This lab covers the use of static taps for the measurement of static and stagnation pressures, area ratio, and Mach number in nominally steady, compressible (subsonic and supersonic) flows. In addition, it includes the use of the Schlieren technique to visualize shock and expansion waves in supersonic flows. The various measurements are acquired in an intermittent supersonic wind tunnel that utilizes a converging-diverging nozzle. The experiments also allow the student to experimentally investigate many interesting and important aspects of the behavior of supersonic flows.
The supersonic wind tunnel used in this experiment is the AF300 Intermittent Supersonic Wind Tunnel designed by TecQuipment. This is a two-dimensional wind tunnel, meaning that the test articles span the width of the entire test section. It has the capability of accelerating flow to three different speeds: subsonic, Mach 1.4, and Mach 1.8 depending on the liner that is put into the wind tunnel. A picture of the wind tunnel is shown in Figure 1 below.
The wind tunnel works using the concept of induction. A downstream flow of high pressure compressed air induces an upstream airflow. This method of flow induction through the tunnel reduces turbulence and increases flow steadiness, allowing us to have a better comparison to theory. Figure 2 below shows a diagram of the wind tunnel and the direction of induced flow.
The working section of the wind tunnel is where the test article sits. Compressed air is introduced into the flow via the injector block downstream of the working section. Because of the high pressure air that is introduced aft of the working section and directed downstream, this creates a pressure drop in the working section of the wind tunnel and thus inducing air through it. This air moves around the duct, through the flow straightener, through the contraction cone, and then into the working section of the wind tunnel. After passing through the working section, the flow mixes with the compressed air flow and moves around the tunnel again in a cycle. Any excess air that is in the system exits through the outlet filter until the compressed air supply is shut off.
Compressed air is supplied to the wind tunnel by a compressor. The compressor used for this wind tunnel is the Renner Kompressoren RS-PRO 15. This is a 15 kW 220V belt-driven compressor with the capability of compressing air up to 15 bar (about 215 PSI). Air from this compressor then pases through a dryer and a filter system. For this tunnel, we use a Nano D-Series Heatless Compressed Air Dryer. The purpose of this dryer is to remove moisture from the compressed air and filter out contaminants that may otherwise cause unintentional icing and unnecessary wear on the wind tunnel.
Once through the dryer, the air enters 3 receiver tanks. Each of these tanks has a capacity of 500 liters for a total air supply of 1500 liters or just under 400 gallons. The compressor automatically shuts off when the air in these receiver tanks reaches 13 bar. There is a flexible hose that connects these receiver tanks to the the injector block, where compressed air is introduced into the wind tunnel. Also on this line is a compressed air ball valve which can be opened or closed in order to introduce compressed air into the wind tunnel. When open, air is supplied to the wind tunnel from the receiver tanks and when closed, the compressed air supply is stopped. Figure 3 below shows the complete setup of the supersonic wind tunnel in the Low Turbulence Wind Tunnel room in MK 104.
The working section of this wind tunnel is the black section of the wind tunnel aft of the contraction cone seen in Figure 4 below. It is a precision engineered converging-diverging rectangular-section nozzle. That is to say that the the outside of the working section is a rectangle block but the inside of it, where the air actually flows, is a converging-diverging nozzle.
Near the back of the working section is the test section. The test section is where the test article sits. There are two removable glass plates (portals) on either side of the working section that gives access to the test section, allowing us to swap out the test articles. The test articles are held in place via slots in the portals--see above image. The portals are made of glass so that light can be shined through it for Schlieren imaging. Downstream of the working section is the brass injector block which injects compressed air into the wind tunnel.
There is a knob connected to the portals that can be turned to change the angle of attack of the test article. This can be seen in both Figure 4 above and Figure 5 below. On the outer diameter of the portal, there is a gear ring that is connected to a gear on the angle of attack adjustment knob. Therefore when the knob is turned, the gears on the knob turn the gear ring on the portals, thus changing the angle of attack of the test article. This angle of attack is recorded by an angle encoder is read by the data acquisition system.
The bottom of the working section is flat and has 25 static pressure taps along the length of it. The first tap is 19.5mm downstream of the contraction cone and each subsequent tap is 25mm away from the previous tap. A diagram of this is shown in Figure 5 below. Each of these static pressure taps is connected to an electronic transducer that is read by the data acquisition system. Essentially, each of these taps reads pressures along the length of the converging-diverging nozzle.
The top surface of the working section consists of an interchangeable liner. This tunnel is supplied with 3 different liners, which are shown in Figure 6 below. Depending on which liner is inserted into the working section, a different flow speed can be achieved. If the subsonic liner is put into the working section, no supersonic flow can be achieved because the throat area is not large enough to choke the flow to sonic speed (hence the name subsonic liner).
The Mach 1.4 liner can be placed into the working section to achieve flow at a speed of Mach 1.4. This is because the throat area is converged more than the subsonic liner and is reduced enough to choke the flow. It then diverges to an area ratio that gives a flow speed of Mach 1.4. Similarly, the Mach 1.8 liner accelerates the flow to a speed of Mach 1.8 by further reducing the throat area (and thus increasing the diverging area ratio).
The data acquisition system used by this wind tunnel is called VDAS, or Versatile Data Acquisition System. VDAS connects to a software called TecQuipment VDAS that is used to record and export data for this wind tunnel. VDAS reads and records all the relevant data for this wind tunnel and feeds it into the software so we can visualize and export data:
Time
Tunnel Reference Pressure
Supply Pressure
Atmospheric Pressure
Angle of Attack
Up to 32 Static Pressure Taps
This system has the capability of reading up to 32 static pressures from taps labeled 1 through 32. For our purposes, we will only be using taps 1-28. The taps are ordered in a specific way and each tap corresponds to a different pressure measurement.
The Schlieren Apparatus is way to visualize shocks using lenses, mirrors, a lamp, and a screen. Schlieren techniqures were first used in the 19th century to detect faults in glass. The concept of Schlieren imaging for the visualization of shock waves relies on the change in the way light is refracted depending on density variations of the atmosphere in liquids. "Schlieren" is a German word that translates to "streaks."
Parallel light is aimed through the material of interest from one side. The light that leaves the material on the opposite side is focused onto a knife edge. Refracted light falls below or above the knife edge and appears lighter or darker that un-refracted light. This then produces an image of varying light intensity. An example of a Schlieren Image of a shock wave can be seen in Figure 7 below.
The Schlieren Apparatus used for this supersonic wind tunnel is the AF300a Schlieren Apparatus designed by TecQuipment. The system begins with a 100 Watt halogen lamp which acts as the light source. The light from the lamp is focused by passing through a condenser lens. The focused light then travels through an optical slit. The light is then turned 90 degrees using a mirror and goes through an achromatic lens that is flush with one side of the working section portal. The light then passes through the test section and through a second achromatic lens that is flush to the other side of the working section portal. A second mirror turns the light another 90 degrees towards the Schlieren edge. The edge enhances the light refracted image by removing light intensity. After passing through the Schlieren edge, the light then passes through a final lens which then displays the Schlieren image on an imaging screen. A diagram of this process is shown in Figure 8 below.
Figure 9 below shows the setup of our Schlieren apparatus that will be used in this lab. There is a camera placed behind that imaging screen that takes a video of the imaging screen. The video feed from this camera is then displayed on the monitor for easy visualization of the imaging. This system yields the best results when the room lights on that side of the room are switched off--essentially it brightens up the Schlieren image for the camera sin the room is relatively much darker.
A brief description of compressible flows (summarizing important details covered in AE 2010) is given below. It is important to note that inviscid flow will be assumed throughout this discussion. Compressible flow, especially at supersonic speeds, results in some remarkable phenomena. For example, it turns out that an adiabatic flow can increase from subsonic to supersonic speeds if a converging-diverging nozzle is used. If you converge subsonic flow and gradually increase Mach number to sonic speed (M=1) at the throat (minimum area) of a CD nozzle, the flow transitions to supersonic in the diverging section of the CD nozzle!
When a gas flows at Mach numbers greater than M>0.3, the density changes of the gas due to changing Mach number become significant (the flow becomes compressible). Bernoulli's equation, which was derived assuming constant density, is no longer valid for compressible flow. Instead, the relationship between flow characteristics becomes a function of Mach number.
Isentropic flow ocurs when the changes in flow variables are slow and gradual. This results in a reversible process, meaning if the process was reversed, the flow conditions revert to their original values. To visualize this, imagine a flow that is decelerated and then accelerated through a nozzle that gradually diverges and then gradually converges back to its original area. In this process, the flow conditions at the beginning and end of the nozzle are the same. In a reversible process, entropy is conserved, hence the name isentropic flow.
For isentropic compressible flow, a group of equations have been derived to describe the flow. These are shown in Figure 10 below.
If a supersonic flow encounters a solid body, part of the flow must be decelerated to stagnation conditions. Since this necessarily involves a transition from supersonic to subsonic flow, a shock stands ahead of the body. A compression shock is a series of compression Mach waves that coalesce to form one finite shock wave which is comparable to the thickness of the mean free path of the gas involved. Through a shock, an abrupt and instantaneous change in the flow conditions occurs. This process is irreversible and therefore, the entropy of the system increases. Because of this increase in entropy, flow through a shock wave is no longer isentropic and the above equations cannot be used.
The shape of the body determines what shape the shock wave will take. These can be seen in Figure 11 above. In the case of a flat blunt body in supersonic flow, a detached normal shock stands ahead of the body. In the case of a thin wedge, an attached oblique shock stands in front of the body. As a third example, in the case of a thin probe in supersonic flow (such as a pitot-static prove), a small, normal shock stands ahead of the probe tip. A conically shaped bow shock trails from the normal shock. The bow shock weakens as one moves away from its leading edge and, eventually, turns into a Mach wave across which the flow properties no longer change significantly.
Through a shock wave, the Mach number decreases, the static pressure increases, the static temperature increases, and the density increases. Furthermore, the stagnation pressure decreases through a shock and the stagnation temperature stays the same. If the shock is a normal shock, the shock wave decelerates the flow from supersonic to subsonic. A series of equations to calculate flow properties after a normal shock have been derived. These are based on the flow properties before the shock and are shown in Figure 12 below.
Oblique shocks form when compression Mach waves coalesce over a concave curve. When the curve is abrupt (a corner) instead of gradual, the oblique shock wave starts right at the corner. This is shown in figure 13 below.
The equations for the flow through an oblique shock are similar to that of a normal shock. They are heavily dependent on the deflection angle and the Mach number upstream of the deflection. The shock angle of the oblique shock changes if either of these two values are altered. Through an oblique shock, the component of the flow that experiences a shock wave is dependent on the velocity component of the flow that is normal to the shock wave. Because of this, it is possible to have supersonic flow after an oblique shock if the shock angle is small enough! Though it is very easy to derive oblique shock equations from the normal shock equations above, they are posted in Figure 14 below.
When supersonic flow hits a concave corner, it has been shown that an oblique compression shock wave occurs. Similarly, when supersonic flow passes over a convex corner, some sort of expansion shock must occur. An expansion shock in its essence cannot occur the way a compression shock does because it would require a decrease in entropy, which is a violation of the second law of thermodynamics. Instead, a series of expansion Mach waves occurs in place of an expansion shock.
Unlike an oblique shock wave where weaker compression Mach waves coalesce to form one finite oblique shock wave, expansion waves across a gradual convex curve do not coalesce to form one finite expansion wave. This is because while compression mach waves converge across a concave curve, expansion mach waves diverge across a convex curve. Therefore, an expansion shock wave is physically impossible (on top of violating the 2nd law of thermodynamics).
When the flow goes through these series of expansion Mach waves across a convex curve, the flow experiences gradual, reversible changes in flow properties rather than an abrupt change like that in a shock wave. Thus, flow going through an expansion fan is isentropic. The stagnation properties of the fluid are constant, but the static properties change according to the isentropic flow equations described above.
If the convex change is abrupt rather than gradual (i.e. a corner), a group of expansion waves centered at the corner are formed. This is called a Prandtl-Meyer Expansion Fan. See Figure 15 below.
The Mach number increases across an expansion fan and therefore, the static pressure of the flow drops across an expansion fan. The increase in Mach number is dependent on the turning angle of the corner and the speed of the flow before the corner. Figure 16 below shows the expansion fan relations.
In this experiment, a 10 degree wedge is used to gather data and visualize shock waves and expansion fans. Furthermore surface static pressure data is also gathered from this wedge. For your reference, a diagram of the test article is shown in Figure 17 below.
The compressed air supply for this experiment is not infinite and the flow rate through the tunnel is substantial. As such, ensure you are aware of all the data you will need to take for each run of the tunnel and that you know how to take the data ahead of time. The run time for this tunnel stands at about 8 seconds before the pressure in the tanks drops too low and results in an inability to complete the experiment. Ensure you have read the procedure thoroughly and you know what data you will need to take before each run of the supersonic tunnel.
Between runs, you will need to wait about 4-6 minutes for the compressor to charge the tanks. The compressor will charge the tanks to about 10-13 bar. The pressure in the tanks will be too low to accelerate the flow to supersonic speeds when it reaches about 7.5 bar.
Take a tour of the equipment with the TAs. In particular, you will look at and discuss:
The Renner Compressor which feeds the receiver tanks. This compressor is set to kick in whenever the pressure in the receivers drops below 10 bar.
The dryer and receiver tanks. Take a look at the pressure relief valves on the tanks and the tank pressure gauges.
The tubing and valves that supply the receivers and the wind tunnel. Keep a note of which ones are closed and which are open. Try to understand why the tubing is the way it is and why certain valves are open or closed. Ask a TA if necessary.
The wind tunnel itself. Take a look at the tunnel duct, the contraction cone, the working section, and the compressed air injector. Can you figure out how the tunnel works? Why is it shaped the way it is?
The various pressure gauges and pressure taps on the VDAS instrument panel and how they correlate to pressure taps in the working section of the wind tunnel.
The working/test section. Look at how the model is held in place and understand the angle of attack adjustment knob.
The Schlieren Apparatus.
PPE and other safety information about the tunnel.
Preparing the wind tunnel for testing:
Locate the two safety switches and ensure both are switched off. One safety switch is located a little to the left of the supersonic wind tunnel. The second is on the wall to the left of the door above the Yellow Educational Wind Tunnel.
Ensure that the compressor power receptacle is plugged into the black wall outlet extension cable. Ensure that the black wall outlet extension cable is plugged into the wall outlet. These are both twist locks. If either is unplugged, have a TA plug them in for you.
The same outlet is used for the EWT and the Supersonic Tunnel so there is a chance it may not be plugged into the proper equipment.
Flip both safety switches from off to on.
Wait a few moments for the compressor to warm up. It will turn on automatically and run its system checks.
Ensure that the E-Stop on the compressor is not switched on. If it is, turn the E-Stop button clockwise until it pops outward. Have a TA remove the E-Stop error that is present upon startup.
Ensure that the tunnel inlet shut-off valve is closed. This valve is closed when the handle is perpendicular to the pipe. The tunnel shut-off valve is located right below the tunnel pressure gauges on the vertical pipe. DO NOT close the valve on the horizontal pipe!
Turn on the compressor by pressing the green button. The compressor is loud, so use hearing protection if you would like.
When the receivers are empty, the compressor takes about 8-10 minutes to fully charge the receivers. The compressor will automatically shut off when the tanks reach about 11.5-13 bar.
Compare the pressure given by the compressor to the values on the receiver pressure gauges and the supply pressure gauge on the wind tunnel. Are they the same?
Turn on the VDAS panel.
Ensure the orange/yellow extension cord is plugged into the wall and ensure the surge protector is turned on.
Turn on the VDAS by toggling the switch to on. This switch is located on the instrument frame on the left side.
Once the compressor shuts off, zero all the static pressure taps by pressing and holding the “Zero” button on the instrument panel for about 5 seconds.
Setting up the software system for data acquisition:
Open "TecQuipment VDAS" software on the computer from the desktop icon.
If prompted to select the Application, select “AF300 Supersonic Wind Tunnel (Intermittent)” from the menu and click OK. The following screen will open up:
Under Operating Conditions, ensure all 4 boxes are checked so that this data gets exported. Ensure the values read the following:
Liner: Mach 1.8
Atmospheric Pressure (mbar): 1013
Uncheck the checkbox across from "Mach Numbers" so that this data does not get exported. Your software screen should now look EXACTLY like the picture above.
Familiarize yourself with the software. Understand where all the static ports are on the wind tunnel and what all the data on the screen means. Familiarize yourself with where the necessary buttons are for running the software, acquiring data, saving/exporting data, and changing the data acquisition intervals. Have a TA help with explaining this.
When beginning your experiment, one of two airfoil models may already be inserted into the wind tunnel. You may start the experiment with either model in the wind tunnel or with the wind tunnel empty. MAKE SURE you are following the proper procedure for whichever model is currently in the wind tunnel!
For this experiment, you will be running the tunnel without a test article in it. You will be obtaining static pressure data from the static taps along the working section of the converging-diverging nozzle in order to understand the area ratio of the nozzle with respect to the throat of the CD nozzle.
If necessary, have a TA open the test section and remove any test article that may be in the wind tunnel.
Fill the mount holes with the end caps provided.
Leaving the mount holes open creates strange 3D shocks and expansion waves that reflect within the test section leaving us with pressure data that is not representative of the phenomenon we are trying to capture
If the end caps are not available, simply put two models on either side of the wind tunnel such that it is sticking out but the mount holes are filled.
The far side panel is currently an acrylic plate that is slightly thinner than the panel on the front side of the working section. For this far side acylic panel, use the resin printed end cap provided.
A TA can assist with this if needed
Running the Experiment:
In the VDAS software, click the red Play button on the top left corner of the screen. You should now see a live reading of the tunnel reference pressure and the model angle.
In the pop-up window, set the following values:
Interval: 0.5 seconds
Stop after: 10 seconds
DO NOT yet press the play button in the pop-up!
When you are ready to run the experiment, do the following. READ STEPS 4-7 CAREFULLY BEFORE RUNNING THE TUNNEL.
Have one student near the tunnel inlet shut-off valve and also keeping an eye on the tunnel reference pressure gauge
Have one student controlling the software
All students and TAs must now be wearing hearing protection!
In the VDAS pop-up window, hit the red Play button. This will start timed data acquisition.
Wait 1-2 seconds, and then open the tunnel inlet shut-off valve to start flow into the tunnel.
When the tunnel reference pressure gauges reads 7.5-8 bar, close the tunnel inlet shut-off valve.
In the VDAS software, you will see a pop-up notification reading "Timed data capture complete". Click OK.
Navigate to C:\Users\Public\Documents\AE3610Data\<Semester Year> and select a relevant file name for your spreadsheet. If your section folder does not already exist, create a new one for yourselves.
Upon hitting save, your file will automatically open. Look through the data and make sure it is good and makes sense. Re-run the experiment if necessary.
When beginning your experiment, one of two airfoil models will already be inserted into the wind tunnel. You may start the experiment with either model in the wind tunnel or with the wind tunnel empty. MAKE SURE you are following the proper procedure for whichever model is currently in the wind tunnel!
In this experiment, you will gather static pressure data from the surface of a double wedge airfoil as angle of attack is changed. Furthermore, you will obtain static pressure data along the working section of the converging-diverging nozzle.
Make sure you are taking lab notes.
If necessary, have a TA open the test section and insert the 10 degree double wedge with two pressure tappings test article into the wind tunnel. Make sure the two static taps are facing upwards!
Connect the two static taps to the open tubes near the test section. These will record static tap data to channels 26 and 27 on VDAS. Have a TA help with this if necessary.
The tubes that connect to the static ports have a tendency to get pinched since they are not that long. Make sure both the tubes are not pinched! You will get bad data if they are!
In the VDAS software, click File-->New. This will open a new file for this experiment. If prompted to clear unsaved data, click OK.
Click the red Play button on the top left corner of the screen. You should now see a live reading of the tunnel reference pressure and the model angle.
Rotate the model such that it is at 0 degrees angle of attack
Rotate the model such that it is at -1.1 degrees angle of attack (+/-0.1 degrees is fine)
The encoder is currently zeroed at at -1.1 degrees, hence why we will put it there. However, this is actually 0 degrees in reality!
In the pop-up window, set the following values:
Interval: 0.5 seconds
Stop after: 10 seconds
DO NOT yet press the play button in the pop-up!
When you are ready to run the experiment, do the following. READ STEPS 8-11 CAREFULLY BEFORE RUNNING THE TUNNEL.
Ensure the static port tubes are not pinched
Have one student near the tunnel inlet shut-off valve and also keeping an eye on the tunnel reference pressure gauge
Have one student controlling the software
All students and TAs must now be wearing hearing protection!
In the VDAS pop-up window, hit the red Play button. This will start timed data acquisition.
Wait 1-2 seconds, and then open the tunnel inlet shut-off valve to start flow into the tunnel.
When the tunnel reference pressure gauges reads 7.5-8 bar, close the tunnel inlet shut-off valve.
In the VDAS software, you will see a pop-up notification reading "Timed data capture complete". Click OK.
Repeat steps 5-12 for angles of attack of 2.5, 5, and 10 degrees. You will need to wait for the compressor to recharge between tests.
Keep in mind that this is actually 2.5, 5, and 10 degrees!!!
Navigate to C:\Users\Public\Documents\AE3610Data\<Semester Year> and select a relevant file name for your spreadsheet. If your section folder does not already exist, create a new one for yourselves.
Upon hitting save, your file will automatically open. Look through the data and make sure it is good and makes sense. There should be 4 data series (one for each angle of attack run) each with 21 data points. Re-run the experiment if necessary.
When beginning your experiment, one of two airfoil models will already be inserted into the wind tunnel. You may start the experiment with either model in the wind tunnel or with the wind tunnel empty. MAKE SURE you are following the proper procedure for whichever model is currently in the wind tunnel!
For this experiment, you will primarily be observing the oblique shock waves and Prandtl-Meyer expansion fans created by a diamond airfoil (double wedge) using the Schlieren apparatus. All data taken during this section will be pictured from either the mounted camera or your camera phone. No pressure data will be obtained.
Make sure you are taking lab notes.
Have a TA remove the two tubes connected to the ports on the wedge.
Setting up the Schlieren Apparatus:
Turn on the Halogen lamp. You can do this by toggling the switch on the VDAS instrument panel.
Turn on the video camera and Monitor.
The camera is turned on by flipping open the camera screen. It is already connected to a power source.
Plug the Micro-HDMI cable into the camera
If the monitor is not already turned on, turn it on by clicking the button on the back of the monitor. You should see an image of the test section and test article from the camera.
Have a TA ensure that the two lenses on either side of the working section are flush to the test section. Do this CAREFULLY.
DO NOT MOVE ANY OTHER LENSES OR MIRRORS ON THE SCHLIEREN TABLE!!!!! These have been PRECISELY set up ahead of time by the lab manager. If anything is moved, have the TA fix it.
Note for TA's: make adjustments to the Schlieren system ONLY if necessary! Call the Head TA or the Lab Manager if you have any issues.
Turn off the room lights so you can see the image better. This light switch is located to the left of the supersonic wind tunnel next to the safety switch that is above the EWT.
Running the Experiment:
Ensure that the VDAS software is running. Press the red play button in the top left if necessary.
Rotate the model such that it is at 0 degrees angle of attack
Rotate the model such that it is at -1.1 degrees angle of attack (+/-0.1 degrees is fine)
The encoder is currently zeroed at at -1.1 degrees, hence why we will put it there. However, this is actually 0 degrees in reality!
When you are ready to run the experiment, do the following. READ STEPS 3-6 CAREFULLY BEFORE RUNNING THE TUNNEL.
Have one student stand near the monitor with their phone camera pointed at the monitor displaying the Schlieren image
Have one student near the tunnel inlet shut-off valve and also keeping an eye on the tunnel reference pressure gauge
All students and TAs must now be wearing hearing protection!
Open the tunnel inlet shut-off valve to start flow into the tunnel. The tunnel will get very loud and you will start to see shock waves on the Schlieren Image.
Take images of the monitor with your phone to capture the oblique shock if you wish.
Alternatively, you may take images with the camera that is mounted and transfer images from the microSD card to your computer using the provided adapter.
This will not be required for your reports
When the tunnel reference pressure gauges reads 7.5-8 bar, close the tunnel inlet shut-off valve.
Wait for the compressor to recharge the tanks.
Repeat steps 2-6 for angles of attack of 2.5, 5, and 10 degrees (again, these are actually 2.5, 5, and 10 degrees!!).
When you are satisfied with all your data,
Turn off the halogen lamp
Turn on the room lights
Turn off the camera by unplugging the micro-HDMI cable and closing the small screen
If you took pictures with the mounted camera, transfer the images to your personal PC using the microSD-->USB adapter provided
Make sure you delete your images from the microSD card after you have transferred them to your PC
Put the microSD card back in the mounted camera
DO NOT leave with the adapter!
This experiment is not necessary for your data report. Do this experiment IF AND ONLY IF (1) you have time remaining at the end of your lab, (2) you have verified that all your data is good and no experiment needs to be redone, and (3) you have no questions for your TAs and nothing to discuss with your group.
For this experiment, you will primarily be observing a detached bow shock created by a blunt body using the Schlieren apparatus. No data or pictures are necessary unless you want to take pictures with your phone.
Have a TA open the test section and insert the blunt body test article into the wind tunnel.
Setting up the Schlieren Apparatus:
Turn on the Halogen lamp. You can do this by toggling the switch on the VDAS instrument panel.
Turn on the video camera and Monitor.
The camera is turned on by flipping open the camera screen. It is already connected to a power source.
Plug the Micro-HDMI cable into the camera
If the monitor is not already turned on, turn it on by clicking the button on the back of the monitor. You should see an image of the test section and test article from the camera.
Have a TA ensure that the two lenses on either side of the working section are flush to the test section. Do this CAREFULLY.
DO NOT MOVE ANY OTHER LENSES OR MIRRORS ON THE SCHLIEREN TABLE!!!!! These have been PRECISELY set up ahead of time by the lab manager. If anything is moved, have the TA fix it.
Note for TA's: make adjustments to the Schlieren system ONLY if necessary! Call the Head TA or the Lab Manager if you have any issues.
Turn off the room lights so you can see the image better. This light switch is located to the left of the supersonic wind tunnel next to the safety switch that is above the EWT.
Running the Experiment:
Ensure that the VDAS software is running. Press the red play button in the top left if necessary.
Rotate the model such that it is at 0 degrees angle of attack (+/-0.1 degrees is fine)
When you are ready to run the experiment, do the following. READ STEPS 3-6 CAREFULLY BEFORE RUNNING THE TUNNEL.
Have one student near the tunnel inlet shut-off valve and also keeping an eye on the tunnel reference pressure gauge
All students and TAs must now be wearing hearing protection!
Open the tunnel inlet shut-off valve to start flow into the tunnel. The tunnel will get very loud and you will start to see shock waves on the Schlieren Image.
Observe the bow shock. Take images if you would like ONLY WITH YOUR PHONE.
When the tunnel reference pressure gauges reads 7.5-8 bar, close the tunnel inlet shut-off valve.
You may wait for the compressor to recharge and run it again if you would like to see the shock again.
Turn off the schlieren system.
Turn off the halogen lamp
Turn on the room lights
Turn off the camera by unplugging the micro-HDMI cable and closing the small screen
Do not start this list if you have any missing or bad data. Go through it with your TAs if you wish.
Hit the red stop button on the compressor panel to turn off the compressor. This button is located right below the green start button. DO NOT HIT THE E-STOP! The compressor will not vent compressed air if you do not hit the proper stop button. It will take about 3 minutes to cool down, vent air, and shut down.
If the E-stop is already engaged, disengage the e-stop, remove the error from the compressor screen, and then start the compressor again
Once the is shown as "On Load" you may hit the red stop button to turn the compressor off properly
Turn off the VDAS by toggling the switch on the instrument panel to off. This switch is different from the Halogen Lamp switch. Make sure BOTH are turn off!
Have a TA upload your data to canvas.
Turn off the TV. DO NOT switch off the surge protector or unplug the orange/yellow extension cable from the wall.
Once the compressor finishes its cool down and turns off, flip both safety switches to off. This will cut power to the compressor.
With hearing protection on, vent the receivers of compressed air by opening the tunnel inlet shut-off valve.
Once the tunnel reference pressure gauge drops to zero, it means the air in the tanks has been discharged. Close the tunnel inlet shut-off valve.
Throw out any used hearing protection and plastic wrappers into the trash bin.
If you are the last section of the day, have a TA empty the water collection tray near the dryer and place it back there.
If the schlieren system is on:
Turn off the halogen lamp
Turn on the room lights
Turn off the camera by unplugging the micro-HDMI cable and closing the small screen
If you took pictures with the mounted camera, transfer the images to your personal PC using the microSD-->USB adapter provided
Make sure you delete your images from the microSD card after you have transferred them to your PC
Put the microSD card back in the mounted camera
DO NOT leave with the adapter!
From the wind tunnel area characterization experiment:
An excel spreadsheet containing 1 data series with 21 data points: the working section converging-diverging nozzle static pressures from the static taps
From the 10 degree double wedge with two pressure tappings experiment:
An excel spreadsheet containing 4 data series each with 21 data points: the working section converging-diverging nozzle static pressures and airfoil surface static pressures at 4 angles of attack (0, 2.5, 5, 10 degrees)
For the wind tunnel area characterization experiment:
Parse your data series to get rid of erroneous data rows (e.g. rows that recorded data before the tunnel was turned on)
From the remaining data, select the row of data you believe is the best row of data
Note: We are leaving this up to you to pick. What is good vs. bad data in this scenario?
Using the row of data selected from step 1.2, calculate the Mach number at each of the 25 stations based on the recorded pressure ratio and given that A* is somewhere between stations 7 and 8
Calculate the area ratio at each of the 25 stations
For the 10 degree double wedge experiment:
Obtain the shock angle, wedge angle, and expansion angle of the airfoil's top surface at 0, 2.5, and 5 degrees angle of attack. Do this using the wedge angle and supersonic flow theory.
Using the Mach number at station 19 as your upstream mach number and using the shock angle information obtained from step 2.1, calculate the mach number and stagnation pressure downstream of the oblique shock.
Using the expansion angle information from step 2.1 and the Mach number and stagnation pressure calculated from step 2.2, calculate the mach number and stagnation pressure downstream of the expansion fan using supersonic flow theory.
Do Steps 2.2 and 2.3 for 0, 2.5, and 5 degrees angle of attack.
For the 10 degree double wedge with two pressure tappings experiment:
Parse your data series to get rid of erroneous data rows (e.g. rows that recorded data before the tunnel was turned on)
From the remaining data, select the row of data you believe is the best row of data for each of the 4 angles of attack
Note: We are leaving this up to you to pick. What is good vs. bad data in this scenario? Your selection criteria to justify this may be a good discussion point!
You should now have 4 rows of data--one for each angle of attack
Calculate the mach number on the top surface of the airfoil downstream of the oblique shock using the stagnation pressure obtained in step 2.2 and the static pressure from the surface tap.
Hint: your pressure ratio for taps 26 and 27 are different from what the software gave you
Calculate the mach number on the top surface of the airfoil downstream of the expansion fan using the stagnation pressure obtained in step 2.3 and the static pressure from the surface tap.
Hint: your pressure ratio for taps 26 and 27 are different from what the software gave you
Do Steps 3.3 and 3.4 for 0, 2.5, and 5 degrees angle of attack.
It is on you to figure out which airfoil surface pressure tap comes first in the flow! Tap 26 could be either after the oblique shock or after the expansion fan depending on how the test article was oriented during your test. Refer to Figure 17 if needed.
For the wind tunnel area characterization experiment at all 25 stations:
A table containing the station number, static pressure, pressure ratio, calculated mach number, and calculated area ratio (A/A*) for the chosen row of data.
A plot of the mach number in the converging-diverging nozzle with the pressure tap distance as the abscissa and the mach number as the ordinate
A plot of the area ratio in the converging-diverging nozzle with the pressure tap distance as the abscissa and the area ratio as the ordinate
For the 10 degree double wedge experiment:
A table containing:
Shock angle on the top surface of the airfoil at 0, 2.5, and 5 degrees angle of attack
Wedge angle on the top surface of the airfoil at 0, 2.5, and 5 degrees angle of attack
Expansion angle on the top surface of the airfoil at 0, 2.5, and 5 degrees angle of attack
Mach number downstream of the shock wave and Mach number downstream of expansion wave at 0, 2.5, and 5 degrees angle of attack calculated from Step 2 of the Data Reduction
If you have them, you may include the four images of the shock at 0, 2.5, 5, and 10 degrees angles of attack. This is not necessary for your report
For the 10 degree double wedge data with two pressure tappings:
For all 27 stations, a table containing the station number, static pressure, and pressure ratio for 0, 2.5, and 5 degrees angles of attack (hint: your pressure ratio for taps 26 and 27 are different from what the software gave you)
A table containing the angle of attack, pressure ratio downstream of the shock, pressure ratio downstream of the expansion wave, Mach number downstream of the shock calculated from pressure ratio, and Mach number downstream of the expansion wave calculated from the pressure ratio
A plot with the angle of attack as the abscissa and the pressure ratio downstream of the shock as the ordinate. On the same figure, plot the pressure ratio downstream of the expansion wave as a function of angle of attack as a different series
A plot with the angle of attack as the abscissa and the Mach number downstream of the shock calculated from pressure ratio as the ordinate (Data Reduction Step 3). On the same figure, plot the Mach number downstream of the shock obtained from the shock and wedge angles (Data Reduction Step 2).
A plot with the angle of attack as the abscissa and the Mach number downstream of the expansion wave calculated from pressure ratio as the ordinate (Data Reduction Step 3). On the same figure, plot the Mach number downstream of the expansion wave obtained from the expansion angle (Data Reduction Step 2).
| Pressure Tap | Corresponding Tap Number on VDAS |
|---|---|
In the VDAS software, click the "Start timed acquisition of data" button
Export this data to a .xlsx file by clicking the "Export recorded data to Excel XLSX file" button
In the VDAS software, click the "Start timed acquisition of data" button
Export this data to a .xlsx file by clicking the "Export recorded data to Excel XLSX file" button
Working Section Static Tap 1-25 in order
Taps 1-25 in order
Airfoil Top Surface Pressure Taps 1 and 2
Tap 26 and 27 (not necessarily in order)
Static Pressure at Contraction Cone
Tap 28
