A fundamental investigation of the flow in the endwall corner region of a linear compressor cascade wind tunnel and the effect of boundary layer blowing in this region was conducted using blade surface pressure tap measurements and five - hole prism probe measurements taken downstream of the cascade. The results are presented as a series of velocity vector plots, loss contour plots, and pitchwise mass - averaged loss coefficient plots. The angle of attack test range was from 5 to 29 degrees. For the corner region boundary layer blowing investigations, two slots were machined into the ends of a set of cascade blades, and an external air source was used as the blade slot jet air source. Tests were done for 19, 21, and 23 degrees angle of attack. The main effect of corner boundary layer blowing was a significant reduction in total pressure losses in the region along the blade span between the exterior portion of the corner boundary layer flow and the blade profile boundary layer flow. / Master of Science

No description available.

The purpose of this paper was to design a simplified wind tunnel force balance for use in elementary aerodynamics/wind tunnel laboratory courses. The applied loads and moments were determined to be as follows: The design lift force is ± 20 lbs, the design drag force is ± 10 pounds, the design pitching moment is ± 20 inch-pounds. The force balance output accuracy was arbitrarily set at ± 5% since this would be sufficient for preliminary student work. The results of this work are as follows: The force balance is fabricated from a single bar of 2024 aluminum, machines and bent into an "L" shape, the applied forces and moments are sensed by strain gages bonded to machined surfaces on the bar, the output of the strain indicator equipment is transformed into uncorrected forces and moments by a system of three equations, the uncorrected forces and moments are transformed into true forces and moments by a system of force balance interaction equations, the design values for lift, drag, and pitching moment remain the same as originally proposed, the output error in lift is determined to be ± 3.5%, the output error in drag and pitching moment are determined to be ± 10%. The prototype has been in use for several months and its operation has been completely satisfactory.

The primary object of the investigation reported in this paper was to obtain information that would aid in the design of a more efficient diffuser for FTU's tunnel, and thus increase the run time. Presently FTU's four inch supersonic wind tunnel uses a constant area, normal shock, diffuser to recover the fluid pressure after the test section. Also, FTU's tunnel is of the intermittent blowdown type, which provides only a relatively short test time before the storage pressure decreases to a limiting value at which flow in the test section ceases to be supersonic. The use of a constant area diffuser and normal shock pressure recovery has the disadvantage of always entailing a large loss in stagnation pressure. These losses increase as the test section Mach number increases. Since a diffuser employing a system of oblique shocks should have a better pressure recovery than one with a single normal shock, efforts were made to improve FTU's wind tunnel along these lines. Variable area diffusers whose throats can be closed after flow has been established were of interest in this report because of their higher pressure recovery. The maximum run time of FTU's wind tunnel is limited by the overall operating pressure ratio required to maintain supersonic flow in the test section area. If one can reduce the losses in the tunnel, the operating pressure ratio can be reduced. The reduction in operation pressure can result in an increase in run time. In FTU's tunnel, the majority of losses occurs in the second throat area or the supersonic diffuser. Tunnel run time improvement may be required to conduct heat transfer studies or to conduct force, moment and pressure tests. The results of the one-dimensional analyses of a variable geometry supersonic diffuser are very promising in that they show a longer run time can be obtained for FTU's tunnel. By using a variable geometry diffuser, an intermittent blowdown wind tunnel run time can be increased two to three times that of a constant are diffuser at high Mach numbers. At the design Mach number of 4.0, the theoretical run time can be increased 321 percent over the run time of a constant area diffuser. The references cited made it possible to geometrically design a relatively simple, yet efficient contractible wall (convergent-constant area-divergent) type diffuser. Three flagellates were chosen to form the side walls of the adjustable diffuser. The length of the plates were a compromise between mechanical construction requirements and the need to keep the wall convergent angle relatively small for the Mach number range of FTU's tunnel and to minimize energy losses. The first adjustable diffuser plate has an overall length of 14.5 inches. The angle of convergent for design was chosen to be 7 degrees at the design Mach number of 4.0. The second diffuser plate that forms the constant area passage has an overall length of 12 inches. The third diffuser plate that forms the divergent section has an overall length of 13.5 inches.

The purpose of this thesis project was to design, construct and evaluate the performance of a high temperature continuous-run electric arc wind tunnel. A pilot model of such a facility was designed assuming that equilibrium air was the working gas. The pilot model facility was constructed and consisted of the following components; arc chamber, stagnation chamber, nozzle section, test and diffusor sections. In the arc chamber, the air passes through the positive column of an electric arc there-by raising its stagnation temperature before entering the stagnation chamber. Also included in the design and construction were water cooling and waste disposal systems, air supply and vacuum systems, and electric arc power supply system and control. An examination of tests performed in the electric-arc facility showed that a low density supersonic flow with a stagnation temperature of approximately 10,000 F could be produced. The power level for this flow was 36 kw; however, with an expected increase of power to 72 kw the stagnation temperature should be raised to 15,000 F. Since no valid technique for measuring temperatures of this magnitude has been perfected to the author's knowledge, these temperatures were calculated according to a method as outlined in the present thesis. The present facility will present an opportunity for study in such topic areas as (1) Aerodynamic Ablation, (2) Magnetoaerodynamic studies and (3) Qualitative studies of chemically reacting gas flows. / Master of Science

The theory used to develop curved flow as a method of obtaining dynamic stability derivatives is presented including an analysis of the flows involved in the curved flow wind tunnel and in curved flight. Equations for the forces and moments for each of these flows are presented and then used to develop equations for the corrections to the forces and moments obtained in curved flow wind tunnel tests. An example of the physical setup and of the testing procedure for curved flow testing is also presented with some of the results from a typical test. The principles involved in several other methods of testing that are also used to obtain the dynamic stability derivatives are discussed so that a comparison may be made with the curved flow method. / Master of Science

Various sources of error can cause discrepancies among flight test results, experimental measurements and numerical predictions in the transonic regime. For unsteady flow, the effects of wind tunnel walls or a finite computational domain are the least understood and perhaps the most important. Although various techniques can be used in steady wind tunnel testing to minimize wall reflections, e.g., using slotted walls with ventilation, wind tunnel wall effects remain in unsteady wind tunnel testing even when they have been essentially eliminated from the steady flow. Even when the walls are ten chord lengths or more from the airfoil being tested, they can have a substantial effect on the unsteady aerodynamic response of the airfoil. In this study we compare numerical computations of two- and three-dimensional unsteady transonic flow with one another, and with experimental measurements, to isolate and examine the effects of tunnel walls. An extension of the time-linearized code developed by Fung, Yu and Seebass (1978) is used to obtain numerical results in two dimensions for comparison with one another and with the experimental measurments of Davis and Malcolm (1980). The steady flow which is perturbed by small unsteady airfoil motions is found numerically by specifying the pressure distribution rather than the airfoil coordinates using the procedure provided by Fung and Chung (1982). This provides results that are nearly free from effects caused by the small perturbation approximation; it also simulates the viscous effects present in the experimental measurements. A similar algorithm, developed especially for this study, is used for the related investigations in three dimensions. Different wall conditions are simulated numerically. Aside from a shift of frequency due to nonlinear effects, our numerical predictions of resonance conditions in two dimensions agree very well with those of linear acoustic theory. A substantial discrepancy between unconfined computations and wind tunnel experiments is observed in the low frequency range. This discrepancy highlights the importance of wall interference and wind tunnel measurements of unsteady transonic flows and delineates the conditions required to suppress them satisfactorily.

A wing with a Wortmann FX-63-137-ESM airfoil section has been used to study some unique problems encountered in wing aerodynamics in the range of Reynolds numbers between 50,000 and 500,000. The wind-tunnel testing conducted in the 6'x 6' Stability tunnel included strain-gauge data, pressure data, and flow-visualization studies. The laminar separation bubble which frequently occurs on the upper surface of the wing is found to dominate its performance and gives rise to a hysteresis loop for lift and drag. Changes in airfoil performance due to positive flap or control surface deflections resemble changes witnessed at higher Reynolds numbers. Negative deflections are seen to considerably change the stall behavior and the flow over the airfoil. This is due to the considerably greater effect on the separation bubble for negative flap deflections. The structure and mechanism of the laminar separation bubble can also be altered by the introduction of selected acoustic disturbance and increased free-stream turbulence. The wind-tunnel test-section environment is, therefore, capable of considerably altering wing performance in this regime. / Ph. D. / incomplete_metadata

Master of Science

Call number: LD2668 .T4 1962 H37