Engineering lab report rewrite(paraphrasing)

ME 491

THERMAL FLUIDS LAB

EXPERIMENT:

LIFT AND DRAG MEASUREMENT

TEST PERFORMED: MARCH 24, 2018

REPORT SUBMITTED: MAY 5, 2018

BY

GROUP 1

ABSTRACT

The purpose of this lab was to display the effects of the angle of attack. By using a small wind tunnel and a model wing, the forces of lift and drag were measured in order to further quantify the effects of angle attack. At +10 degrees, the wing had effectively stalled, with the values of lift decreasing dramatically while the drag had increased. At +8 degrees, the maximum ratio of the coefficient of lift and drag had been revealed to be 9. This meant that the force of lift was the greatest in comparison to the drag, making it the most effective angle for this model plane to rise, should it be experiencing air speed around 40 MPH. Overall, the experiment was successful and can be easily repeated in order to verify correct data. The biggest concern was the leak in the pitot tube, which was vital to obtaining the dynamic pressure of the small wind tunnel.

INTRODUCTION

The concept of an airfoil is to produce lift with the least amount of drag. The forces of lift () allows heavy planes to fly. The forces of drag () pull the plane down and seriously limit the speed and fuel consumption. The angle of attack that the aircraft wings experience is due to the angle that the chord of the wing (A) is exposed to as well as the air direction and air speed (V). Should a plane experience too high of an angle of attack, the lift would drop, and the drag would increase exponentially, known as stalling.

(1)

(2)

The value of air density () would remain constant, but the lift and drag coefficients (would change depending on the angle of attack. By measuring the quantities to find the lift and drag coefficients, we will be able to quantify the effects of the viscous forces in the flow around the wing

DESCRIPTION OF WORK

In order to begin measuring the forces that the angle of attack has on the wing, we had to analyze the resolution of each instrument. After, the fan was turned on, we took the pitot tube and measured the dynamic pressure of the system, which was shown on the manometer. We set the angle of attack to -10 degrees and adjusted it up 2 degrees until we had reached the stall point of +10 degrees. We had taken three more measurements past the stall point to verify the sustained change in drag and lift. For each measurement, we would record the lift and drag shown on the force balance

Figure 1: The angle of attack was adjusted by turning the adjusting screw.

Figure 2: The force balance used to find the differences in lift and drag, depending on the angle of attack

Figure 3: The manometer that displayed the readings of the dynamics pressure of the small wind tunnel from the pitot tube. (.72 in-water)

RESULTS AND DISCUSSION

By examining the results, we clearly see that the stall point is at +10 degrees. Also, at +8 degrees, the best ratio of CL/CD was 9. This meant that the benefits of lift far outweigh the effects of the drag on the chord area.

Figure 4: The lift and drag plotted against the angle of attack. The stall point is clearly seen at +10 degrees.

Figure 5: The coefficients of lift and drag plotted on a polar diagram. The maximum ratio was found to be 9 (+8 degrees)

CONCLUSIONS

This experiment clearly shows the effects that the angle of attack has on an air foil. According to the wing shape, the coefficients of lift and drag are important in measuring the viscous forces on the wing and at which angle that the wings will create the most lift. In the case of our experiment, that angle was +8 degrees that gave us a nine-to-one ratio of lift to drag. By quantifying the lift and drag, the most efficient wing shape can be made to advance travel possibilities and limit fuel consumption.

Error that was relevant to this experiment pertained to the leak in the pitot tube. If the dynamic pressure reading was affected drastically, the resulting values based off of calculation would have a severe change. I recommend that instead of trying to cover up the leak, the university should fund a new device.

APPENDIX

Model Wing Data
Angle ()	Lift (lbs)	CL	Drag (lbs)	CD
-10	0.18	0.333	0.09	0.166
-8	0.2	0.370	0.09	0.166
-6	0.31	0.573	0.085	0.157
-4	0.45	0.832	0.085	0.157
-2	0.51	0.943	0.085	0.157
0	0.6	1.109	0.085	0.157
2	0.66	1.220	0.085	0.157
4	0.78	1.442	0.09	0.166
6	0.85	1.572	0.095	0.176
8	0.9	1.664	0.1	0.185
10	0.91	1.683	0.13	0.240
12	0.59	1.091	0.24	0.444
14	0.58	1.072	0.26	0.481
16	0.59	1.091	0.28	0.518

Dynamic Pressure
Inches-water	0.72
Lbf-in^2	0.0259857

Force Variables
Chord Length (in)	2.3125
Chord Width (in)	9
Chord Area (in^2)	20.8125
Air Density (lb-s^2/in^4)	1.15E-07
Air Speed (in/s)	673.347

Design Uncertainty
Instrument	Uo
Lift (lbs)	0.01
Drag (lbs)	0.01
Manometer (in-water)	0.01
Angle (degrees)	1
Uc	0.05
Ud	1.00139902

REFERENCES

Dr. Robert Ryan, Lift and Drag Measurement Experiment

ME 491 Laboratory Manual, 2007 Version

Lift and Drag Vs. Angle of Attack

Lift Coefficient -10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16 0.33282338550235896 0.36980376166928774 0.57319583058739598 0.83205846375589743 0.94299959225668373 1.1094112850078632 1.2203524135086496 1.4422346705102222 1.5716659870944729 1.6641169275117949 1.6826071155952593 1.0909210969243988 1.0724309088409343 1.0909210969243988 Drag Coefficient -10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16 0.16641169275117948 0.16641169275117948 0.1571665987094473 0.1571665987094473 0.1571665987094473 0.1571665987094473 0.1571665987094473 0.16641169275117948 0.17565678679291169 0.18490188083464387 0.24037244508503702 0.44376451400314526 0.48074489017007405 0.51772526633700289

Angle of Attack (Degrees)

Force Coefficients

Polar Diagram 0.16641169275117948 0.16641169275117948 0.1571665987094473 0.1571665987094473 0.1571665987094473 0.1571665987094473 0.1571665987094473 0.16641169275117948 0.17565678679291169 0.18490188083464387 0.24037244508503702 0.44376451400314526 0.48074489017007405 0.51772526633700289 0.33282338550235896 0.36980376166928774 0.57319583058739598 0.83205846375589743 0.94299959225668373 1.1094112850078632 1.2203524135086496 1.4422346705102222 1.5716659870944729 1.6641169275117949 1.6826071155952593 1.0909210969243988 1.0724309088409343 1.0909210969243988

Drag Coefficient

Lift Coefficient

1