SIZING OF AIRCRAFT ACCORDING TO TAKEOFF DISTANCE REQUIREMENT

SIZING OF AIRCRAFT ACCORDING TO    TAKEOFF           DISTANCE REQUIREMENT

The take off ground run (S_TOG) of an aircraft is proportional to takeoff  wing loading (W/S)_TO, or  take off power loading (W/P)_TO . and maximum take off  lift C_Lmax  .

i.e,             S_TO a  ( (W/S)_TO* (W/P)_TO ) / (sC_Lmax) = TOP    ——-(1)

     this equals to TOP (take off parameter ) and its dimension  is  newton²/ m²hp .

As we know that the coefficient of lift at take off is typically 1.21 times maximum lift coefficient.

So we can write it as a C_TO = C_LmaxTO  / 1.21

Below Fig relates S_TOG to the  take off parameter . for a range of aircraft of single and twin engine (source by aircraft flight dynamics and automotive flight control , Roskam.).

This graph suggest following relationship

                     S_TOG  = 4.9 TOP +0.009TOP²               _{———————}(2)

And , fig  Below  implies that ,

S_TO = 1.66 S_TOG                               ————(3)

   NOW from above two equation we get,

                          STO = 8.134 TOP + 0.0149 TOP²         ———–(4)

THIS EQUATION are generated on the basis of data from FAR 23 aircraft.

For the calculation for jet aircraft we might replace (W/P) TO (W/T).

Here is one example so that the above concept will clear to everyone.

For a given aircraft (take propeller driven) it is specified that S_TOG LESS THAN 1000m and S_TO IS LESS THAN 1500m at an altitude of 5000 ft  in standard atmosphere.

Equation (3) stipulates that ,

S_TO < 1660m

This clearly violate the second requirement. So for both takeoff requirement to be met it is necessary that ,

1500 = 8.134TOP + 0.0149 TOP²

FROM this we get TOP = 145.6

Since s = 0.8616 at 5000 ft , now put this values in eq—-(1) it translate into ( (W/S)_TO* (W/P)_TO ) / (C_Lmax) < 145.6 *0.8616

Now by putting the values of  (C_Lmax)   and  (W/P)_TO we get approximate value of (W/S)_TO  

SIZING OF AIRCRAFT ACCORDING TO MANEUVARING REQUIREMENT

SIZING OF AIRCRAFT ACCORDING TO MANEUVARING REQUIREMENT

In this post I am going to introduce about the sizing of aircraft according to their maneuvering requirement .

Specific requirement for sustained maneuvering capability (including sometimes specific turn rate) are often contained in the mission specification for agricultural ,aerobatic or for military aircrafts.

Sustained maneuvering requirement are usually formulated in terms of a combination of sustained load factor (n) at some speed and altitude.

The sustained maneuvering capability of an aircraft depends strongly on its maximum lift coefficient and on its trust.

For equilibrium in vertical direction of flight , it Is necessary that:-

n*W =C_L*q*S       ———–(1)

where n= load factor and, q= dynamic pressure

the maximum load factor can be capability of an aircraft (n_max)can be found from above equation(1) ;

n_max = C_L*q /(W/S)

this load factor can be sustained as long as there is sufficient thrust, since:-

THRUST = DRAG

T= C_DO + KC_L²,

T= (q*C_D0*S )+( C_L²*q*S/∏e*A.R)

After dividing above equation by W  we get :-

T/W = (q*C_DO/(W/S)) + (W/S)(n_max)²/(∏*e*A.R )

NOW,

If some maximum load factor is desired on a sustained basis at a given speed and altitude then , the above equation can be used to find out the relation between (T/W) and (W/S) for a given value of c_d0.

so we can find out the wing loading for a specific thrust by weight ratio and hence design the wing /aircraft accordingly.

If requirement is includes some specific turn rate then the following equation might be helpful:-

                            Turn rate= g(n²-1)^1/2/V

EFFECT OF DIFFERENT LOCATION OF AIRCRAFT CENTER OF GRAVITY

EFFECT OF DIFFERENT LOCATION OF AIRCRAFT

                                               CENTER OF GRAVITY

In this post I am going to share my knowledge about the effect of  different locations of center of gravity(away from the margin specified by designer ) in different directions  , so that you can identify the probable cause in your model if your craft show below given character .

 

1. 1.      IMPACT OF LONGITUDINAL  POSITION OF C.G

a)      When c.g is too forward

• When c.g is forword the elevator authority will decrease by large magin. This effect we can see at the time of landing and take off .
• ON LANDING  aircraft is unable to pitch up at low speed , the aircraft is too nose heavy.
• ON TAKE OFF  aircraft unable to produce enough moment to rotate the nose resulting aircraft need longer runway for taking off.

 

• EFFECT ON LONGITUDINAL STABILITY :- The forword c.g location will increase aircraft longitudinal stability (why? Explanation below ).
• The forward c.g has a greater distance to the neutral point, means greater the static margin, resulting better aircraft attitude after a disturbance due to gust .

 

• PERFORMANCE VIEW. :- aircraft with forward c.g has very poor performance at any given airspeed (why ? reason below).
• Due to too forward c.g the downward force on the tail to resist the nose drop tendency will increased ,it increased the trim angle of attack of the aircraft, resulting overall drag will increased.
• Also Because o the above reason, the cruise speed for a given thrust and weight will reduced.

b)      When c.g is too aft

• Because of  too aft  c.g location , aircraft always have nose up tendency .and aircraft become more likely to stall(its too dangerous, believe me )
• ON LANDING the effect is exactly opposite to the what  we see in forward c.g location i.e the more nose down elevator is required to counter the nose up tendency of aircraft.
• ON TAKEOFF aircraft is likely to nose up prematurely , and reduced the climb performance  
• Spin recovery becomes more difficult as the CG moves rearward. Reason:-Centrifugal forces acting about the CG, during a “flat” spin, pull the aircraft out of its axis of spinning, making it difficult to nose down and recover.

 

• Possibility of a tip over if the CG is far aft the Neutral Point.

 2.     EFFECT OF LETERAL POSITION OF C.G

• Laterally imbalanced c.g will cause unexpected roll during flight that required an extra effort to make the aircraft balanced .

• It has little or no effect on longitudinal stability but affect the aircraft’s lateral stability.

 

• The roll control surface (ailerons, spoilers ) sizing is extremely depends on the expected  variation of  lateral c.g that arise due to factors like weight asymmetry   , one side engine failure  etc.

 

• The effect of lateral c.g imbalance on aircraft handling is typically assessed by:-

–          Fuel imbalance test

–          One engine failure test (for multiengine aircraft  ).

 

The combination  of both scenarios provides the maximum rolling moment for the aircraft that is crucial in aileron and spoilers sizing.

3.      EFFECT OF VERTICAL POSITION OF C.G

• The vertical location of the CG changes the amount of dihedral effect (for more on dihedral effect follow my 22^nd  JAN  post ). As the  vertical CG  moves lower, dihedral effect increases. This is caused by the center of lift and drag being further above the CG and having a longer moment arm.
•  So, the same forces that change as sideslip changes (primarily sideforce, but also lift and drag) produce a larger moment about the CG of the aircraft. This is sometimes referred to as the pendulum effect.
• On lateral stability – a c.g elevation leads to a deterioration of spiral stability, dutch roll and roll maneuverability.

AIRCRAFT WING TIP SHAPE

Aircraft wing tip shape

As I have introduced in last post about winglet (an wing tip device) . how it improves the aircraft stall characteristics , and also increases the effective span and effective aspect ratio of the wing.

SOME MORE POINT ABOUT WINGLET

The winglet is cambered and twisted so that the rotating vortex flow at the tip creat a lift force on the winglet that has forward component. This forward component act as a negative drag, reducing the total wng drag.

A properly designed winglet can potentially provide an effective span increase up to double that bought by adding the winglets height to the wing span.

Winglets provide the greatest benefit when the wing tip vortex is strong, so a low aspect ratio wing will see more advantages fram the use of winglet than an already  highly efficient high aspect ratio wing.

Now  here , I am going to discuss about the type of wing tip shape that affect the aerodynamics of aircraft and its advantages and disadvantages.

TYPES OF WING TIP DEVICES

• End plates
• Classic Winglet (Whitcomb)
•  Blended Winglet
• Hoerner Tips
• Upswept and Drooped Tips
•  Wing Grid
•  Sail Tips
• Spiroid Tips
• Tip Turbines

1)      END PLATES

 

• End plates are most common and most effective type.
• It work same as like T –tail.
• Above fig shows how it affects the span and aspect ratio of wing.

2)      CLASSIC WINGLET (WHITCOMB

• It Defined by Whitcomb
• The Upper winglet begins at the point of maximum  thickness of wing airfoil.
•  It having a sweep same as In wing.
• The Span of it(upper part) equals to the tip chord of the wing.
• Its camber is generally higher  than wing.
• Lower winglet contributes little to drag hence it often omitted.

3)      BLENDED WINGLET

• Greatly reduces the adverse flow conditions at winglet junction.
• The Blended Winglet incorporates a large radius and a smooth chord variation in the wing-to-winglet transition area. This allows optimum aerodynamic loading and avoids vortex concentrations that produce drag.
• High AR blended winglet can be up to 60% more effective than a conventional winglet.
• Most important parameter in design is the ratio of winglet high to wing span.
• Blended Winglets enhance longitudinal and directional stability, thereby providing better handling in turbulent atmospheric conditions.

4)      HOERNER TIP

• this is a sharp edged wing tip with the upper surface continuing the upper surface of the wing.
• The lower surface is “undercut” and canted approximately 30 deg to the horizontal. It may also be undercambered   (concave).
• It is somewhat similar to conventional tip.

 

5)      UPSWEPT  AND DROOPERD TIP

• Similar to Hoerner Tips but curve either up or down to increase  the wing’s effective span.
• This effect is similar to that employed by endplates.

6)      WING GRIDS

 

•  The circulation is taken over by the wing grid along the chord of thE main wing.

•  The segmented circulation is transferred to the end of the wing grid, increasing the far field vortex spacing.
• The lift distribution on several winglets results in a reduction of the far field vortex energy.
•  Induced drag is reduced by the wing grid up to 60%, that corresponds to span efficiencies of up to over 3.0, that means that total drag can be reduced up to 50% depending on velocity and design.
•  The winggrid has two distinct operating regimes:

1) Below a critical angle of attack (above a specific design speed) span efficiency is between 2.0 and 3.0 with full wing grid effect.

2) Above a critical angle of attack (below a specific design speed) the effect of reduced induced drag fades out, the winggrid perates as a slit wing with very high stall resistance.

7)      SAIL TIP

 

• Sail tip is defined by multiple high Aspect Ratio lifting elements at several dihedral angles.
• This kind are more complex to design and analyze .

 

8)      SPIROID WING TIP

• Spiroid wing tip produces a reduction in induced drag, much like that of a winglet. However, its closed planform shape may make it possible to reduce local lift coefficient often a problem for winglets.
• Extensive optimization is necessary.
• The major concern with it is fluttering of tip.
• It reduce fuel consumption by 6-10% as compared to conventional tip.

9)      TIP TURBINE

• Reduce the strength of vortices.
• Recover energy required to overcome the drag.
• It is estimated that a similar system on  B747 would result in the recovery of 400hp(source).

IMPROVING AIRCRAFT STALL CHARACTERISTICS WITH FIXED DEVICES

IMPROVING AIRCRAFT STALL CHARACTERISTICS WITH  FIXED                                                                     DEVICES 

Stalls and Spins

Stalls and resulting spins have caused aircraft accidents since the beginning of flight. Even though airplanes have evolved to have better stall characteristics, stalls and spins continue to be a leading cause of accidents.

A stall occurs when airflow separates from all or part of the upper surface of a wing, resulting in sudden loss of lift. This is caused by exceeding the critical angle of attack (angle of attack is the angle between the relative wind and the chord line of the airfoil).SEE FIG.

Below the critical angle, airflow over the wing surface is relatively smooth. Above the critical angle, the thin layer of air above the wing, or “boundary layer,” becomes turbulent and separates from the airfoil  Lift is destroyed and drag increases, causing the aircraft to rapidly lose altitude. Pilots are trained to recover from this condition. However, if the stall occurs too low to the ground, there may not be enough altitude to recover.Stalls are usually associated with slow flight in a nose-up attitude, but they can occur at any airspeed or attitude.

FIG- aircraft under spin

Spins are of even greater concern because recovery requires more altitude and more actions on the part of the pilot. Simply stated, a spin is an autorotation resulting from one side of the wing stalling more than the other. Spins cause rapid loss of altitude. If the pilot does not recover, the aircraft will spin into the ground. Aircraft design affects the ease of entering and recovering from spins. Adding devices to improve stall characteristics will generally reduce spin-related accidents; preventing a stall or making it gentler can reduce accidental spins. Straight-wing aircraft must stall before they will spin (swept-wing aircraft do not necessarily have to stall first) .

Favorable Stall Characteristics

• Aircraft are designed to have the most favorable stall characteristics possible given the compromises involved. A good aircraft should give the pilot adequate warning of the stall, stall gradually, and tend not to spin after the stall .

• This means the aircraft wing should stall at the roots first, rather than at the tips where the ailerons are located. (Stalling at the tips first renders the ailerons ineffective for roll control.) Usually a NEGATIVE twist, or “washout,” is built into the wings so that the tips are always at a lower angle of attack. However, this is not always enough to create good stall characteristics. Some aircraft require further modifications to the wing.

•  In the interest of safety, several types of fixed devices can modify a basic wing in order to improve stall characteristics. These include winglets, leading edge cuffs, stall strips, stall fences, slots, and vortex generators.

Winglets

Winglets, which are vertical extensions of the wingtips, improve stall characteristics by reducing induced drag. This induced drag comes from high pressure air under the wings flowing around the tips to the lower pressure area above, creating vortices. Winglets redistribute the intensity of wingtip vortices over a larger area. They increase the maximum coefficient of lift, resulting in a lower stall speed.

NASA studied the effect of winglets on the performance and handling qualities of aircraft. Using windtunnels, NASA tested two versions of a model airplane: one with winglets and one without. At stall, the airplane without winglets tended to “roll off” and “drop a wing”. The airplane with winglets demonstrated improved stall characteristics. According to a NASA report, “The winglets appeared to prevent the wing tip from stalling first… reducing the tendency to roll off” . In another study during the fuel crisis of the 1970s, NASA aerodynamicist Richard Whitcomb found that winglets reduced drag by 20 percent. Less drag means faster stall recovery.

Winglets have disadvantages that may outweigh their advantages in some cases. They create interference drag at the junction between the wing and winglet. This drag offsets the induced drag reduction. Winglets also have a tendency to flutter.

Leading Edge Cuffs

• Leading edge cuffs are extensions that project forward and droop from the outboard sections of an aircraft’s wings.

• They divide a wing into two different airfoils. The outboard section, with the cuffs, has a lower angle of attack and continues flying while the inboard section stalls. This allows the ailerons to continue to be effective in the stall. Wing cuffs also increase the stalling angle of attack .

Cuff Creates a secondary vortex over the wing Prevents separated flow from propagating down the span Attached flow is maintained over the tip and aileron see fig below

The disadvantages of leading edge cuffs are that they complicate the wing design and cannot be easily added to an existing wing.

Stall Strips

• A simple way to make the inboard section of a wing stall first is to install a stall strip on the inboard leading edge.

• Stall strips are lengths of wedge-shaped metal, wood, or other material that run parallel to the leading edge.

•  At high angles of attack, the strips disrupt the boundary layer behind them and cause a stall in that area . They have the added benefit of causing a more pronounced stall buffet, providing more warning to the pilot.

• Stall strips are very common, especially in homebuilt aircraft, because they are easily removed and reattached. Homebuilders can experiment by putting the strips in different locations and noting their effect on stall characteristics.

•  Stall strips are also found on many production aircraft, such as the Mooney series and the Cirrus SR20 and SR22. Some aircraft use stall strips on only one wing in order to eliminate an asymmetric stalling pattern .

While stall strips often improve stall characteristics, they are not always a quick fix for an airplane with poor handling qualities.

Stall Fences

• Stall fences are another device to prevent wing tips from stalling before the roots.

• They are thin plates which project up from the wing and lie parallel to the aircraft’s axis of symmetry. Without these fences, a spanwise airflow along wings causes the boundary layer to thicken toward the wingtips, especially on swept-wing aircraft. This results in early boundary layer separation at the wingtips and loss of aileron control.

• Fences block spanwise airflow, preventing boundary layer buildup over the ailerons and thus improving stall characteristics .

• Stall fences are primarily found on swept-wing aircraft like business jets and fighters, but are also found on straight-winged general aviation aircraft. The Eagle 150 has a unique design featuring a main wing, a forward wing, and a horizontal stabilizer. Each side of the main wing has a stall fence, which according to the manufacturer “redirects the airflow to the ailerons, creating a re-energizing effect. This allows the pilot or student pilot full control at minimum speed, and even at the point of stall” (Eagle Aircraft).

Some aircraft have aerodynamic stall fences. These are not the typical vertical plates, but instead other devices which create the same effect. The Questair Venture kitplane has a small vertical slot on the leading edge of each wing; the airplane looks like it ran into a bandsaw. At high angles of attack, air flowing through this slot creates a trailing vortex that acts like a stall fence .the  “rooster tail” of air creates turbulent, high pressure air that can “impede the spanwise advance of the stall” .

Whether the stall fence is a physical plate or an airflow barrier, this device combats the progression of a stall across a wing and helps keep air flowing over the ailerons.

Slots

• Fixed slots in aircraft wings are used to increase the maximum coefficient of lift and delay the stall.
• Slots are long holes near the leading edge of a wing that run parallel to the leading edge. At high angles of attack, slots route high pressure air near the stagnation region under the wing to the lower pressure region on top . (This rerouted air energizes the boundary layer and delays its separation. Slots allow the aircraft to reach a higher angle of attack before stalling.

• Slots greatly improve the performance of aircraft at high angles of attack, but they have disadvantages as well. They must be designed into the wing from the beginning; they cannot be retrofitted to improve a design.

• ·         They cause a higher stall angle, so the plane must approach in a nose-up attitude that decreases visibility .

•  The main disadvantage is that slots create excessive drag during normal cruising flight . (A solution to that problem is the slat, which is a slot with a moveable cover.

Vortex Generators

The most common boundary layer control devices are vortex generators. These are tiny plates mounted near the leading edge of the wing, perpendicular to the surface . Vortex generators energize the boundary layer by mixing in high-energy air from outside the boundary layer . They delay airflow separation .

Conclusion

The devices discussed in this section can help warn the pilot of a stall, cause the airplane to stall more gently, and make the airplane resist spinning. They are added measures of safety that improve a basic wing. Small changes in airflow patterns over a wing can greatly improve aircraft handling qualities at high angle of attack.

 

WING IN GROUND EFFECT

 

WING IN GROUND EFFECT

The conventional practical use of lifting bodies, are wings on aircraft. In very broad terms, aircraft fly because the movement of the wing through the air produces a greater static pressure on the lower surface of the wing than on the upper surface of the wing. The pressure differential equates to a resultant force upward which supports the weight of the aircraft.

Aircraft normally fly in a freestream, that is the air around the wing is not bounded in any way.

 

WIG(wing in ground effect) craft make use of a phenomenon known as ‘ground effect’. Ground effect is the common name for the phenomenon where a boundary is placed below (and near) the lower surface of the wing. This results in an effective increase in the static pressure below the wing and increases the lift to drag ratio. In practice, the boundary is the earth’s surface, whether it is terrain or water.

• When an aircraft is flying at an altitude that is approximately at or below the same distance as the aircraft’s wingspan or helicopter’s rotor diameter, there is, depending on airfoil and aircraft design, an often noticeable ground effect. This is caused primarily by the ground interrupting the wing tip vortices and downwash behind the wing. When a wing is flown very close to the ground, wingtip vortices are unable to form effectively due to the obstruction of the ground. The result is lower induced drag, which increases the speed and lift of the aircraft.

 

• A wing generates lift, in part, due to the difference in air pressure gradients between the upper and lower wing surfaces. During normal flight, the upper wing surface experiences reduced static air pressure and the lower surface comparatively higher static air pressure. These air pressure differences also accelerate the mass of air downwards. Flying close to a surface increases air pressure on the lower wing surface, known as the “ram” or “cushion” effect, and thereby improves the aircraft lift-to-drag ratio. As the wing gets lower, the ground effect becomes more pronounced. While in the ground effect, the wing will require a lower angle of attack to produce the same amount of lift. If the angle of attack and velocity remain constant, an increase in the lift coefficient will result, which accounts for the “floating” effect. Ground effect will also alter thrust versus velocity, in that reducing induced drag will require less thrust to maintain the same velocity.
  •  Low winged aircraft are more affected by ground effect than high wing aircraft. Due to the change in up-wash, down-wash, and wingtip vortices there may be errors in the airspeed system while in ground effect due to changes in the local pressure at the static source.
  • Another important issue regarding ground effect is that the makeup of the surface directly affects the intensity; this is to say that a concrete or other smooth hard surface will produce more effect than water or broken ground.
  • When a wing approaches the ground two phenomena are actually involved in the increase of the lift force and reduction of the drag. The ground effect is a common name for both effects, which is sometimes confusing. These two phenomena are sometimes referred to as span dominated and chord dominated ground effect. The former results mainly in the reduction of the induced drag (D) and the latter in the increase of the lift (L). The designations span dominated and chord dominated are related to the the fact that the main parameter in the span dominated ground effect (reduces drag) is the height-to-span ratio ,whereas in the chord dominated ground effect (increases lift)it is the height-to-chord ratio.
• an empirical formula has been derived for practical and preliminary calculations of the chord-dominate ground effect. It has the following form
•               CLground = CL(h/c)-0.11

where CLground represents the wing lift coefficient in the ground effect mode,

CL – lift coefficient in open air, h indicates height of the wing and c is the chord length.

GROUND EFFECT

. Figure  depicts a wing in ground effect.

The boundary creates an alteration of the flow field that is caused by the boundary not allowing the flow under the wing to expand as it would in free air. In terms of the total pressure of the flow, the additional lift is due to a rise in static pressure under the wing.

The total pressure of the flow field can be divided between the static pressure (surface pressure) and dynamic pressure (the pressure associated with velocity). As the total pressure remains constant throughout the flow field, the sum of the static and dynamic pressure must also remain constant.

As the flow is forced into the region between the wing and the boundary, the decrease in dynamic pressure is transformed into a rise in the static pressure. This rise in the Static pressure is often referred to as ‘ram pressure’. The resulting altered pressure distribution causes a net increase in the lift and a change to many of the other aerodynamic characteristics of the wing.

GROUND EFFECT’s Effect on lift ,drag, downwash

the boundary near the wing alters the flow field about the wing. The effect is

demonstrated in Figure

 

The change in flow field has the effect of reducing the downwash angle and therefore increasing the effective angle of incidence at a given geometric angle of attack. This causes a corresponding rotation of the resultant force vector and changes to the component of lift an drag forces. The effect is to increase the lift component and reduce the induced drag component, thus increasing the lift to drag ratio.

The increased lift to drag ratio provides a net gain in efficiency and the reduction in drag provides the benefit of a reduced thrust requirement in cruise flight.

EFEECT ON PITCHING MOMENT

In addition to creating lift and drag, the movement of a wing through the air creates a moment about the aerodynamic centre of the wing. This moment is known as the pitching moment and is the result of the pressure distribution on the wings surface. In a moving craft this pitching moment needs to be balanced in order to keep the craft stable. WE typically add another lifting surface to overcome pitching moment, either at the rear of the aircraft (tailplane) or at the front of the aircraft (canard).

Ground effect alters the pitching moment generated by a wing. The altered flow about the wing moves the aerodynamic centre of the wing and therefore the pitching moment generated by the wing. The effect is the result of the pressure distribution changes over the lower surface of the wing. The ram pressure in extreme ground effect causes a near uniform pressure distribution over the under surface of the wing, while not significantly altering the upper surface pressure distribution (see Figure ).

Wings generally create a nose down pitching moment in cruise flight. Ground effect causes an increase in this moment, resulting in a greater stabilizing force being required to balance the pitching moment. To remain stable, a craft in ground effect will generally require a larger tailplane or canard.

This larger surface creates greater drag and therefore reduces the efficiency of the craft as a whole. It also creates structural and weight penalties that reduce the efficiency of the craft.

An additional complication of pitching moment in ground effect is that the pitching moment changes with height above the boundary. In freestream flight, the aerodynamic centre is generally considered to be approximately one quarter of the chord back from the leading edge.

Flight in extreme ground effect may move the aerodynamic centre to the half chord position. This movement of the aerodynamic centre with the height of the wing above the boundary may cause considerable configuration design difficulties. In addition, the need to be able to control the craft over a large pitching moment range increases the drag, structural and weight penalties discussed earlier.

Considerable research has been Considerable research has been conducted into overcoming the variation of pitching moment with height. Many designers have claimed to overcome the effect by the use of unique wing sections and/or craft configurations. Different shaped wing sections should be able to limit this effect by altering the pressure distribution over the lower surface so the change from IGE(in ground effect) to OGE(out of ground effect) is not large. Such a section is the S-shaped section used on the Amphistar. However, these sections may be dramatically inefficient in OGE flight or incapable of operating OGE and this is a likely area for further research. Planform shapes differing from conventional aircraft may provide another method to reduce the change in pitching moment.

Effect of Height above the Ground

From the modelling viewpoint, three separate models have emerged, each modelling a certain zone above the boundary.

The first zone is the region in which the wing is operating between the boundary and a height of 20% of the chord of the wing. This region has a high level of constriction of the flow in the vertical direction and the flow becomes two dimensional with the vertical degree of freedom of the flow is restrained.

The second zone is the region between the height of one chord length of the wing to ten span lengths. In this region, the model is dominated by the span of the wing. Inviscid flow models are used in this region and show a marginal increase in the L/D to that of OGE flight.

For awing flying in the region between 20% of the chord and one chord height, a combination of the two models are required. Above ten span lengths, free flight models currently used in aerodynamic theory for aircraft design, are used.

for wings operating in ground effect the addition of end plates is more efficient because they increase the lift to drag ratio more than if they where used to increase the wing’s span. It is also noted the end plates are more effective on wings of low aspect ratios, which are more likely to be found on WIG aircraft.

Theoretical Benefits of Ground Effect

The theoretical efficiencies of airborne craft can be expressed in terms of their ability to carry a given payload over a given distance. This efficiency is directly related to the craft’s lift to drag ratio. WIG craft’s higher lift to drag ratio, provide them with the potential for greater efficiencies than aircraft.

The resulting increase in the lift to drag ratio of a WIG craft results in an increase in the craft efficiency. One measure of efficiency is to consider the distance a specific payload can be transported.

Airborne craft are governed by the Bréguet range equation, for which the representation for propeller driven craft is shown below:

Range= (ηp*L/ Cp*D)*ln(wi /wi+wf)

• ηp propeller efficiency

• Cp specific fuel consumption

• L/D lift to drag ratio

• Wi initial weight

• Wf fuel weight

• The drag of the craft and the most efficient speed for the operation of the propulsion system dictate the best cruise speed.
•  While the maximum level speed is determined by the drag of the craft and the maximum thrust generated by the propulsion system.
• A reduction in the drag of the craft will see a corresponding increase in the craft’s maximum speed and optimal cruise speed.

Height Stability

• WIG craft height stability can be explained by considering the effect on lift with changes in height. The stable case is achieved when a decrease in height results in an increase in lift and vice versa.
• Under these conditions the increased lift has the effect of restoring the craft to the original height. Thus, if the craft is disturbed in height the lift force will act to restore the craft to the original height.
• In the opposite case, the craft will be unstable in height if the lift force acts to amplify the change in height. In this case a decrease in height will result in a decrease in lift. The decrease in lift will result in the aircraft accelerating further towards the ground, a result          enforced by the variation of lift with height.

• In WIG craft, the lift coefficient is a function of both height and incidence. For WIG craft the response is defined by the position of the centre of gravity. Dependent on the position of the centre of gravity, a change in speed may result in a change in incidence or a change in height. Pure speed changes resulting in height changes; occur at one extreme of the centre of gravity envelope.

• At the other, speed changes will result in pure changes of incidence. Between these extremes, speed changes will result in a combination of both height and incidence.

• For low angles of attack: There is no “ground effect” lift bonus at low angles of attack, as and lift (as well as induced drag) is less per unit airspeed for low angles of attack.

deep stall landing

SOME POINTS ON DEEPSTALL LANDING OF UAV’S

BRIEF ABOUT UAV’S

A small unmanned aerial vehicle (UAV) for aerial photography and remote sensing has been developed. In general, the size of a small UAV can range from several tens of centimeters to a few meters, and the propulsion is mostly by an electric powered motor. Ground control software can run on a laptop computer, so the ground station can be set up practically everywhere. Generally, small UAVs are deployed in environments where no access to conventional runways is available, because the UAV can take off by hand launch or catapult, and land at a short distance.

WHAT IS DEEP STALL LANDING ?

Deepstall landing is one of the landing methods by which airplanes can decrease flight speed while maintaining a deep glide path angle. This is because the wing can generate large air drag in the deep stall state. Therefore, the method is suitable for small unmanned aerial vehicle to land within narrow areas surrounded by tall obstacles.

 

WHERE IT REQUIRED ?

At confined areas surrounded by trees or buildings. It is very difficult to land UAVs at these areas adopting normal glide landing methods, especially for untrained operators. Hence, some effective landing methods must be devised so that the small UAV system can develop into a user-friendly technology. In these environments, the horizontal speed needs to be decreased and the deep path angle has to be maintained to land safely.

To meet this requirement, several solutions have been considered, e.g., parachute landing and vertical take off and landing (VTOL) mechanism. The parachute landing is the most common method because it is easy to operate and the UAV can be retrieved at any time and anywhere. However, it requires a considerable skill to properly fold parachutes and store them in the aircraft body.

Furthermore, the parachute itself has large air drag, which makes it difficult to precisely land the UAV at the target point in strong wind conditions. On the other hand, adopting the VTOL method, we can handle UAVs in narrow spaces. However, payload and endurance are limited and high performance controller is necessary as the VTOL system is complicated and originally unstable. Another way which doesn’t suffer from these disadvantages is “deepstall landing”,which is discussed in this post.

Deep stall is the state in which the angle of attack is much higher than the stall angle.

Even though the flow is unstable around the stall angle, it is stable in post-stall. Therefore, the wing can generate constant lift and drag forces when the angle of attack exceeds the stall angle by a certain value. Fig.1 and Fig.2 show the wind tunnel data of a NACA0012 wing .According to Fig.1, the lift coefficient C_l and drag coefficient  C_d are larger than at cruise.And Fig.2 means that the value of the lift to drag ratio L/D is smaller in the post stall state than in normal flight.

`            FIG-1=   C_L , C_D DATA FOR NACA 0012

FIG-2 = L/D RATIO OF NACA 0012

 

DEEPSTALL LANDING MECHAHISM

The only mechanism to get the plane into deep stall is quickly tilt of the horizontal tail plane (HTP). When the HTP tilts up suddenly at cruise, the plane pitches up largely and the angle of attack gets much higher than the stall angle. Then the plane trims at the deep stall state because the lift of the HTP cancels the wing and body moments. Furthermore, lift of the HTP acts as a restoring force because the HTP is not in stall. Consequently, the plane can keep this stable trim state. In this trimmed deep stall state,  C_L and C_D are so large that flight speed is decreased effectively. Besides, the L/D is so small that the plane can keep the deep glide path angle.

ADVANTAGES OF DEEPSTALL

The deepstall landing method can be implemented with a simple mechanism of HTP and it has a good stability. Therefore, untrained operators can easily handle it. This is an important point because small UAV system is becoming a platform for aerial photography and remote sensing and in many cases non-skilled operators handle the system. Furthermore, the path angle and speed can be set arbitrarily, controlling the HTP to tilt up by an angle . Hence, the landing point can be adjusted in strong wind conditions.

On the other hand, because touch down speed is higher than for the conventional methods, its application may be limited to small UAV only.

phase of DEEPSTALL landing were ESTIMATED . The TRIM STATE of an aircraft can be expressed by the

following three equations:

         

      FIG —–3

Variables are defined in Fig.3. Constant number g is a gravitation constant. Variables with suffix h are the value for the HTP. Lift and drag forces and moment can be modeled as functions of the angle of attack referring to NACA0012 wind tunnel data which includes after stall data. The moment balance equation (3) is expressed as below applying HTP volume  V_h  as variable.

 

Therefore, the trim state can be determined once  V_h and d are fixed.

Center of Gravity Position and its effect on stability

 

Center  of  Gravity  Position and its effect on stability

An aircraft’s horizontal tail size and position, and the CG position are the dominant factors controlling the aircraft’s pitch stability, which is the tendency to automatically maintain an angle of attack and airspeed.

The basic effects of moving the CG position are:

Decrease xcg/c    (move CG fwd.):  increased stability; more resistance to α and V changes.

Increase  xcg/c   (move CG back):  decreased stability; less resistance to α and V changes.

There is one particular CG position which gives neutral stability, which is called the Neutral Point (NP). This is shown as xnp in Figure (above). The degree of pitch stability or instability is traditionally specified by the Stability Margin.

S.M= (x_np – x_cg)/c

 

Figure 1 illustrates the natural behaviors of an airplane after a pitch disturbance, for different values of S.M. The unstable behavior occurs when S.M. is negative, i.e. when the CG is behind the NP. Because pitch instability makes the aircraft very difficult or impossible to control, the NP position is considered to be a practical aft  CG  limit.

Figure 1: Natural aircraft responses to a pitch disturbance, for different amounts of pitch stability.

Making the S.M. strongly positive by moving the CG far forward will give plenty of pitch stability and a strong resistance to pitch upsets, but it also has undesirable side effects. One large drawback    of a large S.M. is that it causes large (and annoying) pitch trim changes with changing airspeed. Figure 2 shows the flight paths of airplanes with different nonnegative S.M., immediately after an airspeed increase caused by a power increase. The straight-ahead acceleration of the weakly stable or neutral airplane is more desirable for the pilot.

        Figure 2: Pitch-up behavior from an airspeed increase, for large and small Static Margin.

More specifically, the strongly stable airplane shown in Figure 3 requires a relatively large

elevator angle change commanded by the pilot to restore it to level flight. Figure 4 compares the situation for the strongly and weakly stable airplanes. In effect, a large positive S.M. degrades the pitch trim authority of the elevator, since large trim deflections are needed to

maintain level flight in response to airspeed changes.

Figure 3: Elevator trim adjustment with changing airspeed, for large and small Static Margin

This situation illustrates the benefit of reducing the S.M., by moving the CG closer to the NP. However, if the CG is moved behind the NP, the airplane will now have a negative S.M., and be unstable in pitch to some degree, with the results illustrated in Figure 2. This makes it difficult or even impossible to fly. In general, the small positive S.M. suggested by expression given below  is the ideal situation.

.

SM= xnp-xcg/C  ==== 0.05…..0.15

ASPECT RATIO AND ITS EFFECT

ASPECT RATIO AND ITS EFFECT ON AIRCRAFT

As many early wing were in rectangular shape, the aspect ratio was initially defined as simply the span divided by the chord, for tapper wing it is defined as the span divided by the wing area.

ASPECT RATIO AND WING TIP VORTICES

                                                      When wing is generating lift it has to reduce pressure on the upper surface and an increased pressure on the lower surface , the would like like to escape from the bottom of the wing, moving to the top .

Air escaping around the wing tip lower the pressure difference between the upper and the lower surface , this reduces lift near the tip and also air flowing around the tip flows in a circular path, this reduces the effective angle of attack of the wing airfoil and this phenomenon is called as wing tip vortices .

Now , by keeping the area of the wing constant , tip of high aspect ratio wing is farther apart than low aspect ratio wing its mean the area or part of the wing affected by vortices is less in case oh high aspect ratio wing ,thus high aspect ratio wing does not experienced much loss of lift due wing tip effect as compared to same area low aspect ratio wing.

ASPECT RATIO AND STALING ANGLE

Another effect of changing aspect ratio is change in stalling angle.

Due to reduced effective angle of attack at the tip , a lower aspect ratio wing will stall at higher angle of attack than high aspect ratio wing . this is one reason why tail tends to be lower aspect ratio than wing , delaying the tail stall until well after the wing stall and assures adequate control .

Conversely a canard can be made to stall before the wing by making it a very high aspect ratio surface. This prevent the pilot from stalling the wing .

EFFECT ON STRUCTURE

A long wing has higher bending stress for a given load than a short wing and therefore requires higher structural design specification.

EFFECT ON MAUVERABILITY

A low aspect ratio wing will have a higher roll angular acceleration than one of high aspect ratio because high aspect ratio wing  has higher moment of inertia to overcome.

In steady roll the longer wing will gives a higher roll moment because of large moment arm of aileron,

low aspect ratio wing usually used on fighter jet , not only for the higher roll rate but especially for longer chord and thinner airfoil involved in supersonic flight

EFFECT ON INDUCED DRAG

only the one relation will tell everything in this section ,that is the induce drag coefficient (C_Di) :-

C_Di= C_L²/∏ e A.R

C_L= coefficient of lift

_e = span efficiency factor

A.R= aspect ratio

 

 

DIHEDRAL EFFECT

When an aircraft is disturbed from an upright position, it will sideslip toward the down-going Wing, increasing airflow along the length of the wing from tip to root. The dihedral angle increases angle of attack to this lateral flow, generating additional lift to restore the aircraft to a level attitude If the center of gravity is below the wing, the weight tends to restore the upright position.This is known as pendulum stability or the keel effect. If the CG is above the wing, the weight is destabilizing.

Sweepback of the wing, especially the leading edge, causes greater drag and greater lift on the wing panel that is rotated forward into the relative wind, increasing the roll still further – three to ten degrees of sweepback is approximately equivalent to one degree of dihedral for most model aircraft.

Dihedral bestows stability at the expense of lift. Only the vertical component of lift in level flight actually supports the airplane. It is proportional to the cosine of the dihedral angle. The horizontal component of wing lift, proportional to the square of the sine of the dihedral angle, is wasted. But the effect is small if the angle is small. A wing of 3 degrees dihedral, for example, wastes only 0.137% of its total lift (cosine 0° – cosine 3° = 0.00137). That may be insignificant in most models, but important in a competition sailplane or the long-term fuel costs of an airliner.

• Dihedral in a multi-engine airplane adds to undesirable roll when one engine quits.

•  Dihedral reduces stability in inverted flight and varies roll rate during the inverted part of a slow roll. Rolls become corkscrew shaped instead of axial.

• Dihedral makes an airplane more vulnerable to turbulence, Especially side gusts.

Where?

Rudder only-(aircraft without aileron)- radio controlled airplanes need lots of dihedral. Their only means of turning is by yawing with the rudder. The wing panel that swings forward presents a greater angle-of-attack to the relative wind, increasing lift. The greater lift banks and turns the airplane.

For efficiency, this method of turning is best implemented by adding extra dihedral to the wingtips, reducing the total dihedral.

The three or four-panel wing, typical of free-flight and rudder-only airplanes, is known as polyhedral. When the nose pulls up, the angle of attack of the outer panels increases at a slower rate than the inner panels. The inner panels stall first. Polyhedral wings do not require the negative twist known as washout.

The gull-wing variant is typically used to increase pendulum stability by raising the wing without the drag of cabane struts. The inverted gull-wing of the F4U Corsair was used to shorten the landing gear struts and lower the height of the airplane for below-deck storage.

Dihedral Sizing Criteria – Spiral Stability

 The dihedral angle of the wing, denoted by Υ in above Figure, provides some degree of natural spiral stability. A spirally-unstable aircraft tends to constantly increase its bank angle at some rate, and therefore requires constant attention by the pilot.Conversely, a spirally-stable aircraft will tend to roll upright with no control input from the pilot, and thus make the aircraft easier to fly. Figure 6 compares the two types of behavior.

Whether an aircraft is spirally stable or unstable can be determined via the spiral parameter B (named after its originator Blaine Rawdon, from Douglas Aircraft):

    B = L_F*Y/b*C_L            (Y in degree) 

B > 5	spirally stable
B=5	spirally neutral
B < 5	spirally unstable

 

The main parameter which is used to adjust B in the design phase is the dihedral angle Υ.

Spiral stability is not a hard requirement, and most aircraft are in fact spirally unstable. Level flight is then ensured either by the pilot, or by a wing-levelling autopilot, provided the instability is slow enough. RC aircraft which can fly stably hands-off must be spirally stable, although a small amount of instability (B=3…4, say) does not cause major difficulties for an experienced pilot.

Dihedral Sizing – Roll Control

On rudder/elevator aircraft, the rudder acts to generate a sideslip angle β, which then combines with dihedral to generate a roll moment and thus provide roll control.

A criterion for adequate roll authority is obtained by the product of Vf and B:

Vf *B =0.10 … 0.20

 

The 0.10 value will likely give marginal roll control, while 0.20 will give very effective control.

For practical reference the value of Vf*B of well famous RC uav SKYLARK by ELBIT SYSTEM Israel is 0.38