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Report describes relatively inexpensive apparatus that creates simulated precipitation from drizzle to heavy rain in flowing air. Small, positive-displacement pump and water-injecting device positioned at low-airspeed end of converging section of wind tunnel 10 in. in diameter. Drops injected by array entrained in flow of air as it accelerates toward narrower outlet, 15 in. downstream. Outlet 5 in. in diameter.
One of the fundamental problems in flight dynamics is the formulation of aerodynamic forces and moments acting on an aircraft in arbitrary motion. Classically, conventional stability derivatives are used for the representation of aerodynamic loads in the aircraft equations of motion. However, for modern aircraft with highly nonlinear and unsteady aerodynamic characteristics undergoing maneuvers at high angle of attack and/or angular rates the conventional stability derivative model is no longer valid. Attempts to formulate aerodynamic model equations with unsteady terms are based on several different wind tunnel techniques: for example, captive, wind tunnel single degree-of-freedom, and wind tunnel free-flying techniques. One of the most common techniques is forced oscillation testing. However, the forced oscillation testing method does not address the systematic and systematic correlation errors from the test apparatus that cause inconsistencies in the measured oscillatory stability derivatives. The primary objective of this study is to identify the possible sources and magnitude of systematic error in representative dynamic test apparatuses. Sensitivities of the longitudinal stability derivatives to systematic errors are computed, using a high fidelity simulation of a forced oscillation test rig, and assessed using both Design of Experiments and Monte Carlo methods.
Given here is a description of the design of a scientific recording wind tunnel balance. It was decided that the most satisfactory arrangement would be a rigid ring completely surrounding the tunnel or wind stream, so that the model could be supported from it by wires or any arrangement of spindles. The forces and moments acting on this ring can then be recorded by suitable weighing apparatus. The methods available for recording forces on the arms are explained. The proposed type of balance will support the model rigidly in a variety of ways, will make a complete test without attention, and will plot the results so that all computations are avoided.
Two wind tunnel techniques for determining part of the aerodynamic information required to describe the dynamic bahavior of various types of vehicles in flight are described. Force and moment measurements are determined with a rotary-balance apparatus in a coning motion and with a Magnus balance in a high-speed spinning motion. Coning motion is pertinent to both aircraft and missiles, and spinning is important for spin stabilized missiles. Basic principles of both techniques are described, and specific examples of each type of apparatus are presented. Typical experimental results are also discussed.
A tunnel ultra-light (or TUL) is a duct composed of acoustically transparent cloth designed to transform an open-jet wind tunnel into a closed-jet wind tunnel. This concept is of interest (a priori) for anechoic wind tunnels because it improves the aerodynamic quality without hindering the measurement of sound in the far field. A full scale device designed for the 3 m diameter test section of CEPRA 19 was described. The apparatus installation did not develop any significant problems, and the mechanical support turned out to be excellent. Aerodynamic and acoustic tests are discussed. Certain imperfections in the installation as tested - instabilities above 25 m/s and acceptable cloth transmission up to 4kHz were revealed. The system as tested could eventually be used in certain applications, for example, in ground based transport. However, the concept of TUL must be developed further to arrive at a reliable mechanism for use in a large number of applications.
A two-dimensional oscillating airfoil test apparatus is presented as a method of measuring unsteady aerodynamic forces on an airfoil or rotor blade section. The oscillating airfoil test rig, which was built for use in NASA Ames Research Center's 11x11-foot Transonic Wind Tunnel (speed range M = 0.4 - 1.4), allows determination of unsteady loadings and detailed pressure distributions on representative airfoil sections undergoing simulated pitching and flapping motions. The design details of the motion generating system and supporting structure are presented.
A method for obtaining two-dimensional aerodynamic force coefficients at low Reynolds numbers using a three-component external platform balance is presented. Regardless of method, however, the importance of understanding the possible influence of the test facility and instrumentation on the final results cannot be overstated. There is an uncertainty in the ability of the facility to simulate a two-dimensional flow environment due to the confinement effect of the wind tunnel and the method used to mount the airfoil. Additionally, the ability of the instrumentation to accurately measure forces and pressures has an associated uncertainty. This paper focuses on efforts taken to understand the errors introduced by the techniques and apparatus used at the University of Notre Dame, and, the importance of making an earnest estimate of the uncertainty. Although quantitative estimates of facility induced errors are difficult to obtain, the uncertainty in measured results can be handled in a straightforward manner and provide the experimentalist, and others, with a basis to evaluate experimental results.
New optical method enables determination of sizes of approximately-micron-sized particles. Arrays of particles act as diffraction gratings. Simple and inexpensive apparatus replaces high-magnification microscopes in measurements of diameters of small polystyrene latex spheres. Device being used to determine monodispersity and diameters of polystyrene latex microspheres used as seeding material for laser-velocimetry flow measurements in NASA's wind tunnels. Characterization of particle sizes important in wide range of fields, including manufacture of plastics and ceramics, biochemistry, and medicine.
Recent research conducted at the NASA Langley Research Center on the effects of large amplitude pitching motions on the aerodynamic characteristics of modern fighter aircraft configurations is highlighted. Wind tunnel tests were conducted on simple flat-plate wings to gain understanding of the complex flow phenomena during unsteady motions at high angles of attack. Studies then progressed to a representative modern fighter configuration. Using a computer controlled dynamic apparatus, tests were conducted to investigate effects of pitch rate and motion time history and to determine the persistence of unsteady effects. Data were also obtained in sideslip and with control surface deflections to investigate dynamic effects on lateral stability and available control power. Force and moment data were obtained using a 6-component internal strain-gage balance. To aid in the interpretation of the results, flow visualization using a laser light-sheet system was also obtained. Results of these tests are discussed, along with their implications on the maneuverability of future advanced airplanes designed to operate in the highly dynamic, high angle-of-attack environment.
An infrared (IR) camera provides a way of examining temperature trends associated with simulated microgravity flame spread in the Narrow Channel Apparatus (NCA). The IR camera measures the surface temperature of solid poly methyl methacrylate (PMMA) fuel. These tests examine the forward conduction of heat ahead of the flame front in the non-thermally thin fuel.The NCA is a combustion wind tunnel that simulates a microgravity flame spread environment by employing a narrow gap between the fuel and ceiling of the device, limiting the effects of buoyancy. Test conditions of a 5 mm gap, mean opposed flow velocity of 15 cm/s, and fuel thickness of 3 mm are used.PMMA is selected as the fuel due to repeatability of test results, ease of computational modeling, and known combustion mechanics. Using specific lens and bandpass filter combinations the PMMA can be imaged as effectively opaque. The spectral emissivity for PMMA was calculated and incorporated into the calibration of the camera.Surface temperatures from the IR camera are compared to results from thermocouples embedded in the surface of the fuel. The IR camera results show that nontrivial forward conduction occurs during tests, and therefore must be included in computational models of the process.
A two-dimensional wind-tunnel investigation of the pressure distributions over several NACA 16-series airfoils with thicknesses of 4, 6, 9, and 12 percent of the chord and design lift coefficients of 0, 0.2, 1 and 0.5 has been conducted in the Langley airfoil test apparatus at transonic Mach numbers from 0.7 to 1.25. The tests ranged in Reynolds number from 2.4 x 10 (exp 6) to 2.8 x 10 (exp 6) and in angle of attack from -10 to 12 degrees. Chordwise pressure distributions and schlieren flow photographs are presented without analysis.
The design of viable, low-mass thermostructural concepts for future space transportation systems requires accurate definition of the localized heat flux, pressures, and flow-surface interaction parameters for complex flow regions of detailed areas of large vehicles, such as wing-elevon coves, thermal protection system tile gaps, and corrugated metallic surfaces. This paper discusses investigations recently conducted in and planned for two high-energy wind tunnels at the Langley Research Center - the 8-foot High-temperatures Tunnel and the Thermal Protection System Test Facility. The data obtained on large-to-full-scale vehicle components, comparisons of the experimental data with theory, and the large, reusable, generalized test apparatus available at or planned for these facilities are discussed.
Two dimensional airfoil testing was conducted at the Wichita State University Beech Wind Tunnel for a number of years. The instrumentation developed and adapted during this period of testing for determination of flow fields along with traversing mechanisms for these probes are discussed. In addition, some of the techniques used to account for interference effects associated with the apparatus used for this two dimensional testing are presented. The application of a minicomputer to the data reduction and presentation is discussed.
The investigation reported is related to studies of the noise generated by a lifting jet in forward motion such as in the case of a V/STOL aircraft. The investigation involved experiments in a subsonic wind tunnel. The two-microphone cross-correlation method used for measuring the acoustic intensity provides data concerning the noise directly radiated from the source, practically free from any reflection effects. The theoretical basis of the two-microphone method is discussed along with the experimental apparatus, the characteristics of the flow inside the jet, and the features of the acoustic field.
One of the important parameters in the analysis of sediment entrainment and transport is gravitational attraction. The availability of a laboratory in Earth orbit would afford an opportunity to conduct experiments in zero and variable gravity environments. Elimination of gravitational attraction as a factor in such experiments would enable other critical parameters (such as particle cohesion and aerodynamic forces) to be evaluated much more accurately. A Carousel Wind Tunnel (CWT) is proposed for use in conducting experiments concerning sediment particle entrainment and transport in a space station. In order to test the concept of this wind tunnel design a one third scale model CWT was constructed and calibrated. Experiments were conducted in the prototype to determine the feasibility of studying various aeolian processes and the results were compared with various numerical analysis. Several types of experiments appear to be feasible utilizing the proposed apparatus.
This report presents a discussion of the analysis and the results of sensitive micro-g accelerometry measurements on the Shuttle Orbiter in the rarefied flow flight regime, that is, from altitudes of 60 km to 160 km. A brief description of the experimental apparatus called High Resolution Accelerometer Package (HiRAP) is given along with a description of the ground processing to obtain the lift-to-drag ratio (L/D) from the flight measurements. The flight results of L/D from six missions are presented and discussed. Comparisons are made with preflight wind tunnel data in the near continuum viscous interaction regime and preflight analytical estimates in the rarefied transition regime. Excellent agreement is obtained between flight and wind tunnel data in the viscous regime, but differences are noted between flight and preflight predictions in the transition regime. Free molecule flow F/D data are also obtained on each mission. Comparisons of flight data with classical theory show that complete diffuse reflection is not an appropriate model for rarefied aerodynamic coefficient calculations, and, for the Shuttle, some specular fraction is required. Surface reflection laws are explored to define the sensitivity of the L/D measurements to gas-surface molecule interactions.
One of the important parameters in the analysis of sediment entrainment and transport is gravitational attraction. The availability of a laboratory in earth orbit would afford an opportunity to conduct experiments in zero and variable gravity environments. Elimination of gravitational attraction as a factor in such experiments would enable other critical parameters (such as particle cohesion and aerodynamic forces) to be evaluated much more accurately. A Carousel Wind Tunnel (CWT) is proposed for use in conducting experiments concerning sediment particle entrainment and transport in a space station. In order to test the concept of this wind tunnel design a one third scale model CWT was constructed and calibrated. Experiments were conducted in the prototype to determine the feasibility of studying various aeolian processes and the results were compared with various numerical analysis. Several types of experiments appear to be feasible utilizing the proposed apparatus.
Coefficient of convective transfer of heat at spot on surface of wind-tunnel model computed from measurements acquired by developmental laser-induced-heat-flux technique. Enables non-intrusive measurements of convective-heat-transfer coefficients at many points across surfaces of models in complicated, three-dimensional, high-speed flows. Measurement spot scanned across surface of model. Apparatus includes argon-ion laser, attenuator/beam splitter electronic shutter infrared camera, and subsystem.