Originally Posted by Floatdog

Be careful using this chart, it's log-log.

The chart shows that air flow perpendicular to a cylinder has a Cd of about 1.2 for Reynold Numbers below about 1E05. Dr. Hoerner's book (Chapter 3) indicates that these results were measured with cylinders which went from wall to wall. So an experiment was run where a 5.5 mm rod that went from the floor to within about 2 mm of the tunnel ceiling.
The results are as follows; (Max frontal is the reference area)

Original Cd = 0.86, length =19.8 cm, diameter = 5.5 mm , Ref area = 11.44 cm2, V = 19.2m/s, Pressure = 177.6 Pa, Drag =0.175 N, rho = .0967 kg/m3, Re = 6,000

Wall to Wall Cd = 0.95, length = 24.3 cm, diameter = 5.5 mm, Ref area = 11.44 cm2, V = 19.2 m/s, Pressure = 178.2 Pa, Drag = 0.227 N, rho = 0.967 kg/m3, Re = 6,000

A check was made on the tunnel and test instrumentation by using a 50 mm acrylic sphere mounted on a rear sting. Literature indicates that the Cd should be in the region of 0.43 to 0.45 at an Reynolds Number about 53,000 The measured value is 0.48, about 10% too high.

Diameter = 50.0 mm, Max frontal area (ref) = 13.37 cm2, V = 19.3 m/s, Pressure = 177.8 Pa, Drag = 0.168 N, rho = 0.956 kg/m3, Re = 53, 800

Given that the instrumentation and tunnel are operating properly, why is the rod result so much less than shown in the literature? Hoerner (Figure 14, page 3-10) indicates that roughness can cause that magnitude of difference but the rod was polished to a dull shine.
It was noted that there was some rod vibration while running.

Bill

Can you check the velocity profile in the test area?

Perhaps thats causing some issues

Originally Posted by Floatdog

Can you check the velocity profile in the test area?

Perhaps thats causing some issues

As part of the tunnel test program a vertical survey was conducted from top to bottom in the tunnel center. The variation was less than =+/- 1% rms. The rods were tested in the same vertical plane. See graph.

xx

Holy mackerel, nothing wrong there.

Excellent work!

Perhaps a larger diameter rod will help produce more 'real' numbers, and may eliminate the vibration.

Originally Posted by Floatdog

Holy mackerel, nothing wrong there.

Excellent work!

Yea, I spent a lot of time in designing and building this tunnel: honeycomb, dual screens, concentrator section contours, smooth surfaces, small diffuser angle and exit shaping to match the fan.

I will try a much larger diameter cylinder; mailing tube covered with film, about 2" inches in diameter and 20 cm length.
The vibration may be an issue as the sting is mounted to the force sensor mount using servo tape which is somewhat flexible. Something of a problem to fix as I'm reluctant to make a more permanent mount. The mount is wood and the screw threads holding the sting in place tend to work loose with use.
Maybe buy another scale and use epoxy instead.

xx

Originally Posted by aeromodel03

The chart shows that air flow perpendicular to a cylinder has a Cd of about 1.2 for Reynold Numbers below about 1E05. Dr. Hoerner's book (Chapter 3) indicates that these results were measured with cylinders which went from wall to wall. So an experiment was run where a 5.5 mm rod that went from the floor to within about 2 mm of the tunnel ceiling.
The results are as follows; (Max frontal is the reference area)

Original Cd = 0.86, length =19.8 cm, diameter = 5.5 mm , Ref area = 11.44 cm2, V = 19.2m/s, Pressure = 177.6 Pa, Drag =0.175 N, rho = .0967 kg/m3, Re = 6,000

Wall to Wall Cd = 0.95, length = 24.3 cm, diameter = 5.5 mm, Ref area = 11.44 cm2, V = 19.2 m/s, Pressure = 178.2 Pa, Drag = 0.227 N, rho = 0.967 kg/m3, Re = 6,000

If max frontal area is the reference area, how is it that both rods have the same reference area?

Isn't "max frontal" just the rod length times diameter?

If so, shouldn't it be:

- short rod: (19.8)(0.55) = 10.89 cm**2

- long rod: (24.3)(0.55) = 13.37 cm**2

?

Originally Posted by Coupez

If max frontal area is the reference area, how is it that both rods have the same reference area?

Isn't "max frontal" just the rod length times diameter?

If so, shouldn't it be:

- short rod: (19.8)(0.55) = 10.89 cm**2

- long rod: (24.3)(0.55) = 13.37 cm**2

?

Yes, the longer rod should be 13.37 cm2. This error would stand out if there were a way to embed a table in the text.

A search for detailed drag of a cylinder with flow normal to the cylinder axis found the graph in the attachment. The drag for a Reynolds number of 6,000 (the case in question) is between 0.9 and 1.0, compared to the measured value for the 5.5 mm rod of 0.95. The conclusion is that the measured results are correct for the particular conditions of the tests.

Bill

Four RC wheels ranging in size from 50 to 84 mm in diameter and constructed in different ways were tested for stand-alone drag coefficients. The results are presented in the first attachment.

There is considerable variation of CD that appears to be related to materials used (rubber or foam) and the details of the rims and hubs (spokes, cutouts, sharp edges). The values range from 0.27 to 0.5. All are referenced to the diameter x thickness area.

The second attachment was extracted from Professor Hoerner's book "Fluid-Dynamic Drag", 1965. It shows similar variation for different wheel configurations.

Bill

Some advice is needed on measuring the drag coefficient of a motor nacelle mounted on a wing panel and mounted vertically in the wind tunnel. One set of tests were run on a prototype configuration that may have some flaws.

The configuration is shown in the photo. A 22 mmm span, 19.8 mm chord Clark Y type panel was constructed of balsa and film covering. It is not exactly Clark Y but close to it; it also has film wrinkles and dips between the ribs. The test was not intended to measure the true free-steam Cd of the panel; only to provide a base reference for measuring the nacelle induced extra drag.

The panel was mounted vertically in the tunnel with about 2 cm clearances between the top and bottom tunnel surfaces. The initial alignment was such that the chord line was parallel to the airflow, but there was considerable torque induced. The panel was then oriented to minimize the torque (by letting it weather vane) and a test run. No measurements were made but the AOA appeared to be around – 5 degrees. The results indicated that the drag coefficient was about 0.030 at a Reynolds number of about 200k.

A cylindrical motor nacelle was then added, centered on the panel. The OD is 3.5 cm, the length from front to the LE is 4.5 cm and the overall length from front to termination is 9 cm. The maximum frontal area is 9.6 cm2. The sizing was chosen to house a 28 mm diameter motor mounted to the leading edge. The nacelle front edges are rounded and the junctions with the panel slightly faired. The nacelle axis is aligned with the chord line and slightly above it.

A test was then run as before with the panel allowed to weather vane to a neutral position. The AOA was similar to the first test but was not measured.
The total drag coefficient was 0.036 referenced to the panel area. The corresponding Cd referenced to the nacelle max frontal area is about 0.030.

So here are some questions.

1. What is an appropriate alignment of the nacelle axis with the wing chord line?
2. How does the panel alignment with the air flow affect the results?
3. Is the neutral position a suitable reference alignment?
4. If not, what should the alignment be?
5. Does the presence of the wall proximity to the panel ends compromise the test validity?
6. Would a symmetrical airfoil be a better choice, thus avoiding the alignment issue?
7. Does the imperfections in the panel construction cause any problems?
8. Any other observations ?

Bill

Well, if it's meant to represent a nacelle for a prop plane then the tests will not give comparable results to real life situations, due to effects on the leading airflow of the prop wake, either the swirling and acceleration of the airflow of a running prop, the slowing of a windmilling prop, the wake of a stopped and feathered prop etc. but it would be a good baseline to compare windmilliing, stopped and feathered states. Ideally the profile would reach to both ends of the wall to simulate an infinite length wing. It might be possible to reduce interaction artefacts using wing tip plates, but I am not sure, could be another test. And I think a symmetrical profile would be best.

If you want a meaningful comparison of nacelle on vs. off then you need to have the wing section at the same angle. This requires building a mount than can resist torque.

The drag increment due to the nacelle shouldn't be overly sensitive to the airfoil section, so a symmetrical section would probably be easiest to work with.

Alignment with the air flow is very important, because you're essentially varying angle of attack. If AoA is different than the zero-lift value you'll be generating lift - and induced drag.

Can you mount the model on a shaft with bearings, located well forward of the MAC, and thus allow it to weathervane?

Thanks for the advices. The take-aways are as follows:

The interaction between the nacelle and the airfoil section is not very sensitive.
A symmetrical airfoil is preferred.
At any AOA other than zero there is a potential for induced drag that may obscure the effect of nacelle drag.
Tip plates might minimize induced drag but the amount is uncertain as is how to design such plates.

I plan to do the following:
Build a symmetrical airfoil of length selected to minimize the gap between panel and walls.
Find a way to accurately measure the AOA.
Measure the drag as a function of AOA with this panel.
Design tip plates and install.
Measure the effect of the tip plates on total drag.
Evaluate the effectiveness of the plates.

Any advice on the design of tip plates would be greatly appreciated.

Bill