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wind tunnel,apparatus for studying the interaction between a solid body and an airstream. A wind tunnel simulates the conditions of an aircraft in flight by causing a high-speed stream of air to flow past a model of the aircraft (or part of an aircraft) being tested. The model is mounted on wires so that lift and drag forces on it can be measured by measuring the tensions in the wire. The paths of the airstream around the model can also be studied by attaching tufts of wool (which align themselves with the wind direction) to various parts of the model, by injecting thin streams of smoke into the tunnel to render the airflow visible, or by using certain optical devices. Pressures on the model surface are measured through small flush openings in its surface. Forces exerted on the model may be determined from measurement of the airflow upstream and downstream of the model. In wind tunnels operating well below the speed of sound, the airstream is created by large motor-driven vanes. At velocities near or above the speed of sound, the airstream is created either by releasing highly compressed air from a tank at the upwind end of the tunnel or by allowing air to rush through the tunnel into a previously evacuated vacuum tank at its downwind end. Sometimes these methods are combined, especially for the production of hypersonic velocities, i.e., velocities at least five times as great as the speed of sound. The effect of wind on other vehicles, e.g., automobiles, and on stationary objects such as buildings and bridges may also be studied in wind tunnels. In many instances, wind tunnels have been rendered obsolete by computer modeling.
an apparatus that creates a stream of air or gas for the experimental investigation of phenomena which accompany flow around solid bodies. By means of the wind tunnel it is possible to determine the forces that develop during the flight of airplanes, helicopters, rockets, and space vehicles and in the motion of submarines in the submerged state—that is, the stability and maneuverability of these vehicles are studied. The optimal forms of airplanes, rockets, and space and submarine vessels—and also of automobiles and trains—are sought. The wind stresses and stresses from shock waves that act on buildings and installations, bridges, electrical transmission supports, smokestacks, and so on can be determined. In special wind tunnels it is possible to investigate the heating and heat shielding of rockets, space vehicles, and supersonic airplanes.
Experiments in wind tunnels are based on the principle of reversibility of motion, according to which the displacement of a body relative to air can be replaced by the motion of the air flowing over a stationary body. For the modeling of the motion of an object in quiescent air, it is necessary to create a uniform flow in the wind tunnel that has equal and parallel velocities (a uniform velocity field) and identical density and temperature at all points. The flow ordinarily studied in wind tunnels is that around models of a planned object or its parts, and the forces acting on it are determined. Here it is necessary to maintain conditions that guarantee the possibility of carrying over the results obtained for the model under laboratory conditions to the full-scale object. When these conditions are maintained, the aerodynamic coefficients for the model being investigated and the full-scale object are equal, which—when the aerodynamic coefficients in the wind tunnel have been determined—permits the calculation of the force that acts on the full-scale object (for example, on an airplane).
The prototype of a wind tunnel was built in 1897 by K. E. Tsiolkovskii, who used for his experiments the flow of air at the outlet of a centrifugal blower. In 1902, N. E. Zhukovskii built a wind tunnel in which an axial blower created an air current with a velocity of up to 9 m/sec. The first open-circuit wind tunnels were created by T. Stanton in the National Physical Laboratory in London in 1903 and by N. E. Zhukovskii in Moscow in 1906; the first closed wind tunnels were created by L. Prandtl in Gottingen in the period from 1907 to 1909 and by T. Stanton in 1910. The first wind tunnel with a free stream in the working section was built by G. Eiffel in Paris in 1909. The further development of the wind tunnel was principally in the direction of enlarging the dimensions and increasing the flow velocity in the working section (where the model is located), which is one of the basic characteristics of the wind tunnel.
The supersonic wind tunnel, in the working section of which the flow velocity exceeds the velocity of propagation of sound, emerged in connection with the development of artillery, jet aviation, and rocket technology. In the aerodynamics of high velocities, the flow velocity or the flight velocity of aircraft is characterized by the Mach number M = v/a (that is, the ratio of the flow velocity v to the velocity of sound a). In correspondence with the value of this number, wind tunnels are divided into two primary groups: subsonic, M < 1; and supersonic, M > 1.
Subsonic wind tunnels. The constant-action subsonic wind tunnel (see Figure 1) contains the working section, which usually has the form of a cylinder with a circular or rectangular (sometimes elliptical or polygonal) cross section. The working section of the wind tunnel can be closed or open (see Figures 2a and 2b), and if it is necessary to create a wind tunnel with an open working section in which the static pressure is not equal to that of the atmosphere, the stream in the working section is separated from the atmosphere by the so-called Eiffel chamber (altitude chamber). (See Figure 2c.) The model under investigation (see Figure 1) is bracketed to the wall of the working section of the wind tunnel or to the wind-tunnel balance. In front of the working section is located the nozzle, which creates a stream of gas with preset and constant cross-sectional velocity, density, and temperature (the alignment grid equalizes the velocity field). The diffuser reduces the velocity and correspondingly increases the pressure of the stream that flows out from the working section. The compressor (blower), which is driven by a power unit, compensates the energy losses of the stream; the guide vanes reduce the energy losses of the air, preventing the formation of eddies in the turning elbow; and the return channel permits the conservation of a considerable portion of the kinetic energy that is in the stream behind the diffuser. The radiator guarantees the constancy of the gas temperature in the working section of the wind tunnel. If the static pressure in some cross section of the wind-tunnel channel must be equalized to that of the atmosphere, a valve is installed in it.
The dimensions of subsonic wind tunnels vary from large wind tunnels for the investigation of full-scale objects (for example, two-motor airplanes) to miniature table devices.
The wind tunnel of which a schematic diagram is given in Figure 1 refers to the so-called closed wind tunnel. There are also open wind tunnels, in which the gas for the nozzle is
supplied from the atmosphere or from special vessels. A significant feature of subsonic wind tunnels is the possibility of varying the velocity of the gas in the working section by means of varying the pressure differential.
According to the similarity theory, in order for the aerodynamic coefficients in the model and in the object itself (airplane, rocket, and such) to be equal it is necessary to have—in addition to geometric similarity—identical values of the Mach number M and the Reynolds number R in the wind tunnel and in flight (R = pvl/μ where p is the density of the medium, μ is the dynamic viscosity, and I is the characteristic dimension of the body). In order to provide these conditions, the power unit that produces the stream of gas in the wind tunnel must have sufficient power (the power of the power unit is proportional to the Mach number and the square of the Reynolds number and inversely proportional to the static pressure in the working section p8).
Supersonic wind tunnels. The layouts of supersonic and subsonic wind tunnels are analogous in general outline (see Figures 1 and 3). The so-called Laval nozzle, which is a converging-diverging channel, is utilized to obtain a supersonic gas velocity in the working section of the wind tunnel. In the converging region the flow velocity increases, and in the narrowest region of the nozzle it attains the velocity of sound; in the expanding region of the nozzle the velocity becomes supersonic and increases up to a fixed value that corresponds to the Mach number in the working section. A specific contour of the nozzle corresponds to each Mach number. Therefore, for the variation of the Mach number in the working section of supersonic wind tunnels, replaceable or adjustable nozzles, which permit variation of the nozzle’s form, are used.
In the diffuser of the supersonic wind tunnel the velocity of the gas must be reduced and the pressure and density increased; therefore the diffuser, like the nozzle, is made in the form of a converging-diverging channel. In the converging portion the supersonic flow velocity is reduced, and in a certain cross section a compression (shock) wave develops, after which the velocity becomes subsonic. For further slowing of the flow, the contour of the tunnel begins to expand, as in the ordinary subsonic diffuser. In order to reduce losses, the diffusers of supersonic wind tunnels are often made with an adjustable contour that permits the variation of the minimal cross section of the diffuser in the start-up procedure.
In the supersonic wind tunnel the energy losses in shock waves that develop in the diffuser are considerably larger than the frictional and vortex formation losses. In addition, there are considerably greater losses in flow around the model itself; to compensate for these losses, therefore, supersonic wind tunnels have multistage compressors and stronger power units than subsonic wind tunnels.
In the supersonic nozzle the temperature T and pressure p decrease in proportion to the increase in air velocity; in addition, the relative humidity of the air (which ordinarily contains water vapor) increases, and for M ≈ 1.2 there is vapor condensation accompanied by the formation of shock waves—that is, condensation jumps—which significantly disturbs the uniformity of the velocity and pressure fields in the working section of the wind tunnel. For the prevention of condensation jumps, the moisture from the air circulating in the wind tunnel is removed by special driers.
One of the basic advantages of supersonic wind tunnels realized according to the layout of Figure 3 is the possibility of carrying out experiments of considerable duration. However, this advantage is not decisive for many problems of aerodynamics. These wind tunnels are characterized by such shortcomings as the necessity of having energy units of great power and by difficulties that develop for M > 4 as a consequence of the rapid increase in the required compression ratio of the compressor. Therefore there has been widespread use of so-called compressed-air wind tunnels, in which high-pressure containers holding gas at a pressure of 100 meganewtons per sq m (MN/m2), or 1,000 kilograms-force per sq cm (kgf/cm2), are installed to create a pressure differential in front of the nozzle; behind the diffuser, vacuum vessels (gas tanks) evacuated to an absolute pressure of 100 to 0.1 N/m2 (10−3 to 10−6 kgf/cm2) or a system of ejector pumps (Figure 4) are installed.
One of the primary features of the wind tunnels of high Mach number (M > 5) is the necessity of preheating the air in order to avoid its condensation as a result of the temperature reduction with the increase of the Mach number. In contrast to water vapor, air is condensed without appreciable supercooling. The condensation of air considerably varies the parameters of the stream that is flowing out of the nozzle and makes it practically useless for aerodynamic experimentation. Therefore, wind tunnels with high Mach numbers have air preheaters. The temperature T0 to which it is necessary to preheat the air is higher the greater the Mach number in the working section of the wind tunnel and the pressure in front of the nozzle p0. For example, for the prevention of the condensation of air in the wind tunnel with M ≈ 10 and p0 = 5MN/m2 (50 kgf/cm2) it is necessary to preheat the air to an absolute temperature T0 = 1000° K.
Technology is developing in the direction of further increase of flight velocities. The descending cosmic devices Vostok and Voskhod enter the atmosphere of the earth with first cosmic velocity v, cos ≈ 8 km/sec (that is, M > 20). Space vehicles returning to the earth from the moon and other planets will be entering the atmosphere with second cosmic velocity v2cos ≥ 11km/sec (M > 30). At such flight velocities the temperature of the gas behind the shock wave that develops in front of the flying body exceeds 10,000° K, the molecules of nitrogen and oxygen dissociate (decompose into atoms), and ionization of atoms becomes substantial. It is necessary to study the effect of these processes on the forces that develop in the flow around a body and the thermal fluxes that act on its surface. For this it is necessary to obtain in the wind tunnel not only the full-scale M and R but also the appropriate temperature T0. This has led to the creation of new types of wind tunnels that operate with gas heated to high temperatures considerably exceeding the temperature necessary for preventing the condensation of air for a given M number. Among this group of devices are shock tubes, pulse tubes, and electric arc tunnels.
SHOCK TUBE. A shock tube (see Figure 5a) is a multistage cylindrical tunnel that consists of two sections—high pressure and low pressure—which are separated by a membrane. In the high-pressure section there is “propelling” gas (usually He or H), which is heated to a high temperature and compressed to pressure p1. The low-pressure section is filled by the working gas (air) at a low pressure p2. This state, which precedes the start-up of the wind tunnel, corresponds to time t0 in Figure 5b. After the breaking of the membrane, the shock wave begins to move through the working gas, compressing it to pressure p and raising the temperature. Behind the shock wave comes the contact surface, which moves with lower velocity and which separates the propelling and working gases (time t1). The pressure and temperature of the working gas in the space between the shock wave and contact surface are constant. Subsequently, the shock wave passes through the nozzle and the working section of the wind tunnel into the tank, and in the working section a supersonic flow with pressure p4 (time t2) is established.
The study of the flow of gas around the model begins at the moment that the shock wave passes through the cross section in which the model is located, and concludes when the contact surface passes through this cross section. Since the velocity of motion of the shock wave in the tube is greater than the velocity of the contact surface, it is clear that the duration of the experiment in the shock tube is greater the longer the length of the “acceleration” tunnel. In existing shock tubes this length reaches 200–300 m.
The type of shock tube examined affords the possibility of obtaining a temperature of approximately 8000° K for operating times on the order of a millisecond. Using shock tubes with several membranes, it is possible to obtain temperatures up to 18,000° K.
ELECTRIC ARC WIND TUNNELS. The solution of many aerodynamic problems may be carried out at lower temperatures, but considerable experiment time is required—for example, in the investigation of aerodynamic heating or heat-shielding insulation.
In electric arc wind tunnels (see Figure 6), the air that is being supplied to the precompression chamber of the nozzle is heated in an electric arc to a temperature of ∼ 6000° K. The arc, which is formed in a ring-shaped channel between the cooled surfaces of the central electrode and the chamber, is rotated with high frequency by a magnetic field created by an inductive coil (the rotation of the arc discharge is necessary
to decrease the erosion of the electrodes). A wind tunnel of this type permits the attainment of Mach numbers up to 20 for several seconds. However, the pressure in the precompression chamber usually does not exceed 10MN/m2(100kgf/cm2).
PULSE SHOCK TUBES. Higher pressures in the precompression chamber (∼ 60 MN/m2 [600 kgf/cm2]), and, correspondingly, higher Mach numbers, can be obtained in so-called pulse shock tubes, in which the heating of the gas is accomplished by the spark discharge of a set of high-voltage condensers. The temperature in the precompression chamber of a pulse shock tube is 6000° K, and the operating time is several dozen milliseconds.
The drawbacks of a device of this type are the contamination of the stream by the products of electrode and nozzle erosion and the fluctuation of gas pressure and temperature in the experimental procedure.
REFERENCESPankhurst, R., and D. Holder. Tekhnika eksperimenta v areo-dinamicheskikh trubakh. Moscow, 1955. (Translated from English.)
Zaks, N. A. Osnovy eksperimental’noi aerodinamiki, 2nd ed. Moscow, 1953.
Hilton, W. F. Aerodinamika bol’shikh skorostei. Moscow, 1955. (Translated from English.)
Sovremennaia tekhnika aerodinamicheskikh issledovanii pri giperz-vukovykh skorostiakh. Edited by A. Krill. Moscow, 1965. (Translated from English.)
Issledovanie giperzvukovykh techenii. Edited by F. R. Riddell. Moscow, 1965. (Translated from English.)
M. IA. IUDELOVICH
wind tunnel[′win ‚tən·əl]
A duct in which the effects of airflow past objects can be determined. The steady-state forces on a body held still in moving air are the same as those when the body moves through still air, given the same body shape, speed, and air properties. Scaling laws permit the use of models rather than full-scale objects, such as aircraft or automobiles. Models are less costly and may be modified more easily, and conditions may be simulated in the wind tunnel that would be impossible or dangerous in full scale.
Most data are secured from wind tunnels through measurement of forces and moments, surface pressures, changes produced in the airstream by the model, local temperatures, and motions of dynamically scaled models, and by visual studies.
A balance system separates and measures the six components of the total force. The three forces taken parallel and perpendicular to a flight path are lift, drag, and side force. The three moments about these axes are yawing moment, rolling moment, and pitching moment, respectively.
Surface pressures are measured by connecting orifices flush with the model surface to pressure-measuring devices. Local air load, total surface load, moment about a control surface hinge line, boundary-layer characteristics, and local Mach number may be obtained from pressure data.
Measurements of stream changes produced by the model may be interpreted in terms of forces and moments on the model. In two-dimensional tunnels, where an aircraft model spans the tunnel, it is possible to determine the lift and center of pressure by measuring the pressure changes on the floor and ceiling of the tunnel. The parasite drag of a wing section may be determined by measuring the total pressure of the air which has passed over the model and calculating its loss of momentum.
Measurements of surface temperatures indicate the rate of heat transfer or define the amount of cooling that may be necessary.
In elastically and dynamically scaled models used for flutter testing, measurements of amplitude and frequency of motion are made by using accelerometers and strain gages in the structure. In free-flight models, such as bomb or missile drop tests, data are frequently obtained photographically.
At low speeds, smoke and tufts are often used to show flow direction. A mixture of lampblack and kerosine painted on the model shows the surface streamlines. A suspension of talcum powder and a detergent in water is also used.
For aircraft at velocities near or above the speed of sound, some flow features may be made visible by optical devices.
The V/STOL wind tunnel is a newer development of low-speed wind tunnels having a large very-low-speed section to permit testing of aircraft designed for vertical or short takeoff and landing (V/STOL) while operating in the region between vertical flight and cruising flight.