Items 8 - 14 High Speed Wind Tunnel Testing Alan Pope - Free ebook download as PDF File .pdf), Text File .txt) or read book online for free. Book. Low speed wind tunnel testing I by Jewel B. Barlow, William H. Rae, Alan . in High-speed Wind l h n e l Testing by A. Pope and K.L. Goin (John Wiley & Sons. KENNETH GOIN PDF. This is why we suggest you to constantly see this page when you require such book High-Speed Wind Tunnel. Testing By Alan L. Pope, .

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Get this from a library! High-speed wind tunnel testing. [Alan Pope; Kennith L Goin]. Pressure ratio across tunnel determined by vacuum and ambient air . [2] A. Pope and K. Goin, High-Speed Wind Tunnel Testing. New York. The subject of high-speed wind tunnel testing is covered xi xii PREFACE in High- speed Wind l h n e l Testing by A. Pope and K.L. Goin (John Wiley & Sons, New.

The model must be accommodated inside the rhombus formed by the incident and reflected shocks, for proper measurements.

The shock tube is a very useful research tool for investi- gating not only the shock phenomena, but also the behaviour of the materials and objects when subjected to very high pressures and temperatures. A shock tube and its flow process are shown schematically. The diaphragm between the high and low pressure section is ruptured and the high-pressure 4 of 12 SKVA driver gas rushes into the driven section, setting up a shock wave that compresses and heats the driven gas.

The pressure variation through the shock tube and the instant of diaphragm rupture and at two short intervals. The wave diagram simply shows the position of the important waves as the function of time. Figure 5: Pressure and wave diagram for a shock tube. When the shock wave reaches the end of the driven low pressure tube, all of the driven gas will have been compressed and will have a velocity in the direction of a shock wave travel.

Upon striking the end of the tube shock is reflected and starts travelling back upstream. As passes through the driven gas and brings it to rest, additional compression and heating are accomplished. The heated and compressed gas sample at the end of shock tube will retain its state except for heat losses until the shock wave reflected from the end of the tube passes through the driver gas driven gas interface and sends a reflected wave back through the stagnant gas sample, or the rarefaction wave reflected from the end of the driver high-pressure section reaches the gas sample.

The high temperature gas samples that are generated make the shock tube useful for studies of the chemical physics problems of high-speed flight, such as dissociation and ionization. A schematic diagram of a shock tube tunnel, together with a wave diagram, is shown Figure 6: Schematic of shock tunnel and wave diagram As shown in figure, a shock tunnel includes a shock tube, a nozzle attached to the end of the driven section of the shock tube, and a diaphragm between the driven tube and the nozzle.

When the shock tube is fired and the generated shock reaches the end of the driven tube, the diaphragm at the nozzle entrance is ruptured. The shock is reflected at the end of driven tube and heated and compressed air behind the reflected shock is available for operation of the shock tunnel.

As the reflected shock travels back through the driven section, it travels only relatively short distance before striking the contact surface; it will be reflected back towards the end of the driven section. When the reflected and shock reaches the end of the driven section, it will result in a change in pressure and temperature of the gas adjacent to the end of the driven section.

If the change in the conditions of the driven gas is significant, the flow in the nozzle will be unsatisfactory and the useful time will be terminated. The stagnation pressure and temperature in shock tunnels are about MPa and K, respectively to provide test times about 6. It has a high-pressure driver section and a lower-pressure driven section with a diaphragm separating the two as shown.

A piston is placed in the driven section, adjacent to the diaphragm, so that when the di- aphragm ruptures, the piston is propelled through the driven tube, compressing the gas ahead of it. The piston used is so light that it can be accelerated to velocities significantly above the 6 of 12 SKVA speed of sound in the driven gas. This causes a shock wave to precede the piston through the driven tube and heat the gas. The shock wave will be reflected from the end of the driven tube to the piston, causing further heating of the gas.

The piston comes to rest with equal pressure on its two sides, and the heated and compressed driven gas ruptures a diaphragm and flows through the nozzle. Figure 7: A gun tunnel and its wave diagram As can be inferred, gun tunnels are limited in the maximum temperature that can be achieved by the piston design.

The maximum temperatures normally achieved are about K. Run times of an order of magnitude higher than the shock tunnels are possible in gun tunnels.

In general, the types of tests that can be carried out in gun tunnels are the same as those in the hotshot tunnels and the shock tunnels. Some of the well-established Ludwieg Tube facilities are the following. The facility has a test time of 80 milliseconds.

This tunnel has a 9. The indraft tunnel can operate at higher Mach numbers before heating is required to prevent the liquefaction of air during the expansion to high speeds. For a given cost, indraft tunnels are larger. Advantages of the blowdown tunnel over the indraft tunnel are: 1. It is possible to vary the Reynolds number widely at a particular Mach number. In some cases the value corresponding to full-scale flight may be reached. Cost is from lower than to less than one-fourth of that of an slightly indraft tunnel of equal Reynolds number.

Short-time burning tests are usually possible. Although we risk antagonizing friends who swear by indraft tunnels by saying this, few such tunnels are built without the stimulus of strong external factors such as gifts of equipment or free vacuum pumps. This assumption is perfectly satisfactory from an engineering standpoint up to Mach numbers of 0.

At Mach num'bers above 0. The reason a marked change in the character for the serious effects is of the air flow which begins as soon as the speed of sound is exceeded. These effects include 1 the localizing of the effects of a body into a zone of action and a zone of silence; 2 a reversal of the subsonic laws governing streamline flow; and 3 the formation of shock waves.

We shall describe these effects, discuss their relation to the flow over an airfoil, and then summarize the laws of flow that govern their application. The Zone of Action' and the Zone of Silence. This effect can be described by Fig. In Fig. At time zero, the particle rushing into the air sets up a disturbance that travels at the speed of sound a in all directions. At a time Ac later, the region affected by the initial disturbance is bounded by a sphere of radius a At.

Meanwhile, the particle moving at half the speed of sound has moved through a distance of only 0.

At time At, another disturbance is created. Between times At and 2 At, the sphere affected by the disturbance initiated at time zero expands to a radius of 2a At, while that iniated at time At is expanding to a radius of a At. The particle is now preceded by both waves.

Continuing in this manner in Fig.

Wind Tunnel Testing

This is typical of subsonic flow. The disturbance waves, which are pressure pulses, go out ahead of the body to prepare the air ahead to move aside. Similarly, pressures at the rear of a body or behind it can reach forward and contribute to the flow pattern.

Looking at Fig. In this case the disturbances coalesce at the bow of the particle and form a concentrated wave front. For this case of Mach 1.

High Speed Wind Tunnel

Thus, the region ahead of the particleis a zone of silence. For the case of the particle moving at a speed greater than the speed of sound, the disturbance propagation pattern is shown in Fig. In this case the radius of each disturbance wave increases at a rate less than the rate of movement of the particle. The region outside this cone is a zone of silence. As indicated in Fig. For air flowing through a duct at Mach numbers less than 1. The velocity of the air increases faster than the density decreases; a 10 per cent increase of velocity, say, yields an 8 per cent decrease in density.

In such a case the number of pounds per second flowing through each square foot of duct cross section increases with increasing Mach number.

Consequently, the duct area must decrease if it is to remain filled with flow. See also Ref. As the Mach number increases, a 10 per cent increase in velocity might yield, say, a 12 per cent decrease in density. In this case, the number of pounds per second flowing through each square foot of duct cross section decreases, so that the duct area accommodate the flow at increasing must be increased to Mach number.

Mach number of a supersonic Conversely, to decrease the flow, the duct area must be decreased, an action that would increase the Mach number of a subsonic flow.

In free air flows, we speak of stream tubes instead of ducts. Stream tubes are imaginary ducts in which constant mass is considered to flow. A primary tool of experimental aerodynamics is the wind tunnel. The proper and productive use of experimental investigations in general and wind tunnels inparticu- lar requires applications of aerodynamic theory and computational methods in the planning of facilities, the planning of experiments, and the interpretation of resulting data.

Those aspects of aerodynamics will be drawn upon heavily in the course of this book.

To answer the question posed above: The invention, use, and ongoing evolution of the wind tunnel has been, and is now, motivated first by awide interest inpractical problems in aerodynamics and second by the fact that theoretical and computational methods have not been, are not now, and will not inthe foreseeable future becapable of providing the full range of results needed to guide detailed design decisions for many practical problems.

The most successful attack on virtually any aerodynamic design problem will be based on application of a combination of results from experimental, theoretical, and computational methods appropriately combined and leavened by experience.

Included in those and other texts are discussions of flow similarity in which definitions of similar flows are given. This is a very important concept that leads to significant advantages in experimental work and in theoretical and computational work as well.

The most common approach to the concept of similarity is through dimensional analysis using some variation of the Buckingham Pi theorem.

A reduction in the number of independent parameters to be manipulated in an investigation is obtained based on the requirement of dimensional homogeneity for any equation expressing avalid relationship among physical variables.

Introductory treatments of dimensional analysis are given by Anderson? David and Nolle,7 and Baker, Westine, and Dodge. Such models are the norm rather than the exception, as becomes apparent when almost any specific wind tunnel program is being planned.

Although the application of dimensional analysis has been of great importance in studies in aerodynamics, that approach will not be elaborated at this point. The fundamental principles from which the equations used to model "low-speed" aerodynamic flows are derived are only three in number.

These are 1 mass is conserved, 2 force and motion are related by Newton's Second Law, and 3 energy exchanges are governed by the First Law of Thermodynamics. In addition to these threeprinciples, certain fluid properties and their variations withpressure and temper- ature must bedescribed mathematically with sufficient accuracy for each application.

The equations expressing thethree principles provide relationships among various quantities such as density, velocity, pressure, rate of strain, internal energy, and viscosity as they vary in space and time. The dependence for aparticular quantity, say velocity, is indicated as VCr, t where r is a three-component position vector and t is time. The details of the function expressing the space and time dependence arestrongly affected by the choice of reference frame while the physical phenomena cannot beaffected by thechoice of reference frame.

It isdesirable tochoose reference frames that lead to relatively simple forms for the functional descriptions of the various quantities. Two kinematic results of importance relate time derivatives from different per- spectives. One relation is between "Lagrangian" and "Eulerian" descriptions of the motion of particles.

The other relation is between the time derivatives of quantities when measurements are made from two reference frames that are moving relative to one another. The Lagrangian and Eulerian perspectives of motion of a field of particles are described in almost every book on aerodynamics.

High-speed wind tunnel testing

The Lagrangian perspective is based on theidea of "tagging" every particle and subsequently describing themotion of each particle as afunction of time with a space coordinate indicating the identity of the particle. The usual choice would be that the space coordinate indicates the position of the particle at time equal to zero.As passes through the driven gas and brings it to rest, additional compression and heating are accomplished. For example, understanding the impact of wind on high-rise buildings, factories, bridges, etc.

High-speed turbulence and vortices can be difficult to see directly, but strobe lights and film cameras or high-speed digital cameras can help to capture events that are a blur to the naked eye. Add a review and share your thoughts with other readers. It is true that computational capability has continued to improve at a substantial pace, but it has not come close to reaching a level sufficient to replace the need for experimental data in development projects.

Tufts can be used to gauge air flow patterns and flow separation. Multiple particle streams released from a grid of many nozzles can provide a dynamic three-dimensional shape of the airflow around a body.

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