differential fluid pressure and hoop strain compare favorably with the experimental data from test V31.1. The computed peak axial stain compares less favorably with the experimental data, probably due to coarseness of the structural grid. 14 refs.
Coroneos, A.C.; Taylor, P.A.; Arnold, W.D. Jr.; Bostick, D.A.; Perona, J.J.
1994-12-01
The purpose of this report is to document the results of bench-scale testing completed to remove {sup 137}Cs and {sup 90}Sr from the Oak Ridge K-25 Site Toxic Substances Control Act (TSCA) Incinerator blowdown at the K-25 Site Central Neutralization Facility, a wastewater treatment facility designed to remove heavy metals and uranium from various wastewaters. The report presents results of bench-scale testing using chabazite and clinoptilolite zeolites to remove cesium and strontium; using potassium cobalt ferrocyanide (KCCF) to remove cesium; and using strontium chloride coprecipitation, sodium phosphate coprecipitation, and calcium sulfate coprecipitation to remove strontium. Low-range, average-range, and high-range concentration blowdown surrogates were used to complete the bench-scale testing.
Application of radial-splitters for improved wide-angle diffuser performance in a blowdown tunnel
NASA Technical Reports Server (NTRS)
Rao, D. M.; Seshadri, S. N.
1976-01-01
Severe flow separation in the 15:1 area-ratio, 38 deg total angle conical diffuser preceding the settling-chamber of an intermittent blowdown wind tunnel was eliminated by the use of a novel radial-splitter arrangement. As a consequence, the operating life of settling-chamber screens was greatly extended and test-section flow steadiness improved, with no penalty in the tunnel running time.
D'Amato, A.W.; Fraver, S.; Palik, B.J.; Bradford, J.B.; Patty, L.
2011-01-01
The role of disturbance in structuring vegetation is widely recognized; however, we are only beginning to understand the effects of multiple interacting disturbances on ecosystem recovery and development. Of particular interest is the impact of post-disturbance management interventions, particularly in light of the global controversy surrounding the effects of salvage logging on forest ecosystem recovery. Studies of salvage logging impacts have focused on the effects of post-disturbance salvage logging within the context of a single natural disturbance event. There have been no formal evaluations of how these effects may differ when followed in short sequence by a second, high severity natural disturbance. To evaluate the impact of this management practice within the context of multiple disturbances, we examined the structural and woody plant community responses of sub-boreal Pinus banksiana systems to a rapid sequence of disturbances. Specifically, we compared responses to Blowdown (B), Fire (F), Blowdown-Fire, and Blowdown-Salvage-Fire (BSF) and compared these to undisturbed control (C) stands. Comparisons between BF and BSF indicated that the primary effect of salvage logging was a decrease in the abundance of structural legacies, such as downed woody debris and snags. Both of these compound disturbance sequences (BF and BSF), resulted in similar woody plant communities, largely dominated by Populus tremuloides; however, there was greater homogeneity in community composition in salvage logged areas. Areas experiencing solely fire (F stands) were dominated by P. banksiana regeneration, and blowdown areas (B stands) were largely characterized by regeneration from shade tolerant conifer species. Our results suggest that salvage logging impacts on woody plant communities are diminished when followed by a second high severity disturbance; however, impacts on structural legacies persist. Provisions for the retention of snags, downed logs, and surviving trees as part
Inverse transonic airfoil design including viscous interaction
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1976-01-01
A numerical technique was developed for the analysis of specified transonic airfoils or for the design of airfoils having a prescribed pressure distribution, including the effect of weak viscous interaction. The method uses the full potential equation, a stretched Cartesian coordinate system, and the Nash-MacDonald turbulent boundary layer method. Comparisons with experimental data for typical transonic airfoils show excellent agreement. An example shows the application of the method to design a thick aft-cambered airfoil, and the effects of viscous interaction on its performance are discussed.
A vapor generator for transonic flow visualization
NASA Technical Reports Server (NTRS)
Bruce, Robert A.; Hess, Robert W.; Rivera, Jose A., Jr.
1989-01-01
A vapor generator was developed for use in the NASA Langley Transonic Dynamics Tunnel (TDT). Propylene glycol was used as the vapor material. The vapor generator system was evaluated in a laboratory setting and then used in the TDT as part of a laser light sheet flow visualization system. The vapor generator provided satisfactory seeding of the air flow with visible condensate particles, smoke, for tests ranging from low subsonic through transonic speeds for tunnel total pressures from atmospheric pressure down to less than 0.1 atmospheric pressure.
MELCOR 1.8.3 assessment: GE large vessel blowdown and level swell experiments
Kmetyk, L.N.
1994-07-01
MELCOR is a fully integrated, engineering-level computer code, being developed at Sandia National Laboratories for the USNRC, that models the entire spectrum of severe accident phenomena in a unified framework for both BWRs and PWRS. As part of an ongoing assessment program, the MELCOR computer code has been used to analyze a series of blowdown tests performed in the early 1980s at General Electric. The GE large vessel blowdown and level swell experiments are a set of primary system thermal/hydraulic separate effects tests studying the level swell phenomenon for BWR transients and LOCAS; analysis of these GE tests is intended to validate the new implicit bubble separation algorithm added since the release of MELCOR 1.8.2. Basecase MELCOR results are compared to test data, and a number of sensitivity studies on input modelling parameters and options have been done. MELCOR results for these experiments also are compared to MAAP and TRAC-B qualification analyses for the same tests. Time-step and machine-dependency calculations were done to identify whether any numeric effects exist in our GE large vessel blowdown and level swell assessment analyses.
The transonic Reynolds number problem. [limitations of transonic aerodynamic test facilities
NASA Technical Reports Server (NTRS)
Jones, J. L.
1977-01-01
Problems in modeling the complex interacting flow fields in the transonic speed regime are reviewed. The limitations of wind tunnel test capabilities are identified, and options for resolving the deficiency are examined. The evolution of the National Transonic Facility, and the various needs for research investigations to be done there are discussed. The relative priorities that should be given within and across subdisciplines for guidance in planning for the most effective use of the facility are considered.
Some iterative schemes for transonic potential flows
NASA Technical Reports Server (NTRS)
Wong, Y. S.; Hafez, M. M.
1985-01-01
The minimal residual (MR) method for the numerical solution of transonic potential flows is closely related to the conjugate gradient method, which has found widespread use in the solution of large sparse, symmetric, and positive-definite linear equations. The primary advantage of the MR method is its applicability to both symmetric and nonsymmetric matrices.
Pyroclastic current dynamic pressure from aerodynamics of tree or pole blow-down
NASA Astrophysics Data System (ADS)
Clarke, A. B.; Voight, B.
2000-07-01
The common occurrence of tree and pole blow-down from pyroclastic currents provides an opportunity to estimate properties of the currents. Blow-down may occur by uprooting (root zone rupture), or flexure or shear at some point on the object. If trees are delimbed before blow-down, each tree or pole can be simulated by a cylinder perpendicular to the current. The force acting on a cylinder is a function of flow dynamic pressure, cylinder geometry, and drag coefficient. Treated as a cantilever of circular cross-section, the strength for the appropriate failure mode (rupture, uprooting or flexure) can then be used to estimate the minimum necessary current dynamic pressure. In some cases, larger or stronger standing objects can provide upper bounds on the dynamic pressure. This analysis was treated in two ways: (1) assuming that the current properties are vertically constant; and (2) allowing current velocity and density to vary vertically according to established models for turbulent boundary layers and stratified flow. The two methods produced similar results for dynamic pressure. The second, along with a method to approximate average whole-current density, offers a means to estimate average velocity and density over the height of the failed objects. The method is applied to several example cases, including Unzen, Mount St. Helens, Lamington, and Merapi volcanoes. Our results compare reasonably well with independent estimates. For several cases, we found that it is possible to use the dynamic pressure equations developed for vertically uniform flow, along with the average cloud density multiplied by a factor of 2-5, to determine average velocity over the height of the failed object.
Investigation of Minimum Film boiling Phenomena on Fuel Rods Under Blowdown Cooling Conditions
Stephen M. Bajorek; Michael Gawron; Timothy Etzel; Lucas Peterson
2003-06-30
Blowdon cooling heat transfer is an important process that occurs early in a hypothetical large break loss-of-coolant accident (LOCA) in a pressurized water reactor. During blowdown, the flow through the hot assembly is a post-critical heat flux dispersed droplet flow. The heat transfer mechanisms that occur in blowdown cooling are complex and depend on droplet and heated surface interaction. In a safety analysis, it is of considerable importance to determine the thermal-hydraulic conditions leading to the minimum film boiling temperature, Tmin. A flow boiling rig for measurement of blowdown cooling heat transfer and quench phenomena on a nuclear fuel rod simulator was designed and constructed for operation at up to 12.4 MPa. The test section consisted of a concentric annulus, with a 9.5 mm OD nuclear fuel rod simulator at the center. The rod was contained within a 0.85 mm thick, 19 mm OD 316 stainless steel tube, forming the flow channel. Two types of rods were tested; one type was sheathed with Inconel 600 while the other was clad with Zircaloy-2. Water was injected into the test section at the top of the heated length through an injection header. This header was an annular sign that fit around the fuel rod simulator and within the stainless steel tube. Small spacers aligned the injection header and prevented contract with either the heater rod or the tube. A series of small diameter holes at the bottom of the header caused the formation of droplets that became entrained with the steam flow. The test section design was such that quench would take place on the rod, and not along the channel outer annulus.
Blow-down and blow-in of Inland`s No. 7 blast furnace
Ricketts, J.; Quisenberry, P.; Carter, W.
1995-12-01
After extensive and detailed planning, a mini-reline of the 13.7 meter No. 7 Blast Furnace was executed in November 1993. The furnace lining had 18 million metric tons of production and the bosh, belly and lower stack lining were being maintained through a scheduled grouting practice. The mini-reline was planned for 33 days and the reline work included (a) replacing the bosh, belly and lower stack alumina lining with graphite brick, (b) gunning the middle and upper stack, (c) rebuilding the furnace top, stove burners and tapholes and (d) minor repairs to other auxiliary equipment. During this 33 day reline period the two 8 meter furnaces could only produce 40% of the normal production requirement, therefore the blow-down, quench, salamander tap and blow-in activities were critical to meeting the planned schedule. The planning of these activities was started in the spring of 1993 and included review of Inland`s past blow-down and blow-in performance as well as bench marking the performance of other large blast furnaces in North America, Japan and Europe. The development of the 1993 procedures focused on opportunities to accelerate the blow-down, quench, salamander tap and blow-in as well as having a clean hearth and stack which could also save time during the demolition phase of the reline. Any time that could be saved in these activities directly translated to an early start-up and more plantwide production. This paper will cover the successful planning and implementation of these activities which resulted in a 2 day reduction in the reline schedule, an accelerated production curve and an earlier than planned use of PCI during blow-in.
Transonic Symposium: Theory, Application, and Experiment, volume 1, part 2
NASA Technical Reports Server (NTRS)
Foughner, Jerome T., Jr. (Compiler)
1989-01-01
In order to assess the state of the art in transonic flow disciplines and to glimpse at future directions, NASA-Langley held a Transonic Symposium. Emphasis was placed on steady, three dimensional external, transonic flow and its simulation, both numerically and experimentally. The symposium included technical sessions on wind tunnel and flight experiments; computational fluid dynamic applications; inviscid methods and grid generation; viscous methods and boundary layer stability; and wind tunnel techniques and wall interference. This, being volume 1, is unclassified.
Transonic Symposium: Theory, Application and Experiment, volume 2
NASA Technical Reports Server (NTRS)
Foughner, Jerome T., Jr. (Compiler)
1989-01-01
Papers presented at the Transonic Symposium are compiled. The following subject areas are covered: National Transonic Facility status; transonic aerodynamics of slender wing-body configuration; laminar flow flight experiments; laminar flow wind tunnel experiments; computational support of X-29A flight experiment; transition location on a clean-up glove installed on a F-14 aircraft; and design studies for a laminar glove for the X-29 aircraft.
A theoretical and numerical investigation of turbulent steam jets in BWR steam blowdown.
NguyenLe, Q.
1998-06-26
The preliminary results of PHOENICS and RELAP5 show that the current numerical models are adequate in predicting steam flow and stratification patterns in the upper Drywell of a BWR containment subsequent to a blow-down event. However, additional modeling is required in order to study detailed local phenomena such as condensation with non-condensables, natural convection, and stratification effects. Analytically, the intermittence modified similarity solutions show great promise. Once {gamma} is accounted for, the jet's turbulent shear stress can be determined with excellent accuracy.
Transonic turbine blade cascade testing facility
NASA Technical Reports Server (NTRS)
Verhoff, Vincent G.; Camperchioli, William P.; Lopez, Isaac
1992-01-01
NASA LeRC has designed and constructed a new state-of-the-art test facility. This facility, the Transonic Turbine Blade Cascade, is used to evaluate the aerodynamics and heat transfer characteristics of blade geometries for future turbine applications. The facility's capabilities make it unique: no other facility of its kind can combine the high degree of airflow turning, infinitely adjustable incidence angle, and high transonic flow rates. The facility air supply and exhaust pressures are controllable to 16.5 psia and 2 psia, respectively. The inlet air temperatures are at ambient conditions. The facility is equipped with a programmable logic controller with a capacity of 128 input/output channels. The data acquisition system is capable of scanning up to 1750 channels per sec. This paper discusses in detail the capabilities of the facility, overall facility design, instrumentation used in the facility, and the data acquisition system. Actual research data is not discussed.
Kuechemann Carrots for transonic drag reduction.
NASA Astrophysics Data System (ADS)
Bechert, D. W.; Hage, W.; Stanewsky, E.
1999-11-01
Wave drag reduction bodies on the suction side of transonic wings are investigated. Following the original invention by O. Frenzl (1942), subsequently, such bodies have been suggested by Kuechemann and Whitcomb. These devices have been used sucessfully on various TUPOLEV aircraft and on the CONVAIR 990 airliner. New transonic wind tunnel data from an unswept wing with an array of Kuechemann Carrots are presented (airfoil: CAST 10/DOA-2). In a certain parameter range (M= 0.765-0.86) the measurements exhibit a significant reduction of the shock strength on a wing between the Kuechemann Carrots. This entails a dramatic reduction of drag, in a certain Mach number and angular regime up to 50-60%.
Viscous Transonic Airfoil Workshop compendium of results
NASA Technical Reports Server (NTRS)
Holst, Terry L.
1987-01-01
Results from the Viscous Transonic Airfoil Workshop held at the AIAA 25th Aerospace Sciences Meeting at Reno, NV in January 1987, are compared with each other and with experimental data. Test cases used in this workshop include attached and separated transonic flows for three different airfoils: the NACA 0012 airfoil, the RAE 2822 airfoil, and the Jones airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical methods used vary widely and include: 16 Navier-Stokes methods, 2 Euler/boundary-layer methods, and 5 full-potential/boundary-layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately-separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.
Analysis of a theoretically optimized transonic airfoil
NASA Technical Reports Server (NTRS)
Lores, M. E.; Burdges, K. P.; Shrewsbury, G. D.
1978-01-01
Numerical optimization was used in conjunction with an inviscid, full potential equation, transonic flow analysis computer code to design an upper surface contour for a conventional airfoil to improve its supercritical performance. The modified airfoil was tested in a compressible flow wind tunnel. The modified airfoil's performance was evaluated by comparison with test data for the baseline airfoil and for an airfoil developed by optimization of leading edge of the baseline airfoil. While the leading edge modification performed as expected, the upper surface re-design did not produce all of the expected performance improvements. Theoretical solutions computed using a full potential, transonic airfoil code corrected for viscosity were compared to experimental data for the baseline airfoil and the upper surface modification. These correlations showed that the theory predicted the aerodynamics of the baseline airfoil fairly well, but failed to accurately compute drag characteristics for the upper surface modification.
Buffet test in the National Transonic Facility
NASA Technical Reports Server (NTRS)
Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.
1992-01-01
A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk to the facility. This paper presents the test results from a structural dynamics and aeroelastic response point of view and describes the activities required for the safety analysis and risk assessment. The test was conducted in the same manner as a flutter test and employed onboard dynamic instrumentation, real time dynamic data monitoring, automatic, and manual tunnel interlock systems for protecting the model. The procedures and test techniques employed for this test are expected to serve as the basis for future aeroelastic testing in the National Transonic Facility. This test program was a cooperative effort between the Boeing Commercial Airplane Company and the NASA Langley Research Center.
Vector processor algorithms for transonic flow calculations
NASA Technical Reports Server (NTRS)
South, J. C., Jr.; Keller, J. D.; Hafez, M. M.
1979-01-01
This paper discusses a number of algorithms for solving the transonic full-potential equation in conservative form on a vector computer, such as the CDC STAR-100 or the CRAY-1. Recent research with the 'artificial density' method for transonics has led to development of some new iteration schemes which take advantage of vector-computer architecture without suffering significant loss of convergence rate. Several of these more promising schemes are described and 2-D and 3-D results are shown comparing the computational rates on the STAR and CRAY vector computers, and the CYBER-175 serial computer. Schemes included are: (1) Checkerboard SOR, (2) Checkerboard Leapfrog, (3) odd-even vertical line SOR, and (4) odd-even horizontal line SOR.
Transonic and supersonic ground effect aerodynamics
NASA Astrophysics Data System (ADS)
Doig, G.
2014-08-01
A review of recent and historical work in the field of transonic and supersonic ground effect aerodynamics has been conducted, focussing on applied research on wings and aircraft, present and future ground transportation, projectiles, rocket sleds and other related bodies which travel in close ground proximity in the compressible regime. Methods for ground testing are described and evaluated, noting that wind tunnel testing is best performed with a symmetry model in the absence of a moving ground; sled or rail testing is ultimately preferable, though considerably more expensive. Findings are reported on shock-related ground influence on aerodynamic forces and moments in and accelerating through the transonic regime - where force reversals and the early onset of local supersonic flow is prevalent - as well as more predictable behaviours in fully supersonic to hypersonic ground effect flows.
Flow instabilities in transonic small disturbance theory
NASA Technical Reports Server (NTRS)
Williams, M. H.; Bland, S. R.; Edwards, J. W.
1985-01-01
The dynamics of unsteady transonic small disturbance flows about two-dimensional airfoils is examined, with emphasis on the behavior in the region where the steady state flow is nonunique. It is shown that nonuniqueness results from an extremely long time scale instability which occurs in a finite Mach number and angle of attack range. The similarity scaling rules for the instability are presented and the possibility of similar behavior in the Euler equations is discussed.
Magnus effects on spinning transonic missiles
NASA Technical Reports Server (NTRS)
Seginer, A.; Rosenwasser, I.
1983-01-01
Magnus forces and moments were measured on a basic-finner model spinning in transonic flow. Spin was induced by canted fins or by full-span or semi-span, outboard and inboard roll controls. Magnus force and moment reversals were caused by Mach number, reduced spin rate, and angle of attack variations. Magnus center of pressure was found to be independent of the angle of attack but varied with the Mach number and model configuration or reduced spin rate.
Recent Enhancements to the National Transonic Facility
NASA Technical Reports Server (NTRS)
Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W.; Underwood, P.
2003-01-01
The National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the restoration of reliability and improved performance of the heat exchanger systems resulting in the expansion of the NTF air operations envelope. Additionally, results are presented from a continued effort to reduce model dynamics through the use of a new stiffer balance and sting
Recent Enhancements to the National Transonic Facility
NASA Technical Reports Server (NTRS)
Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W.; Underwood, P.
2003-01-01
The National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the restoration of reliability and improved performance of the heat exchanger systems resulting in the expansion of the NTF air operations envelope. Additionally, results are presented from a continued effort to reduce model dynamics through the use of a new stiffer balance and sting.
Transonic airfoil and axial flow rotary machine
Nagai, Naonori; Iwatani, Junji
2015-09-01
Sectional profiles close to a tip 124 and a part between a midportion 125 and a hub 123 are shifted to the upstream of an operating fluid flow in a sweep direction. Accordingly, an S shape is formed in which the tip 124 and the part between the midportion 125 and the hub 123 protrude. As a result, it is possible reduce various losses due to shook, waves, thereby forming a transonic airfoil having an excellent aerodynamic characteristic.
Studying Transonic Gases With a Hydraulic Analog
NASA Technical Reports Server (NTRS)
Wagner, W.; Lepore, F.
1986-01-01
Water table for hydraulic-flow research yields valuable information about gas flow at transonic speeds. Used to study fuel and oxidizer flow in high-pressure rocket engines. Method applied to gas flows in such equipment as furnaces, nozzles, and chemical lasers. Especially suitable when wall contours nonuniform, discontinuous, or unusually shaped. Wall shapes changed quickly for study and evaluated on spot. Method used instead of computer simulation when computer models unavailable, inaccurate, or costly to run.
Transonic cryogenic test section for the Goettingen tube facility
NASA Technical Reports Server (NTRS)
Hornung, H.; Hefer, G.; Krogmann, P.; Stanewsky, E.
1983-01-01
The design of modern aircraft requires the solution of problems related to transonic flow at high Reynolds numbers. To investigate these problems experimentally, it is proposed to extend the Ludwieg tube facility by adding a transonic cryogenic test section. After stating the requirements for such a test section, the technical concept is briefly explained and a preliminary estimate of the costs is given.
ATRAN3S: An unsteady transonic code for clean wings
NASA Technical Reports Server (NTRS)
Guruswamy, G. P.; Goorjian, P. M.; Merritt, F. J.
1985-01-01
The development and applications of the unsteady transonic code ATRAN3S for clean wings are discussed. Explanations of the unsteady, transonic small-disturbance aerodynamic equations that are used and their solution procedures are discussed. A detailed user's guide, along with input and output for a sample case, is given.
Nathan D. Jerred; Robert C. O'Brien; Steven D. Howe; James E. O'Brien
2013-02-01
Recent developments at the Center for Space Nuclear Research (CSNR) on a Martian exploration probe have lead to the assembly of a multi-functional variable atmosphere testing facility (VATF). The VATF has been assembled to perform transient blow-down analysis of a radioisotope thermal rocket (RTR) concept that has been proposed for the Mars Hopper; a long-lived, long-ranged mobile platform for the Martian surface. This study discusses the current state of the VATF as well as recent blow-down testing performed on a laboratory-scale prototype of the Mars Hopper. The VATF allows for the simulation of Mars ambient conditions within the pressure vessel as well as to safely perform blow-down tests through the prototype using CO2 gas; the proposed propellant for the Mars Hopper. Empirical data gathered will lead to a better understanding of CO2 behavior and will provide validation of simulation models. Additionally, the potential of the VATF to test varying propulsion system designs has been recognized. In addition to being able to simulate varying atmospheres and blow-down gases for the RTR, it can be fitted to perform high temperature hydrogen testing of fuel elements for nuclear thermal propulsion.
Comparison of the calculation of HDR RPV-1 blowdown loads for test V32 with the experimental data
Mortensen, G.A.; Santee, G.E. Jr.; Chang, F.H.; Brockett, G.F.; Gross, M.; Belytschko, T.
1984-08-01
Using the three-dimensional STEALTH/WHAMSE computer program, researchers calculated blowdown loads for HDR test V32 on a large-scale reactor vessel. The calculation, which compared well with test results, provides a good example of the predictive capability of the computer program. The program can be applied to nuclear and nonnuclear fluid-structure interaction problems.
Transonic potential flow in hyperbolic nozzles
NASA Technical Reports Server (NTRS)
Park, M.; Caughey, D. A.
1986-01-01
The full potential equation for the classical problem of transonic flow through a hyperbolic nozzle (with or without a shock wave) is solved in conservation form using the finite volume method of Jameson and Caughey (1977). Either a firstor a second-order numerical viscosity is added in the direction of the flow, explicitly, in conservation form. A multigrid alternating direction implicit method is used to solve the difference equations, and the results obtained are compared with analytical and numerical results from previous researches.
Buffet test in the National Transonic Facility
NASA Technical Reports Server (NTRS)
Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.
1992-01-01
A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk for the facility. Presented here are the test results from a structural dynamics and aeroelastic response point of view. The activities required for the safety analysis and risk assessment are described. The test was conducted in the same manner as a flutter test and employed on-board dynamic instrumentation, real time dynamic data monitoring, and automatic and manual tunnel interlock systems for protecting the model.
Computational methods for unsteady transonic flows
NASA Technical Reports Server (NTRS)
Edwards, John W.; Thomas, James L.
1987-01-01
Computational methods for unsteady transonic flows are surveyed with emphasis upon applications to aeroelastic analysis and flutter prediction. Computational difficulty is discussed with respect to type of unsteady flow; attached, mixed (attached/separated) and separated. Significant early computations of shock motions, aileron buzz and periodic oscillations are discussed. The maturation of computational methods towards the capability of treating complete vehicles with reasonable computational resources is noted and a survey of recent comparisons with experimental results is compiled. The importance of mixed attached and separated flow modeling for aeroelastic analysis is discussed and recent calculations of periodic aerodynamic oscillations for an 18 percent thick circular arc airfoil are given.
Computational methods for unsteady transonic flows
NASA Technical Reports Server (NTRS)
Edwards, John W.; Thomas, J. L.
1987-01-01
Computational methods for unsteady transonic flows are surveyed with emphasis on prediction. Computational difficulty is discussed with respect to type of unsteady flow; attached, mixed (attached/separated) and separated. Significant early computations of shock motions, aileron buzz and periodic oscillations are discussed. The maturation of computational methods towards the capability of treating complete vehicles with reasonable computational resources is noted and a survey of recent comparisons with experimental results is compiled. The importance of mixed attached and separated flow modeling for aeroelastic analysis is discussed, and recent calculations of periodic aerodynamic oscillations for an 18 percent thick circular arc airfoil are given.
Numerical simulation of small perturbation transonic flows
NASA Technical Reports Server (NTRS)
Seebass, A. R.; Yu, N. J.
1976-01-01
The results of a systematic study of small perturbation transonic flows are presented. Both the flow over thin airfoils and the flow over wedges were investigated. Various numerical schemes were employed in the study. The prime goal of the research was to determine the efficiency of various numerical procedures by accurately evaluating the wave drag, both by computing the pressure integral around the body and by integrating the momentum loss across the shock. Numerical errors involved in the computations that affect the accuracy of drag evaluations were analyzed. The factors that effect numerical stability and the rate of convergence of the iterative schemes were also systematically studied.
0.3 Meter Transonic Cryogenic Tunnel
NASA Technical Reports Server (NTRS)
1985-01-01
Full Description: The Langley 0.3-Meter Transonic Cryogenic tunnel (0.3-m TCT) is used for testing two-dimensional airfoil sections and other models at high Reynolds numbers. The tunnel can operate continuously over a range of Mach numbers from about 0.1 to above 1.2, with a stagnation pressure from 14.7 to 88.0 psia (1 to 6 atmospheres) and a stagnation temperature from -320F to 130F (78 K to 328 K). This results in a maximum Reynolds number capability in excess of 100 x 106 per foot. The adaptive walls, floor, and ceiling in the 13-in. by 13-in. (33-cm by 33-cm) test section can be moved to the free-stream streamline shape, eliminating or reducing the wall effects on the model. The combination of flight Reynolds numbers capability and minimal wall interference makes the 0.3-m TCT a powerful tool for aeronautical research at transonic speeds. The Mach number, pressure, temperature, and adaptive wall shape are automatically controlled. The test section has computer-controlled angle of attack and traversing wake survey-probe systems. The facility has been modified to also use alternate test media--a heavy gas (sulfur hexafluoride, SF6), or air, both with a newly installed heat exchanger.
Geometrical acoustics and transonic helicopter sound
NASA Technical Reports Server (NTRS)
Isom, Morris; Purcell, Timothy W.; Strawn, Roger C.
1987-01-01
A new method is presented for predicting the impulsive noise generated by a transonic rotor blade. The method is a combined approach involving computational fluid dynamics and geometrical acoustics. A full-potential finite-difference method is used to obtain the pressure field close to the blade. A Kirchhoff integral formulation is then used to extend these finite-difference results into the far field. This Kirchhoff formula is based on geometrical acoustics approximations. It requires initial data across a plane at the sonic radius in a blade-fixed coordinate system. This data is provided by the finite-difference solution. Acoustic pressure predictions show good agreement with hover experimental data for cases with hover tip Mach numbers of 0.88 through 0.96. The cases above 0.92 tip Mach number are dominated by non-linear transonic effects seen as strong shocks on and off the blade tip. This paper gives the first successful predictions of far-field acoustic pressures for high-speed impulsive noise over a range of Mach numbers after delocalization.
High-transonic-speed transport aircraft study
NASA Technical Reports Server (NTRS)
Kulfan, R. M.
1974-01-01
An initial design study of high-transonic-speed transport aircraft has been completed. Five different design concepts were developed. These included fixed swept wing, variable-sweep wing, delta wing, double-fuselage yawed-wing, and single-fuselage yawed-wing aircraft. The boomless supersonic design objectives of range = 5560 km (3000 nmi), payload = 18,143 kg (40,000 lb), Mach = 1.2, and FAR Part 36 aircraft noise levels were achieved by the single-fuselage yawed-wing configuration with a gross weight of 211,828 kg (467,000 lb). A noise level of 15 EPNdB below FAR Part 36 requirements was obtained with a gross weight increase to 226,796 kg (500,000 lb). The off-design subsonic range capability for this configuration exceeded the Mach 1.2 design range by more than 20%. Although wing aeroelastic divergence was a primary design consideration for the yawed-wing concepts, the graphite-epoxy wings of this study were designed by critical gust and maneuver loads rather than by divergence requirements. The transonic nacelle drag is shown to be very sensitive to the nacelle installation. A six-degree-of-freedom dynamic stability analysis indicated that the control coordination and stability augmentation system would require more development than for a symmetrical airplane but is entirely feasible. A three-plane development plan is recommended to establish the full potential of the yawed-wing concept.
Numerical calculations of two dimensional, unsteady transonic flows with circulation
NASA Technical Reports Server (NTRS)
Beam, R. M.; Warming, R. F.
1974-01-01
The feasibility of obtaining two-dimensional, unsteady transonic aerodynamic data by numerically integrating the Euler equations is investigated. An explicit, third-order-accurate, noncentered, finite-difference scheme is used to compute unsteady flows about airfoils. Solutions for lifting and nonlifting airfoils are presented and compared with subsonic linear theory. The applicability and efficiency of the numerical indicial function method are outlined. Numerically computed subsonic and transonic oscillatory aerodynamic coefficients are presented and compared with those obtained from subsonic linear theory and transonic wind-tunnel data.
Computational, unsteady transonic aerodynamics and aeroelasticity about airfoils and wings
NASA Technical Reports Server (NTRS)
Goorjian, Peter M.; Guruswamy, Guru P.
1987-01-01
Research in the area of computational, unsteady transonic flows about airfoils and wings, including aeroelastic effects is reviewed. In the last decade, there have been extensive developments in computational methods in response to the need for computer codes with which to study fundamental aerodynamic and aeroelastic problems in the critical transonic regime. For example, large commercial aircraft cruise most effectively in the transonic flight regime and computational fluid dynamics (CDF) provides a new tool, which can be used in combination with test facilities to reduce the costs, time, and risks of aircraft development.
Transonic separated solutions for an augmentor-wing
NASA Technical Reports Server (NTRS)
Flores, J.; Van Dalsem, W. R.
1985-01-01
The viscous transonic flow about a multielement airfoil (augmentor-wing) is simulated by coupling full-potential and direct/inverse differential boundary-layer algorithms. Solutions have been obtained for a variety of conditions and are in fair agreement with available experimental data. Typical results from this transonic augmentor-wing code (TAUG-V) require approximately three minutes of CRAY-XMP CPU time. Since this viscous transonic code accounts for most of the important flow physics, yet is still economical, it is a practical tool for the design aerodynamicist.
Wing analysis using a transonic potential flow computational method
NASA Technical Reports Server (NTRS)
Henne, P. A.; Hicks, R. M.
1978-01-01
The ability of the method to compute wing transonic performance was determined by comparing computed results with both experimental data and results computed by other theoretical procedures. Both pressure distributions and aerodynamic forces were evaluated. Comparisons indicated that the method is a significant improvement in transonic wing analysis capability. In particular, the computational method generally calculated the correct development of three-dimensional pressure distributions from subcritical to transonic conditions. Complicated, multiple shocked flows observed experimentally were reproduced computationally. The ability to identify the effects of design modifications was demonstrated both in terms of pressure distributions and shock drag characteristics.
Analysis of viscous transonic flow over airfoil sections
NASA Technical Reports Server (NTRS)
Huff, Dennis L.; Wu, Jiunn-Chi; Sankar, L. N.
1987-01-01
A full Navier-Stokes solver has been used to model transonic flow over three airfoil sections. The method uses a two-dimensional, implicit, conservative finite difference scheme for solving the compressible Navier-Stokes equations. Results are presented as prescribed for the Viscous Transonic Airfoil Workshop to be held at the AIAA 25th Aerospace Sciences Meeting. The NACA 0012, RAE 2822 and Jones airfoils have been investigated for both attached and separated transonic flows. Predictions for pressure distributions, loads, skin friction coefficients, boundary layer displacement thickness and velocity profiles are included and compared with experimental data when possible. Overall, the results are in good agreement with experimental data.
NASA Technical Reports Server (NTRS)
Chaparro, Daniel; Fujiwara, Gustavo E. C.; Ting, Eric; Nguyen, Nhan
2016-01-01
The need to rapidly scan large design spaces during conceptual design calls for computationally inexpensive tools such as the vortex lattice method (VLM). Although some VLM tools, such as Vorview have been extended to model fully-supersonic flow, VLM solutions are typically limited to inviscid, subcritical flow regimes. Many transport aircraft operate at transonic speeds, which limits the applicability of VLM for such applications. This paper presents a novel approach to correct three-dimensional VLM through coupling of two-dimensional transonic small disturbance (TSD) solutions along the span of an aircraft wing in order to accurately predict transonic aerodynamic loading and wave drag for transport aircraft. The approach is extended to predict flow separation and capture the attenuation of aerodynamic forces due to boundary layer viscosity by coupling the TSD solver with an integral boundary layer (IBL) model. The modeling framework is applied to the NASA General Transport Model (GTM) integrated with a novel control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF).
Techniques for correcting approximate finite difference solutions. [considering transonic flow
NASA Technical Reports Server (NTRS)
Nixon, D.
1978-01-01
A method of correcting finite-difference solutions for the effect of truncation error or the use of an approximate basic equation is presented. Applications to transonic flow problems are described and examples are given.
Unsteady transonic flow calculations for interfering lifting surface configurations
NASA Technical Reports Server (NTRS)
Batina, J. T.
1985-01-01
Unsteady transonic flow calculations are presented for aerodynamically interfering lifting surface configurations. Calculations are performed by extending the XTRAN3S (Version 1.5) unsteady transonic small-disturbance code to allow the treatment of an additional lifting surface. The research was conducted as a first-step toward developing the capability to treat a complete flight vehicle. Grid generation procedures for swept tapered interfering lifting surface applications of XTRAN3S are described. Transonic calculations are presented for wing-tail and canard-wing configurations for several values of mean angle of attack. The effects of aerodynamic interference on transonic steady pressure distributions and steady and oscillatory spanwise lift distributions are demonstrated. Results due to wing, tail, or canard pitching motions are presented and discussed in detail.
Design optimization of axisymmetric bodies in nonuniform transonic flow
NASA Technical Reports Server (NTRS)
Lan, C. Edward
1989-01-01
An inviscid transonic code capable of designing an axisymmetric body in a uniform or nonuniform flow was developed. The design was achieved by direct optimiation by coupling an analysis code with an optimizer. Design examples were provided for axisymmetric bodies with fineness ratios of 8.33 and 5 at different Mach numbers. It was shown that by reducing the nose radius and increasing the afterbody thickness of initial shapes obtained from symmetric NACA four-digit airfoil contours, wave drag could be reduced by 29 percent for a body of fineness ratio 8.33 in a nonuniform transonic flow of M = 0.98 to 0.995. The reduction was 41 percent for a body of fineness ratio 5 in a uniform transonic flow of M = 0.925 and 65 percent for the same body but in a nonuniform transonic flow of M = 0.90 to 0.95.
Active Suppression of the Transonic Flutter Using Sliding Mode Control
NASA Astrophysics Data System (ADS)
Degaki, Takanori; Suzuki, Shinji
This paper describes two-dimensional active flutter suppression to cope with the transonic dip using the sliding mode control. The airfoil model has plunge and pitch degrees of freedom with leading and trailing edge control surfaces. The aerodynamic forces acting on the airfoil, lift and pitching moment, are calculated by solving Euler's equations using computational fluid dynamics. At a specific altitude, flutter occurs between Mach number of 0.7 and 0.88, which corresponds to the transonic dip. The sliding mode control makes the airfoil to be stable all through the Mach number including the transonic dip. The sliding mode controller gives wider flutter margin than a linear quadratic regulator. These characteristics indicate that the sliding mode control is useful for active flutter suppression in the transonic flight.
2. VIEW SOUTH OF TRANSONIC WIND TUNNEL BUILDING AND SUPERSONIC ...
2. VIEW SOUTH OF TRANSONIC WIND TUNNEL BUILDING AND SUPERSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
1. VIEW SOUTHWEST OF SUBSONIC WIND TUNNEL BUILDING AND TRANSONIC ...
1. VIEW SOUTHWEST OF SUBSONIC WIND TUNNEL BUILDING AND TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
7. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC ...
7. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
5. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC ...
5. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC ...
4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
3. VIEW SOUTHEAST OF TRANSONIC WIND TUNNEL BUILDING TO SUBSONIC ...
3. VIEW SOUTHEAST OF TRANSONIC WIND TUNNEL BUILDING TO SUBSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
TAIR- TRANSONIC AIRFOIL ANALYSIS COMPUTER CODE
NASA Technical Reports Server (NTRS)
Dougherty, F. C.
1994-01-01
The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters
Transonic Unsteady Aerodynamics and Aeroelasticity 1987, part 1
NASA Technical Reports Server (NTRS)
Bland, Samuel R. (Compiler)
1989-01-01
Computational fluid dynamics methods have been widely accepted for transonic aeroelastic analysis. Previously, calculations with the TSD methods were used for 2-D airfoils, but now the TSD methods are applied to the aeroelastic analysis of the complete aircraft. The Symposium papers are grouped into five subject areas, two of which are covered in this part: (1) Transonic Small Disturbance (TSD) theory for complete aircraft configurations; and (2) Full potential and Euler equation methods.
Separated transonic airfoil flow calculations with a nonequilibrium turbulence model
NASA Technical Reports Server (NTRS)
King, L. S.; Johnson, D. A.
1985-01-01
Navier-Stokes transonic airfoil calculations based on a recently developed nonequilibrium, turbulence closure model are presented for a supercritical airfoil section at transonic cruise conditions and for a conventional airfoil section at shock-induced stall conditions. Comparisons with experimental data are presented which show that this nonequilibrium closure model performs significantly better than the popular Baldwin-Lomax and Cebeci-Smith equilibrium algebraic models when there is boundary-layer separation that results from the inviscid-viscous interactions.
Transonic Wing Shape Optimization Using a Genetic Algorithm
NASA Technical Reports Server (NTRS)
Holst, Terry L.; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)
2002-01-01
A method for aerodynamic shape optimization based on a genetic algorithm approach is demonstrated. The algorithm is coupled with a transonic full potential flow solver and is used to optimize the flow about transonic wings including multi-objective solutions that lead to the generation of pareto fronts. The results indicate that the genetic algorithm is easy to implement, flexible in application and extremely reliable.
Shin, Y.W.; Wiedermann, A.H.
1984-02-01
A method was published, based on the integral method of characteristics, by which the junction and boundary conditions needed in computation of a flow in a piping network can be accurately formulated. The method for the junction and boundary conditions formulation together with the two-step Lax-Wendroff scheme are used in a computer program; the program in turn, is used here in calculating sample problems related to the blowdown transient of a two-phase flow in the piping network downstream of a PWR pressurizer. Independent, nearly exact analytical solutions also are obtained for the sample problems. Comparison of the results obtained by the hybrid numerical technique with the analytical solutions showed generally good agreement. The good numerical accuracy shown by the results of our scheme suggest that the hybrid numerical technique is suitable for both benchmark and design calculations of PWR pressurizer blowdown transients.
Mach number effects on transonic aeroelastic forces and flutter characteristics
NASA Technical Reports Server (NTRS)
Mohr, Ross W.; Batina, John T.; Yang, Henry T. Y.
1988-01-01
Transonic aeroelastic stability analysis and flutter calculations are presented for a generic transport-type wing based on the use of the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) finite-difference code. The CAP-TSD code was recently developed for transonic unsteady aerodynamic and aeroelastic analysis of complete aircraft configurations. A binary aeroelastic system consisting of simple bending and torsion modes was used to study aeroelastic behavior at transonic speeds. Generalized aerodynamic forces are presented for a wide range of Mach number and reduced frequency. Aeroelastic characteristics are presented for variations in freestream Mach number, mass ratio, and bending-torsion frequency ratio. Flutter boundaries are presented which have two transonic dips in flutter speed. The first dip is the usual transonic dip involving a bending-dominated flutter mode. The second dip is characterized by a single degree-of-freedom torsion oscillation. These aeroelastic results are physically interpreted and shown to be related to the steady state shock location and changes in generalized aerodynamic forces due to freestream Mach number.
Flutter Analysis of a Transonic Fan
NASA Technical Reports Server (NTRS)
Srivastava, R.; Bakhle, M. A.; Keith, T. G., Jr.; Stefko, G. L.
2002-01-01
This paper describes the calculation of flutter stability characteristics for a transonic forward swept fan configuration using a viscous aeroelastic analysis program. Unsteady Navier-Stokes equations are solved on a dynamically deforming, body fitted, grid to obtain the aeroelastic characteristics using the energy exchange method. The non-zero inter-blade phase angle is modeled using phase-lagged boundary conditions. Results obtained show good correlation with measurements. It is found that the location of shock and variation of shock strength strongly influenced stability. Also, outboard stations primarily contributed to stability characteristics. Results demonstrate that changes in blade shape impact the calculated aerodynamic damping, indicating importance of using accurate blade operating shape under centrifugal and steady aerodynamic loading for flutter prediction. It was found that the calculated aerodynamic damping was relatively insensitive to variation in natural frequency.
The National Transonic Facility: A Research Retrospective
NASA Technical Reports Server (NTRS)
Wahls, R. A.
2001-01-01
An overview of the National Transonic Facility (NTF) from a research utilization perspective is provided. The facility was born in the 1970s from an internationally recognized need for a high Reynolds number test capability based on previous experiences with preflight predictions of aerodynamic characteristics and an anticipated need in support of research and development for future aerospace vehicle systems. Selection of the cryogenic concept to meet the need, unique capabilities of the facility, and the eventual research utilization of the facility are discussed. The primary purpose of the paper is to expose the range of investigations that have used the NTF since being declared operational in late 1984; limited research results are included, though many more can be found in the references.
Analysis of three-dimensional transonic compressors
NASA Technical Reports Server (NTRS)
Bourgeade, A.
1984-01-01
A method for computing the three-dimensional transonic flow around the blades of a compressor or of a propeller is given. The method is based on the use of the velocity potential, on the hypothesis that the flow is inviscid, irrotational and isentropic. The equation of the potential is solved in a transformed space such that the surface of the blade is mapped into a plane where the periodicity is implicit. This equation is in a nonconservative form and is solved with the help of a finite difference method using artificial time. A computer code is provided and some sample results are given in order to demonstrate the influence of three-dimensional effects and the blade's rotation.
Transonic Blunt Body Aerodynamic Coefficients Computation
NASA Astrophysics Data System (ADS)
Sancho, Jorge; Vargas, M.; Gonzalez, Ezequiel; Rodriguez, Manuel
2011-05-01
In the framework of EXPERT (European Experimental Re-entry Test-bed) accurate transonic aerodynamic coefficients are of paramount importance for the correct trajectory assessment and parachute deployment. A combined CFD (Computational Fluid Dynamics) modelling and experimental campaign strategy was selected to obtain accurate coefficients. A preliminary set of coefficients were obtained by CFD Euler inviscid computation. Then experimental campaign was performed at DNW facilities at NLR. A profound review of the CFD modelling was done lighten up by WTT results, aimed to obtain reliable values of the coefficients in the future (specially the pitching moment). Study includes different turbulence modelling and mesh sensitivity analysis. Comparison with the WTT results is explored, and lessons learnt are collected.
Design of a transonically profiled wing
NASA Technical Reports Server (NTRS)
Kiekebusch, B.
1978-01-01
The application of well known design concepts with the combined use of thick transonic profiles to aircraft wing design was investigated. Optimization in terms of weight and operational costs was emphasized. It is shown that the usual design criteria and concepts are too restricted and do not sufficiently represent the physical processes over the wing. Suggestions are made for improving this situation, and a design example given. Compared with a wing design according to previously used criteria, the new design is found to be superior in the most important functions. It is concluded that an isobar concept adjusted to the planform in conjunction with an 'organically' designed wing will lead to the weight optimum solutions of wing profiles.
Transonic Flow Computations Using Nonlinear Potential Methods
NASA Technical Reports Server (NTRS)
Holst, Terry L.; Kwak, Dochan (Technical Monitor)
2000-01-01
This presentation describes the state of transonic flow simulation using nonlinear potential methods for external aerodynamic applications. The presentation begins with a review of the various potential equation forms (with emphasis on the full potential equation) and includes a discussion of pertinent mathematical characteristics and all derivation assumptions. Impact of the derivation assumptions on simulation accuracy, especially with respect to shock wave capture, is discussed. Key characteristics of all numerical algorithm types used for solving nonlinear potential equations, including steady, unsteady, space marching, and design methods, are described. Both spatial discretization and iteration scheme characteristics are examined. Numerical results for various aerodynamic applications are included throughout the presentation to highlight key discussion points. The presentation ends with concluding remarks and recommendations for future work. Overall. nonlinear potential solvers are efficient, highly developed and routinely used in the aerodynamic design environment for cruise conditions. Published by Elsevier Science Ltd. All rights reserved.
Demonstration of PIV in a Transonic Compressor
NASA Technical Reports Server (NTRS)
Wernet, Mark P.
1997-01-01
Particle Imaging Velocimetry (PIV) is a powerful measurement technique which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. PIV data are measured simultaneously at multiple points in space, which enables the investigation of the non-stationary spatial structures typically encountered in turbomachinery. Many of the same issues encountered in the application of LDV techniques to rotating machinery apply in the application of PIV. Preliminary results from the successful application of the standard 2-D PIV technique to a transonic axial compressor are presented. The lessons learned from the application of the 2-D PIV technique will serve as the basis for applying 3-component PIV techniques to turbomachinery.
A transonic rectangular grid embedded panel method
NASA Technical Reports Server (NTRS)
Johnson, F. T.; Bussoletti, J. E.; James, R. M.; Young, D. P.; Woo, A. C.
1982-01-01
A method is presented that has the potential for solving transonic flow problems about the same complex aircraft configurations currently being analyzed by subsonic panel methods. This method does not require the generation of surface fitted grids. Instead it uses rectangular grids and subgrids together with embedded surface panels on which boundary conditions are imposed. Both the Euler and full potential equations are considered. The method of least squares is used to reduce the solution of these equations to the solution of a sequence of Poisson problems. The Poisson problems are solved using fast Fourier transforms and panel influence coefficient techniques. The overall method is still in its infancy but some two dimensional results are shown illustrating various key features.
Nonclassical aileron buzz in transonic flow
NASA Technical Reports Server (NTRS)
Bendiksen, Oddvar O.
1993-01-01
A computational study of inviscid, transonic aileron and trailing-edge buzz instabilities is presented. A mixed Eulerian-Lagrangian formulation is used to model the fluid-structure system and to obtain a system of space-discretized equations that is time-marched to simulate the aeroelastic behavior of the wing-aileron system. Results obtained suggest that shock-induced separation may not be an essential driving force behind all buzz phenomena. Several examples are shown where the shock motion interacts with the aileron motion to extract energy from the flow. If the trailing-edge region is sufficiently flexible and the shocks are at the trailing edge, a trailing-edge buzz instability appears possible.
Transonic flow visualization using holographic interferometry
NASA Technical Reports Server (NTRS)
Bryanston-Cross, Peter J.
1987-01-01
An account is made of some of the applications of holographic interferometry to the visualization of transonic flows. In the case of the compressor shock visualization, the method is used regularly and has moved from being a research department invention to a design test tool. With the implementation of automatic processing and simple digitization systems, holographic vibrational analysis has also moved into routine nondestructive testing. The code verification interferograms were instructive, but the main turbomachinery interest is now in 3 dimensional flows. A major data interpretation effort will be required to compute tomographically the 3 dimensional flow around the leading or the trailing edges of a rotating blade row. The bolt on approach shows the potential application to current unsteady flows of interest. In particular that of the rotor passing and vortex interaction effects is experienced by the new generation of unducted fans. The turbocharger tests presents a new area for the application of holography.
NASA Technical Reports Server (NTRS)
Paryz, Roman W.
2014-01-01
Several upgrade projects have been completed at the NASA Langley Research Center National Transonic Facility over the last 1.5 years in an effort defined as STARBUKS - Subsonic Transonic Applied Refinements By Using Key Strategies. This multi-year effort was undertaken to improve NTF's overall capabilities by addressing Accuracy and Validation, Productivity, and Reliability areas at the NTF. This presentation will give a brief synopsis of each of these efforts.
Vitrification of an incinerator blowdown waste containing both chloride salts and carbon
Resce, J.L.; Ragsdale, R. G.; Overcamp, T.J.
1996-10-01
A low-level, mixed-waste simulant, derived from incinerator blowdown residue, has been vitrified in a series of crucible tests. Major components of this waste simulant included carbon, sodium chloride, silica, and alumina. Hazardous and surrogate radioactive metal complexes were also included. Two different formulations of additives were combined with the waste to facilitate glass formation. These glass formulations included NaNO{sub 3} to serve as a flux and to assist in the oxidation of the carbon. During vitrification, a chloride salt layer was observed on the melt surface which volatilized during the course of heating. Furthermore, significant concentrations of As, Cd, Cs, Ni, Pb, Sb, Se, Tl, and Zn had volatilized during vitrification. It is postulated that this was due, in part, to the formation of volatile metal chlorides. Upon quenching, small metal nuggets, containing Pb, Sb, Ag, and Ni separated from the glasses. This reduction of the metal oxides was caused by the presence of carbon in the simulant. The chemical durability was evaluated by both the Toxicity Characteristic Leaching Procedure (TCLP) and the Product Consistency Test (PCT). The TCLP results were below the limits for the toxicity characteristic for a hazardous waste and also met the Universal Treatment Standards for listed wastes such as F006. 11 refs., 5 tabs.
Recent activities on CCMHD power generation studies with Fuji-1 blow-down facility
Okuno, Y.; Okamura, T.; Yoshikawa, K.
1995-12-31
Recent activities on experimental studies with closed cycle MBD {open_quotes}Fuji-1{open_quotes} blow-down facility at Tokyo Institute of Technology are presented. A new disk generator ( Disk-F4 ) has been installed and a new seed injection system has been introduced from IVTAN (Institute of High Temperature in Russia) in 1994.The design concept of the new generator channel is focused mainly on the reliability of high power generation. The Mach number at the generator inlet and the thermal input are increased up to {approximately}2.8 and {approximately}3.0 MW, respectively. In the new seeding system, a melted seed material is pushed by a piston dozator, instead of gas-pressure-driven in the previous system. The controllability of seed fraction is markedly improved, and the large fluctuation as has been observed previously is diminished, although the spatial non-uniformity of seed fraction which could occur in the all previous experiments is observed. In the power generation experiments with the new components, the maximum power output of 502 kW and the enthalpy extraction ratio of 16.7% have been obtained. These values at the present stage are lower than the maximum values previously achieved in the facility. However, the reliable high power generation can be expected for the new generator.
Liao, Z; Gu, Z; Schulz, M C; Davis, J R; Baygents, J C; Farrell, J
2009-01-01
This research investigated the effectiveness of electrocoagulation using iron and aluminium electrodes for treating cooling tower blowdown (CTB) waters containing dissolved silica (Si(OH)(4)), Ca(2 + ) and Mg(2 + ). The removal of each target species was measured as a function of the coagulant dose in simulated CTB waters with initial pH values of 5, 7, and 9. Experiments were also performed to investigate the effect of antiscaling compounds and coagulation aids on hardness ion removal. Both iron and aluminum electrodes were effective at removing dissolved silica. For coagulant doses < or =3 mM, silica removal was a linear function of the coagulant dose, with 0.4 to 0.5 moles of silica removed per mole of iron or aluminium. Iron electrodes were only 30% as effective at removing Ca(2 + ) and Mg(2 + ) as compared to silica. There was no measurable removal of hardness ions by aluminium electrodes in the absence of organic additives. Phosphonate based antiscaling compounds were uniformly effective at increasing the removal of Ca(2 + ) and Mg(2 + ) by both iron and aluminium electrodes. Cationic and amphoteric polymers used as coagulation aids were also effective at increasing hardness ion removal. PMID:19901466
Soil respiration in pits and mounds following an experimental forest blowdown
Millikin, C.S.; Bowden, R.D.
1996-11-01
Extensive uprooting of trees by windthrow can create areas of severe soil disturbance in temperate forests. Specifically, uprooted trees leave shaded pits and mounds of exposed roots and mineral soil. To assess the contribution of pit and mound microenvironments to overall soil respiration in an experimental hurricane blowdown at the Harvard Forest Long-Term Ecological Research site (MA), summer CO{sub 2} effluxes were measured on pit, mound, and undisturbed microsites. Mean CO{sub 2} effluxes were 45.4, 80.1, and 99.0 mgC m{sup -2} h{sup -1} for pit, mound, and control microsites, respectively. Although soil respiration is lower in areas of disturbed soil than in undisturbed areas, the total efflux contribution (5.3%) form pits and mounds to the overall flux rate at the site was small. The area-weighted soil respiration estimate is 3.1% lower than the estimate obtained using flux measurements from control locations alone. Measurements taken from undisturbed plots represent a small but systematic overestimate of soil respiration across the site. 25 refs., 1 fig.
1. VIEW LOOKING SOUTHEAST AT EXTERIOR OF 8FOOT TRANSONIC PRESSURE ...
1. VIEW LOOKING SOUTHEAST AT EXTERIOR OF 8-FOOT TRANSONIC PRESSURE TUNNEL. NOTE EXPANSION RINGS. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
2. VIEW LOOKING EASTNORTHEAST AT EXTERIOR OF 8FOOT TRANSONIC PRESSURE ...
2. VIEW LOOKING EAST-NORTHEAST AT EXTERIOR OF 8-FOOT TRANSONIC PRESSURE TUNNEL (BUILDING 640). NOTE NACA LOGO OVER DOORWAY. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
5. VIEW LOOKING NORTH AT 8FOOT TRANSONIC PRESSURE TUNNEL PLENUM ...
5. VIEW LOOKING NORTH AT 8-FOOT TRANSONIC PRESSURE TUNNEL PLENUM FLOOR AREA. NOTE SCHLIEREN OPTICAL SYSTEM ON STRUCTURE AT RIGHT CENTER. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
NASA Technical Reports Server (NTRS)
Tanaka, K.; Hirose, H.
1986-01-01
The development of transonic aerodynamic computation methods and specific examples, as well as examples of three-dimensional transonic computation in design, are discussed. The case of the transonic transport and the case of the small transport are analyzed. Requirements for programs of the future are itemized.
Investigations for Supersonic Transports at Transonic and Supersonic Conditions
NASA Technical Reports Server (NTRS)
Rivers, S. Melissa B.; Owens, Lewis R.; Wahls, Richard A.
2007-01-01
Several computational studies were conducted as part of NASA s High Speed Research Program. Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-Speed Research program are given. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin-Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.
Transonic airfoil design for helicopter rotor applications
NASA Technical Reports Server (NTRS)
Hassan, Ahmed A.; Jackson, B.
1989-01-01
Despite the fact that the flow over a rotor blade is strongly influenced by locally three-dimensional and unsteady effects, practical experience has always demonstrated that substantial improvements in the aerodynamic performance can be gained by improving the steady two-dimensional charateristics of the airfoil(s) employed. The two phenomena known to have great impact on the overall rotor performance are: (1) retreating blade stall with the associated large pressure drag, and (2) compressibility effects on the advancing blade leading to shock formation and the associated wave drag and boundary-layer separation losses. It was concluded that: optimization routines are a powerful tool for finding solutions to multiple design point problems; the optimization process must be guided by the judicious choice of geometric and aerodynamic constraints; optimization routines should be appropriately coupled to viscous, not inviscid, transonic flow solvers; hybrid design procedures in conjunction with optimization routines represent the most efficient approach for rotor airfroil design; unsteady effects resulting in the delay of lift and moment stall should be modeled using simple empirical relations; and inflight optimization of aerodynamic loads (e.g., use of variable rate blowing, flaps, etc.) can satisfy any number of requirements at design and off-design conditions.
Subsonic-transonic stall flutter study
NASA Technical Reports Server (NTRS)
Stardter, H.
1979-01-01
The objective of the Subsonic/Transonic Stall Flutter Program was to obtain detailed measurements of both the steady and unsteady flow field surrounding a rotor and the mechanical state of the rotor while it was operating in both steady and flutter modes to provide a basis for future analysis and for development of theories describing the flutter phenomenon. The program revealed that while all blades flutter at the same frequency, they do not flutter at the same amplitude, and their interblade phase angles are not equal. Such a pattern represents the superposition of a number of rotating nodal diameter patterns, each characterized by a different amplitude and different phase indexing, but each rotating at a speed that results in the same flutter frequency as seen in the rotor system. Review of the steady pressure contours indicated that flutter may alter the blade passage pressure distribution. The unsteady pressure amplitude contour maps reveal regions of high unsteady pressure amplitudes near the leading edge, lower amplitudes near the trailing.
Applications of a transonic wing design method
NASA Technical Reports Server (NTRS)
Campbell, Richard L.; Smith, Leigh A.
1989-01-01
A method for designing wings and airfoils at transonic speeds using a predictor/corrector approach was developed. The procedure iterates between an aerodynamic code, which predicts the flow about a given geometry, and the design module, which compares the calculated and target pressure distributions and modifies the geometry using an algorithm that relates differences in pressure to a change in surface curvature. The modular nature of the design method makes it relatively simple to couple it to any analysis method. The iterative approach allows the design process and aerodynamic analysis to converge in parallel, significantly reducing the time required to reach a final design. Viscous and static aeroelastic effects can also be accounted for during the design or as a post-design correction. Results from several pilot design codes indicated that the method accurately reproduced pressure distributions as well as the coordinates of a given airfoil or wing by modifying an initial contour. The codes were applied to supercritical as well as conventional airfoils, forward- and aft-swept transport wings, and moderate-to-highly swept fighter wings. The design method was found to be robust and efficient, even for cases having fairly strong shocks.
Flow Control in a Transonic Diffuser
NASA Astrophysics Data System (ADS)
Gartner, Jeremy; Amitay, Michael
2014-11-01
In some airplanes such as fighter jets and UAV, short inlet ducts replace the more conventional ducts due to their shorter length. However, these ducts are associated with low length-to-diameter ratio and low aspect ratio and, thus, experience massive separation and the presence of secondary flow structures. These flow phenomena are undesirable as they lead to pressure losses and distortion at the Aerodynamic Interface Plane (AIP), where the engine face is located. It causes the engine to perform with a lower efficiency as it would with a straight duct diffuser. Different flow control techniques were studied on the short inlet duct, with the goal to reattach the flow and minimize the distortions at the AIP. Due to the complex interaction between the separation and the secondary flow structures, the necessity to understand the flow mechanisms, and how to control them at a more fundamental level, a new transonic diffuser with an upper ramp and a straight floor was designed and built. The objective of this project is to explore the effectiveness of different flow control techniques in a high subsonic (up to Mach 0.8) diffuser, so that the quasi two-dimensional separation and the formation of secondary flow structure can be isolated using a canonical flow field. Supported by Northrop Grumman.
Computation of Transonic Flows Using Potential Methods
NASA Technical Reports Server (NTRS)
Hoist, Terry L.; Kwak, Dochan (Technical Monitor)
1997-01-01
The proposed paper will describe the state of the art associated with numerical solution of the full or exact velocity potential equation for solving transonic, external-aerodynamic flows. The presentation will begin with a review of the literature emphasizing research activities of the past decade. Next, the various forms of the full or exact velocity potential equation, the equation's corresponding mathematical characteristics, and the derivation assumptions will be presented and described in detail. Impact of the derivation assumptions on simulation accuracy, especially with respect to shock wave capture, will be presented and discussed relative to the more complete Euler or Navier-Stokes formulations. The technical presentation will continue with a description of recently developed full potential numerical approach characteristics. This description will include governing equation nondimensionalization, physical-to-computational-domain mapping procedures, a limited description of grid generation requirements, the spatial discretization scheme, numerical implementation of boundary conditions, and the iteration scheme. The next portion of the presentation will present and discuss numerical results for several two- and three-dimensional aerodynamic applications. Included in the results section will be a discussion and demonstration of a typical grid refinement analysis for determining spatial convergence of the numerical solution and level of solution accuracy. Computer timings for a variety of full potential applications will be compared and contrasted with similar results for the Euler equation formulation. Finally. the presentation will end with concluding remarks and recommendations for future work.
NASA Technical Reports Server (NTRS)
Jones, G. W., Jr.; Unangst, J. R.
1963-01-01
An investigation of the flutter characteristics of a series of thin cantilever wings having taper ratios of 0.6 was conducted in the Langley transonic blowdown tunnel at Mach numbers between 0.76 and 1.42. The angle of sweepback was varied from 0 degrees to 60 degrees on wings of aspect ratio 4, and the aspect ratio was varied from 2.4 to 6.4 on wings with 45 degrees of sweepback. The results are presented as ratios between the experimental flutter speeds and the reference flutter speeds calculated on the basis of incompressible two-dimensional flow. These ratios, designated the flutter-speed ratios, are given as functions of Mach number for the various wings. The flutter-speed ratios were characterized, in most cases, by values near 1.0 at subsonic speeds with large increases in the speed ratios in the range of supersonic speeds investigated. Increasing the sweep effected increases in the flutter-speed ratios between 0 degrees and 30 degrees followed by progressive reductions of the speed ratios to nearly 1.0 as the sweep was increased from 30 degrees to 60 degrees. Reducing the aspect ratio from 6.4 to 2.4 resulted in progressively larger values of the flutter-speed ratios throughout the Mach number range investigated.
Nonlinear Green's function method for unsteady transonic flows
NASA Technical Reports Server (NTRS)
Tseng, K.; Morino, L.
1982-01-01
Advantages to employing Green's function in describing unsteady three-dimensional transonic flows are explored. The development of the function for application to linear subsonic and supersonic unsteady aerodynamics is reviewed. It is shown that unique solutions are possible for external flows, with all functional expressions being defined in Prandtl-Glauert space. The development of methods of using the Green's function for transonic flows is traced, noting the necessity of including the effects of significant nonlinear terms. The steady-state problem is considered to demonstrate the shock-capturing ability of the method and the usefulness of the function in the incompressible, subsonic, transonic, and supersonic areas of potential unsteady three-dimensional flows around complex configurations. Computational time is asserted to be an order of magnitude less than with finite difference methods.
Method to predict external store carriage characteristics at transonic speeds
NASA Technical Reports Server (NTRS)
Rosen, Bruce S.
1988-01-01
Development of a computational method for prediction of external store carriage characteristics at transonic speeds is described. The geometric flexibility required for treatment of pylon-mounted stores is achieved by computing finite difference solutions on a five-level embedded grid arrangement. A completely automated grid generation procedure facilitates applications. Store modeling capability consists of bodies of revolution with multiple fore and aft fins. A body-conforming grid improves the accuracy of the computed store body flow field. A nonlinear relaxation scheme developed specifically for modified transonic small disturbance flow equations enhances the method's numerical stability and accuracy. As a result, treatment of lower aspect ratio, more highly swept and tapered wings is possible. A limited supersonic freestream capability is also provided. Pressure, load distribution, and force/moment correlations show good agreement with experimental data for several test cases. A detailed computer program description for the Transonic Store Carriage Loads Prediction (TSCLP) Code is included.
Transonic flow control by means of local energy deposition
NASA Astrophysics Data System (ADS)
Aul'Chenko, S. M.; Zamuraev, V. P.; Kalinina, A. P.
2011-11-01
Experimental data for the feasibility of transonic flow control by means of energy deposition are generalized. Energy supplied to the immediate vicinity of a body in stream before a compression shock is found to result in the nonlinear interaction of introduced disturbances with the shock and the surface in zones extended along the surface. A new, explosive gasdynamic mechanism behind the shift of the compression shock is discovered. It is shown that the nonlinear character of the interaction may considerably decrease the wave resistance of, e.g., transonic airfoils. It is found that energy supply from without stabilizes a transonic flow about an airfoil—the effect similar to the Khristianovich stabilization effect. The dependence of the energy deposition optimal frequency on the energy source parameters and Mach number of the incoming flow at which the resistance drops to the greatest extent is obtained. The influence of the real thermodynamic properties and viscosity of air is studied.
Solution of steady and unsteady transonic-vortex flows using Euler and full-potential equations
NASA Technical Reports Server (NTRS)
Kandil, Osama A.; Chuang, Andrew H.; Hu, Hong
1989-01-01
Two methods are presented for inviscid transonic flows: unsteady Euler equations in a rotating frame of reference for transonic-vortex flows and integral solution of full-potential equation with and without embedded Euler domains for transonic airfoil flows. The computational results covered: steady and unsteady conical vortex flows; 3-D steady transonic vortex flow; and transonic airfoil flows. The results are in good agreement with other computational results and experimental data. The rotating frame of reference solution is potentially efficient as compared with the space fixed reference formulation with dynamic gridding. The integral equation solution with embedded Euler domain is computationally efficient and as accurate as the Euler equations.
Spindler, B.; Pellissier, M. )
1990-01-01
Independent assessment of the TRAC code was conducted at the Centre d'Etudes Nucleaires de Grenoble of the Commissariate a l'Energie Atomique (France) in the frame of the ICAP. This report presents the results of the assessment of TRAC-PF1/MOD1 version 14.3 using critical flow steady state tests (MOBY-DICK, SUPER-MOBY-DICK), and blowdown tests (CANON, SUPER-CANON, VERTICAL-CANON, MARVIKEN, OMEGA-TUBE, OMEGA-BUNDLE). This document, Volume 1, presents the text and tables from this assessment.
Spindler, B.; Pellissier, M. )
1990-01-01
Independent assessment of the TRAC code was conducted at the Centre d'Etudes Nucleaires de Grenoble of the Commissariate a l'Energie Atomique (France) in the frame of the ICAP. This report presents the results of the assessment of TRAC-PF1/MOD1 version 14.3 using critical flow steady state tests (MOBY-DICK, SUPER-MOBY-DICK), and blowdown tests (CANON, SUPER-CANON, VERTICAL-CANON, MARVIKEN, OMEGA-TUBE, OMEGA-BUNDLE). This document, Volume 2, presents the experimental data and figures from the assessment.
Users Guide for the National Transonic Facility Research Data System
NASA Technical Reports Server (NTRS)
Foster, Jean M.; Adcock, Jerry B.
1996-01-01
The National Transonic Facility is a complex cryogenic wind tunnel facility. This report briefly describes the facility, the data systems, and the instrumentation used to acquire research data. The computational methods and equations are discussed in detail and many references are listed for those who need additional technical information. This report is intended to be a user's guide, not a programmer's guide; therefore, the data reduction code itself is not documented. The purpose of this report is to assist personnel involved in conducting a test in the National Transonic Facility.
Contributions of Transonic Dynamics Tunnel Testing to Airplane Flutter Clearance
NASA Technical Reports Server (NTRS)
Rivera, Jose A.; Florance, James R.
2000-01-01
The Transonic Dynamics Tunnel (TDT) became in operational in 1960, and since that time has achieved the status of the world's premier wind tunnel for testing large in aeroelastically scaled models at transonic speeds. The facility has many features that contribute to its uniqueness for aeroelastic testing. This paper will briefly describe these capabilities and features, and their relevance to aeroelastic testing. Contributions to specific airplane configurations and highlights from the flutter tests performed in the TDT aimed at investigating the aeroelastic characteristics of these configurations are presented.
Numerical studies of transverse curvature effects on transonic flow stability
NASA Technical Reports Server (NTRS)
Macaraeg, M. G.; Daudpota, Q. I.
1992-01-01
A numerical study of transverse curvature effects on compressible flow temporal stability for transonic to low supersonic Mach numbers is presented for axisymmetric modes. The mean flows studied include a similar boundary-layer profile and a nonsimilar axisymmetric boundary-layer solution. The effect of neglecting curvature in the mean flow produces only small quantitative changes in the disturbance growth rate. For transonic Mach numbers (1-1.4) and aerodynamically relevant Reynolds numbers (5000-10,000 based on displacement thickness), the maximum growth rate is found to increase with curvature - the maximum occurring at a nondimensional radius (based on displacement thickness) between 30 and 100.
Implicit, nonswitching, vector-oriented algorithm for steady transonic flow
NASA Technical Reports Server (NTRS)
Lottati, I.
1983-01-01
A rapid computation of a sequence of transonic flow solutions has to be performed in many areas of aerodynamic technology. The employment of low-cost vector array processors makes the conduction of such calculations economically feasible. However, for a full utilization of the new hardware, the developed algorithms must take advantage of the special characteristics of the vector array processor. The present investigation has the objective to develop an efficient algorithm for solving transonic flow problems governed by mixed partial differential equations on an array processor.
National Transonic Facility: A review of the operational plan
NASA Technical Reports Server (NTRS)
Liepmann, H. W.; Black, R. E.; Dietz, R. O.; Kirchner, M. E.; Sears, W. R.
1980-01-01
The proposed National Transonic Facility (NTF) operational plan is reviewed. The NTF will provide an aerodynamic test capability significantly exceeding that of other transonic regime wind tunnels now available. A limited number of academic research program that might use the NTF are suggested. It is concluded that the NTF operational plan is useful for management, technical, instrumentation, and model building techniques available in the specialized field of aerodynamic analysis and simulation. It is also suggested that NASA hold an annual conference to discuss wind tunnel research results and to report on developments that will further improve the utilization and cost effectiveness of the NTF and other wind tunnels.
Slender body theory and Space Shuttle transonic aerodynamics
NASA Technical Reports Server (NTRS)
Malmuth, N. D.; Wu, C. C.; Cole, J. D.
1985-01-01
A computational implementation of transonic slender body theory and the equivalence rule has been utilized to study transonic flow field around the Space Shuttle Orbiter. The far field is described by a nonlinear axisymmetric Karman-Guderley model and the near field by a cross flow Laplace equation boundary value problem. The latter is treated using a source panel method. Preliminary comparisons with experiments give encouraging indications that the model can be useful for quick turnaround estimates. Areas of refinement to obtain more accurate predictions are discussed.
Recent Enhancements to the National Transonic Facility (Mixed Mode Operations)
NASA Technical Reports Server (NTRS)
Kilgore, W. Allen; Chan, David; Balakrishna, S.; Wahls, Richard A.
2006-01-01
The U.S. National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the development of a Mixed-mode of operations that combine Air-mode operations with Nitrogen-mode operations. This implementation and operational results of this new Mixed-mode expands the ambient temperature transonic region of testing beyond the Air-mode limitations at a significantly reduced cost over Nitrogen Mode operation.
Issac, Jason Cherian ses in transonic flow
NASA Technical Reports Server (NTRS)
Issac, Jason Cherion; Kapania, Rakesh K.
1993-01-01
Flutter analysis of a two degree of freedom airfoil in compressible flow is performed using a state-space representation of the unsteady aerodynamic behavior. Indicial response functions are used to represent the normal force and moment response of the airfoil. The structural equations of motion of the airfoil with bending and torsional degrees of freedom are coupled to the unsteady air loads and the aeroelastic system so modelled is solved as an eigenvalue problem to determine the stability. The aeroelastic equations are also directly integrated with respect to time and the time-domain results compared with the results from the eigenanalysis. A good agreement is obtained. The derivatives of the flutter speed obtained from the eigenanalysis are calculated with respect to the mass and stiffness parameters by both analytical and finite-difference methods for various transonic Mach numbers. The experience gained from the two degree of freedom model is applied to study the sensitivity of the flutter response of a wing with respect to various shape parameters. The parameters being considered are as follows: (1) aspect ratio; (2) surface area of the wing; (3) taper ratio; and (4) sweep. The wing deflections are represented by Chebyshev polynomials. The compressible aerodynamic state-space model used for the airfoil section is extended to represent the unsteady aerodynamic forces on a generally laminated tapered skewed wing. The aeroelastic equations are solved as an eigenvalue problem to determine the flutter speed of the wing. The derivatives of the flutter speed with respect to the shape parameters are calculated by both analytical and finite difference methods.
Haynes, H.D.
1988-01-01
In support of the NRC-funded Nuclear Plant Aging Research (NPAR) program, Oak Ridge National Laboratory (ORNL) has carried out a comprehensive aging assessment of Motor-Operated Valves (MOVs). As part of this work, ORNL participated in the Gate Valve Flow Interruption Blowdown (GVFIB) tests carried out in Huntsville, Alabama. The GVFIB tests were intended primarily to determine the behavior of motor-operated gate valves under the temperature, pressure, and flow conditions expected to be experienced by isolation valves in Boiling Water Reactors (BWRs) during a high energy line break (blowdown) outside of containment. In addition, the tests provided an excellent opportunity to evaluate signature analysis methods for determining the operational readiness of the MOVs under those accident conditions. ORNL acquired motor current and torque switch shaft angular position data on two test MOVs during various times of the GVFIB tests. The reduction in operating ''margin'' of both MOVs due to the presence of additional valve running loads imposed by high flow was clearly observed in motor current and torque switch angular position signatures. In addition, the effects of differential pressure, fluid temperature, and line voltage on MOV operations were observed and more clearly understood as a result of utilizing signature analysis techniques. 1 ref.; 16 figs.
ACT Missile Model In Langley 16 Foot Transonic Tunnel
NASA Technical Reports Server (NTRS)
1994-01-01
The photograph shows a 15-percent scale model of the ACT advanced missile concept in the Langley 16-Foot Transonic Tunnel. The model featured independently controlled reaction jets near the nose and the tail of the model. Aerodynamic control was provided by four fins that were located near the tail.
Computed Aeroelastic Motions Of Wings In Transonic Flows
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; Obayashi, Shigeru
1995-01-01
Report describes computational simulations of aeroelastic motions of delta and swept wings in transonic flows. Study directed toward understanding aerodynamic behavior and enhancing maneuverability of fighter airplanes equipped with such wings. Also has implications for gas pumps and turbines, in which flows near tips of vanes and blades reach supersonic speeds.
Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix
NASA Technical Reports Server (NTRS)
Li, Wesley Waisang; Pak, Chan-Gi
2010-01-01
A technique for approximating the modal aerodynamic influence coefficients [AIC] matrices by using basis functions has been developed and validated. An application of the resulting approximated modal AIC matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO(TradeMark) flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing [ATW] 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle
Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix
NASA Technical Reports Server (NTRS)
Li, Wesley W.; Pak, Chan-gi
2011-01-01
A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.
Unsteady transonic algorithm improvements for realistic aircraft applications
NASA Technical Reports Server (NTRS)
Batina, John T.
1987-01-01
Improvements to a time-accurate approximate factorization (AF) algorithm were implemented for steady and unsteady transonic analysis of realistic aircraft configurations. These algorithm improvements were made to the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code developed at the Langley Research Center. The code permits the aeroelastic analysis of complete aircraft in the flutter critical transonic speed range. The AF algorithm of the CAP-TSD code solves the unsteady transonic small-disturbance equation. The algorithm improvements include: an Engquist-Osher (E-O) type-dependent switch to more accurately and efficiently treat regions of supersonic flow; extension of the E-O switch for second-order spatial accuracy in these regions; nonreflecting far field boundary conditions for more accurate unsteady applications; and several modifications which accelerate convergence to steady-state. Calculations are presented for several configurations including the General Dynamics one-ninth scale F-16C aircraft model to evaluate the algorithm modifications. The modifications have significantly improved the stability of the AF algorithm and hence the reliability of the CAP-TSD code in general.
Intermittent Flow Regimes in a Transonic Fan Airfoil Cascade
NASA Technical Reports Server (NTRS)
Lepicovsky, J.; McFarland, E. R.; Chima, R. V.; Capece, V. R.; Hayden, J.
2002-01-01
A study was conducted in the NASA Glenn Research Center linear cascade on the intermittent flow on the suction surface of an airfoil section from the tip region of a modern low aspect ratio fan blade. Experimental results revealed that, at a large incidence angle, a range of transonic inlet Mach numbers exist where the leading-edge shock-wave pattern was unstable. Flush mounted high frequency response pressure transducers indicated large local jumps in the pressure in the leading edge area, which generates large intermittent loading on the blade leading edge. These measurements suggest that for an inlet Mach number between 0.9 and 1.0 the flow is bi-stable, randomly switching between subsonic and supersonic flows. Hence, it appears that the change in overall flow conditions in the transonic region is based on the frequency of switching between two stable flow states rather than on the continuous increase of the flow velocity. To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the byproduct of an endwall restricted linear cascade.
Transonic Shock Problem for the Euler System in a Nozzle
NASA Astrophysics Data System (ADS)
Xin, Zhouping; Yan, Wei; Yin, Huicheng
2009-10-01
In this paper, we study the well-posedness problem on transonic shocks for steady ideal compressible flows through a two-dimensional slowly varying nozzle with an appropriately given pressure at the exit of the nozzle. This is motivated by the following transonic phenomena in a de Laval nozzle. Given an appropriately large receiver pressure P r , if the upstream flow remains supersonic behind the throat of the nozzle, then at a certain place in the diverging part of the nozzle, a shock front intervenes and the flow is compressed and slowed down to subsonic speed, and the position and the strength of the shock front are automatically adjusted so that the end pressure at exit becomes P r , as clearly stated by Courant and Friedrichs [Supersonic flow and shock waves, Interscience Publishers, New York, 1948 (see section 143 and 147)]. The transonic shock front is a free boundary dividing two regions of C 2,α flow in the nozzle. The full Euler system is hyperbolic upstream where the flow is supersonic, and coupled hyperbolic-elliptic in the downstream region Ω+ of the nozzle where the flow is subsonic. Based on Bernoulli’s law, we can reformulate the problem by decomposing the 3 × 3 Euler system into a weakly coupled second order elliptic equation for the density ρ with mixed boundary conditions, a 2 × 2 first order system on u 2 with a value given at a point, and an algebraic equation on ( ρ, u 1, u 2) along a streamline. In terms of this reformulation, we can show the uniqueness of such a transonic shock solution if it exists and the shock front goes through a fixed point. Furthermore, we prove that there is no such transonic shock solution for a class of nozzles with some large pressure given at the exit.
Space Shuttle flutter as affected by wing-body aerodynamic interaction
NASA Technical Reports Server (NTRS)
Chipman, R. R.; Rauch, F. J.; Shyprykevich, P.; Hess, R. W.
1974-01-01
In the NASA Langley Research Center 26-inch transonic blowdown wind-tunnel, flutter speeds were measured on 1/80-th scale semispan models of the orbiter wing, the complete Space Shuttle, and intermediate component combinations. Using the doublet lattice method combined with slender body theory to calculate unsteady aerodynamic forces, subsonic flutter speeds were computed for comparison. Aerodynamic interaction was found by test and analysis to raise the flutter speed in some configurations while lowering it in others. Although at Mach number less than 0.7, predicted speeds correlated to within 6% of those measured, rapid deterioration of the agreement occurred at higher subsonic Mach numbers, especially on the more complicated configurations. Additional analysis showed that aerodynamic forces arising from body flexibility potentially can have a large effect on flutter speed, but that the current shuttle design is not so affected.
NASA Technical Reports Server (NTRS)
Chipman, R. R.; Rauch, F. J.
1975-01-01
The effects on flutter of the aerodynamic interaction between the space shuttle bodies and wing, 1/80th-scale semispan models of the orbiter wing, the complete shuttle and intermediate component combinations were tested in the NASA Langley Research Center 26-inch Transonic Blowdown Wind Tunnel. Using the double lattice method combined with slender body theory to calculate unsteady aerodynamic forces, subsonic flutter speeds were computed for comparison. Using calculated complete vehicle modes, flutter speed trends were computed for the full scale vehicle at an altitude of 15,200 meters and a Mach number of 0.6. Consistent with findings of the model studies, analysis shows the shuttle to have the same flutter speed as an isolated cantilevered wing.
Shelton, M.L.; Gregory, B.A. ); Doughty, R.L.; Kiss, T.; Moses, H.L. . Mechanical Engineering Dept.)
1993-07-01
In aircraft engine design (and in other applications), small improvements in turbine efficiency may be significant. Since analytical tools for predicting transonic turbine losses are still being developed, experimental efforts are required to evaluate various designs, calibrate design methods, and validate CFD analysis tools. However, these experimental efforts must be very accurate to measure the performance differences to the levels required by the highly competitive aircraft engine market. Due to the sensitivity of transonic and supersonic flow fields, it is often difficult to obtain the desired level of accuracy. In this paper, a statistical approach is applied to the experimental evaluation of transonic turbine airfoils in the VPI and SU transonic cascade facility in order to quantify the differences between three different transonic turbine airfoils. This study determines whether the measured performance differences between the three different airfoils are statistically significant. This study also assesses the degree of confidence in the transonic cascade testing process at VPI and SU.
Report of the panel on dynamics and aeroelasticity. [transonic tunnel capabilities
NASA Technical Reports Server (NTRS)
Houbolt, J.
1977-01-01
Model scaling for flutter analysis is reviewed. Characteristics of the Langley Transonic Dynamics Tunnel (TDT) are described and several features are recommended for inclusion in the National Transonic Facility. Problem areas suggested for the NTF include: Reynolds number effects on control surface unsteady aerodynamics; effects of Reynolds number on buffet onset and loads; transonic unsteady aerodynamics; and Reynolds number effects on flutter characteristics of wing planforms and airfoils.
NASA Technical Reports Server (NTRS)
Ballhaus, W. F.; Jameson, A.; Albert, J.
1977-01-01
Implicit approximate-factorization algorithms (AF) are developed for the solution of steady-state transonic flow problems. The performance of the AF solution method is evaluated relative to that of the standard solution method for transonic flow problems, successive line over-relaxation (SLOR). Both methods are applied to the solution of the nonlinear, two-dimensional transonic small-disturbance equation. Results indicate that the AF method requires substantially less computer time than SLOR to solve the nonlinear finite-difference matrix equation for a transonic flow field. This increase in computational efficiency is achieved with no appreciable increase in computer storage or coding complexity.
Emerging technology for transonic wind-tunnel-wall interference assessment and corrections
NASA Technical Reports Server (NTRS)
Newman, P. A.; Kemp, W. B., Jr.; Garriz, J. A.
1988-01-01
Several nonlinear transonic codes and a panel method code for wind tunnel/wall interference assessment and correction (WIAC) studies are reviewed. Contrasts between two- and three-dimensional transonic testing factors which affect WIAC procedures are illustrated with airfoil data from the NASA/Langley 0.3-meter transonic cyrogenic tunnel and Pathfinder I data. Also, three-dimensional transonic WIAC results for Mach number and angle-of-attack corrections to data from a relatively large 20 deg swept semispan wing in the solid wall NASA/Ames high Reynolds number Channel I are verified by three-dimensional thin-layer Navier-Stokes free-air solutions.
Chambers, Jeffrey Q; Robertson, Amanda L; Carneiro, Vilany M C; Lima, Adriano J N; Smith, Marie-Louise; Plourde, Lucie C; Higuchi, Niro
2009-05-01
Advanced recruitment and neutral processes play important roles in determining tree species composition in tropical forest canopy gaps, with few gaps experiencing clear secondary successional processes. However, most studies are limited to the relatively limited spatial scales provided by forest inventory plots, and investigations over the entire range of gap size are needed to better understand how ecological processes vary with tree mortality events. This study employed a landscape approach to test the hypothesis that tree species composition and forest structural attributes differ between large blowdown gaps and relatively undisturbed primary forest. Spectral mixture analysis on hyperspectral satellite imagery was employed to direct field sampling to widely distributed sites, and blowdown plots were compared with undisturbed primary forest plots. Tree species composition and forest structural attributes differed markedly between gap and non-gap sites, providing evidence of niche partitioning in response to disturbance across the region. Large gaps were dominated by classic Neotropical pioneer genera such as Cecropia and Vismia, and average tree size was significantly smaller. Mean wood density of trees recovering in large gaps (0.55 g cm(-3)) was significantly lower than in primary forest plots (0.71 g cm(-3)), a difference similar to that found when comparing less dynamic (i.e., tree recruitment, growth, and mortality) Central Amazon forests with more dynamic Western Amazon forests. Based on results, we hypothesize that the importance of neutral processes weaken, and niche processes strengthen, in determining community assembly along a gradient in gap size and tree mortality intensity. Over evolutionary time scales, pervasive dispersal among colonizers could result in the loss of tree diversity in the pioneer guild through competitive exclusion. Results also underscore the importance of considering disturbance processes across the landscape when addressing
A linearized Euler analysis of unsteady transonic flows in turbomachinery
Hall, K.C.; Clark, W.S.; Lorence, C.B. . Dept. of Mechanical Engineering and Materials Science)
1994-07-01
A computational method for efficiently predicting unsteady transonic flows in two- and three-dimensional cascades is presented. The unsteady flow is modeled using a linearized Euler analysis whereby the unsteady flow field is decomposed into a nonlinear mean flow plus a linear harmonically varying unsteady flow. The equations that govern the perturbation flow, the linearized Euler equations, are linear variable coefficient equations. For transonic flows containing shocks, shock capturing is used to model the shock impulse (the unsteady load due to the harmonic motion of the shock). A conservative Lax-Wendroff scheme is used to obtain a set of linearized finite volume equations that describe the harmonic small disturbance behavior of the flow. Conditions under which such a discretization will correctly predict the shock impulse are investigated. Computational results are presented that demonstrate the accuracy and efficiency of the present method as well as the essential role of unsteady shock impulse loads on the flutter stability of fans.
Reynolds Number Effects on a Supersonic Transport at Transonic Conditions
NASA Technical Reports Server (NTRS)
Wahls, R. N.; Owens, L. R.; Rivers, S. M. B.
2001-01-01
A High Speed Civil Transport configuration was tested in the National Transonic Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. The primary purposes of the tests were to assess Reynolds number scale effects and the high Reynolds number aerodynamic characteristics of a realistic, second generation supersonic transport while providing data for the assessment of computational methods. The tests included longitudinal and lateral/directional studies at low speed high-lift and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to near flight conditions. Results are presented which focus on both the Reynolds number and static aeroelastic sensitivities of longitudinal characteristics at Mach 0.90 for a configuration without an empennage.
A finite-difference method for transonic airfoil design.
NASA Technical Reports Server (NTRS)
Steger, J. L.; Klineberg, J. M.
1972-01-01
This paper describes an inverse method for designing transonic airfoil sections or for modifying existing profiles. Mixed finite-difference procedures are applied to the equations of transonic small disturbance theory to determine the airfoil shape corresponding to a given surface pressure distribution. The equations are solved for the velocity components in the physical domain and flows with embedded shock waves can be calculated. To facilitate airfoil design, the method allows alternating between inverse and direct calculations to obtain a profile shape that satisfies given geometric constraints. Examples are shown of the application of the technique to improve the performance of several lifting airfoil sections. The extension of the method to three dimensions for designing supercritical wings is also indicated.
National Transonic Facility model and model support vibration problems
NASA Technical Reports Server (NTRS)
Young, Clarence P., Jr.; Popernack, Thomas G., Jr.; Gloss, Blair B.
1990-01-01
Vibrations of models and model support system were encountered during testing in the National Transonic Facility. Model support system yaw plane vibrations have resulted in model strain gage balance design load limits being reached. These high levels of vibrations resulted in limited aerodynamic testing for several wind tunnel models. The yaw vibration problem was the subject of an intensive experimental and analytical investigation which identified the primary source of the yaw excitation and resulted in attenuation of the yaw oscillations to acceptable levels. This paper presents the principal results of analyses and experimental investigation of the yaw plane vibration problems. Also, an overview of plans for development and installation of a permanent model system dynamic and aeroelastic response measurement and monitoring system for the National Transonic Facility is presented.
Three-dimensional shock structure in a transonic flutter cascade
NASA Technical Reports Server (NTRS)
Boldman, D. R.; Buggele, A. E.; Decker, A. J.
1982-01-01
Rapid double-pulse holography was employed to obtain detailed, two-dimensional images of the shock forming during simulated flutter in a transonic flowfield. The experiment comprised a linear cascade of airfoils externally oscillated in torsion and viewed tangentially at the shock surface. Three biconvex airfoils were subjected to harmonic pitching motion about the midchord axis at a frequency of 0.53 while immersed in a Mach 0.81 flow. Failure to produce observable shocks led to use of choked flow with a Mach number near one, of which 50 holograms were taken. The images revealed a narrow shock surface with a spanwise variation in the shock properties. The method is concluded to be useful for examining transonic flowfield shocks in the presence of airfoil flutter.
Assessment of the National Transonic Facility for Laminar Flow Testing
NASA Technical Reports Server (NTRS)
Crouch, Jeffrey D.; Sutanto, Mary I.; Witkowski, David P.; Watkins, A. Neal; Rivers, Melissa B.; Campbell, Richard L.
2010-01-01
A transonic wing, designed to accentuate key transition physics, is tested at cryogenic conditions at the National Transonic Facility at NASA Langley. The collaborative test between Boeing and NASA is aimed at assessing the facility for high-Reynolds number testing of configurations with significant regions of laminar flow. The test shows a unit Reynolds number upper limit of 26 M/ft for achieving natural transition. At higher Reynolds numbers turbulent wedges emanating from the leading edge bypass the natural transition process and destroy the laminar flow. At lower Reynolds numbers, the transition location is well correlated with the Tollmien-Schlichting-wave N-factor. The low-Reynolds number results suggest that the flow quality is acceptable for laminar flow testing if the loss of laminar flow due to bypass transition can be avoided.
The transonic multi-foil Augmentor-Wing
NASA Technical Reports Server (NTRS)
Farbridge, J. E.; Smith, R. C.
1977-01-01
The paper describes the development of a transonic blown multi-foil Augmentor-Wing airfoil section that has a thickness/chord (t/c) value of 0.18. In comparison with an unblown single-foil supercritical section of the same overall t/c the new multi-foil section is characterized by an increased drag rise Mach number, increased buffet boundaries, and a reduction in 'effective' drag due to blowing. Potential advantages of the Augmentor-Wing are considered and the testing of three high-speed models in a trisonic pressurized wind tunnel (possessing a two-dimensional transonic insert) is discussed. The data indicate that a very thick wing is feasible since separations toward the rear of the main foil can be controlled both by shroud location and augmentor blowing.
Experimental transonic flutter characteristics of supersonic cruise configurations
NASA Technical Reports Server (NTRS)
Durham, Michael H.; Cole, Stanley R.; Cazier, F. W., Jr.; Keller, Donald F.; Parker, Ellen C.; Wilkie, W. Keats
1990-01-01
The flutter characteristics of a generic arrow-wing supersonic transport configuration are studied. The wing configuration has a 3 percent biconvex airfoil and a leading-edge sweep of 73 deg out to a cranked tip with a 60 deg leading-edge sweep. The ground vibration tests and flutter test procedure are described. The effects of flutter on engine nacelles, fuel loading, wing-mounted vertical fin, wing angle-of-attack, and wing tip mass and stiffness distributions are analyzed. The data reveal that engine nacelles reduce the transonic flutter dynamic pressure by 25-30 percent; fuel loadings decrease dynamic pressures by 25 percent; 4-6 deg wing angles-of-attack cause steep transonic boundaries; and 5-10 percent changes in flutter dynamic pressures are the result of the wing-mounted vertical fin and wing-tip mass and stiffness distributions.
Geared-elevator flutter study. [transonic flutter characteristics of empennage
NASA Technical Reports Server (NTRS)
Ruhlin, C. L.; Doggett, R. V., Jr.; Gregory, R. A.
1976-01-01
The paper describes an experimental and analytical study of the transonic flutter characteristics of an empennage flutter model having an all-movable horizontal tail with a geared elevator. Two configurations were flutter tested: one with a geared elevator and one with a locked elevator with the model cantilever-mounted on a sting in the wind tunnel. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter boundary was nearly flat at transonic speeds for both configurations. It was found that an analysis which treated the elevator as a discrete surface predicted flutter dynamic pressure levels better than analyses which treated the stabilizer and elevator as a warped surface. Warped-surface methods, however, predicted more closely the experimental flutter frequencies and Mach number trends.
Recent developments in finite element analysis for transonic airfoils
NASA Technical Reports Server (NTRS)
Hafez, M. M.; Murman, E. M.
1979-01-01
The prediction of aerodynamic forces in the transonic regime generally requires a flow field calculation to solve the governing non-linear mixed elliptic-hyperbolic partial differential equations. Finite difference techniques were developed to the point that design and analysis application are routine, and continual improvements are being made by various research groups. The principal limitation in extending finite difference methods to complex three-dimensional geometries is the construction of a suitable mesh system. Finite element techniques are attractive since their application to other problems have permitted irregular mesh elements to be employed. The purpose of this paper is to review the recent developments in the application of finite element methods to transonic flow problems and to report some recent results.
Transonic wind-tunnel tests of a lifting parachute model
NASA Technical Reports Server (NTRS)
Foughner, J. T., Jr.; Reed, J. F.; Wynne, E. C.
1976-01-01
Wind-tunnel tests have been made in the Langley transonic dynamics tunnel on a 0.25-scale model of Sandia Laboratories' 3.96-meter (13-foot), slanted ribbon design, lifting parachute. The lifting parachute is the first stage of a proposed two-stage payload delivery system. The lifting parachute model was attached to a forebody representing the payload. The forebody was designed and installed in the test section in a manner which allowed rotational freedom about the pitch and yaw axes. Values of parachute axial force coefficient, rolling moment coefficient, and payload trim angles in pitch and yaw are presented through the transonic speed range. Data are presented for the parachute in both the reefed and full open conditions. Time history records of lifting parachute deployment and disreefing tests are included.
A tomographic technique for aerodynamics at transonic speeds
NASA Technical Reports Server (NTRS)
Lee, G.
1985-01-01
Computer aided tomography (CAT) provides a means of noninvasively measuring the air density distribution around an aerodynamic model. This technique is global in that a large portion of the flow field can be measured. A test of the applicability of CAT to transonic velocities was studied. A hemispherical-nose cylinder afterbody model was tested at a Mach number of 0.8 with a new laser holographic interferometer at the 2- by 2-Foot Transonic Wind Tunnel. Holograms of the flow field were taken and were reconstructed into interferograms. The fringe distribution (a measure of the local densities) was digitized for subsequent data reduction. A computer program based on the Fourier-transform technique was developed to convert the fringe distribution into three-dimensional densities around the model. Theoretical aerodynamic densities were calculated for evaluating and assessing the accuracy of the data obtained from the tomographic method.
Initial Assesment of Space Launch System Transonic Unsteady Pressure Environment
NASA Technical Reports Server (NTRS)
Sekula, Martin K.; Piatak, David J.; Rausch, Russ D.; Florance, James R.; Ramey, James M.
2015-01-01
A series of wind tunnel tests were conducted at the NASA Langley Research Center Transonic Dynamics Tunnel to assess the transonic buffet environment for the Space Launch System (SLS) launch vehicle. An initial test, conducted in 2012, indicated an elevated buffet environment prompting a second test to provide further insight into the buffet phenomena and assess potential solutions to reduce the response levels of these environments. During the course of the test program, eight variants of the SLS-10000 configuration were examined. The effect of these configuration variants on the coefficient of the root-mean-square fluctuation of pressure about the mean as a function of test condition indicates that the maximum fluctuating pressure levels are extremely sensitive to the geometry of the forward attachment of the solid rocket boosters (SRBs) to the SLS Core. The addition of flow fences or changes to the SRB nose cone geometry can alleviate the unsteady pressure environment.
Shockless design and analysis of transonic blade shapes
NASA Technical Reports Server (NTRS)
Dulikravich, D. S.; Sobieczky, H.
1981-01-01
A fast computer program was developed to eliminate the shocks by slightly altering portions of the contour of a given airfoil in the cascade. The program can be used in two basic modes: (1) An analysis for steady, transonic, potential flow through a given planar cascade of airfoils and (2) a design for converting a given cascade into a shockless transonic cascade. The design mode can automatically be followed by the analysis mode, which confirms that the flow field is shock free. The program generates its own multilevel boundary conforming computational grids and solves a full potential equation in a fully conservative form. The shockless design is performed by implementing Sobieczky's fictitious-gas elliptic continuation concept.
Prediction of unsteady transonic flow around missile configurations
NASA Technical Reports Server (NTRS)
Nixon, D.; Reisenthel, P. H.; Torres, T. O.; Klopfer, G. H.
1990-01-01
This paper describes the preliminary development of a method for predicting the unsteady transonic flow around missiles at transonic and supersonic speeds, with the final goal of developing a computer code for use in aeroelastic calculations or during maneuvers. The basic equations derived for this method are an extension of those derived by Klopfer and Nixon (1989) for steady flow and are a subset of the Euler equations. In this approach, the five Euler equations are reduced to an equation similar to the three-dimensional unsteady potential equation, and a two-dimensional Poisson equation. In addition, one of the equations in this method is almost identical to the potential equation for which there are well tested computer codes, allowing the development of a prediction method based in part on proved technology.
A parametric study of transonic blade-vortex interaction noise
NASA Technical Reports Server (NTRS)
Lyrintzis, A. S.
1991-01-01
Several parameters of transonic blade-vortex interactions (BVI) are being studied and some ideas for noise reduction are introduced and tested using numerical simulation. The model used is the two-dimensional high frequency transonic small disturbance equation with regions of distributed vorticity (VTRAN2 code). The far-field noise signals are obtained by using the Kirchhoff method with extends the numerical 2-D near-field aerodynamic results to the linear acoustic 3-D far-field. The BVI noise mechanisms are explained and the effects of vortex type and strength, and angle of attack are studied. Particularly, airfoil shape modifications which lead to noise reduction are investigated. The results presented are expected to be helpful for better understanding of the nature of the BVI noise and better blade design.
Convergence acceleration and shock fitting for transonic aerodynamics computations
NASA Technical Reports Server (NTRS)
Hafez, M. M.; Cheng, H. K.
1975-01-01
Two problems in computational fluid dynamics are studied in the context of transonic small-disturbance theory - namely, (1) how to speed up the convergence for currently available iterative procedures, and (2) how a shock-fitting method may be adapted to existing relaxation procedures with minimal alterations in computer programming and storage requirements. The paper contributes to a clarification of error analyses for sequence transformations based on the power method (including also the nonlinear transforms of Aitken, Shanks, and Wilkinson), and to developing a cyclic iterative procedure applying the transformations. Examples testing the procedure for a model Dirichlet problem and for a transonic airfoil problem show that savings in computer time by a factor of three to five are generally possible, depending on accuracy requirements and the particular iterative procedure used.-
Unsteady transonic potential flow over a flexible fuselage
NASA Technical Reports Server (NTRS)
Gibbons, Michael D.
1993-01-01
A flexible fuselage capability has been developed and implemented within version 1.2 of the CAP-TSD code. The capability required adding time dependent terms to the fuselage surface boundary conditions and the fuselage surface pressure coefficient. The new capability will allow modeling the effect of a flexible fuselage on the aeroelastic stability of complex configurations. To assess the flexible fuselage capability several steady and unsteady calculations have been performed for slender fuselages with circular cross-sections. Steady surface pressures are compared with experiment at transonic flight conditions. Unsteady cross-sectional lift is compared with other analytical results at a low subsonic speed and a transonic case has been computed. The comparisons demonstrate the accuracy of the flexible fuselage modifications.
Analysis of transonic flow about lifting wing-body configurations
NASA Technical Reports Server (NTRS)
Barnwell, R. W.
1975-01-01
An analytical solution was obtained for the perturbation velocity potential for transonic flow about lifting wing-body configurations with order-one span-length ratios and small reduced-span-length ratios and equivalent-thickness-length ratios. The analysis is performed with the method of matched asymptotic expansions. The angles of attack which are considered are small but are large enough to insure that the effects of lift in the region far from the configuration are either dominant or comparable with the effects of thickness. The modification to the equivalence rule which accounts for these lift effects is determined. An analysis of transonic flow about lifting wings with large aspect ratios is also presented.
Transonic Flows of Bethe-Zel'dovich-Thompson Fluids
NASA Astrophysics Data System (ADS)
Cramer, Mark; Andreyev, Aleksandr
2013-11-01
We examine steady transonic flows of Bethe-Zel'dovich-Thompson (BZT) fluids over thin turbine blades or airfoils. BZT fluids are ordinary fluids having a region of negative fundamental derivative over a finite range of pressures and temperatures in the single phase regime. We present the transonic small disturbance equation, shock jump conditions, and shock existence conditions capable of capturing the qualitative behavior of BZT fluids. The flux function is seen to be quartic in the pressure or density perturbation rather than the quadratic (convex) flux function of the perfect gas theory. We show how this nonconvex flux function can be used to predict and explain the complex flows possible. Numerical solutions using a successive line relaxation (SLR) scheme are presented. New results of interest include shock-splitting, collisions between expansion and compression shocks, two compressive bow shocks in supersonic flows, and the observation of as many as three normal stern shocks following an oblique trailing edge shock.
Design of transonic airfoil sections using a similarity theory
NASA Technical Reports Server (NTRS)
Nixon, D.
1978-01-01
A study of the available methods for transonic airfoil and wing design indicates that the most powerful technique is the numerical optimization procedure. However, the computer time for this method is relatively large because of the amount of computation required in the searches during optimization. The optimization method requires that base and calibration solutions be computed to determine a minimum drag direction. The design space is then computationally searched in this direction; it is these searches that dominate the computation time. A recent similarity theory allows certain transonic flows to be calculated rapidly from the base and calibration solutions. In this paper the application of the similarity theory to design problems is examined with the object of at least partially eliminating the costly searches of the design optimization method. An example of an airfoil design is presented.
Validation of Blockage Interference Corrections in the National Transonic Facility
NASA Technical Reports Server (NTRS)
Walker, Eric L.
2007-01-01
A validation test has recently been constructed for wall interference methods as applied to the National Transonic Facility (NTF). The goal of this study was to begin to address the uncertainty of wall-induced-blockage interference corrections, which will make it possible to address the overall quality of data generated by the facility. The validation test itself is not specific to any particular modeling. For this present effort, the Transonic Wall Interference Correction System (TWICS) as implemented at the NTF is the mathematical model being tested. TWICS uses linear, potential boundary conditions that must first be calibrated. These boundary conditions include three different classical, linear. homogeneous forms that have been historically used to approximate the physical behavior of longitudinally slotted test section walls. Results of the application of the calibrated wall boundary conditions are discussed in the context of the validation test.
Studies in a transonic rotor aerodynamics and noise facility
NASA Technical Reports Server (NTRS)
Wright, S. E.; Lee, D. J.; Crosby, W.
1984-01-01
The design, construction and testing of a transonic rotor aerodynamics and noise facility was undertaken, using a rotating arm blade element support technique. This approach provides a research capability intermediate between that of a stationary element in a moving flow and that of a complete rotating blade system, and permits the acoustic properties of blade tip elements to be studied in isolation. This approach is an inexpensive means of obtaining data at high subsonic and transonic tip speeds on the effect of variations in tip geometry. The facility may be suitable for research on broad band noise and discrete noise in addition to high-speed noise. Initial tests were conducted over the Mach number range 0.3 to 0.93 and confirmed the adequacy of the acoustic treatment used in the facility to avoid reflection from the enclosure.
A computational design method for transonic turbomachinery cascades
NASA Technical Reports Server (NTRS)
Sobieczky, H.; Dulikravich, D. S.
1982-01-01
This paper describes a systematical computational procedure to find configuration changes necessary to modify the resulting flow past turbomachinery cascades, channels and nozzles, to be shock-free at prescribed transonic operating conditions. The method is based on a finite area transonic analysis technique and the fictitious gas approach. This design scheme has two major areas of application. First, it can be used for design of supercritical cascades, with applications mainly in compressor blade design. Second, it provides subsonic inlet shapes including sonic surfaces with suitable initial data for the design of supersonic (accelerated) exits, like nozzles and turbine cascade shapes. This fast, accurate and economical method with a proven potential for applications to three-dimensional flows is illustrated by some design examples.
Numerical studies of unsteady transonic flow over an oscillating airfoil
NASA Technical Reports Server (NTRS)
Chyu, W. J.; Davis, S. S.
1984-01-01
A finite-difference solution to the Navier-Stokes equations combined with a time-varying grid-generation technique was used to compute unsteady transonic flow over an oscillating airfoil. These computations were compared with experimental data (obtained at Ames Research Center) which form part of the AGARD standard configuration for aeroelastic analysis. A variety of approximations to the full Navier-Stokes equations was used to determine the effect of frequency, shock-wave motion, flow separation, and airfoil geometry on unsteady pressures and overall air loads. Good agreement is shown between experiment and theory with the limiting factor being the lack of a reliable turbulence model for high-Reynolds-number, unsteady transonic flows.
Space Shuttle Model In The 16 Foot Transonic Tunnel
NASA Technical Reports Server (NTRS)
1978-01-01
What may appear at first glance to be a swimming shark is a wind tunnel model of the Space Shuttle Orbiter, being tested at NASA's Langley Research Center in Hampton,VA. The Orbiter model is 5.5 feet long (1/20th of the real Orbiter's length) and has remotely operated control surfaces. Inside Langley's 16 foot Transonic Wind Tunnel, the model simulated Orbiter re-entry into the Earth's atmosphere, when it must fly through the transonic speed range (the range that crosses the sound barrier). Information on Orbiter stability and control, collected and analyzed during the tests, were integrated with other data to become part of computerized flight simulation programs.
Laser velocimetry applied to transonic and supersonic aerodynamics
NASA Technical Reports Server (NTRS)
Johnson, D. A.; Bachalo, W. D.; Moddaress, D.
1976-01-01
Measurements obtained with laser velocimetry in a Mach 2.9 separated turbulent boundary layer and in the transonic flow past a two-dimensional airfoil section are presented and compared to data realized by conventional techniques. Agreement in mean velocities was realized where the pressure measurements could be considered reliable; however, in regions of instantaneous reverse velocities, the laser results were found to be consistent with the physics of the flow whereas the pressure data were not. Streamwise turbulence intensities are also presented. In the transonic airfoil study, velocity measurements obtained immediately outside the upper surface boundary layer of a 6-inch chord NACA 64A010 airfoil are compared to edge velocities inferred from surface pressure measurements. For free-stream Mach numbers of 0.6 and 0.8, the agreement in results was very good. "Dual scatter" optical arrangements in conjunction with a single particle, counter-type signal processor were employed in these investigations.
National Transonic Facility Fan Blade prepreg material characterization tests
NASA Technical Reports Server (NTRS)
Klich, P. J.; Richards, W. H.; Ahl, E. L., Jr.
1981-01-01
The test program for the basic prepreg materials used in process development work and planned fabrication of the national transonic facility fan blade is presented. The basic prepreg materials and the design laminate are characterized at 89 K, room temperature, and 366 K. Characterization tests, test equipment, and test data are discussed. Material tests results in the warp direction are given for tensile, compressive, fatigue (tension-tension), interlaminar shear and thermal expansion.
Some examples of unsteady transonic flows over airfoils
NASA Technical Reports Server (NTRS)
Ballhaus, W. F.; Magnus, R.; Yoshihara, H.
1975-01-01
A finite difference flutter analysis is presented for the NACA 64A-410 airfoil at M equals 0.72, where the incidence is abruptly changed from 2 to 4 degrees. The effect of gust loads is studied, and the unsteady flow adjusting process is displayed. The semi-implicit procedure of Ballhaus and Lomax (1974) is used to solve the small disturbance transonic potential equation. The physical aspects of the results, rather than the numerical details, are emphasized.
Preliminary calibration and test results from the National Transonic Facility
NASA Technical Reports Server (NTRS)
Mckinney, Linwood W.; Fuller, Dennis E.
1986-01-01
The National Transonic Facility (NTF) was operated to design condition of 120 million Reynolds number at a Mach number of 1.0. All systems were checked out except plenum isolation valves; modifications are being made to heaters on the actuators. Initial steady-state calibration indicates excellent steady flow characteristics. The first test of the Pathfinder 1 model indicated significant Reynolds number effects. Some effect of temperature on instrumentation were obtained. The cause of these effects is being evaluated.
Spectral multigrid methods with applications to transonic potential flow
NASA Technical Reports Server (NTRS)
Streett, C. L.; Zang, T. A.; Hussaini, M. Y.
1983-01-01
Spectral multigrid methods are demonstrated to be a competitive technique for solving the transonic potential flow equation. The spectral discretization, the relaxation scheme, and the multigrid techniques are described in detail. Significant departures from current approaches are first illustrated on several linear problems. The principal applications and examples, however, are for compressible potential flow. These examples include the relatively challenging case of supercritical flow over a lifting airfoil.
Spectral multigrid methods with applications to transonic potential flow
NASA Technical Reports Server (NTRS)
Streett, C. L.; Zang, T. A.; Hussaini, M. Y.
1985-01-01
Spectral multigrid methods are demonstrated to be a competitive technique for solving the transonic potential flow equation. The spectral discretization, the relaxation scheme, and the multigrid techniques are described in detail. Significant departures from current approaches are first illustrated on several linear problems. The principal applications and examples, however, are for compressible potential flow. These examples include the relatively challenging case of supercritical flow over a lifting airfoil.
Calculation of transonic flows using an extended integral equation method
NASA Technical Reports Server (NTRS)
Nixon, D.
1976-01-01
An extended integral equation method for transonic flows is developed. In the extended integral equation method velocities in the flow field are calculated in addition to values on the aerofoil surface, in contrast with the less accurate 'standard' integral equation method in which only surface velocities are calculated. The results obtained for aerofoils in subcritical flow and in supercritical flow when shock waves are present compare satisfactorily with the results of recent finite difference methods.
Cryogenic Balance Technology at the National Transonic Facility
NASA Technical Reports Server (NTRS)
Parker, P. A.
2001-01-01
This paper provides an overview of force measurement at the National Transonic Facility (NTF). The NTF has unique force measurement requirements that dictate an integration of all aspects of balance design, production, and calibration. An overview of current force measurement capabilities is provided along with new balance development efforts. Research activities in the areas of thermal compensation and balance calibration are presented. Also, areas of future research are detailed.
Multiple Solutions of Transonic Flow over NACA0012 Airfoil
NASA Astrophysics Data System (ADS)
Xiong, Juntao; Liu, Ya; Liu, Feng; Luo, Shijun; Zhao, Zijie; Ren, Xudong; Gao, Chao
2012-11-01
Multiple solutions of the small-disturbance potential equation and full potential equation were known for the NACA0012 airfoil in a certain range of transonic Mach numbers and at zero angle of attack. However the multiple solutions for this airfoil were not observed using Euler or Navier-Stokes equations under the above flow conditions. In the present work, both the Unsteady Reynolds-Averaged Navier-Stokes (URANS) computations and transonic wind tunnel experiments are performed under certain Reynolds numbers to further study the problem. The results of the two methods reveal that buffet appears in a narrow Mach number range where the potential flow methods predict multiple solutions. Boundary layer displacement thickness computed from URANS at the same flow condition is used to modify the geometry of the airfoil. Euler equations are then solved for the modified geometry. The results show that the addition of the boundary layer displacement thickness creates multiple solutions for the NACA0012 airfoil. Global linear stability analysis is also performed on the original and the modified airfoils. This shows a close relationship between the viscous unsteady shock buffet phenomenon of transonic airfoil flow and the existence of multiple solutions of the external inviscid flow. Postdoctoral Research Assistant.
Unsteady transonic flow calculations for wing-fuselage configurations
NASA Technical Reports Server (NTRS)
Batina, J. T.
1986-01-01
Unsteady transonic flow calculations are presented for wing-fuselage configurations. Calculations are performed by extending the XTRAN3S unsteady transonic small-disturbance code to allow the treatment of a fuselage. Details of the XTRAN3S fuselage modeling are discussed in the context of the small-disturbance equation. Transonic calculations are presented for three wing-fuselage configurations with leading edge sweep angles ranging from 0 deg to 46.76 deg. Simple bending and torsion modal oscillations of the wing are calculated. Sectional lift and moment coefficients for the wing-alone and wing-fuselage cases are compared and the effects of fuselage aerodynamic interference on the unsteady wing loading are revealed. Tabulated generalized aerodynamic forces used in flutter analyses, indicate small changes in the real in-phase component and as much as a 30% change in the imaginary component when the fuselage is included in the calculation. These changes result in a 2 to 5% increase in total magnitude and a several degree increase in phase.
Fourier time spectral method for subsonic and transonic flows
NASA Astrophysics Data System (ADS)
Zhan, Lei; Liu, Feng; Papamoschou, Dimitri
2016-06-01
The time accuracy of the exponentially accurate Fourier time spectral method (TSM) is examined and compared with a conventional 2nd-order backward difference formula (BDF) method for periodic unsteady flows. In particular, detailed error analysis based on numerical computations is performed on the accuracy of resolving the local pressure coefficient and global integrated force coefficients for smooth subsonic and non-smooth transonic flows with moving shock waves on a pitching airfoil. For smooth subsonic flows, the Fourier TSM method offers a significant accuracy advantage over the BDF method for the prediction of both the local pressure coefficient and integrated force coefficients. For transonic flows where the motion of the discontinuous shock wave contributes significant higher-order harmonic contents to the local pressure fluctuations, a sufficient number of modes must be included before the Fourier TSM provides an advantage over the BDF method. The Fourier TSM, however, still offers better accuracy than the BDF method for integrated force coefficients even for transonic flows. A problem of non-symmetric solutions for symmetric periodic flows due to the use of odd numbers of intervals is uncovered and analyzed. A frequency-searching method is proposed for problems where the frequency is not known a priori. The method is tested on the vortex shedding problem of the flow over a circular cylinder.
MAVRIC Flutter Model Transonic Limit Cycle Oscillation Test
NASA Technical Reports Server (NTRS)
Edwards, John W.; Schuster, David M.; Spain, Charles V.; Keller, Donald F.; Moses, Robert W.
2001-01-01
The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations of clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including LCO behavior.
MAVRIC Flutter Model Transonic Limit Cycle Oscillation Test
NASA Technical Reports Server (NTRS)
Edwards, John W.; Schuster, David M.; Spain, Charles V.; Keller, Donald F.; Moses, Robert W.
2001-01-01
The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including Limit Cycle Oscillation behavior.
Fourier time spectral method for subsonic and transonic flows
NASA Astrophysics Data System (ADS)
Zhan, Lei; Liu, Feng; Papamoschou, Dimitri
2016-01-01
The time accuracy of the exponentially accurate Fourier time spectral method (TSM) is examined and compared with a conventional 2nd-order backward difference formula (BDF) method for periodic unsteady flows. In particular, detailed error analysis based on numerical computations is performed on the accuracy of resolving the local pressure coefficient and global integrated force coefficients for smooth subsonic and non-smooth transonic flows with moving shock waves on a pitching airfoil. For smooth subsonic flows, the Fourier TSM method offers a significant accuracy advantage over the BDF method for the prediction of both the local pressure coefficient and integrated force coefficients. For transonic flows where the motion of the discontinuous shock wave contributes significant higher-order harmonic contents to the local pressure fluctuations, a sufficient number of modes must be included before the Fourier TSM provides an advantage over the BDF method. The Fourier TSM, however, still offers better accuracy than the BDF method for integrated force coefficients even for transonic flows. A problem of non-symmetric solutions for symmetric periodic flows due to the use of odd numbers of intervals is uncovered and analyzed. A frequency-searching method is proposed for problems where the frequency is not known a priori. The method is tested on the vortex shedding problem of the flow over a circular cylinder.
Transonic flows of dense gases over finite wings
NASA Astrophysics Data System (ADS)
Cinnella, P.
2008-04-01
Transonic inviscid flows of dense gases of the Bethe-Zel'dovich-Thompson (BZT) type over finite wings are numerically investigated. BZT gases are fluids of the retrograde type (i.e., that superheat when expanded), which exhibit a region of negative values of the fundamental derivative of gas dynamics Γ. As a consequence, they display, in the transonic and supersonic regime, nonclassical gas dynamic behaviors, such as rarefaction shock waves and mixed shock/fan waves. The peculiar properties of BZT fluids have received increased interest in recent years because of their possible application in energy-conversion cycles. The present research aims at providing insight about the transonic aerodynamics of BZT fluids past finite wings, roughly representative of isolated turbine blades with infinite tip leakage. This represents an important step toward the design of advanced turbine blades by using organic working fluids. An investigation of the flow patterns and aerodynamic performance for several choices of the upstream thermodynamic conditions is provided, and the advantages of using BZT working fluids instead of classical ones are discussed.
A numerical study of flutter in a transonic fan
Isomura, K.; Giles, M.B.
1998-07-01
The bending mode flutter of a modern transonic fan has been studied using a quasi-three-dimensional viscous unsteady CFD code. The type of flutter in this research is that of a highly loaded blade with a tip relative Mach number just above unity, commonly referred to as transonic stall flutter. This type of flutter is often encountered in modern wide chord fans without a part span shroud. The CFD simulation uses an upwinding scheme with Roe`s third-order flux differencing, and Johnson and King`s turbulence model with the later modification due to Johnson and Coakley. A dynamic transition point model is developed using the e{double_prime} method and Schubauer and Klebanoff`s experimental data. The calculations of the flow in this fan reveal that the source of the flutter of 1H1 transonic fan is an oscillation of the passage shock, rather than a stall. As the blade loading increases, the passage shock moves forward. Just before the passage shock unstarts, the stability of the passage shock decreases, and a small blade vibration causes the shock to oscillate with a large amplitude between unstarted and started positions. The dominant component of the blade excitation force is due to the foot of the oscillating passage shock on the blade pressure surface.
Parametric Evaluation of Thin, Transonic Circulation-Control Airfoils
NASA Technical Reports Server (NTRS)
Schlecht, Robin; Anders, Scott
2007-01-01
Wind-tunnel tests were conducted in the NASA Langley Transonic Dynamics Tunnel on a 6 percent-thick, elliptical circulation-control airfoil with upper-surface and lower-surface blowing capability. Results for elliptical Coanda trailing-edge geometries, biconvex Coanda trailing-edge geometries, and leading-edge geometries are reported. Results are presented at subsonic and transonic Mach numbers of 0.3 and 0.8, respectively. When considering one fixed trailing-edge geometry, for both the subsonic and transonic conditions it was found that the [3.0:1] ratio elliptical Coanda surface with the most rounded leading-edge [03] performed favorably and was determined to be the best compromise between comparable configurations that took advantage of the Coanda effect. This configuration generated a maximum. (Delta)C(sub 1) = 0.625 at a C(sub mu) = 0.06 at M = 0.3, alpha = 6deg. This same configuration generated a maximum (Delta)C(sub 1) = 0.275 at a C(sub mu) = 0.0085 at M = 0.8, alpha = 3deg.
Transonic flow past a wedge profile with detached bow wave
NASA Technical Reports Server (NTRS)
Vincenti, Walter G; Wagoner, Cleo B
1952-01-01
A theoretical study has been made of the aerodynamic characteristics at zero angle of attack of a thin, doubly symmetrical double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis utilizes the equations of the transonic small-disturbance theory and involves no assumptions beyond those implicit in this theory. The mixed flow about the front half of the profile is calculated by relaxation solution of boundary conditions along the shock polar and sonic line. The purely subsonic flow about the rear of the profile is found by means of the method of characteristics specialized to the transonic small-disturbance theory. Complete calculations were made for four values of the transonic similarity parameter. These were found sufficient to bridge the gap between the previous results of Guderley and Yoshihara at a Mach number of 1 and the results which are readily obtained when the bow wave is attached and the flow is completely supersonic.
4. VIEW LOOKING NORTHNORTHEAST AT TEST SECTION OF 8FOOT TRANSONIC ...
4. VIEW LOOKING NORTH-NORTHEAST AT TEST SECTION OF 8-FOOT TRANSONIC PRESSURE TUNNEL SHOWING ACCESS PORT TO TEST SECTION (RIGHT) AND PLENUM SURROUNDING AREA. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
Design features and operational characteristics of the Langley pilot transonic cryogenic tunnel
NASA Technical Reports Server (NTRS)
Kilgore, R. A.
1974-01-01
A fan-driven transonic cryogenic tunnel was designed, and its purging, cooldown, and warmup times were determined satisfactory. Cooling with liquid nitrogen is at the power levels required for transonic testing. Good temperature distributions are obtained by using a simple nitrogen injection system.
On the application of transonic similarity rules to wings of finite span
NASA Technical Reports Server (NTRS)
Spreiter, John R
1953-01-01
The transonic aerodynamic characteristics of wings of finite span are discussed from the point of view of a unified small perturbation theory for subsonic, transonic, and supersonic flows about thin wings. This approach avoids certain ambiguities which appear if one studies transonic flows by means of equations derived under the more restrictive assumption that the local velocities are everywhere close to sonic velocity. The relation between the two methods of analysis of transonic flow is examined, the similarity rules and known solutions of transonic flow theory are reviewed, and the asymptotic behavior of the lift, drag, and pitching-moment characteristics of wings of large and small aspect ratio is discussed. It is shown that certain methods of data presentation are advantageous for the effective display of these characteristics.
NASA Technical Reports Server (NTRS)
Yates, E. C., Jr.; Wynne, E. C.; Farmer, M. G.; Desmarais, R. N.
1981-01-01
Use of a supercritical airfoil can adversely affect wing flutter speeds in the transonic range. As adequate theories for three dimensional unsteady transonic flow are not yet available, the modified strip analysis was used to predict the transonic flutter boundary for the supercritical wing. The steady state spanwise distributions of section lift curve slope and aerodynamic center, required as input for the flutter calculations, were obtained from pressure distributions. The calculated flutter boundary is in agreement with experiment in the subsonic range. In the transonic range, a transonic bucket is calculated which closely resembles the experimental one with regard to both shape and depth, but it occurs at about 0.04 Mach number lower than the experimental one.
NASA Technical Reports Server (NTRS)
Howell, Robert R.; Braslow, Albert L.
1955-01-01
A comparison of the zero-lift drag coefficients at Mach numbers from 0.81 to 1.41 of a fin-stabilized parabolic body of revolution as measured in the Langley transonic blowdown tunnel has been made with measurements obtained in free-flight on a larger but geometrically similar model. The absolute values of drag coefficient obtained in the slotted wind tunnel were equivalent to the free-flight drag-coefficient values up to a Mach number of 1.4 when adjustments were made for the effect on viscous drag of differences in Reynolds number between the two test conditions. Excellent agreement was obtained between the two tests for the pressure-drag variation with Mach number, regardless of whether the scale effect on skin friction was considered. Favorable agreement was also obtained between the pressure-drag increments due t o the presence of the stabilizing fins as determined in the wine tunnel from fins-on and fins-off tests and as obtained by a different method in free flight. Tests of a specific airplane configuration to obtain an indication of the problems involved in the construction and tests of small-scale (approximately 7-inch span) complete airplane configuration with internal air flow indicated that reliable zero-lift drag-coefficient measurements at Mach numbers up to 1.4 can be attained with such models, provided the model is constructed with a high but not an unreasonable degree of accuracy.
NASA Technical Reports Server (NTRS)
Sewall, W. G.
1982-01-01
A transonic similarity rule which accounts for the effects of attached sidewall boundary layers is presented and evaluated by comparison with the characteristics of airfoils tested in a two dimensional transonic tunnel with different sidewall boundary layer thicknesses. The rule appears valid provided the sidewall boundary layer both remains attached in the vicinity of the model and occupies a small enough fraction of the tunnel width to preserve sufficient two dimensionality in the tunnel.
Adjoint-based airfoil shape optimization in transonic flow
NASA Astrophysics Data System (ADS)
Gramanzini, Joe-Ray
The primary focus of this work is efficient aerodynamic shape optimization in transonic flow. Adjoint-based optimization techniques are employed on airfoil sections and evaluated in terms of computational accuracy as well as efficiency. This study examines two test cases proposed by the AIAA Aerodynamic Design Optimization Discussion Group. The first is a two-dimensional, transonic, inviscid, non-lifting optimization of a Modified-NACA 0012 airfoil. The second is a two-dimensional, transonic, viscous optimization problem using a RAE 2822 airfoil. The FUN3D CFD code of NASA Langley Research Center is used as the ow solver for the gradient-based optimization cases. Two shape parameterization techniques are employed to study their effect and the number of design variables on the final optimized shape: Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD) and the BandAids free-form deformation technique. For the two airfoil cases, angle of attack is treated as a global design variable. The thickness and camber distributions are the local design variables for MASSOUD, and selected airfoil surface grid points are the local design variables for BandAids. Using the MASSOUD technique, a drag reduction of 72.14% is achieved for the NACA 0012 case, reducing the total number of drag counts from 473.91 to 130.59. Employing the BandAids technique yields a 78.67% drag reduction, from 473.91 to 99.98. The RAE 2822 case exhibited a drag reduction from 217.79 to 132.79 counts, a 39.05% decrease using BandAids.
A transonic-small-disturbance wing design methodology
NASA Technical Reports Server (NTRS)
Phillips, Pamela S.; Waggoner, Edgar G.; Campbell, Richard L.
1988-01-01
An automated transonic design code has been developed which modifies an initial airfoil or wing in order to generate a specified pressure distribution. The design method uses an iterative approach that alternates between a potential-flow analysis and a design algorithm that relates changes in surface pressure to changes in geometry. The analysis code solves an extended small-disturbance potential-flow equation and can model a fuselage, pylons, nacelles, and a winglet in addition to the wing. A two-dimensional option is available for airfoil analysis and design. Several two- and three-dimensional test cases illustrate the capabilities of the design code.
Initial research program for the National Transonic Facility
NASA Technical Reports Server (NTRS)
Gloss, B. B.
1984-01-01
The construction and checkout of the National Transonic Facility (NTF) have been completed, and detailed calibration is now in progress. The initial NTF research program covers a wide range of study areas falling into three major elements: (1) the assessment of Reynolds number sensitivities for a broad range of configurations and flow phenomena; (2) validation of the ability of NTF to simulate full-scale aerodynamics; and (3) the development of test techniques for improved test simulations in existing wind tunnels. This paper, therefore, is a status report on these various elements of the initial NTF research program.
A hybrid algorithm for transonic airfoil and wing design
NASA Technical Reports Server (NTRS)
Campbell, Richard L.; Smith, Leigh A.
1987-01-01
The present method for the design of transonic airfoils and wings employs a predictor/corrector approach in which an analysis code calculates the flowfield for an initial geometry, then modifies it on the basis of the difference between calculated and target pressures. This allows the design method to be straightforwardly coupled with any existing analysis code, as presently undertaken with several two- and three-dimensional potential flow codes. The results obtained indicate that the method is robust and accurate, even in the cases of airfoils with strongly supercritical flow and shocks. The design codes are noted to require computational resources typical of current pure-inverse methods.
An inverse method with regularity condition for transonic airfoil design
NASA Technical Reports Server (NTRS)
Zhu, Ziqiang; Xia, Zhixun; Wu, Liyi
1991-01-01
It is known from Lighthill's exact solution of the incompressible inverse problem that in the inverse design problem, the surface pressure distribution and the free stream speed cannot both be prescribed independently. This implies the existence of a constraint on the prescribed pressure distribution. The same constraint exists at compressible speeds. Presented here is an inverse design method for transonic airfoils. In this method, the target pressure distribution contains a free parameter that is adjusted during the computation to satisfy the regularity condition. Some design results are presented in order to demonstrate the capabilities of the method.
Refined numerical solution of the transonic flow past a wedge
NASA Technical Reports Server (NTRS)
Liang, S.-M.; Fung, K.-Y.
1985-01-01
A numerical procedure combining the ideas of solving a modified difference equation and of adaptive mesh refinement is introduced. The numerical solution on a fixed grid is improved by using better approximations of the truncation error computed from local subdomain grid refinements. This technique is used to obtain refined solutions of steady, inviscid, transonic flow past a wedge. The effects of truncation error on the pressure distribution, wave drag, sonic line, and shock position are investigated. By comparing the pressure drag on the wedge and wave drag due to the shocks, a supersonic-to-supersonic shock originating from the wedge shoulder is confirmed.
Interaction of multiple supersonic jets with a transonic flow field
NASA Technical Reports Server (NTRS)
Seginer, A.; Manela, J.
1983-01-01
The influence of multiple high pressure, supersonic, radial or tangential jets, that are injected from the circumference of the base plane of an axisymmetric body, on its longitudinal aerodynamic coefficients in transonic flow is studied experimentally. The interaction of the jets with the body flow field increases the pressures on the forebody, thus altering its lift and static stability characteristics. It is shown that, within the range of parameters studied. This interaction has a stabilizing effect on the body. The contribution to lift and stability is significant at small angles of attack and decreases nonlinearly at higher angles when the crossflow mechanism becomes dominant.
Aeroelastic Tailoring of Transport Wings Including Transonic Flutter Constraints
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Wieseman, Carol D.; Jutte, Christine V.
2015-01-01
Several minimum-mass optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic stress and panel buckling constraints are imposed across several trimmed static maneuver loads, in addition to a transonic flutter margin constraint, captured with aerodynamic influence coefficient-based tools. Tailoring with metallic thickness variations, functionally graded materials, balanced or unbalanced composite laminates, curvilinear tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.
Viscous transonic flow computation over Space Shuttle configuration
NASA Technical Reports Server (NTRS)
Fujii, K.; Kutler, P.
1984-01-01
A thin-layer Navier-Stokes code capable of predicting steady-state viscous flows is applied to the transonic flow over a Space Shuttle configuration. The code is written in the generalized coordinate system, and the grid-generation code of Fujii (1983) is used for the discretization of the flow field. The flow-field computation is done using the CRAY 1S computer at NASA Ames. The computed result is physically reasonable, even though no experimental data is available for the comparison purpose.
Fast Euler solver for transonic airfoils. I - Theory. II - Applications
NASA Technical Reports Server (NTRS)
Dadone, Andrea; Moretti, Gino
1988-01-01
Equations written in terms of generalized Riemann variables are presently integrated by inverting six bidiagonal matrices and two tridiagonal matrices, using an implicit Euler solver that is based on the lambda-formulation. The solution is found on a C-grid whose boundaries are very close to the airfoil. The fast solver is then applied to the computation of several flowfields on a NACA 0012 airfoil at various Mach number and alpha values, yielding results that are primarily concerned with transonic flows. The effects of grid fineness and boundary distances are analyzed; the code is found to be robust and accurate, as well as fast.
Numerical calculation of transonic flow about slender bodies of revolution
NASA Technical Reports Server (NTRS)
Bailey, F. R.
1971-01-01
A relaxation method is described for the numerical solution of the transonic small disturbance equation for flow about a slender body of revolution. Results for parabolic arc bodies, both with and without an attached sting, are compared with wind-tunnel measurements for a free-stream Mach number range from 0.90 to 1.20. The method is also used to show the effects of wind-tunnel wall interference by including boundary conditions representing porous-wall and open-jet wind-tunnel test sections.
Transonic wall interference effects on bodies of revolution.
NASA Technical Reports Server (NTRS)
Couch, L. M.
1972-01-01
Efforts to develop a near sonic transport have placed renewed emphasis on obtaining accurate aerodynamic force and pressure data in the near sonic speed range. Comparison of wind-tunnel and flight data obtained for a blunt-nose body of revolution showed significant discrepancies in drag levels near Mach 1 - apparently due to wind-tunnel wall interference. Subsequent tests of geometrically similar bodies of revolution showed that increasing the model-to-test-section blockage ratio from 0.00017 to 0.0043 resulted in altered drag curve shapes, delayed drag divergence, and 'transonic creep' from subsonic drag levels due to increased wall interference.
Operational manual for two-dimensional transonic code TSFOIL
NASA Technical Reports Server (NTRS)
Stahara, S. S.
1978-01-01
This code solves the two-dimensional, transonic, small-disturbance equations for flow past lifting airfoils in both free air and various wind-tunnel environments by using a variant of the finite-difference method. A description of the theoretical and numerical basis of the code is provided, together with complete operating instructions and sample cases for the general user. In addition, a programmer's manual is also presented to assist the user interested in modifying the code. Included in the programmer's manual are a dictionary of subroutine variables in common and a detailed description of each subroutine.
An analysis method for two-dimensional transonic viscous flow
NASA Technical Reports Server (NTRS)
Bavitz, P. C.
1975-01-01
A method for the approximate calculation of transonic flow over airfoils, including shock waves and viscous effects, is described. Numerical solutions are obtained by use of a computer program which is discussed in the appendix. The importance of including the boundary layer in the analysis is clearly demonstrated, as well as the need to improve on existing procedures near the trailing edge. Comparisons between calculations and experimental data are presented for both conventional and supercritical airfoils, emphasis being on the surface pressure distribution, and good agreement is indicated.
Finite element analysis of periodic transonic flow problems
NASA Technical Reports Server (NTRS)
Fix, G. J.
1978-01-01
Flow about an oscillating thin airfoil in a transonic stream was considered. It was assumed that the flow field can be decomposed into a mean flow plus a periodic perturbation. On the surface of the airfoil the usual Neumman conditions are imposed. Two computer programs were written, both using linear basis functions over triangles for the finite element space. The first program uses a banded Gaussian elimination solver to solve the matrix problem, while the second uses an iterative technique, namely SOR. The only results obtained are for an oscillating flat plate.
Transonic Turbulent Flow Predictions With Two-Equation Turbulence Models
NASA Technical Reports Server (NTRS)
Liou, William W.; Shih, Tsan-Hsing
1996-01-01
Solutions of the Favre-averaged Navier-Stokes equations for two well-documented transonic turbulent flows are compared in detail with existing experimental data. While the boundary layer in the first case remains attached, a region of extensive flow separation has been observed in the second case. Two recently developed k-epsilon, two-equation, eddy-viscosity models are used to model the turbulence field. These models satisfy the realizability constraints of the Reynolds stresses. Comparisons with the measurements are made for the wall pressure distribution, the mean streamwise velocity profiles, and turbulent quantities. Reasonably good agreement is obtained with the experimental data.
Aerodynamic optimum design of transonic turbine cascades using Genetic Algorithms
NASA Astrophysics Data System (ADS)
Li, Jun; Feng, Zhenping; Chang, Jianzhong; Shen, Zuda
1997-06-01
This paper presents an aerodynamic optimum design method for transonic turbine cascades based on the Genetic Algorithms coupled to the inviscid flow Euler solver and the boundary-layer calculation. The Genetic Algorithms control the evolution of a population of cascades towards an optimum design. The fitness value of each string is evaluated using the flow solver. The design procedure has been developed and the behavior of the genetic algorithms has been tested. The objective functions of the design examples are the minimum mean-square deviation between the aimed pressure and computed pressure and the minimum amount of user expertise.
Evaluation of flow field approximations for transonic compressor stages
Dorney, D.J.; Sharma, O.P.
1997-07-01
The flow through gas turbine compressors is often characterized by unsteady, transonic, and viscous phenomena. Accurately predicting the behavior of these complex multi-blade-row flows with unsteady rotor-stator interacting Navier-Stokes analyses can require enormous computer resources. In this investigation, several methods for predicting the flow field, losses, and performance quantities associated with axial compressor stages are presented. The methods studied include: (1) the unsteady fully coupled blade row technique, (2) the steady coupled blade row method, (3) the steady single blade row technique, and (4) the loosely coupled blade row method. The analyses have been evaluated in terms of accuracy and efficiency.
Optimum Transonic Airfoils Based on the Euler Equations
NASA Technical Reports Server (NTRS)
Iollo, Angelo; Salas, Manuel, D.
1996-01-01
We solve the problem of determining airfoils that approximate, in a least square sense, given surface pressure distributions in transonic flight regimes. The flow is modeled by means of the Euler equations and the solution procedure is an adjoint- based minimization algorithm that makes use of the inverse Theodorsen transform in order to parameterize the airfoil. Fast convergence to the optimal solution is obtained by means of the pseudo-time method. Results are obtained using three different pressure distributions for several free stream conditions. The airfoils obtained have given a trailing edge angle.
Calculation of unsteady transonic flows using the integral equation method
NASA Technical Reports Server (NTRS)
Nixon, D.
1978-01-01
The basic integral equations for a harmonically oscillating airfoil in a transonic flow with shock waves are derived; the reduced frequency is assumed to be small. The problems associated with shock wave motion are treated using a strained coordinate system. The integral equation is linear and consists of both line integrals and surface integrals over the flow field which are evaluated by quadrature. This leads to a set of linear algebraic equations that can be solved directly. The shock motion is obtained explicitly by enforcing the condition that the flow is continuous except at a shock wave. Results obtained for both lifting and nonlifting oscillatory flows agree satisfactorily with other accurate results.
Langley 16- Ft. Transonic Tunnel Pressure Sensitive Paint System
NASA Technical Reports Server (NTRS)
Sprinkle, Danny R.; Obara, Clifford J.; Amer, Tahani R.; Leighty, Bradley D.; Carmine, Michael T.; Sealey, Bradley S.; Burkett, Cecil G.
2001-01-01
This report describes the NASA Langley 16-Ft. Transonic Tunnel Pressure Sensitive Paint (PSP) System and presents results of a test conducted June 22-23, 2000 in the tunnel to validate the PSP system. The PSP system provides global surface pressure measurements on wind tunnel models. The system was developed and installed by PSP Team personnel of the Instrumentation Systems Development Branch and the Advanced Measurement and Diagnostics Branch. A discussion of the results of the validation test follows a description of the system and a description of the test.
Vectorizable multigrid algorithms for transonic-flow calculations
NASA Technical Reports Server (NTRS)
Melson, N. D.
1986-01-01
The analysis and the incorporation into a multigrid scheme of several vectorizable algorithms are discussed. von Neumann analyses of vertical-line, horizontal-line, and alternating-direction ZEBRA algorithms were performed; and the results were used to predict their multigrid damping rates. The algorithms were then successfully implemented in a transonic conservative full-potential computer program. The convergence acceleration effect of multiple grids is shown, and the convergence rates of the vectorizable algorithms are compared with those of standard successive-line overrelaxation (SLOR) algorithms.