1.Problems in operating Transonic wind tunnel Testing at transonic speeds presents additional problems, mainly due to the reflection of the shock waves from the walls of the test section . Therefore, perforated or slotted walls are required to reduce shock reflection from the walls Supersonic wind tunnel 1)The power required to run a supersonic windtunnel is enormous, of the order of 50 MW per square meter of test section.:2)adequate supply of dry air 3)wall interference effects 4)high-quality instruments capable of rapid measurements due to short run times on intermittent tunnels Hypersonic Tunnels 1)supply of high temperatures and pressures for times long enough to perform a measurement 2)reproduction of equilibrium conditions 3)structural damage produced by over-heating 4)fast instrumentation 5) power requirements to run the tunnel .2. Schlieren is German for ‘striations’. The term was coined by Albert Töpler, who developed the technique in 1906 from a related technique used to identify figuring errors in telescope mirrors. Schlieren photography is a way of visualizing density variations in a gas, and is useful in wind tunnel studies and investigations into heat flow. It employs ashadowgraph principle. A collimated (i.e. parallel) beam of light passes through the test space and is brought to a focus at a knife edge; it then diverges on to a screen or a camera system. Any gas density gradient with a component perpendicular to the knife edge will deviate the light from the region, so that it
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9.Transonic - Web viewThe key word in that last sentence was stationary. ... radius as compared to traditional airfoil shapes. The supercritical airfoils were designed in the 1960s,
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1.Problems in operating
Transonic wind tunnel
Testing at transonic speeds presents additional problems, mainly due to the reflection of the shock waves from the walls of the test section . Therefore, perforated or slotted walls are required to reduce shock reflection from the walls
Supersonic wind tunnel
1)The power required to run a supersonic windtunnel is enormous, of the order of 50 MW per square meter of test section.:2)adequate supply of dry air 3)wall interference effects 4)high-quality instruments capable of rapid measurements due to short run times on intermittent tunnels
Hypersonic Tunnels
1)supply of high temperatures and pressures for times long enough to perform a measurement2)reproduction of equilibrium conditions 3)structural damage produced by over-heating4)fast instrumentation 5) power requirements to run the tunnel
.2.
Schlieren is German for ‘striations’. The term was coined by Albert Töpler, who developed the technique in 1906 from a related technique used to identify figuring errors in telescope mirrors. Schlieren photography is a way of visualizing density variations in a gas, and is useful in wind tunnel studies and investigations into heat flow. It employs ashadowgraph principle. A collimated (i.e. parallel) beam of light passes through the test space and is brought to a focus at a knife edge; it then diverges on to a screen or a camera system. Any gas density gradient with a component perpendicular to the knife edge will deviate the light from the region, so that it either clears the edge, giving a bright area on the screen, or is intercepted by it, giving a dark area. The resolution can be improved by a further knife edge at the first focus of the system. Where large spaces are to be imaged, off-axis parabolic mirrors are used rather than lenses to collimate and focus the beam . An alternative to the knife edge is a band of three colour filters, red above and blue below, with a narrow strip of green in between.
Schlieren photography is sensitive enough to record the pattern of warm air rising from a human hand, but a more sensitive test uses interferometry, in a kind of hybrid of Schlieren photography and holography. A laser beam replaces the white light beam, and a beamsplitter and beam combiner form a Mach-Zehnder interferometer set-up . This shows density differences directly, rather than density gradients
3.A shock tube is a device used primarily to study gas phase combustion reactions. Shock tubes (and
related impulse facilities: shock tunnels, expansion tubes, and expansion tunnels) can also be used to
study aerodynamic flow under a wide range of temperatures and pressures that are difficult to obtain in
other types of testing facilities.
A simple shock tube is a tube, rectangular or circular in cross-section, usually constructed of metal, in
which a gas at low pressure and a gas at high pressure are separated using some form of diaphragm, see
for instance texts by Soloukhin, Gaydon and Hurle, and Bradley.[1][2][3] This diaphragm suddenly bursts
open under predetermined conditions to produce a wave propagating through the low pressure section.
The shock that eventually forms increases the temperature and pressure of the test gas and induces a flow
in the direction of the shock wave. Observations can be made in the flow behind the incident front or take
advantage of the longer testing times and vastly enhanced pressures and temperatures behind the reflected
wave.
The low-pressure gas, referred to as the driven gas, is subjected to the shock wave. The high pressure gas
is known as the driver gas. The corresponding sections of the tube are likewise called the driver and
driven sections. The driver is usually chosen to have a low molecular weight, hydrogen or helium, for
safety reasons, with high speed of sound, but may be slightly diluted to 'tailor' interface conditions across
the shock. To obtain the strongest shocks the pressure of the driven gas is well below atmospheric.
The test being conducted begins with the bursting of the diaphragm. Three methods are in common use to
Superior design for propulsion and smoke visualization. There is no accumulation of exhaust
products in an open tunnel.
Smaller loads on model during startup because of faster starts.
Disadvantages of the Blowdown Tunnel
Shorter test times require faster (often more expensive) instrumentation.
Need for pressure regulator valves.
Noisy operation.
5.Supersonic Wind Tunnels
Supersonic wind tunnels operate differently than subsonic and transonic wind tunnels. First, because fans are inefficient at supersonic speeds, they must run subsonic and the air must make a transition from subsonic to supersonic speeds. Second, supersonic wind tunnels require an enormous amount of power. Supersonic wind tunnels can require so much power that if run during periods of peak electricity demands they can cause a regional brown-out. Very few facilities have continuous supersonic wind tunnels for this reason. The key to making a supersonic wind tunnel is to employ a supersonic venturi. Figure 8.18 shows a schematic of a closed-circuit supersonic wind tunnel. The fan moves the air in a subsonic channel. During startup the subsonic section has been pressurized while the test section remains at a static pressure of 1 atmosphere. The air accelerates in the first venturi until the speed at the throat becomes Mach 1. As the channel opens up, since the air is flowinginto a region of lower pressure it accelerates, producing the supersonic flow in the test section. After the test section the airflow goes through a second venturi. Here the speed decreases until it becomes Mach 1 at the throat. Since the air is going into a region of higher pressure, as the channel opens up the flow slows down, becoming subsonic again. The supersonic wind tunnel has an additional source of power loss. In addition to the friction on the walls and the drag on the models, now there are losses associated with the inevitable shock waves. All of these losses mean a lot of heat is being generated. In order to run continuously, a supersonic wind tunnel must have a large cooler, which is placed in the airflow in the subsonic section. The great amount of power required for supersonic wind tunnels
means there are very few continuous wind tunnels and they are not very large. A 3 _ 3 foot (1 _ 1 m) test section is considered very large and requires half a million horsepower (375 megawatts) to operate at Mach 3. But there are other methods to test supersonic aircraft. One method is the “blowdown” supersonic wind tunnel depicted in Figure 8.19. A huge tank is filled with high-pressure air and then exhausted through a venturi. This kind of wind tunnel works quite well but will
only allow a few minutes of testing. However, a carefully planned test can gather a tremendous amount of data in a very short time. With this technique the energy required is generated and stored over time. This type of wind tunnel requires very little power but requires quite a long time between tests. The NASA Hypersonic Tunnel Facility at Plum Brook can generate speed up to Mach 7. This blowdown facility can accommodate a 5-minute test every 24 hours. The Twenty-Inch Supersonic Wind Tunnel at the Langley Research Center can generate flows with Mach numbers from 1.4 to 5 for 1.5 to 5 minutes. Another option, which is more common, is the vacuum supersonic wind tunnel shown schematically in Figure 8.20. Rather than pump a chamber to a high pressure, which is dangerous, the chamber is evacuated and the airflow is in the other direction through the test section. Thus, the upstream reservoir of air is just the atmosphere and the air is being drawn through the throat and test section into a vacuum. In all supersonic venturis, the air expands on the high-speed sideand thus cools. For continuous supersonic wind tunnels this is not a concern because all the energy losses cause the air to be hot to start with. For the blowdown wind tunnels the air is often heated before it reaches the venturi so that the test section remains at a reasonable temperature. Vacuum wind tunnels have a problem that the room air is used and thus it is not practical to preheat the air. Therefore, the test section is very cold. For
example, a Mach 3 test section would be _274°F (_170°C) if the air supply were at room
temperature
6.Hypersonic TestingWith the incredible power required for supersonic wind tunnels, how can anyone expect to create hypersonic flow conditions, typically above a Mach 5, in a test environment? The only effective method to do this with a stationary model is with the blowdown method, lots of preheating of the air, and a very small test section. The key word in that last sentence was stationary. Some hypersonic facilities actually use a combustion gun, where gases combust in the breach to propel the model. The problem with this technique is that the desired measurements must be made on a nonstationary model, one that is moving very fast. But there is another trick up an engineer’s sleeve. Hypersonic flight implies that the Mach number is typically greater than Mach 5. Up to this point we implicitly assumed that to achieve hypersonic speeds we have to increase the speed in the test section or of the model. What if we were to decrease the speed of sound instead? Sound speed differs for different gases. The speed of sound decreases as the weight of the gas molecules increases. So, instead of using air for our working gas, we could look for a heavier gas, like carbon dioxide, although this will only decrease the sound speed by 14 percent. The advantage of using an alternate gas is that the true speeds can be kept reasonable, while the Mach number is fairly high.
Hypersonic Wind TunnelsSince air is stored in the high pressure air flasks at ambient temperature, it must be heated in
order to avoid condensation during operation of the hypersonic tunnels. This is accomplished by
passing the air through a bed of aluminum oxide pebbles which is enclosed in a silicon carbide
cylinder surrounded by 12 electrical heating elements (Globars); the bed is maintained at the
desired temperature by radiation from these elements. The heater was designed for a maximum
operating pressure and temperature of 600 psia and 2500F, respectively. For effective utilization
of the tunnels and instrumentation equipment available in this facility, several hypersonic tunnels
are permanently connected to the heater.
At present there are three tunnels connected to the collector; these are the Mach 4.4, the Mach 8
and the Mach 12 tunnels. However, additional access ports are available on the collector which
have been used in the past for low speed, high temperature tests involving thermal ablation,
studies of thermal stresses, and heat transfer in tubes. At the present time the hypersonic tunnels
and the heater system are in a stand-down mode. Due to the resurgence of interest in hypersonics
there is an ongoing effort to reactivate this part of the facility.
The Mach 8 tunnel has a test section diameter of 2 ft. and consists of an axisymmetric inner
contoured nozzle surrounded by a pressure shell. Test models are supported from the horizontal
access ports, the vertical windows being used for flow visualization. Free stream static pressures
between 0.2 and 3 mm Hg can be obtained with a test duration of up to 90 seconds. Test
programs have been carried out in this tunnel relating to near wake studies, nose cone
configurations, mass transfer cooling, hypersonic boundary layers, and low density shock layers.
The Mach 12 tunnel is also axisymmetric in the throat region, but at the test section the tunnel
has a decagon cross section. The test section diameter is 4 ft. and therefore permits testing of
relatively large models. The decagon shape was chosen for economy and ease of construction
since an axisymmetric contoured nozzle of this size would be extremely expensive. As a result,
the tunnel was manufactured with a monocoque structural design; ribs and stringers were formed
with the proper contour and thin plates welded to these ribs form the nozzle. The resulting
contour has the same cross sectional area at any position as the corresponding axisymmetric
nozzle. A transition between the axisymmetric throat region and the decagon cross section is
achieved in the vicinity of the throat at fairly low Mach numbers. Due to the large boundary
layer thickness of the flow in this nozzle the test section flow is essentially axisymmetric and
quite uniform. The pressure in the test section varies between 0.1 and 0.3 mm Hg and run times
on the order of 3 to 4 seconds can be achieved in this tunnel. Experimental studies which have
been conducted in this tunnel include the investigation of the near and far wake of blunt and
slender bodies and viscous/inviscid interactions occurring in a low density, high Mach number
free stream environment. Both the Mach 8 and Mach 12 tunnels exhaust into the vacuum sphere
through two large butterfly valves which can be used to isolate either tunnel from the sphere.
7.Shock Tunnels –( Hydrogen and Helium Tunnels )
How Free Piston Shock Tunnels Work
In order to understand the methods of testing and limitations for free piston shock tunnels
their method of operation must be known. The type of shock tunnel discussed will be the Free Piston Shock Tunnel (FPST) which was developed in Australia. . In order to explain how a free piston shock tunnel operates the driving principles behind a conventional shock tunnel must be understood. A regular shock tunnel has 6 major components. The main structure of a shock tunnel is of a thick walled pipe. This structure must be strong enough to withstand the high temperatures and pressures of shock tunnel operation. This is the tunnel where the shock system is created and travels through. The length of this tunnel is dependent upon the size of the test section and the duration of test wanted. Inside this tunnel there are three main sections. The first section is the compression chamber. The compression chamber is filled with a high pressure gas. This gas can be of any composition as it does not flow over the test section. The choice of driver gas is usually driven the gases thermodynamic properties. Nitrogen (N2), helium (He) and hydrogen (H2) are often used as driver gasses. The type of driver gas used dictates the resulting enthalpy of the flow and hence the speed of the flow. Hydrogen generates the highest enthalpy flow but it is dangerous to use due to its high flammability. A more common selection of driver gas is helium,
this is because it behaves as a perfect gas up to high temperatures, less energy is required to raise the temperature than hydrogen and it is safe to use. The second section of the tunnel is the expansion chamber. The gas in the expansion chamber is the gas which flows through the test section. The third component in the shock tunnel is the diaphragm. This diaphragm separates the compression and expansion sections of the tunnel. It is made of strong metal and will rupture at a predefined pressure and temperature. This is important to make this out of a light weight material so only a small amount of energy is needed to accelerate the particles to the flow velocity of the gas. The shattered diaphragm absorbs energy in the rupturing process and also tends to impede the flow through the tube. 3 Connected to the end of the tunnel adjacent to the expansion section is the nozzle. In this section enthalpy in the flow is converted into kinetic energy. The test gas expands through the nozzle and as a result accelerates. This motion is governed by the continuity equation and thermodynamic properties of the flow. The nozzle must be specially designed such that the boundary layer does not separate through the nozzle. If there is boundary layer separation the flow over the test subject may not be as desired.
The final component of a shock tunnel is the test section. This is located downstream from the nozzle. This area has mounts and wiring such that a model can be placed inside and instrumented. This section usually has windows such that the flow can be visualised using Schlieren or shadow photography techniques. In a free piston shock tunnel a “piston” which is a
large piece of metal (often steel) is placed inside the compression tube. As the name suggests this piston is free to move within the compression tube. A reservoir tank is connected to a FPST upstream of the compression tube. This reservoir is filled with high pressure gas, typically 100 atm. When the reservoir is opened the piston is accelerated down the compression tube. The speed which the piston moves is determined by the reservoir pressure, initial compression tube pressure, L/D ratio and the piston mass. As the piston moves down the compression tube energy is transferred from the reservoir gas to the compression gas. The result of this energy transfer is an increase of temperature and pressure of the compression gas. When the compression gas reaches a predefined pressure (typically 900atm, 4600K) the diaphragm separating the compression and expansion chambers will rupture. As the pressure in the expansion tube is less than that of the compression tube the He in the compression tube will flow into the expansion tube. As the pressure of the compression gas is much higher than the test gas the expansion happens rapidly. In fact the contact surface between the driver and test gas creates a shockwave. This shockwave moves faster than the contact surface between the gasses. During this stage of the test both the contact surface and the shock wave move down the expansion tube towards the nozzle.When the shockwave reaches the nozzle a secondary diaphragm in the throat of the nozzle is ruptured. This diaphragm is to separate the expansion tube from the test section. The test section is held at a pressure lower than that of the expansion tube to ensure the nozzle starts properly. When the flow hits the nozzle a shock is reflected back down the tube. This reflected shock raises the enthalpy of the flow. The test gas expands through the nozzle at supersonic speeds. As this is the case the nozzles allows the flow to expand and hence accelerate. During this expansion process the static pressure and temperature reduce. The flow velocity and the Mach number increase in the expansion process.
As the flow escapes out through the nozzle the reflected shock wave moves through the expansion tube back towards the piston. The shockwave will travel through the contact surface and be partially reflected back towards the nozzle, this shock reflection causes a negligible rise in flow enthalpy. When the shock travels through the contact surface the contact surface stops moving down the expansion tube. Between the shock being reflected and the shock travelling through the contact surface an expansion wave is propagated from the piston end of the expansion tube. The expansion wave towards the nozzle. The test is completed when the expansion wave reaches the nozzle.
8..Wind Tunnel Balance BasicsA wind tunnel balance is a device that measures the aerodynamic loads a model experience during a wind tunnel test. A balance is just a multiple axis force transducer. Balances are designed to measure some or all of the three forces and three moments a model experience. In aerodynamics terms, these forces and moments are called: Normal, Side, and Axial Force and Pitch, Yaw, and Rolling moment.
Balances come in many different designs and configurations. Most balances use strain gauged elements that relate applied loads to voltage signals. In the past, wind tunnel loads where measured using weight scales, much like the ones that existed in doctor's offices, and that's why today they're called balances.
Variations in Wind Tunnel Balances
Size and Shape How it Attaches to the Model and to the Support System The Number of Forces and Moments it can Measure The Electronics, Type of Strain Gauges, and Wiring Composed of Single or Multiple Assembled Pieces Designed Operating Load Ranges
Common Balance Types( Strain Gauged )
Internal Multiple Component Balance, with a tapered end, measures six axis loads Internal Single Piece Balance, with a cylindrical end, measures six axis loads Semi-Span Balance, Single Piece, measures five axis loads Ring or Rotor Balance
Flow through Balance
How a Strain Gauged Balance Works
Physical Elements Balances are made of flexures that deflect with load is applied. These flexures are designed to respond to load in a particular axis. Balance that can measure multiple loads and moments have individual flexures that each measure load in one axis. Strain gauges are bonded to these flexures to measure the deflections.
Electrical Elements Applied loads cause the bonded strain gauges to stretch. When a strain gauge changes length its electrical resistance changes. Individual strain gauges are wired in a whetstone bridge so that these small resistance changes can be measured as voltage signals.
Balance Inspection
Ames Balance Calibration Lab does a basic inspection for each balance before use. It is recommended that customers do an inspection before using a balance. Although not comprehensive, a basic check will avoid aggravation that could result if a faulty balance is installed in a wind tunnel model.
NOTE See the detailed explanation and working principles in Aerodynamics by Clancy9.Transonic tunnel
High subsonic wind tunnels (0.4 < M < 0.75) or transonic wind tunnels (0.75 < M < 1.2) are designed on
the same principles as the subsonic wind tunnels. Transonic wind tunnels are able to achieve speeds
close to the speeds of sound. The highest speed is reached in the test section. The Mach number is
approximately one with combined subsonic and supersonic flow regions. Testing at transonic speeds
presents additional problems, mainly due to the reflection of the shock waves from the walls of the test
section (see figure below or enlarge the thumb picture at the right). Therefore, perforated or slotted walls
are required to reduce shock reflection from the walls. Since important viscous or inviscid interactions
occur (such as shock waves or boundary layer interaction) both Mach and Reynolds number are
important and must be properly simulated. Large scale facilities and/or pressurized or cryogenic wind
tunnels are used.
UNIT 4(Theory parts) for derivations see class notes
The drag divergence Mach number is the Mach number at which the aerodynamic drag on an
airfoil or airframe begins to increase rapidly as the Mach number continues to increase [1]. This increase
can cause the drag coefficient to rise to more than ten times its low speed value.
The value of the drag divergence Mach number is typically greater than 0.6; therefore it is a transonic
effect. The drag divergence Mach number is usually close to, and always greater than, the critical Mach
number. Generally, the drag coefficient peaks at Mach 1.0 and begins to decrease again after the
transition into the supersonic regime above approximately Mach 1.2.
The large increase in drag is caused by the formation of a shock wave on the upper surface of the airfoil,
which can induce flow separation and adverse pressure gradients on the aft portion of the wing. This
effect requires that aircraft intended to fly at supersonic speeds have a large amount of thrust. In early
development of transonic and supersonic aircraft, a steep dive was often used to provide extra
acceleration through the high drag region around Mach 1.0. In the early days of aviation, this steep
increase in drag gave rise to the popular false notion of an unbreakable sound barrier, because it seemed
that no aircraft technology in the foreseeable future would have enough propulsive force or control