Investigates subsonic and supersonic air flow, including flow around two dimensional models report aeronautical lab 1

Objectives:

• 1. Pressure distribution along a convergent/divergent (Laval) nozzle with subsonic and
• supersonic airflow
• 2. To compare the experimental results with that calculated from 1-D frictionless flow.
• 3. To comparison of actual and theoretical area ratio of a nozzle at supersonic air velocities
• (Mach numbers).
• 4. Pressures around a two-dimensional model in subsonic and supersonic flow conditions, at
• different angles of incidence Angle.
• 5. Lift coefficient for aerodynamic models in supersonic flow

Description

A compressed air supply induces a flow in the working section of the wind tunnel. This gives a less turbulent and

more stable flow for accurate results and comparison with theory. Students use a delivery valve to allow

compressed air to enter the wind tunnel. The wind tunnel includes two analogue pressure gauges. One measures

the compressed air pressure available from the supply (for reference); the other measures the pressure delivered

to the wind tunnel and includes an analogue transducer that connects to record the pressure.

The working section of the wind tunnel is a convergent divergent nozzle with a removable top part (‘liner’). The

shape of the liner controls the maximum air velocity at the divergent part of the working section. Included are

three different liners.

High optical-quality glass windows (‘portals’) are at each side of the divergent part of the working 

section. 

Spaced at precise intervals along the working section of the wind tunnel are pressure taps. Two extra taps

connect to one of the models when in use. Hg manometer displays the pressures and you must transmits them to

paper for instant recording and calculations of pressure ratios and Mach numbers.

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Demonstrates the thermodynamics and fluid mechanics of the adiabatic expansion of air through subsonic and supersonic nozzles report aeronautical lab 1

Abstract

In this experiment we studies the pressure distribution over a convergent divergent nozzle for subsonic and supersonic flow, the nozzle was inside a reservoir, we calculated the pressure ratios and put them in the isentropic relations to obtain the Mach number .

Introduction

A nozzle is a relatively simple device, just a specially shaped tube through which hot gases flow. Ramjets and rockets typically use a fixed convergent section followed by a fixed divergent section for the design of the nozzle. This nozzle configuration is called a convergent-divergent, or CD, nozzle. In a CD nozzle, the hot exhaust leaves the combustion chamber and converges down to the minimum area, or throat, of the nozzle. 

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Boundary layers report aeronautical lab 1

Abstract

In this experiment we studied the effect of smoothness and roughness of a surface on the boundary layer thickness and Reynolds number. And the relation of friction with the streamlines (effect of viscosity). By the end of this experiment we calculated Reynolds number, boundary layer thickness, and the momentum thickness both experimentally and theoretically and for smooth and rough surfaces.

Introduction

All experimental observations indicate that a fluid in motion comes to a complete stop at the surface and assumes a zero velocity relative to the surface. That is, a fluid in direct contact with a solid “sticks” to the surface due to viscous effects, and there is no slip. This is known as the no-slip condition.

Therefore, the no-slip condition is responsible for the development of the velocity profile. The flow region adjacent to the wall in which the viscous effects (and thus the velocity gradients) are significant is called the boundary layer.

A fluid layer adjacent to a moving surface has the same velocity as the surface.

A consequence of the no-slip condition is that all velocity profiles must have zero values with respect to the surface at the points of contact between a fluid and a solid surface

The x-coordinate is measured along the plate surface from the leading edge of the plate in the direction of the flow, and y is measured from the surface in the normal direction. The fluid approaches the plate in the x-direction with a uniform velocity V, which is practically identical to the free-stream velocity over the plate away from the surface

Results

For the two cases the flow was laminar (Re<Re.cr)

The smooth surface had much higher velocity thus higher Reynolds number

For the smooth surface the boundary layer thickness and the momentum thickness was less the for the rough surface

The values of the momentum thickness experimentally was not very close to the ones calculated theoretically, also for the boundary layer thickness there was a large error

Error in momentum thickness=67.7%, 72.7%

Conclusion

Friction force per unit area is called shear stress, and is denoted by τ. Experimental studies indicate that the shear stress for most fluids is proportional to the velocity gradient

The values of friction coefficient and the shear were as expected, higher values for the rough surface.

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The effect of high lifting devices on aerodynamic forces report aeronautical lab 1

Abstract

In this experiment we studied the effect of a flaps and a slat on a chambered wing on the lift, drag and moment coefficients. Also we studied their effect on flight performance and the effect of other high lift devices and their working principle. We calculated the coefficients of lift, drag and moment for different angles of attack and different flaps angle.

Introduction

High-lift device is a component or mechanism which increase lift beyond that obtainable from the main aircraft components. The device may be a fixed component or a movable mechanism which is deployed when required. Common high lift devices include wing flaps and slats. Leading edge root extensions and boundary layer control systems are less commonly used.

Discussion

The higher the angle of the flaps the more lift we get for large angles of attack and the drag and moment coefficients increases

Also, the higher angle of the flaps with the slot is open the more lift we get for higher angles of attack and the drag and moment coefficients increases

For the same angle of the flaps the values of lift coefficients were higher when the slat is opened for higher angles of attack only

Conclusion

The flaps are used to increase the lift for high angles of attack

The slats and slots are used to increase the stall angle

LERX are used in advanced jet fighter to produce more lift

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Bernoulli’s Equation Applied to A Convergent-Divergent Passage report aeronautical lab 1

Abstract

In this experiment we measured the distribution of the total pressure and the static pressure along

a convergent divergent duct and compared these with Bernoulli’s equation finding the dynamic

pressure then the velocity and compare them with the velocity by the width ratio.

Introduction

This experimental module illustrates Bernoulli’s equation as applied to a convergentdivergent duct.

A Pitot static tube measures both the total pressure and the static pressure independently. 

The tube traverses along the axis of the duct and connects to the Manometer via flexible tubes. A clear scale printed on the duct helps to show the probe position. We should see the constant total pressure while observing the rise and fall of the static pressure. We compare the velocity-area ratio as 

calculated from Bernoulli’s equation to the experimental results.

Discussion

The velocity of the flow begins with an inlet speed of 37 m/s and increases at as it reaches the

throat till it reaches a maximum speed of 53.3 m/s then it decreases gradually as till it reaches 35

m/s at the outlet.

The dynamic pressure also increases as it reaches the throat area at reaches its’ maximum then it

will decrease as it reaches the outlet.There are a variation of the velocity for the lower surface if we see the graph of the velocity over

the duct.

Conclusion

The relation between the dynamic pressure and the velocity is proportional

The relation between the static pressure and the velocity is inversely proportional

The relation between the static pressure and the dynamic pressure is inversely proportional

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Drag measurement on circular Cylinder report aeronautical lab 1

Abstract

In this experiment we measured the drag in a circular cylinder by direct measurement and by a wake survey method at a specific Reynold number. We took several readings using the Pitot tube to find the velocity and with certain relations we plot those with MS excel to gain the values of the drag coefficient either by the slope for the direct measurement and the area under the curve for the wake measurement.

Introduction

In fluid dynamics, drag (sometimes called air resistance, a type of friction, or fluid resistance, another type of friction or fluid friction) refers to forces acting opposite to the relative motion of any object moving with respect to a surrounding fluid. This can exist between two fluid layers (or surfaces) or a fluid and a solid surface. Unlike other resistive forces, such as dry friction, which are nearly independent of velocity, drag forces depend on velocity.

Types of drag

• 1.       Parasite drag (skin friction drag) due to shear stress
• 2.       Pressure drag : due to pressure distribution and normal stress (form drag)
• 3.       Profile drag : drag due to skin friction drag and pressure drag
• 4.       Induced drag : drag due to lift
• 5.       Wave drag : drag due to shock waves
• 6.       Interference drag : drag due to cross flow

In other references the skin friction drag and the pressure drag and the interference drag are called Parasite drag.

 (Found in Wikipedia)

The total drag is the combination of drag due to lift and zero lift drag which affects the velocity of the airplane for determining the gliding speed, maximum range, maximum (L/D).

Discussion

The drag coefficient measured in the direct measurements is 0.626984 which is very close to the value measured in the last experiment (Cd=0.63554)

Error = 1.3% (negligible)

The drag coefficient measured in the wake measurements is 0.05461 which is far from the value measured in the last experiment (Cd=0.63554)

Error= 91.4% (very huge)

Conclusion

The direct method is more accurate the wake measurement method and it is very close to the pressure distribution method.

The wake method is very sensitive which define the large error in addition to the human error and 
calculations.

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Pressure Distribution over an Airfoil using Air Floe Bench (Airfoil with Tappings) report aeronautical lab 1

Abstract

In this experiment we used the pressure distribution to calculate the center of pressure, the lift coefficient, drag coefficient and the moment coefficient for different angles of attack for symmetrical 0020 NACA airfoil using an air flow bench and an airfoil with tappings, and their plots with the angles of attack and compared them with different data such as XFlR5 and standard data from NACA.

Introduction

A finite wing is the sum of an infinite airfoil cross sections in a two dimension. A pressure imbalance in produced over the wing, it’s responsible of generating lift. This pressure distribution is simply the pressure at all points around an airfoil.

The main forces applied over a wing are pressure and shear. Normal and shear force are the resultant of the forces over the wing. The normal force in perpendicular to the chord line and the shear force in parallel to the chord line.

We can derive these forces to obtain the total aerodynamic forces; the lift and the drag. The lift component is perpendicular to the relative wind that makes an angle of attack with the chord line. The drag component is parallel to the relative wind.

Unlike the cambered airfoil, a symmetrical airfoil produces no lift on an angle of attack of zero degree.

“The center of pressure of an aircraft is the point where all of the aerodynamic pressure field may be represented by a single force vector with no moment. A similar idea is the aerodynamic center which is the point on an airfoil where the pitching moment produced by the aerodynamic forces is constant with angle of attack “^[Wikipedia]

For symmetrical airfoil, when angle of attack varies, the pressure distribution will also vary. But for any angle of attack if we consider the resultant lift force position or center of pressure it will lie on the aerodynamic center. It won’t vary with angle of attack. Because aerodynamic center is at constant location and usually lies at 0.25 of the chord.

 Aerodynamic center and Center of Pressure are same in symmetrical airfoil.

Discussion

The moment coefficient increases as the angle of attack increases

The center of pressure should not be changed due to the symmetry of the airfoil, and it should be located at 25% of the chord

Minimum drag coef. was between -5 and 5 degrees (AoA)

Maximum lift coef. was between 5 and 10 degrees for Re= 50,000 / 100,000

Conclusion

The data from the XFlR5 and NACA tools are identical stall angle 10 at cl max 1.1

The data taken showed that the angle of stall was at 20 at cl max 2.7

Error is large 145%, 100%

The data taken from the experiment was hard to solve to get the coefficient, using the xflr is more convenient.

Also human error should be taken into consideration in this experiment.

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Lift characteristics of a low-speed airfoil using pressure distribution report aeronautical lab 1

Abstract

In this experiment we used the pressure distribution to calculate the lift coefficient and drag coefficient for different angles of attack for 0015 NACA airfoil and their plots with the angles of attack and compared them with standard data.

Introduction

A finite wing is the sum of an infinite airfoil cross sections in a two dimension. A pressure imbalance in produced over the wing, it’s responsible of generating lift. This pressure distribution is simply the pressure at all points around an airfoil.

The main forces applied over a wing are pressure and shear. Normal and shear force are the resultant of the forces over the wing. The normal force in perpendicular to the chord line and the shear force in parallel to the chord line.

We can derive these forces to obtain the total aerodynamic forces; the lift and the drag. The lift component is perpendicular to the relative wind that makes an angle of attack with the chord line. The drag component is parallel to the relative wind.

By using a micro-manometer and the pressure distribution over the airfoil we can obtain the pressure coefficient which is used to calculate the lift characteristics of out NACA 0015 airfoil.

Discussion

The values for the lift coefficient for the small angles of attack are not accurate but acceptable as experimental data for higher angles of attack

Our data show a stall angle of 10 degrees at Cl max around 0.22, however the standard data shows about 14 degrees at cl max around 1.2

The distance between the curves is small for small angles (must be zero for angle zero) and increase as AoA increase.

Cd max is at Cl max as seen in the graphs

Conclusion

Large errors are found in this experiment mainly human errors in reading the data from the manometer and the lack of accuracy.

The pressure distribution is very important to examine the airfoil characteristics and how lift is 
generated by the pressure imbalance. 

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Airfoil Characteristics report aeronautical lab 1

Abstract

In this experiment we will understand how an airfoil can generate lift and lift theories and will obtain a NACA 2412 characteristics using a model airplane with a slab at different speed and angles of attack experimentally using the wind tunnel. And then we are going to compare these results with the same conditions but from different sources using XFLR5 and standard data from NACA.

Introduction

“An airfoil-shaped body moved through a fluid produces an aerodynamic force. The component of this force perpendicular to the direction of motion is called lift. The component parallel to the direction of motion is called drag.” ^[Wikipedia]

Airfoils are found in helicopter rotor blades, propellers, fans, compressors and turbines. Sails are also airfoils, and the underwater surfaces of sailboats. Also it is found in swimming and flying creatures and even plants and sessile organism and the bodies of fishes. Also used in automobiles and spoilers to create a down force improving traction.

Airfoils tents in the wind tunnel are simplified assumptions near to the theoretical results; they ignore the effect of viscosity, nonlinearities in the equation of motion, three-dimensional effects, unsteady flow, free stream and turbulence and wing surface roughness.

Various terms are related to airfoils are defined below

• ·         The mean camber line
• ·         The chord line
• ·         The Chord
• ·         Leading edge
• ·         Trailing edge
• ·         The aerodynamic center
• ·         The center of pressure
• 
  

             :main tybe of airfoil is

• ·   Symmetric / cambered
• Subsonic/ supersonic/ supercritical

Discussion

As angle of attack increase the lift coefficient increase till stall.

As the speed increase (Reynold number increase) the drag increase and CL max decrease.

As the speed increase the zero lift angle decrease (the slope shifts to the left)

How can a negative camber or a negative angle of attack produce lift?

The angle of attack is a combination of an induced angle and an effective angle of attack.

The wing is attached to the fuselage at a certain angle more than zero. So it has by default an angle of attack.

A negative angle of attack produces lift because the effective angle of attack is indeed positive

Conclusion

The standard data are the most accurate and then the data from the XFLR5 with slight variance in the data from the software.

Both the software and the experimental are close and acceptable in comparing between them.

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