My Aerodynamics lab report

Force Measurement (FM) Lab – MAE 2404
Aerodynamics 1
Monash University
Department of Aerospace and Mechanical Engineering
Attn: Mr. David Smith (Tutor)
Group F
Conducted at 11am on Monday, 14th September 2015
David MK Chan
[email protected]
Please refer to general report guidelines for content.
In this experiment we use wind tunnels to determine lift and drag forces on three types of
aerofoil: flat, symmetric and cambered. Wind tunnels have been a traditional means for
aerodynamicists to test models and through analysis of diagnostic information, enable them
to predict fluid movement around and performance of the full- scale aircraft.
The aims of this experiment are: to determine the lift coefficient as a function of positive
and negative angle- of- attack; stall characteristics among the three; and drag coefficient as
a function of lift coefficient.
Is lift a linear function of AOA, up to a point? Is stall able to be identified? are some of the
questions posed.
Lift and drag in wind tunnels can be obtained either by force balance or pressure
measurements. Here we use the former approach.
The aforementioned aerofoils are tested, one by one. There is a separate wind tunnel for
each. Drag and lift force measurements are obtained, from AOA settings between -2° and
+22°, in increments of 2. It is possibly useful to mention that parallax error in scale- reading
should be avoided.
Before the wind tunnel is activated, adjustment screws are used to ensure that the lift and
drag scales are set to zero. For negative AOA values, which expectedly produce negative lift,
the lift scale is set to an arbitrary (positive) value, again before the wind tunnel is activated.
A throttle adjacent to the fan housing adjusts the velocity of the ingested air.
Its value is determined using a provided chart, using a “calibrated sloping- tube”
manometer. Velocity2 is a function of the liquid length(cm). Here we obtained 3.5cm, which
gives v2=300ms-1.
The Reynolds number obtained in the experiment environment is Re= 74, 813. This indicates
turbulent flow. The density and viscosity values are based on a temperature value of 20°C,
which is reasonable in the operating environment. Characteristic (chord) length is calculated
based on the provided aspect ratio of 4.63 and wing area 195.65cm2. Kindly refer to related
calculations in the Appendix. We will later discuss the effect of turbulent flow on the
accuracy of aerodynamic measurements.
Results and Discussion:
This is Graph 1, for lift coefficient CL against angle of attack AOA:
As illustrated, the cambered aerofoil stalls at approximately 16°, as does the flat wing. The
symmetric aerofoil has a stall angle of ≈9°. The negative gradient demonstrated at
AOAcambered ≈5° may be attributed to the outlier point preceding it. All the following points
indicate positive gradient till ≈16°.
The symmetrical wing demonstrates a gradual loss of lift following AOAsymmetrical≈8.5°,
although the last three points indicate a slightly positive gradient. The cambered aerofoil
demonstrates a slight negative gradient between ≈16° and 19°, and subsequently a large
loss of lift is experienced. The stalling AOA of the flat and symmetrical wing is less decisive.
This may be owing to the limited number of experimental values taken. It is impossible to
correctly predict the nature of the curves after AOA≈22°.
Nonetheless, we can illustrate that all three curves level out or experience change in
gradient sign after the highest recorded, or critical point, after which flow separation occurs,
which is a precursor to a stall, and that lift is no longer increasing linearly with AOA.
To summarise :
AOAstall, cambered ≈ 16°
AOAstall, symmetric ≈ 8.5°
AOAstall, flat
≈ 16°
Results and Discussion:
We can assert that the “best angle of attack” for each airfoil is the critical one, i.e. that
produces the highest CL and therefore the highest lift (assuming fluid density and velocity,
wing area and chord length remain constant). In practice, on an aircraft for instance, it may
be inadvisable to maintain a critical AOA, especially at low altitude, for existential reasons.
The lift coefficient CL at AOA=0° is zero for the flat wing, ≈0.25 for the symmetrical, and ≈0.5
for the cambered. This implies that the flat wing produces no lift at AOA=0°.
From tables 1.2 and 1.3, Liftsymm,0°= 0.9N and Liftcamb,0°= 1.9N.
These findings are in agreement with MAE2404 Lecture 21 (Thin Airfoil Theory II), which
asserts that thin wings are aerodynamically inefficient.
In that Lecture we also discover that thin wings have a separation point much closer to the
leading- edge than do cambered aerofoils, where it is much farther aft. This means that
stalls in thin wings are abrupt, having little buffeting prior to a full stall.
Unfortunately this finding was not replicated in this experiment, since neither the flat plate
wing nor the symmetrical aerofoil demonstrate a dramatic decrease in CL beyond any point.
This is Graph 2, for drag coefficient CD against lift coefficient CL:
Results and Discussion:
Lift- induced drag occurs whenever any 3-D wing produces lift. When the AOA ≠ 0°, or larger
than ±5° (small angles), the frontal area of the wing and boundary layer thickness increases,
leading to increased drag.
Consulting Tables 1.1, 1.2 and 1.3, we see this effect very clearly with the flat and symmetric
wings. The Drag values increase rapidly above AOAflat ≈ 10° and AOAsymmetrical≈ 8°.
It is much less apparent with the cambered wing, which maintains fairly consistent drag
values from AOAcamb≈0° till the critical AOA.
This theory is borne out experimentally by Graph 2. We see an almost level line for the
Cambered, while those of Flat plate and Symmetrical are very steep beyond a small angle.
High levels of drag are inefficient as they require larger amounts of power to maintain level
flight and constant speed in an aircraft.
Random errors:
On at least one occasion, a drag counter was stuck in a fixed position and recalibrated while
the wind tunnel was in operation midway through the experiment. Nonetheless, once
unstuck, it was reset to its previous position at constant fluid velocity and AOA. Owing to
time constraints, it was deemed imprudent to repeat the data collection. This may explain
the appearance of outlier points.
Systematic errors:
Reference 1) alludes to the possibility of gross inconsistencies in wind tunnel testing with
high turbulence (high Reynolds number).
This would affect all the data in the same scale, possibly causing a deviation between
experimental results and theoretical expectations.
Loose access panels on the sides of the wind tunnel are another possibility. These may
cause flow disruption and therefore affect force measurements on the aerofoil.
This experiment, while inherently simple, reinforced expectations of superior performance
and efficiency of Cambered Aerofoils, relative to the others. The large quantity of data
calculated required thoroughness and commitment to accuracy.
For future purposes, it is proposed that AOA values up to 30° be taken, since this would
more conclusively display critical values while allowing for gradient variation in the graphs.
4) Fundamentals of Aerodynamics, 5th Edition (Anderson)
ISBN 0-07-237335-0
Sample Calculations:
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