Foil Sim Exercise PDF

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FoilSim Exercise

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Exercise Aims • Access through straightforward simulation a basic knowledge on aerodynamics and flight principles • Get acquainted with the main parameters defining an aerofoil and wing • Study the effects of the various geometric and aerodynamic parameters on aerofoil and wing lift and drag

Introduction FoilSim is a Java applet that simulates the aerodynamics of unswept untwisted rectangular wings. The effects on aerodynamic forces of geometrical parameters such as the chord and span of the wing or the camber and thickness of the airfoil can be analysed together with flight parameters like altitude, velocity and angle of attack.

FoilSim Q ⇒ Open the Foil.html applet in a browser equipped with the JAVA plugin1

Applet description The upper-left panel shows the wing embedded in the fluid flow. Several visualisations may be chosen:

View:

Display:

Edge Top Side-3D Find Streamlines Moving Frozen Geometry

Wing visualisation window airfoil view wing planform view three dimensional view of the wing automatic zoom adjustment to fit the wing within the window flow representation through streamlines flow representation through particle tracking freezes particle tracks shows parameter definitions on current view

The top-right panel is meant for user input of aerodynamic models choice, parameter values and required output selection. The Input buttons on the top-right panel select the family of parameters to be edited in the bottom-left panel, while the Output buttons select the output visualisation option for the bottom right panel. The following tables summarise the use of the various buttons: Units Reset

Switch between imperial and metric units systems Reset all values to default

1 The FoilSim applet was developed by NASA’s Glenn Research Center and may be downloaded for free at https://www.grc.nasa.gov/www/K-12/airplane/foil3.html

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Flight:

Shape:

Size:

Analysis:

Select Plot:

Plots:

Probe:

Gages: Geometry: Data:

Input Set flight parameters Flight Test Atmosphere choice Speed Free-stream flow velocity (V ) Altitude Altitude above sea level (h) Press Ambient static pressure (p) Temp Ambient static temperature (T ) Density Static density (ρ) Viscosity Fluid dynamic viscosity (µ) Set airfoil geometric parameters Airfoil Shape Airfoil/Ellipse/Plate... Angle Angle of attack (α) Camber Camber (c) Thick Thickness (t) Set wing geometric parameters Chord Chord (l) Span Span (b) Area Wing projected area (S) Aspec Rat Aspect ratio (λ) Select aerodynamic models Lift Analysis Ideal Flow or Stall Model AR Lift Correction Correction for finite Aspect Ratio (λ) Induced Drag Drag Polar Model Re Correction Correction for Reynolds number Type of plot to be shown on bottom-right panel when selecting Output: Plots Output Select type of plot to be shown in bottom-right panel: Surface: Pressure Pressure distribution along lower and upper surface Surface: Velocity Velocity distribution along lower and upper surface Drag Polar Drag polar (Drag coeff. vs lift coeff.) Lift vs: lift (L) as a function of: Cl vs: lift coefficient (CL ) as a function of: Drag vs: drag (D) as a function of: Cd vs: drag coefficient (C D) as a function of: Angle airfoil angle of attack (α) Camber airfoil camber (c) Thickness airfoil thickness (t ) Speed free stream velocity (V ) Altitude flight height (h) Density fluid density (ρ) Wing Area wing projected area (S) Output indicators Velocity point velocity Pressure point pressure Smoke emission lines Drag and lift indicators Detailed geometric data Detailed flight data

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Finally, the bottom of the top-right panel outputs the current values of Lift (or lift coefficient Cl), Drag (or drag coefficient Cd), Reynolds number and aerodynamic efficiency (L/D ratio). Q ⇒ Initialise the applet with the following basic options: View: Display: Units: Analysis: Flight Test:

Side-3D Geometry Metric Stall Model + AR Off + ID Off + Re Off Earth - Average Day

Airfoil and wing analysis Figure 1 shows the light airplane whose flight will be simulated. It is the Cessna 172 Skyhawk, a widespread airplane of common use in flight schools and leisure general aviation due to its good per formance, high stability and ease of piloting. Its wing planform can be approximated by a rectang le and the dimensions are as follows: Camber Thickness Chord Span

c = 2% l t = 12% l l = 1.49m b = 11 m

Figure 1: Cessna 172 Initialise the simulator with the data above and answer the questions below. Use the space provided to briefly explain how the simulator must be used to extract the desired information. e What are the projected area and aspect ratio of the wing? S=

16.433619

λ=

7.367

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Switch to Edge/Geometry view and check the relation between lift and angle of attack. e For what angle of attack do we get the highest lift? What is the highest lift achievable? αS =

15.12

CL max =

1.793

e What is the angle of attack for no lift? What is the lift for zero angle of attack? α0 =

-2.04

C L0 =

0.248

e Sketch the lift characteristic curve (C L vs. α) indicating all the values computed above. Approximate the curve by a truncated linear model and justify under which circumstances is the model accurate. CL =

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Place the wing at sea level and immersed in a flow with free stream velocity V = 200 km/h. e What is the lift generated by the wing if the angle of attack is α = 5 o? L=

26598 N

Assume the airplane is operating at its maximum take off weight (MTOW) with m = 1000 kg and consider a gravity pull g = 9.8 m/s2. e What must the angle of attack be in order to sustain straight and level flight at a speed V = 200 km/h at sea level? and at an altitude of h = 2000 m? αSL =

αh =

1.1143 °

0.5531°

e How would α be modified for a symmetric airfoil? α jSL =

12.5778 °

αjh =

3.1365 °

Back to the actual cambered airfoil: e What is the stall speed at sea level? and at h = 2000 m? VS SL =

23.31 m/s

V Sh =

25.717 m/s

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We want to make a coordinated turn with bank angle φ = 60 ◦ . e What is the lift required to perform this turn and preserve the altitude? What is the load factor of such a manoeuver? Lφ =

nφ =

2

19600 N

e What are the stall speeds associated with this manoeuver at sea level and h = 2000 m? V SSL =

1118.82

Vh= Sφ

130.95 m/s

e What would the stall speeds be, both at sea level and altitude h = 2000 m, for a straight and level flight (no longer turning) if the airplane’s mass was mj = 750 kg? VSSL mt =

72.74 m/s

VShmt =

77.62 m/s

We want now to engage in a perfect dive, the mass being again m = 1000 kg. e What is the lift required to perform this manoeuver? How must the angle of attack be set to sustain the dive? Ldive =

ON

αdive =

-2.032 °

Fianlly, we want to attempt an inverted flight at sea level and speed V = 200 km /h. e What is the lift required to perform this manoeuver? How must the angle of attack be set to sustain the inverted flight at constant altitude? Linv =

α inv =

9800 N

5.19 °

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Set the angle of attack to α = 2◦, the flow speed to V = 100 km/h and the altitude to sea level. Visualise the plot showing the airflow speed distribution on the upper and lower surfaces of the airfoil. e Sketch and comment the plot.

P1-Upper P2-Lower

e What happens as the angle of attack is increased?

1. When the stage angle increases, the upper surface initially is higher and the decrease is constant. Instead, what happens with the lower surface is that his velocity initially is very high but gradually it decreases. ( Parable in the graphic ).

e Comment on what happens as the angle of attack exceeds α = 4.8 o?

1. The upper Surface increase gradually at the lower percentage of chord but the lower surface has a gradually decrease at that lower chord percentage.

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Visualise now the plot showing the pressure distribution along the upper and lower surfaces of the airfoil. e Sketch the plot and comment its relation with the velocity distribution.

The Lift is created between the the difference of the pressure and at low % this difference is greater. It happens the same with the velocity, at greater speeds, greater difference.

e Describe the effects on pressure and velocity distributions of varying the curvature and thickness of the airfoil.

1. If there’s a variation in the curvature, the difference of the pressure is noticed over the whole chord. 2. Instead, if there’s a variation in the thickness, there’s also a variation in the yoxis but the difference between the upper/lower in this case remains though to the velocity.

e What are the effects on the lift characteristics C L(α )?

1. If we vary the curvature we can see how that the couber % increases with the αS and the same happens with the αS – couber %.

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Turn the model “Aspect Ratio Lift Correction” on: Analysis:

Stall Model + AR On + ID Off + Re Off

e How does this affect the lift characteristics CL ( α)? And the drag polar C D(α)? Describe qualitatively what the effects would be on all previous computations?

In the model “Aspect Ratio Lift Correction”in the CL with α we can see a increase (graphic).

Instead, the CD with α does not vary.

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Finally, turn the model “Induced Drag” on: Analysis:

Stall Model + AR On + ID On + Re Off

e How does this affect the lift characteristics CL ( α)? And the drag polar C D(α)? Describe qualitatively what the effects would be on all previous computations?

In this ‘’case”the CL does not vary.

If we turn the model “Induced Drag”on, we can see that in the angle -2° or higher, the Co tends to increase more gradually.

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Assume now that all the drag comes from the wing and that the fuselage and gear have negligible drag (I know, that’s too optimistic an assumption). e Estimate the best glide angle of the cessna (or, rather, its wing). Compare the values for infinite aspect ratio (AR Off / Id Off) and finite aspect ratio AR On / Id On. Comment the results.

1. The best glide angle, we can see that is 2.96 ° when it’s finite, and when it’s not finite the glide angle is 7.18 °. Also in this “case” (Infinite angle), we can see that the α is very higher and is quite same as αs . Is like a similar approximation in the real life....