Nozzle wall boundary layers at Mach numbers 20 to 47
Nozzle wall hypersonic boundary layers in helium tunnel, presenting skin friction measurements and heat transfer rates
Kemp, J. H., Jr.↗
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Progress on the development of skin friction sensors.
A skin friction measuring system is described which will operate to 2000 F for short periods of time and continuously to 1300 F. A prototype system has been constructed and is being placed in operation. Information on alternative feasible systems is presented. Emphasis is placed on presenting research results in the form of useful design information and experimental data.
Mcvey, E. S.↗
Ground simulation and tunnel blockage for a jet-flapped, basic STOL model tested to very high lift coefficients
Ground effects experiments and large/small-tunnel interference studies were carried out on a model with a 20-inch (50.8 cm) span wing. The wing, which includes a highly deflected knee-blown flap can be fitted with unflapped tips and slats. A low-mounted tailplane can be fitted to the aft fuselage. Three-component balance meaurements, made with a fixed ground equipped with a single boundary-layer blowing slot, were compared with datum, moving-ground results. Good comparisons were obtained up to model blowing momentum coefficients of approximately two, after which the particular floor blowing settings used proved insufficient to prevent floor separation in the vicinity of the model. Skin friction measurements, taken routinely along the floor centerline, proved invaluable during the analysis of results, and their use is recommended as input to determination of floor BLC setting. A careful investigation was made of pitching moments, including tail-on, close-to-ground cases, with favorable results. Drag proved the most sensitive to the change from a moving to the boundary-layer controlled ground.
Hackett, J. E.↗
Fundamental studies of antimisting fuels
Research is currently being undertaken to determine the various physical characteristics of antimisting fuel. It is an endeavor to prevent post-crash fuel mist fires. The following studies were done: (1) drop and jet breakup studies using an image processing technique, (2) flammability of fuel mist under simulated plane crash conditions; (3) skin friction measurements; (4) determination of the tensile viscosity of non-Newtonian antimisting fuel as a function of the extensional rate; and (5) water compatibility measurements. The results indicate that the mechanism which prevents antimisting fuel breakup is related to the time-dependent tensile viscosity of antimisting fuel under stress. Drag reduction phenomenon was also observed at Reynolds numbers higher than 20,000. Uptake of water by the antimisting fuel is higher than that for the neat fuel. In addition, its rate of absorption depends on the concurrent degree of agitation.
Sarohia, V.↗
Design of high-Reynolds-number flat-plate experiments in the NTF
The design of an experiment to measure skin friction and turbulent boundary-layer characteristics at Reynolds numbers exceeding one billion is described. The experiment will be conducted in a zero-pressure-gradient flow on a flat plate in the National Transonic Facility. The development of computational codes to analyze the aerodynamic loads and the blockage is documented. Novel instrumentationn techniques and models, designed to operate in cryogenic environments, are presented. Special problems associated with aerodynamic loads, surface finish, and hot-wire anemometers are discussed.
Saric, W. S.↗
Computations for the 16-foot transonic tunnel, NASA, Langley Research Center, revision 1
The equations used by the 16 foot transonic tunnel in the data reduction programs are presented in eight modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: tunnel parameters; jet exhaust measurements; skin friction drag; balance loads and model attitudes calculations; internal drag (or exit-flow distributions); pressure coefficients and integrated forces; thrust removal options; and turboprop options. This document is a companion document to NASA TM-83186, A User's Guide to the Langley 16 Foot Transonic Tunnel, August 1981.
Mercer, Charles E.↗
A simplified oil-film skin-friction meter
The oil-film method of skin friction measurement, which does not require calibration, nevertheless entails that interferometric measurements of the oil film be obtained. The oil-film method is presently simplified by eliminating interferometry and its requisite polished surface, by taking direct and dynamic measurements of the oil-film slope with a small position-sensing photodiode. This technique has undergone verification in incompressible turbulent and laminar flows in flat-plate boundary layers and pipe flows; the meter is judged to be inexpensive, simple, and robust.
Bandyopadhyay, P. R.↗
Design of high-Reynolds-number flat-plate experiments in the NTF
The design of an experiment to measure skin friction and turbulent boundary layer characteristics at Reynolds numbers exceeding 1 x 10 to the 9th is described. The experiment will be conducted in a zero-pressure-gradient flow on a flat plate in the National Transonic Facility (NTF). The development of computational codes to analyze the aerodynamic loads and the blockage is documented. Novel instrumentation techniques and models, designed to operate in cryogenic environments, are presented. Special problems associated with aerodynamic loads, surface finish, and hot-wire anemometers are discussed.
Saric, William S.↗
Simultaneous detection of separation and transition in surface shear layers
Flush-mounted hot-film gages have proved effective in detecting boundary-layer transition and in measuring skin friction but with limited success in detecting laminar separation and reattachment. The development of multielement micro hot-film sensors, and the recent discovery of the phase reversal phenomena associated with low-frequency dynamic shear stress signals across regions of laminar separation and turbulent reattachment, have made it possible to simultaneously and unambiguously detect these surface shear layer characteristics. Experiments were conducted on different airfoils at speeds ranging from low subsonic to transonic speeds to establish the technique for incompressible and compressible flow applications. The multielement dynamic shear stress sensor technique was successfully used to detect laminar separation, turbulent reattachment, as well as, shock induced laminar and turbulent separation.
Mangalam, Siva M.↗
Internally mounted thin-liquid-film skin-friction meter - Comparison with floating element method with and without pressure gradient
A new, robust oil film skin friction meter was designed and constructed. This enables skin friction measurements remotely and from within the model, as well as avoiding the need to know the location of the leading edge of the film. The instrument was tested by comparing measurements with those given by a floating element gage in a zero pressure gradient flat plate turbulent boundary layer. Both instruments agreed satisfactorily with the well-known curve for this case. Significant discrepancies between the two instruments were observed in the case of adverse and favorable pressure gradients. The discrepancies were of opposite sign for opposite-sign pressure gradients as is consistent with the error expected from floating-element gages. Additional confidence in the oil film technique is supplied by the good agreement of the behavior of the film profile with predictions from lubrication theory.
Hornung, Hans↗
Wing Leading Edge Joint Laminar Flow Tests
An F-104G aircraft at NASA's Dryden Flight Research Center has been equipped with a specially designed and instrumented test fixture to simulate surface imperfections of the type likely to be present near the leading edge on the wings of some laminar flow aircraft. The simulated imperfections consisted of five combinations of spanwise steps and gaps of various sizes. The unswept fixture yielded a pressure distribution similar to that of some laminar flow airfoils. The experiment was conducted at cruise conditions typical for business-jets and light transports: Mach numbers were in the range 0.5-0.8, and unit Reynolds numbers were 1.5-2.5 million per foot. Skin friction measurements indicated that laminar flow was often maintained for some distance downstream of the surface imperfections. Further work is needed to more precisely define transition location and to extend the experiments to swept-wing conditions and a broader range of imperfection geometries.
Drake, Aaron↗
Optimal and robust control of transition
Optimal and robust control theories are used to determine feedback control rules that effectively stabilize a linearly unstable flow in a plane channel. Wall transpiration (unsteady blowing/suction) with zero net mass flux is used as the control. Control algorithms are considered that depend both on full flowfield information and on estimates of that flowfield based on wall skin-friction measurements only. The development of these control algorithms accounts for modeling errors and measurement noise in a rigorous fashion; these disturbances are considered in both a structured (Gaussian) and unstructured ('worst case') sense. The performance of these algorithms is analyzed in terms of the eigenmodes of the resulting controlled systems, and the sensitivity of individual eigenmodes to both control and observation is quantified.
Bewley, T. R.↗
Wall Cooling Effects on Hypersonic Transitional/Turbulent Boundary Layers at High Reynolds Numbers
A 4 degree wedge was used to produce a thick turbulent boundary layer with an edge Mach number of 11. By using a two-dimensional model, the boundary layer was nearly free from upstream history effects associated with nozzle wall turbulent boundary layers. Heat-transfer distributions were used to define regions of laminar, transitional, and turbulent flow at several values of T(sub w)/T(sub t) for an edge unit Reynolds number of 0.47 x lot per cm. Pitot and total temperature profiles and skin-friction measurements were obtained at selected stations along the model. Turbulence parameters (mixing length/sigma and epsilon) were derived from the fully turbulent profiles and used to more completely define the "low Reynolds number" effect. Turbulent Prandtl number distributions are also presented.
Watson, Ralph D.↗
Summary of Research Report Cooperative Agreement
Several areas of work related to commercialization of technology developed at NASA Ames Research Center (ARC) are discussed in this report. The areas are: (1) perform a feasibility study to develop a software commercialization center is at ARC; (2) perform preliminary work for formation of joint development of sensor technology for telemedicine applications; (3) development of a discovery interview process and staff training to assist the commercialization of technology developed at Ames, specifically aimed at working with researchers; (4) develop partners to further develop and commercialize image compression technology developed at AMES; (5) assist efforts to commercialize a software technology which imparts the ability to establish relevance-based retrieval in the handling of large repositories of information; (6) explore the development of cryocooler technology using pulse tube refrigeration; (7) assess interest in commercialization of a new method of measuring skin friction drag on wind tunnel models using liquid crystal material; (8) attempt to incorporate emerging technologies in the infrastructure of natural hazards mitigation; and (9) forming a nonprofit organization, "The Bootstrap Alliance", whose mission is to promote the use of digital technologies for collaborative problem solving. The results of these initiatives are discussed.
Source record↗
Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds
Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.
John E Lamar↗
Introduction: Prediction of F-16XL Flight Flow Physics
This special section is the result of fruitful endeavors by an international group of researchers in industry, government laboratories and university-led efforts to improve the technology readiness level of their CFD solvers through comparisons with flight data collected on the F-16XL-1 aircraft at a variety of test conditions. These 1996 flight data were documented and detailed the flight-flow physics of this aircraft through surface tufts and pressures, boundary-layer rakes and skin-friction measurements. The flight project was called the Cranked Wing Aerodynamics Project (CAWAP), due to its leading-edge sweep crank (70 degrees inboard, 50 degrees outboard), and served as a basis for the International comparisons to be made, called CAWAPI. This highly focused effort was one of two vortical flow studies facilitated by the NATO Research and Technology Organization through its Applied Vehicle Panel with a title of Understanding and Modeling Vortical Flows to Improve the Technology Readiness Level for Military Aircraft. It was given a task group number of AVT-113 and had an official start date of Spring 2003. The companion part of this task group dealt with fundamentals of vortical flow from both an experimental and numerical perspective on an analytically describable 65 degree delta-wing model for which much surface pressure data had already been measured at NASA Langley Research Center at a variety of Mach and Reynolds numbers and is called the Vortex Flow Experiment - 2 (VFE-2). These two parts or facets helped one another in understanding the predictions and data that had been or were being collected.
Lamar, John E.↗
Hot-Film and Hot-Wire Anemometry for a Boundary Layer Active Flow Control Test
Unsteady active flow control (AFC) has been used experimentally for many years to minimize bluff-body drag. This technology could significantly improve performance of rotorcraft by cleaning up flow separation. It is important, then, that new actuator technologies be studied for application to future vehicles. A boundary layer wind tunnel was constructed with a 1ft-x-3ft test section and unsteady measurement instrumentation to study how AFC manipulates the boundary layer to overcome adverse pressure gradients and flow separation. This unsteady flow control research requires unsteady measurement methods. In order to measure the boundary layer characteristics, both hot-wire and hot-film Constant Temperature Anemometry is used. A hot-wire probe is mounted in the flow to measure velocity while a hot-film array lays on the test surface to measure skin friction. Hot-film sensors are connected to an anemometer, a Wheatstone bridge circuit with an output that corresponds to the dynamic flow response. From this output, the time varying flow field, turbulence, and flow reversal can be characterized. Tuning the anemometers requires a fan test on the hot-film sensors to adjust each output. This is a delicate process as several variables drastically affect the data, including control resistance, signal input, trim, and gain settings.
Control Test↗
Measurement and visualization of skin friction on the leeside of delta wings in supersonic flow
The reported investigation was conducted with the aid of the NASA Ames High Reynolds Number Wind Tunnel. The flow conditions considered involved free stream Mach numbers of 2 and 3, and a Reynolds number of approximately 10,000,000 p. m. The employed model was a delta wing with an angle of 70 deg. A method reported by Tanner (1977) was used for the measurement of the skin friction. This method involves the use of a laser interferometer to determine the change in the thickness of an oil film. The procedure can also be employed for a visualization of skin friction in the form of interferograms. The investigation shows that skin friction measurements can provide a significant contribution to a physical understanding of the flow processes at the delta wing.
Szodruch, J.↗