Classification of wind tunnel experiments

The main types of wind tunnel experiments in fluid mechanics are force measurement experiment, pressure measurement experiment, heat transfer experiment, dynamic model experiment and flow observation experiment. Force measurement and pressure measurement experiments are used to measure the aerodynamic force and surface pressure distribution acting on the model or model components (such as wings in aircraft models), and are mostly used to provide aerodynamic characteristics data for aircraft design. The heat transfer experiment is mainly used to study the aerodynamic heating phenomenon on supersonic or hypersonic aircraft. Dynamic model experiments include flutter, buffeting and dynamic stability experiments. It is required that the model can not only satisfy the geometric similarity, but also simulate the structural stiffness, mass distribution and deformation of real objects. Flow observation experiment is widely used to study the basic phenomenon and mechanism of flow. The application of high-speed computer in the above-mentioned wind tunnel experiment greatly improves the automation, high efficiency and high precision of the experiment. Force measurement experiment is a wind tunnel experiment that uses wind tunnel balance to measure the aerodynamic force and torque acting on the model (see wind tunnel test instrument). It is one of the most important experimental items in wind tunnel experiment. Force measurement experiments mainly include: longitudinal and transverse force measurement experiments of the whole model and components, jet experiments, static aeroelasticity experiments, external force measurement and release trajectory experiments.

The longitudinal and transverse force measurement experiments of the whole model and parts are experiments to measure the force along three mutually perpendicular axes and the moment around three axes on the model, in which the experiment without sliding measurement is a longitudinal experiment and the experiment with sliding measurement is a transverse experiment. The model is supported by the abdomen strut or the tail strut in the wind tunnel (Figure 1 and Figure 2).

In order to study the contribution and interference of each component, in addition to the disassembly experiment of the whole machine and components, a more accurate method is to install multiple scales in the model and measure the aerodynamic force of the whole machine and components at the same time. For the aircraft with symmetrical surface, under the condition of flow symmetry, the wind tunnel wall or reflector can be used as the symmetrical surface, and half of the model can be used for experiments. This kind of experiment is called semi-model experiment, and its advantages are that the model can be made larger, the Reynolds number can be higher, there is no interference from the tail strut, and it is convenient and economical to manufacture. The disadvantage is that there are boundary layers and gaps in the tunnel wall, so only longitudinal experiments can be carried out. Jet experiment is an experiment to measure the influence of jet from aircraft engine on aerodynamic characteristics of aircraft body. It is difficult to simulate jet accurately in wind tunnel. In addition to simulating free-flow Mach number Mα∞, specific heat ratio γ and nozzle geometry, other similar parameters such as static pressure ratio pj/p∞, outlet Mach number Mαj, specific heat ratio γ 1, product ratio of universal gas constant and thermodynamic temperature (RT)j/(RT)∞ should also be simulated. Usually, only some of these projects can be selectively simulated. For example, when the nozzle is at the bottom of the plane, cold air can be used to simulate the jet. When the nozzle is located upstream of the aircraft bottom, γ 1 and (RT)j/(RT)∞ should also be simulated. It is suitable to simulate the jet flow of rocket engine with scale rocket engine. The key of jet experiment is to develop a high-precision balance, a bracket with little interference and a gas sealing system without force transmission.

Static aeroelasticity experiment is an experiment to measure the influence of model stiffness on aerodynamic characteristics. Usually, the models in wind tunnel experiments are made of metal with high strength and stiffness, but the stiffness of real aircraft is much lower than that of the model. Therefore, it is necessary to make an elastic model with metal as the skeleton and balsa wood or plastic as the filler, which can simulate the bending and torsion stiffness of aircraft components, put it in a wind tunnel to simulate flight conditions, measure the influence on the model stiffness, and correct the data of rigid body model experiments.

The experiment of external force measurement and release trajectory is an experiment to measure the aerodynamic force and release trajectory of external fuel tank, bomb or other objects of aircraft. Due to the limitation of the size of the wind tunnel, the external model in the wind tunnel is very small and it is difficult to measure. The early experiment was to design a special shop balance. The balance can be placed on the store model or its hanger for direct measurement. The trajectory of the store is to record the free release model with high-speed photography or multiple exposure technology. Fig. 3 is a photo of the store release trajectory taken by multiple exposures in a low-speed wind tunnel. This method is simple and intuitive, but it is difficult to design and adjust the model because of the need to simulate Froude number. Since the 1960s, a double-balance measuring system has been developed, in which the master model and stores are supported on their own scales. In the experiment, the aerodynamic forces of the store and the main engine are measured first, then input into the computer, and the next position of the store under aerodynamic forces is calculated through the motion equation and the given time interval, and then the store is manipulated to move to the position calculated before measurement. Until the desired trajectory is measured. At this time, all instantaneous aerodynamic forces of the main engine and the external store are also measured at the same time. This method does not require the models to be similar in dynamics, and the models can be used many times. At the same time, the device can also be used for other two-body experiments or trajectory measurement after stall at high angle of attack. The disadvantages are high precision and high manufacturing cost.

In addition to the above experiments, there are some special force measurement experiments, such as hinge torque measurement, friction measurement, intake resistance measurement, Magnus force and torque measurement (see Magnus effect) and so on. All these require specially designed scales.

Measurement of local pressure parameters of wind tunnel wall, model surface point and airflow midpoint in pressure measurement experiment. There is a total pressure p0 and a static pressure p∞ corresponding to each point of the flow field. Total pressure is the pressure that can be reached when the air flow is assumed to be isentropic and adiabatic, and the final speed drops to zero. Static pressure is the normal force between interacting layers in airflow. In an incompressible fluid, the difference between the total pressure and the static pressure, that is, the pressure increase (P0-P ∞) caused by the aerodynamic effect of the flow point, is called dynamic pressure or rapid pressure q∞. The measurement of airflow pressure is one of the most basic measurement items in aerodynamic experiments.

1738, Daniel I Bernoulli established the relationship between pressure and velocity in inviscid incompressible fluid, which is the later Bernoulli theorem. This theorem was later extended to compressible fluids. Because it is easy to measure the pressure of airflow, the velocity is often calculated by measuring the pressure of airflow in wind tunnel experiments.

The pressure pi at a point on the surface of an object (such as point I) is often expressed as the dimensionless pressure coefficient Cρii. If p∞ and q∞ respectively represent the static pressure and dynamic pressure of undisturbed airflow on the far front, then Cρii is the ratio of the residual pressure (pi-q∞) at that point to the dynamic pressure Q ∞.

The most common pressure measurement experiment in wind tunnel is the pressure distribution measurement on the model surface. Pressure taps are directly opened on the surface of the model. Through the experiment, we can understand the local flow characteristics and synthesize the overall aerodynamic characteristics. Commonly used are aircraft pressure measurement, automobile pressure measurement and building pressure measurement. The inlet pressure measurement experiment is to obtain the flow-total pressure recovery characteristics of the inlet through the pressure measurement hole on the surface of the inlet and the pressure measurement of the exhaust pipe in the pipeline. The velocity field, pressure field and direction field in wind tunnel flow field calibration are also measured by pressure measurement. In addition, boundary layer pressure measurement is also a regular experimental project. Sometimes, the drag of a binary object is calculated by measuring the wake pressure. Therefore, wind tunnel pressure measurement experiment has been widely used in engineering design and research.

A total pressure and static pressure detection tube and a pressure gauge or pressure sensor used to measure the total pressure and static pressure of the airflow in the wind tunnel. Figs. 4 and 5 show the structures of general full pressure pipes and static pressure pipes. Full pressure or static pressure exhaust pipe can obtain many pressure measurement data at the same time. But the interaction between tubes is very small. The pressure measuring hole on the model surface should be perpendicular to the local surface, and the hole edge should be smooth and burr-free. It is particularly important to choose the location of the static pressure hole on the static pressure detection tube, so that it should be minimally affected by the head of the static pressure tube and the support handle. The pressure transmission pipeline in pressure measuring equipment should not be too long, otherwise the pressure in the pipeline will take a long time to reach equilibrium. Under the condition that the airflow and the model are moving at relatively high speed, the experiment of measuring the aerodynamic heating of the model surface by the airflow flowing around the model is carried out. When the flight Mach number is greater than 3, the influence of aerodynamic heating on the shape, surface roughness and structure of the aircraft must be considered. The purpose of heat transfer experiment in wind tunnel is to provide reliable thermal environment data for aircraft thermal protection design. The experimental items include: heat flow experiments on smooth and rough surfaces, experiments on the influence of boundary layer transition and mass injection on heat flow, and heat flow experiments on separated flows such as steps, gaps, shock waves and boundary layers. In wind tunnel heat transfer experiments, thermal radiation is generally ignored and only convection heating is considered. The parameters to be simulated include Mach number, Reynolds number, wall temperature ratio, relative roughness (the ratio of roughness to boundary layer displacement thickness), mass injection rate, free turbulence and so on. Generally, hypersonic wind tunnel, pulse wind tunnel, shock wind tunnel, arc heater, low-density wind tunnel and ballistic target can all carry out heat transfer experiments, but the above parameters cannot be completely simulated. Therefore, it is necessary to comprehensively analyze the experimental data of different equipment. There are two methods for wind tunnel heat transfer experiment: one is thermal mapping technology to determine the distribution of heat flux, such as coating phase change materials on the model surface and recording the expansion process of isotherms with time; Another example is that the mixture of paint and powder phosphor is coated on the model surface, and the heat flux distribution is obtained by recording the brightness distribution of phosphor (the latter method has fast response and high sensitivity). Thermal map technology can provide rich pneumatic heating data, but its accuracy is low. The other is thermal measurement technology, which uses calorimeter to measure the heat of scattered points. Generally, under the assumption of one-dimensional heat conduction, the heat flux density is measured by measuring the rate of change of temperature with time. There are two kinds of calorimeters commonly used in hypersonic wind tunnels: ① Thin-walled calorimeter, which requires the model wall to be made very thin, so that the temperature of the inner and outer surfaces of the model is almost equal when heated, and a thermocouple is installed on the inner surface to measure the temperature change with time and calculate the heat flux density. (2) Gaden meter was put forward by R. Gaden in 1953, which is based on the relationship between the temperature gradient between the center and the edge of the heated element and the heat flux. Thin-walled calorimeter and Gardiner can not be used in pulse wind tunnel because it takes a long time to reach temperature equilibrium. In pulse wind tunnel, plug calorimeter and thin film resistance thermometer can be used for measurement. Plug calorimeter uses a calorimetric element to absorb the heat introduced into it, then measures the average temperature change rate of the element and calculates the surface heat flux.

Wind tunnel heat transfer experiment should not only properly solve the problems of model design, protection, cooling and signal transmission, but also study the simulation technology, reduce the size of the sensor, solve the stability problem of the sensor, and determine the influence of various uncertain factors on the accuracy of the experimental results. The experiments to determine the relative motion of the model to airflow and the change of aerodynamic force on the model with time include flutter experiment, buffeting experiment, dynamic stability experiment, control surface humming experiment, unsteady pressure measurement and so on.

Flutter experiment Flutter is a self-excited vibration produced by the absorption of energy from airflow under the joint action of aerodynamic force, structural elastic force and inertial force. Once it happens, it is likely to cause structural damage. The purpose of wind tunnel flutter test is to select a structural scheme that is beneficial to flutter prevention (see flutter test).

Buffeting experiment Buffeting is the vibration of aircraft structure caused by airflow separation. When flying at low speed and high angle of attack, the airflow separation on the lifting surface will produce buffeting, which is called lift buffeting. In transonic flight, the initial attack angle of buffeting is obviously reduced due to the induction of shock wave. In addition, there are non-lifting buffeting caused by airflow separation. Buffeting affects the structural strength and fatigue life of aircraft, which will make the weapon system and electronic instruments work abnormally and make passengers uncomfortable. The curve of the lift coefficient (see lift) corresponding to the initial angle of attack of buffeting with Mach number is called buffeting boundary. The higher the buffeting boundary, the lower the minimum flight speed of the aircraft, and the better the maneuverability and safety in flight. Buffeting test is to measure buffeting boundary and buffeting load. Buffeting boundary can be determined by root mean square moment method and trailing edge static pressure divergence method. The so-called root-mean-square bending moment method is a method of sticking a strain gauge on the wing root of the model, measuring the root-mean-square horizontal value proportional to the root bending moment at different angles of attack at a certain Mach number, and determining the angle of attack corresponding to the transition point where the horizontal value begins to increase sharply as the initial angle of attack of buffeting. The so-called trailing edge static pressure divergence method is a method to measure the trailing edge static pressure of airfoil by using the principle that the trailing edge static pressure increases rapidly after airflow separation. In addition to maintaining the aerodynamic similarity between the model and the real object, it is also required to simulate the first-order bending frequency. Buffeting experiment has strict requirements on wind tunnel noise level, turbulence degree and boundary layer state on model surface.

Experiment of measuring dynamic derivative with dynamic stability experiment. Dynamic derivative is the derivative of aerodynamic force and moment to the time change rate of motion parameters, such as the derivative of rolling moment mx to rolling angular velocity ωx, which usually plays a damping role, and is also called rolling damping derivative. The dynamic derivative experiment generally adopts a rigid model, and the degassing dynamics are similar. It also requires that the reduction frequency ωL/v is the same as the real thing, where ω is the vibration frequency; L is the characteristic length; V is the air speed. Generally, free vibration method or forced vibration method is used to measure dynamic derivative in wind tunnel. The free vibration method is to release the model after giving it a certain initial displacement, so that it can freely attenuate the vibration in the airflow, and determine the dynamic derivative according to the recorded model displacement time history. This method has simple equipment, but it is greatly influenced by external interference such as wind tunnel background noise and its accuracy is not high. Forced vibration method is to apply a sinusoidal excitation torque with a certain frequency to the model system. In this process, the phase difference between the excitation torque and the angular displacement of the model vibration is measured by measuring instruments, so as to determine the dynamic derivative. In addition, the dynamic derivative can be measured by the free flight of the wind tunnel model.

Control surface humming experiment Control surface humming is a one-degree-of-freedom unstable motion caused by the interaction of shock wave on the wing, boundary layer separation after wave and control surface deflection when the aircraft flies transonic. Control surface hum is sensitive to Mach number. Hum will reduce the control efficiency and even make the control ineffective, which will lead to the fatigue damage of the structure in serious cases. Through the buzzer test, the vibration characteristics of aircraft control surface can be determined, and the methods to eliminate vibration and determine stiffness index are provided. The buzzer experimental model consists of rigid main wing and control surface, and the stiffness of the control system can be simulated by spring leaf. The structural damping of the control system should be roughly equivalent to the physical object. In the experiment, the vibration waveform is measured by the strain measurement system, and the vibration intensity can also be measured by the root mean square level recorder.

Unsteady pressure measurement is a basic means to study unsteady aerodynamics. There are two measuring methods: one is to measure the unsteady pressure at multiple points at the same time with a miniature pressure sensor embedded in the model; The other is that many pressure pipes are placed in the model, and the unsteady pressure is measured through the pressure pipes, which are connected with the sensor through the scanning valve. Using the latter method, it is necessary to correct the dynamic transfer characteristics of the pipeline under the condition of blowing.

In dynamic experiments, wind tunnel background noise has a great influence on the accuracy of experimental results. Therefore, in addition to limiting the level of wind tunnel noise, it is necessary to reduce the influence of wind tunnel noise in experimental technology, such as using correlation filtering and overall averaging in data processing. Equipped with dynamic analysis equipment with fast Fourier transform ability, the dynamic experiment ability can be significantly improved and real-time analysis can be realized.

Flow observation experiment is an experimental method to change colorless and transparent airflow into visible airflow in wind tunnel by physical and chemical means. By using this technology, the physical image of gas flow can be directly observed with naked eyes or other auxiliary means, so as to deepen the understanding of gas flow mechanism and find out the problems existing in gas flow in time. The observation results can also be used to verify some theories and assumptions and help to establish a mathematical model of complex flow problems. This technology is the basic method of aerodynamic experiments.

There are many phenomena in nature that can show the flow of fluid. The movement of floating objects on the water surface often indicates the direction of water flow; The smoke produced in the fire shows the rising and spreading pattern of hot air. Scientific research with flow visualization technology in the laboratory began at the end of 19. 1883 O. Reynolds introduces a stream of dyeing water into the pipe flow, and judges whether the flow in the pipe is laminar or turbulent according to whether the dyeing water is a regular flow with clear color or a chaotic flow. In 1893, L. Mach observed the airflow and smoke around a vertically placed flat plate in the wind tunnel. With the development of wind tunnel and the progress of science and technology, there are more and more flow observation methods.

The observation methods of flow state in wind tunnel can be roughly divided into two categories: the first category is tracer method; The second category is optical methods.

Tracer method

It is to add colored liquid, smoke, silk thread, solid particles and other substances into the flow field, and observe the movement mode of additives with the fluid through photography or naked eyes. As long as the additive is small enough and its specific gravity is close to that of the flowing medium, the displayed additive motion diagram shows the movement of the air flow. This is an indirect display method, especially suitable for displaying stable flow. There are six commonly used methods: silk thread method, smoke flow method, oil flow method, sublimation method, steam screen method and liquid crystal display method;

(1) silk thread method is to stick silk thread, wool and other fibers on the observed model surface or the grid behind the model. The direction of air flow, the position of separation zone, and the position and direction of spatial vortex can all be identified by the movement of silk thread (rotation, shaking or inversion of silk thread). Fig. 6 is a diagram showing the airflow of the wing in the model experiment. Now it has been developed to use nylon thread thinner than silk thread, sometimes so thin that it is not clear to the naked eye. Nylon yarn is treated with fluorescent dye and then stuck to the model. This kind of silk thread is displayed under ultraviolet radiation and can be photographed. The glue line is very thin, which has no effect on the model and can be used for force measurement experiments. This method is called fluorescent silk thread method.

(2) The smoke flow method shows the flow pattern of gas around the model by using the smoke flow released by a special smoke pipe or model in the wind tunnel. This is a good observation method. Many smoke wind tunnels have been built around the world. Usually, smoke generated by heating nonflammable mineral oil by wires outside the wind tunnel is introduced into the wind tunnel; There are also oiled stainless steel or tungsten wires placed in front of the model, and the tungsten wires are electrically heated to produce fine smoke during the experiment. In order to keep the flue gas flowing smoothly, it is necessary to install an anti-turbulence net at the entrance of the contraction section, stable section or wind tunnel with large contraction ratio, and make the tunnel wall with good vibration absorption materials to keep the flue gas flowing in a laminar state. Besides observing the flow around the model, the smoke flow method can also be used to measure the transition point of the boundary layer and study the vortex structure. Fig. 7 is a photo taken in the model smoke flow experiment.

(3) By oil flow method, an appropriate amount of indicator (such as carbon black) is mixed into viscous oil, and oil acid is dripped to make a paste-like liquid, which is evenly coated on the surface of the model. In the experiment, the flow pattern on the model surface is displayed by indicating the texture structure formed by particles along the flow direction. If a small amount of fluorescent dye is added to the oil, a fluorescent stripe pattern can be displayed under ultraviolet irradiation, which is called fluorescent oil flow pattern. It can display the air flow direction on the model surface, the position of boundary layer transition point, air separation zone, the interaction between shock wave and boundary layer and other flow phenomena. Fig. 8 is an experimental photograph of simulated oil flow.

④ Sublimation method sprays volatile liquid or easily sublimed solid on the surface of the model, and according to the principle that the loss rate of paint from the model is related to the boundary layer state (the evaporation or sublimation amount in the turbulent boundary layer is greater than that in the laminar layer due to the irregular movement of airflow), the boundary layer state can be judged, and the position of the turning point can be determined.

⑤ The steam screen method forms supersaturated steam in the wind tunnel. On the cross section to be observed, parallel light is injected perpendicular to the direction of air flow. When the airflow passes through a smooth surface, due to the centrifugal force, the water vapor content inside and outside the vortex is different, so the refractive index of light is different, and the position of the vortex core can be displayed. This method is often used to observe the position of detached vortex at high angle of attack.

⑥ Liquid crystal display method uses the characteristics of liquid crystal color changing with temperature to identify laminar flow, turbulent boundary layer and shock wave. Liquid crystal is an oily organic substance, which is colorless and transparent at low temperature. With the increase of temperature, it changes in the order of red, yellow, green, blue and colorless, which can distinguish laminar and turbulent boundary layer flows with slight temperature difference and the temperature difference before and after shock wave. Suitable for high-speed and supersonic flow observation. Liquid crystal is coated in a way similar to paint, which is diluted first and then sprayed. Liquid crystal is sensitive to dirt and impurities, so the surface of the model must be clean when spraying. According to the principle that the refractive index of light beam in gas changes with the density of gas flow, optical instruments such as shadow instrument, schlieren instrument, interferometer (see wind tunnel test instrument) and holographic device can be used to observe the flow pattern of gas. This method does not add other substances to the flow field and interfere with the gas flow, and can collect a large number of spatial data in a short time. It is an intuitive display method, especially suitable for observing compressible flow and unsteady flow, such as shock wave, wake and boundary layer transition.

In addition to the above two methods, another method is to inject energy into the flow field. For example, in a low-density wind tunnel, an electron beam is emitted to the airflow, so that gas molecules excite fluorescence, and the luminous flux of fluorescence is related to the airflow density. According to the change of luminous flux, the change of airflow density can be displayed. This method can display the position and shape of shock wave in hypersonic rarefied gas flow, and can be used for quantitative measurement of flow field density.

At the end of 1970s, the flow observation technology of color photography graphics was developed. It scans the measured flow field area with a total pressure probe and converts the sensed pressure into a voltage value. According to different voltages, different colors of light are triggered and irradiated on the camera. The vortex distribution and vortex image behind the aircraft model can be clearly seen through the isobar diagram of the flow field recorded by the multi-color signal lamp. Recently, this technology has been developed, which directly records the pressure signal sensed by the sensor on the magnetic tape and inputs it into the computer for processing. The sensor probe can be a pressure probe, a hot wire, a hot film or other probes. The processed data can be displayed by color TV. Because the camera device is not used, it is replaced by a computer, which brings great convenience: it can process a large amount of data at one time (which can be the data sensed by one probe or several probes); There are as many as 4096 colors displayed (but due to the limitation of human resolution, only 20 ~ 30 colors are commonly used); For areas of special interest, the detailed display of colors can be enlarged and increased; In addition, the displayed data plane can be rotated as required to obtain color display images of the flow field observed from different angles. For example, it can be observed in a plane perpendicular to the axis of the wind tunnel, or in a plane parallel to the axis of the wind tunnel or any other plane. A high-resolution color TV screen can indicate the flow direction with colors and arrows.