Aerodynamics Lab

NEAR ITPB, CHANNASANDRA, BENGALURU – 560 067
Affiliated to VTU, Belagavi

Approved by AICTE, New Delhi
Recognized by UGC under 2(f) & 12(B)
Accredited by NBA & NAAC
DEPARTMENT OF AERONAUTICAL ENGINEERING

V SEMESTER
17AEL57 – AERODYNAMICS LAB
ACADEMIC YEAR 2019 – 2020
LABORATORY MANUAL
NAME OF THE STUDENT :
BRANCH :
UNIVERSITY SEAT NO. :
SEMESTER & SECTION :
BATCH :
Department of Aeronautical Engineering MVJ College of Engineering
VISION AND MISSION OF THE INSTITUTION
Institution Vision:
To become an institute of academic excellence with international standards
Institution Mission:
1. Impart quality education along with industrial exposure

2. Provide world class facilities to undertake research activities relevant to industrial & professional needs
3. Promote entrepreneurship & value added education that is socially relevant with economic benefits.
Program Outcome:
1. Engineering knowledge: Apply the knowledge of mathematics, science, engineering

fundamentals, and an engineering specialization to the solution of complex engineering
problems.
2. Problem analysis: Identify, formulate, research literature, and analyze complex engineering
problems reaching substantiated conclusions using first principles of mathematics, natural
sciences, and engineering sciences.
3. Design / development of solutions: Design solutions for complex engineering problems
and design system components or processes that meet the specified needs with appropriate
consideration for the public health and safety, and the cultural, societal, and environmental
considerations.
4. Conduct investigations of complex problems: Use research-based knowledge and
research methods including design of experiments, analysis and interpretation of data, and
synthesis of the information to provide valid conclusions.
5. Modern tool usage: Create, select, and apply appropriate techniques, resources, and
modern engineering and IT tools including prediction and modeling to complex engineering
activities with an understanding of the limitations.
6. The engineer and society: Apply reasoning informed by the contextual knowledge to
assess societal, health, safety, legal and cultural issues and the consequent responsibilities
relevant to the professional engineering practice.
7. Environment and sustainability: Understand the impact of the professional engineering
solutions in societal and environmental contexts, and demonstrate the knowledge of, and need
for sustainable development.
8. Ethics: Apply ethical principles and commit to professional ethics and responsibilities and
norms of the engineering practice.
AERODYNAMICS LAB 2
(17AEL57) V Semester
9. Individual and team work: Function effectively as an individual, and as a member or
leader in diverse teams, and in multidisciplinary settings.
10. Communication: Communicate effectively on complex engineering activities with the
engineering community and with society at large, such as, being able to comprehend and write
effective reports and design documentation, make effective presentations, and give and receive
clear instructions.
11. Project management and finance: Demonstrate knowledge and understanding of the
engineering and management principles and apply these to one’s own work, as a member and
leader in a team, to manage projects and in multidisciplinary environments.
12. Life-long learning: Recognize the need for, and have the preparation and ability to engage
in independent and life-long learning in the broadest context of technological change.
VISION AND MISION OF THE DEPARTMENT
Department Vision
To have an international standing for imparting quality technical education in the field of
aeronautical engineering and technology, to a more sustainable and socially responsible
future.
Department Mission
Knowledge & Innovation: The department aims in dissemination of knowledge to develop

innovative solutions to the various problems in Aeronautical Engineering and related fields.
Professional Skills: To mould students in to successful Aeronautical Engineers by

maintaining best teaching and learning environment in which faculty grow professionally and
students receive unsurpassed knowledge, skills, insights and tools for lifelong learning.
Research Inculcation: To inculcate the state-of-the-art technologies and R & D to design the
next generation of high performance, efficient air & space transportation.
Socio-Ethical responsibility: To nurture Aeronautical Engineers to be sensitive to ethical,

societal and environmental issues while conducting their professional work.
PROGRAM EDUCATIONAL OBJECTIVE (PEO)
PEO-01: Employability & Skills: Graduates will be successful Aeronautical Engineers in

Industry, Research and Academic sectors by applying the basic principles of Mathematics,
Science and Engineering, with high quality communication and interpersonal skillsto work
effectively in multidisciplinary teams, both as team members and as leaders.
AERODYNAMICS LAB 3
PEO-02: Professional Development: Graduates will be able to synthesize data & derive
technical specifications and also design and develop innovative solutions to the various
problems in Aeronautical Engineering by engaging in lifelong learning and professional
development.
PEO-03: Social & Ethical values: Graduates will use modern engineering techniques, skill
and tools with high degree of professional ethics and standards, to fulfill the societal and
personal needs.
PROGRAM SPECIFIC OUTCOMES (PSO)

PSO-1: DGCA/FAR/FAA/MIL/JAR/DEFSTD-regulations: Use the standard government
regulations/specifications for design, manufacturing and testing purposes of civil and military
aircrafts and perform various maintenance related works and synthesize information / data
from various sources of Aircraft operations.
PSO-2: Design, Development & Manufacturing of aircraft and related systems: Carry
out the preliminary design and development of aircraft and manufacturing of various systems
involved and Predict performance characteristics along with the stability analysis.
COURSE OBJECTIVES
This course will enable students to
 CALIBRATE THE WIND-TUNNEL FOR VARIOUS MOTOR RPM.

 ANALYZE THE RESULTS OF SMOKE AND TUFT FLOW VISUALIZATION
 SKETCH THE PRESSURE DISTRIBUTION AROUND DIFFERENT AIRFOILS AND CIRCULAR
CYLINDERS USING PITOT-STATIC PROBES.
 ESTIMATE THE DRAG COEFFICIENT FOR 2-D OBJECTS USING PITOT-STATIC WAKE
SURVEY METHOD.
 PREDICT THE BOUNDARY LAYER VELOCITY PROFILE ON WIND TUNNEL WALL USING
PITOT-STATIC WAKE SURVEY METHOD.
COURSE OUTCOMES
CO307.1 APPLY THE FLOW VISUALIZATION TECHNIQUES.

CO307.2 ESTIMATE THE PRESSURE DISTRIBUTION OVER THE BODIES.
CO307.3 CALCULATE THE LIFT AND DRAG
DRAG COEFFICIENT ESTIMATION FOR 2-D OBJECTS USING PITOT-STATIC WAKE SURVEY
CO307.4
METHOD
BOUNDARY LAYER VELOCITY PROFILE PREDICTION IN WIND TUNNEL WALL USING PITOT-
CO307.5
STATIC WAKE SURVEY METHOD
CONTENTS
AERODYNAMICS LAB 4
RBT PAGE
SI. NO NAME OF THE EXPERIMENT
LEVEL NO.
Calibration of a subsonic wind tunnel: test section static pressure and L1, L2,
1 6
total head distributions. L3, L4
Smoke flow visualization studies on a two-dimensional circular L1, L2,
2 9
cylinder at low speeds. L3, L4
Smoke flow visualization studies on a two dimensional airfoil at different L1, L2,
3 12
angles of incidence at low speeds L3, L4
Smoke flow visualization studies on a two dimensional multi element L1, L2,
4 14
airfoil with flaps and slats at different angles of incidence at low speeds L3, L4
Tuft flow visualization on a wing model at different angles of incidence L1, L2,
5 at low speeds: identify zones of attached and separated flows. L3, L4, 16
L5
Surface pressure distributions on a two-dimensional smooth circular L1, L2,
6 cylinder at low speeds and calculation of pressure drag. L3, L4, 18
L5
Surface pressure distributions on a two-dimensional rough circular L1, L2,
7 21
cylinder at low speeds and calculation of pressure drag. L3, L4
Surface pressure distributions on a two-dimensional symmetric airfoil L1, L2,
8 24
L3
Surface pressure distributions on a two-dimensional cambered airfoil at L1, L2,
9 26
different angles of incidence and calculation of lift and pressure drag. L3
Calculation of total drag of a two-dimensional circular cylinder at L1, L2,
10 29
low speeds using pitot-static probe wake survey. L3
. Calculation of total drag of a two-dimensional cambered airfoil at L1, L2,
11 31
low speeds at incidence using pitot-static probe wake survey. L3
. Measurement of a typical boundary layer velocity profile on the L1, L2,
12 tunnel wall (at low speeds) using a pitot probe and calculation of L3 33
boundary layer displacement and momentum thickness.
Calculation of aerodynamic coefficients forces acting on a model L1, L2,
13 aircraft using force balance at various angles of incidence, L3 35
speed.
Measurement of a typical boundary layer velocity profile on the airfoil L1, L2,
14 40
at various angles of incidence from leading edge to trailing edge L3
AERODYNAMICS LAB 5
Experiment No:1 Date:
Calibration of a subsonic wind tunnel: test section static pressure

and total head distributions.
Aim:
To calibrate the subsonic wind tunnel by preparing a calibration chart drawn between motor
speeds (RPM) and test section velocity.
Equipment:
Subsonic Wind Tunnel, Manometer, Pitot - static tube mounted just ahead the Test Section
Theory:
The calibration of wind tunnel is done to measure the tunnel speed, which can be measured
through Pitot-static tube. The tunnel speed is the mean speed at the test section when the tunnel
is empty. The tunnel speed is measured in terms of the difference between a total head and a
static pressure reading. Calibration also ensures the uniformity of flow parameters in the region to
be used for model testing. A Pitot-static tube is mounted just ahead of test section for this
purpose.
Procedure:
1. Check the wind tunnel for any loose parts.
2. Check the level of manometer liquid in multitube manometer.
3. Connect the total and static pressure port of the Pitot-static tube to the manometer tubes
for measuring the total and static pressure.
4. Note down the initial reading in the manometer.
5. Run the tunnel at different speeds and note down the manometer reading in multitube
manometer.
6. Calculate the velocity of air flow as per the formula given below.
7. Draw the calibration chart between Flow velocity and Motor Speed.
Formula:
1
pT − p s= ρV 2
2
Where, PT - Total Pressure of tunnel, N/m2
Ps - Static pressure of tunnel, N/m2
ρ - Density of Air, Kg/m3
V - Velocity of Air Flow
As this is because it is a open tunnel, therefore

PT = Pa = P
pa =ambient pressure
(The total pressure in the tunnel remains constant irrespective of speed of flow in Tunnel; only the
static pressure of the tunnel varies)
2 ( pT − pS ) 2 Δp
= =V 2 ⇒ V =3 . 77 √ hT −hS =3 . 77 √ Δh
ρ ρ
Since
p − p = Δp=ρ
T s alchol gΔh , ∆h=manometer differential height in mm
Density of air at Bangalore is taken as 1.1x103 Kg/m3. The density of alcohol
AERODYNAMICS LAB 6
(
ρalchol ) is 800 kg/m3.
2 ρ alchol gΔh
V=
√ ρ
=
√2×0 . 8×9. 81×Δh
1 .1×10 3
=3 . 77 √ Δh
Tunnel static pressure concept
Inclined Manometer:
V =3.77 √ Δh
V =3.77 √ Δhsin θ
For inclined manometer case, where θ is inclination of
manometer (θ=30°)
Tabular Column:
Flow
Pitot -static Velocity
manometer V=3.77 √ Δh
Sl. Speed reading m/sec
No. (RPM)
∆
h0 h∞ h
1
2
3
AERODYNAMICS LAB 7
Flow
Inclined Velocity
manometer V=3.77 √ Δhsinθ
Sl. Speed reading m/sec
No. (RPM)
∆
h0 h∞ h
1
2
3
Graph:
Speed Vs Velocity
Result:
Thus the wind tunnel is calibrated by using the pitot-static probe by measuring the total and
static pressure.
Note:-Static pressure plot wrt speed may be drawn for study purpose
AERODYNAMICS LAB 8
Experiment No: 2 Date:
Smoke flow visualization studies on a two dimensional

circular cylinder at low speeds
Aim:
To carry out the smoke flow visualization on a two dimensional circular cylinder and to draw the
flow pattern observed at different speeds.
Equipment:
Subsonic wind tunnel, Circular cylinder model with support mount, Smoke generation apparatus,
liquid paraffin, manometer.
Theory:
In general, flow visualization is an experimental means of examining the flow pattern around a
body or over its surface. The flow is "visualized" by introducing dye, smoke or pigment to the
flow in the area under investigation. The primary advantage of such a method is the ability to
provide a description of a flow over a model without complicated data reduction and analysis.
Smoke flow visualization involves the injection of streams of vapor into the flow. The vapor
follows filament lines (lines made up of all the fluid particles passing through the injection point).
In steady flow the filament lines are identical to streamlines (lines everywhere tangent to the
velocity vector). Smoke-flow visualization can thus reveal the entire flow pattern around a body.
Procedure:
1. Mount the circular cylinder model with its support in the tunnel test section securely.
2. Ensure that the tunnel is not having any loose components.
3. Generate the smoke for the flow visualization through smoke generator.
4. Adjust the amount of smoke generated by adjusting heater control provided with smoke
generator.
5. Observe the flow pattern around the body and infer the location of stagnation point, flow
separation, formation of eddies and vortex shedding nature at different speeds (at
different Reynolds Number).
6. Tabulate the observed flow pattern at different Reynolds Number with a neat sketch with
inference.
7. Gradually shutdown the tunnel.
A typical image of smoke past a circular cylinder is shown below:
Smoke past a cylinder
AERODYNAMICS LAB 9
Reynolds number has influence on nature of flow past objects as shown below:
Formula:
1. Velocity:
2 ( PT −PS )
=V 2 ⇒ V =3 . 77 √h t −h s
ρ
2. Reynolds Number:
ρVD
Re=
μ
where ρ - Density of air=1.2 Kg/m3
V - Velocity of air in m/s
D - Diameter of cylinder=0.050 m
μ - Dynamic viscosity of air = 1.7e-5 Ns/m2
Tabular Column:
Manometer Velocity= Re. No.

Speed* ρVD Flow
Sl. No. Diff height 3.77√ Inference
(RPM)
∆h Δh
Re= pattern
μ
1.
2.
3.
* Maintain a very Low Speed for Flow Visualization.
Result:
AERODYNAMICS LAB 10
Thus the flow visualization is carried out and the flow pattern around the body at different
Reynolds number and velocity is observed
AERODYNAMICS LAB 11
Smoke flow visualization studies on a two dimensional airfoil at

different angle of incidence at low speeds
Aim:
To carry out the smoke flow visualization on a two dimensional airfoil and to draw the flow
pattern at different angle of incidence.
Equipment:
Subsonic wind tunnel, two dimensional airfoil model with support mount, Smoke generation
apparatus, liquid paraffin, manometer.
Theory:
Smoke flow visualization involves the injection of streams of vapor into the flow. The vapor
A picture below shows smoke flow patterns over aerofoilat;a) zero incidence and b) higher angle
of attack.
Note- the separated flow at higher angles of attack.
(a) Low Incidence b) High Incidence

Smoke flow past at aerofoil at different incidences
Procedure:
1. Mount the airfoil model with its support in the tunnel test section securely.
generator.
5. Observe the flow pattern around the body and infer the location of stagnation point,
flow separation, formation of eddies and vortex shedding nature at different speeds (at
different Reynolds Number).
AERODYNAMICS LAB 12
inference.
Formula:
1. Velocity:
2 ( PT −PS )
=V 2 ⇒ V =3 . 77 √ Δh
ρ
2. Reynolds Number:
ρVD
Re=
μ
V - Velocity of air, m/s
D - Chord length=0.150 m
μ - Dynamic viscosity of air, 1.7e-5 Ns/m2
Tabular Column:
Angle of attack=………
Speed* ρVD Flow
(RPM)
∆h Δh
Re= pattern
μ
1.
2.
3.
* Maintain a very Low Speed for Flow Visualization.

Vary angle of attack and take more readings
Result:
Reynolds number and velocity is observed.

Smoke flow visualization studies on a two dimensional multi
element airfoil with flaps and slats at different angle of incidence
at low speeds
Aim:
To carry out the smoke flow visualization on a two dimensional multi element airfoil with flaps
and slats and to draw the flow pattern at different angle of incidence.
Equipment:
Subsonic wind tunnel, two dimensional multi element airfoil model with support mount, Smoke
generation apparatus, liquid paraffin, manometer.
Theory:
Smoke flow visualization involves the injection of streams of vapour into the flow. The vapour
AERODYNAMICS LAB 13
Procedure:
1. Mount the multi element airfoil model with its support in the tunnel test section securely.
generator.
5. Observe the flow pattern around the body and infer the location of stagnation point, flow
separation, formation of eddies and vortex shedding nature at different speeds (at different
Reynolds Number).
inference.
Formula:
1. Velocity:
2 ( PT −PS )
=V 2 ⇒ V =3 . 77 √ Δh
ρ
2. Reynolds Number:
ρVD
Re=
μ
D - Chord length=0.150 m
Tabular Column:
Angle of attack=………

Speed* ρVD Flow
(RPM)
∆h Δh
Re= pattern
μ
1.
2.
3.
Result:
Reynolds number and velocity is observed.
AERODYNAMICS LAB 14
Tuft flow visualization on a wing model at different angles of

incidence at low speeds:
Identify zones of attached and separated flows.
Aim:
To carry out the tuft flow visualization on a two dimensional wing model at different angles of
incidence.
Equipment:
Subsonic wind tunnel, a wing model with support mount, Tufts, Scotch tape.
Theory:
Tuft flow visualization is a type of flow visualization and the tufts readily show where the flow is
steady and where the flow is unsteady. Regions of complete separation and buffeting flow are
readily identified. Tufts are light, flexible material that will align with the local surface flow. The
most commonly used material is light yarn, and the weight and length are chosen according to
model size and test speeds
Procedure:
1. Cut equal sized tufts and place them at equidistant along the span of the wing.
2. Mount the tuft attached wing model to the test section with the help of supporting
mount.
3. Ensure for any loose parts in the tunnel and run the tunnel at low speeds
4. Observe the flow pattern and gradually increase the speed there by varying the Reynolds
number.
5. Observe the change in the flow pattern
6. Tabulate the inference of flow pattern at various Reynolds number.
Formula:
1. Velocity:
2 ( PT −PS )
=V 2 ⇒ V =3 . 77 √ Δh
ρ
2. Reynolds Number:
ρVD
Re=
μ
D - Cord Length=0.150 m
Tabular column:
Manomete Re. No.

Speed Velocity=
Sl. r Diff ρVD Flow
* 3.77√ Inference
No.
(RPM)
height
Δh
Re= pattern
∆h μ
1
AERODYNAMICS LAB 15
Typical Test set up of Tufts is shown below:
Result:
Thus the tuft flow visualization is conducted and the flow pattern at different angles of incidence
is observed.
AERODYNAMICS LAB 16
Surface Pressure distributions on a two-dimensional smooth
circular cylinder at low speeds and calculation of pressure drag
Aim:
To Measure the pressure distribution on a two-dimensional smooth circular cylinder and to
estimate the drag of the cylinder.
Equipment:
Low speed wind tunnel, Multi tube manometer, Cylinder model with pressure tapings and with
support mount, Pitot - static tube.
Diagram:
Ideal flow and Actual static pressure distribution over a circular cylinder
Theory:
There are various methods by which the drag of the bluff body can be measured. One such
method is estimating the drag of the body by measuring the pressure distribution over the body.
Here the pressure distribution over the cylinder is measured which comes from the pressure force
created by the free stream flow over the cylinder. Then in turn by suitable formula the drag
generated by the cylinder is calculated.
Procedure:
1. Assemble the cylinder with pressure tapings in the test section with the help of support.
Connect the pressure tapping to manometer.
2. Rotate the cylinder such that the static holes form the upper or lower surface of the
cylinder.
3. Ensure the tunnel for any loose components and start the tunnel.
4. Run the tunnel at various desired speeds and note down the manometer reading which
measures the surface pressure distribution of the cylinder.
5. Also note down the Pitot-Static tubes manometer reading.
6. Since the cylinder is axially symmetric the pressure distribution is measured for half the
surface and the same trend follows for another half portion.
7. Gradually shut down the tunnel.
Data Reduction:
1. Dynamic Pressure,q:
1
q= pT − p S = ρV 2
2
ps = tunnel static pressure
AERODYNAMICS LAB 17
2. Pressure Coefficient:
pi −Ps ρal gΔh i
C p= =
q q
Where, pi= Static pressure values measured around cylinder
ps= Tunnel static pressure
∆hi =hi - h∞, manometerdifferential column height(wrt tunnel static)
3. Pressure Coefficient (theoretical value):
C p=1−4 Sin 2 θ
Where , θ= Angular location of static ports around the cylinder
4. DragCoefficient:
We do not have continuous pressure distribution; therefore we evaluate this with a numerical
summation.
Tabular Column:
Differential
height Cp Interval
Dynamic Differential Experimen
∆hi between
Sl. pressure pressure tal θ Cpcosθ ×
q= ∆P = ports
No ΔP (rad) dθ
dθ
h ρal g (ho −h∞ ) ρal g (hi−h∞ ) C p=
h0 hi q
∞
1
2
3
CD= ∑Cpcosθdθ =
Interval Cp
betwee Theoretical
Sl. θ
No (rad)
n ports C p= Cpcosθ × dθ
dθ
1−4 Sin2 θ
1
2
3
CD= ∑Cpcosθdθ
Graph:
Cp Vs θ theoretical and compare with experimental values
Results:
1. The pressure distribution around the cylinder is measured and the drag of the cylinder is
estimated.
2. The coefficient of drag of cylinder,CD=
AERODYNAMICS LAB 18
AERODYNAMICS LAB 19
Surface Pressure distributions on a two-dimensional rough circular
cylinder at low speeds and calculation of pressure drag
Aim:
To Measure the pressure distribution on a two-dimensional rough circular cylinder and to
estimate the drag of the cylinder.
Equipment:
Low speed wind tunnel, Multi tube manometer, Rough cylinder model with pressure tapings and
with support mount, Pitot - static tube.
Diagram:
Ideal flow and Actual static pressure distribution over a circular cylinder
Theory:
There are various methods by which the drag of the bluff body can be measured. One such
method is estimating the drag of the body by measuring the pressure distribution over the body.
Here the pressure distribution over the cylinder is measured which comes from the pressure force
created by the free stream flow over the cylinder. Then in turn by suitable formula the drag
generated by the cylinder is calculated.
Procedure:
1. Assemble the rough cylinder with pressure tapings in the test section with the help of
support. Connect the pressure tapping to manometer.
2. Rotate the rough cylinder such that the static holes form the upper or lower surface of the
cylinder.
4. Run the tunnel at various desired speeds and note down the manometer reading which
measures the surface pressure distribution of the rough cylinder.
6. Since the cylinder is axially symmetric the pressure distribution is measured for half the
surface and the same trend follows for another half portion.
Data Reduction:
1. Dynamic Pressure, q:
1
q= pT − p S = ρV 2
2
AERODYNAMICS LAB 20
ps = tunnel static pressure
2. Pressure Coefficient:
C p= =
q q
Where, pi= Static pressure values measured around cylinder
∆hi =hi - h∞, manometerdifferential column height(wrt tunnel static)
3. Pressure Coefficient (theoretical value):
2
C p=1−4 Sin θ
Where , θ= Angular location of static ports around the cylinder
4. DragCoefficient:
summation.
Tabular Column:
Differential
height Cp Interval
Dynamic Differential Experimen
∆hi θ between
Sl. pressure pressure tal Cpcosθ ×
q= ∆P = (ra ports
No ΔP dθ
d) dθ
h ρal g (ho −h∞ ) ρal g (hi−h∞ ) C p=
h0 hi q
∞
1
2
3
CD= ∑Cpcosθdθ =
Interval Cp
betwee Theoretical
Sl. θ
No (rad)
n ports C p= Cpcosθ × dθ
dθ
1−4 Sin2 θ
1
2
3
CD= ∑Cpcosθdθ
Graph:
Cp Vs θ
Results:
1. The pressure distribution around the cylinder is measured and the drag of the cylinder is
estimated.
AERODYNAMICS LAB 21
2. The coefficient of drag of cylinder, CD=
AERODYNAMICS LAB 22
Surface Pressure distributions on a two-dimensional
Symmetric airfoil
Aim:
To Measure the pressure distribution on a two-dimensional symmetric airfoil at low speeds at
different angle of attacks.
Equipment:
Low speed wind tunnel, Multi tube manometer, wing model with pressure tapings and with
Theory:
A symmetric airfoil is one which has same shape on both sides of the chord line i.e. the chord line
and camber line for the symmetric airfoil coincides. The pressure distribution and shear stress
distribution over the airfoil generates the aerodynamic forces. For a symmetric airfoil no lift is
produced for zero angle of attack.
Procedure:
1. Assemble the wing model with pressure tapings in the test section with the help of
support.
2. Rotate the wing model such that the chord line is horizontal, thereby keeping the wing at
zero angle of incidence.
4. Run the tunnel at a desired speed and note down the manometer reading which
measures the surface pressure distribution.
7. Again repeat the experiment for various angles of attack and tabulate the readings.
Port points and nomenclatures

C p= =
q q
Where, pi = Static pressure values measured around cylinder (note it is also the
stagnation pressure at the leading edge and is equal to tunnel total)
AERODYNAMICS LAB 23
∆hi = manometer differential column height wrt tunnel total
1
q= ρV 2
2
Tabular Column:
Angle of attack:……. Speed:………(RPM)
Dynamic Differential Cp
Differential pressure pressure Experimental Chord wise
height q= ∆P = ΔP % location of
Port. ∆hi C p=
ρal g (ho −h∞ ) ρal g (hi−h∞ ) port
No q
Graph:
Cp Vs X/C
Result:
Explain the nature of pressure distribution

Surface Pressure distributions on a two-dimensional cambered airfoil at
different angles of incidence and calculation of lift and pressure drag
Aim:
The purpose of the experiment is to measure the surface pressure distribution and calculate the
aerodynamic coefficients from those pressure measurements for a cambered airfoil at a specified
Reynolds number.
Equipment:
Low speed wind tunnel, Multi tube manometer, wing model with pressure tapings and with
Theory:
An airfoil is a two dimensional cross section of a wing, sliced in the general direction of the flow.
The airfoil displays the aerodynamic shape used to produce a pressure imbalance. The net force
of the pressure imbalance (in a real fluid, frictional forces are also present), summed over the
wing, is resolved into lift and drag. By definition, lift is the net force component perpendicular to
the flow and drag is the net force component parallel to the flow.
The curvature in the airfoil shape is called camber. Note that if the upper and lower surfaces are
identical in shape, the mean camber line and the chord line coincide and the airfoil is symmetric.
A cambered airfoil will produce lift, even at α= 0 degree.
Procedure:
AERODYNAMICS LAB 24
1. Assemble the cambered wing model with pressure tapings in the test section with the
help of support.
2. Rotate the wing model such that the chord line is horizontal, thereby keeping the wing at
zero angle of incidence.
4. Run the tunnel at a desired speed and note down the manometer reading which
measures the surface pressure distribution.
7. Again repeat the experiment for various angles of attack and tabulate the readings.
Port points and nomenclatures

p i− ps Δhi
Cp= =
i q q
where, pi = Static pressure values measured around cylinder
ps = Tunnel staticl pressure
∆hi = manometer differential column height wrt tunnel total
The lift and drag force are given as:
L=F y =∫ − p ((cos α+ϕ ) ) dA

D=F x =∫ − p ((sin α +ϕ ) ) dA
summation as below.
L=F y =∑ − pi ( cos ( α + φi ) ΔA i
i
D=F x =∑ − pi ( sin( α + φi )ΔA i
i
AERODYNAMICS LAB 25
Tabular Column:
Angle of attack:……. Speed:………
− pi cos( α +ϕi ) − pi sin(α + ϕi ) Chord

wise %
Pressure
Differential ¿ ΔA i ¿ ΔA i location
Coefficien
Height of port
t
Port
No
Cp X/c
∆hi Δh i
¿
q
1
L=F y =∑ − pi ( cos ( α + φi ) ΔA i
i
D=F x =∑ − pi ( sin( α + φi )ΔA i
i
Graph:
Cp Vs X/C for different angles of attack
Lift and Pressure drag versus angle of attack
Result:
AERODYNAMICS LAB 26
Calculation of total drag of a two-dimensional circular cylinder at

low speeds using Pitot-Static probe wake survey
Aim:
To determine the drag of a two-dimensional circular cylinder using Pitot-Static probe wake survey.
Equipment:
Subsonic wind tunnel, two-dimensional circular cylinder, pitot-static probe rake, Multitube
manometer.
Theory:
Drag can be determined experimentally by mounting a model on a balance and measuring the
force directly, it can be determined by integrating a measured static pressure distribution over
theentire surface, or it can be determined from a momentum balance on a control volume which
contains a model. This momentum balance would require velocity measurements both
upstreamand downstream from the model. This is the method which will be utilized in this
experiment.
Undisturbed flow enters the control volume containing the bluff body. When the only flow
disturbance in the control volume is the bluff body, any loss of fluid momentum is realized as a
force on the body. An application of the momentum equation to the control volume will yield
the drag force when analyzed in the stream wise direction.
Diagram:
Velocity profile in the wake region

Procedure:
1. Assemble the cylinder model in the test section securely with the help of support
mounting
2. Place the Pitot-static wake rake behind the cylinder at a distance of 1D from the cylinder
such that the probe is in the wake region of cylinder.
3. Connect the tubing to multitube manometer.
4. Start the tunnel and run at a constant speed
5. Note down the manometer reading and tabulate to find the drag coefficient.
1. Drag coefficient:
C d =2∫
[√ qwake
q freestream
−
q wake
q freestream ] dy
c
AERODYNAMICS LAB 27
Where qwake ¿ (hi-h∞)w (calculated in wake region)

qfreestream ¿ (hi-h∞)f (calculated for free stream)
dy = Lateral spacing between tubes of wake probe rake
c = Circumference of the cylinder
i.e Cd =2×∑XY, where ∑XY represents the entire integral term in the above equation
Note: In this experiment h∞ and p∞ are taken and not hs and ps , since the measurements are for
velocity and not pressure difference against tunnel static.
2. Wake Velocity:
2 ( pi − p∞ ) wake 2
Tabular Column: =V ⇒V =3 .77 √( hi −h∞ )wake
ρ
PitotProbe Wake probe XY=MN× Wake
A MN
Reading reading dy/c velocity
S.
No
hi h∞ hi-h∞ hi h∞ hi-h∞ (hi-h∞)w/(hi-h∞)f √A - A
V
∑XY=
Graph:
dy Vs V
Result:
Thus the drag of the two-dimensional cylinder is measured by the pitot-static probe wake survey
method.
The value of Cd is………….
Calculation of total drag of a two-dimensional cambered airfoil at

low speeds at an incidence using Pitot-Static probe wake survey
Aim:
To determine the drag of a two-dimensional circular cylinder using Pitot-Static probe wake survey.
Equipment:
Subsonic wind tunnel, wing model, pitot-static probe rake, Multitube manometer.
Theory:
Drag can be determined experimentally by mounting a model on a balance and measuring the
force directly, it can be determined by integrating a measured static pressure distribution over
theentire surface, or it can be determined from a momentum balance on a control volume which
contains a model. This momentum balance would require velocity measurements both
upstreamand downstream from the model. This is the method which will be utilized in this
experiment.
AERODYNAMICS LAB 28
Undisturbed flow enters the control volume containing the bluff body. When the only flow
disturbance in the control volume is the bluff body, any loss of fluid momentum is realized as a
force on the body. An application of the momentum equation to the control volume will yield
the drag force when analyzed in the stream wise direction.
Procedure:
1. Assemble the wing model in the test section securely with the help of support mounting
2. Place the Pitot-static wake rake behind the wing at a distance of 1 chord from the wing
such that the probe is in the wake region of wing.
3. Connect the tubing to multitube manometer.
4. Start the tunnel and run at a constant speed
5. Note down the manometer reading and tabulate to find the drag coefficient.
Formula:
1. Drag coefficient:
C d =2∫
[√ q
qwake
freestream
−
q
q wake
freestream
] dy
c
Where qwake ¿ (hi-h∞)w (calculated in wake region)
qfreestream ¿ (hi-h∞)f (calculated for free stream)
dy = Lateral spacing between tubes of wake probe rake
c = Chord of aerofoil
i.e Cd =2×∑XY, where ∑XY represents the entire integral term in the above equation
Note: In this experiment h∞ and p∞ are taken and not hs andps , since the measurements are for
velocity and not pressure difference against tunnel static.
2. Wake Velocity:
2 ( pi − p∞ ) wake 2
=V ⇒V =3 .77 √ (hi −h∞ )wake
ρ
Tabular Column:
Wake
Pitot Probe Wake probe XY=MN×
A MN velocit
Reading reading dy/c
Sl y
No
hi h∞ hi-h∞ hi h∞ hi-h∞ (hi-h∞)w/(hi-h∞)f √A - A
V
∑XY=
Graph:
dy Vs V
Result:
Thus the drag of the two-dimensional wing is measured by the pitot-static probe wake survey
method.
The value of Cd is………….
AERODYNAMICS LAB 29
Measurement of a typical boundary layer velocity profile on the

tunnel wall (at low speeds) using a pitot probe and calculation of
boundary layer displacement and momentum thickness
Aim:
To determine the boundary layer displacement thickness and momentum thickness by using pitot
probe.
Equipment:
Subsonic wind tunnel, boundary layer rake, multitube manometer
Theory:
Boundary layer thickness (δ): It is defined as the distance from the boundary of the solid body
measured in the y-direction to the point, where the velocity of fluid approximately reaches the
free stream velocity of fluid.
Displacement thickness (δ*) : It is the distance measured perpendicular to the boundary of the
solid body, by which the boundary should be displaced to compensate for the reduction in flow
rate on account of boundary layer formation.
Momentum thickness (θ): Momentum thickness is defined as the distance measured
perpendicular to the boundary of the solid body, by which the boundary should be displaced to
compensate for the reduction in momentum of the flowing fluid on account of boundary layer
formation.
Procedure:
1. Insert the boundary layer probe in the tunnel.
2. Connect the probe to manometer.
3. Connect the total head and pitot static tubes to manometer.
4. Note down the manometer reading and tabulate to get the boundary layer displacement
thickness and momentum thickness.
Formula:
1. Displacement thickness (δ*):
δ
Δh r
¿
δ =∫ 1−
0
( u
U)dy= 1−
( √ )
Δh f
dy
Where, y = distance of elemental strip from the plate

dy = thickness of elemental strip
u = velocity of fluid at the elemental strip
U = free stream velocity
Δhr = height of liquid column from probe rake with respect to
tunnel total
Δhf = height of liquid column from pitot probe with respect to
tunnel total
2. Momentum thickness (θ):
√
δ
Δh r Δhr
θ=∫
0
u
U
u
(
1− dy=
U Δh f )
×¿ 1−
Δh f
dy ¿
( √ )
3. Wake velocity (u):
AERODYNAMICS LAB 30
2 ( pi − p∞ ) wake
=u2 ⇒u=3 .77 √( hi−h∞ )rake
ρ
Tabular Column:
Wake
Probe Probe rake Pitot probe δ* θ
Velocity
No
hi h∞ Δhr = hi- h∞ hi h∞ Δhf =hi - h∞ u
Graph:
1. Probe no. vs U
Result:
The boundary layer displacement thickness is……………
The boundary layer momentum thickness is…………
AERODYNAMICS LAB 31
Calculation of aerodynamic coefficients forces acting on a model

aircraft using force balance at various angles of incidence, speed.
Aim:
To understand the principle and method of calculation of the aerodynamic forces coefficients
acting on a model aircraft using force balance.
Theory:
Force balances are used to directly measure the aerodynamic forces and moments on the model.
As shown in the figure, a wing model is mounted in the tunnel on a bar that passes through the
side of the tunnel. The bar passes over a fulcrum and the force generated by the model
is balanced by weights placed on a plate hung from the end of the bar. With the tunnel turned
off and no air passing through the test section, weights are added to the plate to balance the
weight of the model. The tunnel is then turned on and air flows over the model. The model
generates an aerodynamic lift force. In the figure, the leading edge of the wing model is lower
than the trailing edge. In this case, the wing will generate a lift force that is in the same direction
as the weight of the model. The bar is no longer in balance; the end outside the tunnel is higher
than the previous setting. Additional weights are added to the plate to re-balance the bar. The
amount of the added weight is equal to the magnitude of the aerodynamic lift of the model. If
the model had been mounted with a positive angle of attack, one would have to remove weights
from the plate to bring the bar back into balance.
Figure 1: Simplified one component force balance
There are more sophisticated, three-component balances that can measure lift, drag and pitching

moment. A six-component balance is required to measure all three forces (lift, drag, and side) and
three moments (pitch, roll, and yaw) that determine an aircraft's motion through the air.
In a six component force balance six strain gauges are present. Each gauge measures a force by
the stretching of an electrical element in the gauge. Each electrical element is connected in a
different Wheatstone bridge circuit. The stretching changes the resistance of the element which
changes the measured current through the element according to Ohm's law. The change in
resistance can be measured by the Wheatstone bridge principle.
AERODYNAMICS LAB 32
If the Wheatstone bridge is in balanced

condition i.e. V g = 0, then
R1 R3
=
R2 Rx
Figure 2: Wheatstone bridge circuit

TYPES OF FORCE BALANCE: There are two types of force balance:
1. Internal force balance: In internal force balance, the measuring devices are
placed inside the model.
Figure 3: Internal force balance connected to a model fighter aircraft
As shown in the lower left of the figure 3, an idealized fighter plane model is attached to a
sting and placed in the test section of a wind tunnel. At the top of the figure 3, the details of
the model attachment to the sting are shown. The model is actually attached to a three-
component balance system and the balance system is attached to the sting. The three-
component balance can detect the axial and normal forces and the bending along an axis
perpendicular to the axial and normal axes. From these measurements one can derive
the lift , drag and pitch of the model, but cannot determine the side force, roll, or yaw.
Forces on the model are detected by strain gages located on the balance. Each gage
measures a force by the stretching of an electrical element in the gage. The stretching
changes the resistance of the gage which changes the measured current through the gage
according to Ohm's law. Wires carry electricity to the gages through the hollow sting and
carry the resulting signal back through the sting to recording devices in the control room.
AERODYNAMICS LAB 33
Multiple strain gages are arranged on the balance to account for temperature changes on the
model during the test. A Wheatstone bridge electrical circuit is used to provide temperature
compensation. Because this example uses only a three-component balance, the model must
be aligned with the flow in the tunnel to eliminate the side force, roll, and yawing moments.
We can only vary the angle of attack of the model, as shown at the left bottom of the figure.
With an internal balance, the forces are measured in a co-ordinate system attached to the
model. The resulting measurements must be corrected to produce the lift, drag, and pitching
moment in the tunnel co-ordinate system. When using a sting with an internal balance, the
aft geometry of the model is often modified to accept the sting. Additional tests may be
required to determine the final aircraft drag with the true aircraft geometry.
2. External force balance: In case of external force balance, the measuring devices are
located external to the model and the test section. The location of the device affects the
choice of mounting system for the model.
As shown in the figure 4, an idealized fighter plane model is attached to a platform located
beneath the test section by a two strut mount. There are six strain gauges, labeled A through
F, that are connected to the platform. Each gage measures a force by the stretching of an
electrical element in the gage. The stretching changes the resistance of the element which
changes the measured current through the element according to Ohm's law. The model can
be rotated in pitch and roll by its connections to the struts, and rotated in yaw by the circular
section in the floor of the test section.
Figure 4: External force balance connected to a model fighter aircraft
A test is conducted in the following manner. With the tunnel turned off and no air
passing through the test section, the weight (W) of the model and mounting system is
determined as the sum of the forces from gages A, B, and C. The tunnel is then turned on
AERODYNAMICS LAB 34
and air flows over the model. The model generates aerodynamic forces and moments that
changes the readings on the strain gages.
The lift (L) is given by:
L=A+B+C-W
The drag (Dr) is given by:
Dr = E + D
The side force Y is:
Y=F
If there is no rolling moment, the values of A and B are equal. If there is a rolling
moment ( M r ), the value is equal to:
M r = (A - B) * a / 2
Similarly, the yawing moment, ( M y ) is equal to:
M y = (D - E) * c / 2
The pitching moment, ( M p) is equal to
Mp = C * b
Formula:
1. Lift Coefficient:
2L
C l= 2
ρV S
2. Drag Coefficient:
2D
C d=
ρV2S
3. Side Force Coefficient:
2F
Cf =
ρV2S
where,
ρ - Density of air (1.126 Kg/m3)
S - Planform surface area (m2)
L, D, S - Lift, Drag and Side force respectively (N)
C l, C d, C f - Lift, Drag and Side force coefficient respectively
AERODYNAMICS LAB 35
Measurement of a typical boundary layer velocity profile on the

airfoil at various angles of incidence from leading edge to trailing
edge
Aim:
To measure the boundary layer velocity profile on an airfoil at various angles of incidence from
leading edge to trailing edge
Equipment:
Subsonic wind tunnel, airfoil model, boundary layer rake, multitube manometer.
Theory:
As an object moves through a fluid, or as a fluid moves past an object, the molecules of the fluid
near the object are disturbed and move around the object. Aerodynamic forces are generated
between the fluid and the object. These aerodynamic forces depend in a complex way on the
viscosity of the fluid. As the fluid moves past the object, the molecules right next to the surface
stick to the surface. The molecules just above the surface are slowed down in their collisions with
the molecules sticking to the surface. These molecules in turn slow down the flow just above
them. The farther one moves away from the surface, the fewer the collisions affected by the
object surface. This creates a thin layer of fluid near the surface in which the velocity changes
from zero at the surface to the free stream value away from the surface. This layer is termed as
the boundary layer because it occurs on the boundary of the fluid.
The details of the flow within the boundary layer are very important for many problems in
aerodynamics, including wing stall, the skin friction drag on an object, and the heat transfer that
occurs in high speed flight.
Boundary layers may be either laminar, or turbulent depending on the value of the Reynolds
number. For lower Reynolds numbers, the boundary layer is laminar and the streamwise velocity
changes uniformly as one move away from the wall. For higher Reynolds numbers, the boundary
layer is turbulent and the streamwise velocity is characterized by unsteady swirling flows inside
the boundary layer.
Laminar Boundary Layer Flow:

The laminar boundary is a very smooth flow, while the turbulent boundary layer contains swirls or
"eddies." The laminar flow creates less skin friction drag than the turbulent flow, but is less stable.
Boundary layer flow over a wing surface begins as a smooth laminar flow. As the flow continues
back from the leading edge, the laminar boundary layer increases in thickness.
Turbulent Boundary Layer Flow:

At some distance back from the leading edge, the smooth laminar flow breaks down and
transitions to a turbulent flow. From a drag standpoint, it is advisable to have the transition from
laminar to turbulent flow as far aft on the wing as possible, or have a large amount of the wing
surface within the laminar portion of the boundary layer. The low energy laminar flow, however,
tends to break down more suddenly than the turbulent layer.
Procedure:
5. Mount the airfoil model in the tunnel test section carefully.
6. Connect the boundary layer probe at the 25 % of chord of the airfoil.
7. Connect the probe to manometer.
8. Connect the total head and pitot static tubes to manometer.
9. Note down the manometer reading and tabulate to get the boundary layer profile.
AERODYNAMICS LAB 36
10. Repeat the same procedure to measure the boundary layer profile at 50%, 75% and 100%
of the chord length.
11. Change the angle of incidence, and repeat the same procedure to find out the boundary
layer profile.
Formula:
1. Velocity:
2(Pi −P ∞ )
U ∞=
√ ρair
2 ρalcohol g(hi −hf )

U ∞=
√ ρair
Tabular Column:
Angle of Location of Manometer Reading Boundary

Probe Velocity (U)
incidence probe
hi (mm) h f (mm) ∆h Layer
No. (m/s)
(Degree) (% of chord) (mm) Thickness
1
2
25
3
4
1
2
50
3
4
1
5
2
75
3
4
1
2
3
100
10 1
2
25
3
4
1
2
50
3
4
1
2
75
3
4
100 1
AERODYNAMICS LAB 37
2
3
4
Graph:
1. Boundary layer velocity profile at 5 degree angle of incidence
2. Boundary layer velocity profile at 10 degree angle of incidence
Result:
The velocity profile from leading edge to trailing edge on airfoil at various angles of
incidence was measured successfully.
AERODYNAMICS LAB 38
APPENDIX- 1
OPEN CIRCUIT WIND TUNNEL FACILITY
Introduction:
Schematic of the wind tunnel at the aerodynamic laboratory is shown below:
PARTS
1. Bell mouthed section

2. Honey Comb
3. Settling Chamber, and screen section
4. Contraction cone
5. Test Section
6. Transition (square to circular)
7. Diffuser
8. Fan Duct
9. Motor and Stand
Performance of the Facility: The tunnel top speed is 50m/sec at 1500

r.p.m. The fan is driven by 3 phase AC motor. The speed in the tunnel is
worked out as below:
1
p0 − p= ρV 2
Bernoulli’s theorem: 2 (1)
AERODYNAMICS LAB 39
p0 = Static pressure in the settling chamber, and p= Static pressure in
the test section
ρ = Density of air
v= velocity of air
( p 0−p) , measured value is given by

ρw gh , where
ρw is the density
of the liquid used in the manometer (Methyl alcohol is used at present),
`g` is acceleration due to gravity and `h` is the vertical length of the
liquid column sustaining the pressure measured ( the difference between
tunnel static and total). The total pressure of tunnel remains at
atmospheric pressure at any speed of tunnel, since it is an open type
tunnel.
Density of air at Bangalore is taken as 1.2. The density of alcohol is 800
kg/m3. If `h` is measured in mm of alcohol column, the velocity is given
by the following relationship.
V (m /sec )=3.77 √h( mm) (2)
If the measured liquid column length on the inclined manometer is h m

then h= { hm-hm (initial)}/2 because the inclination is 30 0 to the
horizontal.
Tunnel Specifications:
Test section size = Cross Section: 600mm x 600mm

Length = 2 meters
Maximum Speed = 50 m/sec
Max RPM = 1500
Contraction ratio = 9:1
Contraction length = 1.8m
Test section has four Perspex windows for viewing inside the rest test
section. The top can be opened for easy access to the test section so that
it becomes convenient to set up experiment.
AERODYNAMICS LAB 40
The fabrication of tunnel is done using teak wood and water proof
plywood.
Tunnel Performance:
The table below indicates measured performance of tunnel
R.P.M. Velocity
(Controller) m/sec
100 3.77
200 7.31
300 10.51
500 17.81
700 25.55
900 33.00
1100 39.9
1400 50.78
AERODYNAMICS LAB 41
PRECAUTIONS TO BE TAKEN WHILE RUNNING THE TUNNEL:
1. Do not allow anyone to stand behind the motor while the tunnel is
being run.
2. While staring tunnel motor, see that fan is clear and that no one is
around that area. Only competent people should handle the
controller. Main power must be highly secured.
3. Before starting the tunnel, check whether any loose parts are in the
test section and remove these before start. See that test section is
secured with C clamps.
4. Do not run tunnel below 100 r.p.m., as this will result in heating
up of motor. Intermittent running at lower speeds is allowed. But do
not exceed more than a minute or two.
5. As far as possible do not run the tunnel for long time at higher
speeds.
6. It is recommended that blade angle setting be checked regularly
once in few months. While checking the blade angle setting, check
also the gaps between the blades and the surface of the fan section
of the diffuser. Check also if any blade has become lose.
AERODYNAMICS LAB 42
Smoke Generator
A picture of the smoke generator is shown in the figure 1 and 2, various parts of the
smoke generator are numbered and nomenclature of those parts is given below:
1) Smoke generator module made of glass
2) Heating coil
3) Kerosene or liquid paraffin reservoir jar
4) Silicone tube connecting smoke generator and reservoir jar
5) Traverse to traverse the oil reservoir up and down
6) T connector -1 for by pass of pressurized air
6a) By pass valve
7) T connector-2 connecting pressurized air to reservoir jar as well as smoke
generator module
8) T connector-3 for connecting the pressurized air to the two inlets A and B of
smoke generator module
9) Oil drain flask
10) Smoke collector flask
11) Outlet tube for smoke generator
12) Spike buster extension/junction box with four sockets and individual switches for
these sockets
13) Heater control
14) Centrifugal blower with inlet control disc
AERODYNAMICS LAB 43
Front view of Smoke Generator with variousparts numbered
Back view of the Smoke Generator withvarious parts numbered.

There is an air blower at the bottom of the stand. The air from the blower is
connected to a house pipe. The connection is made such that the pressurized air goes
through a bypass T connector-1. The pressurized air through the systems can be
controlled by opening the bypass valve so that a part of air is bleed out.
OPERATION OF THE SMOKE GENERATOR
Check all the connections of the tubes as shown in figure 1 and 2. Pour liquid paraffin
into the reservoir so that half of it is filled. Then raise or lower the reservoir such that
the liquid level in the bottom tube of the smoke generator modules is about 50mm below
the nozzle outlet. Connect the heater through the heater control which is a 400W
controller. Keep the controller at the minimum and switch on the heater using the
junction box. Slowly increase the heating up to the ½ the capacity. Observe the liquid
paraffin in the tube. It will start slowly boiling. The liquid level increases in the tube and
the bubbles of liquid paraffin start reaching the nozzle exit. At this point of time, turn on
the blower to send pressurized air. The cold air mixes with the vaporized oil and forms
dense smoke. By properly controlling the heating as well the liquid level in the tube, a
good dense white smoke can be generated. The out flow of smoke can be controlled by
the bypass valve as well as the inlet control disc at the inlet to the bowler.
PRECAUTIONS
1. Do not switch on the heater without the liquid paraffin being present in the tube
level indicated already.
2. Unless the smoke is required, do not generate it and allow it to the atmosphere.
Prolonged breathing of the smoke may be very disturbing.
3. Sometimes overheating may not produce the smoke. At these times restart the
smoke generator from low heat again.
AERODYNAMICS LAB 44
APPENDIX-2
The symmetric Aerofoil available for experiment has the following data.
150 mm chord length.
15% thick Aerofoil: NACA 662-015
X,mm Y,mm
0 0
.75 1.683
1.125 2.015
1.875 2.513
3.75 3.353
7.50 4.65
11.25 5.672
15.00 6.537
22.50 7.929
30.00 8.993
37.50 9.815
45.0 10.434
52.50 10.875
60.00 11.145
67.50 11.243
75.00 11.175
82.50 10.925
90.00 10.439
97.50 9.558
105.00 8.364
112.50 6.448
120.00 5.397
127.50 3.795
135.00 2.22
142.50 0.849
150.00 0.
Note: - Source of Data `Abott and Van-Doenoff`- Theory of wings
AERODYNAMICS LAB 45
Location of pressure port holes along the chord the symmetric Aerofoil
Upper Surface Lower Surface
Distance Distance
from from
Hole no. leading X/C x 10 Hole no. leading X/C x 10
edge edge
X (mm) X (mm)
1 0 0 14 6.0 0.4
2 1.5 0.1 15 12.0 0.8
3 3.0 0.2 16 22.5 1.5
4 6.0 0.4 17 40.5 2.7
5 12.0 0.8 18 58.5 3.9
6 22.5 1.5 19 78.0 5.2
7 40.5 2.7 20 97.0 6.5
8 58.5 3.9 21 117.0 7.8
9 78.0 5.2 22 135.0 9.0
10 97.0 6.5
11 117.0 7.8
12 135.0 9
13 150.0 10
AERODYNAMICS LAB 46
APPENDIX- 3
The Cambered Aerofoil available for experiment has the following data.
150 mm chord length, 15% thickness,
Basic Aerofoil: NACA 662-015
CL (Incompressible) 1.0 at 4.560 angle of attack and
Cm at quarter chord point= 0.083
Bottom Upper
X,mm
Y,mm Y,mm
-0.990 0 0
-0.993 0.75 2.373
-1.054 1.125 2.9771
-1.067 1.875 3.959
-0.892 3.75 5.815
-0.611 7.50 8.699
-0.412 11.25 10.933
-0.296 15.00 12.779
-0.243 22.5 15.615
-0.373 30.00 17.614
-0.644 37.50 18.986
-1.019 45.00 19.850
-1.466 52.50 20.285
-1.950 60.00 20.340
-2.437 67.5 20.050
-2.901 75.00 19.449
-3.304 82.50 18.547
-3.568 90.00 17.311
-3.51 97.50 15.606
-3.197 105.00 13.532
-2.694 112.50 11.202
-2.072 120.00 8.723
-1.389 127.50 6.201
-.701 135.00 3.740
-0.149 142.50 1.550
0 150.00 0
AERODYNAMICS LAB 47
Note: - Source of Data `Abott and Van-Doenoff`- Theory of wings
AERODYNAMICS LAB 48
Location of pressure port holes along the chordfor the cambered airfoil
Upper Surface Lower Surface
Distance Distance
from from
Hole no. leading X/C x 10 Hole no. leading X/C x 10
edge edge
X (mm) X (mm)
1 0 0 14 6.0 0.4
2 1.5 0.1 15 12.0 0.8
3 3.0 0.2 16 22.5 1.5
4 6.0 0.4 17 40.5 2.7
5 12.0 0.8 18 58.5 3.9
6 22.5 1.5 19 78.0 5.2
7 40.5 2.7 20 97.0 6.5
8 58.5 3.9 21 117.0 7.8
9 78.0 5.2 22 135.0 9.0
10 97.0 6.5
11 117.0 7.8
12 135.0 9
13 150.0 10
AERODYNAMICS LAB 49
APPENDIX- 4
The Wake Rake used has the following data:
DISTANCE FROM CENTER OF TUBE No 1
1 0.0
2 3.7
3 6.7
4 9.2
5 12.7
6 15.7
7 18.7
8 21.7
9 24.7
10 28.2
11 30.7
12 32.7
13 36.2
14 39.7
15 42.7
16 44.7
17 48.2
18 50.7
19 53.7
20 56.7
21 59.7
22 63.2
23 66.7
24 69.2
25 71.7
26 74.7
27 77.7
28 81.2
29 84.5
30 87.5
APPENDIX- 5
AERODYNAMICS LAB 50
The Boundary Layer Rake data for calculation of dy is as below:
Distance from the Wall of RAKE tapings
Distance
Tube no.
(mm)
1 0.3
2 1.0
3 2.0
4 3.0
5 4.5
6 5.5
7 6.5
8 9.0
9 11.5
10 13.5
11 16.0
12 17.5
13 20.5
14 23.0
15 25.5
Rake must be aligned with flow direction and the tube no. (1) Should touch the wall of
tunnel.
AERODYNAMICS LAB 51
QUESTIONS
1. What is angle of attack?

2. What is critical angle of attack?
3. Explain stagnation point.
4. Explain coefficient of pressure.
5. Name any three dimensionless coefficients.
6. What is an Incompressible Flow?
7. Explain Speed of Sound?
8. Define Mach number?
9. What is a Potential Flow?
10. Define Stream Function, Velocity Potential?
11. What is Drag?
12. What is the significance of Coefficient of Drag?
13. Explain different types of Drag?
14. What is Reynolds Number and Explain its significance wrt Drag?
15. What is Wake?
16. Explain Laminar and Turbulent Flow?
17. What is NACA?
18. Explain NACA 4 Digit Series Airfoil?
19. What is Cambered Airfoil?
20. What is Lift and Drag and Explain significance of Lift and Drag coefficient?
21. What is Center of Pressure and Aerodynamic Center?
22. What is a Wind Tunnel?
23. Explain Different Flow Regimes?
24. What is Continuum and Free Molecular Flow? Explain its significance?
25. What do you mean by Oblique Shock wave and Expansion Fan/Wave?
26. What is Pitot Static Tube?
27. What is Static Pressure?
28. What is streakline flow?
29. What is the significance of honeycomb structure?
30. What are the different types of drag?
AERODYNAMICS LAB 52

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NEAR ITPB, CHANNASANDRA, BENGALURU – 560 067

Affiliated to VTU, Belagavi

DEPARTMENT OF AERONAUTICAL ENGINEERING

17AEL57 – AERODYNAMICS LAB

ACADEMIC YEAR 2019 – 2020

NAME OF THE STUDENT :

UNIVERSITY SEAT NO. :

SEMESTER & SECTION :

VISION AND MISSION OF THE INSTITUTION

To become an institute of academic excellence with international standards

1. Impart quality education along with industrial exposure

1. Engineering knowledge: Apply the knowledge of mathematics, science, engineering

VISION AND MISION OF THE DEPARTMENT

Knowledge & Innovation: The department aims in dissemination of knowledge to develop

Professional Skills: To mould students in to successful Aeronautical Engineers by

Socio-Ethical responsibility: To nurture Aeronautical Engineers to be sensitive to ethical,

PROGRAM EDUCATIONAL OBJECTIVE (PEO)

PEO-01: Employability & Skills: Graduates will be successful Aeronautical Engineers in

PROGRAM SPECIFIC OUTCOMES (PSO)

This course will enable students to

 CALIBRATE THE WIND-TUNNEL FOR VARIOUS MOTOR RPM.

CYLINDERS USING PITOT-STATIC PROBES.

PITOT-STATIC WAKE SURVEY METHOD.

CO307.1 APPLY THE FLOW VISUALIZATION TECHNIQUES.

Calibration of a subsonic wind tunnel: test section static pressure

As this is because it is a open tunnel, therefore

Tunnel static pressure concept

Experiment No: 2 Date:

Smoke flow visualization studies on a two dimensional

A typical image of smoke past a circular cylinder is shown below:

Smoke past a cylinder

Manometer Velocity= Re. No.

* Maintain a very Low Speed for Flow Visualization.

Smoke flow visualization studies on a two dimensional airfoil at

Note- the separated flow at higher angles of attack.

(a) Low Incidence b) High Incidence

* Maintain a very Low Speed for Flow Visualization.

Experiment No: 4 Date:

Manometer Velocity= Re. No.

Tuft flow visualization on a wing model at different angles of

Manomete Re. No.

Typical Test set up of Tufts is shown below:

Port points and nomenclatures

Experiment No: 9 Date:

Port points and nomenclatures

The lift and drag force are given as:

L=F y =∫ − p ((cos α+ϕ ) ) dA

− pi cos( α +ϕi ) − pi sin(α + ϕi ) Chord

Calculation of total drag of a two-dimensional circular cylinder at

Velocity profile in the wake region

Where qwake ¿ (hi-h∞)w (calculated in wake region)

The value of Cd is………….

Experiment No: 11 Date:

Calculation of total drag of a two-dimensional cambered airfoil at

Measurement of a typical boundary layer velocity profile on the

Where, y = distance of elemental strip from the plate

Calculation of aerodynamic coefficients forces acting on a model

Figure 1: Simplified one component force balance