Investigating The Lift Produced By Aerofoils Biology Essay

Airfoils impacted the twentieth Century on a monolithic graduated table. Before the twentieth century no human had piloted a powered aircraft and by the terminal of it, it was possible to wing half manner around the universe. Flight has revolutionised the universe, it has enabled economic systems to spread out, trading to diversify and changed the manner wars are fought. Human migration has changed drastically and civilizations mixed all through the development of the airfoil. An airfoil is a device that provides a reactive force when in gesture relation to the environing air. Its construct gives an airplane the indispensable force needed in flight, lift ; generated by the belongingss of the airflow above and below the airfoil.

An airfoil plant by making an environment where the force per unit area above the airfoil is lower than the force per unit area below, ensuing in a net force playing upwards, lift. There are two ways in which this force per unit area difference can be generated, either through planing a cambered airfoil and/or through positioning the airfoil at an angle to the air flow.

The purpose of this experiment is to look into the effects of altering the angle at which the airfoils are positioned in the air flow to happen the angle of onslaught which provides the optimal lift for airfoils.

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A History of Flight

The first recorded flight occurred in 1783 with the Montgolfier brothers ‘ balloon on the 4th June in Annonay, France. Their innovation of the first practical balloon arose from their tests with paper and cloth bags held over a fire, the hot air make fulling the bag and doing it to lift. The brothers so demonstrated flight in Annonay market topographic point with a larger paper lined silk balloon ; the balloon reached a tallness of 6,562 pess.

The first riders to see flight did so in another Montgolfier hot air balloon on September 19th 1783 ; in forepart of Louis XVI, Mary Antoinette and the Gallic Court, a sheep, a cock and a duck flew for a sum of eight proceedingss. These riders were shortly followed by Pilatre de Rozier and Marquis d’Arlandes in an unteathered balloon in free flight. In early 1784 seven riders were carried to a tallness of 3000 pess in a Montgolfier balloon, thought to be powered by ‘Montgolfier gas ‘ ; simply hot air but thought to be a freshly discovered gas igniter than air.

In the late 1700s Sir George Cayley was the first to place the four forces of flight, weight, lift, retarding force and push and the relationship between them. He was besides the first to construct a successful human transporting sailplane. All old attempts had focussed on miming bird flight with flapping wings, these machines, called orthopters, generated both lift and forward gesture at the same clip. Cayley was the first to bring forth programs that resemble today ‘s aircraft with a fuselage, fixed wings, cockpit and rudder controls, he besides built assorted theoretical accounts to show his thoughts.

After analyzing the manner birds glide he noted that they soar for long periods of clip through writhing their arched wing surfaces. His designs introduced cambered wings and his tests became the first scientific testing of airfoils, the beginning of lift generated in modern aircraft. He even noticed the part of low force per unit area created above the sailplane wings which is what causes the lift generated. By 1849 his designs had progressed such that the ten-year-old of one of his retainers became the first individual to wing a sailplane successfully, pull offing a short flight in a Cayley sailplane.

Fifty old ages before the first powered flight Cayley designed and constructed a tri-plane sailplane ( shown in Figure 1.1 ) with three horizontal flying subdivisions, this was successfully piloted by his coachman over 275 meters before crashing.

Figure 1.1 – The design for Cayley ‘s triplane sailplane which successfully carried his coachman 275 meters.

In 1901 after many tests, clangs and designs Orville and Wilbur Wright built a air current tunnel in the dorsum of their bike store to prove wings that had antecedently produced 30 % less lift than they had calculated. Using balance and spring graduated tables to mensurate lift and force per unit area on assorted different airfoils. They discovered that much of the informations about airfoils at the clip was wrong and did non take into history of import factors impacting flight. They investigated the manner the Centre of force per unit area alterations with fluctuations in the angle of onslaught and they found out what the control devices would hold to make to give control to the pilot. After proving over two hundred different flying subdivisions the brothers designed their Glider III.

The brothers took the sailplane to their regular testing land at Kitty Hawk in North Carolina, America, and over one 1000 glides found that the sailplane worked good and most significantly how they had predicted it to. Now the brothers thought that they were on the brink of a interruption through in powered flight. In 1903 the Wrights filed for a patent on the design and turned their focal point to turning the sailplane into a flying machine.

Figure 1.2 – The Wright Brother ‘s air current tunnel

They set about seeking to add a power beginning to Glider III and ended up re-designing the whole construction. The motor was to be fixed to the bottom wing of the plane next to the pilot ; the propellors, which had besides been developed and perfected with the usage of the air current tunnel, would force the aircraft instead than drawing it and would revolve in antagonistic waies. For the trade to be a circular it could non utilize the dunes at Kitty Hawk to take off, as the sailplanes did, they needed a mechanism to take off. This was comprised of a auto on a length of path on which the Flyer sat, when the circular reached the right velocity it would so take off and go on into the air. The auto was powered by the aircrafts propellors as it was to be a flight wholly under the power of the aircraft itself.

The motor itself they had hoped would be designed and built in the automotive industry but a motor with the right specifications ne’er arose and so finally they had to build their ain with the aid of their mechanic, Charles Taylor. The engine was n’t rather what they had wanted but they decided to continue with the proving anyhow to avoid farther delaies.

After many yearss of tweaking the machine was ready to be flown and on the 14th December the brothers tossed a coin to see who would acquire the first flight. Wilbur won and so took his topographic point on the lower wing. On the first effort the lift was set excessively low and so at the terminal of the auto path the circular crashed nose first into the sand. After three yearss of madly seeking to repair the aircraft and the conditions endangering the flights Orville took his bend in the Flyer. At 10:35 AM on December 17th in forepart of several informants the circular took off into a 21 stat mi per hr air current with Wilbur running aboard the wingtip to do certain it did non drag along the land.

Having felt lucky about the twenty-four hours the brothers had placed a camera at the terminal of the path and instructed a informant, John T. Daniels to shut the shutter when the aircraft left the path. From the camera came one of the most celebrated exposure in the history of air power. ( Figure 1.3 )

Figure 1.3 – One of the most celebrated exposure in the history of air power, Orville Wright aboard the Flyer and Wilbur Wright holding run alongside the Flyer during take-off.

The Flyer flew for a sum of 12 seconds and landed in the sand 120 pess off, the first powered flight in the history of air power.

The Wright brothers so went back to their place town of Drayton to better designs and by the fall of 1905 had flown Flyer III for over half an hr in figure of eight forms set downing back where they had started from.

The development of airfoils was non greatly investigated until 1915 when the National Advisory Committee on Aeronautics ( NACA ) stated the demand for “ the development of more efficient flying subdivisions of practical signifier, incarnating suited dimensions for an economical construction, with moderate travel of the Centre of force per unit area and still affording a big angle of onslaught combined with efficient action. ”[ 1 ]

The development of flying designs ( Figure 1.4 ) up until 1917 had been merely test and mistake and it was n’t until NACA published its first work on “ Aerofoils and Aerofoil Structural Combinations ” when mathematical theory started to be applied to the airfoil design. One of the jobs in the design of wings up until the 1930s was that there was non a air current tunnel big plenty to house a full wing, this meant that informations could merely be recorded on little subdivisions of the wing and so extrapolated out to pattern a whole wing. This meant that data about the flow of air around the wingtips was non able to be studied until a air current tunnel was available.

Figure 1.4 – Diagram demoing the development of airfoils in the first half of the twentieth Century.

Uncertainty and Calibration

Restrictions in the truth of the measurings occur due to the quality of the instruments being used. Instruments can merely mensurate to a certain declaration, below which the alterations in the value being measured are excessively little to be detected. As the sensitiveness of the instruments is increased the preciseness of measurings additions leting better, more accurate readings to be recorded.

Calibration of mensurating instruments is besides required earlier usage in order to avoid systematic mistake or prejudice in the readings measured. The impairment of instruments from their original scenes over clip can bring forth readings that differ from the true values by a changeless prejudice. To look into for bias uncertainness, a known mass was placed on the balance and checked for consistence, this was besides compared to reading taken from other balances. The readings finally settled and read to within A±0.01g of each other which was sufficient as I would merely be reading to one denary topographic point.

In order to avoid zero mistake before each reading was taken the fan was switched away and the balance tared, this ensured that readings were taken from the same nothing point each clip and improved the truth of the consequences.

Whilst taking the consequences there was rather a batch of uninterrupted fluctuations whilst the balance was settling down, this response clip could hold caused uncertainness in the consequences if non accounted for. In order to make this the balance was allowed to settle to within A±0.1g before trying to take a reading, this besides increased the stableness of the consequences taken as this was the greatest beginning of mistake due to the non-uniform air flow from the fan.

How Aerofoils Create Lift

In order for something to wing there must be forces moving on the organic structure to antagonize gravitation, this force is called lift. In aircraft lift is produced by fast traveling air go throughing over an airfoil, the air traveling over the top of the airfoil moves faster than that going underneath. This velocity difference creates a difference in force per unit area which creates the lift needed to wing.

To understand how an airfoil creates lift it is necessary to see a few of import equations. Equation 3.1 and Equation 3.2 shows Bernoulli ‘s equation.

Equation 3.1 – Bernoulli ‘s Equation for syrupy flow, where: P = Pressure, I? = Density of fluid, v = speed, g = acceleration due to gravitation, H = tallness.

Equation 3.2 – Height remains more or less changeless with an airfoil ; therefore Bernoulli ‘s equation can be re-written, independent of tallness.

Another of import equation that is necessary to understand how an airfoil plant is the Continuity Equation, shown in Equation 3.3

Equation 3.3 – The Continuity Equation, where: I? = denseness, v = speed, A = cross sectional country of flow.

The design of an airfoil is such that the top surface is curved ( as shown in Figure 3.5 ) this means that the distance over the top of the airfoil is larger than the country below. When air is going towards the airfoil it is going at a changeless speed, as it hits the taking border of the airfoil some of the air passes over the top and some underneath, this point is called the stagnancy point. The air which passes over the top of the wing must therefore travel faster than the air underneath as it has to cover a larger distance in the same clip. This occurs as the air reaches the draging border of the wing at the same clip regardless of the path it takes past the wing. From Bernoulli ‘s rule it is hence possible to see that the force per unit area must be lower on the top of the wing if the air is going faster. The differences in force per unit area above and below the wing produce the net upwards force which we call lift. ( Figure 3.6 )

Figure 3.5 – Diagram demoing the dimensions of an airfoil.

Figure 3.6 – Diagram demoing the parts of high and low force per unit area responsible

For a wing subdivision when the angle of onslaught varies, the air flow patterns over the wing subdivision alteration. The lift generated from thin airfoils starts off with a additive relationship for little angles of onslaught, and so as the angle increases the relationship becomes more complex. This is because as an object moves through the air, air molecules stick to the surface of the object. This creates a bed of air near the surface of the object called a boundary bed that changes the form of the object. The air flow over the object is disrupted as the air responds to this boundary bed as it would a physical surface of the object. The boundary bed can besides sometimes separate from the physical object and cause larger breaks to the air flow. This separation consequence of the boundary bed is the cause of a flying stall and it occurs when the angle of onslaught is excessively big ( shown in Mistake: Reference beginning non found ) .

Figure 3.7 – Diagram demoing the consequence of a flying stall when the angle of onslaught becomes excessively big, the boundary bed offprints and whirls are created doing turbulency.

Experiment A

Overview

I designed this experiment to present the air current tunnel in order to try to cut down the effects of turbulency environing the airfoil and make a more controlled environment to prove in. Turbulent air from the environing environment could make whirls around the wing subdivision interrupting the air flow over the wing and knock abouting the wing subdivision. It would do perturbations in the wing ‘s public presentation, cut downing the effects of lift and so causing inaccuracies in the information.

Diagram

Figure 4.8 – Diagram of equipment apparatus for experiment Angstrom

Method

I set up the equipment as shown in Mistake: Reference beginning non found and tared the balance to zero ; I so made certain that the balance reading remained at zero to look into that the support and balance pan were non in contact with any other equipment as this would hold caused big prejudice ( systematic mistake ) and big nothing mistake. After look intoing for mistakes I carefully set the angle of the airfoil to 0A° and made sure that the index acerate leaf was demoing 0A° on the protractor.

The following measure was to look into that the readings to be obtained were within the balance ‘s scope for negative readings as it was lift that I was traveling to be mensurating. I did this by puting the fan to maximum velocity and so changing the angle of onslaught in unsmooth intervals to guarantee that readings across the whole scope of angles from 0-90A° were suited.

Now that I had checked all the equipment and conditions for the experiment I was able to get down taking consequences. I reset the airfoil place and waited for the balance to settle at 0.00g, I so set the fan to the maximal air current velocity and recorded readings at 5A° increases, each clip waiting for the balance to settle. I repeated this procedure 3 times to hold adequate readings for an norm and so checked to see that the consequences were consistent and that there was no demand to reiterate the experiment a farther clip to obtain better consequences to extinguish outliers.

The same process was so repeated with 3 farther flying subdivisions with differing cambers to prove the fluctuation in lift provided by a larger difference in upper and lower air flows.

To obtain the lift force in Newtons, the alteration detected by the top pan balance was so multiplied by the acceleration due to gravitation experienced at the Earth ‘s surface ( g )

Angle of onslaught, I±

Figure 4.9 – Diagram demoing the angle of onslaught of the wingwhich will be vaired during the experiments.

Consequences and Observations

Whilst obtaining consequences it was hard to read the balance to more than one denary topographic point as there was so much fluctuation in the last figure, this was to be expected as the air flow over the wing subdivision would be far from consistent with a simple apparatus as detailed above. Further experiments would hold provided efforts to cut down the fluctuations in the air flow around the wing and supply a steadier watercourse of air flow. This is detailed in ulterior subdivisions on how I would better the cogency of the consequences in farther experiments.

Flying 1

Angle / A° ( A±1.0 )

Average Lift / manganese ( A±0.6 )

0

1.6

5

6.2

10

10.1

15

15.0

20

19.9

25

24.5

30

29.7

35

34.3

40

36.9

45

39.6

50

39.6

55

38.9

60

36.9

65

33.7

70

28.4

75

23.2

80

17.0

85

7.8

90

0.0

Table 4.1 – Table demoing the mean lift in milli-Newtons of a wing subdivision with camber Flying 1 xmm with changing angles of onslaught ( Plotted in Figure 4.10. See full tabular array in Appendix, Table 7.2 )

Figure 4.10 – Plot of informations in Table 4.1 demoing the relationship between lift in manganese and angle of onslaught for an airfoil of camber Wing 1 xmm The uncertainness bars represent A± twice the standard divergence.

Flying 2

Angle / A° ( A±1.0 )

Average Lift / manganese ( A±0.6 )

0

2.6

5

5.9

10

11.8

15

16.7

20

21.6

25

26.8

30

32.0

35

36.3

40

38.6

45

38.9

50

37.3

55

37.3

60

33.7

65

29.1

70

24.2

75

17.0

80

9.5

85

2.3

90

0.3

Table 4.2 – Table demoing the mean lift in milli-Newtons of a wing subdivision with camber Flying 2 xmm with changing angles of onslaught ( Plotted in Figure 4.11. See full tabular array in Appendix Table 7.3 )

Figure 4.11 – Plot of informations in Table 4.2 demoing the relationship between lift in manganese and angle of onslaught for an airfoil of camber Wing 2 xmm The uncertainness bars represent A± twice the standard divergence.

Flying 3

Angle / A° ( A±1.0 )

Average Lift / manganese ( A±0.6 )

0

3.3

5

8.2

10

12.1

15

17.0

20

20.6

25

25.8

30

31.1

35

35.3

40

38.9

45

40.5

50

41.5

55

38.2

60

35.3

65

31.4

70

27.5

75

21.9

80

16.7

85

8.8

90

0.7

Table 4.3 – Table demoing the mean lift in milli-Newtons of a wing subdivision with camber Flying 3 xmm with changing angles of onslaught ( Plotted in Figure 4.12. See full tabular array in Appendix Table 7.4 )

Figure 4.12 – Plot of informations in Table 4.3 demoing the relationship between lift in manganese and angle of onslaught for an airfoil of camber Wing 3 xmm The uncertainness bars represent A± twice the standard divergence.

Flying 4

Angle / A° ( A±1.0 )

Average Lift / manganese ( A±0.6 )

0

5.9

5

9.5

10

13.1

15

18.3

20

22.6

25

26.8

30

31.4

35

35.6

40

38.6

45

39.6

50

37.3

55

36.3

60

34.0

65

31.4

70

27.1

75

21.6

80

14.7

85

8.5

90

0.7

Table 4.4 – Table demoing the mean lift in milli-Newtons of a wing subdivision with camber Flying 4 xmm with changing angles of onslaught ( Plotted in Figure 4.13. See full tabular array in Appendix Table 7.5 )

Figure 4.13 – Plot of informations in Table 4.4 demoing the relationship between lift in manganese and angle of onslaught for an airfoil of camber Wing 4 xmm The uncertainness bars represent A± twice the standard divergence.

Figure 4.14 – Plot of informations in Table 4.5 demoing the relationship between lift in manganese and angle of onslaught for airfoils of different cambers. The uncertainness bars represent A± twice the standard divergence.

Decision

From the graphs it is possible to see that the maximal lift was obtained at an angle of onslaught of between 45A° and 50A° . This seems to be consistent across all flying subdivisions of differing cambers. There is no obvious stall angle, which occurs where the lift reaches a upper limit and so beads quickly off, explained further in the subdivision How Aerofoils Create Lift. This is surprising as most airfoils have a stall angle between 15 and 25 but this greatly depends on the velocity of the air flow over the wing and the airfoil design. It would be hard to explicate accurately as to why the consequences from the experiment differ greatly from the mean wing but through mensurating the air current velocity it may be possible to derive a clearer image.

The Reynold ‘s figure ( given in Equation 4.4 ) is a numerical value which describes the ratio of the inertial forces ( forces to make with the mass of the object ) to the syrupy forces ( the resistive forces provided by the medium through which it is going ) and the importance of these in given flow conditions. It is used to compare the conditions found in wind tunnel trials on little graduated table theoretical accounts to the conditions found on full graduated table wings. It is good known that some features such as the retarding force and maximal lift coefficients vary with the size of the wing for a given air flow. As a consequence the wing subdivisions used in wind tunnel trials do non see the same effects of flow every bit would a full graduated table wing, if the wing was to be scaled down by a factor of 4, the flow velocity would hold to be increased by a factor of 4 to fit the Reynold ‘s figure in both conditions.

Equation 4.4 – The equation for Reynold ‘s figure where I? = fluid denseness, V = mean fluid speed, L = characteristic additive dimension and I? = fluid viscousness. The characteristic additive dimension for an airfoil is its chord length.

Objects in flow conditions with a low Reynold ‘s figure experience laminar flow whilst those with high Reynold ‘s figure create disruptive flow with random Eddies and whirls.

Laminar flow is a “ smooth ” uninterrupted flow of air over the surface of the wing. With any type of flow, near to the flying a boundary bed is stationary relation to the wing ; as the distance from the flying surface additions, the comparative speed additions. If the boundary bed flows in parallel beds and there are no perturbations doing energy transportation between the beds, the flow is said to be laminar.

From mensurating the air current velocity and obtaining values for the viscousness and denseness of air it would be possible to cipher the Reynold ‘s figure for the airfoils being tested and see how these compared to the Reynold ” s Numberss of full graduated table wings. Investigating farther into this may demo that the air current velocities that were tested were excessively slow to obtain similar air flow over the wing and may give an account as to why the angle of greatest lift is so much higher than expected.

Experiment B

Overview

In order to seek and detect the effects of the turbulent flow over the wing subdivision I attempted to utilize methods in the air current tunnel to demo lines of flow as they pass over the airfoil. To demo this I attempted to utilize lines of fume injected into the air current tunnel trial chamber and gaining control the flow lines utilizing high definition picture recordings.

Diagram

Figure 5.1 – Diagram demoing the apparatus for the Experiment B, presenting the fume to demo airflow.

Method

I used a similar apparatus to that in Experiment A but in order to try to unbend the air flow and better the consistence, I used little subdivisions of imbibing straw arranged to make a honeycomb construction placed in the funnel. I so made a little hole in the roof of the air current tunnel and used a gas syringe connected to a thin tubing and nose to shoot a all right watercourse of fume into the way of the air flow. This should hold shown the flow of the air above and below the wing subdivision and with a high definition picture camera positioned to enter the flow I would so hold played back the picture and analysed the flow forms. I would hold looked for the turbulency that occurs when the wing is about to procrastinate and the splitting of the boundary bed ( shown in Figure 5.2 ) which causes the force per unit area difference to drop and the lift to diminish dramatically.

Figure 5.2 – Image demoing the lines of flow at the point of procrastinating when the boundary bed splits from the wing subdivision[ 2 ]

Decision

Attempts to see the fume as it passed over the wing subdivision were unsucessful, the jobs arose when shooting the fume into the chamber. The fume used was either non heavy plenty or merely spread far excessively rapidly in the air flow to be picked up by the camera or by oculus. To seek and better the visibleness of the fume I increased the strength of the visible radiation used to light the fume by presenting a stroboscope visible radiation at the terminal of the air current tunnel. This made the fume a small seeable to the bare oculus but still it was non picked up by the camera. If I were to make the experiment once more I would seek and utilize dry ice to make the streeamlines and demo the flow over the wing. This I think would hold a greater contrast to the background and demo the lines of flow and the turbulency much more clearly. This would besides state me whether the conditions in the air current tunnel were similar to those found in true flight as it would be possible to see how the flow alterations for different angles of onslaught. When the airfoil approaches the stall angle of onslaught the lines of flow separate as they pass over the wing, I could hold compared this information with the informations recorded in Experiment A and so tested to see whether it was really the wing procrastinating which created the loss of lift or if it was another factor. The stall angle of onslaught for existent wings is normally around 15-25A° nevertheless the consequences from Experiment A suggested that this was non the instance in this apparatus.

It would hold been interesting to happen out whether there was merely excessively much turbulency and perturbation in the air flow for there to be a noticeable point at which the wing stables. If the wing was in changeless turbulent flow at that place may non hold been a point at which it specifically stalled.

Further Experiments

From analyzing the consequences in Experiment A I would hold liked to execute this following experiment to see what consequence the external conditions outside of the air current tunnel would hold on the lift. The disruptive air may do random mistake to the consequences which may demo fluctuations in the lift created. The air flow over the wing subdivision would be disrupted and so little whirls and currents would be created, impacting the lift generated by the wing subdivision. This sort of turbulency could happen in conditions systems were the air currents are lifting and falling with different force per unit area systems created on Earth. Failure of the flying design in a state of affairs like this could hold ruinous effects with loss of control of the aircraft and potentially unsafe state of affairss.

Diagram

Protractor

Figure 6.1 – Diagram of equipment apparatus for experiment B

Method

I would put up the equipment as shown in Figure 5.1 taking the same safeguards to understate uncertainness in the consequences, as in Experiment A. I would so continue to take three sets of consequences increasing the angle of onslaught by 5A° each clip and mensurating the lift to one denary topographic point after the balance had clip to settle. Repeating 3 times would better the dependability of the consequences and cut down the hazard of outliers, the mean of these consequences would so be taken and plotted to give a graph of the consequences to demo how lift varied with the angle of onslaught.

Decisions

From the consequences I would anticipate to see a extremum at a similar angle but for there to be more fluctuation in the three repeated sets of consequences. This is because I would anticipate there to be more fluctuation in the readings taken from the balance due to the excess turbulency created from external conditions.

Appendix

Risk Assessment Table

Hazard

Hazard

Degree of Hazard

Control Measures

Mains electricity supply

Electrocution

Low

Avoid contact with bare wires if present, do non utilize defective equipment, do non utilize equipment when custodies are wet

Open backed fan

Web of vesture or equipment in fan blades

Low

Turn off when non in usage, do certain all long hair tied back and maintain electrical leads etc. clear from rear of fan.

Obstructions

Stumbling and falling

Low

Keep work country free of obstructors ( i.e. chairs, bags, etc )

Falling Objects

Damage to organic structure parts

Low

Ensure that all heavy equipment ( power supplies, fan ) are safely on tabular array and steady before get downing experiments.

Table 7.1- Risk appraisal tabular array for work in the research lab during experimental work.

Full Results Tables

Angle / A° ( A±1.0 )

Lift / manganese ( A±0.1 )

Standard Deviation / manganese

Error / manganese

1

2

3

Average ( A±0.6 )

0

2.0

2.0

1.0

1.6

0.57

1.13

5

6.9

5.9

5.9

6.2

0.57

1.13

10

10.8

9.8

9.8

10.1

0.57

1.13

15

15.7

14.7

14.7

15.0

0.57

1.13

20

20.6

19.6

19.6

19.9

0.57

1.13

25

25.5

24.5

23.5

24.5

0.98

1.96

30

30.4

29.4

29.4

29.7

0.57

1.13

35

35.3

34.3

33.3

34.3

0.98

1.96

40

39.2

36.3

35.3

36.9

2.04

4.08

45

41.2

39.2

38.2

39.6

1.50

3.00

50

40.2

39.2

39.2

39.6

0.57

1.13

55

39.2

39.2

38.2

38.9

0.57

1.13

60

37.3

37.3

36.3

36.9

0.57

1.13

65

34.3

33.3

33.3

33.7

0.57

1.13

70

28.4

28.4

28.4

28.4

0.00

0.00

75

23.5

22.6

23.5

23.2

0.57

1.13

80

16.7

16.7

17.7

17.0

0.57

1.13

85

6.9

7.8

8.8

7.8

0.98

1.96

90

0.0

0.0

0.0

0.0

0.00

0.00

Table 7.2 – Full moon consequences table for the mean lift in milli-Newtons of a wing subdivision with camber WING 1 xmm with changing angles of onslaught

Angle / A° ( A±1.0 )

Lift / manganese ( A±0.1 )

Standard Deviation / manganese

Error / manganese

1

2

3

Average ( A±0.6 )

0

2.0

2.9

2.9

2.6

0.57

1.13

5

5.9

3.9

7.8

5.9

1.96

3.92

10

11.8

10.8

12.7

11.8

0.98

1.96

15

15.7

16.7

17.7

16.7

0.98

1.96

20

20.6

21.6

22.6

21.6

0.98

1.96

25

27.5

25.5

27.5

26.8

1.13

2.26

30

32.4

31.4

32.4

32.0

0.57

1.13

35

36.3

35.3

37.3

36.3

0.98

1.96

40

39.2

38.2

38.2

38.6

0.57

1.13

45

39.2

39.2

38.2

38.9

0.57

1.13

50

38.2

37.3

36.3

37.3

0.98

1.96

55

38.2

37.3

36.3

37.3

0.98

1.96

60

34.3

34.3

32.4

33.7

1.13

2.26

65

29.4

29.4

28.4

29.1

0.57

1.13

70

24.5

24.5

23.5

24.2

0.57

1.13

75

17.7

16.7

16.7

17.0

0.57

1.13

80

9.8

9.8

8.8

9.5

0.57

1.13

85

2.0

3.9

1.0

2.3

1.50

3.00

90

0.0

1.0

0.0

0.3

0.57

1.13

Table 7.3 – Full moon consequences table for the mean lift in milli-Newtons of a wing subdivision with camber WING 2 xmm with changing angles of onslaught

Angle / A° ( A±1.0 )

Lift / manganese ( A±0.1 )

Standard Deviation / manganese

Error / manganese

1

2

3

Average ( A±0.6 )

0

3.9

2.9

2.9

3.3

0.57

1.13

5

8.8

7.8

7.8

8.2

0.57

1.13

10

12.7

11.8

11.8

12.1

0.57

1.13

15

17.7

16.7

16.7

17.0

0.57

1.13

20

20.6

20.6

20.6

20.6

0.00

0.00

25

26.5

25.5

25.5

25.8

0.57

1.13

30

32.4

30.4

30.4

31.1

1.13

2.26

35

36.3

35.3

34.3

35.3

0.98

1.96

40

39.2

39.2

38.2

38.9

0.57

1.13

45

41.2

40.2

40.2

40.5

0.57

1.13

50

42.2

41.2

41.2

41.5

0.57

1.13

55

39.2

38.2

37.3

38.2

0.98

1.96

60

36.3

35.3

34.3

35.3

0.98

1.96

65

31.4

31.4

31.4

31.4

0.00

0.00

70

27.5

27.5

27.5

27.5

0.00

0.00

75

21.6

22.6

21.6

21.9

0.57

1.13

80

17.7

16.7

15.7

16.7

0.98

1.96

85

9.8

8.8

7.8

8.8

0.98

1.96

90

1.0

1.0

0.0

0.7

0.57

1.13

Table 7.4 – Full moon consequences table for the mean lift in milli-Newtons of a wing subdivision with camber WING 3 xmm with changing angles of onslaught

Angle / A° ( A±1.0 )

Lift / manganese ( A±0.1 )

Standard Deviation / manganese

Error / manganese

1

2

3

Average ( A±0.6 )

0

5.9

5.9

5.9

5.9

0.00

0.00

5

9.8

9.8

8.8

9.5

0.57

1.13

10

13.7

13.7

11.8

13.1

1.13

2.26

15

18.6

18.6

17.7

18.3

0.57

1.13

20

23.5

22.6

21.6

22.6

0.98

1.96

25

27.5

27.5

25.5

26.8

1.13

2.26

30

31.4

31.4

31.4

31.4

0.00

0.00

35

35.3

36.3

35.3

35.6

0.57

1.13

40

39.2

38.2

38.2

38.6

0.57

1.13

45

40.2

39.2

39.2

39.6

0.57

1.13

50

38.2

37.3

36.3

37.3

0.98

1.96

55

37.3

36.3

35.3

36.3

0.98

1.96

60

34.3

34.3

33.3

34.0

0.57

1.13

65

32.4

31.4

30.4

31.4

0.98

1.96

70

27.5

27.5

26.5

27.1

0.57

1.13

75

22.6

20.6

21.6

21.6

0.98

1.96

80

15.7

12.7

15.7

14.7

1.70

3.40

85

9.8

6.9

8.8

8.5

1.50

3.00

90

2.0

0.0

0.0

0.7

1.13

2.26

Table 7.5 – Full moon consequences table for the mean lift in milli-Newtons of a wing subdivision with camber WING 4 xmm with changing angles of onslaught

Angle / A° ( A±1.0 )

Lift / manganese ( A±0.1 )

Standard Deviation / manganese

Error / manganese

1

2

3

Average ( A±0.6 )

0

-12.7

-12.7

-11.8

-12.4

0.57

1.13

5

9.8

8.8

10.8

9.8

0.98

1.96

10

26.5

27.5

29.4

27.8

1.50

3.00

15

49.0

47.1

53.0

49.7

3.00

5.99

20

65.7

66.7

67.7

66.7

0.98

1.96

25

82.4

83.4

82.4

82.7

0.57

1.13

30

89.2

90.2

92.2

90.5

1.50

3.00

35

77.5

76.5

77.5

77.1

0.57

1.13

40

87.3

86.3

89.2

87.6

1.50

3.00

45

96.1

95.1

96.1

95.8

0.57

1.13

50

102.0

101.0

101.0

101.3

0.57

1.13

55

105.9

104.9

103.0

104.6

1.50

3.00

60

104.0

102.0

98.1

101.3

3.00

5.99

65

98.1

93.2

90.2

93.8

3.96

7.93

70

85.3

82.4

81.4

83.0

2.04

4.08

75

66.7

65.7

62.8

65.1

2.04

4.08

80

47.1

49.0

51.0

49.0

1.96

3.92

85

26.5

25.5

29.4

27.1

2.04

4.08

90

2.0

2.9

2.0

2.3

0.57

1.13

Table 7.6 – Full moon consequences table for the mean lift in milli-Newtons of a wing subdivision with camber WING 2 xmm with changing angles of onslaught with the air flow restricted by the funnel.

Plants Cited

‘Aviation Timeline ‘ . ( n.d. ) . Retrieved April 23, 2010, from Century of Flight: hypertext transfer protocol: //www.century-of-flight.net